Our age-old fascination with the stars and planets has put us on a path of discovery that continues today. Investigate our history in space, from the early Soviet satellites to the Mars Pathfinder Mission and read about upcoming space projects. You can also visit the space photo gallery, which contains over 100 color photographs taken on American and Russian space missions.
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"Space Table of Contents",2,0,0,0
This section of \IWebster's World Encyclopedia\i contains in-depth text and pictorial information on space history as well as upcoming space exploration and projects.
It is divided into eight parts:
\JEarly Space Missions\j
\JMajor Moon Missions\j
\JNew and Upcoming Space Projects\j
\JSoviet/Russian Space Programs\j
\JSpace Photo Gallery\j
\JSpace Shuttle Information\j
\JYear in Space\j
\JMonth in Space\j
Click on the entry of your choice to move to that topic.
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"Early Space Missions",3,0,0,0
This part covers information on the following space missions:
\JClementine Project\j
\JGalileo Project Information\j
\JGiotto Mission\j
\JHubble Space Telescope (HST)\j
\JMagellan Program\j
\JMariner Program\j
\JMars Observer Project\j
\JPioneer Venus Program\j
\JSakigake and Suisei Projects\j
\JSkylab Program\j
\JUlysses Project Information\j
\JViking Mission to Mars\j
\JVoyager Project Information\j
Click on the entry of your choice to move to that part.
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"Major Moon Missions",4,0,0,0
This part covers information on the following space missions:
\JNASA Mercury Project\j
\JNASA Gemini Project\j
\JRanger (1964 - 1965)\j
\JSurveyor (1966 - 1968)\j
\JLunar Orbiter (1966 - 1967)\j
\JApollo 11, Twenty-Five Years On\j
\JApollo Program (1968 - 1972)\j
\JApollo Missions Contents\j
\JApollo 13 Mission\j
\JKey Documents from the Apollo Space Program\j
\JRemarks by the President at the 25th Anniversary of Apollo 11\j
Click on the entry of your choice to move to that part.
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"New and Upcoming Space Projects",5,0,0,0
\JInternational Space Station\j
\JMars Global Surveyor Updates\j
\JMars Pathfinder Status\j
\JMir 23 Status Report\j
\JShuttle Columbia Mission\j
\JSpace Calendar June97 - July98\j
\JSpace Launches 1997 (Worldwide)\j
\JMir Space Station\j
\JShuttle Missions (Upcoming)\j
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"Soviet/Russian Space Programs",6,0,0,0
\JSoviet/Russian Space Programs continued \j
\JSoviet/Russian Space Programs continued 2\j
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"Space Photo Gallery",7,0,0,0
This part provides over 100 color (plus some black and white) pictures of the different planets taken on various American and Russian space missions. It includes photographs of the following planets and planetary bodies:
\JAsteroid Photo Gallery\j
\JComets\j
\JEarth and Moon Photo Gallery\j
\JEarth Photo Gallery\j
\JJupiter Photo Gallery\j
\JMars Photo Gallery\j
\JMercury Photo Gallery\j
\JMoon Photo Gallery\j
\JNeptune Photo Gallery\j
\JPluto Photo Gallery\j
\JSaturn Photo Gallery\j
\JUranus Photo Gallery\j
\JVenus Photo Gallery\j
Click on the entry of your choice to move to that part.
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"Space Shuttle Information",8,0,0,0
\JShuttle CHALLENGER (10) 51-L (25)\j
\JSpace Shuttle, Living in a\j
\JShuttle Location\j
\JNASA Facilities\j
\JShuttle Flights To Date\j
\JSpace Shuttle Launch Team\j
\JSpace Shuttle Mission Summary (1972-1988)\j
\JSpace Shuttle System\j
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"Year in Space",9,0,0,0
\JYear in Space 1994\j
\JYear in Space 1995\j
\JYear in Space 1996\j
\JSpace News Update 1997\j
\JSpace News Update 1998\j
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"Month in Space",10,0,0,0
\JMonth in Space - July\j
\JMonth in Space - June\j
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"International Space Station",11,0,0,0
April 9, 1997
"International Space Station: \JEngineering\j the Future"
The International Space Station program promises a new era of space exploration and space-based scientific research, and will come to fruition through an unprecedented level of international cooperation. When complete, the Space Station will be permanently occupied by a crew of six, and it is anticipated that it will remain fully operational for ten years following its planned completion in June, 2002.
The International Space Station program is divided into three phases, each defined by the inception of new capabilities:
PHASE I (1994-1997) is currently underway, and utilizes existing resources, primarily the Space Shuttle and Russian Mir space station to build experience in the technical aspects of space station construction and occupation, and to usher in a new area of international cooperation in space.
Phase I began with the launch of Discovery for Shuttle mission STS-60, when Russian \Jcosmonaut\j Sergei Krikalev became the first Russian to fly aboard a U.S. \Jspacecraft\j. Phase I will be completed following the ninth Shuttle/Mir docking mission, currently envisioned to be STS-91, in May, 1998, which will carry the Alpha Magnetic Spectrometer (AMS) and other scientific experiments specially selected to maximize the utilization of the space station environment.
PHASE II (November, 1997 - February, 1999) will mark the beginning of the assembly phase of the International Space Station, and will see the first "production" experiments conducted aboard the Station. Phase II will commence upon launch, aboard a Russian Proton launch vehicle, of the Russian-built "FGB" module, a propulsion and attitude control module of a design proven years ago aboard Russian military flights.
In May, 1998, permanent occupation of the station will commence with a crew of three, and the Canadian Mobile Servicing System will be launched in late 1998. The end of Phase II will be marked by the delivery of the U.S. Laboratory Module aboard Shuttle flight STS-94. (Refer to Table)
PHASE III (February, 1999 - June, 2002) will see permanent occupation of the partially completed International Space Station by a crew of six. Construction will continue along with flights tailored to Station utilization. The first Phase III flight will be Shuttle mission STS-96, which will also be the first Space Station Utilization flight.
After the permanent docking of a second Soyuz craft to serve as an escape "lifeboat" and the installation of the U.S. Habitation Module aboard flight 19-A, STS-122, in June, 2002, the Station will be able to support a full-time six-person crew. Phase III will include the addition of the Japanese Experiment Module (JEM) in 2000 and the European Columbus Orbiting Facility (COF) in late 2002 or early 2003.
Phase III will draw to a close with the flight of Shuttle mission STS-121 and the addition of the Station's largest module, the U.S. Habitation Module. With assembly now complete, full operational use of the International Space Station will begin. (Refer to Table)
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"Mars Global Surveyor Updates",12,0,0,0
\JMars Surveyor Report (January)\j
\JMars Surveyor Report (February)\j
\JMars Surveyor Report (March)\j
\JMars Surveyor Report (April)\j
\JMars Surveyor Report (May)\j
\JMars Surveyor Report (June)\j
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"Mars Surveyor Report (January)",13,0,0,0
Friday, 10 January 1997
This week marked the transition from the inner-cruise to the outer- cruise phase of the mission. One of the first transition tasks occurred early Monday morning when the flight team sent a set of commands to change Surveyor's pointing orientation. The commands turned the \Jspacecraft\j from its previous orientation of +X axis pointed 60 degrees away from the Sun to a position where the +X axis is pointed directly at the Earth.
One of the benefits of this new pointing orientation is that Surveyor can now use its high-gain antenna to communicate with the Earth. This antenna is mounted on the \Jspacecraft\j's +X axis and its narrow-beam signal requires that the \Jspacecraft\j point directly at Earth.
Until now, Surveyor was utilizing its wide-beam, low-gain antenna for communications. The high-gain antenna broadcasts with greater power and will allow the \Jspacecraft\j to transmit data at higher data rates.
Before January, it was impossible to use the high-gain antenna because an Earth-pointed orientation would have placed Surveyor at an unfavorable angle with respect to the Sun. The switch from the low-gain to the high-gain antenna occurred early Thursday morning.
The flight team is continuing to diagnose the position discrepancy in Surveyor's -Y solar panel which is deployed, but 20.5 degrees from its proper position. \JEngineering\j data transmitted to Earth during the five solar array "wiggle tests" conducted in December supports the current model regarding the nature of the obstruction keeping the array out of position.
The model suggests that a damper shaft in the solar array's deployment mechanism broke shortly after launch, approximately 43 seconds after the start of the array's deployment. This damper is a device that was installed to minimize the mechanical shock of deployment by slowing the motion of the array during deployment.
The flight team theorizes that the broken shaft caused the damper arm to wedge into the hinge joint connecting the solar panel to the \Jspacecraft\j. Attitude-control telemetry recorded by the \Jspacecraft\j during solar array deployment corroborates this theory.
Plans are currently being developed for three more solar array "wiggle tests" during the week of January 20th. Data from these upcoming tests and the five previous tests in December will assist the flight team in determining the best method to attempt to free the damper arm from the hinge joint.
Today, the flight team transmitted the C4 sequence to Surveyor. C4 contains commands that will control Surveyor for the next five weeks. The first activities in C4 will start on January 13th and will involve using the Mars Orbiter Camera to image stars over four consecutive days. These star images will allow the camera team to refine the camera's focusing capability.
After a mission elapsed time of 64 days from launch, Surveyor is 14.79 million kilometers from the Earth and is moving in an orbit around the Sun with a velocity of 31.32 kilometers per second. This orbit will intercept Mars on September 12th, 1997. All systems on the \Jspacecraft\j continue to be in excellent condition.
Friday, 17 January 1997
On Monday of this week, Surveyor's flight team activated the Mars Orbiter Camera in preparation for four days of star imaging. Once per afternoon from Tuesday through Friday, the \Jspacecraft\j turned to point the camera at a cluster of stars called the Pleiades. Over the course of one hour on each imaging day, the camera observed stars within the cluster in order to perform focus checks.
Communications with the \Jspacecraft\j during star imaging was not possible because the star-pointed orientation resulted in pointing the high-gain antenna away from the Earth. Consequently, all of the data from the camera was stored on Surveyor's solid-state recorders.
This data was transmitted back to Earth approximately three hours after the conclusion of each day's imaging. The daily playback of camera data required 49 minutes. During that time, Surveyor transmitted 250 megabits of data at a downlink rate of 85,333 bits per second.
Next week, the onboard flight computer will activate heaters in the camera that will bake the epoxy structure of the camera to remove residual moisture. A set of four more star images will be taken after the bakeout period ends in late March. The star images taken this week will serve as a reference to assess the focusing capability of the camera after the bakeout.
Other activities this week included a two-hour radio-science \Jcalibration\j that occurred late in the evening on Wednesday. This test involved using the \Jspacecraft\j's ultra-stable \Joscillator\j to control the frequency or "tone" of Surveyor's radio transmissions to the Earth.
Normally, the \Jspacecraft\j listens to a signal transmitted from the Earth as a reference to set the tone of the signal transmitted to Earth. The \Joscillator\j functions as an electronic clock that can precisely control the tone of Surveyor's signal without listening to the Earth-based reference signal.
Future tests of the \Joscillator\j will occur approximately every other week until the \Jspacecraft\j reaches Mars. These tests are important because a stable radio signal as controlled by the \Joscillator\j will be critical toward the collection of scientific data at Mars.
After a mission elapsed time of 71 days from launch, Surveyor is 16.05 million kilometers from the Earth, 136.00 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 30.85 kilometers per second. This orbit will intercept Mars on September 12th, 1997. All systems on the \Jspacecraft\j continue to be in excellent condition.
Friday, 31 January 1997
Early Monday morning, flight controllers sent several commands to Surveyor that deactivated the Mars Orbiter Camera's 53-Watt bakeout heater. This heater was activated on Wednesday, January 22nd to remove residual moisture in the camera's \Jgraphite\j epoxy structure. If the bakeout had not been performed, the moisture in the camera's tube-like structure would have slowly leaked into space and caused its length to gradually change.
As a consequence, this tiny, slow-rate change in the structure's length would have resulted in a gradual shift in the focus of the camera during science operations. The goal of the bakeout was to remove all of the moisture at once in order to stabilize the focus of the camera.
Originally, the bakeout was scheduled to last for 60 days. This duration was subsequently reduced to 14 days last Wednesday when data from the camera suggested that the structure contained significantly less moisture than predicted.
Upon request from the camera team, the flight operations manager made the decision to terminate the bakeout after only six days. The concern is that baking the camera for longer than necessary would be detrimental to the camera's focusing capability.
In several weeks, the camera will image stars over a one-week period for the purpose of acquiring focus \Jcalibration\j images. These images will be compared to the star images taken before bakeout in order to assess the best focus settings for the camera.
Other activities this week included a two-hour radio-science \Jcalibration\j that occurred Thursday morning, just after midnight. This test involved using the \Jspacecraft\j's ultra-stable \Joscillator\j to control the frequency or "tone" of Surveyor's radio transmissions to the Earth.
Later on Thursday, flight controllers sent a command that activated a flange heater located near Surveyor's main rocket engine. The heater will gradually increase the pressure of the \Jnitrogen\j tetroxide inside the oxidizer tank.
As a consequence, the increase in oxidizer pressure will improve the efficiency of the propellant during the second trajectory correction maneuver. This maneuver is currently scheduled for March 20th.
After a mission elapsed time of 85 days from launch, Surveyor is 19.29 million kilometers from the Earth, 116.49 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 29.83 kilometers per second. This orbit will intercept Mars on September 12th, 1997. All systems on the \Jspacecraft\j continue to be in excellent condition.
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"Mars Surveyor Report (February)",14,0,0,0
Friday, 7 February 1997
Today, the flight team sent a command to Surveyor to activate the Mars Orbiter Camera. Over the weekend, the camera team will collect temperature data from the instrument in order to determine the best focus setting for a focus check test that will be performed on Tuesday, February 11th.
Earlier in the week, the flight team completed \Jcalibration\j activities on the gyroscopes in the inertial measurement unit. These gyroscopes are devices that provide critical data to the flight computers regarding Surveyor's pointing orientation in space. Each one of the three gyroscopes on the \Jspacecraft\j has a primary and backup data channel.
Over the course of a several day period, the \Jspacecraft\j team examined data from the backup \Jgyroscope\j channels in order to understand the slight variations between the in-flight performance and the performance as specified by the manufacturer.
The knowledge of these minor variations were incorporated into Surveyor's flight software. This activity was performed to improve the \Jspacecraft\j's ability to maintain a proper orientation in the event that the backup \Jgyroscope\j channels are used.
Throughout this past week, the Magnetometer science instrument has also been active. The data collected during the week will provide the Magnetometer team with an opportunity to conduct further calibrations on the instrument. In addition, the data will provide the team with an opportunity to study the solar wind.
This "wind" is a stream of protons and electrons that are constantly blown out from the Sun at a speed of 100,000 kilometers per second.
After a mission elapsed time of 92 days from launch, Surveyor is 21.51 million kilometers from the Earth, 107.49 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 29.31 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C4 command sequence, and all systems continue to be in excellent condition.
Friday, 14 February 1997
On Tuesday, the Surveyor \Jspacecraft\j rotated to a position that pointed the Mars Orbiter Camera at a cluster of stars called the Pleiades. Over the course of an hour, the camera imaged stars within the cluster. These images were used by the camera team to determine the focus of the narrow-angle camera following the bakeout period that ended two weeks ago.
During that five-day bakeout period, a 53-Watt heater was used to remove residual moisture from the camera's \Jgraphite\j epoxy structure. This moisture affects the camera's focus. Preliminary results from this week's activity indicates that additional bakeout will not be necessary.
Over the next two weeks, the camera will image the Pleiades on four separate opportunities to allow the camera team to make adjustments to the focus settings.
On Wednesday, the \Jspacecraft\j was commanded to spin in the opposite direction for a period of three hours. Under normal conditions during the journey to Mars, Surveyor's high-gain antenna is pointed at the Earth, and the \Jspacecraft\j slowly spins in the clockwise direction as seen from the Earth.
During the three hours, the \Jspacecraft\j spun in a counter- clockwise direction to allow the \Jspacecraft\j team to calibrate the gyroscopes. These devices provide information to Surveyor's flight computers regarding the \Jspacecraft\j's pointing orientation in space.
Today, the flight team transmitted the C5 sequence to Surveyor. C5 contains commands that will control the \Jspacecraft\j for the next four weeks. The first activities in C5 will start on Monday, February 17th.
After a mission elapsed time of 99 days from launch, Surveyor is 24.30 million kilometers from the Earth, 98.95 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 28.78 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C4 command sequence, and all systems continue to be in excellent condition.
Friday, 21 February 1997
Today, in an activity similar to the one that occurred last week, the Surveyor \Jspacecraft\j rotated to a position that pointed the Mars Orbiter Camera at a cluster of stars called the Pleiades. Over the course of an hour, the camera imaged stars within the cluster.
Images from today's opportunity, combined with three image sets that will be taken between February 24th and February 28th, will allow the camera team to determine settings to control the instrument's focus.
Other major events this week included a complete memory read-out of Surveyor's on-board flight computers on Monday. During this activity, the flight team commanded the \Jspacecraft\j's computers to transmit the contents of its memory banks back to Earth.
The read-out was performed to allow the flight team to verify the values of critical flight software parameters that control the \Jspacecraft\j. Because some of these parameters are periodically updated, the results of the memory read-out were entered into a tracking system that provides a historical record of the changes. Monday's activity was only the second time during the mission that the memory has been completely read out.
After a mission elapsed time of 106 days from launch, Surveyor is 27.71 million kilometers from the Earth, 90.93 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 28.25 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
Friday, 28 February 1997
On Monday, Wednesday, and Friday of the week that began on February 24th, the Surveyor \Jspacecraft\j rotated to a position that allowed the Mars Orbiter Camera to obtain images within a cluster of stars called the Pleiades. Images were gathered over the course of one hour on each day's opportunity. These images, combined with the images obtained on February 21st, will allow the camera team to determine settings to control the instrument's focus.
Late in the afternoon on Friday, the \Jspacecraft\j experienced a minor glitch with the star scanner. Normally, this device constantly scans a set of reference stars in deep space. These distant stars serve as fixed reference points that allow the \Jspacecraft\j to determine its proper pointing orientation relative to the Earth and Sun. This process is called attitude control and is not related to the camera's star imaging for focus determination purposes.
This glitch occurred during Friday's playback of Mars Orbiter Camera data from Surveyor's recorders. At that time, the star scanner began misidentifying stars. As a consequence, the flight team transmitted a command to the flight software to reset the portion of the attitude control software that controls the star scanner. After several hours, all conditions returned to normal.
Although the cause of the glitch has not yet been determined, the flight team suspects that the star scanner was fooled by sunlight reflecting off of dust particles in the vicinity of the \Jspacecraft\j. In order to further investigate this event, a playback of \Jspacecraft\j \Jengineering\j data recorded during the glitch will occur later this week.
After a mission elapsed time of 113 days from launch, Surveyor is 31.76 million kilometers from the Earth, 83.40 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 27.74 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
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"Mars Surveyor Report (March)",15,0,0,0
Friday, 7 March 1997
On Monday, the on-board command sequence controlling Surveyor executed a test called the "Solar Array Feather." During the several-hour test, the solar arrays were rotated back and forth several times in a similar fashion to the motion that a person makes when rotating the wrist joint.
This activity was performed for the benefit of the Magnetometer science team. The test simulated the rotation of the solar arrays that will occur as the arrays automatically track the Sun during Mars mapping operations. Because the Magnetometer sensors sit at the end of the solar arrays, the data collected from the test will allow the science team to determine the effect of the solar array rotation on the quality of their data.
On Tuesday, the flight team loaded new parameters to Surveyor's attitude control software. These parameters deal with the performance of the star scanner that controls the \Jspacecraft\j's ability to point at targets in space. With this parameter update, the \Jspacecraft\j will be able to point its science instruments at objects with better accuracy than previously possible.
Later on Tuesday, the Ka-band communications team accomplished a major milestone in their experiment. Over a several hour time period, an antenna at the Goldstone tracking station recorded data transmitted simultaneously from Surveyor's X-band and Ka-band transmitters.
Normally, the \Jspacecraft\j utilizes the 25-Watt, X-band transmitter for communicating with the Earth. The main difference between the two signals is that the 1-Watt, Ka-band transmitter operates at a frequency near 32 gigaHertz versus 8 gigaHertz for X-band.
An analysis of the experiment indicated that no disagreements existed between the X-band and Ka-band data for all 12 million data bits observed on Tuesday. This positive result marks the first verified data transmission by an interplanetary \Jspacecraft\j using a Ka-band signal.
The result affirms a long-held belief that the use of Ka-band signals can allow a \Jspacecraft\j to transmit information at faster data rates with transmitters that consume much less power.
After a mission elapsed time of 120 days from launch, Surveyor is 36.46 million kilometers from the Earth, 76.39 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 27.23 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
Friday, 14 March 1997
On Monday of this week, the flight team loaded new parameters to Surveyor's attitude control software. These parameters deal with the alignment of the Inertial Measurement Unit. This device contains three gyroscopes that provide the flight computers with critical information regarding the \Jspacecraft\j's pointing orientation in space.
The new parameters, combined with the new parameters for the star scanner that were loaded last week, will enable Surveyor to point its science instruments at objects with better accuracy than previously possible.
Today marked the first day since the launch of both Mars Pathfinder and Mars Global Surveyor that Pathfinder's distance to Mars was less than Surveyor's. However, because the two \Jspacecraft\j are on different types of flight paths to Mars, they did not physically fly past each other.
At the time of closest approach, Pathfinder and Surveyor were separated by 4.7 million kilometers. Pathfinder was launched after Surveyor, but will reach Mars first because it is traveling on a shorter, more direct flight path.
This week was a relatively quiet week as the flight team prepared for next week's trajectory correction maneuver. This engine firing will refine Surveyor's flight path to Mars and will take place on Thursday, March 20th at 10:00 a.m. PST.
After a mission elapsed time of 127 days from launch, Surveyor is 41.78 million kilometers from the Earth, 69.86 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 26.74 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C5 command sequence, and all systems continue to be in excellent condition.
Friday, 21 March 1997
On Wednesday, the flight team transmitted the C6 sequence to Surveyor. This sequence contains commands that will control the \Jspacecraft\j for the next four weeks. C6 became active on Thursday at 6:00 a.m. PST.
The first major event in C6 occurred at 10:00 a.m. PST on Thursday. At that time, the onboard flight computer commanded the \Jspacecraft\j's main rocket engine to fire for six seconds in order to make minor corrections to Surveyor's flight path.
During this trajectory correction maneuver, the main engine burned a propellant combination of hydrazine fuel and \Jnitrogen\j tetroxide oxidizer. In total, the \Jspacecraft\j expended approximately 1.4 kilograms of propellant.
Immediately before the six-second burn was performed, Surveyor ignited eight of its 12 attitude-control thrusters for 20 seconds. These tiny thruster rockets are normally used to stabilize the \Jspacecraft\j during main engine firings.
The initial, 20-second thruster firing settled the liquid in the \Jspacecraft\j's tanks to ensure a smooth flow of propellant to the more powerful main rocket engine that was used to perform the correction maneuver.
At this time, the navigation team is busy analyzing the accuracy of yesterday's trajectory correction maneuver. However, preliminary results from the accelerometer onboard the \Jspacecraft\j show that the engine firing provided a velocity change of 3.875 meters per second. This value was within 0.5% of the predicted change of 3.857 meters per second.
Yesterday's maneuver was the second in a series of four trajectory correction maneuvers that are designed to refine the \Jspacecraft\j's flight path to Mars. The first maneuver occurred shortly after launch last November. The third and fourth are currently scheduled for April 21st and August 25th, respectively.
After a mission elapsed time of 134 days from launch, Surveyor is 47.69 million kilometers from the Earth, 63.84 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 26.27 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 28 March 1997
No major activities occurred onboard the Mars Global Surveyor \Jspacecraft\j this week. Meanwhile, at the Jet Propulsion Laboratory in Pasadena, Surveyor's navigation team has completed their preliminary assessment of the trajectory correction maneuver that took place on March 20th. This short firing of the \Jspacecraft\j's main rocket engine resulted in a velocity change of 3.875 meters per second and refined Surveyor's flight path to Mars.
Initial analysis provided by the navigation team indicates that the \Jspacecraft\j performed the maneuver with an accuracy of greater than 99%. Consequently, the \Jspacecraft\j is now on a flight path that will come within 630 kilometers of the Martian surface at the point of closest approach on September 12th. Additional trajectory correction maneuvers scheduled for April 21st and August 25th will reduce this approach altitude to 500 and 380 kilometers, respectively.
After a mission elapsed time of 141 days from launch, Surveyor is 54.12 million kilometers from the Earth, 58.29 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 25.82 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
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"Mars Surveyor Report (April)",16,0,0,0
Friday, 4 April 1997
On Saturday, March 29th, the flight team performed a several-hour communications test to measure low-level interference between Surveyor's ultra-stable-oscillator-generated X-band signal and the Ka-band signal. Normally, the \Jspacecraft\j utilizes the 25-Watt, X-band transmitter for communicating with the Earth.
The main differences between the two are that the 1-Watt, Ka-band transmitter is experimental and operates at a frequency near 32 gigaHertz versus 8 gigaHertz for X-band.
During last Saturday's test, the \Jspacecraft\j simultaneously activated both the X- and Ka-band signal sources. The test was designed to determine the effect of the Ka-band signal on the purity of the X-band signal as generated by the ultra-stable \Joscillator\j.
Understanding the performance of the \Joscillator\j under potential interference conditions is important because this device functions as an electronic clock that precisely controls the tone of Surveyor's radio signal. Precision control of the signal's tone is important for gathering data regarding the Martian atmosphere.
On Monday, March 31st, the \Jspacecraft\j passed through a major milestone on its way to Mars. For the first time in the mission, Surveyor was closer to Mars than to Earth. This equidistant point was approximately 57 million kilometers between the two planets. The half-way point measured in terms of days to reach Mars will occur on April 10th.
After a mission elapsed time of 148 days from launch, Surveyor is 61.05 million kilometers from the Earth, 53.20 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 25.36 kilometers per second. This orbit will intercept Mars on September 12th, 1997. The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 11 April 1997
This week, the Mars Global Surveyor science team received an unexpected bonus from the Sun due to a solar flare eruption that took place on Monday. Eruptions of solar flares occur when disturbances deep within the Sun's interior cause streams of electrically charged atomic particles to be ejected from the solar surface. These charged particles move through the solar system at speeds in excess of 1,000,000 kilometers per hour.
In order to allow the science team to study this event, the flight team sent commands to Surveyor that enabled the \Jspacecraft\j to record solar flare data gathered from the Magnetometer science instrument. These commands activated the \Jspacecraft\j's data recorders late Wednesday afternoon, about half a day before the stream of charged particles from Monday's eruption reached Surveyor.
Although past occurrences of solar flares have both disrupted space communications and damaged \Jspacecraft\j, Monday's eruption was relatively mild in comparison. The Mars-bound Surveyor \Jspacecraft\j sustained no damage from the solar flare.
Late Thursday afternoon, the navigators on the project canceled the trajectory correction maneuver that was planned for later this month. This maneuver would have refined the flight path to Mars by slightly altering the \Jspacecraft\j's speed and velocity.
However, analysis showed that this month's maneuver involves a velocity change of only 40 millimeters per second (less than one-tenth of a mile per hour). The maneuver was canceled because with such a small velocity change, the errors in executing the maneuver are comparable to the size of the maneuver.
This canceled maneuver would have been the third of four planned maneuvers during the journey to Mars. The first two occurred in November 1996 and March 1997. The fourth trajectory correction maneuver will take place on August 25th, 1997.
Yesterday marked the halfway point in the journey to Mars with respect to time of flight. As of April 10th, Surveyor has completed 154 of the 308 days required to reach the red planet. The halfway point in terms of distance between the Earth and Mars occurred last week on Monday, March 31st. This difference in halfway dates arises from the fact that the positions of the two planets constantly change during the \Jspacecraft\j's journey to Mars.
After a mission elapsed time of 155 days from launch, Surveyor is 68.43 million kilometers from the Earth, 48.55 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 24.98 kilometers per second.
This orbit will intercept Mars 153 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 18 April 1997
No major mission activities occurred this week onboard the Mars Global Surveyor \Jspacecraft\j. Back at the Jet Propulsion Laboratory in Pasadena, \JCalifornia\j, the project management has made a decision not to attempt any more efforts to free debris that is currently keeping the -Y-side solar array slightly out of position. This solar panel is currently deployed and fully functional, but is 20.5 degrees from its proper position.
The flight team believes that the position discrepancy was caused when a damper shaft in the array's deployment mechanism broke shortly after launch. This damper is a device that was installed to minimize the mechanical shock of deployment by slowing the motion of the array during deployment. The flight team theorizes that the broken shaft caused the damper arm to wedge into the hinge joint connecting the solar panel to the \Jspacecraft\j.
An important aspect of this position discrepancy is that the solar panels will be used at Mars not only to produce electrical power, but also to help the \Jspacecraft\j attain its final mapping orbit. Over the course of a four-month period following Mars orbit insertion, Surveyor will be dipped into the upper Martian atmosphere on every orbit.
During these atmospheric passes, air resistance generated by the solar panels will slow the \Jspacecraft\j and gradually lower its orbit. Surveyor will use this "aerobraking" technique to lower the high point of its orbit from an initial 56,000 kilometer altitude to just under 400 kilometers.
For the last few months, the flight team has been considering several options to free the debris and allow the panel to latch and lock into its proper position. One idea involved a short firing of Surveyor's main rocket engine to provide a small force to dislodge the damper arm.
However, such efforts will not be necessary because an extensive analysis has indicated that aerobraking with the -Y solar panel slightly out of position is feasible with a few minor modifications to the original plan.
One of the minor changes involves rotating the panel into a position where the front side will face into the air flow instead of the back side. This orientation will keep the unlatched panel from folding up on itself when it encounters the air flow during aerobraking.
Because the front side contains the silicon cells that produce electricity, it is more fragile than the back side and cannot tolerate as much heating from the air flow. As a result, the flight plan will be modified so that Surveyor aerobrakes at a slightly slow pace than previously planned.
After a mission elapsed time of 162 days from launch, Surveyor is 76.20 million kilometers from the Earth, 44.32 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 24.59 kilometers per second.
This orbit will intercept Mars 146 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C6 command sequence, and all systems continue to be in excellent condition.
Friday, 25 April 1997
Last Friday afternoon, the flight team transmitted the C7 sequence to Surveyor. This sequence became active at 7:00 a.m. PDT on Monday, April 21st and contains commands that will control the \Jspacecraft\j for the next 28 days.
Late in the evening on Monday, Surveyor transmitted 1.5 gigabytes of recorded data back to Earth. This data was collected by the Magnetometer science instrument two weeks ago during the solar flare eruption. The playback of the data took five hours to complete and represents nearly 52 hours of recorded science. On Tuesday, the \Jspacecraft\j repeated the five-hour data transmission for redundancy purposes.
No other major activities occurred this week. After a mission elapsed time of 169 days from launch, Surveyor is 84.32 million kilometers from the Earth, 40.49 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 24.23 kilometers per second.
This orbit will intercept Mars 139 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C7 command sequence, and all systems continue to be in excellent condition.
#
"Mars Surveyor Report (May)",17,0,0,0
Friday, 2 May 1997
No major activities took place this week. For the past three weeks, few activities have occurred because the Surveyor \Jspacecraft\j has been configured in a quiet state for a search campaign to detect gravity waves. According to theoretical physics, these waves are gravitational disturbances emitted by all objects in the universe.
However, because gravity is a relatively weak force, detection of these waves is almost impossible unless they are generated by massive objects such as black holes and matter at the center of the Milky Way Galaxy.
To date, nobody has ever detected a gravity wave. If Surveyor encountered these waves, the \Jspacecraft\j would experience an extremely small jolt. This tiny bumping motion would cause a tiny shift in the frequency of the \Jspacecraft\j's radio signal transmitted to Earth. Analysis of the data generated by this experiment will take six months or more.
After a mission elapsed time of 176 days from launch, Surveyor is 92.74 million kilometers from the Earth, 37.03 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 23.89 kilometers per second.
This orbit will intercept Mars 132 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). The \Jspacecraft\j is currently executing the C7 command sequence, and all systems continue to be in excellent condition.
Friday, 9 May 1997
At 4:30 a.m. PDT on Thursday, the flight software onboard Mars Global Surveyor commanded the \Jspacecraft\j into safe mode. Entry into this operational mode placed the \Jspacecraft\j in a safe power, thermal, and communications configuration. This precautionary measure is taken if the \Jspacecraft\j detects an unexpected event in one or more of its subsystems.
The chain of events that resulted in safe mode began Wednesday night. At that time, the flight team was finishing the second of two calibrations of Surveyor's gyroscopes. These calibrations involved commanding the \Jspacecraft\j to rotate in various directions in order to ascertain the performance of the gyroscopes.
Surveyor had just completed the \Jcalibration\j that involved a +Z-axis rotation when the flight software commanded the \Jspacecraft\j into contingency mode. This mode is similar to safe mode, but involves fewer precautionary measures taken to safe the \Jspacecraft\j.
Entry into contingency mode was triggered when the direction to the Sun as measured by Surveyor's Sun sensors disagreed with the predicted direction to the Sun as calculated by the onboard flight software. This discrepancy in Sun position was approximately 5 degrees.
Entry into safe mode occurred about five hours later when a flight software task timed out and failed to report back Surveyor's central processor. At this time, the flight team is identifying the software task that timed out.
The entry into contingency and safe mode resulted in the flight software terminating the execution of the current command sequence, powering off the science payload and non-essential components, and turning the \Jspacecraft\j toward the Sun to guarantee adequate power. Analysis of telemetry transmitted from Surveyor over the last 24 hours indicates that all systems are healthy.
After the exact cause of safe- mode entry is identified and resolved, the flight team will command the \Jspacecraft\j back into its normal operational mode. This process will consume at least the next few days.
Late Thursday night, the flight team transmitted a series of commands to Surveyor for thermal maintenance purposes. One set of commands shut off the secondary set of heaters to avoid overheating the \Jspacecraft\j's 12 attitude-control thruster rockets.
The other set of commands changed Surveyor's pointing orientation from high-gain antenna pointed directly toward the Sun to antenna pointed 10 degrees away from the Sun. This orientation change allowed for more sunlight to maintain warm temperatures on the science instruments.
After a mission elapsed time of 183 days from launch, Surveyor is 101.43 million kilometers from the Earth, 33.90 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 23.58 kilometers per second. This orbit will intercept Mars 125 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th).
Although the \Jspacecraft\j is currently operating in safe mode, all systems are functioning properly, there are no \Jspacecraft\j hardware problems, and there is no threat to the mission.
Friday, 16 May 1997
This week, flight team members concentrated their efforts on determining what event caused the Mars Global Surveyor \Jspacecraft\j to enter safe mode early in the morning on May 8th. Since then, Surveyor has been operating in a configuration that ensures that the \Jspacecraft\j has adequate power, thermal, and communications margins.
Flight software on the \Jspacecraft\j automatically commands entry into this safe mode if it detects an unexpected event in one or more of Surveyor's subsystems.
One of the major diagnostic activities involved commanding the \Jspacecraft\j to transmit portions of its computer memory back to Earth for analysis. An examination of a region of memory called the Audit Queue revealed that entry into safe mode occurred when a flight software task timed out and failed to report back to Surveyor's central processor.
Each software task executed by Surveyor's computer is allocated a certain amount of time to complete. Timeouts occur when a task fails to complete in the allocated time. Members of the flight team at the Lockheed Martin facility in \JDenver\j traced the source of this timeout to an infinite loop that occurred in flight software. A timeout resulted because infinite loops are impossible to complete.
The infinite loop resulted from the corruption of an area of computer memory called the Active Script Table. This table contains a list of programs executed by Surveyor's central processor, and corresponding links to the locations in computer memory where those programs are stored.
A software task that was executing prior to safe- mode entry caused the infinite loop when it incorrectly updated one of the entries in the table by linking that entry back to itself.
Over the last few days, engineers on the flight team reproduced the safe-mode entry conditions in the \Jspacecraft\j simulator. Subsequent analysis indicates that the action that created the infinite loop is uncommon, but predictable. Consequently, the Flight Operations Manager has decided to allow the flight team to begin procedures that will return the \Jspacecraft\j back to its normal operating state.
Commands to perform this recovery will be sent starting in the afternoon on Monday, May 19th. Once recovery is complete, the flight team will transmit modifications to Surveyor's flight software that will prevent this infinite loop condition from occurring again. Normal operations should be restored by mid-week.
In other news not related to safe-mode operations, the flight computer powered down \Jgyroscope\j #2 on Tuesday, May 13th. This power down occurred automatically when the electrical current used by the \Jgyroscope\j exceeded a preset limit.
Gyro #2's functions were automatically assumed by Gyro #1 and #3, the transition was smooth, and there is no performance degradation with respect to Surveyor's ability to point at targets in space. The powered-down \Jgyroscope\j will be reactivated after normal operations recommence.
After a mission elapsed time of 190 days from launch, Surveyor is 110.33 million kilometers from the Earth, 31.07 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 23.29 kilometers per second.
This orbit will intercept Mars 118 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). Although the \Jspacecraft\j is currently operating in safe mode, all systems are functioning properly, and no \Jspacecraft\j hardware problems exist that pose a threat to the mission.
Tuesday, 27 May 1997
Shortly after 9:00 p.m. PDT last Saturday, operators staffing the Goldstone antenna complex in the Mojave desert announced that they had locked up on a signal transmitted from Surveyor at a data rate of 2,000 bits per second.
This milestone marked the transition out of safe-mode and back to normal operating conditions. Since early in the month, Surveyor's safe-mode orientation had limited the maximum data transmission rate to 250 bits per second or less.
The \Jspacecraft\j automatically entered safe-mode on the morning on Thursday, May 7th when the onboard computer encountered an infinite loop in flight software. Entry into safe mode placed Surveyor in a configuration that guaranteed adequate power, thermal, and communications margins. This mode is intended to be a benign operating state favorable for diagnostic and recovery activities if an unexpected event occurs in one or more of the \Jspacecraft\j's systems.
Recovery operations involved a multi-step process that began on Friday. First, the flight team sent a series of instructions to Surveyor's backup flight computer. These instructions initialized the backup computer to begin using its normal flight software rather than the limited software set utilized in safe mode. Then, the flight team commanded the backup computer to control the \Jspacecraft\j while performing the same software initialization procedure on the Surveyor's primary computer.
The next step required reestablishing the \Jspacecraft\j's ability to point at targets in space. In safe mode, the flight computer assumes that its ability to find and point at targets other than the Sun has been compromised.
Restoration of pointing capability involved commanding the \Jspacecraft\j to rotate in a cone-shaped pattern around the Sun for several hours. This action allowed Surveyor's star scanner to lock-up on distant guide stars in space. The \Jspacecraft\j determines its orientation in space by using these stars as reference points.
Pointing capability was restored early Saturday evening. At that time, the flight team commanded Surveyor to rotate from its safe-mode, Sun-pointed orientation to an Earth-pointed orientation.
Aiming the \Jspacecraft\j's antenna directly at the Earth enabled Surveyor to begin transmitting data using any one of its standard rates of 2,000 bits per second or faster. Early next week, the flight team will transmit modifications to Surveyor's flight software to prevent the infinite-loop condition from occurring again.
Surveyor would have been stable in safing for an indefinite period of time even if no corrective action had been taken. However, the flight team worked on restoring standard operations as quickly as possible because normal command sequences, such as those that control science \Jcalibration\j activities, are prohibited from executing in safe mode.
After a mission elapsed time of 201 days from launch, Surveyor is 124.64 million kilometers from the Earth, 27.15 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 22.89 kilometers per second. This orbit will intercept Mars 107 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). All systems continue to be in excellent condition.
#
"Mars Surveyor Report (June)",18,0,0,0
Friday, 6 June 1997
Two weeks after recovery from safe mode and the restoration of standard operations, the Mars Global Surveyor \Jspacecraft\j continues to perform excellently as it cruises toward an encounter with the red planet later this summer. Currently, Surveyor is operating in a quiet state with no major activity sequences programmed in the onboard computer. The flight team will transmit the next major sequence load toward the end of the month.
On Tuesday of this week, the flight team sent a few commands to Surveyor that activated \Jgyroscope\j #2 for a period of one hour. Several weeks ago, this gyro was automatically powered down when its usage of electrical current exceeded a preset limit.
Although gyros help the \Jspacecraft\j keep track of its pointing orientation in space, there was no loss of control because Surveyor's #1 and #3 gyros seamlessly assumed the function of the powered down unit.
During the one hour of operation, the amount of electrical current used by gyro #2 was well below the level that would have resulted in an automatic power down. Although the gyro is functional, the project management has decided to leave it powered off. Flight software code is being developed that will autonomously activate gyro #2 in the unlikely event that an anomalous condition precludes the usage of either the #1 or #3 gyro. This new software will be transmitted to Surveyor in a few weeks.
The only other notable activity this week occurred late Thursday. That evening, the flight team transmitted a short series of commands to Surveyor that modified the onboard software. These minor changes will ensure that the infinite-loop condition that resulted in safe mode entry will never happen again.
After a mission elapsed time of 211 days from launch, Surveyor is 137.88 million kilometers from the Earth, 24.04 million kilometers from Mars, and is moving in an orbit around the Sun with a velocity of 22.57 kilometers per second. This orbit will intercept Mars 97 days from now, slightly after 6:00 p.m. PDT on September 11th (01:00 UTC, September 12th). All systems continue to be in excellent condition.
Due to the dynamic nature of space flight operations, please keep in mind that dates of events can change at a moment's notice. This calendar was last updated on 18 May 1997.
Short-Term Event Schedule
19-May-97
Begin transition out of safe mode and back to normal operations
30-May-97
Uplink command sequence C8 to \Jspacecraft\j
9-May-97
Uplink software patch for Payload Data Subsystem
High-Level Master Schedule
Launch
7-Nov-96
Lift-off occurred at 12:00:50 EST
TCM121-
Nov-96
Optimized maneuver to correct for launch injection errors (1st of 2 burn sequence)
TCM220-
Mar-97
Flight path correction (2nd of 2 burn sequence)
TCM321-
Apr-97
Maneuver to correct for execution errors from TCM2
TCM425-
Aug-97
Maneuver for final adjustment for orbit insertion aim point
MOI12-
Sep-97
Mars orbit insertion burn will last for about 20 minutes, closest approach will occur at about 1:26 a.m.
UTCOrbit Insertion Phase
12-Sep-97 to 14-Mar-98
Lasts for 5 months to reach mapping orbit using aerobraking and propulsive maneuvers
Mapping Phase
15-Mar-98 to 31-Jan-00
Mapping operations will last 1 Mars year, about 687 Earth days in duration
Relay Phase
1-Feb-00 to 1-Jan-03
Communications support for future Mars missions
#
"Mars Pathfinder Status",19,0,0,0
\JPathfinder Status (August96)\j
\JPathfinder Status (September96)\j
\JPathfinder Status (October96)\j
\JPathfinder Status (December96)\j
\JPathfinder Status (January97)\j
\JPathfinder Status (February97)\j
\JPathfinder Status (March97)\j
\JPathfinder Status (April97)\j
\JPathfinder Status (May97)\j
\JPathfinder Status (June97)\j
\JPathfinder: Entry, Descent and Landing\j
#
"Pathfinder Status (August96)",20,0,0,0
Week of August 19, 1996
Once the \Jspacecraft\j arrived and unpacked at KSC, one of the first things we did was re-run several "EDL runs" in several different conditions on the \Jspacecraft\j (EDL stands for Entry Descent and Landing). We wanted to re-verify some minor changes in the flight software that automatically controls the detailed series of events in this critical mission phase. In these tests, the software starts out in the late cruise mode of operation.
Using support equipment that can "stimulate" accelerometers to simulate the effects of the deceleration of entry and landing, as well as other equipment that can simulate the data that comes from the radar \Jaltimeter\j, the entire EDL process can be artificially created without having anything really happen (other than petal and airbag retraction actuators moving).
These tests were quite productive. In several cases we even "pulled the plug" on the computer and let it re-boot, with great success. This was the next-to-last time that the EDL software will be run on the real lander, the next time will be on the Fourth of July, 1997!
#
"Pathfinder Status (September96)",21,0,0,0
Week of September 2, 1996
EDL testing done, the lander was dissembled so that the interior \Jelectronics\j could be accessed. We had two fuses, some relays, and a waveguide transfer switch that needed to be replaced. Also we installed a fresh, fully charged, flight battery. This battery is the best one that we have used to date and will now have to survive to do its job over the next year. Once the \Jelectronics\j were updated, we performed a full functional test to confirm that the fixes worked and that we didn't disturb anything in the process.
Week of September 9, 1996
Early this week, for the final time, we reinstalled the ISA (Integrated Structure Assembly) - that's the white thermal and structural box that surrounds the lander \Jelectronics\j and has the red "JPL" letters on it. We then calibrated the stop positions on the HGA (High Gain Antenna) - that's the lollipop-shaped articulated antenna that sits nest to the camera on the ISA.
We used special theodolites (similar to those used by construction surveyors) to verify that the HGA mechanically points in the direction we want it to point with respect to the lander's base petal. We then reinstalled and checked out the pyro switching \Jelectronics\j and installed the lander thermal batteries.
We use "thermal" batteries to provide power (current) to ignite explosive initiators in the EDL pyrotechnic devices (things like the parachute mortar, separation nuts, cable cutters and rocket ignitors). The batteries are called "thermal" because they get their electrical energy from self-generated chemical heat.
Similar to the pyrotechnic initiators to which they provide, these batteries themselves need to be "lit" seconds before they are used on the \Jspacecraft\j during EDL. Once "lit", these batteries will operate for only a few minutes - plenty of time to do their jobs.
Once the ISA was installed, we performed some radio communication tests between the rover and the lander. We had been uncertain whether or not we needed to launch with an RF (radio frequency) attenuator in series between the lander's rover antenna and the lander's RFD modem used to talk with the rover.
This attenuator was thought some time ago to be needed to allow communication with the lander at close distances. These tests and some others coming up have nearly convinced us that we can live without it. We think that it would be good if we did not use it because the attenuator might reduce the communication range if we ever decided to drive the rover a long way away from the lander in its "extended" mission. Either way, the primary mission is unaffected.
We successfully performed other radio tests as well. Until this week, we had not yet tried to uplink the large software patch files using the real X-band radio and a ground station. The patch files are used in the unlikely event we have to reload large portions of the flight software into the EEPROM memory during the mission. Using the MIL-71 ground station at the Kennedy Space Center, we found that the process works fine.
We also took a few last verification images from each eye of the IMP camera on the lander. This is the last time the IMP camera will be taking interesting pictures until we land on Mars.
There has also been much work on the cruise stage. The HRS (Heat Rejection System) freon pumps have been installed and checked out. The HRS is the system used to flow freon inside the lander and around the perimeter of the cruise stage to keep the lander \Jelectronics\j cool.
We need to keep the battery, the digital \Jelectronics\j, the rover, and the big X-band radio transmitter we call the SSPA (Solid State Power Amplifier) cool during the especially warm early part of the "cruise" phase of the mission as we leave Earth.
We had to replace the pumps that had been installed during this summer's thermal tests because we think that it may have been damaged during one of our electrical tests. Because it is so hard to take apart, we can't tell for sure that it is broken. So just in case it was, we decided to replace it with the flight spare unit.
Later this week we will reattach the petals (the Sojourner Rover is already mounted on its "Y" petal). Next week we begin the long process of installing the flight airbag as well as the many pyrotechnic devices on the lander.
Week of September 16, 1996
As you can see from the live image, the mechanical team has completed installation of the three lander petals. Each of these petals are mounted on a hinge that allows them to open and close by more than 110 degrees of travel. The petals are moved under software command via three actuators mounted along the "hinge line" of each of the three petals.
One or more of these petals open up, after the airbags are retracted, about an hour and fifteen minutes after landing. They are the mechanism for automatically righting the tetrahedral-shaped lander so that no matter how the lander comes to rest inside the inflated airbag \Jcocoon\j, the lander will always end up "top side up" and not the reverse!
Once the petals were installed, the "rock membrane" was placed on each of the petals (except for the rover petal which will be done next week). This is a layer of aluminum sheet metal placed on the outside of the petals (but under the airbags).
The petals are about two inches thick, but for the most part are hollow. The inside of the petals have the lander's solar arrays attached on an aluminum substrate. This layer will provide protection from any tall sharp rocks that might be inclined to try to penetrate a side (or base) petal and damage the delicate solar arrays from below.
When the lander rolls onto its base petal from a side petal during the righting phase, the 270 kg lander (only 223 lbs on Mars) can roll quite hard. The airbags will help some, but aluminum will really do the trick.
The three solid rockets used to stop the lander just before landing are now installed. First they had to be prepared with heaters and thermal blanketing before mounting to the inside of the backshell. The backshell, the heatshield and the cruise stage are not far from the lander, but are out of sight of the camera.
Progress on the cruise stage is also being made. We decided some time ago to separate the fuel tank heater circuits from the propellant line heaters so that we can maintain the option of turning off the tank heaters during battery charging while also not allowing the fuel in the propellant lines to freeze.
Turning the tank heaters on for the few hours of battery charging allows us to gain a few extra precious watts when we need it most during "cruise". The necessary cabling changes on the cruise stage were done this past week and verified.
#
"Pathfinder Status (October96)",22,0,0,0
Week of October 8, 1996
A lot has been accomplished in the last two weeks. For those of you who have been watching the images from KSC regularly, you will have noticed that the airbags have been installed for nearly two weeks.
The airbag and gas generator installation operation is quite complex, but the talented folks (Skip Wilson) from the airbag manufacturer (ILC Dover Inc. of Frederica, DE) and the JPL airbag cognizant engineer (Tom Rivellini) have done this so many times before on other test landers, that they make it look easy.
Once the bags were installed, a detailed "walk through" of the whole lander was performed early last week by some the best \Jspacecraft\j mechanical engineers around. They looked for anything that might appear to be amiss.
For example, they pointed out that the exposed edge of the solar panel honeycomb substrates really ought to be taped closed rather than provide the opportunity for manufacturing particulates to escape and potentially contaminate sensitive surfaces such as the camera \Joptics\j. The edges were taped and bonded closed within two of hours of being mentioned by our very talented mechanical team.
One other item was reviewed and commented upon by the mechanical review team. We all noticed for the first time, that with the fully loaded mass of the lander and the airbags, the petal latches were no longer perfectly aligned when closed.
This meant that when the six petal latch separation bolts are torqued down (they get released after landing released via pyrotechnic separation nuts) there may be some rebound in the structure that would "race" the separation nuts at the moment of release. The engineers were concerned that there was adequate time for the bolts to exit the holes before the structure moved to the point that the bolts could not clear the holes and hang up.
Also we knew that the petal actuators and the mechanical petal stops used to prevent the lander petals from closing in on and making unwanted contact with the internal structures on the base petal, would also contribute some outward \Jtorque\j and angular acceleration at the moment of petal latch release.
So just to be on the safe side, with the help of Jim Baughman, the lander structural engineer and designer, we made the decision to slightly widen the separation bolt holes so that there will be plenty of time for the bolt to exit the hole before the structure moved to the point of impeding bolt motion (the bolts move out of the holes from the energy of the released tension in the bolt plus the bolt has a spring at the head of the bolt that moves it quite quickly out when the nut splits open).
This delayed the final petal closing a day or two and by last Saturday night we had the petals closed for the last time. Fortunately we still have plenty of time built into our schedule to get all of the work still to go completed before our first launch opportunity on Dec. 2.
This coming week, we will lift and weigh the lander and we will install it in to the backshell by raising it from below. Shortly thereafter we will install the heatshield for the last time. Later in the week we will take this "entry" assembly and place it upside down (heatshield up) onto a spin table in the high bay and we will balance the vehicle (much like an \Jautomobile\j tire) so that there is virtually no wobble when it spins at 2 rpm during Mars entry.
For week of October 17, 1996
Spacecraft Status at KSC:
As you can see from the live KSC image, quite a lot has been accomplished over the last two weeks. The petals have been closed for the final time, the lander bolted (using separation nuts) to the "backshell interface plate" inside the backshell. Most recently, the heatshield has been bolted to the aeroshell.
These accomplishments come as a great relief to the entire team. As we and other \Jspacecraft\j builders have discovered, the flight hardware is often more in danger of being damaged inadvertently by human hands than from anything the rigors of outer space could dish out!
For example we had to repair a broken wind sensor in the MET mast after a ground wrist strap brushed past it, damaging a very tiny element. During the mission, nothing will come near this sensor. Fortunately, it was fixed without much trouble. Now that the bulk of the system is neatly packaged inside the backshell, at least we can breathe a partial sigh of relief.
The next step in this delicate process will be to turn the entry vehicle upside down and spin balance it at 70 rpm using a spin table in the facility. This process will remove any wobble from the mass properties. (For the technically inclined, this process aligns the principle axis of \Jinertia\j with the axis of symmetry.) Once done later this week, we will flip the vehicle over again (a tricky process all by itself) and then mount the cruise stage. We will then do yet another spin balance in the launch configuration.
When we started this assembly sequence at KSC two months ago we had about 14 days of schedule "pad" in case anything slowed us down. Well of course, things did slow us down and we now have about 5 days of schedule margin left until we place the \Jspacecraft\j on top of the upper stage in November. However that isn't bad considering the complexity of this operation. And we had expected to use the margin up. We are all very happy with the progress so far.
Mission Operations Status at JPL:
For the past two weeks those of us on the team working at JPL have been focused on operations. We have a "testbed" at in Pasadena that has a complete duplicate set of all of the flight \Jelectronics\j that are on the \Jspacecraft\j (including a "sim" rover when needed).
In addition to testing flight software on this testbed, we also use it in an operational mode: essentially "flying" the testbed in a simulated mission to Mars! Early last week we "launched" the testbed using the same people, procedures and software that we will use on launch day. We even used the launch team at the Cape who have been doing the \Jelectronics\j \Jintegration\j testing.
Afterward we performed the instrument and rover health checks that we will perform in the weeks after launch. This week, we performed two TCMs (trajectory correction maneuvers - or "burns"). Later this week and this weekend we will "land" the testbed: actually running the "EDL" (entry descent and landing) flight software in the "testbed spacecraft". Saturday we will simulate the first day on Mars.
We are currently in the process of building up a "Mars Room" adjacent to the testbed near our operations area. This room is filled with sand and rocks and will eventually be the home of a test lander and the sim rover.
Although most of the \Jelectronics\j will reside in the testbed a few feet away (so we can more easily access it), the lander's sensors and actuators will reside on the lander in the room. This lander will include the petals, camera, high gain antenna, airbags and rover communication antenna.
The rover ramps will also be installed. In the coming months, we will use this room to simulate the first days and weeks on Mars. These simulations help us learn the finer points of operating this complex little lander and rover.
#
"Pathfinder Status (December96)",23,0,0,0
December 4, 1996
10:30 a.m. Pacific Standard Time
NASA's Mars Pathfinder is reported to be performing well on the first day of a seven-month journey to the red planet following a perfect launch today from Cape Canaveral, FL at 1:58 a.m. Eastern time.
"The \Jspacecraft\j team was ecstatic at seeing good \Jspacecraft\j data," said Brian Muirhead, Pathfinder Deputy Project Manager at NASA's Jet Propulsion Laboratory. "The command and data telecommunications subsystems are working perfectly, sending down data at 1,183 bits per second.
The temperature control and propulsion subsystems reported all temperatures and pressures are within expected ranges. All systems are healthy," he said. Pathfinder is traveling away from Earth at a speed of 3.9 kilometers (2.4 miles) per second.
The Delta II launch vehicle performed flawlessly, placing the \Jspacecraft\j on its trajectory to Mars well within acceptable limits. NASA's Deep Space Network acquired the \Jspacecraft\j telemetry signal on schedule, about five minutes after separation of the Delta's third stage.
When Pathfinder came out of Earth's shadow at one hour and 38 minutes after launch, the solar arrays took over powering the \Jspacecraft\j as planned. "Power from the array looks to be about 10% better than initially predicted," said Muirhead.
Pathfinder engineers continue to analyze data from the \Jspacecraft\j's sun sensor, an instrument that helps the \Jspacecraft\j determine its orientation with respect to the Sun. "The sensor's voltage output is below expected levels but it does appear to be giving good data," said Muirhead.
Navigation data and the sun sensor data agree and show the \Jspacecraft\j to be properly oriented, spinning at the expected 12 rpm and pointed 26 degrees off the Sun. Later today, Muirhead said the \Jspacecraft\j will be commanded to switch to a redundant sensor head to see if it is also performing at a low voltage. "Should the problem persist, we have a number of work around options and there is no risk to the continuation of the mission."
Carried inside the cone-shaped \Jspacecraft\j is Sojourner, the small robotic rover that will roll out to traverse the surface of Mars when the \Jspacecraft\j makes its Martian landing on July 4, 1997.
It will be the first \Jspacecraft\j to land on Mars since NASA's Viking mission soft-landed two \Jspacecraft\j there in 1976. Pathfinder is the second mission in NASA's Discovery program, which is designed to send low-cost \Jspacecraft\j with highly focused mission objectives to explore space.
December 6, 1996
12:00 p.m. Pacific Standard Time
The Mars Pathfinder \Jspacecraft\j continues to perform well in the early part of its cruise to Mars, which is about 209 million kilometers (130 million miles) away today.
Currently the \Jspacecraft\j is 750,000 kilometers (0.5 million miles) from Earth, or about two times the distance that the Moon is from Earth, traveling at a speed of 3.3 kilometers per second (7,400 miles per hour).
The \Jspacecraft\j is performing just as expected, with the exception of the sun sensor. The temperatures of the lander and its \Jelectronics\j are at their predicted levels for this phase of the mission. The cruise stage solar array, propulsion module and \Jelectronics\j are also at their predicted temperatures.
Two of the four segments of the solar array are currently in use, producing approximately 250 watts of power, about 10 percent more power than the original predicts. The battery is charged at 75 percent of its full capacity, and is showing a temperature of 9 \JCelsius\j (48 degrees Fahrenheit), which is approaching the desired steady state of 8 \JCelsius\j (46 degrees Fahrenheit).
The telecommunications system is performing well within its predicted range, indicating that it will be able to maintain higher data rates throughout the mission.
The JPL flight team is continuing its investigation of a lower than expected voltage reading on the sun sensor. However, since the sensor data are good, flight controllers have decided to implement a software update to compensate for this low voltage condition.
The software modification has already been coded and validated in the project's testbed and will be sent to the \Jspacecraft\j this weekend. The software modification will allow Pathfinder's on-board attitude control system to use the sun sensor data in its normal calculations of the \Jspacecraft\j's orientation. Once the attitude control calculations are verified, the planned spin-down maneuver to 2 rpm will be performed, probably early next week.
The \Jspacecraft\j is pointed approximately 55 degrees from Earth and 25 degrees off the Sun. Doppler and ranging data continue to look very good. Because the \Jspacecraft\j is not pointed directly at Earth, flight controllers are able to observe the motion of the antenna as Pathfinder spins about its axis and have confirmed a spin rate of 12.3 rpm.
The latest orbital data from tracking operations at all three Deep Space Network stations around the world indicate that the magnitude of the first trajectory correction maneuver, if performed as scheduled on Jan 4, 1997, would be 29.5 meters per second (96 feet per second).
Mars Pathfinder, the second in NASA's Discovery program of low-cost, highly focused spaceflight missions, is scheduled to land on the surface of Mars on July 4, 1997, and deploy a small rover, called Sojourner, to explore the Martian landscape.
December 10, 1996
12:00 p.m. Pacific Standard Time
The Mars Pathfinder \Jspacecraft\j continues to perform nearly flawlessly on its 203 million kilometer (126 million mile) flight path to Mars. Currently the \Jspacecraft\j is 1.8 million kilometers (1.1 million miles) from Earth, traveling at a speed of 3.2 kilometers per second (7,155 miles per hour). Temperatures and power utilization of the lander and cruise stage remain at predicted levels for this early phase of the mission.
The \Jspacecraft\j's sun sensors are the only issue being watched closely on an otherwise beautifully performing \Jspacecraft\j, the flight team reported. There are five sun sensor heads on board the \Jspacecraft\j, two pointed along the craft's spin axis and three that are equally spaced around the circular cruise stage that look out at about 105 degrees from the spin axis.
Of the five sensor heads, unit #4 on the spin axis is obscured or contaminated to the point of not being useful. Sensor #5, which is also on the spin axis, is providing good sun orientation data, but at a lower voltage than was expected. The other three sensor heads are working fine.
The flight team at JPL uploaded a software modification to the \Jspacecraft\j on Saturday, December 7, which allowed the on-board attitude control system to use the sun sensor data from sensor #5 in its normal calculations of the \Jspacecraft\j's orientation. The software patch was successful and the team was exuberant to see the \Jspacecraft\j's attitude control estimators operating properly.
The team then began to prepare for turning the \Jspacecraft\j more toward Earth to improve the telecommunications link. At the time, Pathfinder was about 58 degrees from the Earth, which is near the edge of the antenna's performance.
Since this was to be the first time flight controllers used the propulsion module, they planned a small turn of two degrees to verify that everything was working properly. Thirty minutes later, they planned to turn the \Jspacecraft\j an additional 20 degrees.
"The turn maneuvers were conducted successfully on Monday morning [December 12]," said Brian Muirhead, Pathfinder flight system manager. "The propulsion and attitude control systems worked properly and the \Jspacecraft\j's spin axis is currently pointed about 44 degrees from the Sun and 37 degrees from Earth. The downlink performance improved as expected and we continue to communicate with Pathfinder at 1,185 bits per second."
The flight team is planning its next maneuver to spin the \Jspacecraft\j down from 12.3 rpm to 2 rpm. The maneuver will be performed in the next few days, Muirhead said. Pathfinder's first trajectory correction maneuver remains on schedule, to take place on January 4, 1997.
December 18, 1996
12:00 p.m. Pacific Standard Time
Sojourner, a 10-kilogram (22-pound) rover tucked away on a petal of the Mars Pathfinder \Jspacecraft\j, got a 'wake up' call on Dec. 17 from flight controllers at NASA's Jet Propulsion Laboratory. After waking up, Sojourner conducted an internal health check and sent data back to the flight team that all was well.
The Pathfinder flight team was ecstatic with the rover data, which showed that all systems within the rover were operating normally. In addition, data from the rover's main science instrument -- the alpha proton x-ray spectrometer -- showed that it was operating properly.
"The rover woke up, did its internal health check, sent the lander its status data and went back to sleep, all as planned," said Art Thompson, rover operations team member. "All subsystems were verified as being in good health."
Pathfinder continues to perform very well on its 500 million-kilometer (310 million-mile) journey to Mars, the team reported. Currently the \Jspacecraft\j is 4 million kilometers (2.5 million miles) from Earth, traveling at a speed of 3.1 kilometers per second (7,000 miles per hour).
Its destination, Mars, is currently about 190 million kilometers (118 million miles) away. All temperatures and power utilization of the lander and cruise stage remain at their predicted levels for this phase of the mission.
The \Jspacecraft\j was spun down from 12.3 rpm to 2 rpm on Dec. 11. Flight controllers first instructed the \Jspacecraft\j to turn to a Sun angle of 50 degrees and an Earth angle of 32 degrees. This allowed them to use all four operating Sun sensors. The \Jspacecraft\j executed the commanded spin down to the normal cruise spin rate of 2 rpm in steps of 2 rpm at a time.
Once the normal spin rate was established, the team turned on the \Jspacecraft\j's star scanner on Dec. 12. Star scanner data allows the \Jspacecraft\j to establish full, three-axis knowledge of its orientation in space. This is the normal cruise attitude control mode and the one in which all trajectory correction maneuvers will be performed.
While Sun sensor #5 continues to work well after a software fix, the flight team continues to investigate the cause of the loss of Sun sensor head #4. The team expects to reach a likely conclusion on the cause of the problem within the next month or two.
Dave Gruel, Pathfinder flight director at JPL, conducted the Dec.16 health check of the lander science instruments, including the atmospheric sensor instrument and \Jmeteorology\j (ASI/MET) package and the imager.
Temperature, pressure and accelerometer readings from the atmospheric/meteorology instrument verified it was in normal working order. Power and dark current measurements received from the imager while it was imaging the darkness around it, confirmed that the instrument was working properly, Gruel said.
Richard Cook, Pathfinder mission operations manager at JPL, reported today that Pathfinder has been fully checked out for this phase of the mission and that all subsystems are "go" for a successful seven-month cruise to Mars.
The next major in-flight event will be Pathfinder's first trajectory correction maneuver, which is scheduled for Jan. 4, 1997.
31 December 1996
The \Jspacecraft\j is currently 7.2 million kilometers from Earth and traveling at 32.6 kilometers per second. All \Jspacecraft\j subsystems continue to operate as expected.
On 30 December 1996 we performed another successful ASI/MET science instrument health check. These are intended to monitor the performance of the pressure \Jtransducer\j that will measure Martian air pressure.
Last Friday, 27 December 1996, we successfully performed our first celestial mode attitude turn. In this turn, the spin axis was turned about 43 degrees, mostly out of the plane of the \Jecliptic\j. The \Jspacecraft\j used the sun sensors and the star scanner to precisely orient the \Jspacecraft\j's spin axis to the new direction, which is about 35 degrees off the Sun and the Earth.
This direction was picked because it is the direction we want the \Jspacecraft\j's thrusters to be in when we perform our first and largest Trajectory Correction Maneuver.
This maneuver is now scheduled for the evening of 3 January 1997 (Important: see update of 2 January 1997). It has been fully designed and tested using the Flight System Testbed \Jspacecraft\j simulation at JPL. Because of the superb performance of the Delta II rocket in getting Mars Pathfinder into a Martian trajectory, this maneuver is expected to use less than 25% of the 93 kilograms of hydrazine on-board the \Jspacecraft\j. It will change the velocity of the \Jspacecraft\j by about 30 m/s over a burn time of about two hours.
#
"Pathfinder Status (January97)",24,0,0,0
2 January 1997
The \Jspacecraft\j is currently 8 million kilometers from Earth and traveling at 32.5 kilometers per second. All \Jspacecraft\j subsystems continue to operate as expected.
Last Friday, we successfully turned the \Jspacecraft\j's spin axis about 43 degrees to the attitude we want the \Jspacecraft\j to be in when we perform our first and largest TCM (Trajectory Correction Maneuver). This had been planned for the evening of 3 January 1997.
However, in the course of testing this TCM using detailed models of the \Jspacecraft\j and the celestial sensors, we discovered that due to the partial obscuration of the sun sensors that occurred shortly after launch, the attitude control software would have unnecessarily fired spin thrusters.
Although in no way dangerous to the \Jspacecraft\j, we would rather not have the thrusters fire any more than absolutely necessary. This problem can be easily solved by changing a parameter in the flight software that will partially reject the bad sun sensor data. Because of the short time before Friday's planned TCM, we decided to postpone it.
Early next week we will send commands that will change the flight software parameters and later in the week we will do the TCM (now tentatively scheduled for 9 January 1997).
Starting tomorrow the team will resume its normal schedule.
10 January 1997
The Mars Pathfinder \Jspacecraft\j is currently 10 million km (6.2 million miles) from Earth, traveling at 32 km/s on its trajectory to Mars. All \Jspacecraft\j subsystems continue to operate as expected.
Earlier in the week we successfully updated attitude control software to further compensate for the apparent obscuration of the sun sensors. Miguel San Martin, lead Attitude Control Subsystem Engineer, stated: "This software update should allow the \Jspacecraft\j to perform all turns and maneuvers needed to get it to Mars. We are very pleased with the subsystem's performance in spite of the sun sensor obscuration."
At 7:40 PM PST on January 9, we successfully performed our first and largest Trajectory Correction Maneuver (TCM). This maneuver, planned by lead Flight Engineer Rob Manning, used two of the \Jspacecraft\j's eight 1 pound thrusters. These thrusters were fired continuously for an hour and a half, and changed the velocity of the \Jspacecraft\j by 31 meters/second.
The purpose of this maneuver was to correct small launch vehicle targeting errors and reduce the planetary protection trajectory bias. Later we turned the \Jspacecraft\j's spin axis 35 degrees back toward Earth so that we can perform radio navigation more effectively. We will leave the \Jspacecraft\j in this attitude until the next TCM scheduled for early February.
From the Doppler radio signature, it appears that last night's TCM performance was well within our expectations. In the next few days, radio ranging data will allow us to precisely gauge the quality of the maneuver.
Next week we will turn on the backup heat rejection system pump and allow it to operate for a while in parallel with the primary pump. These pumps circulate Freon around the perimeter of the cruise stage and down into the lander to keep the lander and rover \Jelectronics\j cool during the 7 month cruise to Mars.
24 January 1997
The \Jspacecraft\j continues to function well and is currently 14 million kilometers from Earth. The major activity for last week was starting the performance testing of the K=15 R=1/6 convolutional code. Early tests results indicate that the Block III MCD is operating as expected and that the expected telecom link improvements match predicts.
The project is investigating some minor anomalies which occurred this week involving the \Jspacecraft\j Command Detector Unit. The most serious of these occured on Monday, January 20 when the CDU transitioned to a lock state during a period when we were not uplinking to the \Jspacecraft\j.
The CDU lost lock as expected when we transmitted an uplink signal, but this "self-lock" behavior is not expected. In addition, we experienced two other episodes where commands were rejected by the \Jspacecraft\j uplink hardware for unexplained reasons.
The project has started a tiger team activity to further investigate these problems and determine potential causes. The possibility of solar flare induced SEUs is being assessed, but we are not ruling out other potential causes.
The flight team is also investigating an attitude control fault that we experienced on Sunday, January 19. This fault occured when estimates of the attitude \Jcovariance\j matrix inexplicably jumped by several orders of magnitude. The resulting fault response reinitialized attitude control flight software and turned off the Propulsion Drive \JElectronics\j.
Analysis of telemetry before and after the fault did not show a definitive cause. We are currently planning to perform additional diagnostic tests to assess memory and software integrity in case SEUs or numerical divergence problems may have caused this problem.
Continuing EDL and surface operations planning. The Rover team completed a successful Rover Operations Readiness Test, and planning is nearly complete for the first full team Surface Operations Readiness Test on January 27-28.
31 January 1997
The \Jspacecraft\j continues to be in excellent health, and is now about 16 million km from Earth. Key activities completed this week include successful completion of the K=15, R=1/6 convolutional code tests and resolution of the attitude control software glitch detected last week.
Attitude control software has been re-enabled and is currently operating nominally. In addition, we verified that the noise seen during ASI/MET health checks is due to the Propulsion Drive \JElectronics\j. This noise appears to be radiative in nature, and will not be an issue for surface operations because the PDE is located on the cruise stage.
The Uplink Problem Tiger team has developed a plausible explanation for the majority of the command rejections and the CDU In Lock conditions. It involves harmonics from the uplink sweep locking up the CDU and pulling it away from the nominal command frequency.
The team is developing a test plan to confirm this hypothesis and is also gathering information about the incidents where the CDU went into lock while we were not tracking.
An Operational Readiness Test (ORT) of the Sol 1-2 sequences was run on Jan 27 and 28. The sequences used were identical to the last pre-launch surface ORT. The ORT was successful in that all of the sequences were executed properly by the simulated lander and rover.
However, a number of relatively minor problems were logged during the test. These problems were reviewed and action has been assigned in all cases for problem resolution.
Nineteen investigators have been selected by NASA Headquarters in response to the Announcement of Opportunity for selection of Mars Pathfinder Participating Scientists and a Facility Instrument Science Team for the Atmospheric Structure Instrument/Meteorology Package.
An "All Hands" Pathfinder Science Team meeting has been set up for Feb. 5-7, 1997 at JPL to begin integrating the new investigators into the Experiment Operations Team.
#
"Pathfinder Status (February97)",25,0,0,0
4 February 1997
The Mars Pathfinder \Jspacecraft\j is currently 19 million km (11 million miles) from Earth traveling at 30 km/s on its trajectory to Mars . All \Jspacecraft\j subsystems continue to operate as expected.
At 5:00 PM PST on February 3, we successfully completed our second Trajectory Correction Maneuver. This maneuver was designed to correct errors in the first TCM performed on January 9, and move us closer to our final trajectory.
The \Jspacecraft\j will not be placed on a Mars atmospheric entry trajectory until after TCM-3 (currently scheduled for May 5) because of planetary quarantine requirements. The TCM-2 design team, led by Flight Engineer Guy Beutelschies, developed a two part approach to perform the maneuver. In the first part, the \Jspacecraft\j fired two of its forward facing thrusters continuously for five minutes. The change in velocity for this "axial" component was about 1.5 m/s.
The second part of the maneuver was a smaller velocity correction of 0.1 m/s performed in the "lateral" mode. In this mode, the \Jspacecraft\j pulses all four thrusters on one side of the \Jspacecraft\j for five seconds. This pulse causes a small change in the \Jspacecraft\j velocity in the direction perpendicular to the \Jspacecraft\j spin axis.
This mode will be used for all future maneuvers, so TCM-2 was a good proof-of-concept test. Early analysis of tracking data from NASA's Deep Space Network indicates that both components were completed successfully.
Upon completing the maneuver, the \Jspacecraft\j's spin axis was turned 15 degrees back toward Earth so that we can perform radio navigation more effectively. The \Jspacecraft\j is currently pointed about 5 degrees from Earth and 2 degrees from the Sun. We will remain in this attitude until late March.
The \Jspacecraft\j will remain in a relatively quiescent mode for the next two to three months. The flight team is currently working hard to complete planning for Mars entry and surface operations.
7 February 1997
Successfully completed Trajectory Correction Maneuver #2 on February 3. The purpose of this maneuver was to clean up TCM-1 execution errors and had a magnitude of about 1.6 m/s. The maneuver consisted of two parts, an axial component of 1.5 m/s and a lateral component of 0.1 m/s.
All \Jspacecraft\j subsystems performed as expected, and the resulting maneuver execution error was less than 2%. The \Jspacecraft\j was turned back to Earth after the maneuver, and will remain in this attitude until late March.
The \Jspacecraft\j computer reset during an off-track period on Wednesday, February 5. Analysis of post reset telemetry indicates that it was caused by a divide-by-zero fault in the attitude control flight software (ACS). The \Jspacecraft\j responded as expected to this anomaly, and correctly enforced the early cruise boot configuration.
This configuration idles ACS and reduces the uplink and downlink data rates from nominal cruise values. Several commands were sent on February 6 to increase the data rates and return diagnostic data, but attitude control is still idle. The \Jspacecraft\j is in a safe attitude, however, so there is no compelling reason to restart ACS.
An all-hands science team meeting was held on February 5-7. All existing investigators plus the newly selected ASI/MET FIST scientists and participating scientists were invited. The purpose of this meeting was to organize the science teams into a set of science operations groups and to review the current baseline plans for surface operations. A number of useful suggestions were made for modifying the nominal Sol 1-2 plans.
21 February 1997
The \Jspacecraft\j remains in excellent health and is currently about 25 million kilometers from Earth. No significant operational activities were conducted this week, and we have adopted a policy of unattended operations for most of our tracking sessions.
The DSN Radiometric Tiger Team reported the cause of the ~500 Range Unit bias seen in DSS 15 range data between TCM-1 and TCM-2 was caused by an incorrect value for the signal inversion parameter in the SRA configuration table. The error has been fixed and subsequent DSS 15 data looks good.
A set of improvements to the Flight System Testbed Ground Support Equipment Software have been completed which will allow higher fidelity testing of the EDL flight software. Robustness testing of the EDL software has now been started and will continue for the next two months.
28 February 1997
The \Jspacecraft\j remains in excellent health and is currently about 32 million kilometers from Earth. No significant operational activities were conducted this week. Continued investigation of the recent reset and attitude control software problems now indicates that they are related to the Command Detector Unit erroneous lock problem.
A bug was found in the codeblock error detection software which causes a corruption to the floating point registers. The reset and attitude control problems were caused by floating point error conditions and can be tied directly to this corruption.
We are currently in the process of developing a patch to correct this problem. Congratulations to Steve Stolper and Glenn Reeves for quickly identifying the problem and developing the required fix.
A combined Project Science Group meeting involving Mars Pathfinder, Mars Global Surveyor, and Mars Surveyor '98 was held on Thursday and Friday, February 27-28. Although there is not a great deal of overlap between Pathfinder and these other projects, there are some synergistic investigations that can be performed.
The Rover Operations Team completed a Rover Operational Readiness Test in the Mars Yard this week. Preparations are proceeding for next week's project wide surface Operations Readiness Test in the Pathfinder Sandbox. The current plan is to conduct nominal Sol 1 and 2 operations.
#
"Pathfinder Status (March97)",26,0,0,0
7 March 1997
The \Jspacecraft\j is currently about 37 million kilometers from Earth and continues to function as expected. The total travel distance covered since launch is 248 million kilometers, which means that the \Jspacecraft\j has reached the halfway point to Mars.
A set of Entry, Descent, and Landing communications tests were started this week using the \Jspacecraft\j and the Deep Space Network \JGalileo\j Telemetry recorders at Goldstone. These tests are meant to simulate the open loop strategy that we intend to use during entry to record significant events. The first test was successfully completed on March 3, and three additional tests will be performed during the next week.
The project completed Surface Operational Readiness Test #3 on March 7-8. This test was the first formal operations test after launch, and was designed to test the nominal Sol 1 and 2 sequences. Although there were a few start up problems, the test was generally successful. All elements of the project worked well together to complete the critical Sol 1 operations and re-plan Sol 2.
The Rover Operations Team performed remote field testing on Monday and Tuesday. With the SIM Rover at Amboy Crater, the Operations Team ran four Martian sol sequences from JPL. The sequences included navigation and traverse activities, and science and technology experiments. The Pathfinder Science Team also participated in the testing.
14 March 1997
The \Jspacecraft\j remains in good health and is currently about 44 million kilometers from Earth. We experienced a minor command error this week when a command was sent to activate a sequence which was not on-board. The \Jspacecraft\j rejected the command as expected, but we have tightened up our command approval process to prevent future incidents.
The set of Entry, Descent, and Landing (EDL) communications tests started last week were completed this week. All of the Deep Space Network hardware and software elements required to support EDL communications are performing as expected.
EDL Flight software testing is progressing well, and has been the primary focus of testbed activities this week. A full scale airbag retraction test is planned for Friday, March 14 in the Mars sandbox in Building 230.
21 March 1997
The \Jspacecraft\j remains in good health and is currently about 49 million kilometers from Earth. No significant \Jspacecraft\j operations were performed this week.
Entry, Descent, and Landing Flight Software Testing is progressing well. We completed an airbag retraction test in the Building 230 Sandbox last week, and have finished a set of parachute deploy and rocket ignition \Jalgorithm\j robustness tests this week.
A couple of significant issues have come up in this testing which are likely to cause us to change flight software. A flight software change board meeting will be held on April 1 to determine what changes to make and what regression tests to perform.
28 March 1997
The \Jspacecraft\j remains in good health and is currently about 55 million kilometers from Earth. The most significant \Jspacecraft\j activity performed this week was to turn the \Jspacecraft\j to a 5 degree Earth leading attitude. Regular attitude turns will be required for the remainder of cruise to keep the \Jspacecraft\j pointed within 5 degrees of Earth. The propulsion and attitude control subsystems functioned flawlessly after a 7 week hiatus.
Successfully completed a set of sun recognition tests using the Prototype IMP at the University of \JArizona\j. These tests have boosted our confidence that we can successfully perform sun search and point the High Gain Antenna after landing.
An IMP Science Team meeting was held at the University of \JArizona\j on March 25, 26. The sequence changes for Sols 1 and 2 were reviewed. The mission plan for sols 3-5 was discussed and a scenario for the imaging observations was adopted. Discussions of image processing plans and data distribution were held and a number of contentious issues were resolved.
#
"Pathfinder Status (April97)",27,0,0,0
4 April 1997
The \Jspacecraft\j remains in good health and is currently about 64 million kilometers from Earth. The most significant \Jspacecraft\j activity performed this week was to switch to a new convolutional code, K=15, Rate 1/6, on our downlink.
This code gives us a significant increase in our downlink capability. Mars Pathfinder is the first \Jspacecraft\j to operationally use this code with the DSN and we are very pleased with the results.
The project conducted a meeting to discuss Flight Software changes for Entry, Descent and Landing (EDL) and Surface phases of the mission. These changes are to correct bugs found during testing performed since launch. We expect to finalize these changes and prepare for loading them onboard within the next two months.
11 April 1997
The \Jspacecraft\j remains in good health and is currently about 71 million kilometers from Earth. The only \Jspacecraft\j activities performed this week were a regular ASI/MET health check, a Heat Rejection System Pump B cycle, and some modification to Attitude Control Subsystem fault protection parameters. The total flight time since launch is now 128 days, and we have 85 days until Mars arrival.
We successfully completed an operational readiness test of all activities from Mars entry -2 days through Sol 2. This test was performed using the Pathfinder Testbed and Mars Sandbox. The Mars approach phase included periodic updates to the Entry, Descent, and Landing flight software parameter set and execution of a contingency Trajectory Correction Maneuver #5.
We are investigating a minor problem which occurred during airbag retraction, which did not significantly effect the test. The surface operations test included a planned failure of the High Gain Antenna and subsequent Low Gain Antenna operations.
John Wellman and Matt Golombek attended a special Project Science Group meeting at NASA Headquarters to discuss data rights issues and the role of the participating scientists. A number of useful discussions were held and a draft policy on these issues was developed.
18 April 1997
The \Jspacecraft\j remains in good health and is currently about 80 million kilometers from Earth. The only activity performed this week was to continue gathering sun sensor data to characterize its performance at large sun angles. No additional degradation of the sun sensor has been observed since launch. The total flight time since launch is now 135 days, and we have 78 days until Mars arrival.
The flight team is completing preparations for next week's Operational Readiness Test #5. This test is a five day simulation of surface operations using the flight system testbed and Mars sandbox. A significant number of science team members and participating scientists will be in attendance, and are already here this week conducting test and training activities.
The project completed the first of two sessions on lessons learned during the Mars Pathfinder development effort. The first session focused on system level issues and was well received by a lab-wide audience. The second session is scheduled for April 28, and will cover subsystem lessons. Both sessions are being videotaped, and the presentation material will be gathered into a single "book".
25 April 1997
The \Jspacecraft\j remains in good health and is currently about 88 million kilometers from Earth. Major activities performed this week included a regularly scheduled attitude turn to maintain Earth point. We also transitioned to the Late Cruise mission phase and switched in the fourth and final solar panel quadrant. The total flight time since launch is now 142 days, and we have 71 days until Mars arrival.
We successfully completed a week long surface Operational Readiness Test (ORT #5). The purpose of this test was to train team members on operational processes and procedures and verify nominal surface operations plans. Although we had some early difficulties deploying the rover, the test was a great learning experience and an overall success. We did discover several issues with our tools and processes that we will correct prior to ORT #6 (scheduled for May 19).
#
"Pathfinder Status (May97)",28,0,0,0
2 May 1997
The \Jspacecraft\j remains in good health and is currently about 98 million kilometers from Earth. No significant \Jspacecraft\j activities were performed this week. The total flight time since launch is now 149 days, and we have 63 days until Mars arrival.
The EDL planning team completed a very successful Sequence of Events Peer Review. This review covered the detailed flight software sequence used during EDL, recent changes in the EDL software, robustness and regression testing performed since launch, and detailed operational plans for TCM-5 and EDL approach.
The review board, led by Jim Marr, agreed that no significant holes exist in the EDL sequence and that we are well on our way to being prepared for EDL operations.
9 May 1997
The \Jspacecraft\j remains in good health and is currently about 107 million kilometers from Earth. We performed our third Trajectory Correction Maneuver as scheduled on May 6. The purpose of this maneuver was to make a small correction in the trajectory and to test our implementation approach for a potential contingency maneuver just before Mars entry.
The maneuver consisted of three components: a 0.4 m/s lateral burn away from Mars, a 0.1 m/s axial mode burn to correct arrival time, and a 0.5 m/s burn back towards Mars. All three burns were performed without incident, and placed us on a trajectory that is closed to our desired Mars entry trajectory.
We will perform the final planned maneuver (TCM-4) on June 24 to clean up TCM-3 execution errors. The \Jspacecraft\j is now back into quiescent cruise mode. The total flight time since launch is now 156 days, and we have 56 days until Mars arrival.
16 May 1997
The \Jspacecraft\j remains in good health and is currently about 115 million kilometers from Earth (26 million km from Mars). The only major \Jspacecraft\j activity performed this week was a battery heating and solar array characterization test. The total flight time since launch is now 163 days, and we have 49 days until Mars arrival.
Successfully completed the Entry, Descent, and Landing (EDL) and Surface Operations Readiness Review. The review board, led by Mike Sander, asked many useful questions and generated several advisories, but agreed that the project will be ready for pre-entry and surface operations by July 4.
Completed the second of three EDL Operations Readiness Tests (ORT). This test used the testbed to simulate all pre-EDL operations, including TCM-5a. In addition, we conducted a simultaneous EDL communications ORT using the actual \Jspacecraft\j to simulate the telecom behavior during EDL.
Several members of the EDL data acquisition team traveled to Madrid to support this test. Although a number of lessons were learned during the ORT, both the data acquisition and pre-EDL operations teams completed the test successfully.
Completed a set of surface operations mini-ORTs which tested our strategy for petal movements after landing, end-to-end image processing and rover target designation, and the ramp deployment decision process. These three issues were among the most significant concerns in ORT #5, and appear to be resolved at this point.
23 May 1997
The \Jspacecraft\j remains in good health and is currently about 129 million kilometers from Earth. The only significant \Jspacecraft\j activity performed this week was a turn to maintain Earth point attitude. In addition, we have resumed nearly continuous DSN coverage. The total flight time since launch is now 170 days, and we have 42 days until Mars arrival.
Completed the sixth Surface Operations Readiness Test (ORT). The purpose of this test was to validate our low power and no battery contingency scenarios and correct problems from ORT #5. The test was hampered by a recurring testbed hardware problem involving the Imager for Mars Pathfinder (IMP) and the power support equipment.
The problem caused the testbed flight computer to reset several times and caused the flight team to invoke reset recovery procedures instead of performing normal operations. The problem was fixed after Sol 3, at which time we restarted the test and ran through Sol 1-2 operations. In spite of this problem, the test was generally a success in that we exercised all of our core operational processes. Several other minor problems occurred, and all will be corrected by ORT #7.
30 May 1997
The \Jspacecraft\j remains in good health and is currently about 140 million kilometers from Earth (17 million km from Mars). Major \Jspacecraft\j activities performed this week included a turn to maintain Earth point attitude and starting battery charge. The total flight time since launch is now 175 days, and we have 35 days until Mars arrival.
The most significant project activity completed this week was to begin charging the flight battery. Approximately 22 amp-hours of capacity has been taken out of the battery since installation, and the objective of charging was to replace as much as possible.
Degradation of the total battery capacity has occured over the six months since launch, but a total capacity of at least 40 amp-hours should still be possible (compared to a pre-launch capacity of 56 amp-hours). A total of 7 amp-hours has been added after two days of charging, so we are already above the 40 amp-hour level. The flight team is currently assessing whether additional charging is warranted.
Completed updates to the final flight software load. We are currently performing a final set of regression tests prior to patching the software next week.
Deep Space Network personnel corrected the final open ranging data accuracy problem which has been a concern for the last several months. This problem involved larger than expected range jitter in the data acquired at Goldstone. The problem is related to a faulty board in the Sequential Ranging Assembly. All ranging data acquired for the project now meets the pre-launch accuracy specifications.
#
"Pathfinder Status (June97)",29,0,0,0
6 June 1997
The \Jspacecraft\j remains in good health and is currently about 148 million kilometers from Earth (14 million km from Mars). The total flight time since launch is now 182 days, and we have 28 days until Mars arrival.
The flight team completed uploading the final flight software patch to the \Jspacecraft\j on June 5-6. The patch corrects a number of problems in the Entry, Descent, and Landing control software. The total volume of the patch was approximately 400 Kbytes.
The decision was made to delay using the new software until after Operational Readiness Test #7 (ORT #7) because it will be a good final regression test. The \Jspacecraft\j will begin using the new software on June 16.
Completed two mini-ORTs which tested our petal move and low gain antenna contingency plans for Sol 1. All flight sequences for pre-entry and surface operations have now been completed (except for some minor tweaks to the backup mission load), and will be loaded on the \Jspacecraft\j on June 18. We have also completed all preparations for the mission dress rehearsal (ORT #7), which will start on June 9.
Completed a detailed review of Public Outreach plans involving project, program office, and JPL PIO personnel. No major issues or concerns exist, and the detailed implementation activities are proceeding well.
13 June 1997
The \Jspacecraft\j remains in good health and is currently about 160 million kilometers from Earth (10 million km from Mars). The total flight time since launch is now 190 days, and we have 21 days until Mars arrival.
The Operations Team completed the final Operational Readiness Test in preparation for Entry, Descent, and Landing (EDL) and surface operations. This test simulated activities from two days before landing through two days after landing and was very successful.
Major activities completed during this ORT included a simulated Trajectory Correction Maneuver (TCM)-5, a nominal EDL, a series of simulated press conferences for Sol 1 & Sol 2, a successful rover deployment of Sol-1, and extended rover traverses over the course of the following two days.
20 June 1997
The \Jspacecraft\j remains in good health and is currently about 168 million kilometers from Earth (7 million km from Mars). The total flight time since launch is now 196 days, and we have 14 days until Mars arrival.
Completed the final flight software load process including resetting the flight computer. The \Jspacecraft\j is now operating using the new software, including two small changes made as a result of Operational Readiness Test #7. The patch process worked flawlessly and we recovered from the reset within two hours.
Completed the rover flight software patch and health check. A number of changes to the rover software have been identified as a result of post-launch testing, so a small patch was performed this week in conjunction with the planned rover health check. The rover woke up as expected based on lander command, the code patch was accepted, and all \Jengineering\j telemetry measurements were normal.
Completed validating all sequences required for pre-EDL and initial surface operations. Approximately 370 sequences will be loaded on the \Jspacecraft\j starting this weekend.
TMOD conducted a readiness review for Mars Pathfinder EDL and Surface Operations on June 13. No major issues were identified, and all DSN and MGSO elements will be ready to support project operations.
27 June 1997
Mars Pathfinder, now eight days away from landing on the surface of Mars, performed the last of its scheduled trajectory correction maneuvers at 10 a.m. Pacific Daylight Time on Wednesday, June 25.
The correction maneuver was performed in two phases occurring 45 minutes apart. The first burn, lasting just 1.6 seconds, involved firing four thruster engines on one side of the vehicle. The second burn lasted 2.2 seconds and involved firing two thrusters closest to the heat shield.
The combined effect of both burns changed Pathfinder's velocity by 0.018 meters per second (0.04 miles per hour), which places the \Jspacecraft\j on target for a July 4 landing in an ancient flood basin called \JAres\j Vallis. Pathfinder is scheduled to land at 10:07 a.m. PDT (in Earth-received time). The one-way light time from Mars to Earth is 10 minutes, 35 seconds, so in actuality, Pathfinder lands at 9:57 a.m. PDT.
If necessary, a fifth trajectory correction maneuver may be performed just before Pathfinder hits the upper atmosphere of Mars. The maneuver would be carried out either 12 hours or six hours before Pathfinder reaches the atmosphere at 10 a.m. PDT in Earth-received time. The flight team will make a decision to proceed with the final correction maneuver the evening before landing.
A final health check of the \Jspacecraft\j and rover was performed on June 20. All \Jspacecraft\j systems, including science instruments and the critical radar \Jaltimeter\j, remain in excellent health from the last check about six months ago.
The rover received a "wake up" call, woke up on command from the lander, then accepted a software upgrade. Flight controllers next loaded the 370 command sequences that will be required by Pathfinder to carry out its surface operations mission.
The \Jspacecraft\j is now ready to begin its entry, descent and landing phase. It will be commanded into that mode at 1:42 p.m. PDT on June 30 by an onboard sequence.
Mars Pathfinder is currently about 180 million kilometers (111 million miles) from Earth and about 3.5 million kilometers (2.2 million miles) from Mars. After 202 days in flight, the \Jspacecraft\j is traveling at about 18,000 kilometers per hour (12,000 miles per hour) with respect to Mars.
#
"Pathfinder: Entry, Descent and Landing",30,0,0,0
The entry, descent and landing (EDL) process for Mars Pathfinder will begin days before landing when controllers at JPL will send commands to the \Jspacecraft\j to tell it precisely when and how to begin the complex autonomous series of steps necessary to safely land on the surface of Mars. These commands are sent periodically right up to a few hours before landing, when controllers on the Earth will have the most precise knowledge of where the \Jspacecraft\j is relative to Mars (the effect of Mars' gravity well is not felt until the \Jspacecraft\j is less than 48 hours away).
Landing occurs at about 3:00 am local time on Mars, which will be about 10:00 am PDT on Friday, July 4, 1997. From an hour and a half before landing until about 3 and a half hours later, the \Jspacecraft\j is under control of autonomous on-board software that precisely controls the many events that must occur.
The fast-paced approach of Pathfinder at Mars begins with venting of the heat rejection system's cooling fluid about 90 minutes prior to landing. This fluid is circulated around the cruise stage perimeter and into the lander to keep the lander and rover cool during the 7 month cruise phase of the mission.
Its mission fulfilled, the cruise stage is then jettisoned from the entry vehicle about one-half hour prior to landing at a distance of 8500 km from the surface of Mars.
Several minutes before landing, the \Jspacecraft\j begins to enter the outer fringes of the atmosphere about 125 km. (80 mi.) above the surface. Spin stabilized at 2 rpm, and traveling at 7.5 km/sec, the vehicle enters the atmosphere at a shallow 14.8 deg angle. A shallower entry angle would result in the vehicle skipping off the atmosphere, while a steeper entry would not provide sufficient time to accomplish all of the entry, descent and landing tasks.
A Viking-derived aeroshell (including the heatshield) protects the lander from the intense heat of entry. At the point of peak heating the heatshield absorbs more than 100 megawatts of thermal energy. The Martian atmosphere slows the vehicle from 7.5 km/sec to only 400 m/sec (900 mph).
Then entry deceleration of up to 20 gees, detected by on-board accelerometers, sets in motion a sequence of preprogrammed events that are completed in relatively quick succession.
Deployment of the single, 24-ft. diameter parachute occurs 2-3 min. after atmospheric entry at an altitude of 5-11 km. (3-7 mi.) above the surface, eventually slowing the vehicle down to 65 meters/sec. The parachute is similar in design to those used for the Viking program but has a wider band around the perimeter which helps minimize swinging.
The heatshield is pyrotechnically separated from the lander 20 sec. later and drops away at an altitude of 2-9 km. (3-6 mi.). The lander soon begins to separate from the backshell and "rappels" down a metal tape on a centrifugal braking system built into one of the lander petals.
The slow descent down the metal tape places the lander into position at the end of a braided Kevlar tether, or bridle, without off-loading the parachute or placing excessive loads on the backshell. The 20 m bridle provides space for airbag deployment, distance from the solid rocket motor exhaust stream and increased stability.
Once the lander has been lowered into position at the end of the bridle, the radar \Jaltimeter\j is activated and aids in the timing sequence for airbag \Jinflation\j, backshell rocket firing and the cutting of the Kevlar bridle.
The lander's Honeywell radar \Jaltimeter\j is expected to acquire the surface about 32 sec. prior to landing at an altitude of about 1.5 km. The airbags are inflated about 8 sec. before landing at an altitude of 300 meters above the surface.
The airbags have two pyro firings, the first of which cuts the tie cords and loosens the bags. The second, 0.25 sec. later, and 4 sec. before the rockets fire, ignites three gas generators that inflate the three 5.2 m (17-ft) dia. bags to a little less than 1 psi. in less than 0.3 sec.
The conical backshell above the lander contains three solid rocket motors each providing about a ton of force for over 2 seconds. They are activated by the computer in the lander. Electrical wires that run up the bridle close relays in the backshell which ignite the three rockets at the same instant.
The brief firing of the solid rocket motors at an altitude of 80-100 meters is intended to essentially bring the downward movement of the lander to a halt some 12 meters (▒10 m) above the surface. The bridle separating the lander and heatshield is then cut in the lander, resulting in the backshell driving up and into the parachute under the residual impulse of the rockets, while the lander, encased in airbags, falls to the surface.
Because it is possible that the backshell could be at a small angle at the moment that the rockets fire, the rocket impulse may impart a large lateral velocity to the lander/airbag combination. In fact the impact could be as high as 25 m/sec (56 mph) at a 30 deg grazing angle with the terrain.
It is expected that the lander may bounce at least 12 m about the ground and soar 100-200 m between bounces. (Tests of the airbag system verified that it was capable of much higher impacts and longer bounces.)
Once the lander has settled on the surface, pyrotechnic devices in the lander petal latches are blown to allow the petals to be opened. The latches locking the sturdy side petals in place are necessary because of the pulling forces exerted on the lander petals by the deployed airbag system.
In parallel with the petal latch release, a retraction system will begin slowly dragging the airbags toward the lander, breaching vent ports on the side of each bag, in the process deflating the bags through a cloth filter. The airbags are drawn toward the petals by internal lines extending between attachments within the airbags and small winches on each of the lander sides. It takes about 64 minutes to deflate and fully retract the bags.
There is one high-torque motor on each of the three petal hinges. If the lander comes to rest on its side, it will be righted by opening a side petal with a motor drive to place the lander in an upright position. Once upright, the other two petals are opened.
About 3 hours is allotted to retract the airbags and deploy the lander petals. In the meantime, the lander's X-band radio transmitter will be turned off for the first time since before launch on December 4, 1996. This saves battery power and will allow the transmitter \Jelectronics\j to cool down from being warmed up during entry without the cooling system.
It also allows time for the Earth to rise well above the local horizon and be in a better position for communications with the lander's low-gain antenna later in the morning.
Normal digital data transmissions will cease near the time of cruise stage separation due to the dynamics of EDL. Instead, the transmitter's carrier signal and sidebands will be recorded by the Deep Space Network's Madrid station so that the effects of the many events on the signal may be discerned. The digital data downlink will automatically resume 3.5 hours after landing, long after the airbags have been retracted and the petals opened.
#
"Mir 23 Status Report",31,0,0,0
Mission Control Center, Korolev
\BApril 29, 1997\b
U.S. \Jastronaut\j Jerry Linenger and Mir 23 Commander Vasily Tsibliev conducted a successful five-hour spacewalk today, the first joint U.S.-Russian spacewalk ever undertaken, to attach and retrieve several experiments designed to collect data on the environment around the orbiting space complex.
Linenger and Tsibliev opened the airlock hatch on the Kvant-2 module at 12:10 a.m. Central time this morning and the two spacewalkers went right to work, testing the mobility and design of new Orlan-M spacesuits earmarked for eventual use in the assembly of the International Space Station.
Linenger and Tsibliev reported that new visors in the spacesuit helmets to protect them from the harsh effect of the Sun worked to perfection, and prevented their visors from fogging during the most strenuous periods of activity. Linenger was congratulated by Russian ground controllers at the start of his first spacewalk as he and Tsibliev used a telescoping cargo crane to move themselves and their equipment from the Kvant-2 module to the Mir's Docking Module for the installation of the Optical Properties Monitor (OPM).
The device, which is designed to collect data on the environment around the Mir, was installed near a pair of similar experiments attached to the Docking Module by STS-76 spacewalkers Linda Godwin and Rich Clifford 13 months ago. The so-called Mir Environmental Experiment Packages (MEEPS) will be retrieved by veteran \Jcosmonaut\j Vladimir Titov and \Jastronaut\j Scott Parazynski during a spacewalk outside Atlantis during the STS-86 mission to the Mir in September.
Titov will become the first Russian to conduct a spacewalk wearing a U.S. suit during that excursion. A short time after its installation, the OPM was activated and was reported to be in good working order.
With their first task completed, Linenger and Tsibliev returned to the cargo crane and slowly swung back to the Kvant-2 module, where they installed a meter to monitor radiation levels around the Mir. Video of the spacewalk, downlinked to the Russian Mission Control Center by Flight Engineer Alexander Lazutkin inside Mir, showed the two spacewalkers operating with care near the Mir's delicate solar arrays as they worked to the timeline crafted over the past year.
Linenger and Tsibliev then retrieved a pair of micrometeorite and debris particle collection experiments from the exterior of Kvant-2 which had been left outside by Mir 21 cosmonauts Yuri Onufrienko and Yuri Usachev last year. The experiments were returned to the Kvant-2 airlock where they will be stowed before being brought back to Earth.
Finally, at 5:08 a.m. Central time, after five hours outside Mir, Linenger and Tsibliev returned to Kvant-2 and repressurized the airlock to complete the spacewalk. It was Tsibliev's sixth excursion outside Mir in his two flights dating back to 1993.
Gen. Yuri Glazkov, the Deputy Director of the Gagarin \JCosmonaut\j Training Center in Star City, outside Moscow, congratulated Linenger and Tsibliev for their performance following the completion of the spacewalk for which they had trained for over a year. In a post-spacewalk debriefing with flight controllers, Tsibliev again praised the new spacesuits, particularly the helmet visors and the flexibility of the shoulders, arms and knees, which enabled him and Linenger to move with relative ease outside the station.
Tsibliev plans two more spacewalks outside Mir with Lazutkin in late June and early July to erect an experiment platform on the Spektr module and to prepare valves on the outside of the Core Module for later work designed to add a second carbon dioxide removal system to the outpost.
The crewmembers plan to relax on Wednesday before resuming their scientific agenda and their search and repair of a small cooling loop leak in the Kvant-1 module. Otherwise, the Mir's systems continue to operate normally as plans proceed for the launch of Atlantis in mid-May to deliver U.S. \Jastronaut\j Mike Foale to the complex to replace Linenger.
\BJuly 21 1997\b
Mir Spacewalk Officially Rescheduled to August
Top Russian space officials met Monday and officially announced that the next Mir crew, Commander Anatoly Solovyev and Flight Engineer Pavel Vinogradov, will perform the internal spacewalk to try to restore power from solar arrays outside the damaged Spektr module. The pair is scheduled to be launched Aug. 5, arriving two days later at the Russian station.
Mir 23 cosmonauts Vasily Tsibliev and Alexander Lazutkin are expected to return to Earth on Aug. 14, after the completion of a one-week handover with Solovyev and Vinogradov. The internal spacewalk into Spektr will occur no earlier than Aug. 20.
#
"Shuttle Columbia Mission",32,0,0,0
Space Shuttle Columbia
Last Updated: June 27, 1997, 11:50 a.m. CDT
Preparations for the fifth space shuttle mission of 1997 are on schedule at the Kennedy Space Center in \JFlorida\j, where the shuttle Columbia was rolled out to launch pad 39-A last week.
Ground crews remain on track for a targeted July 1st launch of mission STS-94, the reflight of April's \Jmicrogravity\j science laboratory mission, which was curtailed after four days due to indications of a faulty power-producing fuel cell. Mission managers plan to start Columbia's fuel cells earlier than usual on this mission to allow for additional monitoring prior to liftoff. An official launch date for STS-94 should be set at next Thursday's shuttle program flight readiness review at the Kennedy Space Center.
Why STS-94?
A shuttle flight takes approximately two years to prepare, including all the safety reviews of the planned operations, \Jengineering\j \Jintegration\j to connect the payloads up properly in the payload bay, and development of the flight plan to meet the various payload requirements.
When a flight gets a number, all the documentation that has been prepared carries that designation. If a flight slips for some reason, it would take a lot of time and money to change all the documentation just to keep the order right. Those resources are better spent preparing to go to work in space.
STS-83 landed early in April because of indications of a fuel cell problem. That same flight will fly again reusing all the STS-83 products, with their original designation. The same crew will fly again. But because STS-83 has already flown, NASA will designate the July reflight STS-94, which was the next available number.
The Crew:
James D. Halsell Jr., Commander
Susan Still, Pilot
Janice E. Voss , Mission Specialist
Michael L. Gernhardt, Mission Specialist
Donald Thomas, Mission Specialist
Roger K. Crouch, Payload Specialist
Gregory T. Linteris, Payload Specialist
Columbia Launches
Last Update: 1 July 1997, 6:30 p.m. CDT (-0500 GMT)
After a short delay due to bad weather, Columbia launched at 1:02 p.m. CDT.
The Space Shuttle Columbia and seven astronauts are back in space, resuming a mission of scientific research in \Jmicrogravity\j that was cut short earlier this year by a suspect fuel cell. Less than three months after returning to Earth, the same crew and the \JMicrogravity\j Science Laboratory-1 payloads are once again on orbit to study how certain materials behave and change in the absence of gravity.
This is the twenty-third flight of the oldest orbiter in the NASA fleet, the fifth shuttle mission of this year. The turnaround took only 82 days after the premature landing of Columbia on April 8th. Mission managers decided at that time to bring Columbia home after just four days on orbit due to indications of a problem with that power-producing fuel cell.
After extensive analysis of the fuel cell, the shuttle program concluded that an undetermined and isolated incident caused a slight voltage change in about one-fourth of the 96 cells that make up each fuel cell. To ensure the pre- launch health of those cells this time, the fuel cells were activated earlier than usual in the countdown to provide additional time for monitoring prior to lift-off.
Today's launch, at 1:02 p.m. CDT, was slightly delayed due to weather concerns at the Kennedy Space Center. Shortly after reaching orbit pilot Susan Still reported an unusual reading during shutdown of one of the orbiter's three auxiliary power units, but space shuttle launch \Jintegration\j manager Loren Shriller says mission managers don't believe that reading indicates any problem.
The shuttle's auxiliary power units will be used to supply the hydraulic power to run the orbiter's flight control surfaces during entry and landing. Mission managers also report no indication of anything unusual with Columbia's power-producing fuel cells.
This 85th mission in space shuttle program history is the first time that an entire crew from one mission has flown together a second time. These seven astronauts, operating in two teams to support around-the-clock operations in the Scapulae module, were already on their separate schedules when they started launch day activities very early today.
Columbia is orbiting at 160 nautical, with all systems operating in good shape and the red team astronauts already sent off to bed.
Leading the red team is commander Jim Hales, a 40-year-old Lieutenant Colonel in the U.S. Air Force on his fourth trip to space. Also on the red team is the pilot on this mission, Susan Still. STS-94 is the second spaceflight for the 35-year-old Lieutenant Commander in the U.S navy.
The other two red team members are mission specialist-3 Dr. Don Thomas, a 42-year-old materials scientist on his fourth trip to space, and payload specialist-2 Dr. Greg Linnets, a 39-year-old engineer on his second spaceflight. Linnets is the crew specialist on the combustion experiments in the Scapulae module.
Leading the blue team of astronauts is payload commander Dr. Janice Voss. The 40-year-old aeronautical engineer, making her fourth shuttle flight, has entered the Scapulae module to begin activating its systems. Working with her is payload specialist-1 Dr. Roger Crouch, a 56-year-old physicist on his second spaceflight. Crouch is the chief scientist of NASA's \JMicrogravity\j Science and Applications Division.
The final blue team member is mission specialist-2 Dr. Mike Gernhardt, a 41- year-old bioengineer on his third spaceflight. Gernhardt is seeing to the orbiter systems as flight engineer on this mission.
Orbiter operations for Columbia are being overseen from the Mission Control Center in Houston, and the science operations for the \JMicrogravity\j Science Laboratory are being managed from the Scapulae Operations Control Center, located at the Marshall Spaceflight center in \JHuntsville\j, \JAlabama\j.
A few minutes after 6:00 CDT this evening payload commander Janice Voss will go to work in the spacelab module activating payload systems. At 11:32 p.m. CDT the red team astronauts will get their wake-up call from Mission Control, and half an hour later they'll begin taking a handover from the blue team astronauts. Voss, mission specialist Mike Gernhardt, and payload specialist Roger Crouch begin an eight-hour sleep period shortly after 1:00 a.m.
At 2:17 a.m. Thomas and Linnets will begin their day's work with the experiments in the Scapulae module. At 5:32 a.m. Still will set up the wireless data acquisition system, a risk mitigation experiment for the international space station that will gather remote data on the health of payload hardware. At 9:02 a.m. the wake-up call will sound for the three blue team astronauts.
#
"Space Calendar June97 - July98",33,0,0,0
\JSpace Calendar (June97)\j
\JSpace Calendar (July97)\j
\JSpace Calendar (August97)\j
\JSpace Calendar (September97)\j
\JSpace Calendar (October97)\j
\JSpace Calendar (November97)\j
\JSpace Calendar (December97)\j
\JSpace Calendar (January98)\j
\JSpace Calendar (February98)\j
\JSpace Calendar (March98)\j
\JSpace Calendar (April98)\j
\JSpace Calendar (May98)\j
\JSpace Calendar (June98)\j
\JSpace Calendar (July98)\j
#
"Space Calendar (June97)",34,0,0,0
Jun 13 - \JComet\j C/1997 L3 (SOHO) Perihelion (0.00844 AU)
Jun 13 - Moon Occults Mars
Jun 13-18 - 13th North American Workshop on Cataclysmic Variables, Jackson Hole, Wyoming
Mir represents a unique capability -- an operational space station that can be permanently staffed by two or three cosmonauts. Visiting crews have raised Mir's population to six for up to a month.
#
"Mir Station - Progress-M",51,0,0,0
Progress-M is the unmanned supply ship used to send food and other supplies to the Astronauts and Cosmonauts aboard Mir. It is launched atop a Russian Soyuz SL-4 rocket.
#
"Mir Station - Core Module",52,0,0,0
Mir is the first space station designed for expansion. The 20.4-ton Core Module, Mir's first building block, was launched in February 1986. The Core Module provides basic services (living quarters, life support, power) and scientific research capabilities. It has two axial docking ports, fore and aft, for Soyuz-TM manned transports and automated Progress-M supply ships, plus four radial berthing ports for expansion modules.
The 20.4-ton Core Module, Mir's first building block, was launched in February 1986. The Core Module provides basic services (living quarters, life support, power) and scientific research capabilities. It has two axial docking ports, fore and aft, for Soyuz-TM manned transports and automated Progress-M supply ships, plus four radial berthing ports for expansion modules.
The working compartment is the main habitable volume on Mir and is made up of two concentric cylinders connected by a tapered conical section. The interior of the working compartment is divided into an operations zone and a living area.
Mir crews prefer a spatial orientation of floor and ceiling with the sides arranged in a bottom-to-top orientation despite the formal irrelevance of the terms in the absence of gravity. The floor of the operations area is covered with dark green carpet, the walls are light green and the ceiling is white with fluorescent lamps.
The arrangement of equipment and the interior finish of the working compartment are designed to reinforce this bottom-to-top orientation. The living area uses the same spatial orientation concepts, but soft pastel colors are used to imply a home-like atmosphere.
The living area of the working compartment provides the necessities for long-term human missions. The living area contains a galley area with a table, cooking elements, and trash storage. Individual crew cabins, which include a porthole, hinged chairs and a sleeping bag are found next as one moves axially through the working compartment. The aft end of the working compartment contains the personal hygiene area with toilet, sink, and shower.
#
"Mir Station - Priroda",53,0,0,0
On April 23, 1996, \JRussia\j launched the Mir module Priroda (Nature) on a Proton launcher for rendezvous and docking with the space station on April 26. Weighing nearly 20 tons, the unit carries more than a ton of U.S. cargo for \Jastronaut\j Shannon Lucid aboard the space station. Other Priroda equipment includes optical systems to survey Earth's resources.
When docked, the new module completes the Mir construction complex started ten years earlier; four other modules -- Kristall, Kvant, Kvant-2, and Spektr -- have been launched and attached to the core unit before. Unlike them, Priroda has no solar power arrays but must rely on its on-board batteries as long as it is not docked to Mir.
ò Modular Optoelectronic Multispectral Scanner (MOMS-2P) experiment aboard Priroda
Priroda was the last module to be added to the Mir. After its launch from Baikonur on April 23, 1996, it docked to the space station as scheduled on April 26. Its primary purpose is to add Earth remote sensing capability to Mir. It also contains the hardware and supplies for several joint U.S.-Russian science experiments.
Its Earth remote sensing capabilities include:
ò monitoring the ecological situation of large industrial areas, estimation of anthropogenous effects on ecological systems
ò measuring concentration and spacial distribution of small gaseous components in atmosphere of ozone and anthropogenous impurities
ò determining temperature fields on the ocean surface and researching the process of energy and mass exchange between ocean and atmosphere affecting the weather
ò receiving data on classification, structure, and moisture of clouds, including their optical characteristics
ò receiving data for plotting geological structure maps on refinement of mineral reserves, water reserves, erosion of soil and conditions of forests and crops
ò acquiring emergency information from buoys in areas of nuclear power stations, seismically dangerous and other zones to create an integrated monitoring and warning system (Kentavr)
ò performing measurements in order to obtain data for working out ecological and economic theory of natural resources utilization
#
"Mir Station - Spektr",54,0,0,0
Launched on a Russian Proton rocket from the Baikonur launch center in central Asia, Spektr was lofted into orbit on May 20, 1995. The module was berthed at the radial port opposite Kvant 2 after Kristall was moved out of the way. Spektr carries four solar arrays and scientific equipment (including more than 1600 pounds of U.S. equipment).
Spektr was badly damaged at 5:18 a.m. EDT on June 25, 1997, when Progress M-34, an unmanned resupply vessel, crashed into the module during tests of the new TORU Progress guidance system. The module lost pressure and electricity and had to be shut completely down and sealed off from the remainder of the Mir complex. It had served as the living quarters for U.S. astronauts aboard Mir. The possibility of repairing the module is under consideration by the Russian Mir team.
The focus of scientific study for this module is Earth observation, specifically natural resources and atmosphere. The equipment onboard is supplied by both Russian and the United States.
Scientific equipment on Spektr includes:
ò Pion, Lira and Buton equipment for atmospheric research.
ò Faza and Feniks equipment for surface studies.
ò Astra-2 equipment for atmospheric trace constituent monitoring.
ò Taurus and Grif equipment for monitoring Mir's induced x-ray and gamma-ray background.
#
"Mir Station - Soyuz-TM",55,0,0,0
Soyuz-TM is the Russian manned \Jspacecraft\j that ferries Cosmonauts and Astronauts to and from Mir. It also serves as an escape "lifeboat" in the event Mir should experience any life-threatening condition.
#
"Mir Station - Kristall",56,0,0,0
Berthed opposite Kvant 2 in 1990, Kristall weighs 19.6 tons and carries two stowable solar arrays, science and technology equipment, and a docking port equipped with a special androgynous docking mechanism designed to receive heavy (up to about 100 tons) \Jspacecraft\j equipped with the same kind of docking unit.
The androgynous unit was originally developed for the Russian Buran Shuttle program. Atlantis used the androgynous docking unit on Kristall during mission STS-71.
Added in 1990, Kristall carries scientific equipment, retractable solar arrays, and a docking node equipped with a special androgynous docking mechanism designed to receive \Jspacecraft\j weighing up to 100 tons.
The purpose of the Kristall module is to develop biological and materials production technologies in the space environment. One component of the Kristall is a radial docking port. Originally designed as a potential means of docking the Russian Buran reusable shuttle orbiter, this port is now attached to the Docking Module.
#
"Mir Station - Docking Module",57,0,0,0
The Docking Module was launched in the payload bay of Atlantis and berthed at Kristall's androgynous docking port during the STS-74 mission. The Docking Module will provide clearance for future Shuttle dockings with Mir and will carry two solar arrays -- one Russian and one jointly developed by the U.S. and \JRussia\j -- to augment Mir's power supply.
#
"Mir Station - Kvant-2",58,0,0,0
Berthed at a radial port since 1989, the module weighs 19.6 tons and carries an EVA airlock, two solar arrays, and science and life support equipment.
Kvant 2, added in 1989, carries an EVA airlock, solar arrays, and life support equipment. The 19.6-ton module is based on the transport logistics \Jspacecraft\j originally intended for the Almaz military space station program of the early 1970s.
The purpose of Kvant-2 is to provide biological research data, Earth observation data, and EVA capability. It adds additional system capability to Mir. Kvant-2 includes additional life support system, drinking water, and oxygen provisions, motion control systems, and power distribution, as well as shower and washing facilities.
Kvant-2 is divided into three pressurized compartments: instrumentation/cargo, science instrument, and airlock.
The airlock not only provides EVA capability, but also contains a self-sustained \Jcosmonaut\j maneuvering unit that increases the range and complexity of tasks that can be attempted via EVA. For instance, various construction materials and electronic components can be placed on the outside of the Mir Complex modules via EVA. The effects of space environment exposure on these construction materials can later be investigated.
#
"Mir Station - Kvant",59,0,0,0
Berthed at the core module's aft axial port in 1987, the module weighs 11 tons and carries telescopes and equipment for attitude control and life support.
Kvant was added to the Mir core's aft port in 1987. This small, 11-ton module contains astrophysics instruments, life support and attitude control equipment.
The purpose of the Kvant-1 module is to provide data and observations for research into the physics of active galaxies, quasars, and neutron stars. This data is gathered with devices which measure electromagnetic spectra and x-ray emissions. The Kvant-1 also supports biotechnology experiments in the areas of antiviral preparations and fractions.
The Kvant-1 module is divided into a pressurized laboratory compartment and a nonpressurized equipment compartment. The laboratory compartment is further divided into an instrumentation area and a living area, which are separated by an interior partition. A pressurized transfer chamber connects the Passive Docking Unit with the laboratory chamber. The nonpressurized equipment compartment contains power stabilizers.
#
"Shuttle Missions (Upcoming)",60,0,0,0
(Refer to Table)
#
"Apollo Missions Contents",61,0,0,0
\JApollo Missions, The\j
\JApollo Flights, The\j
\JSkylab Missions\j
\JApollo-Soyuz Mission Outline\j
\JApollo 7\j
\JApollo 8\j
\JApollo 9\j
\JApollo 10\j
\JApollo 11 - 'One Giant Leap'\j
\JApollo 12\j
\JFlight Of Apollo 13, The\j
\JApollo 13 Accident, Chronology of Events\j
\JApollo 14\j
\JApollo 16\j
\JApollo 17 - Farewell\j
\JFlight Of Apollo-Soyuz, The\j
\JApollo-Soyuz\j
#
"Apollo Missions, The",62,0,0,0
The Apollo Program began before the first American was launched into space. In July, 1960, NASA announced that a program to fly Astronauts around the moon would follow the planned Mercury program, but with President Kennedy's famous speech on May 25, 1961, the focus on the Apollo missions shifted to a lunar landing and came into sharper focus with the concrete goal of achieving this before the decade's end.
Many people feel that the Apollo program stands as mankind's greatest technological achievement. In all, six missions landed on the surface of the moon, and three others orbited the moon without landing, including the ill-fated Apollo 13.
The \Jspacecraft\j was in three parts: The conical Command Module where the crew ate and slept on its way to the moon and home; the Service Module, supplying electricity, maneuvering power and thrust to get home from lunar orbit, and water to the \Jspacecraft\j; and the Lunar Module, or LM, a two-part, totally self-contained \Jspacecraft\j that used its own rockets to land on and take off from the surface of the moon, and even served as its own launch pad.
Apollo missions were launched atop two different boosters, the Saturn 1B used for the Earth orbiting missions (including Skylab and Apollo-Soyuz), and the mighty Saturn V, the rocket to the moon.
Apollo started in tragedy, when a fire on the launch pad in the Command Module of Apollo 1 claimed the lives of our second man in space and first Gemini \Jastronaut\j, Virgil I. "Gus" Grissom, our first space walker, Edward White, and rookie \Jastronaut\j Roger Chaffee on January 27, 1967, in a routine training exercise for what had been scheduled to be the first Apollo mission.
A detailed description of the ill-fated Apollo 1 mission is available from KSC, and images are available from JSC's image library.
Although the \Jspacecraft\j had to be modified to prevent any chance of a recurrence, Apollo 7 was readied for flight by October, 1968, after an unoccupied test (named Apollo 4, the first flight of the Saturn V). Following its success, Apollo 8 , the first human flight of the Saturn V, was launched around the moon in December, and by the following July, Apollo 11 actually placed a man on another celestial body and brought him home again.
Twelve men in all walked on the moon before Apollo was done. The last three missions featured the Lunar Rover, which permitted the astronauts to drive about and explore various terrains too rough for the LM to attempt to land upon. On the last Apollo mission to the moon, the astronauts spent 22 hours in moon walks and camped out on the moon for three days total.
Sadly, Apollo 18, 19 and 20 were canceled due to budget limitations. One of these missions had been scheduled to explore the scientifically intriguing crater Aristarchus, where astronomers through the ages had witnessed geological (or, more properly, "lunalogical") activity through their telescopes and wondered whether or not it might be volcanism. We are still wondering.
We have never yet returned to the moon. How sad. The Apollo \Jspacecraft\j was used for four later missions, the three long-duration Skylab missions and the final Apollo flight, the Apollo-Soyuz linkup with the Soviet Soyuz 19.
#
"Apollo Flights, The",63,0,0,0
(Refer to Table)
#
"Skylab Missions",64,0,0,0
(Refer to Table)
#
"Apollo-Soyuz Mission Outline",65,0,0,0
(Refer to Table)
#
"Apollo 7",66,0,0,0
\BLaunched:\b October 11, 1968\b
\BSplashed Down:\b October 22, 1968\b
\BCrew:\b
ò Walter M. Schirra. Jr.
ò Donn F. Eisele
ò Walter Cunningham
Following the tragedy of the January, 1967, Apollo 1 launchpad fire, the success of this mission was a badly needed confidence builder. The Command Module had been extensively redesigned to eliminate the risk of another conflagration, and the program had fallen behind schedule.
This mission was primarily intended as a shakedown of the new Command Module and the Service Module, including the Service Propulsion System (SPS) that would be relied upon to place Apollo into and out of Lunar orbit.
The mission was a roaring success. Because it was launched without the Lunar Module, it was able to be placed into orbit by the smaller, more extensively tested Saturn 1B rather than the magnificent but still brand-new Saturn V. The SPS performed almost flawlessly on eight separate firings, and the larger and more commodious cabin was significantly more comfortable than the cramped quarters of the Gemini flights.
The eleven days in orbit, though, took their toll: The food was bad, and all three astronauts caught colds. But Apollo 7 demonstrated the spaceworthiness of the basic Apollo vehicle, and truly great deeds were very near in its future.
On this flight, Wally Schirra became the only \Jastronaut\j to fly on Mercury, Gemini and Apollo missions, and will forever remain the one and only man with that honor.
#
"Apollo 8",67,0,0,0
\BLaunched:\b December 21, 1968\b
\BSplashed Down:\b December 27, 1968\b
\BCrew:\b
ò Frank Borman
ò James A. Lovell, Jr.
ò William A. Anders
For the first time in human existence, man finally broke the bounds of the earth when the three Apollo 8 astronauts took man's first trip to the moon. This flight was initially planned as another earth orbiting checkout of the Apollo hardware, but rumors that the Soviets were plotting to beat us into orbit around the moon caused a little last-minute change in plans.
This was the first human spaceflight atop the wondrous Saturn V launch vehicle, which had flown only twice before. There could be no thought of a lunar landing; like Apollo 7, no Luner Module had been launched with them.
The booster worked flawlessly, as did the Service Propulsion Module. Who can forget Christmas Eve, when the astronauts passed behind the moon, and fired their engines while outside communications range?
The whole nation, the whole world, sat and waited on the edge of our seats, waiting for the craft to come out from behind the moon. Did the engines fire properly? Could they come home, or would they remain, forever trapped in the icy cold of space?
It all worked, and Apollo 8 proved at last that our horizons are truly limitless. Man had, at last, seen the Earth rise above the horizon of another world, and brought back the pictures to prove it!
#
"Apollo 9",68,0,0,0
\BLaunched:\b March 3, 1969\b
\BSplashed Down:\b March 13, 1969\b
\BCrew:\b
ò James A. McDivitt
ò David R. Scott
ò Russell L. Schweickart
This mission was the first flight test of the Lunar Module, the third critical piece of Apollo hardware. During ten days in earth orbit, the crew undocked, maneuvered and docked the LM and the Command Modules, simulating as closely as possible the conditions that would be encountered when men finally would land on the moon itself.
This mission also tested the Apollo spacesuit, the first to carry self-contained life support rather than being dependent on an umbilical connection.
This was the first Apollo mission where the astronauts were granted the honor, often exercised rather whimsically, of naming their \Jspacecraft\j. The Apollo 9 astronauts named the Command Module "Gumdrop" and the Lunar Module "Spider." Oh, well, they can't all be Eagles, now, can they?.
After Apollo 8, Apollo 9 was a bit of an anticlimax, but it was another absolutely necessary step that we needed to take if we were going to place a man onto the moon.
#
"Apollo 10",69,0,0,0
\BLaunched:\b May 18, 1969\b
\BSplashed Down:\b May 26, 1969\b
\BCrew:\b
ò Thomas P. Stafford
ò John W. Young
ò Eugene A. Cernan
This was the "dress rehearsal," the penultimate in space exploration. Apollo 10 entered actual orbit around the moon, separated from the Lunar Module, and the LM (named "Snoopy") with Stafford and Cernan aboard descended to within nine miles of the lunar surface before returning and redocking with the waiting John Young in the Command Module (named "Charlie Brown").
The astronauts tested the LM's radar and ascent engine and surveyed Apollo 11's eventual landing site, the Sea of Tranquility. This mission also served as a test of the extensive new Apollo tracking and control network on earth.
Another first for this mission: For the first time, live color TV pictures were broadcast into our homes from space. What a country!
#
"Apollo 11 - 'One Giant Leap'",70,0,0,0
\BLaunched:\b July 16, 1969\b
\Blanded:\b July 20, 1969, Sea of Tranquility\b
\BSplashed Down:\b July 24, 1969\b
\BCrew:\b
ò Neil A. Armstrong
ò Michael Collins
ò Edwin E. "Buzz" Aldrin, Jr.
On July 20, 1969, the human race accomplished its single greatest technological achievement of all time when a man first set foot on another celestial body. We entered a new era, no longer bound by the circles of the earth that had held us so jealously so close to its surface for so long.
On that day we evolved from lowly, ape-like homo sapiens to homo universalis, Man of the Universe, through the power of our minds and the strength of our indomitable will.
Six hours after landing at 4:17 p.m. Eastern Standard Time (with less than 30 seconds of fuel remaining), Neil A. Armstrong took the "Small Step" into our greater future when he stepped off the Lunar Module, named "Eagle", onto the surface of the moon, from which he could look up and see the earth in the heavens as no man had done before him.
He was shortly joined by "Buzz" Aldrin, and the two astronauts spent 21 hours on the lunar surface and returned 46 pounds of lunar rocks. Their liftoff from the surface of the moon was (partially) captured on a TV camera they left behind, and they successfully docked with Michael Collins, patiently orbiting the cold but no longer lifeless moon alone in the Command Module "Columbia."
The moon walkers left behind a plaque on the lunar surface that read:
"Here Men From \JPlanet\j Earth First Set Foot Upon The Moon. July 1969 A.D. We Came In Peace For All Mankind."
Since that day the world's wars have all wound down, implacable foes have become allies, and we have known the longest period of near universal peace and prosperity in recorded history. Coincidence? Perhaps. But then perhaps we all could now see, along with the astronauts, the world as it truly exists, nestled in the heavens, without borders or boundaries, and with wealth beyond the reckoning of kings. And perhaps we became more cognizant of the value of life beyond the need for such petty strife.
In any event, we all grew that day, that fateful day, that greatest of days, back in 1969.
#
"Apollo 12",71,0,0,0
\BLaunched:\b November 14, 1969\b
\BLanded:\b November 18, 1969, Ocean of Storms\b
\BSplashed Down:\b November 24, 1969\b
\BCrew:\b
ò Charles "Pete" Conrad, Jr.
ò Richard F. Gordon
ò Alan L. Bean
This was an exercise of precision targeting. The Lunar Module Intrepid was brought to the surface of the moon automatically by radar and computer, needing only a few manual corrections by the pilot Pete Conrad. It landed only 183 meters from its target in the Ocean of Storms, where the old Surveyor 3 robot \Jspacecraft\j had soft-landed on the surface of the moon back in 1967.
Conrad and Bean brought back pieces of the old Surveyor and took two moon walks, each lasting almost four hours, during which they set up seismic and magnetism experiments and some experiments to test the effects of the solar wind. When the crew, back aboard the Command Module "Yankee Clipper", launched the LM's ascent stage into the moon, the seismometers they had left behind recorded the vibrations of its impact for over an hour.
They had such a good time that they stayed an extra day in lunar orbit taking pictures.
#
"Flight Of Apollo 13, The",72,0,0,0
\BMission Data:\b
ò Pad 39-A (7)
ò Saturn-V AS-508 ()
ò High Bay 1
ò MLP 3
ò Firing Room 1
\BFlight Crew:\b
ò James A. Lovell, Jr.
ò John L. Swigert, Jr.
ò Fred W. Haise, Jr.
\BBackup Crew:\b
ò John W. Young
ò Thomas K. Mattingly II
ò Charles M. Duke, Jr.
\BMilestones:\b
ò 06/13/69 - S-IVB ondock at Kennedy Space Center (KSC)
ò 06/16/69 - S-1C Stage ondock at KSC
ò 06/29/69 - S-II Stage ondock at KSC
ò 07/07/69 - S-IU ondock at KSC
ò 04/11/70 - Launch
\BPayload:\b
ò Odyssey (CM-109) Block II Command and Service Module
ò Aquarius (LM-7) Lunar Module
\BLaunch:\b
ò April 11, 1970 , 19:13:00 Greenwich Mean Time (GMT)
\BEarth Orbit:\b
ò Translunar injection performed over Pacific Ocean before completion of second revolution
\BMission Duration:\b
ò 05 Days, 22 hours, 54 minutes
\BLanding:\b
ò April 17, 1970
\BMission Highlights:\b
Third lunar landing attempt. Mission aborted after rupture of service module oxygen tank. Classed as "successful failure" because of experience in rescuing crew. Spent upper stage successfully impacted on the Moon.
\BMission Narrative:\b
Apollo 13, the third human lunar landing and exploration mission, had been tentatively scheduled in July 1969 for launch in March 1970, but by the end of the year the launch date had been shifted to April. In August 1969 crew assignments for Apollo 13 were announced, eventually James A. Lovell commanded the mission, with Fred Haise and John Swigert.
The target for the mission was the Fra Mauro Formation, a site of major interest to scientists, specifically a spot just north of the crater Fra Mauro, some 550 kilometers (340 miles) west-southwest of the center of the Moon's near side.
On March 24, 1970, during the countdown demonstration test for Apollo 13, Kennedy Space Center test engineers encountered a problem with an oxygen tank in the service module. The \Jspacecraft\j carried two such tanks, each holding 320 pounds (145 kilograms) of supercritical oxygen. They provided the oxygen for the command module atmosphere and (along with two tanks of hydrogen) three fuel cells, which were the \Jspacecraft\j's primary source of electrical power.
Besides power, the chemical reaction in the cells produced water, which not only supplied the crew's drinking water but was circulated through cooling plates to remove heat from certain critical electronic components. The tanks were designed to operate at pressures of 865 to 935 pounds per square inch (psi) (6,000 to 6,450 kilopascals) and temperatures between-340 degrees F and +80 degrees F (-207 degrees C to +27 degrees C). Inside each spherical tank were a quantity gauge, a thermostatically controlled heating element, and two stirring fans driven by electric motors.
The fans were occasionally operated to homogenize the fluid in the tank; it tended to stratify, leading to erroneous quantity readings. All wiring inside the tank was insulated with Teflon, a fluorocarbon plastic that is ordinarily noncombustible.
Each tank was fitted with a relief valve designed to open when the pressure rose above 1,000 psi (6,900 kilopascals); the tanks themselves would rupture at pressures above 2,200 psi (15,169 kilopascals). Both tanks were mounted on a shelf in the service module between the fuel cells and the \Jhydrogen\j tanks.
The countdown demonstration test called for the tanks to be filled, tested, and then partially emptied by applying pressure to the vent line, thus forcing oxygen out through the fill line. Number one tank behaved normally in this test, but number two released only 8 percent of its contents, not 50 percent as required.
Test engineers decided to proceed with the rest of the test and investigate the problem later. The next day, after KSC engineers had discussed the problem with colleagues at the Manned \JSpacecraft\j Center (MSC, the Johnson Space Center after 1973), North American Rockwell (builders of the service module), and Beech \JAircraft\j (manufacturers of the oxygen tanks), they tried emptying the tank again, with no success.
Further talks led to the conclusion that the tank probably contained a loose-fitting fill tube, which could allow pressure to escape without emptying the tank.
When normal procedures again failed to empty the tank, engineers decided to use its internal heaters to boil off the contents and applied direct-current power at 65 volts to the heaters. This was successful but slow, requiring eight hours of heating. It was then decided that if the tank could then be filled normally it would not cause a problem in flight. A third test gave the same result as the second, requiring heating to empty the tank.
In view of the difficulty of replacing the oxygen shelf a job that would take at least 45 hours and the possibility that other components might be damaged in the process and the launch delayed for a month, NASA and contractor officials decided not to replace the tanks.
The \Jspacecraft\j was launched on April 11, 1970, and the mission was quite routine for the first two days. At 30 hours and 40 minutes after launch (30:40 ground elapsed time), the crew ignited their main engine to put the \Jspacecraft\j on a hybrid trajectory, a flight path that saved fuel in reaching the desired lunar landing point.
At 46:40 the crew routinely switched on the fans in the oxygen tanks briefly. A few seconds later the quantity indicator for tank number two went off the high end of the scale, where it stayed. The tanks were stirred twice more during the next few hours; and at 55:53, after a master alarm had indicated low pressure in a \Jhydrogen\j tank, the Mission Control Center (MCC) directed the crew to switch on all tank stirrers and heaters.
Shortly thereafter the crew heard a loud banging and felt unusual vibrations in the \Jspacecraft\j. Mission controllers noticed that all telemetry readings from the \Jspacecraft\j dropped out for 1.8 seconds. In the command module (CM), the caution and warning system alerted the crew to low voltage on d.c. main bus B, one of two power distribution systems in the \Jspacecraft\j. At this point command module pilot Jack Swigert told Houston, "Hey, we've had a problem here."
Because of the interruption of telemetry that had just occurred, flight controllers in the MCC had difficulty for the next few minutes determining whether they were getting true readings from the \Jspacecraft\j sensors or whether the sensors had somehow lost power.
Before long, however, both MCC and the crew realized that oxygen tank number two had lost all of its contents, oxygen tank number one was slowly losing its contents, and the CM would soon be out of oxygen and without electrical power. Among the first actions taken were shutting down one fuel cell and switching off nonessential systems in the CM to minimize power consumption; shortly after, the second fuel cell was shut down as well. When the remaining oxygen ran out, the CM would be dead; its only other power source was three reentry batteries providing 120 ampere-hours, and these had to be reserved for the critical reentry period.
An hour and a half after the "bang," MCC notified the crew that "we're starting to think about the lifeboat" using the lunar module (LM) and its limited supplies to sustain the crew for the rest of the mission. Plans for such a contingency had been studied for several years, although none had anticipated a situation as grave as that of Apollo 13. Many of these studies were retrieved and their results were adapted to the situation as it developed.
Shortly after the accident, mission commander James Lovell reported seeing a swarm of particles surrounding the \Jspacecraft\j, which meant trouble. Particles could easily be confused with stars, and the sole means of determining the \Jspacecraft\j's attitude was by locating certain key stars in the onboard \Jsextant\j. Navigational sightings from the lunar module (LM) were difficult in any case as long as it was attached to the command module, and this would only complicate matters.
Flight controllers decided to align the lunar module's guidance system with that in the command module while the CM still had power. That done, the last fuel cell and all systems in the command module were shut down, and the crew moved into the lunar module. Their survival depended on this craft's oxygen and water supplies, guidance system, and descent propulsion engine (DPS).
Normally all course corrections were made using the service propulsion system (SPS) on the service module, but flight controllers ruled out using it, partly because it required more electrical power than was available and partly because no one knew whether the service module had been structurally weakened by the explosion. If it had, an SPS burn might be dangerous. The DPS would have to serve in its place.
When word got out that Apollo 13 was in trouble, off-duty flight controllers and \Jspacecraft\j systems experts began to gather at MSC, to be available if needed. Others stood by at NASA centers and contractor plants around the country, in touch with Houston by \Jtelephone\j. Flight directors Eugene Kranz, Glynn Lunney, and Gerald Griffin soon had a large pool of talent to help them solve problems as they arose, provide information that might not be at their fingertips, and work on solutions to problems they could anticipate further along in the mission.
Astronauts occupied the CM and LM training simulators at Houston and at Kennedy Space Center, testing new procedures as they were devised and modifying them as necessary. MSC director Robert R. Gilruth, Dale D. Myers, director of manned space flight, and NASA administrator Thomas O. Paine were all on hand at Mission Control to provide high-level authority for changes.
Soon after the explosion, the assessment of life-support systems determined that although oxygen supplies were adequate, the system for removing carbon dioxide (CO2) in the lunar module was not. The system used canisters filled with \Jlithium\j \Jhydroxide\j to absorb CO2 as did the system in the command module.
Unfortunately the canisters were not interchangeable between the two systems, so the astronauts were faced with plenty of capacity for removing CO2 but no way of using it. A team in Houston immediately set about improvising a way to use the CM canisters, using materials available in the \Jspacecraft\j.
Flight controllers, meanwhile, were addressing operational problems. Their first critical decision was to put the crippled \Jspacecraft\j back on a free-return trajectory, which was accomplished by firing the LM descent engine at 61:30.
Mission Control then had some 18 hours to consider the remaining problems; the next was a possible adjustment to change the \Jspacecraft\j's landing point on Earth. If this was to be done, it was scheduled for "PC + 2" two hours after pericynthion (closest approach to the Moon), after the \Jspacecraft\j emerged from behind the Moon. In the interval, Houston worked out a new flight plan that would minimize the consumption of oxygen, water, and electricity while keeping vital systems operating.
The alternatives for the PC +2 maneuver were worked out by about 64 hours into the flight. A major consideration was the total time to splashdown. Left on its free-return course the command module would return at about 155 hours after launch to a landing in the Indian Ocean. Three options would bring it back in the mid-Pacific and could reduce the total mission time to as little as 118 hours.
The fifth possibility returned the \Jspacecraft\j in 133 hours, but to the South Atlantic.
For one reason or another, all but one of these choices were discarded. The free-return (no course correction) choice was abandoned, since there was no known reason not to use the LM descent propulsion system. Recovery in either the Atlantic or the Indian Ocean was far from ideal; the main recovery force was deployed in the mid-Pacific and there was not enough time to move it or to make adequate arrangements elsewhere.
Two options giving the shortest return time (118 hours) had other drawbacks. Both would require using virtually all of the available propellant, and it was not prudent to assume that no additional course corrections would be required. One of them involved jettisoning the service module, which would expose the CM heat shield to the cold of space for 40 hours and raise questions about its integrity on reentry.
After five and a half hours of weighing the choices and their consequences, flight directors met with NASA and contractor officials and presented their findings and recommendations. The decision, made some ten hours before the scheduled engine burn, was to go for mid-Pacific recovery at 143 hours.
During all of these deliberations the atmosphere in the lunar module was gradually accumulating carbon dioxide as the absorbers in the environmental control system became saturated. Members of MSC's Crew Systems Division devised a makeshift air purifier by taping a plastic bag around one end of a CM \Jlithium\j \Jhydroxide\j cartridge and attaching a hose from the portable life-support system, allowing air from the cabin to be circulated through it.
After verifying that this jury rig would function, they prepared detailed instructions for building it from materials available in the \Jspacecraft\j and read them up to the crew. For the rest of the mission the improvised system kept the CO2 content of the atmosphere well below hazardous levels.
The decision to recover in the Pacific fixed the time line for the remainder of the mission and imposed some rigid constraints on preparations for reentry. The final course correction had to be made with the LM engine; command module systems had to be turned on and the guidance system aligned; the service module had to be discarded; and when all preparations had been made, the lunar module would be cut loose.
In all these preparations the power available from the CM's reentry batteries was a limiting factor. From the PC + 2 burn until about 35 hours before reentry the sequence of activation of CM systems was worked out, checked in the simulators, and modified. Fifteen hours before beginning reentry the revised sequence of activities was read to the crew, to give them time to review and practice it.
The husbanding of expendable resources, particularly electrical power, paid off on the morning of landing, when it was discovered that power reserves in the LM were adequate to allow use of it in the CM. Some of the early CM activities could then be done at a less hurried pace.
The Apollo 13 command module splashed down within a mile of the recovery carrier with about 20 percent of its battery power remaining. Three weary, chilled astronauts came aboard the U.S.S. Iwo Jima on April 17 and were flown to Hawaii for an emotional reunion with their families.
Mission Control teams and their hundreds of helpers were no less drained. The usual cigars were lit up after recovery, but the splashdown parties that evening were subdued: most of those who went quit early and went home to bed. Their efforts were recognized the next day when President Richard M. Nixon, on his way to Hawaii, stopped in Houston to present the Presidential Medal of Freedom, the nation's highest civilian award, to the entire team.
NASA immediately convened an investigation board to determine the cause of the accident and postponed Apollo 14 until its results were in. Lacking the \Jspacecraft\j itself, the service module had been jettisoned before reentry, and the crew had been able to take only a few rather poor photographs of it, the board initially had only the data from inflight telemetry to work with.
When it became clear that the fault lay in oxygen tank number two, the board carefully reviewed its entire history, from fabrication to launch, as recorded in the detailed documentation that followed every piece of equipment from plant to launch pad. Under the board's direction, MSC and other NASA centers conducted tests under simulated mission conditions to verify its findings. The investigation, which concluded in a few weeks, turned up a highly improbable sequence of human error and oversight that led inexorably to the failure in flight.
Board Chairman Edgar M. Cortright, director of Langley Research Center, explained the board's findings to congressional committees in June 1971. The accident, he reported, was not a random malfunction but resulted from an unusual combination of mistakes as well as "a somewhat deficient and unforgiving design."
As the board's report reconstructed the events leading up to the accident, the tank left Beech \JAircraft\j's plant on May 3, 1967, after passing all acceptance tests. It was installed as part of a shelf assembly in service module no. 106 on June 4, 1963, having passed all tests conducted at North American Rockwell during assembly. Design changes in the service module, however, necessitated removing the entire shelf from SM 106 for modification.
During removal, which was accomplished by use of a special fixture that fit under the shelf to lift it upward, workmen overlooked one bolt that held down the back of the shelf, with the result that the removal fixture broke, dropping the shelf two inches. The board concluded that this incident might have jarred loose a poorly fitting fill tube. Subsequent tests did not detect any flaws, and after modification the shelf was shipped to Kennedy Space Center for installation in SM 109, the Apollo 13 \Jspacecraft\j.
What was not known was that this oxygen tank was fitted with obsolete thermostatic switches protecting its heating elements. Original specifications for the switches called for operation on 28 volts d.c.; in 1965 this was changed to 65 volts d.c. to match the test and checkout equipment at the Cape. Later tanks conformed to the new specifications, but this one, which should have been modified, was not, and the discrepancy was overlooked at all stages thereafter.
In a normal checkout of a normal tank, this would not have mattered, because the switches would not have opened during normal operation. But the improvised procedure used when this tank failed to empty (the result of a loose fitting, as noted above) raised the temperature in the tank above 80 degrees F (27 degrees C), at which point the switches opened.
Tests conducted during the investigation showed that the higher current produced by the 65-volt power source caused an arc between the contact points as they separated, welding them together and preventing their opening when the temperature dropped. This went undetected during the detanking procedure at the Cape; it could have been noticed if anyone had monitored the heater current, which would have shown that the heaters were operating when they should not have been.
But all attention was on the specific malfunction, and no one was aware that the heaters were on continuously for eight hours on two separate occasions. The result, as tests showed, was that the heater tube reached 1,000 degrees F (538 degrees C) in spots, damaging the Teflon \Jinsulation\j on the adjacent fan-motor wiring and exposing bare wire. From that point on, the board concluded, the tank was hazardous when filled with oxygen and electrically powered. Teflon can be ignited at a high enough temperature in the presence of pure oxygen, and the tank contained small amounts of other combustibles as well.
Unfortunately for Apollo 13, the tank functioned normally for the first 56 hours of the mission, when the heaters and the fans were energized during routine operations. At that point an arc from a short circuit probably ignited the Teflon, and the rapid pressure rise that followed either ruptured the tank or damaged the conduit carrying wiring into the tank, expelling high-pressure oxygen.
The board could not determine exactly how the tank failed or whether additional combustion occurred outside the tank, but the pressure increase blew off the panel covering that sector of the service module and damaged the directional antenna, causing the interruption of telemetry observed in Houston. It also evidently damaged the oxygen distribution system, or the other oxygen tank, as well, leading to the loss of all oxygen supplies and aborting the mission.
The board pointed out that although the circumstances of the tank failure were highly unusual and that the system had worked flawlessly on six successful missions, Apollo 13 was a failure whose causes had to be eliminated as completely as possible. It recommended that the oxygen tanks be modified to remove all combustible material from contact with oxygen and that all test procedures be thoroughly reviewed for adequacy.
Compared to the AS-204 fire in 1967, Apollo 13 was only a frightening near-miss, and because its cause was localized and comparatively easy to discover, it had fewer adverse effects on the program. Only the skill and dedication of hundreds of members of the often-celebrated "manned space flight team" saved it, however, and the accident served to remind NASA and the public that human flight in space, no matter how commonplace it seemed to the casual observer, was not a routine operation.
#
"Apollo 13 Accident, Chronology of Events",73,0,0,0
The following includes events from 2.5 minutes before the accident to about 5 minutes after. Times given are in Ground Elapsed Time (G.E.T.), that is, the time elapsed since liftoff of Apollo 13 on April 11, 1970, at 2:13 PM Eastern Standard Time (EST). 55:52:00 G.E.T. is equal to 10:05 PM EST on April 13, 1970.
55:52:31 - Master caution and warning triggered by low \Jhydrogen\j pressure in tank no. 1
55:52:58 - CapCom (Charlie Duke): "13, we've got one more item for you, when you get a chance. We'd like you to stir up the cryo tanks. In addition, I have shaft and trunnion .....
55:53:06 - Swigert: "Okay."
55:53:07 - CapCom: ".... for looking at \JComet\j Bennett, if you need it."
55:53:12 - Swigert: "Okay. Stand by."
55:53:18 - Oxygen tank No. 1 fans on.
55:53:19 - Oxygen tank No. 2 pressure decreases 8 psi.
55:53:20 - Oxygen tank No. 2 fans turned on.
55:53:20 - Stabilization control system electrical disturbance indicates a power transient.
55:53:21 - Oxygen tank No. 2 pressure decreases 4 psi.
55:53:22.718 - Stabilization control system electrical disturbance indicates a power transient.
55:53:22.757 - 1.2 Volt decrease in ac bus 2 voltage.
55:53:22.772 - 11.1 \Jamp\j rise in fuel cell 3 current for one sample.
55:53:26 - Oxygen tank No. 2 pressure begins rise lasting for 24 seconds.
55:53:38.057 - 11 volt decrease in ac bus 2 voltage for one sample.
55:53:38.085 - Stabilization control system electrical disturbance indicates a power transient.
55:53:41.172 - 22.9 \Jamp\j rise in fuel cell 3 current for one sample
55:53:41.192 - Stabilization control system electrical disturbance indicates a power transient.
55:54:00 - Oxygen tank No. 2 pressure rise ends at a pressure of 953.8 psia.
55:54:15 - Oxygen tank No. 2 pressure begins to rise.
55:54:30 - Oxygen tank No. 2 quantity drops from full scale for 2 seconds and then reads 75.3 percent.
55:54:31 - Oxygen tank No. 2 temperature begins to rise rapidly.
55:54:43 - Flow rate of oxygen to all three fuel cells begins to decrease.
55:54:45 - Oxygen tank No. 2 pressure reaches maximum value of 1008.3 psia.
55:54:51 - Oxygen tank No. 2 quantity jumps to off-scale high and then begins to drop until the time of telemetry loss, indicating failed sensor.
55:54:52 - Oxygen tank No. 2 temperature sensor reads -151.3 F.
55:54:52.703 - Oxygen tank No. 2 temperature suddenly goes off-scale low, indicating failed sensor.
55:54:52.763 - Last telemetered pressure from oxygen tank No. 2 before telemetry loss is 995.7 psia.
55:54:53.182 - Sudden accelerometer activity on X, Y, Z axes.
55:54:53.220 - Stabilization control system rate changes begin.
55:54:53.323 - Oxygen tank No. 1 pressure drops 4.2 psi.
55:54:53.500 - 2.8 \Jamp\j rise in total fuel cell current.
55:54:53.542 - X, Y, and Z accelerations in CM indicate 1.17g, 0.65g, and 0.65g.
55:54:53.555 - Master caution and warning triggered by DC main bus B undervoltage. Alarm is turned off in 6 seconds. All indications are that the cryogenic oxygen tank No. 2 lost pressure in this time period and the panel separated.
55:54:54.741 - \JNitrogen\j pressure in fuel cell 1 is off-scale low indicating failed sensor.
55:54:55.350 - Telemetry recovered.
55:54:56 - Service propulsion system engine valve body temperature begins a rise of 1.65 F in 7 seconds. DC main A decreases 0.9 volts to 28.5 volts and DC main bus B 0.9 volts to 29.0 volts. Total fuel cell current is 15 amps higher than the final value before telemetry loss. High current continues for 19 seconds. Oxygen tank No. 2 temperature reads off-scale high after telemetry recovery, probably indicating failed sensors. Oxygen tank No. 2 pressure reads off-scale low following telemetry recovery, indicating a broken supply line, a tank pressure below 19 psi, or a failed sensor. Oxygen tank No. 1 pressure reads 781.9 psia and begins to drop.
55:54:57 - Oxygen tank No. 2 quantity reads off-scale high following telemetry recovery indicating failed sensor.
55:55:01 - Oxygen flow rates to fuel cells 1 and 3 approached zero after decreasing for 7 seconds.
55:55:02 - The surface temperature of the service module oxidizer tank in bay 3 begins a 3.8 F increase in a 15 second period. The service propulsion system \Jhelium\j tank temperature begins a 3.8 F increase in a 32 second period.
55:55:09 - DC main bus A voltage recovers to 29.0 volts, DC main bus B recovers to 28.8.
55:55:20 - Swigert: "Okay, Houston, we've had a problem here."
55:55:28 - Duke: "This is Houston. Say again please."
55:55:35 - Lovell: "Houston, we've had a problem. We've had a main B bus undervolt."
55:55:42 - Duke: "Roger. Main B undervolt."
55:55:49 - Oxygen tank No. 2 temperature begins steady drop lasting 59 seconds indicating a failed sensor.
55:56:10 - Haise: "Okay. Right now, Houston, the voltage is--is looking good. And we had a pretty large bang associated with the caution and warning there. And as I recall, main B was the one that had an \Jamp\j spike on it once before.
55:56:30 - Duke: "Roger, Fred."
55:56:38 - Oxygen tank No. 2 quantity becomes erratic for 69 seconds before assuming an off-scale low state, indicating a failed sensor.
55:56:54 - Haise: "In the interim here, we're starting to go ahead and button up the tunnel again."
55:57:04 - Haise: "That jolt must have rocked the sensor on -- see now -- oxygen quantity 2. It was oscillating down around 20 to 60 percent. Now it's full-scale high."
55:57:39 - Master caution and warning triggered by DC main bus B undervoltage. Alarm is turned off in 6 seconds.
55:57:40 - DC main bus B drops below 26.25 volts and continues to fall rapidly.
55:57:44 - Lovell: "Okay. And we're looking at our service module RCS \Jhelium\j 1. We have -- B is barber poled and D is barber poled, \Jhelium\j 2, D is barber pole, and secondary propellants, I have A and C barber pole." AC bus fails within 2 seconds.
55:57:45 - Fuel cell 3 fails.
55:57:59 - Fuel cell current begins to decrease.
55:58:02 - Master caution and warning caused by AC bus 2 being reset.
55:58:06 - Master caution and warning triggered by DC main bus undervoltage.
55:58:07 - DC main bus A drops below 26.25 volts and in the next few seconds levels off at 25.5 volts.
55:58:07 - Haise: "AC 2 is showing zip."
55:58:25 - Haise: "Yes, we got a main bus A undervolt now, too, showing. It's reading about 25 and a half. Main B is reading zip right now."
56:00:06 - Master caution and warning triggered by high \Jhydrogen\j flow rate to fuel cell 2.
#
"Apollo 14",74,0,0,0
\BLaunched:\b January 31, 1971\b
\BLanded:\b February 3, 1971, Fra Mauro Region\b
\BSplashed Down:\b February 9, 1971\b
\BCrew:\b
ò Alan B. Shepard, Jr.
ò Stuart A. Roosa
ò Edgar D. Mitchell
Following the frightening problems of Apollo 13, almost ten months elapsed before we returned to the moon with Apollo 14. This flight marked the return to space of America's first spaceman, Alan B. Shepard, who had first flown aboard Freedom 7 a decade earlier.
Shepard and Mitchell had to scrap a planned rock-collecting trip to the 1,000 foot wide Cone Crater when they became disoriented and almost got lost. Interestingly, it was later discovered that they were only a little over 30 yards from the crater's rim when they gave up the search.
After their return aboard the Command Module "Kitty Hawk,", the three Apollo 14 astronauts became the last to be required to undergo a period of quarantine.
#
"Apollo 16",75,0,0,0
\BLaunched:\b April 16, 1972\b
\BLanded:\b April 20, 1972, Descartes Highlands\b
\BSplashed Down:\b April 27, 1972\b
\BCrew:\b
ò John W. Young
ò Thomas K. Mattingly II
ò Charles M. Duke, Jr.
A malfunction in the main propulsion system of the Lunar Module "Orion" almost scrubbed the landing, but after its success Young and Duke spent three days exploring the geologically interesting Descartes Highlands region while Mattingly circled overhead in the Command Module "Casper."
It was thought that Descartes may be an area of active volcanism, but this proved not to be the case. Among other specimens, the astronauts returned the largest moon rock ever, a 23-pound chunk that turned out to contain not even a single gram of green cheese.
During this flight the moon racers also had a bit of fun testing out the capabilities of the Lunar Rover, at one point getting it up to almost 11 miles per hour!
#
"Apollo 17 - Farewell",76,0,0,0
\BLaunched:\b December 7, 1972\b
\BLanded:\b December 11, 1972, Taurus-Littrow Valley\b
\BSplashed Down:\b December 19, 1972\b
\BCrew:\b
ò Eugene A. Cernan
ò Ronald E. Evans
ò Harrison H. "Jack" Schmitt
Schmitt and Cernan left behind the Lunar Module "Challenger" and drove around almost 34 kilometers of lunar ground during this, the last mission to the moon. Schmitt was the first scientist, a geologist by trade, to visit the Moon, and became the last man (so far) to set foot on another celestial body.
The astronauts collected a record 108.86 kilograms of lunar rocks, including some that were orange in color, during their three moon walks and drives through the Taurus-Littrow Valley.
The last three scheduled Apollo missions had already been cancelled because of budgetary shortfalls by the time of this flight. The astronauts left behind a plaque that read:
"Here Man completed his first exploration of the Moon, December 1972 A.D. May the spirit of peace in which we came be reflected in the lives of all mankind."
#
"Flight Of Apollo-Soyuz, The",77,0,0,0
\BFlight Crew:\b
ò Apollo:
ò Thomas P. Stafford
ò Vance D. Brand
ò Donald K. Slayton
ò Soyuz:
ò Alexey A. Leonov
ò Valery N. Kubasov
\BLaunch:\b
ò Apollo: July 15, 1975
ò Soyuz: July 15, 1975
\BLanding:\b
ò Apollo: July 24, 1975
ò Soyuz: July 21, 1975
\BMission Duration:\b
ò 09 days, 07 hours, 28 minutes
ò July 15-24, 1975
\BMission Highlights:\b
The Soyuz was launched just over seven hours prior to the launch of the Apollo CSM. Apollo then maneuvered to rendezvous and docking 52 hours after the Soyuz launch. The Apollo and Soyuz crews conducted a variety of experiments over a two-day period. After separation, Apollo remained in space an additional 06 days. Soyuz returned to Earth approximately 30 hours after separation.
\BMission Narrative:\b
The final flight of the Apollo program was the first spaceflight in which \Jspacecraft\j from different nations docked in space. In July 1975, a U.S. Apollo \Jspacecraft\j carrying a crew of three docked with a Russian Soyuz \Jspacecraft\j with its crew of two.
For the Apollo-Soyuz Test Project (ASTP), the United States used an Apollo Command and Service Module (CSM) modified to provide for experiments to be conducted during the mission, extra propellant tanks and the addition of controls and equipment related to the Docking Module. Launch was accomplished with a Saturn IB.
The Docking Module was designed jointly by the United States and Soviet Union, and built in the United States. Its purpose was to enable a docking between the dissimilar Soyuz \Jspacecraft\j and the U.S. Apollo. It was a three meter long cylinder 1.5 meters in diameter, and in addition to serving as a docking device, also served as an airlock module between the different atmospheres of the two ships (the U.S. ship with 100% oxygen at 260 millimeters of mercury; the Soyuz with a mixed oxygen-nitrogen atmosphere at 520 mm HG--lowered from its usual 760 mm Hg for this mission).
Prior to the conduct of ASTP, the astronauts and cosmonauts visited each other's space centers and became familiar with the \Jspacecraft\j of the other country. The first visit was by the Russians to Johnson Space Center in July 1973, followed by a U.S. visit to Moscow in November 1973. In late April and early May 1974, the Russian flight crews returned to Johnson Space Center, and the U.S. crews went to Moscow in June and July 1974.
The Russian crew made a third trip to the United States in September 1973 and came for the fourth and last time in February 1975. The U.S. crew visited the Soviet Union in late April and early May 1975 and became the first Americans to see the Russian launch facilities at Tyuratam on April 28, 1975.
Three simulation sessions were conducted between flight controllers and the ASTP crew in Houston and Moscow on May 13, 15 and 18, 1975 involving communications links between the two control centers, and fully occupied control center facilities. A final simulation was conducted from June 30-July 1, 1975. Additionally, in December 1974, the Russians made a human flight of the modified version of the Soyuz spaceship for system tests (Soyuz 16).
One of the most difficult problems to overcome was that of language differences. To alleviate this problem as much as possible, the Americans learned Russian and the Russians learned English. It was found that the best scenario was for the Russians to speak English and for the Americans to speak Russian.
Soyuz Launch: Soyuz 19, carrying cosmonauts Aleksey A. Leonov and Valery N. Kubasov, was launched into sunny skies from Baykonur Cosmodrome at 5:20 pm local time (8:20 am EDT) July 15, 1975. The \Jspacecraft\j entered orbit with a 221.9-km apogee, 186.3-km perigee, 88.5-min period, and 51.8 inclination.
Foreign correspondents, barred from the launch site, watched the launch on color TV sets in a Moscow press center. The first Soviet launch to be televised live, it was transmitted to viewers throughout the Soviet Union, the U.S., and eastern and western Europe. President Ford watched from a U.S. State Dept. auditorium with Soviet Ambassador to the U.S, Anatoly P. Dobrynin and NASA Administrator James C. Fletcher, before Dr. Fletcher and Ambassador Dobrynin flew to Kennedy Space Center to watch the Apollo launch.
On the third orbit the Soyuz 19 crew established contact with U.S. mission control in Houston, putting into operation the global Moscow and Houston Soyuz-Apollo communications system. On the fifth orbit the cosmonauts made the first of two maneuvers to place Soyuz 19 into a circular docking orbit. New orbital parameters were 231.7-km apogee and 192.4-km perigee. The \Jspacecraft\j was spin-stabilized at 3 per sec with all systems operating normally.
Apollo Launch: At 3:50 pm EDT July 15, 1975, 7 hr, 30 min, after the Soyuz launch-a Saturn IB flawlessly lifted the Apollo \Jspacecraft\j from Kennedy Space Center's launch complex 39, carrying Apollo commander Thomas P. Stafford, command-module pilot Vance D. Brand, and docking-module pilot Donald K. Slayton. The \Jspacecraft\j entered orbit with a 173.3-km apogee, 154.7-km perigee, 87.6-min period, and 51.8 inclination.
The \Jspacecraft\j's launch-vehicle adapter was jettisoned at 9 hr 4 min ground elapsed time (9:04 GET, counted from the Soyuz 19 launch) and the crew maneuvered the Apollo 180 to dock with the adapter and extract the docking module. These events were videotaped and transmitted to earth later via ATS 6 (NASA's Applications Technology Satellite launched 30 May 1974). A maneuver 2 hr later at 7:35 pm circularized the orbit at 172 km. The Saturn S-IVB stage was deorbited into the Pacific Ocean 1 hr 30 min later.
A second Soyuz 19 circularization burn of 18.5 sec at 8:43 am EDT July 16 placed that \Jspacecraft\j in a circular orbit of 229 km, with all systems functioning normally.
Rendezvous and Docking: A series of Apollo maneuvers, with the final braking maneuver at 8:51 am EDT July 17, put the Apollo \Jspacecraft\j in a 229.4-km circular orbit matching the orbit of Soyuz 19. A few minutes later Brand reported, "We've got Soyuz in the sextant." Voice contact was made soon after. Hello. Soyuz, Apollo," Stafford said in Russian. Kubasov replied in English, "Hello everybody. Hi to you, Tom and Deke. Hello there, Vance."
All communications among the five crew members during the mission were made in the language of the listener, with the Americans speaking Russian to the Soviet crew and the Soviet crew speaking English to the Americans. Contact of the two \Jspacecraft\j, 51 hr, 49 min, into the mission (12:09 pm July 17) was transmitted live on TV to the earth, and Stafford commented, "We have succeeded. Everything is excellent." "Soyuz and Apollo are shaking hands now," the cosmonauts answered. Hard docking was completed over the Atlantic Ocean at 12:12 pm, 6 min earlier than the prelaunch flight plan watched by millions of TV viewers worldwide.
"Perfect. Beautiful. Well done, Tom. It was a good show. We're looking forward to shaking hands with you in board [sic] Soyuz," Leonov said. Tass later reported that Kubasov told Moscow ground controllers that "we felt a slight jolt at the moment of docking" but that all went according to plan.
Joint Activities: At 3:17 pm hatch 3 opened; Apollo commander Stafford and Soyuz commander Leonov shook hands 2 min later. "Glad to see you," Stafford told Leonov in Russian. "Glad to see you. Very, very happy to see you," Leonov responded in English. "This is Soyuz and the United States," Slayton told TV viewers around the world. Both Soviet Communist Party General Secretary Leonid I. Brezhnev and President Ford congratulated the crews and expressed their confidence in the success of the mission.
Stafford then presented Leonov with "five flags for your government and the people of the Soviet Union" with the wish that "our joint work in space serves for the benefit of all countries and peoples on the earth." Leonov presented the U.S. crew with Soviet flags and plaques. The men signed international certificates and exchanged other commemorative items. After nearly 4 hrs of joint activities, including a meal aboard the Soyuz, the Americans returned to the Apollo and the hatch was closed at 6:51 pm.
An integrity check of the hatches indicated an atmospheric leak on the Soviet side. Ground controllers later attributed the indication to temperature changes in the sealed docking module that were detected by the sensitive Soviet instrumentation. Future integrity checks of the hatches would be more rigorous, however.
Following a sleep period, the crews prepared for another day of joint activity. Kubasov described the mission to Soviet TV viewers while the rest of the crews performed experiments in their respective \Jspacecraft\j. At 5:05 am July 18, 1975, Brand entered the Soviet \Jspacecraft\j; Leonov joined Stafford and Slayton in Apollo, greeting them with "Howdy partner."
Kubasov gave American TV viewers a tour of his Soyuz, and Stafford followed with a tour of the Apollo. Then both Kubasov and Brand videotaped scientific demonstrations for transmission to earth later. Kubasov and Brand ate lunch in the Soyuz while Leonov ate with Stafford and Slayton in Apollo.
During a third transfer, Stafford and Leonov went into the Soyuz and Kubasov and Brand joined Slayton in Apollo. Brand gave Soviet viewers a Russian-language tour of the eastern U.S. as seen from space. Further speeches and exchanges of commemorative items were made for both U.S. and Soviet viewers before the final handshakes at 4:49 pm EDT July 18, when the crews returned to their respective \Jspacecraft\j.
The hatches were closed after Brand told Leonov and Kubasov, "We wish you the host of success. I'm sure that we've opened up a new era in history. Our next meeting will be on the ground." Total time for all transfers and joint activities was 19 hr 55 min. Stafford had spent 7 hr 10 min aboard Soyuz; Brand, 6 hr 30 min; Slayton, 1 hr 35 min. Leonov spent 5 hr 43 min in the Apollo, Kubasov 4 hr 57 min. During nearly 2 days of joint activities, the five men carried out five joint experiments.
Undocking and Separation: The Apollo and Soyuz \Jspacecraft\j undocked at 95:42 GET (8:02 am EDT July 19, 1975). While the \Jspacecraft\j were in station-keeping mode, the crews photographed them and the docking apparatus, transmitting the pictures live on TV to earth.
The Apollo \Jspacecraft\j then served as an occulting disk, blocking the sun from the Soyuz and simulating a solar eclipse the first man-made eclipse. Leonov and Kubasov photographed the solar corona as the Apollo backed away from the Soyuz and toward the sun.
The two \Jspacecraft\j then redocked at 8:34 am EDT with the Apollo maneuvering and the Soyuz docking system active while good quality TV was transmitted to earth. The second docking was not as smooth as the first because a slight misalignment of the two \Jspacecraft\j caused both to pitch excessively at contact.
Final undocking also with the Soyuz active went smoothly and was completed at 11:26 am. As the \Jspacecraft\j separated, the two crews performed the ultraviolet atmospheric absorption experiment, making unsuccessful data measurements at 150 m and then moving to a distance of 500 and 1,000 m, where data were successfully collected.
The Apollo maneuvered to within 50 m of Soyuz and took intensive still photography of the Soyuz. Separation maneuvers to put the two \Jspacecraft\j on separate trajectories began at 2:42 pm with a reaction-control system burn. With the maneuvers completed, Leonov told the Apollo crew, "Thank you very much for your very big job....It was a very good show." Brand answered, "Thank you, also. This was a very good job."
Soyuz Orbit and Landing: Soyuz 19 remained in orbit nearly 30 hrs after the undocking. The cosmonauts conducted biological experiments with microorganisms and zone-forming fungi. At 2:39 am EDT 21 July the Soyuz crew closed hatch 5 between their orbital vehicle and descent module and began depressurizing the orbital module.
Braking burns of the descent engines began at 6:06 am when the \Jspacecraft\j was 772 km from the Apollo. The 194.9-sec burn slowed the \Jspacecraft\j to 120 km per sec. After another burn to stabilize the \Jspacecraft\j the orbital and descent modules separated over Central Africa.
While Soviet viewers watched the first landing of a Soviet \Jspacecraft\j televised in real time, the main parachute deployed at 7 km and jettisoned before the soft-landing engines fired. Soyuz 19 landed about 11 km from the target point northeast of Baykonur Cosmodrome at 6:51 am EDT July 21, after a 142-hr 31-min mission.
The rescue \Jhelicopter\j approached the capsule immediately and specialists opened hatch 5. Kubasov stepped out waving to rescue-team members, followed by Leonov, both cosmonauts in apparent good health and spirits. The cosmonauts returned to Baykonur for medical checks and debriefings.
Apollo Postdocking Orbital Activities: Apollo remained in orbit while its crew continued U.S science experiments begun during predocking. Searching for extreme ultraviolet radiation, the ASTP crew marked the birth of a new branch of \Jastronomy\j when they found, for the first time, extreme ultraviolet sources outside the solar system; some scientists had believed that such sources could never be found.
One of the newly discovered sources turned out to be the hottest known white dwarf star. The Apollo detector also revealed the existence of the first pulsar discovered outside the Milky Way. About 200 000 light years from earth's galaxy, in the Small Magellanic Cloud, it was the most luminous pulsar known to astronomers, 10 times brighter than any discovered so far. After repairing some malfunctioning equipment, the astronauts also mapped x-ray sources throughout the Milky Way.
The crew completed nearly all the 110 earth-observation tasks assigned. Coordinated investigations had been made simultaneously by six groups of scientists on the ground, on ships at sea and in \Jaircraft\j. The astronauts looked at ocean currents, ocean \Jpollution\j, desert \Jgeography\j, shoreline erosion, volcanoes, \Jiceberg\j movements, and vegetation patterns.
On 23 July the command-module tunnel was vented and the crew put on spacesuits to jettison the docking module. The command and service module unlocked from the DM at 3:43 pm EDT, and a 1-sec engine firing put the CSM into a higher orbit (232.2-km apogee, 219.0 km perigee) so that the DM could move ahead. A second maneuver put the CSM in a 223.2-km by 219.0-km orbit. Deorbit began at 4:38 p.m.
The command module and service module separated, the drogue and main parachutes deployed normally, and the Apollo splashed down at 224:58 GET (5:18 p.m. EDT July 24) in the Pacific Ocean 163W and 22N, 500 km west of Hawaii. This was the last ocean landing planned for U.S. human space flights; future flights on the Space Shuttle would be wheeled touchdowns at land bases.
The CM landed in "stable 2" position (upside down 7.4) km from the prime recovery ship, U.S.S. New Orleans. After swimmers from the rescue \Jhelicopter\j righted the \Jspacecraft\j and attached a flotation collar, the Apollo was lifted by crane on to the deck of the recovery ship and Stafford, Brand, and Slayton stepped out to the cheers of the ship's crew. President Ford telephoned congratulations.
During the welcome, the crew was evidently experiencing eye and lung discomfort; subsequent conversations and \Jspacecraft\j data revealed that, during reentry, the earth landing system had failed to jettison the apex cover and drogues as scheduled and had to be fired manually, without first disabling the reaction-control system thrusters.
With the CM oscillating, the thrusters began firing rapidly to compensate, and combustion products including a small amount of \Jnitrogen\j tetroxide entered through the cabin-pressure relief valves. As soon as the RCS system had been disabled, fresh air was once again drawn into the cabin. The crew members told flight officials that they had put on oxygen masks once the \Jspacecraft\j had landed, and then activated the postlanding vent system.
Because of the crew's discomfort, further shipboard ceremonies had been canceled and the crew had been sent to sick bay and then to Tripler \JHospital\j in Hawaii for observation until August 8, 1975.
Primary ASTP mission objectives were to evaluate the docking and undocking of an Apollo \Jspacecraft\j with a Soyuz, and determine the adequacy of the onboard orientation lights and docking target; evaluate the ability of astronauts and cosmonauts to make inter-vehicular crew transfers and the ability of \Jspacecraft\j systems to support the transfers: evaluate the Apollo's capability of maintaining attitude-hold control of the docked vehicles and performing attitude maneuvers; measure quantitatively the effect of weightlessness on the crews' height and lower limb volume, according to length of exposure to zero-g; and obtain relay and direct synchronous-satellite navigation tracking data to determine their accuracy for application to Space Shuttle navigation-system design. The objectives were successfully completed, and the mission was adjudged successful on August 15, 1975.
#
"Apollo-Soyuz",78,0,0,0
\BApollo 18 and Soyuz 19 Launched:\b July 15, 1975\b
\BMeeting in Space:\b July 17, 1975\b
\BSoyuz 19 Landed:\b July 21, 1975\b
\BApollo 18 Splashed Down:\b July 24, 1975\b
\BDuration:\b
ò Apollo 18: 217 hours, 30 minutes
ò Soyuz 19: 143 hours, 31 minutes
\BOrbits:\b (Apollo 18) 136; (Soyuz 19) 96\b
\BAstronaut Crew:\b
ò Thomas P. Stafford
ò Vance D. Brand
ò Donald K. "Deke" Slayton
\BCosmonaut Crew:\b
ò Alexei Leonov
ò Valeri Kubasov
ò The Flight of Apollo-Soyuz from the NASA History Office.
This, the final flight of the Apollo \Jspacecraft\j, was the first docking of \Jspacecraft\j built by different nations and presaged the era of cooperation between the Russians and the Americans that is now such an essential part of our efforts to build a permanently occupied space station.
The American crew included three-flight veteran Thomas P. Stafford, rookie Vance Brand, and the last of the original seven Mercury astronauts to make it into orbit, Donald K. "Deke" Slayton, whose heart murmur had previously kept him grounded. The Soviet crew included the first space walker, Alexei Leonov, and rookie Valeri Kubasov.
While this mission is generally remembered as a political/public relations venture, it resulted in some major technological advancements necessitated by the requirement to dock the two extremely variant \Jspacecraft\j, neither of which had been built for the purpose, together.
The two \Jspacecraft\j were launched within seven and a half hours of one another, and, three hours after they docked two days later, the Astronauts and Cosmonauts met in the middle and shook hands in orbit, exchanged flags and gifts (including the seeds of trees that were later planted in each others' countries) and conversed haltingly with one another in each other's native tongues.
It would be six long years before another American \Jastronaut\j would fly in space, this time aboard the reusable Space Shuttle. The Apollo era, an era of the greatest achievements in mankind's history, had ended.
#
"NASA Mercury Project",79,0,0,0
\JNASA Mercury Project Summary\j
\JFlights of Project Mercury\j
\JMercury-Redstone 4: Liberty Bell 7\j
\JMercury-Atlas 6: Friendship 7\j
\JMercury-Atlas 8: Sigma 7\j
\JMercury-Atlas 9: Faith 7\j
#
"NASA Mercury Project Summary",80,0,0,0
At the time of the announcement of Project Apollo by President Kennedy in May 1961, NASA was still consumed with the task of placing an American in orbit through Project Mercury. Stubborn problems arose, however, at seemingly every turn. The first space flight of an \Jastronaut\j, made by Alan B. Shepard, had been postponed for weeks so NASA engineers could resolve numerous details and only took place on 5 May 1961, less than three weeks before the Apollo announcement.
The second flight, a suborbital mission like Shepard's, launched on 21 July 1961, also had problems. The hatch blew off prematurely from the Mercury capsule, Liberty Bell 7, and it sank into the Atlantic Ocean before it could be recovered. In the process the \Jastronaut\j, "Gus" Grissom, nearly drowned before being hoisted to safety in a \Jhelicopter\j. These suborbital flights, however, proved valuable for NASA technicians who found ways to solve or work around literally thousands of obstacles to successful space flight.
As these issues were being resolved, NASA engineers began final preparations for the orbital aspects of Project Mercury. In this phase NASA planned to use a Mercury capsule capable of supporting a human in space for not just minutes, but eventually for as much as three days. As a launch vehicle for this Mercury capsule, NASA used the more powerful Atlas instead of the Redstone. But this decision was not without controversy. There were technical difficulties to be overcome in mating it to the Mercury capsule to be sure, but the biggest complication was a debate among NASA engineers over its propriety for human spaceflight.
When first conceived in the 1950s many believed Atlas was a high-risk proposition because to reduce its weight Convair Corp. engineers under the direction of Karel J. Bossart, a pre-World War II immigrant from \JBelgium\j, designed the booster with a very thin, internally pressurized fuselage instead of massive struts and a thick metal skin. The "steel balloon," as it was sometimes called, employed \Jengineering\j techniques that ran counter to a conservative \Jengineering\j approach used by Wernher von Braun for the \JV-2\j and the Redstone at \JHuntsville\j, \JAlabama\j.
Von Braun, according to Bossart, needlessly designed his boosters like "bridges," to withstand any possible shock. For his part, von Braun thought the Atlas too flimsy to hold up during launch. He considered Bossart's approach much too dangerous for human spaceflight, remarking that the \Jastronaut\j using the "contraption," as he called the Atlas booster, "should be getting a medal just for sitting on top of it before he takes off!" The reservations began to melt away, however, when Bossart's team pressurized one of the boosters and dared one of von Braun's engineers to knock a hole in it with a sledge hammer. The blow left the booster unharmed, but the recoil from the hammer nearly clubbed the engineer.
Most of the differences had been resolved by the first successful orbital flight of an unoccupied Mercury-Atlas combination in September 1961. On 29 November the final test flight took place, this time with the chimpanzee Enos occupying the capsule for a two-orbit ride before being successfully recovered in an ocean landing. Not until 20 February 1962, however, could NASA get ready for an orbital flight with an \Jastronaut\j.
On that date John Glenn became the first American to circle the Earth, making three orbits in his Friendship 7 Mercury \Jspacecraft\j. The flight was not without problems, however; Glenn flew parts of the last two orbits manually because of an autopilot failure, and left his normally jettisoned retrorocket pack attached to his capsule during reentry because of a loose heat shield.
Glenn's flight provided a healthy increase in national pride, making up for at least some of the earlier Soviet successes. The public, more than celebrating the technological success, embraced Glenn as a personification of heroism and dignity. Hundreds of requests for personal appearances by Glenn poured into NASA headquarters, and NASA learned much about the power of the astronauts to sway public opinion.
The NASA leadership made Glenn available to speak at some events, but more often substituted other astronauts and declined many other invitations. Among other engagements, Glenn did address a joint session of Congress and participated in several ticker-tape parades around the country. NASA discovered in the process of this hoopla a powerful public relations tool that it has employed ever since.
Three more successful Mercury flights took place during 1962 and 1963. Scott Carpenter made three orbits on 20 May 1962, and on 3 October 1962 Walter Schirra flew six orbits. The capstone of Project Mercury was the 15-16 May 1963 flight of Gordon Cooper, who circled the Earth 22 times in 34 hours. The program had succeeded in accomplishing its purpose: to successfully orbit a human in space, explore aspects of tracking and control, and to learn about \Jmicrogravity\j and other biomedical issues associated with spaceflight.
#
"Flights of Project Mercury",81,0,0,0
Project Mercury began on October 7, 1958, one year and three days after the launch of \JSputnik\j 1 by the Soviet Union heralded the beginning of the Space Age. The challenges were seemingly insurmountable; to devise a vehicle light enough to be launched into Earth orbit yet sturdy enough to withstand the forces of liftoff and splashdown, while protecting the pilot from the vacuum of space and the intense heat of re-entry into the atmosphere.
To meet these goals, they designed the \Jspacecraft\j as a wingless capsule, complete with an ablative heat shield that would literally be burned off during re-entry. This capsule was to sit atop a liquid-fueled ballistic missile.
The first, for suborbital flights, was the Redstone, designed by Wernher von Braun's \JHuntsville\j team. The second vehicle, designed to actually rocket the capsule into orbit, was the Atlas-D, with steel skin so thin that it would collapse like a plastic bag if it were not kept under constant pressure from within.
Three weeks after Alan Shepard's first manned suborbital flight, on May 5, 1961, President John F. Kennedy announced the goal of landing a man on the moon within the decade. On July 20, 1969, Neil Armstrong became that man when he stepped off Apollo 11's LEM onto the Moon's surface, taking the last "small step" on that Giant Leap begun a decade earlier by the Mercury astronauts.
The six Mercury flights totaled only two days, six hours in space. (Refer to Table)
#
"Mercury-Redstone 4: Liberty Bell 7",82,0,0,0
Astronaut: Virgil I. "Gus" Grissom
Launched: July 21, 1961
Landed: July 21, 1961 (15 Minutes later)
Orbits: 0
This mission was basically a repeat of Freedom 7, with some improvements, including a window and hand controls, added to the capsule design. Also added was an explosive escape hatch, which almost proved disastrous.
While in the ocean off the \JBahamas\j waiting to be picked up, the side hatch blew, causing the capsule to fill with water and sink, almost taking the \JAstronaut\j with it. A soaking wet Gus Grissom was safely rescued, although the Liberty Bell 7 sank to the bottom and was lost.
#
"Mercury-Atlas 6: Friendship 7",83,0,0,0
Astronaut: John H. Glenn, Jr.
Launched: February 20, 1962
Landed: February 20, 1962 (4 Hours, 55 Minutes later)
Orbits: 3
Apogee: 161.75 Miles
John Glenn became the first American to orbit the Earth aboard Friendship 7, circling the Earth three times. He became the first American to see a sunrise and sunset from space, and became the the first space photographer, using a Minolta camera he bought in a \JCocoa\j Beach drugstore to take some holiday snaps through the window.
There were some exciting times on this flight, particularly when a malfunctioning sensor indicated that the heat shield had come loose. During reentry, Glenn ran out of fuel trying to compensate for the capsule's bucking motions, but he nevertheless came to a safe landing about 40 miles from the target, a miss attributed to having overlooked the loss of consumable items when computing the \Jspacecraft\j's weight.
Those were simpler times!
#
"Mercury-Atlas 8: Sigma 7",84,0,0,0
Astronaut: Walter M. Schirra, Jr.
Launched: October 3, 1962
Landed: October 3, 1962 (9 Hours, 13 Minutes later)
Orbits: 6 Apogee: 175.73 Miles
On this flight, Wally Schirra became the first man to conduct a live TV broadcast from space. After Carpenter's overshoot of the landing site, this flight was intended to improve the fine-tuning of the \Jspacecraft\j's controls.
Much of the flight was spent with the craft on autopilot, although Schirra attempted to do some steering by the stars, a task he found difficult. He observed \Jlightning\j bolts in the atmosphere and took photographs with a Hasselblad camera he brought on-board.
This, at a little over 175 miles, was the highest flight of the Mercury program.
#
"Mercury-Atlas 9: Faith 7",85,0,0,0
Astronaut: L. Gordon Cooper, Jr.
Launched: May 15, 1963
Landed: May 16, 1963 (34 Hours, 19 Minutes later)
Orbits: 22.5
By far the longest of the Mercury flights, it was during this flight that the first satellite was released from a \Jspacecraft\j when Gordon Cooper detached a 152.4 mm sphere containing a beacon as part of a test of the \Jastronaut\j's visual tracking capabilities. Cooper also spotted a 44,000-watt \Jxenon\j lamp shining up from the ground, and he claimed to be able to discern individual buildings and smoke rising from chimneys.
He became the first American to sleep in space, and he took the best space pictures up to that time. His mission was such a roaring success that the planned seventh Mercury flight was canceled and we went instead directly into the somewhat more ambitious Gemini program.
#
"NASA Gemini Project",86,0,0,0
\JNASA Gemini Project Summary\j
\JFlights of Project Gemini\j
\JGemini 3: The Unsinkable Molly Brown\j
\JGemini IV\j
\JGemini V\j
\JGemini VI\j
\JGemini VII\j
\JGemini VIII\j
\JGemini IX\j
\JGemini XI\j
\JGemini XII\j
#
"NASA Gemini Project Summary",87,0,0,0
Even as the Mercury program was underway, and work took place developing Apollo hardware, NASA program managers perceived a huge gap in the capability for human spaceflight between that acquired with Mercury and what would be required for a Lunar landing. They closed most of the gap by experimenting and training on the ground, but some issues required experience in space.
Three major areas immediately arose where this was the case. The first was the ability in space to locate, maneuver toward, and rendezvous and dock with another \Jspacecraft\j. The second was closely related, the ability of astronauts to work outside a \Jspacecraft\j. The third involved the collection of more sophisticated physiological data about the human response to extended spaceflight.
To gain experience in these areas before Apollo could be readied for flight, NASA devised Project Gemini. Hatched in the fall of 1961 by engineers at Robert Gilruth's Space Task Group in cooperation with McDonnell \JAircraft\j Corp. technicians, builders of the Mercury \Jspacecraft\j, Gemini started as a larger Mercury Mark II capsule but soon became a totally different proposition. It could accommodate two astronauts for extended flights of more than two weeks. It pioneered the use of fuel cells instead of batteries to power the ship, and incorporated a series of modifications to hardware.
Its designers also toyed with the possibility of using a paraglider being developed at Langley Research Center for "dry" landings instead of a "splashdown" in water and recovery by the Navy. The whole system was to be powered by the newly developed Titan II launch vehicle, another ballistic missile developed for the Air Force. A central reason for this program was to perfect techniques for rendezvous and docking, so NASA appropriated from the military some Agena rocket upper stages and fitted them with docking adapters.
Problems with the Gemini program abounded from the start. The Titan II had longitudinal oscillations, called the "pogo" effect because it resembled the behavior of a child on a pogo stick. Overcoming this problem required \Jengineering\j imagination and long hours of overtime to stabilize fuel flow and maintain vehicle control. The fuel cells leaked and had to be redesigned, and the Agena reconfiguration also suffered costly delays.
NASA engineers never did get the paraglider to work properly and eventually dropped it from the program in favor of a parachute system the one used for Mercury. All of these difficulties shot an estimated $350 million program to over $1 billion. The overruns were successfully justified by the space agency, however, as necessities to meet the Apollo landing commitment.
By the end of 1963 most of the difficulties with Gemini had been resolved, albeit at great expense, and the program was ready for flight. Following two unoccupied orbital test flights, the first operational mission took place on 23 March 1965. Mercury \Jastronaut\j Grissom commanded the mission, with John W. Young, a Naval aviator chosen as an \Jastronaut\j in 1962, accompanying him.
The next mission, flown in June 1965 stayed aloft for four days and \Jastronaut\j Edward H. White II performed the first extra-vehicular activity (EVA) or spacewalk. Eight more missions followed through November 1966. Despite problems great and small encountered on virtually all of them, the program achieved its goals.
Additionally, as a technological learning program, Gemini had been a success, with 52 different experiments performed on the ten missions. The bank of data acquired from Gemini helped to bridge the gap between Mercury and what would be required to complete Apollo within the time constraints directed by the president.
#
"Flights of Project Gemini",88,0,0,0
Somewhat like the current Space Shuttle is a transition between our earliest space exploration and actually living and working in space, Project Gemini was a transitional project between the initial, pioneering Mercury Program and the actual space travel accomplished by the Apollo Program. And also akin to the Shuttle, its success was absolutely critical to achieving our goal of reaching the Moon.
The primary purpose of the Gemini missions was to learn how to "fly" a space vehicle, to maneuver in orbit, to rendezvous and dock with another vehicle, which were essential to the later Apollo missions. Project Gemini also demonstrated that astronauts could endure conditions of weightlessness for the length of time necessary for a lunar mission.
There were ten Gemini missions spanning a period of 20 months. It was during this period that Mission Control was transferred to the Johnson Space Center in Houston. Sixteen new astronauts joined the original seven, and space flight began to become routine.(Refer to Table)
The capsule was larger, almost twice as heavy as the Mercury capsule, but was a much tighter fit for the astronauts, having an interior space only 50% larger than Mercury's. There were many technological improvements, including fuel cells in lieu of batteries, complex maneuvering jets, on-board computers, a modular construction that permitted easy replacement of malfunctioning parts, and ejection seats to replace Mercury's escape rockets.
The launch vehicle, the Titan 2, was far more powerful than the old Atlas-D that had launched Mercury into orbit, and it also served to launch the "Agena" upper stage that contained the docking mechanisms used to train the astronauts for Apollo.
By the time of Gemini's final flight, Lunar Orbiter 2 was already mapping out Apollo landing sites. The Gemini Program provided another rather large step toward our Giant Leap to the moon.
#
"Gemini 3: The Unsinkable Molly Brown",89,0,0,0
Launched: March 23, 1965
Splashed Down: March 23, 1965
Orbits: 3
Duration: 4 hours, 53 minutes
Crew:
ò Virgil I. "Gus" Grissom
ò John W. Young
In a joking reference to the sinking of Liberty Bell 7 on the second suborbital Mercury mission, Gemini 3 became the only one of the Gemini missions to get a "nickname." Subsequent Gemini missions only received numbers, and were numbered with Roman numerals.
This was primarily a testing shakedown for the new, maneuverable Gemini capsule. During the flight, the astronauts used the thrusters to change the shape of their orbit and drop to a lower altitude. Because the aerodynamic behavior of the craft did not match wind tunnel predictions, the splashdown missed the target point by over 80 kilometers, and the shift of the capsule to landing position was so abrupt that Gus Grissom broke the faceplate of his helmet.
#
"Gemini IV",90,0,0,0
Launched: June 3, 1965
Splashed Down: June 7, 1965
Orbits: 62
Duration: 97 hours, 56 minutes
Crew:
ò James A. McDivitt
ò Edward H. White II
This mission was a real learning experience. The plan was to fly in formation with the spent Titan 2 second stage, but the dynamics of maneuvering toward another object in orbit proved to be far different from what flight engineers expected. When the astronauts tried to fly toward the target, the craft got farther and farther away. They had to give up the effort after burning half their fuel.
They discovered that, to catch up with an object ahead of you, you must drop down, and then rise back up after you catch up, rather than speed up, because speeding up puts you into a higher, and therefore slower, orbit.
The highlight of the mission was Ed White's 22-minute space walk, during which he used a hand-held "zip gun" to maneuver at the end of a tether. Gemini IV also marked the first flight controlled out of the new Mission Control Facility in Houston.
#
"Gemini V",91,0,0,0
Launched: August 21, 1965
Splashed Down: August 29, 1965
Orbits: 120
Duration: 190 hours, 55 minutes
Crew:
ò L. Gordon Cooper, Jr.
ò Charles "Pete" Conrad. Jr.
Some mark the Gemini V mission as the point at which America finally took the lead in the Space Race with the Soviets. By far the longest mission to date (surpassing the record of 119 hours, 6 minutes and 81 orbits set by Valery A. Bykovsky on Vostok 5 back in June, 1963), Gemini V, its duration was made possible by the use of new fuel cells for power. Of course, as with so many NEW! IMPROVED!! systems, the fuel cells proved a little bit troublesome at first, and their malfunctions prevented a rendezvous with a pod released from the craft and forced the cancellation of several other experiments, which gave the astronauts so much "free time" in orbit that Conrad complained about the fact that he had not brought along a book.
Medical tests were very positive, indicating that long-duration spaceflight is indeed very feasible, and this was essential in paving the way for the later Apollo lunar missions.
#
"Gemini VI",92,0,0,0
Launched: December 15, 1965
Splashed Down: December 16, 1965
Orbits: 16
Duration: 25 hours, 51 minutes
Crew:
ò Walter M. Schirra, Jr.
ò Thomas P. Stafford
This mission was originally intended to dock with an unoccupied Agena vehicle, but failure of an earlier test of the Agena prompted a rather spectacular change in plans; instead of the Agena, Gemini VI would rendezvous instead with Gemini VII.
This mission was originally scheduled to blast off on December 12, but the Titan 2 launch vehicle shut down right at T-minus-0. The astronauts could have ejected, but Schirra made the command decision to stay; he felt no motion and trusted his senses. Three days later, the lift off was successful.
Schirra used information from his on-board computer (as well as a bit of "seat-of-the-pants" flying) to rendezvous with Gemini VII on the afternoon of December 15. The two craft flew in formation and around one another, at one point coming within a foot without actually touching, for five hours. The astronauts returned the first pictures of another human-occupied craft in space.
#
"Gemini VII",93,0,0,0
Launched: December 4, 1965
Splashed Down: December 18, 1965
Orbits: 206
Duration: 330 hours, 35 minutes
Crew:
ò Frank Borman
ò James A. Lovell, Jr.
This mission set a duration record that remained unbroken until Cosmonauts Georgi Dobrovolsky, Vladislav Volkov and Viktor Patsayev stayed in orbit for 360 revolutions aboard Soyuz 11 in June, 1971, long after American astronauts had walked on the surface of the Moon. In the cramped Gemini quarters it proved to be a real endurance test, during which the astronauts tested a new, lighter space suit (which proved quite uncomfortable if worn for long periods).
This crew conducted the most experiments, 20, of any Gemini mission, inculding studies of \Jnutrition\j in space. The highlight of this mission was, of course, the successful rendezvous with Gemini VI, which was launched after Gemini VII had already been in orbit eleven days. Following the rendezvous, the astronauts just marked time for three more days; fortunately, they heeded the advice of Pete Conrad and brought their books along.
#
"Gemini VIII",94,0,0,0
Launched: March 16, 1966
Splashed Down: March 16, 1966
Orbits: 16
Duration: 10 hours, 11 minutes
Crew:
ò Neil A. Armstrong
ò David R. Scott
This mission was the first to actually dock with another craft in space, linking up to an unmanned Agena target vehicle that had been launched earlier. What followed, though, was one of the most frightening events in the history of the space program. The two craft, still linked together, began to roll continuously.
When the astronauts succeeded in undocking from the Agena, the spinning got faster and faster because the problem was a stuck thruster on the Gemini capsule itself. The dizzying spin got up to one revolution per second, and Armstrong and Scott had to use the reentry control thrusters to stabilize the craft.
The mission had to be cut short; a planned spacewalk by Scott had to be cancelled. The still-nauseous astronauts returned only ten hours after launch, but at least they came back alive, and Neil Armstrong survived to take his "giant leap for mankind" aboard Apollo 11.
#
"Gemini IX",95,0,0,0
Launched: June 3, 1966
Splashed Down: June 6, 1966
Orbits: 45
Duration: 72 hours, 21 minutes
Crew:
ò Thomas P. Stafford
ò Eugene A. Cernan
This mission began on a very sad note; the crew originally scheduled for this flight, Elliott See and Charles Bassett, had died in a plane crash four months earlier. Stafford and Cernan were the first backup crew to fly in space.
Gemini IX was supposed to dock with with the new ATDA (Augmented Target Docking Adapter), a shortened version of the Agena, but the docking had to be cancelled when, after the successful rendezvous, it was discovered that the protective shroud failed to disengage from the ATDA properly, resulting in what came to be known as the "Angry Alligator."
Cernan was also supposed to have test-flown the AMU (Astronaut Maneuvering Unit) during a space walk, but so many hassles ensued, including a fogged visor in his space helmet, that the test had to be cancelled. The device was never tested until Skylab, seven years later.
#
"Gemini XI",96,0,0,0
Launched: September 12, 1966
Splashed Down: September 15, 1966
Orbits: 44
Duration: 71 hours, 17 minutes
Crew:
ò Charles "Pete" Conrad
ò Richard F. Gordon, Jr.
Gemini XI was a race to dock quickly with an already orbiting Agena, to closely simulate the conditions that would soon be encountered by LEM astronauts attempting to dock with the Apollo Command Module in orbit around the Moon. Only 85 minutes after launch the rendezvous was successful and GEMINI XI docked with the Agena target vehicle several times successfully.
There was a chance that this mission would be given a go-ahead for an orbit around the moon, but this was scrapped and the crew had to be satisfied with reaching the highest orbit ever by a human spaceflight, 848.7 miles up.
This mission also succeeded in tethering the Gemini capsule and the Agena vehicle together, and a partially successful attempt to rotate the pair resulted in the first "artificial gravity," although there were problems keeping the tether taut.
This mission also featured the first fully automatic, computer controlled landing, and splashed down only 2.8 miles from the recovery ship.
#
"Gemini XII",97,0,0,0
Launched: November 11, 1966
Splashed Down: November 15, 1966
Orbits: 59
Duration: 94 hours, 35 minutes
Crew:
ò James A. Lovell, Jr.
ò Edwin E. "Buzz" Aldrin, Jr.
Gemini XII finally demonstrated what had proven elusive on all previous Gemini flights; that it WAS indeed possible for man to work effectively outside the protected environment of a \Jspacecraft\j in 0G. Prior to this flight, the new technique of underwater training had prepared Aldrin far better for the rigors of weightlessness, and during his two hour, 20 minute space walk he photographed star fields, retrieved a micrometeorite collector and performed various other tasks.
The docking with the Agena was routine; we had that little chore down pat now. Problems with the Agena vehicle's rockets prevented attaining a high orbit, but two additional space walks were successful and the remaining planned Gemini missions were cancelled; the Gemini program had done its job, preparing us well for the challenges of Apollo.
#
"Ranger (1964 - 1965)",98,0,0,0
The Ranger series was the first US attempt to obtain close-up images of the Lunar surface. The Ranger \Jspacecraft\j were designed to fly straight down towards the Moon and send images back until the moment of impact. This image has a resolution of 5 meters. Ranger 7 impacted in mare terrain modified by crater rays. Ranger 8 also impacted in mare terrain, but this area contained a complex system of ridges. Ranger 9 impacted in a large crater in the lunar highlands.
\BRanger 7
Launched 28 July 1964\b
Impacted Moon 31 July 1964 at 13:25:49 UT
Latitude 10.35 S, Longitude 339.42 E - Mare Cognitum (Sea of Clouds)
\BRanger 8
Launched 17 February 1965\b
Impacted Moon 20 February 1965 at 09:57:37 UT
Latitude 2.67 N, Longitude 24.65 E - Mare Tranquillitatis (Sea of Tranquility)
\BRanger 9
Launched 21 March 1965\b
Impacted Moon 24 March 1965 at 14:08:20 UT
Latitude 12.83 S, Longitude 357.63 E - Alphonsus
Each Ranger \Jspacecraft\j had 6 cameras on board. The cameras were fundamentally the same with differences in exposure times, fields of view, lenses, and scan rates. The camera system was divided into two channels, P (partial) and F (full). Each channel was self-contained with separate power supplies, timers, and transmitters. The F-channel had 2 cameras: the wide-angle A-camera and the narrow angle B-camera. The P-channel had four cameras: P1 and P2 (narrow angle) and P3 and P4 (wide angle).
The final F-channel image was taken between 2.5 and 5 sec before impact (altitude about 5 km) and the last P-channel image 0.2 to 0.4 sec before impact (altitude about 600 m). The images provided better resolution than was available from Earth-based views by a factor of 1,000. These highly detailed images showed Apollo planners that finding a smooth landing site was not going to be easy.
#
"Lunar Orbiter (1966 - 1967)",99,0,0,0
Five Lunar Orbiter missions were launched in 1966 through 1967 with the purpose of mapping the lunar surface before the Apollo landings. All five missions were successful, and 99% of the Moon was photographed with a resolution of 60 m or better. The first three missions were dedicated to imaging 20 potential lunar landing sites, selected based on Earth-based observations. These were flown at low inclination orbits. The fourth and fifth missions were devoted to broader scientific objectives and were flown in high altitude polar orbits. Lunar Orbiter 4 photographed the entire nearside and 95% of the farside, and Lunar Orbiter 5 completed the farside coverage and acquired medium (20 m) and high (2 m) resolution images of 36 pre-selected areas.
\BLunar Orbiter 1
Launched 10 August 1966\b
Imaged Moon: 18-29 August 1966
Apollo landing site survey mission
\BLunar Orbiter 2
Launched 06 November 1966\b
Imaged Moon: 18-25 November 1966
Apollo landing site survey mission
\BLunar Orbiter 3
Launched 05 February 1967\b
Imaged Moon: 15-23 February 1967
Apollo landing site survey mission
\BLunar Orbiter 4
Launched 04 May 1967\b
Imaged Moon: 11-26 May 1967
Lunar mapping mission
\BLunar Orbiter 5
Launched 01 August 1967\b
Imaged Moon: 06-18 August 1967
Lunar mapping and hi-res survey mission
The Lunar Orbiters had an ingenious imaging system, which consisted of a dual-lens camera, a film processing unit, a readout scanner, and a film handling apparatus. Both lenses, a 610-mm narrow angle high-resolution (HR) lens and an 80-mm wide-angle medium resolution (MR) lens, placed their frame exposures on a single roll of 70 mm film. The axes of the two cameras were coincident so the area imaged in the HR frames were centered within the MR frame areas. The film was moved during exposure to compensate for the \Jspacecraft\j velocity, which was estimated by an electric-optical sensor. The film was then processed, scanned, and the images transmitted back to Earth.
#
"Surveyor (1966 - 1968)",100,0,0,0
The Surveyor probes were the first US \Jspacecraft\j to land safely on the Moon. The main objectives of the Surveyors were to obtain close-up images of the lunar surface and to determine if the terrain was safe for manned landings. Each Surveyor was equipped with a \Jtelevision\j camera. In addition, Surveyors 3 and 7 each carried a soil mechanics surface sampler scoop which dug trenches and was used for soil mechanics tests and Surveyors 5, 6, and 7 had magnets attached to the footpads and an alpha scattering instrument for chemical analysis of the lunar material.
\BSurveyor 1
Launched 30 May 1966\b
Landed 02 June 1966, 06:17:37 UT
Latitude 2.45 S, Longitude 316.79 E - Flamsteed P
\BSurveyor 2
Launched 20 September 1966\b
Crashed on Moon 22 September 1966
Vernier engine failed to ignite - southeast of Copernicus Crater
\BSurveyor 3
Launched 17 April 1967\b
Landed 20 April 1967, 00:04:53 UT
Latitude 2.94 S, Longitude 336.66 E - \JOceanus\j Procellarum (Ocean of Storms)
\BSurveyor 4
Launched 14 July 1967\b
Radio contact lost 17 July 1967
2.5 minutes from touchdown - Sinus Medii
\BSurveyor 5
Launched 08 September 1967\b
Landed 11 September 1967, 00:46:44 UT
Latitude 1.41 N, Longitude 23.18 E - Mare Tranquillitatus (Sea of Tranquility)
\BSurveyor 6
Launched 07 November 1967\b
Landed 10 November 1967, 01:01:06 UT
Latitude 0.46 N, Longitude 358.63 E - Sinus Medii
\BSurveyor 7
Launched 07 January 1968\b
Landed 10 January 1968, 01:05:36 UT
Latitude 41.01 S, Longitude 348.59 E - Tycho North Rim
#
"Apollo Program (1968 - 1972)",101,0,0,0
The Apollo program was designed to land humans on the Moon and bring them safely back to Earth. Six of the missions (Apollos 11, 12, 14, 15, 16, and 17) achieved this goal. Apollos 7 and 9 were Earth orbiting missions to test the Command and Lunar Modules, and did not return lunar data. Apollos 8 and 10 tested various components while orbiting the Moon, and returned photography of the lunar surface.
Apollo 13 did not land on the Moon due to a malfunction, but also returned photographs. The six missions that landed on the Moon returned a wealth of scientific data and almost 400 kilograms of lunar samples. Experiments included soil mechanics, meteoroids, seismic, heat flow, lunar ranging, magnetic fields, and solar wind experiments.
\BApollo lunar missions
Apollo 8
Launched 21 December 1968\b
Lunar Orbit and Return
Returned to Earth 27 December 1968
\BApollo 10
Launched 18 May 1969\b
Lunar Orbit and Return
Returned to Earth 26 May 1969
\BApollo 11
Launched 16 July 1969\b
Landed on Moon 20 July 1969
Sea of Tranquility
Returned to Earth 24 July 1969
\BApollo 12
Launched 14 November 1969\b
Landed on Moon 19 November 1969
Sea of Storms
Returned to Earth 24 November 1969
\BApollo 13
Launched 11 April 1970\b
Lunar Flyby and Return
Malfunction forced cancellation of lunar landing
Returned to Earth 17 April 1970
\BApollo 14
Launched 31 January 1971\b
Landed on Moon 5 February 1971
Fra Mauro
Returned to Earth 9 February 1971
\BApollo 15
Launched 26 July 1971\b
Landed on Moon 30 July 1971
Hadley Rille
Returned to Earth 7 August 1971
\BApollo 16
Launched 16 April 1972\b
Landed on Moon 20 April 1972
Descartes
Returned to Earth 27 April 1972
\BApollo 17
Launched 07 December 1972\b
Landed on Moon 11 December 1972
Taurus-Littrow
Returned to Earth 19 December 1972
The Apollo mission consisted of a Command Module (CM) and a Lunar Module (LM). The CM and LM would separate after lunar orbit insertion. One crew member would stay in the CM, which would orbit the Moon, while the other two astronauts would take the LM down to the lunar surface. After exploring the surface, setting up experiments, taking pictures, collecting rock samples, etc., the astronauts would return to the CM for the journey back to Earth.
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"Key Documents from the Apollo Space Program",102,0,0,0
This section looks at five key documents from the Apollo Project. They include:
\B\IPresident Kennedy asks the Vice President for Recommendations on the US Space Program
The Vice President Answers
President Kennedy Announces the Apollo Decision
The Kennedy Administration Defends Apollo
Neil Armstrong Sets Foot on the Moon\b\i
\BPresident Kennedy asks the Vice President for Recommendations on the US Space Program\b
The memorandum that follows led directly to the Apollo program. By posing the question:"Is there any . . . space program which promises dramatic results in which we could win?" President Kennedy set in motion a review that concluded that only a crash effort to send Americans to the Moon met the criteria Kennedy had laid out. This memorandum followed a week of discussion within the White House on how best to respond to the challenge to US interests posed by the 12 April 1961 orbital flight of Soviet \JCosmonaut\j Yuri Gagarin.
\BThe White House
April 20, 1961
Memorandum for Vice President\b
In accordance with our conversation I would like for you as Chairman of the Space Council to be in charge of making an overall survey of where we stand in space.
1. Do we have a chance of beating the Soviets by putting a laboratory in space, or by a trip round the moon, or by a rocket to land on the moon, or by a rocket to go to the moon and back with a man. Is there any other space program which promises dramatic results in which we could win?
2. How much additional would it cost?
3. Are we working 24 hours a day on existing programs. If not, why not? If not, will you make recommendations to me as to how work can be speeded up.
4. In building large boosters should we put our emphasis on nuclear, chemical or liquid fuel, or a combination of these three?
5. Are we making maximum effort? Are we achieving necessary results?
I have asked Jim Webb, Dr. Weisner, Secretary McNamara and other responsible officials to cooperate with you fully. I would appreciate a report on this at the earliest possible moment.
[Signed] John F. Kennedy
\BThe Vice President Answers\b
This memorandum, prepared by Edward Welsh, new Executive Secretary of the National Aeronautics and Space Council, and signed by Vice President Johnson, was the first report to President Kennedy on the results of the review he had ordered of the space program on 20 April 1961. It identified a human Lunar landing by 1966 or 1967 as the first dramatic space project in which the United States could beat the Soviet Union. Vice President Johnson identified US "leadership" in the world arena as sufficient justification of this undertaking in space.
\BOffice of the Vice President
April 28, 1961
MEMORANDUM FOR PRESIDENT
Subject: Evaluation of Space Program.\b
Reference is to your April 20 memorandum asking certain questions regarding this country's space program.
A detailed survey has not been completed in this time period. The examination will continue. However, what we have obtained so far from knowledgeable and responsible persons makes this summary reply possible.
Among those who have participated in our deliberations have been the Secretary and Deputy Secretary of Defense; General Schriever (AF); Admiral Hayward (Navy); Dr. von Braun (NASA); the Administrator, Deputy Administrator, and other top officials of NASA; the Special Assistant to the President on Science and Technology; representatives of the Director of the Bureau of the Budget; and three outstanding non-Government citizens of the general public: Mr. George Brown (Brown & Root, Houston, Texas); Mr. Donald Cook (American Electric Power Service, New York, NY); and Mr. Frank Stanton (Columbia Broadcasting System, New York, N.Y.).
The following general conclusions can be reported:
a. Largely due to their concentrated efforts and their earlier emphasis upon the development of large rocket engines, the Soviets are ahead of the United States in world prestige attained through impressive technological accomplishments in space.
b. The US has greater resources than the USSR for attaining space leadership but has failed to make the necessary hard decisions and to marshal those resources to achieve such leadership.
c. This country should be realistic and recognize that other nations, regardless of their appreciation of our idealistic values, will tend to align themselves with the country which they believe will be the world leader -- the winner in the long run. Dramatic accomplishments in space are being increasingly identified as a major indicator of world leadership.
d. The US can, if it will, firm up its objectives and employ its resources with a reasonable chance of attaining world leadership in space during this decade. This will be difficult but can be made probable even recognizing the head start of the Soviets and the likelihood that they will continue to move forward with impressive successes. In certain areas, such as communications, navigation, weather, and mapping, the US can and should exploit its existing advance position.
e. If we do not make the strong effort now, the time will soon be reached when the margin of control over space and over men's minds through space accomplishments will have swung so far on the Russian side that we will not be able to catch up, let alone assume leadership.
f. Even in those areas in which the Soviets already have the capability to be first and are likely to improve upon such capability, the United States should make aggressive efforts as the technological gains as well as the international rewards are essential steps in eventually gaining leadership. The danger of long lags or outright omissions by this country is substantial in view of the possibility of great technological breakthroughs obtained from space exploration.
g. Manned exploration of the moon, for example, is not only an achievement with great propaganda value, but it is essential as an objective whether or not we are first in its accomplishment -- and we may be able to be first. We cannot leapfrog such accomplishments, as they are essential sources of knowledge and experience for even greater successes in space. We cannot expect the Russians to transfer the benefits of their experiences or the advantages of their capabilities to us. We must do these things ourselves.
h. The American public should be given the facts as to how we stand in the space race, told of our determination to lead in that race, and advised of the importance of such leadership to our future.
i. More resources and more effort need to be put into our space program as soon as possible. We should move forward with a bold program, while at the same time taking every practical precaution for the safety of the persons actively participating in space flights.
As for the specific questions posed in your memorandum, the following brief answers develop from the studies made during the past few days. These conclusions are subject to expansion and more detailed examination as our survey continues.
Q.1- Do we have a chance of beating the Soviets by putting a laboratory in space, or by a trip around the moon, or by a rocket to land on the moon, or by a rocket to go to the moon and back with a man. Is there any other space program which promises dramatic results in which we could win?
A.1- The Soviets now have a rocket capability for putting a multi-manned laboratory into space and have already crash-landed a rocket on the moon. They also have the booster capability of making a soft landing on the moon with a payload of instruments, although we do not know how much preparation they have made for such a project. As for a manned trip around the moon or a safe landing and return by a man to the moon, neither the US nor the USSR has such capability at this time, so far as we know. The Russians have had more experience with large boosters and with flights of dogs and man. Hence they might be conceded a time advantage in circumnavigation of the moon and also in a manned trip to the moon. However, with a strong effort, the United States could conceivably be first in those two accomplishments by 1966 or 1967.
There are a number of programs which the United States could pursue immediately and which promise significant world-wide advantage over the Soviets. Among these are communications satellites, and navigation and mapping satellites. These are all areas in which we have already developed some competence. We have such programs and believe that the Soviets do not. Moreover, they are programs which could be made operational and effective within reasonably short periods of time and could, if properly programmed with the interests of other nations, make useful strides toward world leadership.
Q.2- How much additional would it cost?
A.2- To start upon an accelerated program with the aforementioned objectives clearly in mind, NASA has submitted an analysis indicating that about $500 million would be needed for FY 1962 over and above the amount currently requested of the Congress. A program based upon NASA's analysis would, over a ten-year period, average approximately $1 billion a year above the current estimates of the existing NASA program.
While the Department of Defense plans to make a more detailed submission to me within a few days, the Secretary has taken the position that there is a need for a strong effort to develop a large solid-propellant booster and that his Department is interested in undertaking such a project. It was understood that this would be programmed in accord with the existing arrangement for close cooperation with NASA, which Agency is undertaking some research in this field. He estimated they would need to employ approximately $50 million during FY 1962 for this work but that this could be financed through management of funds already requested in the FY 1962 budget. Future defense budgets would include requests for additional funding for this purpose; a preliminary estimate indicates that about $500 million would be needed in total.
Q.3- Are we working 24 hours a day on existing programs? If not, why not? If not, will you make recommendations to me as to how work can be speeded up?
A.3- There is not a 24-hour-a-day work schedule on existing NASA space programs except for selected areas in Project Mercury, the Saturn-C-1 booster, the \JCentaur\j engines and the final launching phases of most flight missions. They advise that their schedules have been geared to the availability of facilities and financial resources, and that hence their overtime and 3-shift arrangements exist only in those activities in which there are particular bottlenecks or which are holding up operations in other parts of the programs. For example, they have a 3-shift 7-day-week operation in certain work at Cape Canaveral; the contractor for Project Mercury has averaged a 54-hour week and employs two or three shifts in some areas; Saturn C-1 at \JHuntsville\j is working around the clock during critical test periods while the remaining work on this project averages a 47-hour week; the \JCentaur\j \Jhydrogen\j engine is on a 3-shift basis in some portions of the contractor's plants.
This work can be speeded up through firm decisions to go ahead faster if accompanied by additional funds needed for the acceleration.
Q.4- In building large boosters should we put our emphasis on nuclear, chemical or liquid fuel, or a combination of these three?
A.4- It was the consensus that liquid, solid and nuclear boosters should all be accelerated. This conclusion is based not only upon the necessity for back-up methods, but also because of the advantages of the different types of boosters for different missions. A program of such emphasis would meet both so-called civilian needs and defense requirements.
Q.5- Are we making maximum effort? Are we achieving necessary results?
A.5- We are neither making maximum effort nor achieving results necessary if this country is to reach a position of leadership.
[signed] Lyndon B. Johnson
\BPresident Kennedy Announces the Apollo Decision\b
John F. Kennedy unveiled the commitment to execute Project Apollo before Congress on 25 May 1961 in a speech on "Urgent National Needs," billed as a second State of the Union message. In the speech he asked for support to accomplish four basic goals in space exploration, only the Lunar landing is usually remembered. In addition, he asked for congressional appropriations for weather satellites, communications satellites, and the Rover nuclear propulsion rocket. Congress agreed to all of them with barely any comment. As seen in this excerpt from the speech, Kennedy couched the space program in the context of the cold war rivalry with the Soviet Union:
. . . Finally, if we are to win the battle that is now going on around the world between freedom and tyranny, the dramatic achievements in space which occurred in recent weeks should have made clear to us all, as did the \JSputnik\j in 1957, the impact of this adventure on the minds of men everywhere, who are attempting to make a determination of which road they should take. Since early in my term, our efforts in space have been under review. With the advice of the Vice President, who is Chairman of the National Space Council, we have examined where we are strong and where we are not, where we may succeed and where we may not. Now it is time to take longer strides -- time for this nation to take a clearly leading role in space achievement, which in many ways may hold the key to our future on earth.
I believe we possess all the resources and talents necessary. But the facts of the matter are that we have never made the national decisions or marshaled the national resources required for such leadership. We have never specified long-range goals on an urgent time schedule, or managed our resources and our time so as to insure their fulfillment.
Recognizing the head start obtained by the Soviets with their large rocket engines, which gives them many months of lead-time, and recognizing the likelihood that they will exploit this lead for some time to come in still more impressive successes, we nevertheless are required to make new efforts on our own. For while we cannot guarantee that we shall one day be first, we can guarantee that any failure to make this effort will make us last. We take an additional risk by making it in full view of the world, but as shown by the feat of \Jastronaut\j Shepard, this very risk enhances our stature when we are successful. But this is not merely a race. Space is open to us now; and our eagerness to share its meaning is not governed by the efforts of others. We go into space because whatever mankind must undertake, free men must fully share.
I therefore ask the Congress, above and beyond the increases I have earlier requested for space activities, to provide the funds which are needed to meet the following national goals:
First, I believe that this nation should commit itself to achieving the goal, before this decade is out, of landing a man on the moon and returning him safely to the earth. No single space project in this period will be more impressive to mankind, or more important for the long-range exploration of space; and none will be so difficult or expensive to accomplish. We propose to accelerate the development of the appropriate lunar space craft. We propose to develop alternate liquid and solid fuel boosters, much larger than any now being developed, until certain which is superior. We propose additional funds for other engine development and for unmanned explorations -- explorations which are particularly important for one purpose which this nation will never overlook; the survival of the man who first makes this daring flight. But in a very real sense, it will not be one man going to the moon -- if we make this judgment affirmatively, it will be an entire nation. For all of us must work to put him there.
Secondly, an additional 23 million dollars, together with 7 million dollars already available, will accelerate development of the Rover nuclear rocket. This gives promise of some day providing a means for even more exciting and ambitious exploration of space, perhaps beyond the moon, perhaps to the very end of the solar system itself.
Third, an additional 50 million dollars will make the most of our present leadership, by accelerating the use of space satellites for world-wide communications.
Fourth, an additional 75 million dollars -- of which 53 million dollars is for the Weather Bureau -- will help give us at the earliest possible time a satellite system for world-wide weather observation.
Let it be clear -- and this is a judgment which the Members of the Congress must finally make -- let it be clear that I am asking the Congress and the country to accept a firm commitment to a new course of action -- a course which will last for many years and carry very heavy costs: 531 million dollars in fiscal '62 -- and estimated seven to nine billion dollars additional over the next five years. If we are to go only half way, or reduce our sights in the face of difficulty, in my judgment it would be better not to go at all.
Now this is a choice which this country must make, and I am confident that under the leadership of the Space Committees of the Congress, and the Appropriating Committees, that you will consider the matter carefully.
It is a most important decision that we make as a nation. But all of you have lived through the last four years and have seen the significance of space and the adventure in space, and no one can predict with certainty what the ultimate meaning will be of mastery of space.
I believe we should go to the moon. But I think every citizen of this country as well as the Members of the Congress should consider the matter carefully in making their judgment, to which we have given attention over many weeks and months, because it is a heavy burden, and there is no sense in agreeing or desiring that the United States take an affirmative position in outer space, unless we are prepared to do the work and bear the burdens to make it successful. If we are not, we should decide today and this year.
This decision demands a major national commitment of scientific and technical manpower, material and facilities, and the possibility of their diversion from other important activities where they are already thinly spread. It means a degree of dedication, organization and discipline which have not always characterized our research and development efforts. It means we cannot afford undue work stoppages, inflated costs of material or talent, wasteful interagency rivalries, or a high turnover of key personnel.
New objectives and new money cannot solve these problems. They could in fact, aggravate them further -- unless every scientist, every engineer, every serviceman, every technician, contractor, and civil servant gives his personal pledge that this nation of freedom, in the exciting adventure of space. . . .
\B\ITo continue please click\b\i \JKey Documents from the Apollo Space Program continued\j
#
"Key Documents from the Apollo Space Program continued",103,0,0,0
\BThe Kennedy Administration Defends Apollo\b
Criticism of the priority assigned to the space program, and particularly Project Apollo, increased in 1963. Congress tried unsuccessfully, for example, to delete $700 million from the NASA appropriation for Apollo. As a result, on 9 April 1963 Kennedy asked Lyndon Johnson as head of the National Aeronautics and Space Council for a careful review of the program. Johnson replied on 13 May with a lengthy report that emphasized the positive results of the space program and noted challenges that it posed. In the end, as these excerpts suggest, Johnson's report reflected the administration's continued commitment to an aggressive Lunar landing program for international prestige, scientific, and cost benefit reasons.
\BII. Benefits to National Economy from NASA Space Programs\b
1. It cannot be questioned that billions of dollars directed into research and development in an orderly and thoughtful manner will have a significant effect upon our national economy. No formula has been found which attributes specific dollar values to each of the areas of anticipated developments, however, the "multiplier" of space research and development will augment our economic strength, our peaceful posture, and our standard of living.
2. Even though specific dollar values cannot be set for these benefits, a mere listing of the fields which will be affected is convincing evidence that the benefits will be substantial. The benefits include:
a. Additional knowledge about the earth and the Sun's influence on the earth, the nature of interplanetary space environment, and the origin of the solar system as well as of life itself.
b. Increased ability and experience in managing major research and development efforts, expansion of capital facilities, encouragement of higher standards of quality production,
c. Accelerated use of liquid oxygen in steelmaking, coatings for temperature control of housing, efficient transfer of chemical energy into electrical energy, and wide-range advances in \Jelectronics\j.
d. Development of effective filters against detergents; increased accuracy (and therefore reduced costs) in measuring hot steel rods; improved medical equipment in human care; stimulation of the use of fiberglass refractory welding tape, high energy metal forming processes; development of new coatings for plywood and furniture; use of frangible tube energy absorption systems that can be adapted to absorbing shocks of failing elevators and emergency \Jaircraft\j landings.
e. Improved communications, improved weather forecasting, improved forest fire detection, and improved navigation.
f. Development of high temperature gas-cooled \Jgraphite\j moderated reactors and liquid metal cooled reactors; development of \Jradioisotope\j power sources for both military and civilian uses; development of instruments for monitoring degrees of radiation; and application of thermoelectric and thermionic conversion of heat to electric energy.
g. Improvements in metals, alloys, and \Jceramics\j.
h. An augmentation of the supply of highly trained technical manpower.
i. Greater strength for the educational system both through direct grants, facilities and scholarships and through setting goals that will encourage young people.
j. An expansion of the base for peaceful cooperation among nations.
k. Military competence. (It is estimated that between $600 and $675 million of NASA's FY 1964 budget would be needed for military space projects and would be budgeted by the Defense Department, if they were not already provided for in the NASA budget.)
\BIII. Problems Resulting from the Space Program\b
1. The introduction of a vital new element into an economy always creates new problems but, otherwise, the nation's space program creates no major complications. The program has, to a lesser magnitude, the same problems which Defense budgets and programs have been creating for several years.
2. Despite claims to the contrary, there is no solid evidence that research and development in industry is suffering significantly from a diversion of technical manpower to the space program. NASA estimates that:
a. The nation's pool of scientists and engineers was 1,400,000 as of January l, 1963.
b. NASA programs employed 42,000 of these scientists and engineers -- only 9,000 directly on NASA payrolls.
c. On this basis, the NASA space program currently draws upon only 3% of the national pool of scientists and engineers.
d. Taking into account anticipated expansion, NASA programs are not expected to absorb more than 7% of our country's total supply of scientists and engineers.
3. The majority of the technical people working for NASA fall in the category of \Jengineering\j. However, NASA's education programs are designed to help the universities train additions to the nation's technical manpower needs.
4. NASA has undertaken to support the annual graduate training of 1,000 Ph.Ds, 1/4 of the estimated overall shortage of 4,000 per year. This program would more than replace those drawn upon by the agency.
5. In overall terms, NASA finds that diversion of manpower and resources is not a major problem arising from the space program. A major problem, however, is the need to minimize waste and inefficiency. To help meet this challenge, turnover of top level Government talent should be reduced and compensation more in line with responsibilities would contribute to this objective.
\BConclusion\b
There is one further point to be borne in mind. The space program is not solely a question of prestige, of advancing scientific knowledge, of economic benefit or of military development, although all of these factors are involved. Basically, a much more fundamental issue is at stake -- whether a dimension that can well dominate history for the next few centuries will be devoted to the social system of freedom or controlled by the social system of \Jcommunism\j.
The United States has made clear that it does not seek to "dominate" space and, in fact, has led the way in securing international cooperation in this field. But we cannot close our eyes as to what would happen if we permitted totalitarian systems to dominate the environment of the earth itself. For this reason our space program has an overriding urgency that cannot be calculated solely in terms of industrial, scientific, or military development. The future of society is at stake.
\BNeil Armstrong Sets Foot on the Moon\b
After eight years of all-out effort, nearly $20 billion expended, and three astronauts' deaths, on 20 July 1969 Apollo 11 landed on the Moon. The two astronauts who set foot on the surface, Neil A. Armstrong and Edwin E. Aldrin, called it in what later astronauts thought of as an understatement, "magnificent desolation." This document contains the radio transmissions of the landing and Armstrong's first venture out onto the Lunar surface. The "CC" in the transcript is Houston Mission Control, CDR is Neil Armstrong, and LMP is Buzz Aldrin.
04 06 45 52....CC....We copy you down, Eagle [the name of the Lunar Module].
04 06 45 57....CDR....Houston, Tranquility Base here.
04 06 45 59....CDR....THE EAGLE HAS LANDED.
04 06 46 04....CC....Roger, Tranquility. We copy you on the ground. You got a bunch of guys about to turn blue. We're breathing again. Thanks a lot.
04 06 46 16....CDR....Thank you.
04 06 46 18....CC....You're looking good here.
04 06 46 23....CDR....Okay. We're going to be busy for a minute. .
04 13 23 38....CDR....[After suiting up and exiting the Lunar Module (LM), Armstrong was ready to descend to the Moon's surface]. I'm at the foot of the ladder. The LM footpads are only depressed in the surface about 1 or 2 inches, although the surface appears to be very, very fine grained, as you get close to it. It's almost like a powder. Down there, it's very fine.
04 13 23 43....CDR....I'm going to step off the LM now.
04 13 24 48....CDR....\BTHAT'S ONE SMALL STEP FOR MAN, ONE GIANT LEAP FOR MANKIND. \b
04 13 24 48....CDR....And the -- the surface is fine and powdery. I can -- I can pick it up loosely with my toe. It does adhere in fine layers like powdered charcoal to the sole and sides of my boots. I only go in a small fraction of an inch, maybe an eighth of an inch, but I can see the footprints of my boots and the treads in the fine, sandy particles.
04 13 25 30....CC....Neil, this is Houston. We're copying.
04 13 25 45....CDR....There seems to be no difficulty in moving around as we suspected. It's even perhaps easier than the simulations at one-sixth g that we performed in the various simulations on the ground. It's actually no trouble to walk around. Okay. The descent engine did not leave a crater of any size. It has about 1 foot clearance on the ground. We're essentially on a very level place here. I can see some evidence of rays emanating from the descent engine, but a very insignificant amount. . . .
#
"Apollo 11, Twenty-Five Years On",104,0,0,0
\BApollo-11
Crew\b
Neil A. Armstrong (2), Commander
Edwin E. Aldrin (2), Jr., Lunar Module Pilot
Michael Collins (2), Command Module Pilot
\BBackup Crew\b
James Lovell (1), Backup Commander
Fred Haise (0), Backup Lunar Module Pilot
William A. Anders (1), Backup Command Module Pilot
\BMilestones\b
01/23/69 -- Hardware ondock at KSC
01/29/69 -- Command and Service Module Mated
04/14/69 -- Rollover of CSM from O&C to VAB
05/20/69 -- Rollout to Pad LC-39A
06/26/69 -- Countdown Demonstration Test
07/16/69 -- Launch
\BPayload\b
CSM-107 (Columbia) and LM-5 (Eagle)
\BMission Objective\b
Perform manned lunar landing and return mission safely. (Achieved.)
\BLaunch\b
July 16, 1969; 09:32:00 am EDT. Launch Complex 39-A Kennedy Space Center, FL. No launch delays.
\BOrbit\b
Altitude: 186km x 183km
Duration: 08 Days, 03 hours, 18 min, 35 seconds
\BLanding\b
July 24, 1969; 12:50 p.m. EDT. Splashdown area 13 deg 19 min North and 169 deg 9 min West; Splashdown at 195:18:35 MET. Crew on board U.S.S Hornet at 01:53 p.m. EDT; \Jspacecraft\j aboard ship at 03:50pm.
\BApollo 11 at Twenty-Five\b
July 1994 marks the twenty-fifth anniversary of the epochal lunar landing of Apollo 11 in the summer of 1969. Although President John F. Kennedy had made a public commitment on 25 May 1961 to land an American on the Moon by the end of the decade, up until this time Apollo had been all promise. Now the realization was about to begin. Kennedy's decision had involved much study and review prior to making it public, and his commitment had captured the American imagination, generating much positive support.
Project Apollo had originated as an effort to deal with an unsatisfactory situation (world perception of the Soviet Union's leadership in space and technology), and it addressed these problems very well. Even though Kennedy's political objectives were essentially achieved with the decision to go to the Moon, Project Apollo took on a life of its own over the years and left an important legacy to both the nation and the proponents of space exploration. Its success was enormously significant, coming at a time when American society was in crisis.
A unique confluence of political necessity, personal commitment and activism, scientific and technological ability, economic prosperity, and public mood made possible the 1961 decision to carry out an aggressive lunar landing program. It then fell to NASA, other organizations of the federal government, and the aerospace community to accomplish the task set out in a few short paragraphs by the president. By the time that the goal was accomplished in 1969, only a few of the key figures associated with the decision were still in leadership positions in the government.
Kennedy fell victim to an assassin's bullet in 1963, and science adviser Jerome B. Wiesner returned to MIT soon afterwards. Lyndon B. Johnson, of course, succeeded Kennedy as president but left office in January 1969 just a few months before the first landing. NASA Administrator James E. Webb resolutely guided NASA through most of the 1960s, but his image was tarnished by, among other things, a 1967 Apollo accident that killed three astronauts. He retired from office in October 1968. Several other early supporters of Apollo in Congress and elsewhere died during the 1960s and never saw the program successfully completed.
The first Apollo mission of public significance was the flight of Apollo 8. On 21 December 1968 it took off atop a Saturn V booster from the Kennedy Space Center. Three astronauts were aboard -- Frank Borman, James A. Lovell, Jr., and William A. Anders -- for an historic mission to orbit the Moon. At first that mission had been planned as a flight to test Apollo hardware in the relatively safe confines of low Earth orbit, but senior engineer George M. Low of the Manned \JSpacecraft\j Center at Houston, \JTexas\j, and Samuel C. Phillips, Apollo Program Manager at NASA headquarters, obtained approval to make it a circumlunar flight. The advantages of this could be important, they believed, both in technical and scientific knowledge gained as well as in a public demonstration of what the US could achieve.
After Apollo 8 made one and a half Earth orbits, its third stage began a burn to put the \Jspacecraft\j on a lunar trajectory. It orbited the Moon on 24-25 December and then fired the boosters for a return flight; it "splashed down" in the Pacific Ocean on 27 December. The public reaction to the Apollo 8 circumlunar mission was enthusiastic. It rekindled the excitement felt in the early 1960s during the first Mercury flights, and set the stage for the Apollo landing missions.
Perhaps most important, the flight was a significant accomplishment because it came at a time when American society was in crisis over Vietnam, race relations, urban problems, and a host of other difficulties. And if only for a few moments the nation united as one to focus on this epochal event. Two Apollo Earth-orbital missions occurred before the climax of the program, but they did little more than confirm that the time had come in mid-1969 for a lunar landing.
That landing came during the flight of Apollo 11, which lifted off on 16 July 1969 and, after confirmation that the hardware was working well, began the three day trip to the Moon. Then, at 4:18 p.m. EST on 20 July 1969 the Lunar Module -- with astronauts Neil A. Armstrong and Edwin E. Aldrin aboard -- landed on the surface of the Moon while Michael Collins orbited overhead in the Apollo Command Module. After checkout, Armstrong set foot on the surface, telling millions who saw and heard him on Earth that it was "one small step for man -- one giant leap for mankind."
Aldrin soon followed him out, and the two plodded around the landing site in the 1/6 lunar gravity, planted an American flag but omitted claiming the land for the US as had been routinely done during European exploration of the Americas, collected soil and rock samples, and set up scientific experiments. The next day they launched back to the Apollo capsule orbiting overhead and began the return trip to Earth, splashing down in the Pacific on 24 July.
The flight of Apollo 11 met with an ecstatic reaction around the globe, as everyone shared in the success of the mission. Ticker tape parades, speaking engagements, public relations events, and a world tour by the astronauts served to create good will both in the US and abroad.
Five more landing missions followed at approximately six month intervals through December 1972, each of them increasing the time spent on the Moon. Three of the latter Apollo missions used a lunar rover vehicle to travel in the vicinity of the landing site, but none of them equaled the excitement of Apollo 11.
The scientific experiments placed on the Moon and the lunar soil samples returned through Project Apollo have provided grist for scientists' investigations of the Solar System ever since. The scientific return was significant, but the Apollo program did not answer conclusively the age-old questions of lunar origins and evolution.
Project Apollo in general, and the flight of Apollo 11 in particular, should be viewed as a watershed in the nation's history. It was an endeavor that demonstrated both the technological and economic virtuosity of the United States and established national preeminence over rival nations -- the primary goal of the program when first envisioned by the Kennedy administration in 1961. It had been an enormous undertaking, costing $25.4 billion (about $95 billion in 1990 dollars) with only the building of the Panama Canal rivaling the Apollo program's size as the largest non-military technological endeavor ever undertaken by the United States and only the Manhattan Project to build the atomic bomb in World War II being comparable in a wartime setting.
There are several important legacies (or conclusions) about Project Apollo that need to be remembered at the twenty-fifth anniversary of the Apollo 11 landing. First, and probably most important, the Apollo program was successful in accomplishing the political goals for which it had been created. Kennedy had been dealing with a Cold War crisis in 1961 brought on by several separate factors -- the Soviet orbiting of Yuri Gagarin and the disastrous Bay of Pigs invasion only two of them -- that Apollo was designed to combat.
At the time of the Apollo 11 landing, Mission Control in Houston flashed the words of President Kennedy announcing the Apollo commitment on its big screen. Those phrases were followed with these: "TASK ACCOMPLISHED, July 1969." No greater understatement could probably have been made. Any assessment of Apollo that does not recognize the accomplishment of landing an American on the Moon and safely returning before the end of the 1960s is incomplete and inaccurate, for that was the primary goal of the undertaking.
Second, Project Apollo was a triumph of management in meeting enormously difficult systems \Jengineering\j and technological \Jintegration\j requirements. James E. Webb, the NASA Administrator at the height of the program between 1961 and 1968, always contended that Apollo was much more a management exercise than anything else, and that the technological challenge, while sophisticated and impressive, was largely within grasp at the time of the 1961 decision. More difficult was ensuring that those technological skills were properly managed and used.
Webb's contention was confirmed in spades by the success of Apollo. NASA leaders had to acquire and organize unprecedented resources to accomplish the task at hand. From both a political and technological perspective, management was critical. For seven years after Kennedy's Apollo decision, through October 1968, James Webb politicked, coaxed, cajoled, and maneuvered for NASA in Washington. In the process he acquired for the agency sufficient resources to meet its Apollo requirements.
More to the point, NASA personnel employed a "program management" concept that centralized authority over design, \Jengineering\j, procurement, testing, construction, manufacturing, spare parts, logistics, training, and operations. The systems management of the program was recognized as critical to Apollo's success in November 1968, when Science magazine, the publication of the American Association for the Advancement of Science, observed:
In terms of numbers of dollars or of men, NASA has not been our largest national undertaking, but in terms of complexity, rate of growth, and technological sophistication it has been unique. . . . It may turn out that [the space program's] most valuable spin-off of all will be human rather than technological: better knowledge of how to plan, coordinate, and monitor the multitudinous and varied activities of the organizations required to accomplish great social undertakings.
Understanding the management of complex structures for the successful completion of a multifarious task was a critical outgrowth of the Apollo effort.
Third, Project Apollo forced the people of the world to view the planet Earth in a new way. Apollo 8 was critical to this fundamental change, for on its outward voyage the crew focused a portable \Jtelevision\j camera on Earth and for the first time humanity saw its home from afar, a tiny, lovely, and fragile "blue marble" hanging in the blackness of space.
When the Apollo 8 \Jspacecraft\j arrived at the Moon on Christmas Eve of 1968 the image of Earth was even more strongly reinforced when the crew sent images of the planet back while reading the first part of the \JBible\j -- "And God created the heavens and the Earth, and the Earth was without form and void" -- before sending holiday greetings to humanity.
Writer Archibald MacLeish summed up the feelings of many people when he wrote at the time of Apollo, that "To see the Earth as it truly is, small and blue and beautiful in that eternal silence where it floats, is to see ourselves as riders on the Earth together, brothers on that bright loveliness in the eternal cold -- brothers who know now that they are truly brothers." The modern environmental movement was galvanized in part by this new perception of the planet and the need to protect it and the life that it supports.
Finally, the Apollo program, while an enormous achievement, left a divided legacy for NASA and the aerospace community. The perceived "golden age" of Apollo created for the agency an expectation that the direction of any major space goal from the president would always bring NASA a broad consensus of support and provide it with the resources and license to dispense them as it saw fit. Something most NASA officials did not understand at the time of the Moon landing in 1969, however, was that Apollo had not been conducted under normal political circumstances and that the exceptional circumstances surrounding Apollo would not be repeated.
The Apollo decision was, therefore, an anomaly in the national decision-making process. The dilemma of the "golden age" of Apollo has been difficult to overcome, but moving beyond the Apollo program to embrace future opportunities has been an important goal of the agency's leadership in the recent past. Exploration of the Solar System and the universe remains as enticing a goal and as important an objective for humanity as it ever has been. Project Apollo was an important early step in that ongoing process of exploration.
#
"Apollo 13 Mission",105,0,0,0
This mission was designed to land on the moon but after an on board explosion the only concern was how to get the astronauts back to earth safely. This aborted moon project held the world spellbound for days and finally resulted in a safe landing back on earth.
The following is a summary of the mission, the problems encountered, and the magnificent effort performed by the ground crew and the astronauts to bring everyone home safely.
\BCrew\b
James A Lovell, Jr; John L Swigert, Jr; Fred W Haise, Jr.
\BMission Objective\b
Apollo 13 was supposed to land in the Fra Mauro area. An explosion onboard forced Apollo 13 to circle the moon without landing.
The Fra Mauro site was reassigned to Apollo 14.
\BLaunch\b
Saturday, April 11, 1970 at 13:13 CST.
At five and a half minutes after liftoff, Swigert, Haise, and Lovell felt a little vibration. Then the center engine of the S-II stage shutdown two minutes early. This caused the remaining four engines to burn 34 seconds longer than planned, and the S-IVB third stage had to burn nine seconds longer to put Apollo 13 in orbit.
Ground tests before launch indicated the possibility of a poorly insulated supercritical \Jhelium\j tank in the LM's descent stage so the flight plan was modified to enter the LM three hours early in order to obtain an onboard readout of \Jhelium\j tank pressure.
\BMission Highlights:\b
Third lunar landing attempt. Mission was aborted after rupture of service module oxygen tank. Classed as "successful failure" because of experience in rescuing crew. Spent upper stage successfully impacted on the Moon.
The first two days the crew ran into a couple of minor surprises, but generally Apollo 13 was looking like the smoothest flight of the program. At 46 hours 43 minutes Joe Kerwin, the CapCom on duty, said, "The \Jspacecraft\j is in real good shape as far as we are concerned. We're bored to tears down here." It was the last time anyone would mention boredom for a long time.
At 55 hours 46 minutes, the crew finished a 49 minute TV broadcast showing how comfortably they lived and worked in weightlessness. Nine minutes later, Oxygen tank No 2 blew up, causing No 1 tank also to fail. The Apollo 13 command modules normal supply of electricity, light, and water was lost, and they were about 200,000 miles from Earth. The message came in the form of a sharp bang and vibration.
The time was 2108 hours on April 13. Next, the warning lights indicated the loss of two of Apollo 13's three fuel cells, which were the space craft's prime source of electricity. With warning lights blinking on, one Oxygen tank appeared to be completely empty, and there were indications that the oxygen in the second tank was rapidly being depleted.
The first thing the crew did, even before discovering the oxygen leak, was to try to close the hatch between the CM and the LM. They reacted spontaneously, like submarine crews, closing the hatches to limit the amount of flooding. First Jack and then Lovell tried to lock the reluctant hatch, but the stubborn lid wouldn't stay shut. Exasperated, and realizing that there wasn't a cabin leak, they strapped the hatch to the CM couch. The pressure in the No 1 oxygen tank continued to drift downward; passing 300 psi, now heading toward 200 psi.
Months later, after the accident investigation was complete, it was determined that, when No. 2 tank blew up, it either ruptured a line on the No. 1 tank, or caused one of the valves to leak. When the pressure reached 200 psi, the crew and ground controllers knew that they would lose all oxygen, which meant that the last fuel cell would also die.
At 1 hour and 29 seconds after the bang, Jack Lousma, then CapCom, said after instructions from Flight Director Glynn Lunney: "It is slowly going to zero, and we are starting to think about the LM lifeboat." Swigert replied, "That's what we have been thinking about too."
Ground controllers in Houston faced a formidable task. Completely new procedures had to be written and tested in the simulator before being passed up to the crew. The navigation problem had to be solved; essentially how, when, and in what attitude to burn the LM descent engine to provide a quick return home.
With only 15 minutes of power left in the CM, CapCom told the crew to make their way into the LM. Fred and Jim Lovell quickly floated through the tunnel, leaving Jack to perform the last chores in the Command Module. The first concern was to determine if there were enough consumables to get home? The LM was built for only a 45 hour lifetime, and it needed to be stretched to 90.
Oxygen wasn't a problem. The full LM descent tank alone would suffice, and in addition, there were two ascent engine oxygen tanks, and two backpacks whose oxygen supply would never be used on the lunar surface. Two emergency bottles on top of those packs had six or seven pounds each in them. (At LM jettison, just before re-entry, 285 pounds of oxygen remained, more than half of what was available after the explosion.)
Power was also a concern. There were 2181 \Jampere\j hours in the LM batteries. Ground controllers carefully worked out a procedure where the CM batteries were charged with LM power. All non-critical systems were turned off and energy consumption was reduced to a fifth of normal, which resulted in having 20 percent of LM electrical power left when \JAquarius\j was jettisoned. There was one electrical close call during the mission. One of the CM batteries vented with such force that it momentarily dropped off the line. Had the battery failed, there would be insufficient power to return the ship to Earth.
Water was the main consumable concern. It was estimated that the crew would run out of water about five hours before Earth reentry, which was calculated at around 151 hours. However, data from Apollo 11 (which had not sent its LM ascent stage crashing into the Moon as in subsequent missions) showed that its mechanisms could survive seven or eight hours in space without water cooling. The crew conserved water. They cut down to six ounces each per day, a fifth of normal intake, and used fruit juices; they ate hot dogs and other wet pack foods when they ate at all.
The crew became dehydrated throughout the flight and set a record that stood up throughout Apollo: Lovell lost fourteen pounds, and the crew lost a total of 315 pounds, nearly 50 percent more than any other crew. Those stringent measures resulted in the crew finishing with 282 pounds of water, about 9 percent of the total.
Removal of carbon dioxide was also a concern. There were enough \Jlithium\j \Jhydroxide\j canisters, which remove carbon dioxide from the \Jspacecraft\j, but the square canisters from the Command Module were not compatible with the round openings in the Lunar Module environmental system. There were four cartridges from the LM, and four from the backpacks, counting backups. However, the LM was designed to support two men for two days and was being asked to care for three men nearly four days. After a day and a half in the LM, a warning light showed that the carbon dioxide had built up to a dangerous level. Mission Control devised a way to attach the CM canisters to the LM system by using plastic bags, cardboard, and tape -- all materials carried on board.
One of the big questions was, "How to get back safely to Earth?". The LM navigation system wasn't designed to help in this situation. Before the explosion, at 30 hours and 40 minutes, Apollo 13 had made the normal midcourse correction, which would take it out of a free return to Earth trajectory and put it on a lunar landing course. Now the task was to get back on a free return course. The ground computed a 35 second burn and fired it 5 hours after the explosion. As they approached the Moon, another burn was computed; this time a long 5 minute burn to speed up the return home. It took place 2 hours after rounding the far side of the Moon.
The Command Module navigational platform alignment was transferred to the LM but verifying alignment was difficult. Ordinarily the alignment procedure uses an on board \Jsextant\j device, called the Alignment Optical \JTelescope\j (AOT), to find a suitable navigation star. Then with the help of the on board computer it verifies the guidance platform's alignment. However, due to the explosion, a swarm of debris from the ruptured service module made it impossible to sight real stars.
An alternate procedure was developed to use the sun as an alignment star. Lovell rotated the \Jspacecraft\j to the attitude Houston had requested and when he looked through the AOT, the Sun was just where it was expected. The alignment with the Sun proved to be less than a half a degree off. The ground and crew then knew they could do the 5 minute PC + 2 burn with assurance, and that would cut the total time of the voyage to about 142 hours.
The trip was marked by discomfort beyond the lack of food and water. Sleep was almost impossible because of the cold. When the electrical systems were turned off, the \Jspacecraft\j lost an important source of heat. The temperature dropped to 38 degrees F and condensation formed on all the walls.
A most remarkable achievement of Mission Control was quickly developing procedures for powering up the CM after its long cold sleep. Flight controllers wrote the documents for this innovation in three days, instead of the usual three months. The Command Module was cold and clammy at the start of power up. The walls, ceiling, floor, wire harnesses, and panels were all covered with droplets of water. It was suspected conditions were the same behind the panels. The chances of short circuits caused apprehension, but thanks to the safe guards built into the command module after the disastrous Apollo 1 fire in January 1967, no arcing took place.
Four hours before landing, the crew shed the service module; Mission Control had insisted on retaining it until then because everyone feared what the cold of space might do to the unsheltered CM heat shield. Photos of the Service Module showed one whole panel missing, and wreckage hanging out, it was a sorry mess as it drifted away. Three hours later the crew left the Lunar Module \JAquarius\j and then splashed down gently in the Pacific Ocean near \JSamoa\j.
\BCause of Explosion\b
The Apollo 13 malfunction was caused by an explosion and rupture of oxygen tank No. 2 in the service module. The explosion ruptured a line or damaged a valve in the No. 1 oxygen tank, causing it to lose oxygen rapidly. The service module bay No. 4 cover was blown off. All oxygen stores were lost within about 3 hours, along with loss of water, electrical power, and use of the propulsion system.
The No. 2 oxygen tank used in Apollo had originally been installed in Apollo 10. It was removed from Apollo 10 for modification and during the extraction was dropped 2 inches, slightly jarring an internal fill line. The tank was replaced with another for Apollo 10, and the exterior inspected. The internal fill line was not known to be damaged, and this tank was later installed in Apollo 13.
The oxygen tanks had originally been designed to run off the 28 volt DC power of the command and service modules. However, the tanks were redesigned to also run off the 65 volt DC ground power at Kennedy Space Center. All components were upgraded to accept 65 volts except the heater thermostatic switches, which were overlooked. These switches were designed to open and turn off the heater when the tank temperature reached 80 degrees F (Normal temperatures in the tank were -300 to -100 F).
During preflight testing, tank No. 2 showed anomalies and would not empty correctly, possibly due to the damaged fill line. It was decided to use the heater to "boil off" the excess oxygen, requiring 8 hours of 65 volt DC power. This may have damaged the thermostatically controlled switches on the heater, designed for only 28 volts. It is believed the switches may have welded shut, allowing the temperature within the tank to rise to over 1,000 degrees F. The high temperature would have resulted in damage to the Teflon \Jinsulation\j on the electrical wires to the power fans within the tank.
56 hours into the mission, at about 03:06 UT on 14 April 1970 (10:06 PM, April 13 EST), the power fans were turned on within the tank. The exposed fan wires shorted and the Teflon \Jinsulation\j caught fire. This fire spread along the wires to the electrical conduit in the side of the tank, which weakened and ruptured under the nominal 1,000 psi pressure within the tank, causing the No. 2 oxygen tank to explode. This damaged the No. 1 tank and parts of the interior of the service module and blew off the bay No. 4 cover.
#
"Remarks by the President at the 25th Anniversary of Apollo 11",106,0,0,0
\BTHE PRESIDENT:\b Thank you very much, Mr. Vice President. Members of Congress. Veterans of the Apollo program. The friends of the space program in America and, most of all, to those whom we honor here today.
Just a day before he died, President Kennedy compared our space program to a boy who comes upon a wall in an orchard. The wall is tall, it looks insurmountable, but the boy is curious about what lies on the other side. So, he throws his cap over the wall and then he has no choice but to go after it.
Twenty-five years ago today, our nation, represented by these three brave men, made that climb. And, so, today we are gathered to celebrate their voyage and I honestly hope to recommit ourselves to their spirit of discovery. Apollo 11, Neil Armstrong Buzz Aldrin and Michael Collins were our guides for the wondrous, the unimaginable at that time, the true handiwork of God. They realized the dreams of a nation. They fulfilled an American destiny. They taught us that nothing is impossible if we set our sights high enough.
Today, we're honored to have them and all the other Apollo astronauts who are here with us. For every American who followed your journey especially for those of us who were young on that fateful day 25 years ago, and for the young Americans who still dream dreams of a future in space, we thank you all.
Looking back on that mission, one thing is clear that we ought to remember today. It wasn't easy. The ship to the heavens measured just 13 feet in diameter. The destination was three days and a world away. On the third day as the tiny module descended to the Moon, it came dangerously close to a crash landing -- that happens around here all the time -- (laughter) -- but Neil Armstrong took over the controls from the computer and landed safely. Man had not been rendered obsolete by the mechanical and that hasn't happened yet.
Not long after that when he stepped on the Moon, Mr. Armstrong marked the outer limit of the human experiment with those simple words, "One small step for man. One giant leap for mankind."
These men and the other astronauts who came before and after have helped us to step into another world right here on Earth. They've shown us that we can harness the technology of space in areas from the economy to the environment to education to information and technology. The products and knowledge that grew out of our space missions has changed our way of life forever and for the better. And in our quest we have relearned a sense of confidence that has always been an essential ingredient of our American Dream.
Today that journey continues. Our commitment to the space program is strong and unwavering. The best way to honor these men and all the others who have helped it so much, is to continue that quest. Many have risked their lives and some have given their lives so that we could go forward.
Today I ask that we remember, especially, the crews of Apollo 1 and the Challenger. On this day of celebration we must never forget the deep debt we owe to those brave Americans. And our thoughts should also be with their families and their loved ones for the sacrifice they have given helped to bring us all to new horizons.
Our space explorations today are important models for cooperation in the new post-Cold War world. The Vice President described that eloquently a moment ago. Sergei's mission was an important first step toward full Russian partnership in what must be our next great mission, the international Space Station.
This permanent orbiting space laboratory, to be built with help from 14 nations, will hasten discoveries in fields from the environment to medicine, to computers. We should also remember that the space station holds great promise for us here at home, as it strengthens our largest export sector, aerospace technology.
All these reasons explain why the House has fully funded already the Space Station. I want to thank many people who are responsible for that bipartisan victory but let me mention especially George Brown, Lou Stokes, Bob Walker and Jerry Lewis. I know the Vice President and Dan Goldin and a lot of other people burned up the phone lines before the House vote.
Let me say that we've fought a lot of battles for the future around here in the last 18 months, and sometimes it seems that the most important ones are decided by the narrowest of margins. The economic plan passed by a vote, the assault weapons ban passed by two votes. Last year the Space Station survived by the vote of a single member of the House of Representatives who changed his mind on the way down the aisle. But this year, thanks to the common endeavors of all of us and thanks to the promise of cooperation with \JRussia\j and with other nations, the House of Representatives voted to fund the Space Station by 122 votes, a bipartisan commitment to America's future. (Applause.)
I thank the members of the Senate who are here today who are pushing for passage. I know they won't miss this great opportunity which is coming on them very soon. I thank you Senator Mikulski and all the other members of the Senate who are here for the work that will be done in the Senate.
As we work toward building a better world, we also have to preserve the one we've got here. William Anders of the Apollo 8 was the first to see the entire Earth at a glance. He said it looked like a fragile "little Christmas tree ornament against an infinite backdrop of space, the only color in the whole universe we could see. It seemed so very finite."
Well, because we are so very finite our responsibility to our planet must not be limited. That's why NASA's "Mission to \JPlanet\j Earth" is also a very important part of our future in space. We have to continue to monitor the global environment from space and to act on what we learn. Above all, let us never forget that all this work is about renewing our hopes and the hopes of generations to come. About the ability of Americans and the ability of human beings everywhere to conquer the seemingly impossible. I don't think anybody can look at the faces of these young people here with us today, and we ought to take a little while and look at them and welcome them here, without seeing again in their eyes dreams that those of us who are older could not have dreamed.
The explorations we continue in space are clear evidence to them that they will grow up in exciting times without limits. Times that demand their imagination, their vision, their courage. Times that will reward them, too, for believing in themselves and their possibilities.
One of our Young Astronauts, 13 year old Wayne Gusman from New Orleans, sees a future where being an \Jastronaut\j will be like, and I quote, "driving a car; everyone will do it."
That's a great dream. But that and our other dreams are clearly the natural extensions of the space program which began a generation ago, the direct descendants of the dreams of the three men we are here to honor today. We can get there.
No one who was alive then will ever forget where they were as Michael Collins traveled his solitary vigil around the Moon and Neil Armstrong and Buzz Aldrin landed that tiny craft on the surface. The world was captivated not only by the risk and the daring, although they were risking and daring, they were captivated because the landing meant again that the human experiment in conquering new and uncharted worlds was reborn.
In that sense it was not an end but a beginning. So, to you, gentlemen, we say for your valor, your courage, your pioneering spirit, and for being here today to remind us again that all things are possible, we are deeply in your debt. Thank you very much. (Applause.)
#
"Magellan Program",107,0,0,0
A US space mission, that between late 1990 and 1992 successfully mapped Venus in its entirety, using an orbiter with side-looking radar to provide sub-kilometer resolution images. Radar is used because of the global cloud cover, the created images being interpreted much as are \Jtelevision\j images. The single Magellan \Jspacecraft\j was managed by NASA's Jet Propulsion Laboratory and was launched using the NASA Shuttle in May 1989.
\BMission Details
Launch Date:\b May 4, 1989
\BDescription\b
The Magellan \Jspacecraft\j was launched on May 4, 1989, arrived at Venus on August 10, 1990 and was inserted into a near-polar elliptical orbit with a periapsis altitude of 294 km at 9.5 degrees. N. Radio contact with Magellan was lost on October 12, 1994. The primary objectives of the Magellan mission were to map the surface of Venus with a synthetic aperture radar (SAR) and to determine the topographic relief of the planet.
At the completion of radar mapping, 98% of the surface was imaged at resolutions better than 100 m, and many areas were imaged multiple times. The mission was divided up into "cycles", each cycle lasted 243 days (the time necessary for Venus to rotate once under the Magellan orbit -- i.e. the time necessary for Magellan to "see" the entire surface once.)
The mission proceeded as follows:
04 May 1989 -- Launch.
10 Aug 1990 -- Venus orbit insertion and \Jspacecraft\j checkout.
15 May 1991 -- Cycle 2: Radar mapping (right-looking).
15 Jan 1992 -- Cycle 3: Radar mapping (left-looking).
14 Sep 1992 -- Cycle 4: Gravity data acquisition.
24 May 1993 -- Aerobraking to circular orbit.
03 Aug 1993 -- Cycle 5: Gravity data acquisition.
30 Aug 1994 -- Windmill experiment.
12 Oct 1994 -- Loss of radio signal.
13 Oct 1994 -- Expected loss of \Jspacecraft\j.
A total of 4,225 usable SAR imaging orbits was obtained by Magellan. Each orbit typically covered an area 20 km wide by 17,000 km long, at a resolution of 75 m/pixel. This raw SAR data was processed into image strips called full-resolution basic image data records (F-BIDRs). Adjacent F-BIDRs were then assembled into full-resolution mosaicked image data records (F-MIDRs). These images were then compressed once (by a factor of 3), twice (9), or 3 times (27), to give C1-, C2-, and C3-MIDRs.
The Magellan mission scientific objectives were to study land forms and tectonics, impact processes, erosion, deposition, chemical processes, and model the interior of Venus. Magellan showed us an Earth-sized planet with no evidence of Earth-like plate tectonics. At least 85% of the surface is covered with volcanic flows, the remainder by highly deformed mountain belts.
Even with the high surface temperature (475║ C) and high atmospheric pressure (92 bars), the complete lack of water makes erosion a negligibly slow process, and surface features can persist for hundreds of millions of years. Some surface modification in the form of wind streaks was observed. Over 80% of Venus lies within 1 km of the mean radius of 6,051.84 km.
The mean surface age is estimated to be about 500 My. A major unanswered question concerns whether the entire surface was covered in a series of large events 500 My ago, or if it has been covered slowly over time. The gravity field of Venus is highly correlated with the surface \Jtopography\j, which indicates the mechanism of topographic support is unlike the Earth, and may be controlled by processes deep in the interior. Details of the global tectonics on Venus are still unresolved.
#
"Ulysses Project Information",108,0,0,0
A joint \JESA\j (European Space Agency)-NASA space exploration mission to observe the Sun and solar wind from a high solar latitude perspective. It requires the \Jspacecraft\j to fly on a trajectory which passes over the poles of the Sun, using a boost from Jupiter's gravity field.
Originally named the 'International Solar Polar Mission', it was planned as a two-spacecraft mission, before the cancellation of the NASA \Jspacecraft\j. It is managed by \JESA\j's Space Research and Technology Center and by NASA's Jet Propulsion Laboratory.
\BUlysses
Launch Date/Time:\b 1990-10-06 at 11:47:16 UTC
\BDescription\b
The primary objectives of Ulysses are to investigate, as a function of solar latitude, the properties of the solar wind and the interplanetary magnetic field, of galactic cosmic rays and neutral interstellar gas, and to study energetic particle composition and acceleration.
The 55 kg payload includes two magnetometers, two solar wind plasma instruments, a unified radio/plasma wave instrument, three energetic charged particle instruments, an interstellar neutral gas sensor, a solar X-ray/cosmic gamma-ray burst detector, and a cosmic dust sensor. The communications systems is also used to study the solar corona and to search for gravitational waves.
Secondary objectives included interplanetary and planetary physics investigations during the initial Earth-Jupiter phase and investigations in the Jovian magnetosphere. The \Jspacecraft\j used a Jupiter swingby in Feb. 1992 to transfer to a heliospheric orbit with high heliocentric inclination, and will pass over the rotational south pole of the sun in mid-1994 at 2 AU, and over the north pole in mid-1995.
A second solar orbit will take Ulysses again over the south and north poles in years 2000 and 2001, respectively. The \Jspacecraft\j is powered by a single radio-isotope generator. It is spin stabilized at a rate of 5 rpm and its high-gain antenna points continuously to the earth. A nutation anomaly after launch was controlled by CONSCAN. The original mission planned for two \Jspacecraft\j, one built by \JESA\j and the other by NASA. NASA canceled its \Jspacecraft\j in 1981.
#
"Viking Mission to Mars",109,0,0,0
\BDescription\b
NASA's Viking Mission to Mars was composed of two \Jspacecraft\j, Viking 1 and Viking 2, each consisting of an orbiter and a lander. The primary mission objectives were to obtain high resolution images of the Martian surface, characterize the structure and composition of the atmosphere and surface, and search for evidence of life. Viking 1 was launched on August 20, 1975 and arrived at Mars on June 19, 1976. The first month of orbit was devoted to imaging the surface to find appropriate landing sites for the Viking Landers. On July 20, 1976, Viking Lander 1 separated from the Orbiter and touched down at Chryse Planitia (22.27 N, 49.97 W, 2 km below the datum elevation).
Viking 2 was launched September 9, 1975 and entered Mars orbit on August 7, 1976. Viking Lander 2 touched down at Utopia Planitia (47.57N, 225.74 W, 3 km below the datum elevation) on September 3, 1976. The Orbiters imaged the entire surface of Mars at a resolution of 150 to 300 meters, and selected areas at 8 meters. The lowest periapsis altitude for both Orbiters was 300 km. Viking Orbiter 2 was powered down on July 25,1978 after 706 orbits, and Viking Orbiter 1 on August 17, 1980, after over 1,400 orbits.
The Viking Landers transmitted images of the surface, took surface samples and analyzed them for composition and signs of life, studied atmospheric composition and \Jmeteorology\j, and deployed seismometers. Viking Lander 2 ended communications on April 11,1980, and Viking Lander 1 on November 13, 1982, after transmitting over 1,400 images of the two sites.
The results from the Viking experiments give our most complete view of Mars to date. Volcanoes, \Jlava\j plains, immense canyons, cratered areas, wind-formed features, and evidence of surface water are apparent in the Orbiter images. The planet appears to be divisible into two main regions, northern low plains and southern cratered highlands. Superimposed on these regions are the Tharsis and Elysium bulges, which are high-standing volcanic areas, and Valles Marineris, a system of giant canyons near the equator.
The surface material at both landing sites can best be characterized as iron-rich clay. Measured temperatures at the landing sites ranged from 150 to 250 K, with a variation over a given day of 35 to 50 K. Seasonal dust storms, pressure changes, and transport of atmospheric gases between the polar caps were observed. The \Jbiology\j experiment produced no evidence of life at either landing site.
\BThe "Face on Mars"
Background: The Viking Images\b
The Viking missions to Mars in the late 1970s produced more information about the Red \JPlanet\j than had been gathered in all the previous centuries of study by Earth-bound astronomers and observers. The primary mission of the Viking program was to search for signs of life on the surface of Mars. Two landers containing sophisticated biological laboratories studied soil samples in a variety of tests which, it was hoped, would prove or disprove the existence of life.
The results of these tests indicated that Mars contained no life, at least at these landing sites. However, Viking gathered volumes of data on the weather, soil chemistry and other surface properties and mapped the surface using low-to-moderate resolution cameras on the two orbiters.
Shortly after mapping began in 1976, an interesting image taken by the Viking 1 Orbiter was received at the Jet Propulsion Laboratory, Pasadena, Calif., which contained a surface feature resembling a human or ape-like face. The photo was immediately released to the public as an interesting geological feature and dubbed the "Face on Mars." Shortly afterwards, other photos of the same area were taken, and some scientists believed that the formation appeared to be a face due to the lighting angles as seen from the Orbiter.
\BOrigin of Features Examined\b
Over the years, some people began to raise questions about the origins of the features. A few ideas and theories arose speculating that the features may have been built by aliens in the distant past. These theories are based largely on the results of computer photo enhancements, and other analytical techniques performed on the Viking images, beginning in the early 1980s.
Most planetary geologists familiar with the set of photos, however, concluded that the natural processes known to occur on Mars -- such as wind erosion, Mars quakes, and erosion from running water in the distant past -- could account for the formation of the complicated fretted terrain of the Cydonia region, including the face.
Because the entire data set includes only nine low-to-moderate resolution photos, scientists say that there just is not enough data available to justify what would be an extraordinary conclusion that the features are not natural in origin (many scientists question whether images alone would be enough to settle the matter). Such a proven discovery of extraterrestrial life or artifacts would be one of the greatest discoveries in human history, and, as such, demand the most rigorous scientific investigation.
However, despite the phenomenal nature of such a potential discovery, no one in the scientific community -- either in the US or worldwide -- has ever proposed an investigation for a mission to study these features. Until more data is gathered, many scientists consider the probability that the features are anything other than natural in origin are just too low to justify the major expenditure of public funds which such an investigation would entail (more on this below).
What is agreed on is that a greater number of high resolution images of this area should be gathered. Following the failure of the Mars Observer mission in August, 1993, NASA proposed a decade-long program of Mars exploration, including orbiters and landers. The program, called Mars Surveyor, would take advantage of launch opportunities about every 2 years to launch an orbiter and a lander to the Red \JPlanet\j.
The first mission, consisting of an orbiter to be launched in 1996, will map the surface and take high- and medium-resolution images of particular features on the Martian surface that are of high interest. NASA intends to make observations of the Cydonia region making the best effort feasible, either with the first orbiter or on follow-on missions, to obtain images of the "face" and nearby landforms.
Quite aside from the interest generated by these curious features, Cydonia has long been regarded as an area of high scientific importance, ever since the first detailed images were returned by NASA's Viking \Jspacecraft\j in the late 1970s. The Cydonia region of Mars is part of the so-called fretted terrain, a belt of landforms that circles Mars at about 30-40 degrees North Latitude. In this region, the ancient crust of Mars has been intensely eroded by weathering processes, leaving high remnants of older crust surrounded by lower plains of eroded debris.
The landforms of Cydonia resemble in some respects those of terrestrial deserts, but they probably have been shaped by a unique range of peculiarly martian agencies: wind, frost and possibly running water in ancient times. Deciphering the geological age and origin of this terrain will yield important insights into the evolution of the martian surface, into the role of ice and water in its development, and into the nature of the martian climate in times past.
\BProposing Investigations\b
The selection of goals and scientific priorities for NASA to undertake on future space science missions starts in the scientific and academic communities, as well as within NASA. Scientific associations, such as the National Academy of Science, determine the research priorities in any given field of science. For instance, the most important questions remaining about Mars include gaining an understanding of the amount of water on the planet; mapping the surface in detail to gain a complete understanding of the geological processes, history and composition; and gaining a global understanding of the atmosphere, including climate and weather.
When NASA receives permission to proceed with a science mission, the Agency publishes an Announcement of Opportunity (AO). The AO solicits interest in providing high priority scientific investigations and instruments that will be part of the new mission. The AO receives the widest possible circulation throughout the university and research communities and industry.
Proposals are submitted and reviewed through a competitive peer review process. In this process, scientists from various institutions and organizations evaluate each proposal's scientific and technical merit, and then rank the relative merit of each. NASA receives the reports of the review panels and makes a final selection as to which instruments will be built and actually flown. This rational selection process ensures that only the most useful research, with a high probability of returning good science, is done at taxpayer expense.
After selection, each Mars Surveyor Principle Investigator (PI) team will develop its instrument, build it, test it and prepare it for launch and the 10-month journey to Mars. They are also charged with developing, testing, and using the software required to properly calibrate their instrument's data. Most of the scientists working on the various Mars Surveyor missions will have several years invested in their instrument before the \Jspacecraft\j arrives at Mars and they can actually receive the bulk of the data they have been waiting for.
\BObtaining Images of the "Face" and other Planetary Data\b
Since the release and subsequent widespread circulation of the "face" images, scientists and individual members of the public have freely drawn their own conclusions about the nature and origin of this feature. NASA encourages anyone seriously interested in this topic to obtain the photo(s) and decide for themselves, just as every day many hundreds of independent researchers and scientists make use of NASA-provided data on a variety of subjects.
The most noteworthy image of the "face" feature is available to the public, for a nominal fee, through Headquarters and JPL. A photo catalogue can be provided to select images.
All imaging data obtained by the Mars Surveyor program, as well as other types of data, will be deposited in open data \Jarchives\j. Two such \Jarchives\j widely used are the Planetary Data System (PDS), an open archive accessible to thousands of scientists and other individuals, and the National Space Science Data Center (NSSDC) where images and other data will be readily available to the general public (generally on CD-ROMs or as hard copy, as appropriate), for a nominal charge that covers the materials and time needed to produce the copies. For information about ordering copies of NASA science mission images, including on CD-ROM format, contact the NSSDC at:
National Space Science Data Center
Request Coordination Center
Goddard Space Flight Center
Greenbelt, MD 20771
Telephone (301) 286-6695
#
"Skylab Program",110,0,0,0
\JSkylab Introduction\j
\JSkylab I\j
\JSkylab II\j
\JSkylab III\j
#
"Skylab Introduction",111,0,0,0
\BLaunch Date/Time:\b 1973-05-14 at 00:00:00 UTC
\BDescription\b
The Skylab (SL) was a manned, orbiting \Jspacecraft\j composed of five parts, the Apollo \Jtelescope\j mount (ATM), the multiple docking adapter (MDA), the airlock module (AM), the instrument unit (IU), and the orbital workshop(OWS). The Skylab was in the form of a cylinder, with the ATM being positioned 90 degrees from the longitudinal axis after insertion into orbit. The ATM was a solar observatory, and it provided attitude control and experiment pointing for the rest of the cluster. It was attached to the MDA and AM at one end of the OWS. The retrieval and installation of film used in the ATM was accomplished by astronauts during extravehicular activity(EVA).
The MDA served as a dock for the command and service modules, which served as personnel taxis to the Skylab. The AM provided an airlock between the MDA and the OWS, and contained controls and instrumentation. The IU, which was used only during launch and the initial phases of operation, provided guidance and sequencing functions for the initial deployment of the ATM, solar arrays, etc. The OWS was a modified Saturn 4B stage suitable for long duration manned habitation in orbit. It contained provisions and crew quarters necessary to support three-person crews for periods of up to 84 days each.
All parts were also capable of unmanned, in-orbit storage, reactivation, and reuse. The Skylab itself was launched on May 14, 1973. It was first manned during the period May 25 to June 22, 1973, by the crew of the SL-2 mission (73-032A). Next, it was manned during the period July 28 to September 25, 1973, by the crew of the SL-3 mission (73-050A). The final manned period was from November 16, 1973, to February 8, 1974, when it was manned by the crew from the SL-4 mission.
#
"Skylab I",112,0,0,0
NASA had studied various concepts for a space station, including inflatable donuts, Chesley Bonestell's magnificent "Wheel," and various other designs since the earliest beginnings of the space program. When the Saturn rocket was developed in the mid-'60s, enabling some heavy lifting into space, the Skylab Program began to take shape. Following cancellation of Apollo 18, 19 and 20, we had a lot of hardware lying around gathering dust, so we put it to some remarkably good use.
At first there were two competing concepts for a space station. The first, called the "Wet Concept," called for launching a Saturn 1B, venting the S IV-B upper stage and refurbishing it, converting it to the space station, while in orbit. The second, or "Dry Concept," called for outfitting the S IV-B while still on the ground and launching it atop a Saturn V. While the Apollo 11 astronauts were actually on the moon in July, 1969, the decision was made to go with the Dry Concept.
Skylab weighed about 100 tons, and its launch marked the last launch of the wonderful Saturn V, the rocket that never failed. It had a volume of 283.17 cubic meters and was separated into two "floors;" the "upper" floor contained storage lockers and a large empty space for conducting experiments, and two airlocks, one pointed "down" toward the earth and the other "up" toward the sun; the "lower" floor was divided into rooms including a dining room with a table, three bedrooms, a work area, a bathroom and a shower. The floors consisted of an open gridwork that fit cleats on the bottom of the astronauts' shoes.
The station was also equipped with an airlock module for the many spacewalks that were required to change film in the external cameras and make repairs to the station. The Apollo Command and Service Modules remained attached to the station's docking mechanism for the duration of the astronauts' stays aboard the station.
The largest piece of scientific equipment was the "Apollo \JTelescope\j Mount" or ATM, which had its own solar panels for electricity generation and was used to make apectrographic analyses of the Sun without interference from Earth's atmosphere.
Launch of the unoccupied Skylab, designated Skylab 1 (the occupied missions were officially designated Skylabs 2, 3 and 4, but are generally referred to as Skylabs I, II and III, and are referred to in that manner here) took place on May 14, 1973, and problems set in early on.
One minute and three seconds into launch, the \Jmeteorite\j shield/sunshade was torn loose by the aerodynamic forces, destroying one of the solar arrays and damaging the other. The first crew, Skylab I, which was supposed to launch the next day, was delayed for ten days while mission personnel devised a method to repair the crippled station.
In all, three crews were launched, each in turn setting a record for longest human space flight; the all-time American record, which stood until Norm Thagard broke it aboard Mir in 1995 (and now held by Shannon Lucid), was set by Skylab III at over 2,017 hours (three months) and 1,214 orbits of the earth.
The Skylab program totalled 513 man-days in orbit, conducted thousands of experiments in many different disciplines, and even viewed the rather disappointing \Jcomet\j Kohoutek from Skylab III.
Skylab's orbit slowly deteriorated and it finally burned up in the atmosphere on July 11, 1979, more than five years after the last crew left for home.
#
"Skylab II",113,0,0,0
Launched: July 28, 1973
Splashed Down: September 25, 1973
Duration: 1,427 hours, 9 minutes, 4 seconds
Orbits: 858
Crew:
ò Alan L. Bean
ò Jack R. Lousma
ò Owen K. Garriott
This mission once again more than doubled the previous endurance record in space, just set by the astronauts of Skylab I just a month earlier. After an early bout with motion sickness, the crew settled down for their two-month mission, deploying a second sun shield on a space walk lasting six hours, 30 minutes.
They conducted many experiments, and brought with them live spiders to conduct a student-designed experiment to see what kinds of webs the spiders would spin in weightlessness.
Also on this mission the astronauts finally got to test the \JAstronaut\j Maneuvering Unit, or AMU, which had initially been carried into space aboard Gemini IX but could not be tested then because of problems with the old Gemini space suit.
The AMU experiments assisted engineers in designing the Manned Maneuvering Unit, which was first flown aboard the Shuttle flight STS 41B in February, 1984, and was still in use until quite recently.
#
"Skylab III",114,0,0,0
Launched: November 16, 1973
Splashed Down: February 8, 1974
Duration: 2,017 hours, 15 minutes, 31 seconds
Orbits: 1,214
Crew
ò Gerald P. Carr
ò William R. Pogue
ò Edward C. Gibson
At 84 days, 1 hour, 15 minutes and 31 seconds, Skylab III (actually officially designated "Skylab 4") remains by far the longest American space flight, a record that will certainly stand until the permanent human occupation of space begins with the international Space Station.
To help keep the crew in physical condition during their almost three months in orbit, they walked treadmills and rode an on-board stationary \Jbicycle\j, and came home in far better condition than had the previous Skylab crews.
Among the thousands of experiments they conducted during this long, long flight, the astronauts took four space walks, including one on Christmas Day to observe the \Jcomet\j Kohoutek.
#
"Galileo Project Information",115,0,0,0
\JGalileo Project Summary\j
\JCallisto Summary\j
\JGalilean Satellites (Jupiter)\j
\JEuropa, Natural and False Color Views\j
\JEuropa's Surface Reshaped\j
\JGalileo Returns To Europa For Another Close Look\j
\JGalileo Mission Status (1)\j
\JGalileo Mission Status (2)\j
\JGalileo Mission Status (3)\j
\JGalileo Images Hint At History For Europa\j
\JGalileo Mission Status (4)\j
\JGalileo Mission Status (5)\j
\JGalileo To Take One Last Close Look At Ganymede\j
\JGalileo Returns New Insights Into Callisto and Europa\j
\JJupiter's Dry Spots and Glowing Auroras To Be Unveiled\j
\JGalileo Finds New View of Jupiter's Light Show\j
\JGalileo Finds Europa Has An Atmosphere\j
\JGanymede G1 & G2 Encounters\j
\JGanymede, Fractures in Transitional Terrain\j
\JGalileo Calendar of Events\j
#
"Galileo Project Summary",116,0,0,0
\BOrbiter Launch Date:\b 12 October 1989
\BProbe Release Date:\b 13 July 1995
\BDescription\b
The \JGalileo\j \Jspacecraft\j flew by the Earth and Moon on Dec. 8, 1990 and Dec. 7, 1992. The primary mission of the \JGalileo\j orbiter and probe is to explore Jupiter and its satellites. Due to the great distance to Jupiter of over 600 million kilometers, and the on board fuel limitations, a series of planetary flybys has taken place in order to give \JGalileo\j a gravity assist to Jupiter:
Launch: 12 Oct 1989
Venus: 10 Feb 1990
Earth/Moon 1: 08 Dec 1990
Gaspra: 29 Oct 1991
Earth/Moon 2: 08 Dec 1992
Ida: 28 Aug 1993
Jupiter arrival: 07 Dec 1995
These flybys gave \JGalileo\j an opportunity to image the Moon at various wavelengths with the Solid State Imaging (SSI) camera. The camera uses a high-resolution, 800 x 800 charge-coupled device (CCD) array with a field of view of 0.46 degrees. Multi-spectral coverage is provided by an eight-position filter wheel on the camera, consisting of three broad-band filters: violet (404 nm), green (559 nm), and red (671 nm); four near-infrared filters: 727 nm, 756 nm, 889 nm, and 986 nm; and one clear filter (611 nm) with a very broad (440 nm) passband.
#
"Callisto Summary",117,0,0,0
With a diameter of over 4,800 km (2,985 miles), Callisto is the third largest satellite in the solar system (only Ganymede and Titan are bigger), and is almost the size of Mercury. Callisto is the outermost of the Galilean satellites, and orbits beyonds Jupiter's main radiation belts.
Callisto has the lowest density of the Galilean satellites (1.86 grams/cubic centimeter). Its interior is probably similar to Ganymede except the inner rocky core is smaller, and this core is surrounded by a large icy mantle. Callisto's surface is the darkest of the Galileans, but it is twice as bright our our own Moon.
Callisto is the most heavily cratered object in the solar system. It is thought to be a long dead world, with a nearly complete absence of any geologic activity on its surface. In fact, Callisto is the only body greater than 1000 km in diameter in the solar system that has shown no signs of undergoing any extensive resurfacing since impacts have molded its surface. With a surface age of about 4 billion years, Callisto has the oldest landscape in the solar system.
There are no large mountains on Callisto, which is probably due to the icy nature of the satellite's surface. The surface features are dominated by impact craters and rings, and the craters are quite shallow. There are two large "bulleye" structure on Callisto, thought to be the result of a massive impact. The largest structure, \JValhalla\j, has a bright patch 600 km across with rings extending out to almost 3000 km. The other ring structure, Asgard, is about 1600 km in diameter.
Seven impact crater chains have been mapped on Callisto. These chains probably formed when fragments of a \Jcomet\j were split apart by Jupiter's gravity and impacted on Callisto. In a similar scenario, \JComet\j Shoemaker-Levy 9 split into 21 fragments and impacted Jupiter in 1994.
No atmosphere has been detected on Callisto.
The \JGalileo\j \Jspacecraft\j will make three close passes by Callisto during its 2 year orbital tour around Jupiter. The first close encounter occurred November 4, 1996.
Callisto Quick-Look Statistics: (Refer to Table)
#
"Galilean Satellites (Jupiter)",118,0,0,0
This composite includes the four largest moons of Jupiter which are known as the Galilean satellites. From left to right, the moons shown are Ganymede, Callisto, Io, and Europa. The Galilean satellites were first seen by the Italian astronomer \JGalileo\j Galilei in 1610. In order of increasing distance from Jupiter, Io is closest, followed by Europa, Ganymede, and Callisto.
The order of these satellites from the planet Jupiter helps to explain some of the visible differences among the moons. Io is subject to the strongest tidal stresses from the massive planet. These stresses generate internal heating which is released at the surface and makes Io the most volcanically active body in our solar system.
Europa appears to be strongly differentiated with a rock/iron core, an ice layer at its surface, and the potential for local or global zones of water between these layers. Tectonic resurfacing brightens terrain on the less active and partially differentiated moon Ganymede. Callisto, furthest from Jupiter, appears heavily cratered at low resolutions and shows no evidence of internal activity.
North is to the top of this composite picture in which these satellites have all been scaled to a common factor of 10 kilometers (6 miles) per picture element.
The Solid State Imaging (CCD) system aboard NASA's \JGalileo\j \Jspacecraft\j obtained the Io and Ganymede images in June 1996, while the Europa images were obtained in September 1996. Because \JGalileo\j focusses on high resolution imaging of regional areas on Callisto rather than global coverage, the portrait of Callisto is from the 1979 flyby of NASA's Voyager \Jspacecraft\j.
Launched in October 1989, \JGalileo\j entered orbit around Jupiter on December 7, 1995. The \Jspacecraft\j's mission is to conduct detailed studies of the giant planet, its largest moons and the Jovian magnetic environment.
#
"Europa, Natural and False Color Views",119,0,0,0
November 12, 1996
The images on this page show different views of the ice-covered satellite, Europa. Dark brown areas represent rocky material derived from the interior, implanted by impact, or from a combination of interior and exterior sources. Bright plains in the polar areas (top and bottom) are shown in tones of blue to distinguish possibly coarse-grained ice (dark blue) from fine-grained ice (light blue).
Long, dark lines are fractures in the crust, some of which are more than 3,000 kilometers (1,850 miles) long. The bright feature containing a central dark spot in the lower third of the image is a young impact crater some 50 kilometers (31 miles) in diameter. This crater has been provisionally named 'Pwyll' for the Celtic god of the underworld.
Europa is about 3,160 kilometers (1,950 miles) in diameter, or about the size of Earth's moon. This image was taken on September 7, 1996, at a range of 677,000 kilometers (417,900 miles) by the solid state imaging \Jtelevision\j camera onboard the \JGalileo\j \Jspacecraft\j during its second orbit around Jupiter. The image was processed by Deutsche Forschungsanstalt fuer Luft- und Raumfahrt e.V., Berlin, \JGermany\j.
Launched in October 1989, \JGalileo\j entered orbit around Jupiter on December 7, 1995. The \Jspacecraft\j's mission is to conduct detailed studies of the giant planet, its largest moons and the Jovian magnetic environment.
#
"Europa's Surface Reshaped",120,0,0,0
January 17, 1997
Ice-spewing volcanoes and the grinding and tearing of tectonic plates have reshaped the chaotic surface of Jupiter's frozen moon Europa, images from NASA's \JGalileo\j \Jspacecraft\j reveal.
\BFlows on Europa\b
The images, captured when \JGalileo\j flew within just 430 miles (692 kilometers) of Europa on Dec. 19, were released at a news briefing today at NASA Headquarters, Washington, DC.
Although the images do not show currently active ice volcanoes or geysers, they do reveal flows of material on the surface that probably originated from them, said \JGalileo\j imaging team member Dr. Ronald Greeley of \JArizona\j State University, Tempe.
"This is the first time we've seen actual ice flows on any of the moons of Jupiter," said Greeley. "These flows, as well as dark scarring on some of Europa's cracks and ridges, appear to be remnants of ice volcanoes or geysers."
The new images appear to enhance Europa's prospects as one of the places in the Solar System that could have hosted the development of life, said Greeley.
"There are three main criteria to consider when you are looking for the possibility of life outside the Earth -- the presence of water, organic compounds and adequate heat," said Greeley. "Europa obviously has substantial water ice, and organic compounds are known to be prevalent in the Solar System. The big question mark has been how much heat is generated in the interior.
"These new images demonstrate that there was enough heat to drive the flows on the surface. Europa thus has a high potential to meet the criteria for exobiology," Greeley added.
"This doesn't prove that there is an ocean down there under the surface of Europa, but it does demonstrate that it is a scientifically exciting place," said \JGalileo\j imaging team member Dr. Robert Sullivan, also of \JArizona\j State University.
\BGalileo Image of Europa\b
The images also reveal a remarkable diversity in the geological age of various regions of Europa's surface. Some areas appear relatively young, with smooth, crater-free terrain, while others contain large craters and numerous pits, suggesting that they are much older.
The icy crust bears the signs of having been disrupted by the motion of tectonic plates. "There appear to be signs of different styles of tectonism," said Greeley. "In many areas we see that the crust was pulled apart in a spreading similar to the processes on the sea floor on Earth. This is different from the tectonic processes at work on, say, Jupiter's moon Ganymede. This suggests that Europa's interior may be different from Ganymede's."
Galileo scientists will have a better chance to understand Europa's interior when the \Jspacecraft\j gathers gravity data on another flyby next November. The gravity field is measured by tracking how the frequency of \JGalileo\j's radio signal changes as it flies past the moon. This was not possible during the recent flyby because radio conditions were degraded as Jupiter passed behind the Sun from Earth's point of view.
\BRidges on Europa\b
Europa is crisscrossed by an amazingly complex network of ridges, according to Sullivan. "Ridges are visible at all resolutions," he explained. "Closely paired ridges are most common. With higher resolution, ridges seen previously as singular features are revealed to be double."
Some of the ridges may have formed by tension in the icy crust: as two plates pull apart slightly, warmer material from below might push up and freeze to form a ridge. Other ridges may have been formed by compression: as two plates push together, the material where they meet might crumple to form the ridge.
In addition to ice flows and tectonics, Greeley and Sullivan noted that some areas on Europa seem to have been modified by unknown processes that scientists are still debating. Greeley said that some areas, for example, seem to have been modified by "sublimation erosion" -- the \Jevaporation\j of water and other volatiles such as ammonia and \Jmethane\j into the vacuum of space. "Something is destroying the topography," said Greeley, "and this sublimation erosion is a good candidate for what is at work."
During last month's encounter, \JGalileo\j flew more than 200 times closer to Europa than the Voyager 2 \Jspacecraft\j did in 1979. After a swing past Jupiter next week in what mission engineers call a "phasing orbit," \JGalileo\j's next targeted flyby will take it again past Europa as it passes within 364 miles (587 kilometers) on Feb. 20.
#
"Galileo Returns To Europa For Another Close Look",121,0,0,0
February 19, 1997
NASA's \JGalileo\j \Jspacecraft\j will make an encore appearance at Jupiter's icy moon, Europa, on Thursday, Feb. 20, marking the closest planned Europa flyby of the initial two-year mission.
The encounter will be \JGalileo\j's closest flyby yet of Europa. The craft will swoop past the Jovian moon at an altitude of 580 kilometers (360 miles) on Thursday, Feb. 20, at 9:06 a.m. Pacific time (12:06 p.m. Eastern time).
Galileo made its first pass of Europa in December 1996, revealing remarkable detail of that moon's terrain. This week's flyby will look at other areas of Europa's surface, which is covered by ice and a series of criss-crossed, dark lines. Europa holds great fascination for scientists because of the possibility that liquid oceans may be hidden underneath the icy surface. The presence of liquid water would boost the odds that Europa could host some form of life.
"I think this flyby may provide additional clues regarding the prospect of liquid water oceans on Europa," said \JGalileo\j Mission Director Bob Mitchell.
With its diameter of 3,138 kilometers (1,946 miles), Europa is just slightly smaller than Earth's moon. Because the \Jgeometry\j of the upcoming flyby will be somewhat different from the path taken by \JGalileo\j's previous Europa encounter, it will yield data and images of different portions of the moon.
"This position will allow for high resolution of different terrain," said Mitchell. "It will help us learn more about Europa's structure and surface and how the surface was formed."
The current Europa encounter phase began on Sunday, Feb. 16, and will continue through Saturday, Feb. 22. The \Jspacecraft\j has already begun returning real-time encounter data, with recorded data scheduled to be transmitted to Earth beginning on the evening of Saturday, Feb. 22 (Pacific time).
This encounter will include the return of magnetospheric measurements from Europa's vicinity. Other science highlights will include the study of surface features of Europa's lineated regions, images of two other, smaller Jovian moons, Thebe and Amalthea, and studies of such Jovian atmospheric features as the south equatorial belt-zone boundary and the aurora borealis.
This flyby provides a period of radio \Joccultation\j, when Europa crosses between Earth and \JGalileo\j, temporarily cutting off the \Jspacecraft\j's radio signal. This affords a prime opportunity for \JGalileo\j to study atmospheric data just before and after radio contact is lost, when the signal passes through the Europa's atmosphere.
"As the fifth encounter in \JGalileo\j's series of 10 flybys, this marks the approximate halfway point for this series, which began in June 1996," said \JGalileo\j Project Manager Bill O'Neil. "It's been eight months since then, and it will be another eight months before the series' final encounter."
A third Europa flyby is planned for Nov. 6, 1997, and JPL has asked NASA to extend the \JGalileo\j mission by two years to include eight more Europa flybys and ultimately a flyby of Io. The proposed extended mission might be shortened if the \Jspacecraft\j's operations were to deteriorate as a result of its continuous exposure to Jupiter's extreme radiation environment.
"NASA has assured us that the extended mission will be funded," said O'Neil. "The $30 million needed for the extension will come from within the existing NASA budget, enabled by cost savings due to improved efficiencies in JPL's \Jspacecraft\j tracking and mission operations."
The 2,223-kilogram (2-1/2 ton) \JGalileo\j orbiter \Jspacecraft\j was launched aboard Space Shuttle Atlantis on October 18, 1989.
#
"Galileo Mission Status (1)",122,0,0,0
February 27, 1997
NASA's \JGalileo\j \Jspacecraft\j has begun returning data from its latest flyby of Jupiter's icy moon, Europa, which took place last Thursday, Feb. 20. The \Jspacecraft\j's closest approach to Europa was at 9:06 a.m. Pacific time, with confirmation received on Earth at 9:56 a.m. \JGalileo\j's flyby was within one kilometer of the target altitude of 586 kilometers.
Galileo engineers will send commands to the \Jspacecraft\j tonight to correct a magnetometer software glitch that was discovered following the flyby. Although the problem caused the loss of flyby magnetometer data, similar data was collected during \JGalileo\j's previous Europa flyby and more will be gathered during the third encounter of that moon in November. Engineers expect the problem to be resolved within a few more days.
Recorded data playback began last Saturday evening and will continue through March 28, with initial data to include recent observations of Jupiter and its moon, Io. Transmission of the new Europa observations will begin later this week and continue through next week. New Europa images may be released within a few weeks.
This flyby, the closest of three Europa encounters for \JGalileo\j's primary mission, recorded data on a linea region which may represent a stress-controlled eruption or intrusion of material from beneath Europa's icy surface. This could be another sign of liquid oceans hidden underneath.
Scientists gathered data about Europa's atmosphere during the flyby when Europa crossed between Earth and \JGalileo\j, temporarily cutting off the \Jspacecraft\j's radio signal. This radio \Joccultation\j experiment uses the disturbances to the \Jspacecraft\j radio signal as it passes close to Europa to infer information about Europa's tenuous atmosphere.
Data to be returned this week include observations of Jupiter's "white ovals" -- huge storms fueled by an unknown energy source; surface changes and volcanic plume activity on Jupiter's moon, Io; and an image of one of Jupiter's smaller moons, Thebe.
Galileo has five more encounters of Jupiter's moons on its two-year primary mission journey through the Jovian system. The next destination is Jupiter's largest moon, Ganymede on Apr. 5, 1997. A planned follow-on mission would run from December 1997 through December 1999 and would include 15 encounters.
#
"Galileo Mission Status (2)",123,0,0,0
March 13, 1997
Playback from the latest Europa encounter on February 20 is proceeding on schedule, returning data from \JGalileo\j's most recent pass close to Jupiter and its satellites. The data include observations of white oval atmospheric features taken by the \Jspacecraft\j's near infrared mapping spectrometer and photopolarimeter-radiometer over a range of solar phase angles.
This week's data return will also include fields and particles instruments' high data rate recording as \JGalileo\j made its closest approach while flying through Jupiter's magnetic equator. Much of the plasma in Jupiter's inner magnetosphere is confined to the equatorial region and is whisked away to the outer boundaries of the magnetosphere through various processes. This data will help scientists understand more about those processes.
Other observations expected to be returned this week include chemical monitoring of volcanic hot spots on Io and a surface map from the Europa encounter. Ganymede observations to be returned are part of multi-orbit efforts to characterize the moon's surface and flesh out information obtained when the Voyager \Jspacecraft\j flew by the Jovian system. New observations of the small moon Amalthea will be used to determine the body's global shape and morphology.
Last week's playback focused on Europa observations, the prime target of last month's flyby. During the approach to Europa, most of the satellite was illuminated by the Sun. This vantage point enabled \JGalileo\j to gather information on the icy moon's surface composition and shape, as well as crater feature observations which may offer clues to what lies underneath Europa's surface. Near infrared mapping spectrometer data included observations of a lineated region and icy regions of varying ages.
A previous magnetometer glitch on \JGalileo\j has been corrected, and it's now believed the failure was caused by radiation effects. Although there are indications that the same radiation-induced faults have occurred with the magnetometer and the spectrometer, in each case, a reloading procedure has corrected the problem.
An orbital trim maneuver will begin today to put \JGalileo\j on track for its next destination, Ganymede. The flyby occurs on April 4 (PST). \JGalileo\j has five additional encounters of Jupiter's moons scheduled during its two-year primary journey through the Jovian system.
#
"Galileo Mission Status (3)",124,0,0,0
March 27, 1997
This is the final week for playback of data gathered during \JGalileo\j's February 20th Europa encounter. Included in the potpourri of Europa information is data on surface composition, an observation expected to help scientists distinguish between new and old ice, images of bright plains and craters that may help explain the formation of these features, and a fields and particles observation of Europa's interaction with Jupiter's magnetosphere.
The playback also features Io observations of surface chemistry, volcanic activity, and the surface while eclipsed from the Sun. Observations of two of Jupiter's white ovals, global observations of Ganymede and Callisto, and a single Amalthea observation have also be transmitted.
With the wrap-up of Europa 6 playback, the end of this week marks the start of preparations for \JGalileo\j's next encounter, a Ganymede flyby at 11:11 p.m. Pacific Standard Time on Friday, April 4. An orbital trim maneuver is planned for Monday, March 31, one day after the beginning of the Ganymede encounter period. Maintenance activities will also be performed, including the flushing of the thruster lines to prevent debris blockage, and conditioning of the tape recorder.
After this next Ganymede flyby, \JGalileo\j has four more encounters of Jupiter's moons scheduled during its two-year primary journey through the Jovian system. A planned two-year continuation of the mission, referred to as the \JGalileo\j Europa Mission (GEM), will include eight more Europa flybys and an Io flyby, as long as the \Jspacecraft\j remains healthy.
#
"Galileo Images Hint At History For Europa",125,0,0,0
April 9, 1997
Chunky ice rafts and relatively smooth, crater-free patches on the surface of Jupiter's frozen moon Europa suggest a younger, thinner icy surface than previously believed, according to new images from \JGalileo\j's \Jspacecraft\j released today.
The images were captured during \JGalileo\j's closest flyby of Europa on February 20, when the \Jspacecraft\j came within 586 kilometers (363 miles) of the Jovian moon. These features, which lend credence to the idea of hidden, subsurface oceans, are also stirring up controversy among scientists who disagree about the age of Europa's surface.
Dr. Ronald Greeley, an \JArizona\j State University geologist and \JGalileo\j imaging team member, said the ice rafts reveal that Europa had, and may still have, a very thin ice crust covering either liquid water or slush.
"We're intrigued by these blocks of ice, similar to those seen on Earth's polar seas during springtime thaws," Greeley said. "The size and \Jgeometry\j of these features lead us to believe there was a thin icy layer covering water or slushy ice, and that some motion caused these crustal plates to break up."
"These rafts appear to be floating and may, in fact, be comparable to icebergs here on Earth," said another \JGalileo\j imaging team member, Dr. Michael Carr, a geologist with the U.S. Geological Survey. "The puzzle is what causes the rafts to rotate. The implication is that they are being churned by convection."
The new images of Europa's surface have also sparked a lively debate among scientists. \JGalileo\j imaging team member Dr. Clark Chapman is among those who believe the smoother regions with few craters indicate Europa's surface is much younger than previously believed.
In essence, Chapman, a planetary scientist at Southwest Research Institute, Boulder, CO, believes the fewer the craters, the younger the region. Chapman based his estimate on current knowledge about cratering rates, or the rate at which astronomical bodies are bombarded and scarred by hits from comets and \Jasteroids\j.
"We're probably seeing areas a few million years old or less, which is about as young as we can measure on any planetary surface besides Earth," said Chapman. "Although we can't pinpoint exactly how many impacts occurred in a given period of time, these areas of Europa have so few craters that we have to think of its surface as young."
Chapman added, "Europa's extraordinary surface \Jgeology\j indicates an extreme youthfulness -- a very alive world in a state of flux."
However, Carr sees things differently. He puts Europa's surface age at closer to one billion years old.
"There are just too many unknowns," Carr said. "Europa's relatively smooth regions are most likely caused by a different cratering rate for Jupiter and Earth. For example, we believe that both Earth's moon and the Jovian moon, Ganymede, have huge craters that are 3.8 billion years old. But when we compare the number of smaller craters superimposed on these large ones, Ganymede has far fewer than Earth's moon. This means the cratering rate at Jupiter is less than the cratering rate in the Earth-moon system."
Scientists hope to find answers to some of the questions surrounding Europa and its possible oceans as the \JGalileo\j \Jspacecraft\j continues its journey through the Jovian system.
"We want to look for evidence of current activity on Europa, possibly some erupting geysers," Greeley said. "We also want to know whether Europa's surface has changed since the Voyager \Jspacecraft\j flyby in 1979, or even during the time of the \JGalileo\j flybys."
The craft will return for another Europa flyby on November 6, 1997, the final encounter of \JGalileo\j's primary mission. However, eight more Europa flybys are planned as part of \JGalileo\j's two-year extended mission, which will also include encounters with two other Jovian moons, Callisto and Io.
#
"Galileo Mission Status (4)",126,0,0,0
April 17, 1997
The \JGalileo\j \Jspacecraft\j is operating normally in its second week of "cruise" following the craft's latest encounter with Ganymede at 11:10pm Pacific Standard Time on April 4, with the signal received on the ground 46 minutes later. \JGalileo\j flew by the satellite at an altitude of 3,102 kilometers.
The fields and particles survey of Jupiter's magnetosphere continues. Other scheduled activities include one of the periodic \Jspacecraft\j turns to keep the antenna pointed near Earth, and transmission of commands to prepare for next week's orbit adjustment.
This week's playback includes observations taken by the \Jspacecraft\j during its non-targeted flyby of the increasingly- popular moon Europa. The playback includes a near-infrared mapping spectrometer (NIMS) observation at regional resolution, part of a plan to map all the Galilean satellites.
Another spectrometer observation was designed to look for differences in the mineral composition of the Tyre Macula region, a circular feature. Return from other instruments will include thermal observations and images of crater features near the terminator, the dividing line between day and night.
The playback schedule also features the return of observations of Jupiter's Great Red Spot, a smaller red spot, and a hot spot near the same latitude as the atmospheric probe entry site.
Galileo will return for another Ganymede flyby on May 7, with two Callisto encounters and another Europa flyby planned for the final orbits of \JGalileo\j's primary mission. A planned two- year continuation of the mission, known as the \JGalileo\j Europa Mission (GEM), will include eight more Europa flybys and one or two Io flybys, as long as the \Jspacecraft\j remains healthy.
#
"Galileo Mission Status (5)",127,0,0,0
April 28, 1997
As the \JGalileo\j \Jspacecraft\j prepares for another encounter with Jupiter's largest moon, Ganymede, playback of data from the craft's previous Ganymede flyby on April 4 Pacific Standard Time is nearing completion.
This final batch of data return from the April flyby will conclude around noon Pacific Daylight Time on Saturday, May 3. It includes observations of Ganymede's bright, dark and dark- rayed regions from the remote sensing instruments, and high resolution fields and particles data on the magnetospheres around Jupiter and Ganymede and the interaction between the two.
Playback also includes observations of Europa's lineated circular regions. This will be used to help determine how these features originated and to construct a global map of Europa at regional resolution. Regional observations of Jupiter to be returned this week will also be used for a global map of Jupiter.
A few Jupiter observations from the remote sensing instruments will be returned, including a study of a small red spot in the Jovian atmosphere, a hot spot on the planet and a single image of Adrastea, one of Jupiter's small inner moons. The fields and particles instruments' survey of Jupiter's magnetosphere will resume on Friday, marking the start of the second magnetospheric "mini-tour."
The \JGalileo\j flight team will transmit the first set of encounter sequence commands to the \Jspacecraft\j later this week, as it prepares for the next encounter with Ganymede on May 7. This will be \JGalileo\j's final close flyby of Ganymede.
#
"Galileo To Take One Last Close Look At Ganymede",128,0,0,0
May 6, 1997
NASA's \JGalileo\j \Jspacecraft\j will fly by Jupiter's largest moon, Ganymede, for the fourth and final time on Wednesday, May 7.
The closest approach will take place at 8:56 a.m. Pacific Daylight Time as the craft travels 1,600 kilometers (994 miles) above Ganymede at a speed of 8.6 kilometers per second (more than 19,000 miles per hour).
During the encounter, \JGalileo\j will collect data on the moon's surface shape and atmosphere. High resolution studies by the craft's remote-sensing instruments will include observations of \JOsiris\j, a dome structure; Tiamat Sulcus, a region of craters, grooves and furrows; a multi-ringed structure; and caldera-like features and dark floor craters.
Galileo will also begin its second "mini-tour" of the Jovian magnetosphere to learn more about the composition and dynamics of this tremendously vast region around Jupiter controlled by the Jovian magnetic field. This second "mini-tour" will continue until the end of \JGalileo\j's primary mission on Dec. 7. The tour, to take place this summer, will include a deep penetration into Jupiter's magnetotail, the region of the magnetosphere opposite the Sun's direction.
During this encounter, Ganymede will block the \Jspacecraft\j from the Earth and the Sun for about seven minutes. This will provide scientists with an opportunity to measure changes in the \Jspacecraft\j's radio signal as it passes very close to Ganymede, but just before it's blocked out by the Jovian moon. These measurements will allow for further study of Ganymede's tenuous atmosphere.
In addition to being the largest of the Jovian satellites, Ganymede is the largest moon in the solar system. Although this marks \JGalileo\j's last encounter with Ganymede, the craft will fly by two other Jovian moons, Callisto and Europa, before its primary mission ends in December. A two-year extension of the \JGalileo\j mission will enable further studies of Europa and Io, depending on the \Jspacecraft\j's health.
Galileo was launched in 1989 and entered orbit around Jupiter on Dec. 7, 1995.
#
"Galileo Returns New Insights Into Callisto and Europa",129,0,0,0
May 23, 1997
Jupiter's icy moon Europa has a metallic core and layered internal structure similar to the Earth's, while the heavily cratered moon Callisto is a mixture of metallic rock and ice with no identifiable central core, according to new results from NASA's \JGalileo\j mission.
In addition, recent plasma wave observations from \JGalileo\j show no evidence of a magnetic field or magnetosphere around Callisto, but do hint at the prospect of a tenuous atmosphere.
These peer-reviewed findings, reported in today's issue of Science magazine and the May 16 issue of Nature magazine, are based on data gathered during \JGalileo\j's Nov. 4, 1996, flyby of Callisto and its Europa encounters on Dec. 19, 1996, and Feb. 20, 1997.
"Before \JGalileo\j, we could only make educated guesses about the structure of the Jovian moons," said Dr. John Anderson, a planetary scientist at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "Now, with the help of the \Jspacecraft\j, we can measure the gravitational fields of the satellites and determine their interior structure and density. We can determine how the matter is distributed inside."
While scientists use seismic waves to study Earth's interior, \JGalileo\j performs remote studies of Jupiter's moons by measuring small changes in the \Jspacecraft\j's trajectory as it passes each body.
"These new results from the gravity data are very consistent with the idea of subsurface oceans on Europa," Anderson said. "We know that Europa has a very deep layer of water in some form, but we don't yet know whether that water is liquid or frozen."
In an article appearing in the May 23 edition of Science, Dr. Margaret Kivelson, principal investigator for \JGalileo\j's magnetometer, reports that during its December 1996 pass by Europa, the magnetometer detected what she described as "a substantial magnetic signature," and also found that Europa's north magnetic pole is pointed in an odd direction.
Based on these observations, Kivelson, a professor at the University of \JCalifornia\j at Los Angeles, said Europa may have a magnetic field about one-quarter the strength of Ganymede's magnetic field.
Although the magnetometer was malfunctioning during \JGalileo\j's Europa flyby in February 1997, Kivelson said the problem is corrected and the device is expected to return valuable data during its upcoming Europa flybys. The next Europa encounter is scheduled for November, with a series of flybys planned during a two-year \JGalileo\j extended mission.
Galileo's findings on the Jovian moon Callisto revealed a much different structure than Europa. Scientists believe that because Callisto is the Galilean moon located farthest from Jupiter, it was never subjected to the same gravitational pull as the inner moons and, therefore, never experienced enough heating to form different layers.
"Callisto had a much more sedate, predictable and peaceful history than the other Galilean moons," Anderson explained, "and, therefore, it is a more typical solar system object." The findings indicate Callisto has no core, but instead has a homogeneous structure, with 60 percent of its ingredients being rock, including iron and iron sulfide, and 40 percent made of compressed ice.
Dr. Donald Gurnett, principal investigator for the \JGalileo\j \Jspacecraft\j's plasma wave instrument, said the instrument displayed a very minor response from Callisto and, consequently, showed no evidence of a magnetic field or magnetosphere. The latest issue of Nature magazine contains these findings, as well as supportive data from magnetometer studies of Callisto, as reported by Dr. Krishan Khurana of UCLA.
However, Gurnett added, "There is some evidence of a plasma source on Callisto, which might indicate a very tenuous atmosphere." Gurnett is a professor at the University of \JIowa\j at \JIowa\j City.
The \JGalileo\j \Jspacecraft\j was launched in October 1989 and entered orbit around Jupiter on Dec. 7, 1995.
#
"Jupiter's Dry Spots and Glowing Auroras To Be Unveiled",130,0,0,0
June 2, 1997
New images from NASA's \JGalileo\j mission revealing dry spots and auroral light patterns on Jupiter will be presented at a press briefing on Thursday, June 5, at 11 a.m. Pacific Standard Time. The briefing will originate from NASA's Jet Propulsion Laboratory, Pasadena, CA, and will be carried live on NASA \JTelevision\j.
The latest images and data reveal the existence of areas where winds converge and cause clouds and moisture to evaporate, but also indicate that the giant, gaseous planet is not as dry as scientists had believed. This may clear up the controversy which arose after \JGalileo\j's probe entered the Jovian atmosphere on Dec. 7, 1995, and found no moisture.
Scientists will also discuss new images and data which show that Jupiter's glowing auroras stretch in a thin, patchy ribbon- like strand near the poles. Scientists believe that despite some similarities, auroras on Jupiter and on Earth are driven by different forces.
NASA \JTelevision\j is available through GE-2, transponder 9C at 85 degrees west longitude, vertical polarization, with a frequency of 3880 Mhz, and audio at 6.8 Mhz.
#
"Galileo Finds New View of Jupiter's Light Show",131,0,0,0
June 5, 1997
Jupiter has both wet and dry regions, just as Earth has tropics and deserts, according to new images and data from the \JGalileo\j \Jspacecraft\j released today. The data may explain why \JGalileo\j's atmospheric probe found much less water than scientists had anticipated when it dropped into the Jovian atmosphere in December 1995.
"We had suspected that the probe landed in the 'Sahara Desert of Jupiter,'" said Dr. Andrew Ingersoll, a professor at the \JCalifornia\j Institute of Technology, Pasadena, CA, and member of the \JGalileo\j science team. "But the new data show there is moisture in the surrounding areas. Jupiter is not as dry overall as we thought."
The area where the probe entered was a clearing in the clouds -- a dry spot through which deeper, warmer layers can be seen. By studying various areas, including those resembling the probe entry site, the \JGalileo\j orbiter has helped scientists understand the probe results.
In fact, the air around a dry spot has 100 times more water than the dry spot itself, according to Dr. Robert Carlson of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, principal investigator for the imaging spectrometer instrument onboard \JGalileo\j.
Such dry spots cover less than one percent of the Jovian atmosphere, and they appear to be regions where the winds converge and create a giant downdraft, according to \JCal\j Tech graduate student Ashwin Vasavada. In fact, the water content of the giant, gaseous planet varies at least as much as the moisture varies from place to place on Earth.
"Winds rise from the deep atmosphere and lose water and ammonia," explained Dr. Glenn Orton, a \JGalileo\j interdisciplinary scientist at JPL and Photopolarimeter-Radiometer co-investigator. "At the top, when they converge and drop back down, nothing is left to condense back into clouds, and a dry clearing is created. These dry spots may grow and diminish, but they recur in the same places, possibly because of the circulation patterns on Jupiter."
Ingersoll said the dry spots are found in a northern hemisphere band at five to seven degrees latitude. When the \JGalileo\j probe was released near the tops of the clouds, it found dry air underneath. But at other locations, the weather might be rather Earth-like.
In the months since the probe's descent, \JGalileo\j mission scientists have debated whether the dry conditions it encountered were due to the downdraft concept, or whether Jupiter's water had somehow been concentrated deep in the gas planet's interior as it formed and evolved four billion years ago. "There was a cosmo- chemical explanation and a meteorological explanation, and our latest analysis clearly favors the idea that the dry spots are a consequence of weather-related activity," Ingersoll said.
"Fifty miles below the cloud top."
#
"Galileo Finds Europa Has An Atmosphere",132,0,0,0
July 18, 1997
NASA's \JGalileo\j \Jspacecraft\j has found an \Jionosphere\j on Jupiter's moon Europa, an indication that the icy moon also has an atmosphere, \JGalileo\j scientists reported today.
"While this discovery does not relate to the question of possible life on Europa, it does show us there is a surface process occurring there, and Europa is not just some dead hunk of material," said lead investigator Dr. Arvydas Kliore of NASA's Jet Propulsion Laboratory, Pasadena, CA. Kliore reports his findings in the July 18 issue of Science magazine.
The \Jionosphere\j was detected through a series of six \Joccultation\j experiments performed during \JGalileo\j's encounters with Europa in December 1996 and February 1997. During \Joccultation\j, Europa was positioned between the \Jspacecraft\j and Earth, causing interruption in the radio signal.
Measurements of the \JGalileo\j radio signal received at the Deep Space Network stations in Goldstone, CA, and \JCanberra\j, \JAustralia\j, showed that the radio beam was refracted by a layer of electrons, or charged particles, in Europa's \Jionosphere\j.
An \Jionosphere\j is a layer of charged particles (ions and electrons) found in the upper levels of an atmosphere, created when gas molecules in the atmosphere are ionized. On Europa, this ionized layer can be caused either by the Sun's ultraviolet radiation or by energetic particles trapped in Jupiter's magnetic field, known as the magnetosphere.
Europa and the other Jovian satellites are immersed in this magnetosphere. "Most likely the charged particles in Jupiter's magnetosphere are hitting Europa's icy surface with great energy, knocking atoms of water molecules off the moon's surface," Kliore said,
Europa's \Jionosphere\j has a maximum density of 10,000 electrons per cubic centimeter, which is significantly lower than the average density of 20,000 to 250,000 electrons per cubic centimeter found in Jupiter's \Jionosphere\j. This indicates that Europa's \Jionosphere\j is tenuous; nonetheless, it is strong enough for scientists to infer the presence of an atmosphere.
The latest \JGalileo\j findings follow last year's observations by NASA's Hubble Space \JTelescope\j of oxygen emissions on Europa, a strong hint that an atmosphere might exist on that moon.
The existence of an \Jionosphere\j and, by inference, an atmosphere, on another Jovian moon, Io, was observed in 1973 during a radio \Joccultation\j conducted by NASA's Pioneer 10 \Jspacecraft\j and confirmed by recent \JGalileo\j occultations.
Io is believed to have an unusual atmosphere affected by sulfur dioxide spewing from the moon's volcanic vents. Kliore and his colleagues are currently studying two of Jupiter's other largest moons, Ganymede and Callisto, to determine whether they also have ionospheres and atmospheres.
"You could say an \Jionosphere\j or some kind of atmosphere has been found on most solar system bodies studied so far," said Kliore.
Participating with Kliore in the Europa radio \Joccultation\j experiments were Dr. David Hinson, professor at Stanford University, Stanford, CA; Dr. Michael Flasar of NASA's Goddard Space Flight Center, Greenbelt, MD; Dr. Andrew Nagy, professor at the University of \JMichigan\j, Ann Arbor, MI, and Dr. Thomas Cravens, professor at the University of Kansas, Lawrence, KS.
The \JGalileo\j mission is managed by JPL for NASA's Office of Space Science, Washington, DC. The \Jspacecraft\j entered the Jovian system on December 7, 1995, and its primary mission will end in November of this year. However, the mission has been extended for two more years so the craft can conduct an intensive study of Europa, with additional flybys of Callisto and Io, depending on \Jspacecraft\j health.
#
"Ganymede G1 & G2 Encounters",133,0,0,0
Voyager images are used to create a global view of Ganymede. The cut-out reveals the interior structure of this icy moon. This structure consists of four layers based on measurements of Ganymede's gravity field and theoretical analyses using Ganymede's known mass, size and density.
Ganymede's surface is rich in water ice and Voyager and \JGalileo\j images show features which are evidence of geological and tectonic disruption of the surface in the past. As with the Earth, these geological features reflect forces and processes deep within Ganymede's interior.
Based on geochemical and geophysical models, scientists expected Ganymede's interior to either consist of: a) an undifferentiated mixture of rock and ice or b) a differentiated structure with a large lunar sized 'core' of rock and possibly iron overlain by a deep layer of warm soft ice capped by a thin cold rigid ice crust.
Galileo's measurement of Ganymede's gravity field during its first and second encounters with the huge moon have basically confirmed the differentiated model and allowed scientists to estimate the size of these layers more accurately. In addition the data strongly suggest that a dense metallic core exists at the center of the rock core.
This metallic core suggests a greater degree of heating at sometime in Ganymede's past than had been proposed before and may be the source of Ganymede's magnetic field discovered by \JGalileo\j's space physics experiments.
Galileo's primary 24 month mission includes eleven orbits around Jupiter and will provide observations of Jupiter, its moons and its magnetosphere.
#
"Ganymede, Fractures in Transitional Terrain",134,0,0,0
This area of dark terrain on Jupiter's moon Ganymede lies near a transitional area between dark and bright terrain. The dark surface is cut by a pervasive network of fractures, which range in width from the limit of resolution up to 2.2 kilometers (1.4 miles).
Bright material is exposed in the walls of the chasms, and dark material fills the troughs. The impurities which darken the ice on the surface of dark terrain may be only a thin veneer over a brighter ice crust. Over time, these materials may be shed down steep slopes, where they collect in low areas.
The image is 68 by 54 kilometers (42 by 33 miles), and has a resolution of 190 meters (623 feet) per picture element (pixel). North is to the top. This image was obtained on September 6, 1996 by the Solid State Imaging (CCD) system aboard NASA's \JGalileo\j \Jspacecraft\j
Launched in October 1989, \JGalileo\j entered orbit around Jupiter on December 7, 1995. The \Jspacecraft\j's mission is to conduct detailed studies of the giant planet, its largest moons and the Jovian magnetic environment.
#
"Galileo Calendar of Events",135,0,0,0
Public events, conferences, symposiums and workshops that members of the \JGalileo\j team will be participating in. For \JGalileo\j's upcoming encounters with Jupiter's satellites, see the Orbital Tour Highlights.
1997
òJan 07-10 - Conference on the Three Galileos: the Man, the \JSpacecraft\j, the \JTelescope\j, Padova, \JItaly\j
òJan 19-24 - Remote Sensing of Volcanoes on Earth and the Planets, Puerto Vallarta, Mexico
òJan 22 - \JGalileo\j Encounter with Europa Event, \JHonolulu\j, Hawaii
òFeb 17 - Carl Sagan Public Memorial, Pasadena, \JCalifornia\j
òSep 22-24 - Io During The \JGalileo\j Era Conference, Flagstaff, \JArizona\j
#
"Clementine Project",136,0,0,0
\BLaunch Date:\b 25 January 1994
\BDescription\b
The objective of the Clementine Mission was to test sensors and \Jspacecraft\j components under extended exposure to the space environment and to make scientific observations of the Moon and the near-Earth asteroid 1620 Geographos. The Clementine mission mapped most of the lunar surface at a number of resolutions and wavelengths from UV to IR. The \Jspacecraft\j was launched on January 25, 1994 at 16:34 and the nominal lunar mission lasted until the \Jspacecraft\j left lunar orbit on May 3. Clementine had five different imaging systems on-board. The UV/Visible camera had a filter wheel with six different filters, ranging from 415 nm to 1000 nm, and including a broad-band filter covering 400 to 950 nm.
The Near Infrared camera also had a six-filter wheel, ranging from 1100 nm to 2690 nm. The Longwave Infrared camera had a wavelength range of 8000 to 9500 nm. The Hi-Res imager had a broad-band filter from 400 to 800 nm and four other filters ranging from 415 to 750 nm. The Star Tracker camera was also used for imaging.
#
"Voyager Project Information",137,0,0,0
A multiple outer-planet flyby mission undertaken by NASA to make the first detailed exploration beyond Mars, and designed to take advantage of a rare (every 175 years) celestial alignment of Jupiter, Saturn, Uranus, and Neptune. Twin \Jspacecraft\j in the Mariner series were launched on Titan-Centaur vehicles in 1977. They flew through the Jovian system (March, July 1979) and, with a boost from Jupiter's gravity, flew on to Saturn (encounters November 1980, August 1981).
Following the Saturn flyby, Voyager 1's trajectory is taking it upward out of the \Jecliptic\j; Voyager 2 used Saturn's gravity to fly on to the historic encounters with Uranus (Jan 1986) and Neptune (August 1989). The \Jspacecraft\j, powered by \Jradioisotope\j thermo-electric generators, were built and are operated by NASA's Jet Propulsion Laboratory. They may send back data about the outermost reaches of the Solar System until well into the 21st-century.
\BMission Details
Launch Date:\b September 5, 1977 (Voyager 1); August 20, 1977 (Voyager 2)
\BDescription\b
The last two \Jspacecraft\j of NASA's Mariner series, Voyager 1 and 2 were the first in that series to be sent to explore the outer solar system. Preceded by the Pioneer 10 (1972 launch) and Pioneer 11 (1973 launch) missions, Voyager 1 and 2 were to make studies of Jupiter and Saturn, their satellites, and their magnetospheres as well as studies of the interplanetary medium. An option designed into the Voyager 2 trajectory, and ultimately exercised, would direct it toward Uranus and Neptune to perform similar studies.
Although launched sixteen days after Voyager 2, Voyager 1's trajectory was a faster path, arriving at Jupiter in March of 1979. Voyager 2 arrived about four months later in July 1979. Both \Jspacecraft\j were then directed on to Saturn with arrival times in November 1980 (Voyager 1) and August 1981 (Voyager 2). Voyager 2 was then diverted to the remaining gas giant, Uranus (January 1986) and Neptune (August 1989).
Data collected by Voyager 1 and 2 were not confined to the periods surrounding encounters with the outer gas giants, with the various fields and particles experiments and the ultraviolet spectrometer collecting data nearly continuously during the interplanetary cruise phases of the mission. Data collection continues as the recently renamed Voyager Interstellar Mission searches for the edge of the solar wind's influence (the heliopause) and exits the solar system.
\BSome Scientific Results of the Voyager Mission\b
A comprehensive list of the achievements of Voyager 1 and 2 would be so extensive that space doesn't permit. Here, then, are a (very) few results that would rank near the top of many such lists:
Discovery of the Uranian and Neptunian magnetospheres, both of them highly inclined and offset from the planets' rotational axes, suggesting their sources are significantly different from other magnetospheres.
The Voyagers found 22 new satellites: 3 at Jupiter, 3 at Saturn, 10 at Uranus, and 6 at Neptune.
Io was found to have active volcanism, the only solar system body other than the Earth to be so confirmed. Triton was found to have active geyser-like structures and an atmosphere.
Auroral zones were discovered at Jupiter, Saturn, and Neptune.
Jupiter was found to have rings. Saturn's rings were found to contain spokes in the B-ring and a braided structure in the F-ring. Two new rings were discovered at Uranus, and Neptune's rings, originally thought to be only ring arcs, were found to be complete, albeit composed of fine material.
At Neptune, originally thought to be too cold to support such atmospheric disturbances, large-scale storms (notably the Great Dark Spot) were discovered.
#
"Giotto Mission",138,0,0,0
\BLaunch Date/Time:\b 1985-07-02 at 11:23:13
\BDescription\b
The Giotto mission was designed to study \JComet\j P/Halley, and also to study \JComet\j P/Grigg-Skjellerup during its extended mission. The major objectives of the mission were to: (1) obtain color photographs of the nucleus; (2) determine the elemental and isotopic composition of volatile components in the cometary coma, particularly parent molecules; (3) characterize the physical and chemical processes that occur in the cometary atmosphere and \Jionosphere\j; (4) determine the elemental and isotopic composition of dust particles; (5) measure the total gas-production rate and dust flux and size/mass distribution and derive the dust-to-gas ratio; and (6) investigate the macroscopic systems of plasma flows resulting from the cometary-solar wind interaction.
The \Jspacecraft\j encountered the \Jcomet\j on March 13, 1986, at a distance of 0.89 AU from the sun and 0.98 AU from the Earth and an angle of 107 degrees from the comet-sun line. The \Jspacecraft\j was based as much as possible on the ESA-GEOS \Jspacecraft\j and was spin stabilized with a rate of 15 rpm. During the encounter with Halley's \Jcomet\j, the spin axis was aligned with the relative velocity vector. The 1.5m dish antenna, operating in the X-band, was inclined and despun in order to point at the Earth (44 degrees with respect to the velocity vector).
The goal was to come within 500 km of Halley's \Jcomet\j at closest encounter. The actual closest approach was measured at 596 km. The \Jspacecraft\j had a dust shield consisting of a front sheet of Aluminum (1 mm thick) and a 12 mm Kelvar near sheet separated by 25 cm, which could withstand impacts of particles up to 0.1 g. The scientific payload was comprised of ten hardware experiments: a narrow-angle camera, three mass spectrometers for neutrals, ions and dust, various dust detectors, a photo polarimeter and a set of plasma experiments. All experiments performed well and returned a wealth of new scientific results, of which perhaps the most important was the clear identification of the cometary nucleus.
Fourteen seconds before closest approach, Giotto was hit by a 'large' dust particle. The impact caused the \Jspacecraft\j angular momentum vector to shift by 0.9 degrees. Scientific data were received intermittently for the next 32 minutes. Some experiment sensors suffered damage during this 32-minute interval.
Other experiments (the camera baffle and deflecting mirror, the dust detector sensors on the front sheet of the bumper shield, and most experiment apertures) were exposed to dust particles regardless of the accident and also suffered damage. Many of the sensors survived the encounter with little or no damage. Questionable or partially damaged sensors included the camera (later proved to not be functional) and one of the plasma analyzers (RPA). Inoperable experiments included the neutral and ion mass spectrometers and one sensor each on the dust detector and the other plasma analyzer (JPA).
During the Giotto extended mission, the \Jspacecraft\j successfully encountered \JComet\j P/Grigg-Skjellerup on July 10, 1992. The closest approach was approximately 200 km. The heliocentric distance of the \Jspacecraft\j was 1.01 AU, and the geocentric distance, 1.43 AU at the time of the encounter. The payload was switched-on in the evening of July 9. Eight experiments were operated and provided a surprising wealth of exciting data.
The Johnstone Plasma Analyser detected the first presence of cometary ions 600,000 km from the nucleus at 12 hours before the closest approach. The Dust Impact Detectors reported the first impact of a fairly large particle at 15:30:56. Bow shocks/waves and acceleration regions were also detected. After the P/Grigg-Skjellerup encounter, the \Jspacecraft\j was retargeted for a possible 1999 encounter pending the existence of sufficient fuel and funding for ground operations support
#
"Hubble Space Telescope (HST)",139,0,0,0
\JHubble Space Telescope (HST) Summary\j
\JHubble's Look At Mars, Conditions For Pathfinder Landing\j
\JHubble's Sharpest View Of Mars\j
\JHubble Captures A Full Rotation Of Mars\j
\JHubble Monitors Weather On Neighboring Planets\j
\JHubble Finds Intergalactic Stars\j
\JBlack Holes Dwell In Most Galaxies, Hubble Census\j
\JSupernova Blast Begins Taking Shape\j
\JProto-Planetary Disks Destruction Explained\j
\JLagoon Nebula, Giant 'Twisters' and Star Wisps \j
\JHubble Space Telescope Check-Out Finds Successes, Concerns\j
\JHubble Images Of Comet Hale-Bopp\j
\JHubble Tracks Fading Optical Counterpart Of Gamma-Ray Burst\j
\JHubble's Upgrades Show Birth and Death Of Stars\j
\JHubble Captures Volcanic Eruption Plume From Io\j
#
"Hubble Space Telescope (HST) Summary",140,0,0,0
\BLaunch Date:\b 1990-04-25
\BDescription\b
The Hubble Space \JTelescope\j (HST) was the first and flagship mission of NASA's Great Observatories program. Designed to complement the wave length capabilities of the other \Jspacecraft\j in the program (CGRO, AXAF, and SIRTF), HST was a 2.4 m, f/24 Ritchey-Chretien \Jtelescope\j capable of performing observations in the visible, near-ultraviolet, and near-infrared(1150 A to 1 mm).
Placed into a low-earth orbit by the space shuttle, HST was designed to be modular so that on subsequent shuttle missions it could be recovered, have faulty or obsolete parts replaced with new and/or improved instruments, and be re-released. HST was roughly cylindrical in shape, 13.1 m end-to-end and 4.3 m in diameter at its widest point.
HST used an elaborate scheme for attitude control to improve the stability of the \Jspacecraft\j during observations. Maneuvering was performed by four of six gyros, or reaction wheels. Pointing could be maintained in this mode (coarse track) or the Fine Guidance Sensors (FGSs) could be used to lock onto guide stars (fine lock) to reduce the \Jspacecraft\j drift and increase the pointing accuracy.
Power to the two on-board computers and the scientific instruments was provided by two 2.4 x 12.1 m solar panels. The power generated by the arrays was also used to charge six nickel-hydrogen batteries which provided power to the \Jspacecraft\j during the roughly 25 minutes per orbit in which HST was within the Earth's shadow.
Communications with the satellite were maintained with the TDRS satellites. Observations taken during the time when neither TDRS was visible from the \Jspacecraft\j were recorded on tape recorder and dumped during periods of visibility. The \Jspacecraft\j also supported real-time interactions with the ground system during times of TDRS visibility, enabling observers to make small offsets in the \Jspacecraft\j pointing to perform their observations. HST was the first scientific \Jspacecraft\j designed to utilize the full capabilities of TDRSS, communicating over either multiple-access or single-access channels at any of the supported transmission rates.
HST was operated in three distinct phases. During the first phase of the mission (Orbital Verification or OV), responsibility for the \Jspacecraft\j was given to Marshall Space Center. OV consisted of an extended, eight-month checkout of the \Jspacecraft\j, including test of the on-board computers, pointing control system, solar arrays, etc. This phase was followed by the Science Verification (SV) phase, lasting nearly another year, during which each of the six science instruments was tested to verify their capabilities and set limits on their safe operations during the remainder of the mission.
Responsibility for the \Jspacecraft\j during SV was given to Goddard Space Flight Center. The last phase of the mission, known as the General Observer (GO) phase, was planned to last from the end of SV through the end of the mission and was the responsibility of the Space \JTelescope\j Science Institute. General observations were phased in gradually, however, during the SV phase because the OV and SV portions of the mission were considerably longer than expected prior to deployment.
The mission was troubled soon after launch by the discovery that the primary mirror was spherically aberrated. In addition, problems with the solar panels flexing as the \Jspacecraft\j passed from the Earth's shadow into sunlight caused problems with the pointing stability. Steps were taken to correct these problems, including replacement of the solar panels, replacement of the Wide Field and Planetary Camera with a second-generation version with built-in corrective \Joptics\j, and replacement of the High-Speed Photometer with COSTAR (Corrective \JOptics\j Space \JTelescope\j Axial Replacement) to correct the aberration for the remaining instruments.
#
"Hubble's Look At Mars, Conditions For Pathfinder Landing",141,0,0,0
July 1, 1997
Hubble Space \JTelescope\j pictures of Mars, taken on June 27 in preparation for the July 4 landing of the Pathfinder \Jspacecraft\j, show a dust storm churning through the deep canyons of Valles Marineris, just 600 miles (1000 km) south of the Pathfinder \Jspacecraft\j landing site.
"Unless the dust storm were to evolve into a massive, global event, its effects on the Pathfinder mission should be minimal, says Steve Lee of the University of \JColorado\j in Boulder, \JColorado\j. "This is something we did not expect to see".
The Hubble astronomers also report the presence of patchy cirrus clouds over the landing site and very thick clouds to the north. Because there are so many clouds (related to low temperatures in the atmosphere causing water vapor to freeze), the dust will probably stay confined to the canyons, they conclude.
If dust rises to the elevations where the water-ice clouds form, ice condenses on dust grains and the heavier ice/dust particles quickly fall back out of the atmosphere. Though the dust could extend at low altitudes over the landing site, researchers say current prevailing winds should not take the dust northward.
"If dust diffuses to the landing site, the sky could turn out to be pink like that seen by Viking," says Philip James of the University of Toledo. Otherwise, Pathfinder will likely show blue sky with bright clouds."
#
"Hubble's Sharpest View Of Mars",142,0,0,0
March 20, 1997
The sharpest view of Mars ever taken from Earth was obtained by the recently refurbished NASA Hubble Space \JTelescope\j (HST). This stunning portrait was taken with the HST Wide Field Planetary Camera-2 (WFPC2) on March 10, 1997, just before Mars opposition, when the red planet made one of its closest passes to the Earth (about 60 million miles or 100 million km).
At this distance, a single picture element (pixel) in WFPC2's Planetary Camera spans 13 miles (22 km) on the Martian surface.
The Martian north pole is at the top (near the center of the bright polar cap) and East is to the right. The center of the disk is at about 23 degrees north latitude, and the central longitude is near 305 degrees.
This view of Mars was taken on the last day of Martian spring in the northern hemisphere (just before summer solstice). It clearly shows familiar bright and dark markings known to astronomers for more than a century.
The annual north polar carbon dioxide frost (dry ice) cap is rapidly sublimating (evaporating from solid to gas), revealing the much smaller permanent water ice cap, along with a few nearby detached regions of surface frost. The receding polar cap also reveals the dark, circular sea' of sand dunes that surrounds the north pole (Olympia Planitia).
Other prominent features in this hemisphere include Syrtis Major Planitia, the large dark feature seen just below the center of the disk. The giant impact basin Hellas (near the bottom of the disk) is shrouded in bright water ice clouds.
Water ice clouds also cover several great volcanos in the Elysium region near the eastern edge of the planet (right). A diffuse water ice haze covers much of the Martian equatorial region as well.
The WFPC2 was used to monitor dust storm activity to support the Mars Pathfinder and Mars Global Surveyor Orbiter Missions, which are currently en route to Mars. Airborne dust is most easily seen in WFPC2's red and near-infrared images.
Hubble's "weather report" from these images in invaluable for Mars Pathfinder, which is scheduled for a July 4 landing. Fortunately, these images show no evidence for large-scale dust storm activity, which plagued a previous Mars mission in the early 1970s.
The WFPC2 was used to observe Mars in nine different colors spanning the ultraviolet to the near infrared. The specific colors were chosen to clearly discriminate between airborne dust, ice clouds, and prominent Martian surface features. This picture was created by combining images taken in blue (433 nm), green (554 nm), and red (763 nm) colored filters.
#
"Hubble Captures A Full Rotation Of Mars",143,0,0,0
Pictures of the planet Mars taken with the recently refurbished NASA Hubble Space \JTelescope\j (HST) will provide the most detailed global view of the red planet ever obtained from Earth.
The images were taken by HST's Wide Field Planetary Camera-2 on March 10, 1997, just before Mars opposition, when the red planet made one of its closest to the Earth (about 60 million miles or 100 million km).
These pictures were taken during three HST orbits that were separated by about six hours. This timing was chosen so that Mars, with its 24-hour 39-minute day, would rotate about 90 degrees between orbits. This imaging sequence therefore covers most of the Martian surface. These observations will be combined with others planned for March 30 to provide complete coverage.
During each orbit, Mars was observed in nine different colors spanning the ultraviolet to the near infrared. The specific colors were chosen to clearly discriminate between airborne dust, ice clouds, and prominent Martian surface features.
The color picture shown here was created by combining images taken in blue (433 nm), green (554 nm), and red (763 nm) colored filters. The Martian north pole is at the top (near the center of the bright polar cap) and East is to the right. The center of the disk is at about 23 degrees north latitude, and the central longitudes are near 160, 210, and 305 degrees.
These images show the planet on the last day of Martian spring in the northern hemisphere (just before summer solstice). The annual north polar carbon dioxide frost (dry ice) cap is rapidly sublimating, revealing the much smaller permanent water ice cap.
This polar cap remnant, along with a few nearby detached regions of surface frost are most obvious in pictures taken through ultraviolet, blue, and green filters. These filters also show numerous bright water ice clouds.
The brightest clouds are in the vicinity of the giant volcanos on the Tharsis Plateau (to right of center on left image), and in the giant impact basin, Hellas (near bottom of right-hand image), but a diffuse haze covers much of the Martian tropics as well.
The familiar bright and dark markings on the Martian surface are most obvious in images taken through red and near-infrared filters. These images clearly reveal the large, dark, circular "sea" of sand dunes (Olympia Planitia) that surrounds the north pole, as well a number of other familiar features, including the giant Tharsis volcanos.
The 16-mile (27 km) high Olympus Mons is near the center of the left-hand image, with Arsia, Povonis, and Ascraeus Mons forming a south-west to north-east line just to its right. The \Jvolcano\j, Elysium Mons is near the center of the middle image. The prominent dark feature just below the center on the disk on the rightmost image is Syrtis Major Planitia.
Hubble is being used to monitor dust storm activity to support the Mars Pathfinder and Mars Global Surveyor Orbiter Missions, which are currently en route to Mars. Airborne dust is most easily seen in WFPC2's red and near-infrared images.
Weather reports derived from these observations are particularly valuable for Mars Pathfinder, which is scheduled for a July 4, 1997 landing on the red planet. A preliminary analysis of these HST data reveals enhanced dust activity over the dark Vastitas Borealis region in the northern hemisphere, and over the Noachis Terra and Terra Tyrrhena regions just south of the Martian equator.
There is also evidence for airborne dust and ice clouds in the Hellas basin. However, these images show no evidence for large-scale dust storm activity.
#
"Hubble Monitors Weather On Neighboring Planets",144,0,0,0
MARS: A COOLER, CLEARER WORLD
Four years, (or two Mars years') worth of Hubble observations show that the Red \JPlanet\j's climate has changed since the mid-1970's. "The Hubble results show us that the Viking years are not the rule, and perhaps not typical. Our early assumptions about the Martian climate were wrong," said Philip James of the University of Toledo.
"There has been a global drop in temperature. The planet is cooler and the atmosphere clearer than seen before," said Steven Lee of the University of \JColorado\j in Boulder. "This shows the need for continuous monitoring of Mars. Space probes provided a close-up look, but it's difficult to extrapolate to long-term conditions based upon these brief encounters."
The researchers attribute the cooling of the Martian atmosphere to diminished dust storm activity, which was rampant when a pair of NASA Viking orbiter and lander \Jspacecraft\j arrived at Mars in 1976. Two major dust storms occurred during the first year of the Viking visits, which left fine dust particles suspended in the Martian atmosphere for longer than normal.
Warmed by the Sun, these dust particles (some only a \Jmicron\j in diameter, about the size of smoke particles) are the primary source of heat in the Martian atmosphere.
"Hubble is showing that our early understanding based on these visits is wrong. We just happened to visit Mars when it was dusty, and now the dust has settled out," Lee said. "We are going to have to look at Mars for many years to truly understand the workings of the climate," said Todd Clancy, of the Space Science Institute, Boulder, \JColorado\j.
Knowledge about the Martian climate has been limited by the fact that ground-based telescopes can only see weather details when Earth and Mars are closest -- an event called opposition -- that happens only once every two years. Though Hubble has observed Mars only for four years, the observations are equivalent to 15 years of ground- based observing because Hubble can follow seasonal changes through most of Mars' orbit.
Though the Mariner and Viking series of flyby, orbiter and lander \Jspacecraft\j that visited Mars in the late 60's and 70's provided a close-up look at Martian weather, these were snapshots of the planet's complex climate.
Hubble provides the advantage of a global view - much like the satellites that monitor Earth's weather, and can follow martian seasonal changes over many years. When Mars is closest to Earth, Hubble returns near-weather satellite resolution.
MARS -- NO LACK OF OZONE
Although there has been concern about a lack of ozone (a form of molecular oxygen created by the effects of sunlight on an atmosphere), dubbed the "ozone hole" over Earth's poles, there are no ozone holes on Mars.
By contrast, the planet has a surplus of ozone over its northern polar cap, as first identified by the Mariner 9 \Jspacecraft\j in 1971. (However the Martian atmosphere is different enough from Earth's that few parallels can be drawn about processes controlling the production and destruction of ozone.)
Hubble's ultraviolet sensitivity is ideal for monitoring ozone levels on a global scale. The Martian ozone is yet another indication the planet has grown drier, because the water in the atmosphere that normally destroys ozone has frozen-out to become ice-crystal clouds.
Spectroscopic observations made with the Faint Object Spectrograph (FOS) show that ozone now extends down from Mars' north pole to mid and lower latitudes. However, the Martian atmosphere is so thin, even this added ozone would offer future human explorers little protection from the Sun's harmful ultraviolet rays.
SEASONS ON MARS
The fourth planet from the Sun, Mars is one of the most intensely scrutinized worlds because of its Earth-like characteristics. Mars is tilted on its axis by about the same amount Earth is, hence Mars goes through seasonal changes.
However, because Mars' atmosphere is much thinner than Earth's, it is far more sensitive to minor changes in the amount of light and heat received from the Sun. This is intensified by Mars' orbit that is more elliptical than Earth's, so it's range of distance from the Sun is greater during the Martian year. Mars is now so distant, the sun is nearly 25% dimmer than average.
This chills Mars' average temperature by 36 degrees Fahrenheit (20 degrees Kelvin). At these cold temperatures, water vapor at low altitudes freezes out to form ice-crystal clouds now seen in abundance by Hubble.
"Clouds weren't considered to be very important to the Martian climate during the Viking visits because they were so scarce," says Clancy. "Now we can see where they may play a role in transporting water between the north and south poles during the Martian year."
Seasonal winds also play a major role is transporting dust across Mars' surface, and rapidly changing the appearance of a region. This gave early astronomers the misperception that Mars' shifting surface color was evidence of vegetation following a season cycle.
As clearly seen in the Hubble images, past dust storms in Mars' southern hemisphere have scoured the plains of fine light dust and transported the dust northward. This leaves behind a relatively coarser, less reflective sand in the southern hemisphere.
VENUS: NO EVIDENCE FOR NEW VOLCANIC ERUPTIONS
Hubble spectroscopic observations of Venus taken with the Goddard High Resolution Spectrograph provide a new opportunity to look for evidence of volcanic activity on the planet's surface. Though radar maps of the Venusian surface taken by the Magellan orbiter revealed numerous volcanoes, Magellan did not find clear cut evidence for active volcanoes.
Hubble can trace atmospheric changes that might be driven by volcanism. An abundance of sulfur dioxide in the atmosphere could be a tell-tale sign of an active volcanos. Sulfur dioxide was first detected by the Venus Pioneer probe in the late 1970s and has been declining ever since.
The Hubble observations show that sulfur dioxide levels continue to decline. This means there is no evidence for the recurrence of large scale volcanic eruptions in the last few years.
Ejected high into Venus' murky atmosphere, this sulfur dioxide is broken apart by sunlight to make an acid rain of concentrated sulfuric acid. This is similar to what happens on Earth above coal-burning power plants - but on a much larger and more intense scale.
FUTURE PLANS
More Hubble observations of Mars and Venus are critical to planning visits by future space probes. In particular, both robotic and human missions to Mars will need to be targeted for times during the Martian year when there is a minimal chance of getting caught in a dust storm.
Knowing whether the atmosphere is relatively hot or cold is crucial to planning aerobraking maneuvers, where \Jspacecraft\j use the aerodynamic drag of an atmosphere to slow down and enter an orbit around the planet.
This reduces the amount of propellant needed for the journey. "If the atmosphere is more extended than expected the added friction could burn up an aerobraking \Jspacecraft\j, just as Earth's atmosphere incinerates infalling meteors," says James.
Ultimately, knowing the Martian climate will be an fundamental prerequisite for any future plans to establish a permanent human outpost on the Red \JPlanet\j.
#
"Hubble Finds Intergalactic Stars",145,0,0,0
January 14, 1997
NASA's Hubble Space \JTelescope\j has found a long sought population of "stellar outcasts" -- stars tossed out of their home galaxy into the dark emptiness of intergalactic space. This is the first time stars have been found more than 300,000 light-years (three Milky Way diameters) from the nearest big galaxy.
The isolated stars dwell in the Virgo cluster of galaxies, about 60 million light-years away. The results suggest this population of "lone stars" accounts for 10 percent of the Virgo cluster's mass, or 1 trillion Sun-like stars adrift among the 2,500 galaxies in Virgo. "Our discovery provides a new tool for studying clusters of galaxies," says Harry Ferguson of Space \JTelescope\j Science Institute in Baltimore, Maryland.
The distribution of the stars in Virgo could help astronomers probe the distribution of dark matter in the cluster. (Dark matter is an unknown type of matter that accounts for most of the mass of the universe.) Another possible spinoff is that the stars detected, which are the brightest members of the red giant class, may serve as "standard candles" (stars that can be used for calibrating distances), providing an independent method for measuring cosmological distances to Virgo.
Such measurements are key to estimating the expansion rate and age of the universe. These results are being presented at the 189th Meeting of the American Astronomical Society in Toronto, Canada, by Ferguson and co-investigators Nial Tanvir (University of Cambridge, Cambridge, United Kingdom) and Ted von Hippel (University of Wisconsin).
Intergalactic stars have been predicted to exist as a result of galaxy interactions and mergers early in a galaxy cluster's history. These close encounters should have ripped stars out of their home galaxies and tossed them into intergalactic space, where they drift free of the gravitational influence of any single galaxy.
It was predicted that the stars should appear as a diffuse excess of light in Virgo, and there have previously been observations from ground-based telescopes that report just such an excess. "However, there are large uncertainties in the ground-based measurements, and it is not clear whether the diffuse light originates from galaxies too faint to detect individually or from a more uniform sea of stars," says Ferguson.
The accidental discovery in 1996 of planetary nebulae (stellar remnants) in Virgo, which are far removed from any galaxy, offered additional evidence that such an intergalactic population really exists.
The Hubble astronomers found the background stars by taking an exposure of a "blank" portion of sky in Virgo. The position is in the vicinity of the giant elliptical galaxy M87 in the center of Virgo, but far enough from the galaxy for the stars not to be members of M87's halo. The Virgo field was compared to the Hubble Deep Field (HDF) image which represents a region of sky devoid of any nearby galaxy cluster. With the HDF serving as the control, the astronomers counted approximately 600 sources down to 27.8 magnitude.
The stars are bright red giants -- stars late in their lives. Presumably there are many fainter stars -- perhaps as many as 10 million -- in the same field but are below Hubble's sensitivity.
"These stars are truly intergalactic because they are so isolated their motion is probably governed by the gravitational field of the cluster as a whole, rather than the pull of any one galaxy," says Ferguson. The Space \JTelescope\j Imaging Spectrograph (STIS) and Near Infrared Camera and Multi-Object Spectrometer (NICMOS) planned for installation on Hubble this February will be used to help understand the history of the stellar outcasts.
Comparison of heavy element abundances in the "loner" stars and in the Virgo galaxies should help to uncover whether the stars wandered off from the outskirts of galaxies that still exist, are the remnants of galaxies that were completely disrupted, or were somehow formed in the dark reaches of intergalactic space.
The Space \JTelescope\j Science Institute is operated by the Association of Universities for Research in \JAstronomy\j, Inc. (AURA) for NASA, under contract with the Goddard Space Flight Center, Greenbelt, Maryland. The Hubble Space \JTelescope\j is a project of international cooperation between NASA and the European Space Agency (ESA).
#
"Black Holes Dwell In Most Galaxies, Hubble Census",146,0,0,0
January 13, 1997
Announcing the discovery of three black holes in three normal galaxies, an international team of astronomers suggests nearly all galaxies may harbor supermassive black holes which once powered quasars (extremely luminous nuclei of galaxies), but are now quiescent.
This conclusion is based on a \Jcensus\j of 27 nearby galaxies carried out by NASA's Hubble Space \JTelescope\j and ground-based telescopes in Hawaii, which are being used to conduct a spectroscopic and photometric survey of galaxies to find black holes which have consumed the mass of millions of Sun-like stars.
The findings, being presented today at the 189th Meeting of the American Astronomical Society in Toronto, Canada, should provide insights into the origin and evolution of galaxies, as well as clarify the role of quasars in galaxy evolution.
The key results are:
Supermassive black holes are so common, nearly every large galaxy has one.
A black hole's mass is proportional to the mass of the host galaxy, so that, for example, a galaxy twice as massive as another would have a black hole that is also twice as massive. This discovery suggests that the growth of the black hole is linked to the formation of the galaxy in which it is located.
The number and masses of the black holes found are consistent with what would have been required to power the quasars.
"We believe we are looking at "fossil quasars" and that most galaxies at one time burned brightly as a quasar," says team leader Doug Richstone of the University of \JMichigan\j, Ann Arbor, \JMichigan\j. These conclusions are consistent with previous Hubble Space \JTelescope\j observations showing quasars dwelling in a variety of galaxies, from isolated normal-looking galaxies to colliding pairs.
Two of the black holes "weigh in" at 50 million and 100 million solar masses in the cores of galaxies NGC 3379 (also known as M105) and NGC 3377 respectively. These galaxies are in the "Leo Spur", a nearby group of galaxies about 32 million light-years away and roughly in the direction of the Virgo cluster.
Located 50 million light-years away in the Virgo cluster, NGC 4486B possesses a 500-million solar mass black hole. It is a small satellite of the galaxy M87, a very bright galaxy in the Virgo cluster. M87 has an active nucleus and is known to have a black hole of about 2 billion solar masses.
Though several groups have previously found massive black holes dwelling in galaxies the size of our Milky Way or larger, these new results suggest smaller galaxies have lower-mass black holes, below Hubble's detection limit. The survey shows the black hole's mass is proportional to the host galaxy's mass. Like shoe sizes on adults, the bigger the galaxy, the larger the black hole.
It remains a challenging puzzle as to why black holes are so abundant, or why they should be proportional to a galaxy's mass. One idea, supported by previous Hubble observations, is that galaxies formed out of smaller "building blocks" consisting of star clusters. A massive "seed" black hole may have been present in each of these protogalaxies. The larger number of building blocks needed to merge and form very luminous galaxies would naturally have provided more seed black holes to coalesce into a single, massive black hole residing in a galaxy's nucleus.
An alternative model is that galaxies start at some early epoch with a modest black hole (not necessarily approaching the masses discussed here), but that the black hole consumes some fixed fraction of the total gas shed by the stars in the galaxy during their normal evolution. If that fraction is around 1 percent, the black holes could easily weigh as much as they do now, and would naturally track the current luminosity of the galaxy.
Critical ground-based observations to identify candidates were obtained for all three of these objects by John Kormendy with the Canada-France-Hawaii \JTelescope\j (CFHT) on Mauna Kea, Hawaii. The NGC 4486b black hole detection was also based on CFHT spectra.
Hubble's high resolution then allowed the team to peer deep into the cores of the galaxies with extraordinary resolution unavailable from ground-based telescopes, and measure velocities of stars orbiting the black hole.
A sharp rise in velocity means that a great deal of matter is locked away in the galaxy's core, creating a powerful gravitational field that accelerates nearby stars. The team is confident their statistical search technique has allowed them to pinpoint all the black holes they expect to see, above a certain mass limit. "However, our result is complicated by the fact that the observational data for the galaxies are not of equal quality, and that the galaxies are at different distances," says Richstone.
One of the features of the February 1997 servicing mission to the Hubble will be the installation of the Space \JTelescope\j Imaging Spectrograph (STIS). This spectrograph will greatly increase the efficiency of projects, such as this black hole \Jcensus\j, that require spectra of several nearby positions in a single object.
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"Supernova Blast Begins Taking Shape",147,0,0,0
January 14, 1997
Though the brightest \Jsupernova\j in four centuries lit up the southern sky almost exactly 10 years ago on Feb. 23, 1987, astronomers have waited a decade for the ballooning fireball to become large enough -- about one-sixth of a light-year -- to be resolved from Earth's orbit with NASA's Hubble Space \JTelescope\j (HST).
Astronomers announced today that a close monitoring of the \Jsupernova\j (designated SN1987A) with HST's sharp view has resolved a one-tenth light-year long dumbbell-shaped structure consisting of two blobs of debris expanding apart at nearly 6 million miles per hour from each other.
"This structure is a bit of a surprise," says Jason Pun of Goddard Space Flight Center, Greenbelt, Maryland. "This is the first time we can see the \Jgeometry\j of the explosion and relate it to the \Jgeometry\j of the large glowing ring system around the \Jsupernova\j, which has an hourglass shape. The images may yield important clues to the dynamics of the \Jsupernova\j explosion and the structure of the progenitor star."
Pun says the dim area between the blobs may be related to the equatorial belt of material seen around the \Jsupernova\j that existed before the star exploded. The ring was illuminated by the \Jsupernova\j in 1987 during the explosion and has been slowly fading since then.
Ever since the star self-destructed in 1987 astronomers realized that it offered a once-in-a-lifetime possibility, because of its close proximity to Earth, to obtain images of the explosion at various stages and look for any changes in the shape and dynamics.
The latest findings are the result of the \JSupernova\j Intensive Study collaboration, headed by Professor Robert Kirshner of the Harvard-Smithsonian Center for Astrophysics in Cambridge, \JMassachusetts\j. Images of SN1987A were taken in September 1994, March 1995, and February 1996 with the Wide Field and Planetary Camera 2 (WFPC2). These results are being presented today at the 189th Meeting of the American Astronomical Society in Toronto, Canada, by co-investigator Pun.
The explosion of the \Jsupernova\j debris appears to be perpendicular to the plane of the inner ring. This suggests that whatever properties that the pre-supernova star has, such as rotation or the existence of a companion star, that is responsible for the formation of the inner ring, may also have influenced the dynamics of the explosion.
The explosion was triggered 10 years ago when the collapse of the star's core sent a blast wave of neutrinos which heated the star's inner layers to 10 billion degrees Fahrenheit. This triggered a shockwave which then ripped the star apart and sent the debris hurtling into space. The fireball has since cooled down (to a few hundred degrees Fahrenheit) and the debris is now heated by nuclear energy from the decay of radioactive nuclei produced in the explosion.
The Space \JTelescope\j Imaging Spectrograph (STIS) and Near Infrared Camera and Multi-Object Spectrometer (NICMOS), planned for installation on Hubble this February, will be used to obtain a spatially resolved velocity map of the debris, providing information on the physical conditions of the two blobs.
The debris is expected to collide with the inner ring as early as the year 2002. This will light up all of the dark nebulosity surrounding the \Jsupernova\j, providing new clues to the nature and evolution of the stellar explosion.
Theoretical models, coupled with NASA Hubble Space \JTelescope\j observations of the Trapezium cluster in the Orion nebula, suggest that disks around young cluster stars may not survive long enough for planets to form within them. This implies that there are certain hostile environments in star-forming regions that may inhibit planet formation.
The findings, presented by an international collaboration of astronomers today at the meeting of the American Astronomical Society in Toronto, Canada, explain the destruction of circumstellar disks in Orion's Trapezium, a star cluster at the very center of the nebula.
The report is being presented by Doug Johnstone, a Natural Sciences and \JEngineering\j Research Council (NSERC) Post-Doctoral Fellow at the Canadian Institute for Theoretical Astrophysics, University of Toronto. "For the first time we have a complete evolutionary picture for the stunning objects observed in the Trapezium," says Johnstone.
Their work provides an innovative technique for analyzing circumstellar disks, determining disk masses, and constraining the gestation period for planet embryos around stars in dense clusters, say researchers. "This may help produce a consensus within the star and planet formation community on standard disk properties," says Johnstone.
The team's results show the disks of dust and gas, which can be several billion kilometers across, are initially similar to the disk which is believed to have formed the planets in our own Solar system, but quickly evaporate in the glare of bright massive neighboring stars in the Trapezium. Radiation from these stars photoionizes, or heats and disperses, the cold gas. Within 1 million years the disk is eroded, a time scale shorter than the 1 to 10 million years it would take for planets to form according to current models.
"The theory of disk destruction predicts most efficient destruction at large distances from the embedded, central star. Near the center of the disk, perhaps even at the same distance as the Earth is from the Sun, the remnant disk might survive long enough to form planet embryos," says Johnstone. "Without a more detailed understanding of planet formation it is not possible to predict the future of these disks, but standard models based on our own Solar system suggest that giant planets like Jupiter and Saturn, at comparable distances from their central star, would be ruled out."
Using the Planetary Camera on the Hubble Space \JTelescope\j (HST), Johnstone's collaborators John Bally and Dave Devine of the University of \JColorado\j, and Ralph Sutherland of the Australian National University observed the Trapezium, a young, million-year-old star-forming region just below the belt of the \Jconstellation\j Orion.
Located nearby, only fifteen hundred light years away, the Trapezium region is the closest star formation site containing both Sun-like stars and stars much more massive than the Sun. While ground-based observations have hinted at extended structures surrounding the Sun-like stars, HST has produced images with incredible detail revealing that these stars are embedded in circumstellar disks and surrounded by diffuse hot ionized gas.
The idea that these young stars are embedded in evaporating disks was first proposed by Ed Churchwell of the University of \JWisconsin\j, based on ground based radio-wave observations. Recognizing that these disks might be the birth sites for planets, C. Robert O'Dell of Rice University in Houston, \JTexas\j confirmed the disk hypothesis using HST, and named the objects "proplyds" as an \Jacronym\j for proto-planetary disk.
"However, until our work there was no satisfactory model detailing the origin of the diffuse cloud of hot gas observed around each of these Sun-like stars," says Johnstone. Johnstone, along with David Hollenbach and Herbert Stoerzer of NASA Ames Research Center in \JCalifornia\j, have developed theoretical models describing the destructive effect of high energy radiation on disks, combining the results with Hubble observations by John Bally's team to produce a coherent evolutionary picture.
They report that the disk surface is initially heated to temperatures in excess of 1000 \JCelsius\j by the impinging radiation, evaporating the surface layer much like steam evaporates from the surface of boiling water.
As this material flows away from the central star and disk, higher energy photons ionize the gas, heating it to temperatures reaching 10000 degrees \JCelsius\j and in the process producing the nebulous glow seen in the images. "We are witnessing the destructive event through the illumination of the evaporated material" according to Johnstone.
Evaporation of the circumstellar disk erodes approximately three moon masses of material per year according to the theoretical model, a number which is verified by the HST observations. The exact \Jevaporation\j rate from the circumstellar disk is directly related to the size of the disk and thus, as the disk evaporates and shrinks, the erosion rate decreases.
By using this knowledge, and fitting the \Jevaporation\j model to the HST observations, the collaboration shows that the original circumstellar disks surrounding stars in the Trapezium were similar in appearance to disks around young stars in other systems, and more importantly to the hypothesized Solar proto-planetary disk from which our own nine planets formed.
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"Lagoon Nebula, Giant 'Twisters' and Star Wisps",149,0,0,0
January 22, 1997
This NASA Hubble Space \JTelescope\j (HST) image reveals a pair of one-half light-year long interstellar "twisters" -- eerie funnels and twisted-rope structures (upper left) -- in the heart of the Lagoon Nebula (Messier 8) which lies 5,000 light-years away in the direction of the \Jconstellation\j \JSagittarius\j.
The central hot star, O Herschel 36 (upper left), is the primary source of the ionizing radiation for the brightest region in the nebula, called the Hourglass. Other hot stars, also present in the nebula, are ionizing the extended optical nebulosity. The ionizing radiation induces photo-evaporation of the surfaces of the clouds (seen as a blue "mist" at the right of the image), and drives away violent stellar winds tearing into the cool clouds.
Analogous to the spectacular phenomena of Earth tornadoes, the large difference in temperature between the hot surface and cold interior of the clouds, combined with the pressure of starlight, may produce strong horizontal shear to twist the clouds into their tornado-like appearance. Though the spiral shapes suggest the clouds are "twisting", future observations will be needed, perhaps with Hubble's next generation instruments, with the spectroscopic capabilities of the Space \JTelescope\j Imaging Spectrograph (STIS) or the Near Infrared Camera and Multi-Object Spectrometer (NICMOS), to actually measure velocities.
This Hubble picture reveals a variety of small scale structures in the interstellar medium, small dark clouds called Bok globules, bow shocks around stars, ionized wisps, rings, knots and jets.
The Lagoon Nebula and nebulae in other galaxies are sites where new stars are being born from dusty molecular clouds. These regions are the "space laboratories" for the astronomers to study how stars form and the interactions between the winds from stars and the gas nearby. By studying the wealth of data revealed by HST, astronomers will understand better how stars form in the nebulae.
These color-coded images are the combination of individual exposures taken in July and September, 1995 with Hubble's Wide Field and Planetary Camera 2 (WFPC2) through three narrow-band filters (red light -- ionized sulphur atoms, blue light -- double ionized oxygen atoms, green light -- ionized hydrogen).
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"Hubble Space Telescope Check-Out Finds Successes, Concerns",150,0,0,0
March 25, 1997
The Servicing Mission Observatory Verification (SMOV) for NASA's Hubble Space \JTelescope\j (HST), currently about halfway through its detailed check-out prior to returning to scientific operations, has found Hubble in overall excellent health, with seven of the eight components replaced or installed during the servicing mission functioning very well to date. However, a few concerns with one of the science instruments are being evaluated.
"The Hubble Space \JTelescope\j is checking out extremely well overall, and the few anomalies we see give us no reason to believe we will not be able to meet all our scientific goals," said Dr. Ed Weiler, HST Program Scientist, NASA Headquarters, Washington, DC.
"I'm very impressed that in just the few weeks since the servicing mission, we've already seen Hubble take the best images of Mars ever obtained from Earth's distance. Every observatory commissioning encounters some problems, but we're on track to clear up all our remaining concerns. That's good news for the many, many astronomers who are lined up for observing time on Hubble."
Earlier this month science observations with the Wide Field and Planetary Camera-2 resumed, and on March 10 the science team obtained images of Mars. Also, further optimization and alignment of the mirrors in the new Fine Guidance Sensor (FGS), installed during the servicing mission, were completed with excellent results following its first star observation. Project management officials say it's clearly the best FGS aboard HST.
Commissioning of the new Space \JTelescope\j Imaging Spectrograph (STIS) has proceeded very well, according to project officials. In the coming two weeks team members will test the instrument's ability to acquire targets in the narrow slits. Once this is demonstrated, the instrument will be ready to begin science operations.
Checkout of the Near Infrared Camera and Multi- Object Spectrometer (NICMOS), installed during the second servicing mission, has provided both excellent results and some areas of concern.
The NICMOS, designed to observe the universe in near-infrared light, contains three cameras and a set of highly advanced light sensors which must be maintained at a very cold temperature -- nominally 58 degrees Kelvin (-355 degrees Fahrenheit). These sensors, along with filters and other components, are housed in a large cryogenic dewar (a high-technology insulated bottle filled with about 225 lbs of solid \Jnitrogen\j embedded in aluminum foam).
The NICMOS Principal Investigator, Dr. Rodger Thompson, University of \JArizona\j, said NICMOS high resolution cameras 1 and 2 have shown excellent images in preliminary focus tests. However, these tests also show that camera 3 focus is currently beyond the range of the NICMOS internal mechanical adjustment capability. Analysis indicates the situation may be due to unexpected thermal contact in the dewar, which results in a slightly warmer cryogen temperature and a subsequent reduction of dewar lifetime.
The most likely explanation is that as the solid \Jnitrogen\j warms up it expands, and exerts pressure on the internal structure of the dewar. This expansion resulted in an unwanted physical contact between two internal structural components of the dewar, providing a pathway for excess heat to travel from the warmer outer structure of the dewar to its colder internal parts, warming the solid \Jnitrogen\j to a higher than desired operating temperature. This expansion also is affecting the performance of Camera 3.
The analysis team expects that the thermal contact might release in the future, returning NICMOS to its nominal state. Under these conditions, analysts predict that camera 3 should move back into the instrument's range of focus. Rearrangement of the NICMOS observing schedule could allow the full implementation of the NICMOS science program.
It will take several weeks or months for team engineers to be able to determine for certain the amount of reduction in the lifetime of the cryogen; however, the reduction can be compensated for by rearrangement of observing schedules.
Current plans call for SMOV activities to continue for the next few weeks with results of the Early Release Observation program available in early May.
During the STS-82 HST Second Servicing Mission in February, astronauts aboard the Space Shuttle Discovery replaced two older science instruments aboard Hubble with STIS and NICMOS, and also replaced a Fine Guidance Sensor, a Reaction Wheel Assembly, a Data Interface Unit, a Solar Array Drive \JElectronics\j package, an Engineering/Science Tape Recorder, and a Solid State Recorder. In addition, the astronauts performed other maintenance on the observatory, including patching of some \Jinsulation\j and installing covers on the Magnetic Sensing System.
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"Hubble Images Of Comet Hale-Bopp",151,0,0,0
March 27, 1997
This is a series of Hubble Space \JTelescope\j observations of the region around the nucleus of Hale-Bopp, taken on eight different dates since September 1995. They chronicle changes in the evolution of the nucleus as it moves ever closer to, and is warmed by, the sun.
The first picture shows a strong dust outburst on the \Jcomet\j that occurred when it was beyond the orbit of Jupiter. Images in the Fall of 1996 show multiple jets that are presumably connected to the activation of multiple vents on the surface of the nucleus.
In these false color images, taken with the Wide Field and Planetary Camera 2, the faintest regions are black, the brightest regions are white, and intermediate intensities are represented by different levels of red. All images are processed at the same spatial scale of 280 miles per pixel (470 kilometers), so the solid nucleus, no larger than 25 miles across, is far below Hubble's resolution.
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"Hubble Tracks Fading Optical Counterpart Of Gamma-Ray Burst",152,0,0,0
April 1, 1997
NASA's refurbished Hubble Space \JTelescope\j has made an important contribution toward solving one of \Jastronomy\j's greatest enigmas by allowing astronomers to continue watching the fading visible-light counterpart of a gamma-ray burst (GRB), one of the most energetic and mysterious events in the universe.
The so-called optical counterpart is presumably a cooling fireball from the catastrophic event that triggered the massive burst of invisible gamma rays -- the highest-energy radiation in the universe. This event may have unleashed as much energy in a few seconds as the Sun does in ten billion years!
The burst was discovered on February 28 by the Gamma-Ray Burst Monitor aboard the Italian-Dutch BeppoSAX satellite. The burst was also within the field of view of one of the SAX Wide Field Cameras. Followup observations were conducted by several other space-based astronomical observatories. The visible GRB counterpart, the first ever detected, was then discovered in a pair of ground-based telescopic images of the region where the burst occurred. Taken a week apart, the later picture showed that an object that could be seen in the first image had disappeared in the field, suggesting it was the decaying fireball from the event. A week after that discovery, astronomers at the New Technology \JTelescope\j and the Keck \Jtelescope\j identified an extended source at the location of the suspected GRB.
Hubble's high resolution and sensitivity were brought in to hunt down the rapidly dimming fireball -- plunging from 21st to below 23rd magnitude in eight days -- after it had grown so faint that it could not be resolved by ground-based telescopes by March 13. On March 26, Hubble allowed astronomers to reacquire the lost remnant, and continue following the behavior of the fading source. The Hubble observation clearly shows that the visible GRB source has two components: a point-like object and an extended feature.
This observation demonstrates Hubble's unique capability for monitoring the aftermath of gamma-ray bursts, long after they have faded from the view of Earth-based telescopes. And there will be no shortage of targets: once a day, a gamma-ray burst occurs somewhere in the universe.
"Now we know that, at least in some cases, we can follow the aftermath of GRBs for several weeks, using a coordinated effort between ground-based telescopes, Hubble and other spacecraft," said Kailash Sahu, leader of a team of scientists at The Space \JTelescope\j Science Institute, Baltimore
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"Hubble's Upgrades Show Birth and Death Of Stars",153,0,0,0
May 12, 1997
Three months after an orbital house call by astronauts, new instruments aboard NASA's Hubble Space \JTelescope\j are helping astronomers probe the universe in greater detail than ever before.
New data released by NASA today include direct evidence of a supermassive black hole and remarkable new details on the explosive life cycle of stars. NASA also reported that all new Hubble instruments and upgrades are generally performing well.
"We're extremely excited about the quality and precision of the images from Hubble," said Wes Huntress, NASA Associate Administrator for Space Science. "Following check-out of the instruments, Hubble will return to full science operations, and we can expect a continuing flow of new and exciting discoveries."
These initial results clearly demonstrate the ability of the new instruments to fulfill their science goals with the Hubble \JTelescope\j, say project astronomers. Project officials are pleased to report that other instruments and \Jelectronics\j installed during the second servicing mission are performing well.
Among Hubble's recent observations:
Jets and Gaseous Disk Around the Egg Nebula -- A new infrared instrument peered deep into the dust-obscured central region around a dying star embedded in the Egg nebula. A nebula is a cloud of dust and gas 3,000 light years from Earth.
The new images provide a clear view of a twin pair of narrow bullet-shaped "jets" of gas and dust blasted into space. The instrument, called the Near Infrared Camera and Multi-Object Spectrometer, also revealed an unusual scalloped edge along a doughnut-shaped molecular \Jhydrogen\j cloud in the nebula.
"Because we can now see these 'missing pieces' in infrared and visible light, we have a more complete view of the dynamic and complicated structure of the star," said Rodger Thompson of the University of \JArizona\j, Tucson, the principal investigator for the infrared instrument. "It also allows us to see a 'fossil record' of the star's late evolutionary stages."
Unveiling Violent Starbirth in the Orion Nebula - The new infrared instrument penetrated the shroud of dust along the back wall of the Orion nebula, located in the "sword" of the \Jconstellation\j Orion. Data revealed what can happen to a stellar neighborhood when massive young stars begin to violently eject material into the surrounding molecular cloud.
Although ground-based infrared cameras have previously observed this hidden region known as OMC-1, the Hubble's new instrument provides the most detailed look yet at the heart of this giant molecular cloud. Hubble reveals a surprising array of complex structures, including clumps, bubbles, and knots of material. Most remarkable are "bullets" composed of molecular \Jhydrogen\j -- the fastest of which travels at more than one million mph (500 km/s). These bullets are colliding with slower-moving material, creating bow shocks, like a speedboat racing across water.
Monster Black Hole in Galaxy M84 - In a single exposure, a new powerful instrument called the Space \JTelescope\j Imaging Spectrograph discovered a black hole at least 300 million times the mass of the Sun. The spectrograph made a precise observation along a narrow slit across the center of galaxy M84, located 50 million light-years away.
This allowed the instrument to measure the increasing velocity of a disk of gas orbiting the black hole. To scientists, this represents the signature of a black hole, among the most direct evidence obtained to date. Due to their nature, it's impossible to directly photograph black holes. Scientists must instead look for clues to show the effects of black holes on surrounding dust, gas and stars.
"Hubble proved the existence of supermassive black holes three years ago," said Bruce Woodgate of the Goddard Space Flight Center, Greenbelt, MD, and principal investigator for the new spectrograph. "With this new instrument, we can do it 40 times faster than we used to."
Composition and Structure of the Ring Around \JSupernova\j 1987A - The new spectrograph also provides an unprecedented look at a unique and complex structure in the universe -- a light-year-wide ring of glowing gas around \JSupernova\j 1987A, the closest \Jsupernova\j explosion in 400 years.
The spectrograph dissects the ring's light to tell scientists which elements are in the ring and helps paint a picture of the physics and stellar processes which created the ring. This gives astronomers better insight into how stars evolve and become a \Jsupernova\j, and into the origin of the chemical elements created in these massive explosions.
Hubble Status -- NASA project officials are encouraged that a problem detected earlier with one of the cameras on the infrared instrument has shown some improvement. The problem stems from the unexpected movement of the dewar -- an insulated vessel containing solid \Jnitrogen\j at extremely cold temperatures.
After launch, the \Jnitrogen\j expanded more than expected as it warmed, moving the dewar into contact with another surface in the mechanism and pushing one of the cameras out of its range of focus. The camera has moved back about one-third of the distance required to be within reach of the instrument's internal focusing mechanism.
This is because the dewar is "relaxing" toward its normal state, as pressure caused by the expansion of the \Jnitrogen\j is reduced. The ice keeps the sensitive infrared detector cooled. Project officials also are considering how to deal with unexpected, excessive coolant loss.
"We are anticipating a shorter lifetime for the instrument, but we don't know how much shorter," said Goddard Hubble Project Scientist David Leckrone. "We are taking steps to work around the problem, and will increase the percentage of time this instrument will be used."
NASA officials also report that other upgrades to Hubble are performing well, including the newly installed solid state recorder, fine guidance sensor and solar array drive \Jelectronics\j. The solid state recorder has significantly improved data storage and playback, and the new fine guidance sensor is by far the best of the three on Hubble.
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"Black Hole Signature Recorded By STIS",154,0,0,0
May 12, 1997
The colorful "zigzag" found is not the work of a flamboyant artist, but the signature of a supermassive black hole in the center of galaxy M84, discovered by Hubble Space \JTelescope\j's Space \JTelescope\j Imaging Spectrograph (STIS).
The image, taken with Hubble's Wide Field Planetary and Camera 2 shows the core of the galaxy where the suspected black hole dwells. Astronomers mapped the motions of gas in the grip of the black hole's powerful gravitational pull by aligning the STIS's spectroscopic slit across the nucleus in a single exposure.
The STIS data shows the rotational motion of stars and gas along the slit. The change in wavelength records whether an object is moving toward or away from the observer. The larger the excursion from the centerline, the greater the rotational velocity. If no black hole were present, the line would be nearly vertical across the scan.
Instead, STIS's detector found the S-shape at the center of this scan, indicating a rapidly swirling disk of trapped material encircling the black hole. Along the S-shape from top to bottom, velocities skyrocket as seen in the rapid, dramatic swing to the left (blueshifted or approaching gas), then the region in the center simultaneously records the enormous speeds of the gas both approaching and receding for orbits in the immediate vicinity of the black hole, and then an equivalent swing from the right, back to the center line.
STIS measures a velocity of 880,000 miles per hour (400 kilometers per second) within 26 light-years of the galaxy's center, where the black hole dwells. This motion allowed astronomers to calculate that the black hole contains at least 300 million solar masses. (Just as the mass of our Sun can be calculated from the orbital radii and speeds of the planets.)
This observation demonstrates a direct connection between a supermassive black hole and activity (such as radio emission) in the nucleus of an active galaxy. It also shows that STIS is ideally suited for efficiently conducting a survey of galaxies to determine the distribution of the black holes and their masses.
Each point on STIS's solid-state CCD (Charge Coupled Device) detector samples a square patch at the galaxy that is 12 light-years on a side. The detection of black holes at the centers of galaxies is about 40 times faster than the earlier Faint Object Spectrograph. STIS was configured to record five spectral features in red light from glowing \Jhydrogen\j atoms as well as \Jnitrogen\j and sulfur ions in orbit around the center of M84. At each sampled patch the velocity of the entrapped gas was measured. Because the patches are contiguous, the astronomers could map the change of velocity in detail.
M84 is located in the Virgo Cluster of galaxies, 50 million light-years from Earth.
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"NICMOS Captures The Heart Of OMC-1",155,0,0,0
May 12, 1997
The infrared vision of the Hubble Space \JTelescope\j's Near Infrared Camera and Multi-Object Spectrometer (NICMOS) is providing a dramatic new look at the beautiful Orion Nebula which contains the nearest nursery for massive stars.
For comparison, Hubble's Wide Field and Planetary Camera 2 (WFPC2) image shows a large part of the nebula as it appears in visible light. The heart of the giant Orion molecular cloud, OMC-1, is included in the relatively dim and featureless area inside the blue outline near the top of the image. Light from a few foreground stars seen in the WFPC2 image provides only a hint of the many other stars embedded in this dense cloud.
NICMOS's infrared vision reveals a chaotic, active star birth region. Here, stars and glowing interstellar dust, heated by and scattering the intense starlight, appear yellow-orange. Emission by excited \Jhydrogen\j molecules appears blue. The image is oriented with north up and east to the left. The diagonal extent of the image is about 0.4 light-years. Some details are as small as the size of our solar system.
The brightest object in the image is a massive young star called BN (Becklin-Neugebauer). Blue "fingers" of molecular \Jhydrogen\j emission indicate the presence of violent outflows, probably produced by a young star or stars still embedded in dust (located to the lower left, southeast, of BN).
The outflowing material may also produce the crescent-shaped "bow shock" on the edge of a dark feature north of BN and the two bright "arcs" south of BN. The detection of several sets of closely spaced double stars in these observations further demonstrates NICMOS's ability to see fine details not possible from ground-based telescopes.
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"Seyfert Galaxy NGC 4151, Fireworks in",156,0,0,0
June 9, 1997
The Space \JTelescope\j Imaging Spectrograph (STIS) simultaneously records, in unprecedented detail, the velocities of hundreds of gas knots streaming at hundreds of thousands of miles per hour from the nucleus of NGC 4151, thought to house a supermassive black hole.
This is the first time the velocity structure in the heart of this object, or similar objects, has been mapped so vividly this close to its central black hole.
The twin cones of gas emission are powered by the energy released from the supermassive black hole believed to reside at the heart of this Seyfert galaxy. The STIS data clearly show that the gas knots illuminated by one of these cones is rapidly moving towards us, while the gas knots illuminated by the other cone are rapidly receding.
The highest velocity material expelled in a cataclysmic stellar explosion ten years ago has been detected for the first time by the Hubble Space \JTelescope\j's Space \JTelescope\j Imaging Spectrograph (STIS).
An image taken with Hubble's Wide Field and Planetary Camera 2 in 1995, shows orange-red rings surrounding \JSupernova\j 1987A in the Large Magellanic Cloud. The glowing debris of the \Jsupernova\j explosion, which occurred in February 1987, is at the center of the inner ring.
The STIS spectrograph viewed the entire inner ring in far-ultraviolet light, spreading it into a spectrum. The Earth's atmosphere completely blocks ultraviolet radiation from reaching the Earth's surface, hence astronomers can study the ultraviolet universe only from orbiting telescopes.
The STIS data shows the presence of glowing \Jhydrogen\j expanding at a speed of 33 million miles per hour (15,000 kilometers per second) coming from an extended area inside the inner ring. In addition to \Jhydrogen\j emission STIS also detected emission from high-velocity ionized \Jnitrogen\j.
This is the first time that astronomers have measured the very fast moving gas ejected by the \Jsupernova\j explosion, which was invisible until observed by Hubble with the STIS ultraviolet detectors. This gas is glowing in the ultraviolet because it is slamming into the remains of the gas lost by the \Jsupernova\j star about 20,000 years before it exploded.
The STIS spectrum also reveals the presence of emissions from hot gasses (oxygen, \Jnitrogen\j, and helium) coming from the inner ring itself. The ring is about 1.2 light-years in diameter.
Supernova 1987A is located 167,000 light-years from Earth in the Large Magellanic Cloud.
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"Gamma-Ray Bursts Common To Normal Galaxies?",158,0,0,0
June 10, 1997
Nature's most powerful explosions, gamma-ray bursts, occur among the normal stellar population inside galaxies scattered across the universe. This means that, on average, a gamma-ray burst goes off once every few million years inside our Milky Way galaxy.
The energy released in such a \Jtitanic\j explosion, which can last from a fraction of a second to a few hundred seconds, is equal to all of the Sun's energy generated over its 10 billion year lifetime.
A team of astronomers, led by Kailash Sahu of the Space \JTelescope\j Science Institute (STScI) in Baltimore, MD, is reporting this conclusion at the 190th Meeting of the American Astronomical Society in \JWinston-Salem\j, NC. They used the Hubble Space \JTelescope\j to study the fading optical counterpart to a burst that happened on February 28, 1997.
The team's results are based on Hubble images taken on March 26 and April 7, 1997 that show the gamma-ray burst is offset from the center of an extended "fuzzy" object that looks like a galaxy. The researchers argue, statistically, that this must be more than a chance alignment. The gamma-ray burst was embedded inside the galaxy.
"Hubble's unmatched resolution was crucial in pinpointing the fact that the gamma-ray burst is away from the center," Sahu says. "This would rule out massive black holes, thought to dwell in the cores of most galaxies, as the source of these incredible explosions."
"These observations definitely represent a huge step forward towards the full understanding of these enigmatic objects," says co-investigator Mario Livio, of STScI. However, he points out that this conclusion assumes the mechanism for creating a gamma-ray burst is basically the same.
Keck \Jtelescope\j spectroscopic measurements of the optical counterpart to another gamma-ray burst, which exploded on May 8, found that its distance from Earth is several billion light-years. The Keck observation establishes that gamma-ray bursts are truly extragalactic in location and origin.
However, a June 2nd Hubble observation of this newer burst source, made with both of the newly installed science instruments, failed to reveal a galaxy adjacent to the optical counterpart.
"If the gamma-ray burster is at the distance indicated by the Keck spectrum, then its host galaxy is far less luminous than is the Milky Way," says Andrew Fruchter, one of the leaders of another STScI team in making the observation. "What is clear is that further observations of new bursters, and of these two bursters at later times, will be required to better understand the nature and location of these astonishing objects."
One possible mechanism for unleashing such a \Jtitanic\j fireball of energy is the collision of a neutron star with another neutron star or a black hole. The Hubble observations support this model because it appears gamma-ray bursts occur in the disk of a galaxy, where there is ongoing stellar formation, and so there should be an abundance of neutron stars from recently exploded supernovae.
The astronomers estimate a gamma-ray burst may explode within a few thousand light-years from Earth every few hundred million years.
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"Fireball From A Cataclysmic Explosion",159,0,0,0
June 10, 1997
The CCD camera (Charge Coupled Device) on the Space \JTelescope\j Imaging Spectrograph, a new instrument on Hubble Space \JTelescope\j, captured a visible fireball from a \Jtitanic\j explosion in deep space, called a gamma-ray burst.
The burst occurred on May 8, and Hubble observations to acquire the fading fireball were made on June 2. No accompanying object, such as a host galaxy, can be found near the burst. This result adds to the puzzlement over of the source of these enigmatic explosions, because a previous Hubble image of another gamma-ray burst counterpart identified a potential host galaxy. If a galaxy is present, and at the distance suggested by Keck \Jspectroscopy\j, it is much fainter than our Milky Way. A few faint galaxies are, however, seen several arcseconds from the source. If one of these is the host, then the gamma-ray burst is very far out in the galaxy's halo, well outside the galaxy's stellar disk.
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"Hubble Captures Volcanic Eruption Plume From Io",160,0,0,0
The Hubble Space \JTelescope\j has snapped a picture of a 400-km-high (250-mile-high) plume of gas and dust from a volcanic eruption on Io, Jupiter's large innermost moon.
Io was passing in front of Jupiter when this image was taken by the Wide Field and Planetary Camera 2 in July 1996. The plume appears as an orange patch just off the edge of Io in the eight o'clock position, against the blue background of Jupiter's clouds. Io's volcanic eruptions blasts material hundreds of kilometers into space in giant plumes of gas and dust. In this image, material must have been blown out of the \Jvolcano\j at more than 2,000 mph to form a plume of this size, which is the largest yet seen on Io.
Until now, these plumes have only been seen by \Jspacecraft\j near Jupiter, and their detection from the Earth-orbiting Hubble Space \JTelescope\j opens up new opportunities for long-term studies of these remarkable phenomena.
Io's volcanic plumes are much taller than those produced by terrestrial volcanos because of a combination of factors. The moon's thin atmosphere offers no resistance to the expanding volcanic gases; its weak gravity (one-sixth that of Earth) allows material to climb higher before falling; and its biggest volcanos are more powerful than most of Earth's volcanos.
This image is a contrast-enhanced composite of an ultraviolet image (2600 Angstrom wavelength), shown in blue, and a violet image (4100 Angstrom wavelength), shown in orange. The orange color probably occurs because of the absorption and/or scattering of ultraviolet light in the plume. This light from Jupiter passes through the plume and is absorbed by sulfur dioxide gas or is scattered by fine dust, or both, while violet light passes through unimpeded. Future HST observations may be able to distinguish between the gas and dust explanations.
#
"Mariner Program",161,0,0,0
A series of \Jspacecraft\j launched by NASA to begin the exploration of the inner and outer Solar System. The program included the first planetary flyby (Mariner 2, Venus, December 1962), the first planetary orbiter (Mariner 9, Mars, November 1971), and the first mission to Mercury (Mariner 10, Mar 1974, September 1974, and Mar 1975). A brief summary of some mission highlights follows.
\BMariner 2
Launch Date/Time:\b 1962-08-27 at 06:53:13 UTC
\BDescription\b
The Mariner 2 \Jspacecraft\j was the second of a series of \Jspacecraft\j used for planetary exploration in the flyby, or nonlanding, mode. Mariner 2 was a backup for the Mariner 1 mission which failed shortly after launch to Venus. The \Jspacecraft\j was attitude-stabilized using the Sun and Earth as references. It was solar powered and capable of continuous telemetry operation.
The \Jspacecraft\j obtained data on the interplanetary medium during the flight to Venus and beyond, and it obtained planetary data during the encounter of Venus. The \Jspacecraft\j passed Venus at a distance of 41,000 km on December 14, 1962.
\BMariner 5
Launch Date/Time:\b 1967-06-14 at 06:01:00 UTC
\BDescription\b
The Mariner 5 \Jspacecraft\j was the fifth in a series of \Jspacecraft\j used for planetary exploration in the flyby mode. Mariner 5 was a refurbished backup \Jspacecraft\j for the Mariner 4 mission and was converted from a Mars mission to a Venus mission. The \Jspacecraft\j was fully attitude stabilized, using the sun and Canopus as references. A central computer and sequencer subsystem supplied timing sequences and computing services for other \Jspacecraft\j subsystems.
The \Jspacecraft\j passed 4,000 km from Venus on October 19, 1967. The \Jspacecraft\j instruments measured both interplanetary and Venusian magnetic fields, charged particles, and plasmas, as well as the radio refractivity and UV emissions of the Venusian atmosphere. The mission was termed a success.
\BMariner 6
Launch Date/Time:\b 1969-02-24 at 01:29:00 UTC
\BDescription\b
Mariner 6 was the sixth in a series of \Jspacecraft\j used for planetary exploration in the flyby mode. Mariner 6 was attitude stabilized in three axes (referenced to the sun and the star, Canopus). The \Jspacecraft\j was solar powered and capable of continuous telemetry transmission. It was fully automatic in operation, although it could be reprogrammed from earth during the mission.
The \Jspacecraft\j was oriented entirely to planetary data acquisition, and no data were obtained during the trip to Mars or beyond Mars. Mariner 6 passed 3,431 km from Mars on July 31, 1969. The \Jspacecraft\j instruments took TV images of Mars and measured the radio refractivity and UV and IR emissions of the Martian atmosphere. The mission was a success, and data from it were used to program Mariner 7.
\BMariner 9
Launch Date:\b 1971-05-30
\BDescription\b
The Mariner 9 Mars mission was planned to consist of two \Jspacecraft\j on complementary missions, but due to the failure of Mariner 8 to launch properly, only one \Jspacecraft\j was available. Mariner 9 combined mission objectives of both Mariner 8 (mapping 70 % of the Martian surface) and Mariner 9 (a study of temporal changes in the Martian atmosphere and on the Martian surface). For the survey portion of the mission, the planetary surface was to be mapped with the same resolution as planned for the original mission, although the resolution of pictures of the polar regions would be decreased due to the increased slant range.
The variable features experiments were changed from studies of six given areas every 5 days to studies of smaller regions every 17 days. Mariner 9 arrived at Mars on November 14, 1971. The \Jspacecraft\j gathered data on the atmospheric composition, density, pressure, and temperature and also the surface composition, temperature, and \Jtopography\j of Mars. After depleting its supply of attitude control gas, the \Jspacecraft\j was turned off October 27,1972.
\BMariner 10
Launch Date/Time:\b 1973-11-03 at 05:45:00 UTC
\BDescription\b
Mariner 10 was the seventh successful launch in the Mariner series and the first \Jspacecraft\j to use the gravitational pull of one planet (Venus) to reach another (Mercury). The \Jspacecraft\j structure was an eight-sided framework with eight \Jelectronics\j compartments. It measured 1.39 m diagonally and 0.457 m in depth. Two solar panels, each 2.7 m long and 0.97 m wide, were attached at the top, supporting 5.1 sq m of solar cell area. The rocket engine was liquid-fueled, with two sets of reaction jets used to stabilize the \Jspacecraft\j on three axes.
It carried a low-gain omnidirectional antenna, composed of a honeycomb-disk parabolic reflector, 1.37 m in diameter, with focal length 55 cm. Feeds enabled the \Jspacecraft\j to transmit at S- and X-band frequencies. The \Jspacecraft\j carried a Canopus star tracker, located on the upper ring structure of the octagonal satellite, and acquisition sun sensors on the tips of the solar panels. The interior of the \Jspacecraft\j was insulated with multilayer thermal blankets at top and bottom. A sunshade was deployed after launch to protect the \Jspacecraft\j on the solar-oriented side.
Instruments on-board the \Jspacecraft\j measured the atmospheric, surface, and physical characteristics of Mercury and Venus. Experiments included \Jtelevision\j photography, magnetic field, plasma, infrared \Jradiometry\j, ultraviolet \Jspectroscopy\j, and radio science detectors. An experimental X-band, high-frequency transmitter was flown for the first time on this \Jspacecraft\j.
Mariner 10 was placed in a parking orbit after launch for approximately 25 minutes, then placed in orbit around the Sun en route to Venus. The orbit direction was opposite to the motion of the Earth around the Sun. Mid-course corrections were made. The \Jspacecraft\j passed Venus on February 5, 1974, at a distance of 4,200 km. It crossed the orbit of Mercury on March 29, 1974, at 2046 UT, at a distance of about 704 km from the surface. The TV and UV experiments were turned on the \Jcomet\j Kohoutek while the \Jspacecraft\j was on the way to Venus.
A second encounter with Mercury, when more photographs were taken, occurred on September 21, 1974, at an altitude of about 47,000 km. A third and last Mercury encounter at an altitude of 327 km, with additional photography of about 300 photographs and magnetic field measurements occurred on March 16, 1975. \JEngineering\j tests were continued until March 24, 1975, when the supply of attitude-control gas was depleted and the mission was terminated.
#
"Mars Observer Project (detailed)",162,0,0,0
US Mars orbiter \Jspacecraft\j launched in September 1992 and designed to acquire global maps of the surface chemistry, \Jmineralogy\j, and elevations, as well as to provide very high resolution (3 m / 7 ft) surface images and measurements of magnetic field and atmosphere dynamics. The \Jspacecraft\j was equipped with a French radio receiver to communicate with balloons and landed Russian \Jspacecraft\j to be launched in 1994 and 1996. Contact with the \Jspacecraft\j was lost on 21 August 1993, three days before the scheduled Mars orbit insertion.
\BMars Observer
Launch Date/Time:\b 1992-09-25 at 17:05:01 UTC
\BDescription \b
Mars Observer, the first of the Observer series of planetary missions, was designed to study the geoscience and climate of Mars. The primary science objectives for the mission were to:
(1) determine the global elemental and mineralogical character of the surface material;
(2) define globally the \Jtopography\j and gravitational field;
(3) establish the nature of the Martian magnetic field;
(4) determine the temporal and spatial distribution, abundance, sources, and sinks of volatiles and dust over a seasonal cycle; and,
(5) explore the structure and circulation of the atmosphere.
The bus and \Jelectronics\j of the Observer series of \Jspacecraft\j, used to study the terrestrial planets and near-Earth \Jasteroids\j, were derived from the Satcom-K and DMSP/TIROS \Jspacecraft\j. The rectangular bus section was 2.1 x 1.5 x 1.1 m. During the cruise phase of the mission, the high-gain antenna and the booms for the magnetometer (MAG/ER) and gamma-ray spectrometer (GRS) were partially deployed.
When fully deployed, the two booms were each 6 m long. The 1.5 m diameter high-gain antenna was, when fully deployed, on a 5.5 m boom to allow for clearance over the solar array when the antenna was pointed toward Earth. Pointing control for the \Jspacecraft\j was maintained through the use of four reaction wheels. Attitude information was provided by a horizon sensor (which defined the direction of the nadir), a star mapper (for inertial attitude), gyros and accelerometers (for measuring angular rates and linear accelerations), and multiple Sun sensors.
Power was provided through a six-panel solar array which, when fully deployed, measured 7.0 x 3.7 m. During the cruise phase, however, only four panels were deployed (due to the proximity of the \Jspacecraft\j to the sun) to reduce the amount of power generated. During periods when the \Jspacecraft\j was in Mars' shadow, energy was provided by two Ni-Cd batteries, each with a capacity of 43 amp-hours.
The interplanetary cruise phase of the mission was intended primarily for \Jspacecraft\j and instrument checkout and \Jcalibration\j. Two periods of data collection for the MAG/ER and GRS and one for the gravity wave experiment were planned for this phase as well. During the four month period from Mars orbital insertion until the \Jspacecraft\j achieved its final mapping orbit, only data collection for the MAG/ER, GRS, and thermal emission spectrometer (TES) were scheduled.
The mapping phase of the mission was scheduled to nominally last one Martian year. Mars Observer also supported the acquisition of data from the Russian Mars 1994 mission through the use of the joint French-Russian-American Mars \JBalloon\j Relay instrument. Contact with Mars Observer was lost on August 21, 1993, three days before scheduled orbit insertion, for unknown reasons and has not been re-established.
It is not known whether the \Jspacecraft\j was able to follow its automatic programming and go into Mars orbit or if it flew by Mars and is now in a heliocentric orbit. Although none of the primary objectives of the mission were achieved, cruise mode data were collected up to loss of contact.
Sponsoring Agencies/Countries included:
ò NASA-Office of Space Science Applications/United States
ò Center National d'Etudes Spatiales/France
ò Russian Space Agency
#
"Pioneer Venus Program",163,0,0,0
The Pioneer Venus mission consisted of two components, launched separately: an Orbiter and a Multiprobe.
\BThe Pioneer Venus Orbiter
Launch Date:\b 20 May 1978
\BLaunch Vehicle:\b Atlas-Centaur
\BDescription\b
The Pioneer Venus Orbiter was inserted into an elliptical orbit around Venus on December 4, 1978. The Orbiter was a flat cylinder 2.5 m in diameter and 1.2 m high. All instruments and \Jspacecraft\j subsystems were mounted on the forward end of the cylinder, except the magnetometer, which was at the end of a 4.7 m boom. A solar array extended around the \Jcircumference\j of the cylinder. A 1.09 m despun dish antenna provided S and X band communication with Earth.
The Pioneer Venus Orbiter carried 17 experiments (with a total mass of 45 kg):
ò a cloud photopolarimeter to measure the vertical distribution of the clouds.
ò a surface radar mapper to determine \Jtopography\j and surface characteristics.
ò an infrared radiometer to measure IR emissions from the Venus atmosphere.
ò an airglow ultraviolet spectrometer to measure scattered and emitted UV light.
ò a neutral mass spectrometer to determine the composition of the upper atmosphere.
ò a solar wind plasma analyzer to measure properties of the solar wind.
ò a magnetometer to characterize the magnetic field at Venus.
ò an electric field detector to study the solar wind and its interactions.
ò an \Jelectron\j temperature probe to study the thermal properties of the \Jionosphere\j.
ò an ion mass spectrometer to characterize the ionospheric ion population.
ò a charged particle retarding potential analyzer to study ionospheric particles.
ò two radio science experiments to determine the gravity field of Venus.
ò a radio \Joccultation\j experiment to characterize the atmosphere.
ò an atmospheric drag experiment to study the upper atmosphere.
ò a radio science atmospheric and solar wind turbulence experiment.
ò a gamma ray burst detector to record gamma ray burst events.
From Venus orbit insertion to July 1980, periapsis was held between 142 and 253 km (at 17 degrees north latitude) to facilitate radar and ionospheric measurements. The \Jspacecraft\j was in a 24 hour orbit with an apoapsis of 66,900 km. Thereafter, the periapsis was allowed to rise (to 2,290 km at maximum) and then fall, to conserve fuel. In 1991, the Radar Mapper was reactivated to investigate previously inaccessible southern portions of the planet.
In May, 1992, Pioneer Venus began the final phase of its mission, in which the periapsis was held between 150 and 250 km until the fuel ran out and atmospheric entry destroyed the \Jspacecraft\j the following August.
\BPioneer Venus Multiprobe
Launch Date:\b 08 August 1978
\BLaunch Vehicle:\b Atlas-Centaur
\BDescription\b
The Pioneer Venus Multiprobe consisted of a bus which carried one large and three small atmospheric probes. The large probe was released on November 16, 1978 and the three small probes on November 20. All four probes entered the Venus atmosphere on December 9, followed by the bus.
The Pioneer Venus large probe was equipped with 7 science experiments, contained within a sealed spherical pressure vessel. This pressure vessel was encased in a nose cone and aft protective cover. After deceleration from initial atmospheric entry at about 11.5 km/s near the equator on the Venus night side, a parachute was deployed at 47 km altitude. The large probe was about 1.5 m in diameter and the pressure vessel itself was 73.2 cm in diameter. The science experiments were:
ò a neutral mass spectrometer to measure the atmospheric composition.
ò a gas chromatograph to measure the atmospheric composition.
ò a solar flux radiometer to measure solar flux penetration in the atmosphere.
ò an infrared radiometer to measure distribution of infrared radiation.
ò a cloud particle size spectrometer to measure particle size and shape.
ò a nephelometer to search for cloud particles.
ò temperature, pressure, and acceleration sensors.
The three small probes were identical to each other, 0.8 m in diameter. These probes also consisted of spherical pressure vessels surrounded by an aeroshell, but unlike the large probe, they had no parachutes and the aeroshells did not separate from the probe. Each small probe carried a nephelometer and temperature, pressure, and acceleration sensors, as well as a net flux radiometer experiment to map the distribution of sources and sinks of radiative energy in the atmosphere.
The radio signals from all four probes were also used to characterize the winds, turbulence, and propagation in the atmosphere. The small probes were each targeted at different parts of the planet and were named accordingly. The North probe entered the atmosphere at about 60 degrees north latitude on the day side. The night probe entered on the night side. The day probe entered well into the day side, and was the only one of the four probes which continued to send radio signals back after impact, for over an hour.
The Pioneer Venus bus also carried two experiments, a neutral mass spectrometer and an ion mass spectrometer to study the composition of the atmosphere. With no heat shield or parachute, the bus survived and made measurements only to about 110 km altitude before burning up. The bus was a 2.5 m diameter cylinder weighing 290 kg, and afforded us our only direct view of the upper Venus atmosphere, as the probes did not begin making direct measurements until they had decelerated lower in the atmosphere.
#
"Sakigake and Suisei Projects",164,0,0,0
The first Japanese interplanetary \Jspacecraft\j, launched to intercept Halley's \Jcomet\j. Sakigake ('forerunner') was launched (January 1985) as a test \Jspacecraft\j for Suisei ('comet', launched August 1985). They were instrumented to measure solar wind interaction and the \Jhydrogen\j cloud of the \Jcomet\j. They were built and operated by the Institute for Space and Aeronautical Science at the University of \JTokyo\j.
\BSakigake
Launch Date/Time:\b 1985-01-08 at 19:26:00 UTC
\BDescription\b
Sakigake is a test \Jspacecraft\j similar to Suisei. It flew by \JComet\j P/Halley on its sunward side at a distance of about 7 million kilometers on March 11, 1986. It carries three instruments to measure plasma wave spectra, solar wind ions, and interplanetary magnetic fields, all of which worked normally. The \Jspacecraft\j is spin-stabilized at two different rates (5 and 0.2 rpm). It is equipped with hydrazine thrusters for attitude and velocity control, star and sun sensors for attitude determination, and a mechanically despun off-set parabolic dish for long-range communication.
Sakigake made an Earth swingby on January 8, 1992. The closest approach was at 23h 08m 47s (JST = UTC+9h) with a geocentric distance of 88,997 km. This was the first planet-swingby for a Japanese \Jspacecraft\j. During the approach, Sakigake observed the geotail. Some geotail passages are scheduled during ISTP's multi-spacecraft investigation of that region.
The second Earth swingby was on June 14, 1993 at 40 Re, and the third on October 28, 1994 at 86 Re. Further mission planning targets a 23.6 km/s, 10,000 km flyby of \JComet\j P/Honda-Mrhos-Pajdusakova on Feb 3, 1996 at about 21.00 GMT (approaching the nucleus along the tail) some 0.17 AU from the Sun, and a 14 million km passage of \JComet\j P/Giacobini-Zinner on Nov 29, 1998.
\BSuisei
Launch Date/Time:\b 1985-08-18 at 23:33:00 UTC
\BDescription \b
Suisei (the Japanese name meaning 'Comet') was launched on March 18, 1985 into heliocentric orbit to fly by \JComet\j P/Halley. It is identical to Sakigake apart from its payload: a CCD UV imaging system and a solar wind instrument. The main objective of the mission was to take UV images of the \Jhydrogen\j corona for about 30 days before and after \JComet\j Halley's descending crossing of the \Jecliptic\j plane.
Solar wind parameters were measured for a much longer time period. The \Jspacecraft\j is spin-stabilized at two different rates (5 and 0.2 rpm). Hydrazine thrusters are used for attitude and velocity control; star and sun sensors are for attitude control; and a mechanically despun off-set parabolic dish is used for long range communication.
#
"Soviet/Russian Space Programs continued",165,0,0,0
This section summarizes some of the major achievements of the Soviet Space Program. This program was replaced by the Russian Space Agency after the demise of the Soviet Union in 1992.
The programs discussed in this section include the following:
\BEarly Soviet Satellites\b - Including \JSputnik\j, Vostok and Soyuz
\BSpace Stations\b - Salyut and Mir
\BSoviet Lunar Missions\b - Luna Series and Zond Series
\BVenera Program\b - Covering Venera missions to Venus
\BVega Project\b - Covering missions to Venus and Halley's \JComet\j
\BPhobos\b - Mission to Mars
A summary of each of the above projects follows.
\B Early Soviet Satellites\b
\B\ISputnik Satellites\b\i
The first Soviet satellite program consisted of four \JSputnik\j satellites. However, the \JSputnik\j launched between \JSputnik\j 2 and 3 failed to reach orbit.
Sputnik 1, launched on October 4, 1957, was designed to send radio signals to Earth and determine the density of the upper atmosphere. However, it only transmitted signals to Earth for a short time after launch. Its orbit decayed and it fell to Earth on January 4, 1958. \JSputnik\j 1 had an orbital period of 98.6 minutes and an orbital altitude of 228-947 km.
Sputnik 2 was launched on November 3, 1957, and carried aboard it a dog, Laika. Biological data was returned for approximately a week (the first data of its kind). However, there was no safe re-entry possible at the time, and Laika was put to sleep after a week in orbit. The satellite itself remained in orbit 162 days. \JSputnik\j 2 had an orbital period of 103.75 minutes and an orbital altitude of 225-1,671 km.
Sputnik 3 was launched on May 15, 1958. It may originally have been intended as the first launch in the \JSputnik\j program, however it was apparently decided to be more cautious in the launch schedule. It was designed to be a geophysical laboratory, performing experiments on the Earth's magnetic field, radiation belt, and \Jionosphere\j. It orbited Earth and transmitted data until April 6, 1960, when its orbit decayed.
All Sputniks were launched using the SS-6, or Sapwood rocket. The SS-6 was originally designed as a ballistic missile, and had its upper stage modified slightly to hold the \JSputnik\j payload. It had two stages, four strap-on booster rockets for the first stage, connected to the second stage rocket. Total mass at launch for \JSputnik\j 3 was 267 tonnes, with a length of 29.17 meters. The primary stage used RD-107 engines, which provided 100,000 kg of thrust. Both stages were powered by LOX/Kerosene.
\B\IVostok Spacecraft\b\i
The first generation of Soviet crewed \Jspacecraft\j, carrying a single member. Vostok 1 carried the first human into space (12 April 1961) -- Yuri Gagarin, who orbited Earth once on a flight of 118 min. Crew were recovered over land after ejection from the capsule at 7,000 m / 23,000 ft altitude after re-entry. The last Vostok flight carried Valentina Tereshkova, the first woman to fly in space (Vostok 6, 16 June 1963).
Voskhod ('Sunrise') was an intermediate-generation Soviet-crewed \Jspacecraft\j following Vostok and preceding Soyuz; it made only two flights, in 1964 and 1965.
\B\ISoyuz\b\i
A Soviet basic space capsule, consisting of three modules (orbiter, descent, and instrumentation), and carrying a crew of one to three. It has been flown on several dozen missions, and is capable of precision targeting to soft-land in Central Asia (contrasting with the US ocean-recovery technique). It has been used to ferry crew to and from Salyut and, later, Mir space stations with which docking takes place. The first flight was in 1967, when its pilot, V Komarov, was killed in a landing accident.
Soyuz 19 docked with the Apollo \Jspacecraft\j in Earth orbit after its rendezvous (July 1975). There was a notable launch accident (September 1983) in which the crew of Soyuz T-10A ejected to safety as the launch vehicle exploded.
\BSpace Stations\b
\B\ISalyut\b\i
The first-generation Soviet space station, capable of docking with the Soyuz crew ferry and Progress resupply vehicle; it provides 100 cubic meters / 3,500 cu ft of living space for up to five cosmonauts. Two versions have been flown in a program spanning 1971 to the present, aimed at accumulating data on long-duration space-flight experience and biomedical experiments, Earth remote sensing, and \Jmicrogravity\j science. The first station was flown in 1971, the last (Salyut 7) in 1982. The station's orbit eventually decays, with the vehicle re-entering the atmosphere and burning up. Crews have accumulated many hundred days of flight experience. A notably dangerous repair mission was undertaken to Salyut 7 in 1985, by V Dzhanibekov and V Savinykh, after the station seriously malfunctioned between crew occupancies.
\B\IMir\b\i
A Soviet space station (launched February 1986) which evolved from Salyut, having more power (solar panels) and more docking ports (five) than previous \Jspacecraft\j, allowing for the build-up of a modular station. It is used for long-duration spaceflight experience, and for biomedical, science, and applications experiments. Yuri Romanenko occupied Mir for 326 days in 1987. The current space endurance record for men is held by Russian \Jcosmonaut\j Valeri Poliakov, who lived in space for 438 days (Jan 1994-Mar 1995), and for women by Yelena Kondakova.
\BSoviet Lunar Missions\b
The Soviet Lunar program had 20 successful missions to the Moon and achieved a number of notable lunar "firsts": first probe to impact the Moon, first flyby and image of the lunar farside, first soft landing, first lunar orbiter, and the first circumlunar probe to return to Earth. The two successful series of Soviet probes were the Luna (15 missions) and the Zond (5 missions).
Lunar flyby missions (Luna 3, Zond 3, 6, 7, 8) obtained photographs of the lunar surface, particularly the limb and farside regions. The Zond 6, 7, and 8 missions circled the Moon and returned to Earth where they were recovered, Zond 6 and 7 in \JSiberia\j and Zond 8 in the Indian Ocean. The purpose of the photography experiments on the lunar landers (Luna 9, 13, 22) was to obtain close-up images of the surface of the Moon for use in lunar studies and determination of the feasibility of manned lunar landings.
\B\IThe Luna Series
Luna 2
Launched 12 September 1959\b\i
Impacted Moon 13 September 1959 at 22:02:04 UT
Latitude 29.10 N, Longitude 0.00 -- Palus Putredinis
\B\ILuna 3
Launched 04 October 1959\b\i
Lunar Flyby
\B\ILuna 9
Launched 31 January 1966\b\i
Landed on Moon 03 February 1966 at 18:44:52 UT
Latitude 7.08 N, Longitude 295.63 E -- \JOceanus\j Procellarum
\B\ILuna 10
Launched 31 March 1966\b\i
Lunar Orbiter
\B\ILuna 11
Launched 24 August 1966\b\i
Lunar Orbiter
\B\ILuna 12
Launched 22 October 1966\b\i
Lunar Orbiter
\B\ILuna 13
Launched 21 December 1966\b\i
Landed on Moon 24 December 1966 at 18:01:00 UT
Latitude 18.87 N, 297.95 E -- \JOceanus\j Procellarum
\B\ILuna 14
Launched 7 April 1968\b\i
Lunar Orbiter
\B\ILuna 16
Launched 12 September 1970\b\i
Landed on Moon 20 September 1970 at 05:18:00 UT
Latitude 0.68 S, Longitude 56.30 E -- Mare Fecunditatis
Lunar Sample Return
\B\ILuna 17
Launched 10 November 1970\b\i
Landed on Moon 17 November 1970 at 03:47:00 UT
Latitude 38.28 N, Longitude 325.00 E -- Mare Imbrium
Lunar Rover -- Lunokhod 1
\B\ILuna 19
Launched 28 September 1971\b\i
Lunar Orbiter
\B\ILuna 20
Launched 14 February 1972\b\i
Landed on Moon 21 February 1972 at 19:19:00 UT
Latitude 3.57 N, Longitude 56.50 E -- Mare Fecunditatis
Lunar Sample Return to Earth 25 February 1972
\B\ILuna 21
Launched 08 January 1973\b\i
Landed on Moon 15 January 1973 at 23:35:00 UT
Latitude 25.51 N, Longitude 30.38 E -- Mare Serenitatis
Lunar Rover -- Lunokhod 2
\B\ILuna 22
Launched 02 June 1974\b\i
Lunar Orbiter
\B\ILuna 24
Launched 14 August 1976\b\i
Landed on Moon 18 August 1976 at 02:00:00 UT
Latitude 12.25 N, Longitude 62.20 E -- Mare Crisium
Lunar Sample Return
\B\IThe Zond Series
Zond 3
Launched 18 July 1965\b\i
Lunar Flyby
\B\IZond 5
Launched 15 September 1968\b\i
Circumlunar
Returned to Earth 21 September 1968
\B\IZond 6
Launched 10 November 1968\b\i
Circumlunar
Returned to Earth 17 November 1968
\B\IZond 7
Launched 07 August 1969\b\i
Circumlunar
Returned to Earth 14 August 1969
\B\IZond 8
Launched 20 October 1970\b\i
Circumlunar
Returned to Earth 27 October 1970
\BVenera Program\b
A highly successful evolutionary series of Soviet space missions to Venus 1961-83 (plus Vega landers and balloons of 1985). Its highlights include: the first successful atmospheric entry probe (Venera 4, 1967); the first complete descent to the surface (Venera 5, 1969); the first science measurements on the surface (Venera 7, 1970); the first TV pictures from the surface (Venera 9, 1975); the first chemical analysis of the soil (Venera 13, 1981); the first high resolution images of the surface from orbit using radar to penetrate clouds (Venera 15/16, 1983); and the first balloons deployed and tracked in the atmosphere (Vega, 1985).
\B\IVenera 4
Launch Date:\b\i 1967-06-12
\B\IOn-orbit dry mass:\b\i 1106.00 kg
\B\IDescription\b\i
Venera 4 was launched from a Tyazheliy \JSputnik\j (67-058B) towards the planet Venus with the announced mission of direct atmospheric studies. On October 18, 1967, the \Jspacecraft\j entered the Venusian atmosphere and released two thermometers, a \Jbarometer\j, a radio \Jaltimeter\j, and atmospheric density gauge, 11 gas analyzers, and two radio transmitters operating in the DM waveband.
The main bus, which had carried the capsule to Venus, carried a magnetometer, cosmic ray detectors, \Jhydrogen\j and oxygen indicators, and charged particle traps. Signals were returned by the \Jspacecraft\j, which braked and then deployed a parachute system after entering the Venusian atmosphere, until it reached an altitude of 24.96 km.
\B\IVenera 5
Launch Date:\b\i 1969-01-05
\B\IOn-orbit dry mass:\b\i 1130.00 kg
\B\IDescription\b\i
Venera 5 was launched from a Tyazheliy \JSputnik\j (69-001C) towards Venus to obtain atmospheric data. The \Jspacecraft\j was very similar to Venera 4 although it was of a stronger design. When the atmosphere of Venus was approached, a capsule weighing 405 kg and containing scientific instruments was jettisoned from the main \Jspacecraft\j. During satellite descent towards the surface of Venus, a parachute opened to slow the rate of descent.
For 53 minutes on May 16, 1969, while the capsule was suspended from the parachute, data from the Venusian atmosphere were returned. The \Jspacecraft\j also carried a medallion bearing the coat of arms of the USSR and a bas-relief of V. I. Lenin to the night side of Venus.
\B\IVenera 6
Launch Date:\b\i 1969-01-10
\B\IOn-orbit dry mass:\b\i 1130.00 kg
\B\IDescription\b\i
Venera 6 was launched from a Tyazheliy \JSputnik\j (69-002C) towards Venus to obtain atmospheric data. The \Jspacecraft\j was very similar to Venera 4 although it was of a stronger design. When the atmosphere of Venus was approached, a capsule weighing 405 kg was jettisoned from the main \Jspacecraft\j. This capsule contained scientific instruments. During descent towards the surface of Venus, a parachute opened to slow the rate of descent.
For 51 minutes on May 17, 1969, while the capsule was suspended from the parachute, data from the Venusian atmosphere were returned. The \Jspacecraft\j also carried a medallion bearing the coat of arms of the USSR and a bas-relief of V. I. Lenin to the night side of Venus.
\B\IVenera 7
Launch Date:\b\i 1970-08-17
\B\IOn-orbit dry mass:\b\i 1180.00 kg
\B\IDescription\b\i
Venera 7 was launched from a Tyazheliy \JSputnik\j in an earth parking orbit towards Venus to study the Venusian atmosphere and other phenomena of the planet. Venera 7 entered the atmosphere of Venus on December 15, 1970, and a landing capsule was jettisoned. After aerodynamic braking, a parachute system was deployed. The capsule antenna was extended, and signals were returned for 35 min.
Another 23 minutes of very weak signals were received after the \Jspacecraft\j landed on Venus. The capsule was the first man-made object to return data after landing on another planet.
\B\IVenera 9 Descent Craft
Description\b\i
On October 20, 1975, this \Jspacecraft\j was separated from the Orbiter, and landing was made with the sun near zenith at 0513 UT on October 22. A system of circulating fluid was used to distribute the heat load. This system, plus precooling prior to entry, permitted operation of the \Jspacecraft\j for 53 minutes after landing.
During descent, heat dissipation and deceleration were accomplished sequentially by protective hemispheric shells, three parachutes, a disk-shaped drag brake, and a compressible, metal, doughnut-shaped, landing cushion. The landing was about 2,200 km from the Venera 10 landing site.
Preliminary results indicated: (A) clouds 30-40 km thick with bases at 30-35 km altitude, (B) atmospheric constituents including HCl, HF, Br, and I, (C) surface pressure about 90 (earth) atmospheres, (D) surface temperature 485 deg C, (E) light levels comparable to those at earth mid-latitudes on a cloudy summer day, and (F) successful TV photography showing shadows, no apparent dust in the air, and a variety of 30-40 cm rocks which were not eroded.
\B\IVenera 11 Descent Craft
Launch Date:\b\i 1978-09-09
\B\IDescription\b\i
The Venera 11 descent craft carried instruments designed to study the detailed chemical composition of the atmosphere, the nature of the clouds, and the thermal balance of the atmosphere. Separating from its flight platform on December 25, 1978, it made a soft landing on the surface after a descent time of approximately 1 hour. During this time, it employed aerodynamic braking followed by parachute braking and ending with atmospheric braking.
The touchdown speed was 7-8 m/s. Information was transmitted to the flight platform for retransmittal to earth. It is unknown whether the Lander Probe carried an imaging system. No mention of it occurs in the Soviet literature examined by the author. Two other experiments on the Lander did fail, and their failure was acknowledged by the Soviets. Some US literature on the subject notes that the imaging system "failed" but did return some data.
\B\IVenera 12 Descent Craft
Launch Date:\b\i 1978-09-14
\B\IDescription \b\i
The Venera 12 descent craft carried instruments designed to study the detailed chemical composition of the atmosphere, the nature of the clouds, and the thermal balance of the atmosphere. Separating from its flight platform on December 21, 1978, it made a soft landing on the surface after a descent time of approximately 1 hour. During this time, it employed aerodynamic braking followed by parachute braking and ending with atmospheric braking.
The touchdown speed was 7-8 m/s. Information was transmitted to the flight platform for retransmittal to earth. It is unknown whether the Lander Probe carried an imaging system. No mention of it occurs in the Soviet literature examined by the author. Two other experiments on the Lander did fail, and their failure was acknowledged by the Soviets. Some US literature on the subject notes that the imaging system "failed" but did return some data.
\B\ITo continue please click\b\i \JSoviet/Russian Space Programs continued 2\j
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"Soviet/Russian Space Programs continued 2",166,0,0,0
\B\IVenera 13 Descent Craft
Launch Date:\b\i 1981-10-30
\B\IDescription\b\i
Venera 13 landed at 7 deg 30 minutes S by 303 deg, just east of the eastern extension of an elevated region known as Phoebe Regio. It survived for 2 h 7 minutes in an environment with a temperature of 457 deg C and a pressure of 89 earth atmospheres.
Venera 13 carried instruments to take chemical and isotopic measurements, monitored the spectrum of scattered sunlight, and recorded electric discharges during its descent phase through the Venusian atmosphere. The \Jspacecraft\j utilized a camera system, an X-ray \Jfluorescence\j spectrometer, and a seismometer to conduct investigations on the surface.
\B\IVenera 14 Descent Craft
Launch Date:\b\i 1981-11-04
\B\IDescription\b\i
Venera 14 landed at 13 deg 15 minutes S by 310 deg, about 950 km southwest of Venera 13. Surface temperature was 465 deg C and pressure was 94 earth atmospheres. Venera 14 carried instruments to take chemical and isotopic measurements, monitored the spectrum of scattered sunlight, and recorded electric discharges during its descent phase through the Venusian atmosphere. The \Jspacecraft\j utilized a camera system, an X-ray \Jfluorescence\j spectrometer, and a seismometer to conduct investigations on the surface.
\B\IVenera 15
Launch Date:\b\i 1983-06-02
\B\IOn-orbit dry mass:\b\i 4000.00 kg
\B\IDescription\b\i
Venera 15 was part of a two \Jspacecraft\j mission (along with Venera 16) designed to use 8 cm band side-looking radar maps to study the surface properties of Venus. The two \Jspacecraft\j were inserted into Venus orbit a day apart with their orbital planes shifted by an angle of approximately 4 degrees relative to one another. This made it possible to reimage an area if necessary. Each \Jspacecraft\j was in a nearly polar orbit with a periapsis at 62 N latitude. Together, the two \Jspacecraft\j imaged the area from the north pole down to about 30 degrees N latitude over the 8 months of mapping operations.
The Venera 15 and 16 \Jspacecraft\j were identical and were based on modifications to the orbiter portions of the Venera 9 and 14 probes. Each \Jspacecraft\j consisted of a 5 m long cylinder with a 6 m diameter, 1.4 m tall parabolic dish antenna for the synthetic aperture radar (SAR) at one end. A 1 meter diameter parabolic dish antenna for the radio \Jaltimeter\j was also located at this end. The electrical axis of the radio \Jaltimeter\j antenna was lined up with the axis of the cylinder. The electrical axis of the SAR deviated from the \Jspacecraft\j axis by 10 degrees.
During imaging, the radio \Jaltimeter\j would be lined up with the center of the planet (local vertical) and the SAR would be looking off to the side at 10 degrees. A bulge at the opposite end of the cylinder held fuel tanks and propulsion units. Two square solar arrays extended like wings from the sides of the cylinder. A 2.6 m radio dish antenna for communications was also attached to the side of the cylinder.
\B\IVenera 16
Launch Date:\b\i 1983-06-07
\B\IOn-orbit dry mass:\b\i 4000.00 kg
\B\IDescription\b\i
Venera 16 was part of a two \Jspacecraft\j mission (along with Venera 15) designed to use 8 cm band side-looking radar maps to study the surface properties of Venus. The two \Jspacecraft\j were inserted into Venus orbit a day apart with their orbital planes shifted by an angle of approximately 4 degrees relative to one another. This made it possible to reimage an area if necessary. Each \Jspacecraft\j was in a nearly polar orbit with a periapsis at 62 N latitude.
Together, the two \Jspacecraft\j imaged the area from the north pole down to about 30 degrees N latitude over the 8 months of mapping operations. In June 1984, Venus was at superior conjunction and passed behind the Sun as seen from Earth. No transmissions were possible, so the orbit of Venera 16 was rotated back 20 degrees at this time to map the areas missed during this period.
The Venera 15 and 16 \Jspacecraft\j were identical and were based on modifications to the orbiter portions of the Venera 9 and 14 probes. Each \Jspacecraft\j consisted of a 5 m long cylinder with a 6 m diameter, 1.4 m tall parabolic dish antenna for the synthetic aperture radar (SAR) at one end. A 1 meter diameter parabolic dish antenna for the radio \Jaltimeter\j was also located at this end. The electrical axis of the radio \Jaltimeter\j antenna was lined up with the axis of the cylinder. The electrical axis of the SAR deviated from the \Jspacecraft\j axis by 10 degrees.
During imaging, the radio \Jaltimeter\j would be lined up with the center of the planet (local vertical) and the SAR would be looking off to the side at 10 degrees. A bulge at the opposite end of the cylinder held fuel tanks and propulsion units. Two square solar arrays extended like wings from the sides of the cylinder. A 2.6 m radio dish antenna for communications was also attached to the side of the cylinder.
\BVega Project\b
Vega was a highly successful Soviet mission to Venus and Halley's \JComet\j undertaken in 1984-86. Two \Jspacecraft\j each deployed a lander and a \Jballoon\j at Venus, then used a Venus gravity assist to fly on to intercept Halley's \JComet\j, passing within an estimated 3,000 and 10,000 km of its nucleus. The mission carried an ambitious science payload, highly international in scope; all elements of the mission were boldly undertaken in public view and proved to be highly successful.
\B\IVega 1 and Vega 2
Launch Date:\b\i 1984-12-15 (Vega 1) and 1984-12-21 (Vega 2)
\B\IOn-orbit dry mass:\b\i 2500.00 kg
\B\IDescription\b\i
This \Jspacecraft\j mission combined a Venus swingby and a \JComet\j Halley flyby. Two identical \Jspacecraft\j, Vega 1 and Vega 2, were launched December 15 and 21, 1984, respectively. After carrying Venus entry probes to the vicinity of Venus (arrival and deployment of probes were scheduled for June 11-15, 1985), the two \Jspacecraft\j were to be retargetted using Venus gravity field assistance to intercept \JComet\j Halley in March 1986.
The first \Jspacecraft\j was to encounter \JComet\j Halley on March 6, 1986, and the second about three days later. The flyby velocity was to be 77.7 km/s. Although the \Jspacecraft\j could be targetted with a precision of 100 km, the position of the \Jspacecraft\j relative to the \Jcomet\j nucleus was estimated to be known only to within a few thousand kilometers. This, together with the problem of dust protection, led to estimated flyby distances of 10,000 km for the first \Jspacecraft\j and 3,000 km for the second.
The \Jspacecraft\j was three-axis stabilized. Its main features were large solar panels, a high-gain antenna dish, and an automatic pointing platform carrying those experiments that required pointing at the \Jcomet\j nucleus. The automatic platform could rotate through + or -110 degrees and + or -40 degrees in two perpendicular directions with a pointing accuracy of 5 arc-min and a stability of 1 arc-min/s. It carried the narrow- and the wide-angle camera, the three-channel spectrometer, and the infrared sounder.
All other experiments were body-mounted, with the exception of two magnetometer sensors and various plasma probes and plasma wave analyzers which were mounted on a 5-m boom. The total scientific payload weighed 125 kg and had a data rate of 65 kbs in fast telemetry mode for encounter. There was also a slow telemetry mode for the cruise mode. The comet-encounter science data-take was from 2.5 h before until 0.5 h after the closest approach, with several periods of data-take before and after, each lasting about 2 h.
Continuous coverage for plasma and dust instruments was provided by an onboard memory (5-megabit tape recorder). The \Jspacecraft\j was shielded from hypervelocity dust impacts by a shield consisting of a 100-micrometer multilayer sheet 20 to 30 cm from the \Jspacecraft\j, and a 1-mm Al sheet 5 to 10 cm from the \Jspacecraft\j.
Approximately half of the VEGA \Jspacecraft\j was devoted to the Halley module, and half to the Venus lander package. The total scientific payload weight was 144.3 kg. The Venus package consisted of a sphere 240 cm in diameter, which was to be separated two days before arrival at Venus and enter the planet's atmosphere on an inclined path, without active maneuvers, as was done on previous Venera missions. The lander probe was identical to those of Venera 9 through 14 and similarly had two objectives, the study of the atmosphere and the study of the superficial crust.
In addition to temperature and pressure measuring instruments, the descent probe carried a UV spectrometer for measurement of minor atmospheric constituents, an instrument dedicated to measurement of the concentration of water, and other instruments for determination of the chemical composition of the condensed phase: a gas-phase chromatograph; an X-ray spectrometer observing the \Jfluorescence\j of grains or drops; and a mass spectrograph measuring the chemical composition of the grains or drops.
The X-ray spectrometer separated the grains according to their sizes using a laser imaging device, while the mass spectrograph separated them according to their sizes using an aerodynamical inertial separator. After landing, a small surface sample near the probe was to be analyzed by gamma \Jspectroscopy\j and X-ray \Jfluorescence\j.
The UV spectrometer, the mass spectrograph, and the pressure- and temperature-measuring instruments were developed in cooperation between French and Soviet investigators. In addition to the lander probe, a constant-pressure instrumented \Jballoon\j was to be deployed immediately after entry into the atmosphere. The \Jballoon\j, with a 5-kg payload and 25-kg total mass, was to float at approximately 50 km altitude in the middle, most active layer of the Venus three-tiered cloud system.
Data from the \Jballoon\j instruments were to be transmitted directly to Earth for the 60-h lifetime of the batteries. Onboard instruments were to measure temperature, pressure, vertical wind velocity, and visibility (density of local aerosols). Very long baseline interferometry was to be used to track the motion of the \Jballoon\j to provide the wind velocity in the clouds.
The balloons were deployed at 54 km altitude and tracked for two days by an international network of antennas, including NASA's Deep Space Network (DSN) - a notable example of international space cooperation. Likewise for the Halley flyby, the Vega project supplied optical navigation data, and DSN supplied radio tracking inputs to the European Space Agency's Giotto Project. This provided the first close-up view of the \JComet\j's nucleus, and the first measurement of gas and dust properties. The \Jspacecraft\j were severely battered by the 75 km impact of Halley's dust, becoming non-operational.
\BPhobos Project\b
\B\ILaunch Date:\b\i July 7, 1988 (Phobos 1) and July 21, 1988 (Phobos 2)
\B\ILaunch Vehicle:\b\i Proton
\B\IMass:\b\i 2,600 Kg (6,220 Kg with orbital insertion hardware attached)
\B\IPower System:\b\i Solar panels
\B\IDescription\b\i
Phobos 1, and its companion \Jspacecraft\j \JPhobos\j 2, were the next-generation in the Venera-type planetary missions, succeeding those last used during the Vega 1 and 2 missions to \Jcomet\j P/Halley. The objectives of the \JPhobos\j missions were to:
(1) conduct studies of the interplanetary environment;
(2) perform observations of the Sun;
(3) characterize the plasma environment in the Martian vicinity;
(4) conduct surface and atmospheric studies of Mars; and,
(5) study the surface composition of the Martian satellite \JPhobos\j.
The main section of the \Jspacecraft\j consisted of a pressurized toroidal \Jelectronics\j section surrounding a modular cylindrical experiment section. Below these were mounted four spherical tanks containing hydrazine for attitude control and, after the main propulsion module was to be jettisoned, orbit adjustment. A total of 28 thrusters (twenty-four 50 N thrusters and four 10 N thrusters) were mounted on the spherical tanks with additional thrusters mounted on the \Jspacecraft\j body and solar panels. Attitude was maintained through the use of a three-axis control system with pointing maintained with sun and star sensors.
Phobos 1 operated normally until an expected communications session on 2 September 1988 failed to occur. The failure of controllers to regain contact with the \Jspacecraft\j was traced to an error in the software uploaded on 29/30 August which had deactivated the attitude thrusters. This resulted in a loss of lock on the Sun, resulting in the \Jspacecraft\j orienting the solar arrays away from the Sun, thus depleting the batteries.
Phobos 2 operated normally throughout its cruise and Mars orbital insertion phases, gathering data on the Sun, interplanetary medium, Mars, and \JPhobos\j. Shortly before the final phase of the mission, during which the \Jspacecraft\j was to approach within 50 m of \JPhobos\j' surface and release two landers, one a mobile 'hopper', the other a stationary platform, contact with \JPhobos\j 2 was lost. The mission ended when the \Jspacecraft\j signal failed to be successfully reacquired on 27 March 1989. The cause of the failure was determined to be a malfunction of the on-board computer.
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"Mercury Photo Gallery",167,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Pluto Photo Gallery",168,0,0,0
Click on the \BCaption\b button to read a description about this image.
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"Uranus Photo Gallery",169,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Neptune Photo Gallery",170,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Venus Photo Gallery",171,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Asteroid Photo Gallery",172,0,0,0
This page contains a series of nine images of \Jasteroids\j taken from the \JGalileo\j \Jspacecraft\j.
The \JPhobos\j and Deimos images were obtained by the Viking Orbiter \Jspacecraft\j in 1977.
The \JGalileo\j project, whose primary mission is the exploration of the Jupiter system in 1995-97, is managed for NASA's Office of Space Science and Applications by the Jet Propulsion Laboratory.
\BPictures:\b Courtesy of NASA
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"Comets",173,0,0,0
This page contains several images. Click on the \BCaption\b button to read a description for each image.
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"Earth and Moon Photo Gallery",174,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Earth Photo Gallery",175,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Jupiter Photo Gallery",176,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Mars Photo Gallery",177,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Moon Photo Gallery",178,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
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"Saturn Photo Gallery",179,0,0,0
This page has several images on it. Click on the \BCaption\b button to read a description about each image.
Planned objectives were deployment of Tracking Data Relay Satellite-2 (TDRS-2) and flying of Shuttle-Pointed Tool for \JAstronomy\j (SPARTAN-203)/Halley's \JComet\j Experiment Deployable, a free-flying module designed to observe tail and coma of Halleys \Jcomet\j with two ultraviolet spectrometers and two cameras.
Other payloads were Fluid Dynamics Experiment (FDE); \JComet\j Halley Active Monitoring Program CHAMP); Phase Partitioning Experiment (PPE); three Shuttle Student Involvement Program (SSIP) experiments; and set of lessons for Teacher in Space Project (TISP).
\BLaunch:\b
January 28, 1986,11:38:00 a.m. EST. First Shuttle liftoff scheduled from Pad B. Launch set for 3:43 p.m. EST, Jan. 22, slipped to Jan. 23, then Jan. 24, due to delays in mission 61-C. Launch reset for Jan. 25 because of bad weather at transoceanic abort landing (TAL) site in \JDakar\j, \JSenegal\j.
To utilize \JCasablanca\j (not equipped for night landings) as alternate TAL site, T-zero moved to morning liftoff time. Launch postponed a day when launch processing unable to meet new morning liftoff time. Prediction of unacceptable weather at KSC led to launch rescheduled for 9:37 a.m. EST, Jan. 27. Launch delayed 24 hours again when ground servicing equipment hatch closing fixture could not be removed from orbiter hatch.
Fixture sawed off and attaching bolt drilled out before closeout completed. During delay, cross winds exceeded return-to-launch-site limits at KSC's Shuttle Landing Facility. Launch Jan. 28 delayed two hours when hardware interface module in launch processing system, which monitors fire detection system, failed during liquid \Jhydrogen\j tanking procedures.
Just after liftoff at .678 seconds into the flight, photographic data show a strong puff of gray smoke was spurting from the vicinity of the aft field joint on the right Solid Rocket Booster. Computer graphic analysis of film from pad cameras indicated the initial smoke came from the 270 to 310-degree sector of the \Jcircumference\j of the aft field joint of the right Solid Rocket Booster.
This area of the solid booster faces the External Tank. The vaporized material streaming from the joint indicated there was not complete sealing action within the joint.
Eight more distinctive puffs of increasingly blacker smoke were recorded between .836 and 2.500 seconds. The smoke appeared to puff upwards from the joint. While each smoke puff was being left behind by the upward flight of the Shuttle, the next fresh puff could be seen near the level of the joint.
The multiple smoke puffs in this sequence occurred at about four times per second, approximating the frequency of the structural load dynamics and resultant joint flexing. As the Shuttle increased its upward velocity, it flew past the emerging and expanding smoke puffs. The last smoke was seen above the field joint at 2.733 seconds.
The black color and dense composition of the smoke puffs suggest that the grease, joint \Jinsulation\j and rubber O-rings in the joint seal were being burned and eroded by the hot propellant gases.
At approximately 37 seconds, Challenger encountered the first of several high-altitude wind shear conditions, which lasted until about 64 seconds. The wind shear created forces on the vehicle with relatively large fluctuations. These were immediately sensed and countered by the guidance, navigation and control system.
The steering system (thrust vector control) of the Solid Rocket Booster responded to all commands and wind shear effects. The wind shear caused the steering system to be more active than on any previous flight.
Both the Shuttle main engines and the solid rockets operated at reduced thrust approaching and passing through the area of maximum dynamic pressure of 720 pounds per square foot. Main engines had been throttled up to 104 percent thrust and the Solid Rocket Boosters were increasing their thrust when the first flickering flame appeared on the right Solid Rocket Booster in the area of the aft field joint.
This first very small flame was detected on image enhanced film at 58.788 seconds into the flight. It appeared to originate at about 305 degrees around the booster \Jcircumference\j at or near the aft field joint.
One film frame later from the same camera, the flame was visible without image enhancement. It grew into a continuous, well-defined plume at 59.262 seconds. At about the same time (60 seconds), telemetry showed a pressure differential between the chamber pressures in the right and left boosters. The right booster chamber pressure was lower, confirming the growing leak in the area of the field joint.
As the flame plume increased in size, it was deflected rearward by the aerodynamic slipstream and circumferentially by the protruding structure of the upper ring attaching the booster to the External Tank. These deflections directed the flame plume onto the surface of the External Tank. This sequence of flame spreading is confirmed by analysis of the recovered wreckage. The growing flame also impinged on the strut attaching the Solid Rocket Booster to the External Tank.
The first visual indication that swirling flame from the right Solid Rocket Booster breached the External Tank was at 64.660 seconds when there was an abrupt change in the shape and color of the plume. This indicated that it was mixing with leaking \Jhydrogen\j from the External Tank.
Telemetered changes in the \Jhydrogen\j tank pressurization confirmed the leak. Within 45 milliseconds of the breach of the External Tank, a bright sustained glow developed on the black-tiled underside of the Challenger between it and the External Tank.
Beginning at about 72 seconds, a series of events occurred extremely rapidly that terminated the flight. Telemetered data indicate a wide variety of flight system actions that support the visual evidence of the photos as the Shuttle struggled futilely against the forces that were destroying it.
At about 72.20 seconds the lower strut linking the Solid Rocket Booster and the External Tank was severed or pulled away from the weakened \Jhydrogen\j tank permitting the right Solid Rocket Booster to rotate around the upper attachment strut. This rotation is indicated by divergent yaw and pitch rates between the left and right Solid Rocket Boosters.
At 73.124 seconds,. a circumferential white vapor pattern was observed blooming from the side of the External Tank bottom dome. This was the beginning of the structural failure of \Jhydrogen\j tank that culminated in the entire aft dome dropping away.
This released massive amounts of liquid \Jhydrogen\j from the tank and created a sudden forward thrust of about 2.8 million pounds, pushing the \Jhydrogen\j tank upward into the intertank structure. At about the same time, the rotating right Solid Rocket Booster impacted the intertank structure and the lower part of the liquid oxygen tank. These structures failed at 73.137 seconds as evidenced by the white vapors appearing in the intertank region.
Within milliseconds there was massive, almost explosive, burning of the \Jhydrogen\j streaming from the failed tank bottom and liquid oxygen breach in the area of the intertank.
At this point in its trajectory, while traveling at a Mach number of 1.92 at an altitude of 46,000 feet, the Challenger was totally enveloped in the explosive burn. The Challenger's reaction control system ruptured and a hypergolic burn of its propellants occurred as it exited the oxygen-hydrogen flames. The reddish brown colors of the hypergolic fuel burn are visible on the edge of the main fireball.
The Orbiter, under severe aerodynamic loads, broke into several large sections which emerged from the fireball. Separate sections that can be identified on film include the main engine/tail section with the engines still burning, one wing of the Orbiter, and the forward fuselage trailing a mass of umbilical lines pulled loose from the payload bay.
The Explosion 73 seconds after liftoff claimed crew and vehicle. Cause of explosion was determined to be an O-ring failure in right SRB. Cold weather was a contributing factor. Launch Weight: 268,829 lbs.
Orbit:
Altitude: 150nm (planned)
Inclination: 28.5 degrees (planned)
Orbits: 0
Duration: 01 min 13 seconds
Distance: 18 miles
Hardware:
SRB: BI-026
SRM: L025(HPM)
ET : 26/LWT-19
MLP : 2
SSME-1: SN-2023
SSME-2: SN-2020
SSME-3: SN-2021
Landing:
None. KSC Landing planned after a 6 day, 34 minute mission.
Mission Highlights:
The planned orbital activities of the Challenger 51-L mission were as follows:
On Flight Day 1, after arriving into orbit, the crew was to have two periods of scheduled high activity. First they were to check the readiness of the TDRS-B satellite prior to planned deployment. After lunch they were to deploy the satellite and its Inertial Upper Stage (IUS) booster and to perform a series of separation maneuvers. The first sleep period was scheduled to be eight hours long starting about 18 hours after crew wake up the morning of launch.
On Flight Day 2, the \JComet\j Halley Active Monitoring Program (CHAMP) experiment was scheduled to begin. Also scheduled were the initial "teacher in space" (TISP) video taping and a firing of the orbital maneuvering engines (OMS) to place Challenger at the 152-mile orbital altitude from which the Spartan would be deployed.
On Flight Day 3, the crew was to begin pre-deployment preparations on the Spartan and then the satellite was to be deployed using the remote manipulator system (RMS) robot arm. Then the flight crew was to slowly separate from Spartan by 90 miles.
On Flight Day 4, the Challenger was to begin closing on Spartan while Gregory B. Jarvis continued fluid dynamics experiments started on day two and day 3. Live telecasts were also planned to be conducted by Christa McAuliffe.
On Flight Day 5, the crew was to rendezvous with Spartan and use the robot arm to capture the satellite and re-stow it in the payload bay.
On Flight Day 6, re-entry preparations were scheduled. This included flight control checks, test firing of maneuvering jets needed for reentry, and cabin stowage. A crew news conferences was also scheduled following the lunch period.
On Flight Day 7, the day would have been spent preparing the Space Shuttle for deorbit and entry into the atmosphere. The Challenger was scheduled to land at the Kennedy Space Center 144 hours and 34 minutes after launch.
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"Space Shuttle, Living in a",181,0,0,0
\JSpace Shuttle Living\j
\JShuttle, The Air Inside\j
\JShuttle Meals\j
\JShuttle Sanitation\j
\JSpace Suit, Unisex\j
\JShuttle Recreation and Sleep\j
\JShuttle Weightlessness\j
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"Space Shuttle Living",182,0,0,0
The idea that ordinary people would someday live and work in space has fascinated science fiction fans as well as serious scientists and engineers. NASA's Space Shuttle is the first step in turning this dream into reality.
The Space Shuttle is a reusable aerospace vehicle that takes off like a rocket, can be maneuvered in space, and lands like an airplane.
The \Jspacecraft\j, called the orbiter, is about the size of a DC-9 commercial jetliner. The orbiter carries people and cargo between the ground and Earth orbit. It can also be used as an observation post in space and as a space platform for a fully equipped laboratory for medical, scientific, \Jengineering\j, and industrial experiments.
One of the key attributes of the Shuttle and its operation is the relatively low g-force exerted on crew and passengers during launch and reentry. Launch and reentry forces are less than 4 g's -- well within the limits which can be tolerated by healthy people.
Orbiter living accommodations are relatively comfortable. They incorporate advances made through nearly 2 decades of experimental manned space missions and an even longer period of ground studies.
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"Shuttle, The Air Inside",183,0,0,0
The orbiter's air is cleaner than Earth's, and hay fever sufferers will welcome its pollen-free atmosphere.
Orbiter air pressure is the same as Earth's at sea level: 1,033 grams per square centimeter (14.7 pounds per square foot). Its air is made up of 80 percent \Jnitrogen\j and 20 percent other gases such as \Jargon\j and neon. The orbiter's environmental control system circulates air through filters to remove carbon dioxide and other impurities. Excess moisture is also removed, keeping \Jhumidity\j at comfortable levels. Temperature in the orbiter can be regulated between 16 and 32 degrees \JCelsius\j (61 and 90 degrees Fahrenheit). The orbiter crew requires only ordinary clothing. People can move about, work, and relax unencumbered by bulky space suits.
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"Shuttle Meals",184,0,0,0
Shuttle meals are tasty and nutritious. They can be eaten anywhere, although crew members normally congregate in the middeck area for their meals. Trays holding the food can be attached to a crew member's legs or to any orbiter surface with adhesive straps, removing the need for a table and chairs at mealtime. Meals are served in a special tray which separates the different food containers and keeps them from lifting off and soaring around in the weightless cabin.
Packages of food that have to be warmed are placed in the galley oven before going into the tray. Hot and cold water are available for preparation of foods or beverages.
Studies have shown that despite zero gravity, most foods can be eaten with ordinary spoons and forks as long as there are no sudden starts, stops, or spinning. As a result, dining in space is almost like dining on Earth.
The orbiter menu includes more than 70 food items and 20 beverages. With so many different items, Shuttle travelers can have varied menus every day for 6 days.
Earth-bound chefs might envy orbiter meal preparation -- one crewmember can ready meals for four people in about 5 minutes.
What are orbiter meals like? A typical day's menus include orange drink, peaches, scrambled eggs, sausage, \Jcocoa\j, and a sweet roll for breakfast; cream of mushroom soup, ham and cheese sandwich, stewed tomatoes, banana, and cookies for lunch; and shrimp cocktail, beefsteak, broccoli au gratin, strawberries, pudding, and \Jcocoa\j for dinner.
Menus provide about 2,700 calories daily. Previous space missions demonstrated that astronauts need at least as many calories in space as they do on Earth.
The orbiter does not have a refrigerator. Most of the Shuttle foods are preserved by \Jdehydration\j, which saves weight and storage space. Water for rehydration is ample since it is a byproduct of the fuel cells which generate electricity. Some foods are thermostabilized, that is, they are heat sterilized and then sealed in conventional cans or plastic pouches. A few, such as cookies and nuts, are available in ready-to-eat form.
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"Shuttle Sanitation",185,0,0,0
Eating utensils are cleaned with wet wipes. The difference between orbiter wet wipes and those used on Earth is that the orbiter's contain a strong \Jdisinfectant\j.
Sanitation is more important in the confines of the orbiter than on Earth. Space studies have shown the population of some microbes can increase extraordinarily in a confined weightless area such as a \Jspacecraft\j cabin. This could potentially spread illness to everyone on board. As a result, not only eating components but also the dining area, the toilet, and sleeping areas are regularly cleaned. Since there are no washing machines in space, trousers (changed weekly), socks, shirts, and underwear (changed every 2 days) are sealed in airtight plastic bags after being worn. Garbage and trash also are sealed in plastic bags.
A favorite question of people interested in space is how the astronauts took care of digestive elimination. The orbiter travelers use a toilet very much the same as one on Earth. Air flow directs waste to the bottom of the toilet, substituting for gravity. Waste goes directly into a sealed container where it is processed and stored.
Some of the waste may be used for post-flight laboratory analyses. Such analyses have told doctors which minerals are lost excessively in space and have helped to increase their understanding of body functions.
Orbiter travelers have facilities and supplies available for sponge baths while in space. They can obtain water from the water dispensing system. Water temperature can be set at any comfortable level from 18 to 35 degrees \JCelsius\j (65 to 95 degrees Fahrenheit).
Because of weightlessness, water droplets would float about in the cabin. This could be not only a nuisance but also potentially hazardous to equipment and crew. To prevent this from happening, an airflow system directs waste water into the orbiter's waste collection system, where the waste water is sealed in plastic watertight bags.
Whiskers cut off in shaving and floating about weightlessly in a cabin could be a nuisance and foul up equipment. This problem is avoided by using conventional shaving cream and a safety razor and cleaning off the face with a disposable towel. Also available is a wind-up shaver that works like an electric razor and contains a vacuum device to prevent the escape of cut whiskers.
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"Space Suit, Unisex",186,0,0,0
In the past, space suits were tailor-made for each \Jastronaut\j, a time-consuming and costly process. The Shuttle space suit is manufactured in small, medium, and large sizes and can be worn by men or women. The suit comes with an upper and lower torso equivalent to a shirt and trousers. Each piece snaps together with sealing rings. A life-support system is built into the upper torso. Previous pressure suits had separate life support systems which had to be connected to the suits.
The Shuttle space suit is lighter, more durable, and easier to move about in than previous space suits. When an \Jastronaut\j has to work outside the space- craft, the Shuttle suit is used for extravehicular activity.
#
"Shuttle Recreation and Sleep",187,0,0,0
Just as on Earth, recreation and sleep are important to good health in space. A scientifically planned exercise program is provided, largely as a countermeasure for cardiovascular deconditioning and atrophy of muscles in a weightless environment. Cards and other games, books, writing material, and tape recorders and tapes to chronicle personal observations or to listen to music, are available.
#
"Shuttle Weightlessness",188,0,0,0
Many of the problems of going into space have been resolved. However, the physiological effects of weightlessness are still not completely understood. Among them are leaching of minerals from bones, reduction in rate of bone formation, atrophy of muscles when not exercised, and motion sickness.
All of the effects of zero gravity have so far been reversed after return to the normal gravity on Earth. In addition, some of the effects have been countered by exercise and food supplements.
However, even vigorous exercise in space does not appear to stop bone loss or decrease in the rate of bone formation. As a result, NASA is engaged in an intense and sustained effort aimed at understanding the causes underlying these changes and developing ways to prevent them. The increased information about body functions derived from this effort will pave the way for prolonged missions in space and contribute to our understanding of the \Jphysiology\j of living things on Earth.
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"Shuttle Location",189,0,0,0
\JShuttle Location Over Earth\j
\JShuttle Position in the Universe\j
#
"Shuttle Location Over Earth",190,0,0,0
When a space shuttle crew wants to know where it is over the Earth, the easiest way to find out is to look out a window. Sometimes, when the shuttle is orbiting over the middle of the Pacific Ocean, or when its windows are facing away from the Earth, this can be a problem.
To help the astronauts with their situational awareness and to provide them with information for documenting their Earth observations with cameras, the flight control team provides them with an application called WORLDMAP which runs on a Payload and General Support Computer (PGSC). These commercially available laptops (at this time, they're using an \JIBM\j Thinkpad 755) are used to run various applications on-board. WORLDMAP is used to display the orbiter's location and to display the location of various Earth observation sites. WORLDMAP uses a state vector to determine its position over the Earth.
The image shown is a snapshot of the same program--developed by the SpOC (Space Operations Computing) team programmers at the Johnson Space Center and running in Microsoft Windows--that the crew on board is using. The image may not be an exact copy of the one used by the astronauts, since they can change the program's preferences, but they could duplicate this view if they so desired.
Both the flight version and the ground version of the program access the orbiter state vector information via the shuttle's pulse code master \Jmodulation\j unit (PCMMU, pronounced "puckamoo"). The PCMMU data stream is read into another commercially available program called PCDeCOM, which extracts specified data parameters and feeds them to the WORLDMAP program.
The small window in the lower left portion of the image shows the overall world view. The largest part of the image shows the Earth below from the same vantage, but magnified eight times. The application's title bar includes the current Greenwich Mean Time (GMT), Mission Elapsed Time (MET), latitude, longitude and altitude.
The map title bar displays the name of the country the shuttle is flying over. The timer window provides a variety of information, including the next S-band and Ku-band acquisition and loss of signal through the Tracking and Data Relay Satellite System, and sunrise and sunset times (remember that the sun rises or sets every 45 minutes on orbit).
Sometimes, areas will be outlined in red and red numerals will be displayed in the large map. These are upcoming Earth observations photography opportunities. The crew on board can click on these red areas and get additional information about the site and the best lenses and film speeds to use when photographing the area.
Flight controllers in the Mission Control Center use a different display called the Distributed Earth Modeling and Orbiter System (DEMOS) to help them visualize the location and attitude of the shuttle.
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"Shuttle Position in the Universe",191,0,0,0
The shuttle's ground track over the Earth's surface is shown on the Distributed Earth Model and Orbiter System (DEMOS).
Flight controllers in NASA's Mission Control Center use this DEMOS computer modeling software to help them visualize exactly where the space shuttle is over the Earth and what attitude it is in (how is it facing in three-dimensional perspective to the Earth). Crewmembers aboard the shuttle can access the same information using a laptop computer and WORLDMAP software.
These computer-generated images are translated into \Jtelevision\j signals that are displayed in the Flight Control Room, and frequently are broadcast on NASA \JTelevision\j. The NASA Shuttle Web uses a frame-grabber to produce still images of the DEMOS display every few minutes. If you are using Netscape as your browser, the image will be updated automatically. If not, you must reload to update the image. Image quality will vary depending on the quality of the \Jtelevision\j signal at the the frame was grabbed.
DEMOS uses the same state vectors used by NASA's Mission Control Center to keep track of the Shuttle. The state vector defines very exactly the position of the \Jspacecraft\j. It contains orbital elements, also called Keplerian elements, such as the tilt of the orbit plane, altitude, and how circular (round) the orbit is.
The Mission Control Center in Houston regularly updates the state vector on the shuttle's General Purpose Computers so that the shuttle's navigation systems will know exactly where the \Jspacecraft\j is and where it will be in the future.
You can use these same orbital elements to track \Jspacecraft\j on your home computer system. There are a number of freeware and shareware tracking programs that will allow you to input the orbital elements and follow a \Jspacecraft\j (or a planet or a satellite for that matter) from home. To use these programs, you will need both the software and the latest orbital elements.
You must update these elements regularly (usually about once every 12 hours, but more often when the \Jspacecraft\j is maneuvering). Use these software programs at your own risk; NASA takes no responsibility for their effect on the stability of your computer system.
The Launch Control Center (LCC) is a four-story building that is the electronic "brain" of Launch Complex 39. Attached to the southeast corner of the Vehicle Assembly Building (VAB) , it is 5,535 meters (18,159 ft) from Pad 39A . At the time it was constructed, advances in \Jelectronics\j had made it unnecessary to continue locating blockhouses adjacent to launch pads.
The first floor contains offices and computer operations. The second floor houses telemetry, RF and tracking, instrumentation, and data reduction and evaluation equipment. The computers of the Central Data Subsystem (CDS), one of the two major components of the Launch Processing System (LPS) that automatically performs most checkout and launch functions, is also on the second floor.
The third floor contains the four firing rooms, and each room contains its own copy of the second major component of the Launch Processing System - the Checkout, Control and Monitor Subsystem (CCMS).
The system consoles of the CCMS system are manned by the team which oversees all aspects of a checkout and launch operation. Firing rooms 1 and 3 are configured for full control of launch and orbiter operations while Firing room 2 is usually used for software development and testing. Firing room 4 is only a partial firing room and is primarily used as an \Jengineering\j analysis and support area for launch and checkout operations.
The fourth floor of the LCC contains conference rooms, offices and mechanical equipment. The LCC is 23.5 meters (77ft) high, 115.2 meters (378ft) long and 55.1 meters (181 ft) wide.
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"Mobile Launch Platforms (MLP)",194,0,0,0
The three Mobile Launchers used in Apollo/Saturn operations were modified for use in Shuttle operations. With cranes, umbilical towers, and swing arms removed, the Mobile Launchers were redesignated Mobile Launcher Platforms (MLP). In place of one large opening in the platform, three smaller openings accommodate flames and hot exhaust gases from the solid rocket boosters and the orbiter engines.
Segments of the dismantled umbilical towers are part of the permanent installation at the launch pad, where they serve as sections of the Fixed Service Structure (FSS). A third Apollo umbilical tower, removed from MLP-3, has been cut into 20 ft sections and placed in a field in the KSC industrial area. It may someday become reconstructed as part of the KSC tour route.
Tail Service Masts (TSM's), one on each side of the main engines exhausts hole provide umbilical connections for fuel and oxidizer, gases, ground electrical power and communications links. These Masts are 4.6m (15ft) long, 2.7m (9ft) wide and 9.4m (31 ft) high.
MLP Statistics
Launch Platform: Two-Story steel structure, 7.6 meters (25ft) high, 49 Meters (160ft) long and 41 meters (135ft)wide.
ò Empty Weight: 4.19 million kg (9.25 million lb)
ò With unfueled Shuttle: 5.45 million kg (12.02 million lb)
ò With fueled Shuttle: 6.22 million kg (13.72 million lb)
Positioned on 6 steel pedestals 7m (22ft) high when in the VAB or at the launch pad. At the pad, 4 extensible columns were used during the Apollo program were used to stiffen the MLP against rebound loads, should engine cutoff occur. They are no longer used for the Shuttle program.
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"Operations and Checkout Building",195,0,0,0
Horizontally integrated payloads are received, assembled and integrated in the Operations and Checkout Building before they are mated with the orbiter at the Orbiter Processing Facility. The Spacelab and its payloads constitute a large majority of the horizontal payloads.
The Operations and Checkout Building is a five-story structure containing 600,000 square feet of offices, laboratories, \Jastronaut\j crew quarters and \Jspacecraft\j assembly areas. It is located in the industrial area immediately east of the KSC headquarters building.
The \JSpacecraft\j Assembly and Encapsulation Facility -2 (SAEF-2) is located at F Avenue and 7th Street in the Hypergol Maintenance Facility Area, KSC Industrial Area. The facility is used for the assembly, test, encapsulation, ordnance work, propellant loading, and pressurization of \Jspacecraft\j. The facility contains approximately 1556.08 meters (m)2 (16,750 feet (ft)2) of usable floor space. Construction is of reinforced concrete and concrete block. The high bay is a steel frame structure with insulated aluminum siding.
Functionally, the building is divided into the following areas: a clean work area (CWA) complex consisting of an airlock, a high bay, and two low bays; a test cell; a sterilization oven (non- operational); support office areas; and mechanical equipment rooms (figure 2-3).
Floors in the airlock, high bay, and test cell are designed for 3175.20 \Jkilogram\j (kg) (7000 pounds (lb)) per wheel plus 20 percent impact loading. The airlock, located at the north end of the building, measures 12.5 m by 17.7 m (41 ft wide by 58 ft long), providing a usable floor area of 221 m\U2\u (2378 ft\U2\u) and is rated as a Class 300,000 CWA. The airlock has a clear ceiling height of 15.9 m (52 ft).
Access is by means of personnel doors, vestibule, and a 6.4 m by 12.2 m (21.5 ft wide by 40 ft high) vertical lift equipment door. A 6.5 m by 12.1 m (21 ft wide by 39.5 ft high) horizontal sliding lift door separates the airlock from the adjacent high bay.
The high bay measures 14.9 m by 30.2 m (49 ft wide by 99 ft long), providing a usable floor area of 450.7 m2 (4851 ft\U2\u); clear ceiling height is 22.6 m (74 ft) and is rated as a Class 100,000 CWA. Personnel and small equipment can enter the high bay through the equipment airlock, equipped with air showers. Clear access is 1.4 m by 2.1 m (4 ft 5 in wide by 6 ft 11 in high).
Two large bays are located along the west side of the high bay. One of the bays measures 5.8 m by 21.9 m (19 ft wide by 72 ft long) with a clear ceiling height of 7.62 m (25 ft); the other bay is 5.8 m by 8.2 m (19 ft wide by 27 ft long) and has a clear ceiling height of 13.3 m (43.5 ft). The combined bay areas provide a usable floor space of 174.8 m2 (1881 ft2). The low bays are also rated as Class 100,000 CWA's.
The test cell is located at the northeast corner of the facility. The cell measures 11.3 m by 11.3 m (37 ft by 37 ft), providing a usable floor area of 127.2 m2 (1369 ft2). The clear ceiling height is 15.9 m (52 ft). Access to the test cell is by means of personnel doors and three 6.7 m by 12.2 m (22 ft wide by 40 ft high) vertical lift doors. The test cell can be used for \Jspacecraft\j and payload support activities not requiring Class 100,000 CWA conditions.
The remainder of the building consists of such support rooms as a mechanical equipment room, communication equipment room, entry and observation room, change room, storage room, miscellaneous equipment rooms, and limited office space.
The sterilization oven is located at the south end of the facility in a 13.1 m by 16.2 m (43 ft wide by 53 ft long) open- sided shed. Ceiling height of the shed in the oven area is 4.9 m (16 ft). (The oven is not operational.)
There is a 13.7 m by 8.5 m (45 ft long by 28 ft wide) conference area located on the second floor above rooms 117 and 119. Room 109 is used as the clean room garment storage and issue room. There are five 3.7 m by 18.3 m (12 ft by 60 ft) office trailers, figure 2-3, available for payload personnel use when their payload is in SAEF-2. Two are on the north side of the building; two, on the west side; and one, on the east side. One of the trailers located on the north side is used by the NASA and PGOC facility managers.
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"Space Station Processing Facility (SSPF)",197,0,0,0
The SSPF is located in the KSC industrial area, just east of the Operations and Checkout Building. It was built for the processing of the International Space Station flight hardware. The three-story SSPF, a 457,000 square foot building, includes two processing bays, an airlock, operational control rooms, laboratories, logistics areas, office space, and a cafeteria. The processing areas, airlock, and laboratories were designed to support non-hazardous Station and Shuttle payloads in 100,000 class clean work areas.
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"Vehicle Assembly Building",198,0,0,0
The Vehicle Assembly Building (VAB) is one of the largest buildings in the world. It was originally built for assembly of Apollo/Saturn vehicles and was later modified to support Space Shuttle operations. High Bays 1 and 3 are used for \Jintegration\j and stacking of the complete Space Shuttle vehicle.
High Bay 2 is used for external tank (ET) checkout and storage and as a contingency storage area for orbiters. High Bay 4 is also used for ET checkout and storage, as well as for payload canister operations and solid rocket booster (SRB) contingency handling.
The Low Bay area contains Space Shuttle main engine maintenance and overhaul shops, and serves as a holding area for SRB forward assemblies and aft skirts.
During Space shuttle build-up operations inside the VAB, integrated SRB segments are transfered from nearby SRB assembly and checkout facilities, hoisted onto a Mobile Launcher Platform in High Bays 1 or 3 and mated together to form two complete SRBs. The ET, after arrival by barge, is inspected and checked out in High Bays 2 or 4 and then transfered to High Bay's 1 or 3 to be attached to the SRBs already in place.
The orbiter is then towed over from the Orbiter Processing Facility to the VAB transfer aisle, raised to a vertical position, lowered onto the Mobile Launcher Platform and then mated to the rest of the stack. When assembly and checkout is complete, the crawler-transporter enters the High Bay, picks up the platform and assembled shuttle vehicle and carries them to the launch pad.
The VAB covers 3.25 hectares (8 acres). It is 160 meters (525 ft) tall, 218 meters (716 ft) long and 158 meters (518 ft) wide. It encloses 3,664,883 cubic meters (129,428,000 cubic feet) of space.
ò Flag & Bicentennial Emblem: Added in 1976, required 6,000 gallons of paint. The flag is 64 x 33.5 meters (209 x 110 ft) in size. Each strip on the flag is as big as the tour buses used to transport visitors around KSC
ò Steel: 89,421 metric tons (98,590 tons)
ò Concrete: 49,696 cubic meters (65,000 cubic yards)
ò Piling: 4,225 open-end steel pipe piles, 0.4 meters (16 inches) in diameter were driven 49 meters (160 ft) into bedrock.
ò Air Conditioning: 9,070 metric tons (10,000 tons), 125 ventilators.
ò Lifting Devices: 71 cranes; two 227 metric ton (250 ton) bridge cranes.
ò Doors: There are 4 High Bay doors. Each opening is 139 meters (456 ft) high. The north entry to the transfer aisle was widened 12.2 meters (40ft) to permit entry of the Orbiter, and slotted at the center to accommodate its vertical stabilizer.
Comparisons:
ò Height: VAB - 160meters (525 ft) Statue of Liberty - 93 meters (305 ft)
ò Volume: VAB - 3,665,013 cu meters (129,428,000 cub ft) Pentagon 2,181,117 cu meters (77,025,000 cu ft).
ò VAB equals 3.75 Empire State Buildings
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"Shuttle Flights To Date",199,0,0,0
(Refer to Table)
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"Space Shuttle Launch Team",200,0,0,0
\JShuttle Launch Team\j
\JShuttle Launch Team Structure\j
\JCountdown To Launch\j
\JShuttle Operation Communication\j
\JShuttle Firing Room Protocol\j
\JLaunch Pad Activities During Countdown\j
\JLaunch: Final Go\j
#
"Shuttle Launch Team",201,0,0,0
Launch Complex 39 at the Kennedy Space Center has served as the springboard for U.S. human spaceflight since the Apollo lunar landing program in the 1960s. The first launch conducted from one of the two LC 39 pads was Apollo 4 on Nov. 9, 1967. Since then, teams staffing the consoles in one of the Launch Control Center firing rooms have sent into space astronauts who walked on the moon, interplanetary explorer \Jspacecraft\j destined for the far reaches of the solar system and Space Shuttle crews to conduct research and to service, deploy and retrieve \Jspacecraft\j.
The first Space Shuttle launch was conducted from Pad 39A on April 12, 1981. Today, Shuttles lift off regularly from either Pad 39A or B under the management of the launch team in the Launch Control Center.
The Space Shuttle launch team is a highly organized and disciplined group of approximately 500 professionals. Membership of the launch team reflects the complexity involved in preparing for and conducting a human space launch. Civil service and contractor personnel from Kennedy occupy central roles, but other NASA centers, contractor personnel and agencies also contribute.
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"Shuttle Launch Team Structure",202,0,0,0
The Shuttle launch team is organized into three groups, according to major functional responsibility. All three groups report to the Shuttle Launch Director, who has overall technical and safety responsibility for the countdown.
1. The prime launch team is responsible for test, checkout and monitoring of the flight hardware and ground support equipment to ensure that all system parameters meet the criteria to commit the vehicle to launch. This team includes the group stationed in the prime firing room in the Launch Control Center, as well as offsite personnel with critical launch support responsibilities. The prime launch team is headed by the NASA Test Director (NTD), who reports to the Shuttle Launch Director.
Approximately 200 of the 300 members of this team are stationed in the prime firing room for the mission at hand. The NTD and other managers overseeing the countdown process are stationed at consoles facing into the firing room. Reporting to the NTD are 11 test conductors, each responsible for a specific subset of requirements to be met prior to launch.
For example, the Orbiter Test Conductor (OTC) is the individual in charge of the system-level engineers monitoring the hardware and software on board the orbiter itself. Other test conductors are responsible for the external tank and solid rocket boosters, the payloads, support operations, landing operations, safety, communications and other functions.
Also stationed in the prime firing room is the Shuttle Project Engineer, who is the lead engineer and reviews all technical issues for the NTD.
The \Jastronaut\j assigned to command the Shuttle mission represents his flight crew as he communicates with the NTD from inside the vehicle at the pad.
Seated at seven console sets facing the windows of the prime firing room and looking toward the two launch pads are the teams of system-level experts who report to the test conductors. They represent all orbiter, external tank and solid rocket booster components. Mission-specific payload experts are assigned to this team also.
Also part of this configuration is an eighth console called the \Jintegration\j console, which includes the Ground Launch Sequencer that automatically controls the final nine minutes of the countdown.
At the rear of the room is the Master Console, which monitors the Launch Processing System computer network with which all Shuttle processing, checkout and launch are performed.
Each console is arranged in its own semicircle and includes up to a dozen individual operator stations. The number of stations dedicated to a major function, such as payloads, varies. Besides the software giving the operator control over a particular system, each console has radio hookups and video displays monitoring Shuttle and pad hardware. Click here for photo of a typical console.
A console chief is assigned to each console. This typically is a senior system engineer. One of the console chief's functions is to communicate with personnel in an adjacent backup firing room so the expertise of both rooms is combined.
The system engineers at the seven console sets have their own method of assigning the various responsibilities that encompass their particular set of requirements. For example, the lead engineer for the liquid \Jhydrogen\j system has approximately eight engineers reporting to him or her, each of whom are responsible for a portion of the entire system. This includes the storage tank at the pad, pressurization systems, the cross-country pipelines used to load the external tank, the \Jhydrogen\j portion of the tank itself, and so on. In turn, each of these elements has its own set of responsibilities and requirements to monitor.
Seated behind the console operators are additional engineers and technicians. They are available to provide technical expertise or to help troubleshoot any glitches that may occur during the countdown.
Approximately 100 members of the primary launch team are located outside the confines of the Launch Control Center, yet oversee systems and hardware critical to the launch process. These include representatives from Johnson Space Center, which manages the Space Shuttle program and is responsible for on-orbit operations. U.S. and overseas contingency landing site readiness is the JSC team's job also. The Eastern Range, managed by the U.S. Air Force, is responsible for ensuring range safety as well as providing weather information. Located nearby is the Merritt Island Launch Area (MILA) Station, which provides voice and data transmission paths to and from the vehicle. Goddard Space Flight Center manages the MILA station.
2. The \Jengineering\j support team has a similar composition and organization to that of the launch team, but is not directly responsible for system management. Acting in more of an oversight capacity, they provide technical support should problems arise. This extra set of eyes is composed of extremely experienced engineers who can assist in solving problems in real time. These engineers are stationed primarily in the backup firing room, Firing Room 2, located between Firing Rooms 1 and 3 on the LCC third floor. They work at computer hardware and software similar to that of the prime team, but perform systems-monitoring only, with no command capability or responsibility.
Because of space limitations, an additional room on the third floor also is used, called the \JEngineering\j Support Area (ESA). The \Jengineering\j support team also includes personnel outside KSC, such as other centers and contractor facilities. All represent a pool of expertise from which the prime team can draw at a moment's notice if need be.
3. The senior government and contractor managers that comprise the Mission Management Team (MMT) are charged with reporting any issues that may affect the safety or success of the countdown or mission. Reportable issues can originate during any phase of the preflight hardware component processing as well as during the countdown itself. All issues raised must be resolved prior to clearing the launch vehicle for flight. During the countdown, the MMT is located in the Operations Support Room area of the prime firing room.
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"Countdown To Launch",203,0,0,0
The launch of the Space Shuttle marks the finale of many thousands of individual tasks performed by highly trained and motivated workers at the Kennedy Space Center and elsewhere. Approximately four months of system tests, refurbishment and unique configurations for the upcoming mission are required to prepare the Shuttle for its next flight.
The countdown formally begins with the call to stations, issued by the NTD from the Launch Control Center. The countdown clock begins ticking at T-43 hours about three days before liftoff. With built-in hold time included, it takes roughly 72 hours to conduct a Shuttle launch countdown.
The reference tool for conducting a Shuttle launch countdown is a five-volume manual encompassing some 5,000 pages of instructions. More commonly known by its identification number, S0007, the Shuttle countdown manual is the lengthiest Operations and Maintenance Instructions (OMIs) used at KSC to document procedures for assembly, processing, testing and launching of the Shuttle. S0007 is reviewed and updated prior to each mission.
The first volume of S0007 contains all the preparations necessary to lead up to the beginning of the three-day countdown. Volume 2 is the actual set of countdown instructions that logically and sequentially configure the vehicle for launch. It is this integrated set of instructions that contains all the requirements necessary to launch.
Volume 3 contains all the instructions to follow in the event a launch is scrubbed. In this case, the vehicle and facilities are recycled for a subsequent launch attempt. This next attempt could be in as little as 24 hours or could be several days later. Volume 4 contains all the specific individual system instructions that are initiated from Volume 2. Volume 5 is a set of preplanned contingency procedures and emergency instructions available in the unlikely event they are required.
This entire set of instructions is performed under the direction of the NTD.
The beginning of a countdown is not unlike a routine power-up of the orbiter. System checks are conducted and the vehicle configured for later operations. Once preparations get under way to load the orbiter's fuel cell power reaction and storage distribution system at around T-28 hours, the spaceship's configuration is getting more oriented toward flight. The remaining countdown milestones focus on pad and vehicle closeouts: closure of the payload bay doors; retraction of the Rotating Service Structure at the launch pad; installation of the crew escape pole; activation of the onboard fuel cells; loading of the external tank, and boarding of the flight crew.
After the countdown clock starts ticking, the prime firing room is staffed around the clock. Personnel are typically selected for this prime launch team at about the time an orbiter is being transferred from the Orbiter Processing Facility to the Vehicle Assembly Building for \Jintegration\j with the other Shuttle flight elements.
As subsequent countdown milestones are met, the composition of the prime firing room team will change. Assigned shifts of teams will report on station at varying times, depending on which system they oversee. For example, a loading crew comes on station at T-9 hours to begin preparing for cryogenic loading of the Shuttle external tank at the T-6 hour mark. Once its task is completed by the T-3 hour mark, this group of individuals will be succeeded by a fresh crew to carry on through launch.
Another example of a specialized countdown crew are the personnel who oversee the ground launch sequencer (GLS). GLS operators may be on hand earlier in the countdown, but they do not have a formal role until the day of launch. The GLS crew -- a primary operator, two backups and a fourth to retrieve data -- are on station by the T-2 hour mark to call up the GLS software and prepare for the final count. At T-20 minutes, the GLS will begin issuing active commands, and at T-9 minutes, it assumes automatic control of the count.
As they work through S0007, the launch team members at the consoles in the prime firing room are monitoring vehicle and support system performance according to acceptable measurements and parameters. The full set of measurements that must be checked totals about 25,000. About two thirds of these originate from the flight vehicle and one third from the ground support equipment and facilities. Measurements generally fall into one of two categories, Launch Commit Criteria or supporting data.
Launch Commit Criteria are those parameters that have safety-related or mission success implications; they define what constitutes a vehicle that is ready to fly as well as the conditions under which it is permissible to launch. Launch Commit Criteria is implemented at T-6 hours (prior to ET load). There are Launch Commit Criteria that guard against flight hardware damage and those designed specifically for \Jastronaut\j safety. For instance, there is a permissible concentration of gaseous \Jhydrogen\j within the orbiter's aft fuselage that deals with \Jastronaut\j safety and a slightly different permissible concentration that may affect shuttle hardware issues. Even the weather must meet specific Launch Commit Criteria requirements in order for a liftoff to proceed.
The other set of data represents supporting information available for the engineers to help maintain a specific hardware configuration or to aid in troubleshooting problems. An example would be the temperature limits for the \Jhydrogen\j vent line that is used to safely transport the gaseous \Jhydrogen\j from the external tank to its flarestack at the pad for disposal by burning.
This vast data base on system performance is part of the Launch Processing System and is available for the engineers to access at any point in the launch countdown. About 2,300 Launch Commit Criteria measurements are monitored. As the countdown progresses toward liftoff, all attention focuses on the Launch Commit Criteria. All the parameters must be met prior to passing a "Go" for launch.
#
"Shuttle Operation Communication",204,0,0,0
Communication in the prime firing room is carefully routed through the Operational Intercommunications System (OIS). It is a closed-loop digital voice system utilizing fiber optic cable. During countdown, the NTD uses one frequency as the command channel for overall countdown \Jintegration\j. The Test Conductors use separate channels to individually lead their specific subset of the countdown. The Test Conductors and the NASA Test Director communicate with each other on an as-needed basis on issues such as status checks, safety and command responses. The use of different channels separates communication traffic and keeps it at a manageable level.
Acronyms are used liberally by the entire team to keep verbiage to a minimum. All Firing Room console positions are assigned unique 'call signs' that are used by the team for quick and positive identification of who is talking. The protocol has the initiator of the dialogue calling the person he wishes to talk to followed by his own call sign. This allows the person called to know who is calling. For instance,
"OTC, NTD, Begin crew module closeout" means the NTD is directing the OTC to begin preparing the crew module for closing the hatch and launch. Other acronyms are used for system descriptions such as PRSD for the Power Reactant Storage and Distribution System which is the liquid oxygen and liquid \Jhydrogen\j fuels that power the orbiter's on-board fuel cells that create electric power. When the flight crew enters the orbiter, the astronauts will hook up to the same communication channel as the OTC. At the T-20 minute mark, the NTD switches over to this channel and it now becomes the Command channel. The \Jintegration\j console in the prime and backup firing rooms maintain communication channels with the other centers, such as the Mission Evaluation Room at Johnson Space Center in Houston, and the \JHuntsville\j Operations Support Center at the Marshall Space Flight Center in \JAlabama\j. The Mission Evaluation Room plans and implements flight data retrieval, processing, exchange analysis, evaluation and reporting, and post-mission evaluation. The \JHuntsville\j center provides technical support on the Shuttle main engines, external tank and solid rocket boosters.
Hookups also are established with the contractors' home offices. Orbiter manufacturer Rockwell International maintains a support room at its Downey, Calif., plant that remains open around the clock throughout a Shuttle mission.
#
"Shuttle Firing Room Protocol",205,0,0,0
A special type of discipline is exercised in the prime firing room, commensurate with its importance. Launch team members undergo training in the rules and regulations governing their conduct. These include limiting conversation to the business at hand, no personal \Jtelephone\j calls except in emergency, and no reading of non-work related materials. During time-critical operations, personnel remain at their assigned stations.
From T-3 hours on, entrance into the prime firing room is restricted. Only personnel with firing room badges are allowed in and movement is minimized. A prime firing room badge is issued only to personnel having a direct console position related to the terminal portion of the count. At T-20 minutes, the door to the prime firing room is locked. The intent is to eliminate distractions and allow the team to focus its attention on the countdown.
#
"Launch Pad Activities During Countdown",206,0,0,0
While countdown activities are controlled from the LCC prime firing room, personnel at the pad perform different tasks required for launch preparations. From T-11 hours to T-6 hours, a great deal of final preparation work occurs at the pad: rollback of the Rotating Service Structure; installation of time-critical flight crew equipment; performance of the pre-ingress switch list; sampling of crew seat oxygen; and installation of the crew escape pole in the orbiter. Overseeing these activities and keeping the NTD informed of their progress is the pad leader.
After the External Tank is loaded, only critical and highly specialized teams will travel to the pad again before liftoff. One of these is the Final Inspection Team, also referred to as the ice team, which conducts a preflight walkdown of the vehicle and pad during the two-hour hold at T-3 hours. Another ice team is stationed in the backup firing room. Its job is to monitor the external tank's \Jinsulation\j and attachment struts for excessive ice formation before, during and after loading of the supercold liquid \Jhydrogen\j and liquid oxygen. Click here for a picture of the ice team on the Mobile Launch Platform.
Another specialized team is the white room closeout crew, which also proceeds to the pad during the two-hour-hold at the T-3 hour mark. Their task is to insure that the orbiter cockpit is properly configured for flight and to assist the astronauts with entry into the orbiter. They also ensure the side hatch is properly closed and that the white room is configured for launch.
Handling of any anomaly at the pad that should occur during or after external tank loading is the responsibility of a red crew. This is not a pre-existing unit, but a team assembled from a pool of specially trained workers with experience in the particular problem area. Members have been specially trained in fire and rescue techniques and must have undergone special certification. Their activity at the pad would be conducted by the system engineer responsible for the anomalous system and under the strict direction of the NASA Test Director..
#
"Launch: Final Go",207,0,0,0
Beginning approximately 15 minutes before launch, readiness polls are conducted by the three teams that together comprise the Shuttle Launch Team. The NTD verifies that the prime launch team is reporting no violation of the Launch Commit Criteria. The \JEngineering\j Director who heads up the \JEngineering\j Support Team verifies no constraints to continuing with the final count. And the Mission Management Team Chairman verifies that there are no open issues with any of the senior element managers.
These three verifications are passed on to the Shuttle Launch Director, who conducts a KSC management poll. Assuming all responsible personnel are in agreement, the Launch Director gives his permission to proceed with the countdown to the NTD. The NTD in turn sets in motion the final nine minutes of the countdown, automatically controlled by the Ground Launch Sequencer.
Once the Shuttle's twin solid rocket boosters ignite at T-0, responsibility for the mission switches from KSC to the Mission Control Center at Johnson Space Center in Houston. KSC once again assumes responsibility after the orbiter has landed and the flight crew has exited the vehicle.
Jan. 5 President Nixon proposes development of a reusable space transportation system, the Space Shuttle.
March 15 NASA selects the three-part configuration for the Space Shuttle -- reusable orbiter, partly reusable SRB and an expendable external tank.
Aug. 9 Rockwell receives NASA contract for construction of the Space Shuttle orbiter.
1975
Oct. 17 First Space Shuttle main engine tested at the National Space Technology Laboratories, Miss.
Sept. 17 Rollout of orbiter Enterprise (OV-101).
1976
July 18 Thiokol conducts 2-minute firing of an SRB at Brigham City, \JUtah\j.
Aug. 12 First free flight Approach and Landing Test (ALT) of orbiter Enterprise from Shuttle carrier \Jaircraft\j at Dryden Flight Research Center, Calif. Flight duration: 5 minutes, 21 seconds. Landing occurred on Runway 17.
Sept. 13 Second Enterprise ALT flight of 5 minutes, 28 seconds; landing on Runway 15. (Three more ALT flights were flown by Enterprise on Sept. 23 Oct. 12 and Oct. 25.)
1978
Jan. 18 Thiokol conducts second test firing of an SRB.
1979
March 8 Orbiter Columbia (OV-102) transported 38 miles overland from Palmdale to Dryden Flight Research Center.
March 20-24 Columbia flown on Shuttle carrier \Jaircraft\j to Kennedy Space Center with overnight stops at El Paso and San Antonio, \JTexas\j, and Eglin AFB, Fla.
June 15 First SRB qualification test firing; 122 seconds.
1980
Feb. 20 Flight readiness firing of Columbia's main engines; 20 seconds.
April 20-21 Columbia returned to KSC by Shuttle carrier \Jaircraft\j via Tinker AFB, Okla.
Aug. 4 Columbia mated with SRBs and external tank for STS-2 mission.
Aug. 26 Space Shuttle vehicle moved to Launch Complex 39A for STS-2 mission.
Nov. 12-14 STS-2, first flight of an orbiter previously flown in space
Nov. 24-25 Columbia transported back to KSC via Bergstrom AFB, \JTexas\j.
Nov. 26 Columbia mated to SRBs and external tank at Vehicle Assembly Building (VAB) for STS-l mission.
Dec. ll Spacelab l arrives at KSC.
Dec. 29 Space Shuttle vehicle moved from VAB to Launch Complex 39A for STS-l mission.1981
1982
Feb. 3 Columbia moved to VAB for mating in preparation for STS-3 mission.
Feb. 16 Assembled Space Shuttle vehicle moved from VAB to launch pad for STS-3 mission.
March 22-30 STS-3 mission; landing at White Sands, N.M.
April 6 Columbia returned to KSC from White Sands.
May 16 Columbia moved to VAB for mating in preparation for STS-4.
May 25 STS-4 vehicle moved to launch pad.
June 27-July 4 STS-4 mission flown; first concrete runway landing at Edwards AFB.
June 30 Orbiter Challenger (OV-099) rolled out at Palmdale.
July l Challenger moved overland to Dryden.
July 4-5 Challenger flown to KSC via Ellington AFB, \JTexas\j.
July 14-15 Columbia flown to KSC via Dyess AFB, \JTexas\j.
Sept. 9 Columbia mated with SRBs and external tank in preparation for STS-5.
Sept. 21 STS-5 vehicle moved to launch pad.
Nov. ll-16 STS-5 mission; landing at Edwards AFB.
Nov. 21-22 Columbia returned to KSC via Kelly AFB, \JTexas\j
Nov. 23 Challenger moved to VAB and mated for STS-6.
Nov. 30 STS-6 vehicle moved to launch pad.
Dec. 18 Flight readiness firing of Challenger's main engines; 20 seconds.
1983
Jan. 22 Second flight readiness firing of Challenger's main engines; 22 seconds.
April 4-9 STS-6 mission, first flight of Challenger.
May 21 Challenger moved to VAB for mating in preparation for STS-7 mission.
May 26 Challenger moved to launch pad for STS-7.
June 18-24 STS-7 mission flown with landing at Edwards AFB.
July 26 Challenger moved to VAB for mating in preparation for STS-8.
June 28-29 Challenger flown back to KSC via Kelly AFB.
Aug. 2 STS-8 vehicle moved to launch pad.
Aug. 30-Sept. 5 STS-8 mission; first night launch and landing at Edwards AFB.
Sep. 9 Challenger returned to KSC via Sheppard AFB, \JTexas\j.
Sept. 23 Columbia moved to VAB for mating in preparation for STS-9.
Sept. 28 STS-9vehicle moved to launch pad.
Oct. 17 STS-9launch vehicle moved back to VAB from pad because of SRB nozzle problem.
Oct. 19 Columbia moved to Orbiter Processing Facility.
Nov. 5 Orbiter Discovery (OV-103) moved overland to Dryden.
Nov. 6 Discovery transported to Vandenberg AFB, Calif.
Nov. 8 STS-9vehicle again moved to launch pad.
Nov. 8-9 Discovery flown from Vandenberg AFB to KSC via Carswell AFB, \JTexas\j.
Nov. 28-Dec. 8 STS-9mission; landing at Edwards AFB.
Dec. 14-15 Columbia flown to KSC via El Paso, Kelly AFB and Eglin AFB.
l984
Jan. 6 Challenger moved to VAB for mating in preparation of STS 41 B mission.
Jan. ll STS 41-B vehicle moved to launch pad.
Feb. 3-ll STS 41-B mission; first landing at KSC.
March 14 Challenger moved to VAB for mating in preparation for STS 41-C mission.
March 19 STS 41-C vehicle moved to launch pad.
April 6-13 STS 41-C mission; landing at Edwards AFB.
April 17-18 Challenger flown back to KSC via Kelly AFB.
May 12 Discovery moved to VAB for mating in preparation for STS 41-D.
May 19 STS 41-D vehicle moved to launch pad.
June 2 Flight readiness firing of Discovery's main engines.
June 25 STS 41-D launch attempt scrubbed because of computer problem.
June 26 STS 41-D launch attempt scrubbed following main engine shutdown at T minus 4 seconds.
July 14 STS 41-D vehicle moved back to VAB for remanifest of payloads.
Aug. 9 STS 41-D vehicle again moved out to the launch pad.
Aug. 30-Sept. 5 STS 41-D mission; first flight of Discovery; landing at Edwards AFB.
Sept. 8 Challenger moved to VAB for mating in preparation for STS 41-G mission.
Sept. 9-10 Discovery returned to KSC via Altus AFB, Okla.
Sept. 13 STS 41-G launch vehicle moved to launch pad.
Oct. 5-13 STS 41-G mission; landing at KSC.
Oct. 18 Discovery moved to VAB for mating in preparation for STS 51-A mission.
Oct. 23 STS 51-A launch vehicle moved to launch pad.
Nov. 7 STS 51-A launch scrubbed because of high shear winds.
Nov. 8-16 STS 51-A mission; landing at KSC.
1985
Jan. 5 Discovery moved to launch pad for STS 51-C mission.
Jan. 24-27 STS 51-C mission landing at KSC.
Feb. 10 Challenger moved to VAB for mating in preparation for STS 51-E mission.
Feb. 15 STS 51-E vehicle moved to launch pad.
March 4 STS 51-E vehicle rolled back to VAB; mission cancelled; payloads combined with STS 51-B.
March 23 Discovery moved to VAB for mating in preparation for STS 51-D mission.
March 28 STS 51-D vehicle moved to launch pad.
April 6 Atlantis (OV-104) rollout at Palmdale.
April 10 Challenger moved to VAB for mating in preparation for STS 51-B mission.
April 12-19 STS 51-D mission; landing at KSC.
April 13 Atlantis ferried to KSC via Ellington AFB, \JTexas\j.
April 15 Challenger moved to launch pad for 51-B missing.
April 29-May 6 STS 51-B mission; landing at Edwards AFB.
May 10 Challenger transported back to KSC via Kelly AFB.
May 28 Discovery moved to VAB for mating in preparation for STS 51-G.
June 4 STS 51-G vehicle moved to the launch pad.
June 17-24 STS 51-G mission; landing Edwards AFB.
June 24 Challenger moved to VAB for mating in preparation for STS 51-F.
June 28 Discovery ferried back to KSC via Bergstrom AFB, \JTexas\j.
June 29 STS 51-F vehicle moved to the launch pad.
July ll Refurbished Columbia moved overland from Palmdale to Dryden.
July 12 STS 51-F launch scrubbed at T-minus 3 seconds because of main engine shutdown.
July 14 Columbia returned to KSC via Offutt AFB, Neb.
July 29-Aug. 6 STS 51-F mission landing at Edwards AFB.
July 30 Discovery moved to VAB for mating in preparation for STS 51-I mission.
Aug. 6 STS 51-I vehicle moved to the launch pad.
Aug. 10-ll Challenger flown to KSC via Davis-Monthan AFB, Ariz.; Kelly AFB; and Eglin AFB.
Aug. 24 STS 51-I mission scrubbed at T minus 5 minutes because of bad weather.
Aug. 25 STS 51-I mission scrubbed at T-minus 9 minutes because of an onboard computer problem.
Aug. 27-Sept. 3 STS 51-I mission; landing at Edwards AFB.
August 29 Atlantis moved to launch pad for the 51-J mission.
Sept. 7-8 Discovery flown back to KSC via Kelly AFB.
Sept. 12 Flight readiness firing of Atlantis' main engines; 20 seconds.
Oct. 3-7 STS 51-J mission; landing at Edwards AFB.
Oct. ll Atlantis returned to KSC via Kelly AFB.
Oct. 12 Challenger moved to VAB for mating in preparation for the STS 61-A mission.
Oct. 16 Challenger vehicle moved to the launch pad for STS 61-A mission.
Oct. 30-Nov. 6 STS 61-A mission; landing at Edwards AFB.
Nov. 8 Atlantis moved to VAB for mating in preparation for the STS 61-B.
Nov. 10-ll Challenger flown back to KSC via Davis-Monthan AFB, Kelly AFB and Eglin AFB.
Nov. 12 STS 61-B vehicle moved to the launch pad.
Nov. 18 Enterprise (OV-101) flown from KSC to Dulles Airport, Washington, D.C., and turned over to the Smithsonian Institution.
Nov. 22 Columbia moved to the VAB for mating in preparation STS 61-C.
Nov. 26-Dec. 3 STS 61-B mission landing at Edwards AFB.
Dec. l STS 61-C vehicle moved to launch pad.
Dec. 7 Atlantis returned to KSC via Kelly AFB.
Dec. 16 Challenger moved to VAB for mating in preparation for the STS 51-L mission.
Dec. 19 STS 61-C mission scrubbed at T minus 13 seconds because of SRB auxiliary power unit problem.
Dec. 22 STS 51-L vehicle moved to Launch Pad 39B.
1986
Jan. 6 STS 61-C mission scrubbed at T minus 31 seconds because of liquid oxygen valve problem on pad.
Jan. 7 STS 61-C mission scrubbed at T minus 9 minutes because of weather problems at contingency landing sites.
Jan. 10 STS 61-C mission scrubbed T minus 9 minutes because of bad weather at KSC.
Jan. 12-18 STS 61-C mission; landing at Edwards AFB.
Jan. 22-23 Columbia returned to KSC via Davis-Monthan AFB, Kelly AFB and Eglin AFB.
Jan. 27-28 STS 51-L launched from Pad B. Vehicle exploded 1 minute, 13 seconds after liftoff resulting loss of seven crew members.
Feb. 3 President Reagan announced the formation of the Presidential Commission on the Space Shuttle Challenger Accident, headed by William P. Rogers, former Secretary of State.
March 24 NASA publishes "Strategy for Safely Returning the Space Shuttle to Flight Status."
May 12 President Reagan appoints Dr. James C. Fletcher NASA Administrator.
July 8 NASA establishes Safety, Reliability Maintainability, and Quality Assurance Office.
July 14 NASA's plan to implement the recommendations of the Rogers commission was submitted to President Reagan.
Aug. 15 President Reagan announced his decision to support a replacement for the Challenger. At the same time, it was announced that NASA no longer would launch commercial satellites, except for those which are Shuttle-unique or have national security or foreign policy implications.
Aug. 22 NASA announced the beginning of a series of tests designed to verify the ignition pressure dynamics of the Space Shuttle solid rocket motor field joint.
Sept. 5 Study contracts were awarded to five aerospace firms for conceptual designs of an alternative or Block II Space Shuttle solid rocket motor.
Sept. 10 \JAstronaut\j Bryan O'Connor was named chairman of Space Flight Safety Panel. This panel, with oversight responsibility for all NASA manned space program activities, reports to the Associate Administrator for Safety, Reliability, Maintainability and Quality Assurance.
Oct. 2 After an intensive study, NASA announced the decision to test fire the redesigned solid rocket motor in a horizontal attitude to best simulate the critical conditions on the field joint which failed during the 51-L mission.
Oct. 30 Discovery moved to OPF where more than 200 modifications are accomplished for STS-26 mission.
Nov. 6 Office of the Director, National Space Transportation System, established in the NASA Headquarters Office of Space Flight.
1987
July 31 Rockwell International awarded contract to build a fifth orbiter to replace the Challenger.
Aug. 3 Discovery in the Orbital Processing Facility is powered up for STS-26 mission.
1988
Mid-Jan. Main engines are installed in Discovery.
March 28 Stacking of Discovery's SRBs gets underway.
May 28 Stacking of Discovery's SRBs completed.
June 10 SRBs and External Tank are mated.
June 14 The fourth full-duration test firing of the redesigned SRB motor is carried out.
June 21 Discovery rolls over from OPF to the VAB.
July 4 Discovery moved to Launch Pad 39B for STS-26 mission.
Aug. 10 Flight Readiness Firing of Discovery's main engines is conducted successfully.
#
"Space Shuttle System",209,0,0,0
\JSpace Shuttle Program\j
\JSpace Shuttle Requirements\j
\JShuttle Launch Sites\j
\JShuttle Background and Status\j
\JShuttle Mission Profile\j
\JShuttle Abort Modes\j
\JOrbiter Ground Turnaround\j
\JOrbiter Operational Improvements\j
\JSpace Shuttle Main Engine Margin Improvement Program\j
\JSSME Flight Program\j
\JShuttle Solid Rocket Motor (SRM)\j
\JShuttle Solid Rocket Boosters\j
\JShuttle External Tank\j
\JShuttle Liquid Oxygen Tank\j
\JShuttle Intertank\j
\JShuttle Liquid Hydrogen Tank\j
\JShuttle ET Thermal Protection System\j
\JShuttle ET Hardware\j
\JShuttle ET Range Safety System\j
\JShuttle Orbiter Structures\j
\JOrbiter Passive Thermal Control\j
\JOrbiter Purge, Vent and Drain System\j
\JOrbiter In-Flight Crew Escape System\j
\JOrbiter Emergency Egress Slide\j
\JOrbiter Secondary Emergency Egress\j
\JOrbiter Side Hatch Jettison\j
\JShuttle Crew Equipment\j
\JSpace Shuttle Orbiter Systems\j
\JShuttle Launch and Flight Operations\j
#
"Shuttle Abort Modes",210,0,0,0
\JShuttle Aborts\j
\JShuttle, Return To Launch Site\j
\JShuttle, Transatlantic Landing Abort\j
\JShuttle, Abort To Orbit\j
\JShuttle, Abort Once Around\j
\JShuttle, Contingency Abort\j
#
"Orbiter Operational Improvements",211,0,0,0
\JOrbital Maneuvering System and Reaction Control System (Modifications)\j
\JOrbiter, Fuel Cell Power Plants (Modifications)\j
\JOrbiter, Auxiliary Power Units (Modifications)\j
\JOrbiter, Main Landing Gear (Modifications)\j
\JOrbiter, Nose Wheel Steering (Modifications)\j
\JOrbiter, Thermal Protection System (Modifications)\j
\JOrbiter, Wing Modification\j
\JOrbiter, Mid-Fuselage Modifications\j
\JOrbiter, General Purpose Computers (Modifications)\j
\JOrbiter, Inertial Measurement Units (Modifications)\j
\JShuttle Payload Deployment and Retrieval System\j
\JShuttle Payload Retention Mechanisms\j
\JSpace Flight Tracking and Data Network\j
\JSatellite System, Tracking and Data Relay\j
\JShuttle Avionics Systems\j
\JShuttle Data Processing System\j
\JShuttle Guidance, Navigation and Control\j
\JShuttle Flight Control System Hardware\j
\JShuttle Navigation Aids\j
\JShuttle Inertial Measurement Units\j
\JShuttle Star Trackers\j
\JShuttle Crewman Optical Alignment Sight\j
\JShuttle Air Data System\j
\JShuttle Microwave Scan Beam Landing System\j
\JShuttle Radar Altimeter\j
\JShuttle Accelerometer Assemblies\j
\JOrbiter Rate Gyro Assemblies\j
\JSolid Rocket Booster Rate Gyro Assemblies\j
\JShuttle Rotational Hand Controller\j
\JShuttle Translational Hand Controller\j
\JShuttle Control Stick Steering Push Button Light Indicators\j
\JShuttle Rudder Pedals\j
\JShuttle Speed Brake/Thrust Controller\j
\JShuttle Body Flap Switch\j
\JShuttle Aerosurface Servoamplifiers\j
\JShuttle Digital Autopilot\j
\JShuttle Rendezvous Thrusting Maneuvers\j
\JShuttle Component Locations\j
\JShuttle, Dedicated Display Systems\j
\JShuttle, Attitude Director Indicator\j
\JShuttle, Horizontal Situation Indicator\j
\JShuttle, Alpha Mach Indicator\j
\JShuttle, Altitude/Vertical Velocity Indicator\j
\JShuttle, Surface Position Indicator\j
\JShuttle, Flight Control System Push Button Light Indicators\j
\JShuttle, Reaction Control System Command Lights\j
\JShuttle, G-Meter\j
\JShuttle, Head-Up Display\j
#
"Shuttle Launch and Flight Operations",217,0,0,0
\JShuttle, Pre-Launch Operations\j
\JShuttle, Pre-Launch Propellant-Loading\j
\JShuttle, Final Pre-Launch Activities\j
\JShuttle, Launch Control Center\j
\JShuttle, Launch Countdown\j
\JShuttle, Mission Control Center\j
\JShuttle, Marshall Payload Operations Control Center\j
\JSpace Tracking and Data Acquisition\j
\JShuttle Support Ground Network\j
\JSpace Network\j
\JShuttle Flight Operations\j
\JShuttle Launch Abort Modes\j
\JShuttle, On-Orbit Operations\j
\JShuttle, Maneuvering In Orbit\j
#
"Space Shuttle Program",218,0,0,0
The Space Shuttle is developed by the National Aeronautics and Space Administration. NASA coordinates and manages the Space Transportation System (NASA's name for the overall Shuttle program), including intergovernmental agency requirements and international and joint projects. NASA also oversees the launch and space flight requirements for civilian and commercial use.
The Space Shuttle system consists of four primary elements: an orbiter \Jspacecraft\j, two Solid Rocket Boosters (SRB), an external tank to house fuel and oxidizer and three Space Shuttle main engines.
The orbiter is built by Rockwell International's Space Transportation Systems Division, Downey, Calif., which also has responsibility for the \Jintegration\j of the overall space transportation system. Both orbiter and \Jintegration\j contracts are under the direction of NASA's Johnson Space Center in Houston, \JTexas\j.
The SRB motors are built by the Wasatch Division of Morton Thiokol Corp., Brigham City, \JUtah\j, and are assembled, checked out and refurbished by United Space Boosters Inc., Booster Production Co., Kennedy Space Center. Cape Canaveral, Fla.
The external tank is built by Martin Marietta Corp. at its Michoud facility, New Orleans, La., and the Space Shuttle main engines are built by Rockwell's Rocketdyne Division, Canoga Park, Calif. These contracts are under the direction of NASA's George C. Marshall Space Flight Center, \JHuntsville\j, Ala.
#
"Space Shuttle Requirements",219,0,0,0
The Shuttle will transport cargo into near Earth orbit 100 to 217 nautical miles (115 to 250 statute miles) above the Earth. This cargo -- or payload -- is carried in a bay 15 feet in diameter and 60 ft long.
Major system requirements are that the orbiter and the two solid rocket boosters be reusable.
Other features of the Shuttle:
The orbiter has carried a flight crew of up to eight persons. A total of 10 persons could be carried under emergency conditions. The basic mission is 7 days in space. The crew compartment has a shirtsleeve environment, and the acceleration load is never greater than 3 Gs. In its return to Earth, the orbiter has a cross-range maneuvering capability of 1,100 nautical miles (1,265 statute miles).
The Space Shuttle is launched in an upright position, with thrust provided by the three Space Shuttle engines and the two SRB. After about 2 minutes, the two boosters are spent and are separated from the external tank. They fall into the ocean at predetermined points and are recovered for reuse.
The Space Shuttle main engines continue firing for about 8 minutes. They shut down just before the craft is inserted into orbit. The external tank is then separated from the orbiter. It follows a ballistic trajectory into a remote area of the ocean but is not recovered.
There are 38 primary Reaction Control System (RCS) engines and six vernier RCS engines located on the orbiter. The first use of selected primary reaction control system engines occurs at orbiter/external tank separation. The selected primary reaction control system engines are used in the separation sequence to provide an attitude hold for separation. Then they move the orbiter away from the external tank to ensure orbiter clearance from the arc of the rotating external tank. Finally, they return to an attitude hold prior to the initiation of the firing of the Orbital Maneuvering System (OMS) engines to place the orbiter into orbit.
The primary and/or vernier RCS engines are used normally on orbit to provide attitude pitch, roll and yaw maneuvers as well as translation maneuvers.
The two OMS engines are used to place the orbiter on orbit, for major velocity maneuvers on orbit and to slow the orbiter for reentry, called the deorbit maneuver. Normally, two OMS engine thrusting sequences are used to place the orbiter on orbit, and only one thrusting sequence is used for deorbit.
The orbiter's velocity on orbit is approximately 25,405 feet per second (17,322 statute miles per hour). The deorbit maneuver decreases this velocity approximately 300 fps (205 mph) for reentry.
In some missions, only one OMS thrusting sequence is used to place the orbiter on orbit. This is referred to as direct insertion. Direct insertion is a technique used in some missions where there are high-performance requirements, such as a heavy payload or a high orbital altitude. This technique uses the Space Shuttle main engines to achieve the desired apogee (high point in an orbit) altitude, thus conserving orbital maneuvering system propellants. Following jettison of the external tank, only one OMS thrusting sequence is required to establish the desired orbit altitude.
For deorbit, the orbiter is rotated tail first in the direction of the velocity by the primary reaction control system engines. Then the OMS engines are used to decrease the orbiter's velocity.
During the initial entry sequence, selected primary RCS engines are used to control the orbiter's attitude (pitch, roll and yaw). As aerodynamic pressure builds up, the orbiter flight control surfaces become active and the primary reaction control system engines are inhibited.
During entry, the thermal protection system covering the entire orbiter provides the protection for the orbiter to survive the extremely high temperatures encountered during entry. The thermal protection system is reusable (it does not burn off or ablate during entry).
The unpowered orbiter glides to Earth and lands on a runway like an airplane. Nominal touchdown speed varies from 184 to 196 knots (213 to 225 miles per hour).
The main landing gear wheels have a braking system for stopping the orbiter on the runway, and the nose wheel is steerable, again similar to a conventional airplane.
There are two launch sites for the Space Shuttle. Kennedy Space Center (KSC) in \JFlorida\j is used for launches to place the orbiter in equatorial orbits (around the equator), and Vandenberg Air Force Base launch site in \JCalifornia\j will be used for launches that place the orbiter in polar orbit missions.
Landing sites are located at the KSC and Vandenberg. Additional landing sites are provided at Edwards Air Force Base in \JCalifornia\j and White Sands, N.M. Contingency landing sites are also provided in the event the orbiter must return to Earth in an emergency.
#
"Shuttle Launch Sites",220,0,0,0
Space Shuttles destined for equatorial orbits are launched from the KSC, and those requiring polar orbital planes will be launched from Vandenberg.
Orbital mechanics and the complexities of mission requirements, plus safety and the possibility of infringement on foreign air and land space, prohibit polar orbit launches from the KSC.
Kennedy Space Center launches have an allowable path no less than 35 degrees northeast and no greater than 120 degrees southeast. These are \Jazimuth\j degree readings based on due east from KSC as 90 degrees.
A 35-degree \Jazimuth\j launch places the \Jspacecraft\j in an orbital inclination of 57 degrees. This means the \Jspacecraft\j in its orbital trajectories around the Earth will never exceed an Earth latitude higher or lower than 57 degrees north or south of the equator.
A launch path from KSC at an \Jazimuth\j of 120 degrees will place the \Jspacecraft\j in an orbital inclination of 39 degrees (it will be above or below 39 degrees north or south of the equator).
These two azimuths - 35 and 120 degrees - represent the launch limits from the KSC. Any \Jazimuth\j angles further north or south would launch a \Jspacecraft\j over a habitable land mass, adversely affect safety provisions for abort or vehicle separation conditions, or present the undesirable possibility that the SRB or external tank could land on foreign land or sea space.
Launches from Vandenberg have an allowable launch path suitable for polar insertions south, southwest and southeast.
The launch limits at Vandenberg are 201 and 158 degrees. At a 201-degree launch \Jazimuth\j, the \Jspacecraft\j would be orbiting at a 104-degree inclination. Zero degrees would be due north of the launch site, and the orbital trajectory would be within 14 degrees east or west of the north-south pole meridian.
At a launch \Jazimuth\j of 158 degrees, the \Jspacecraft\j would be orbiting at a 70-degree inclination, and the trajectory would be within 20 degrees east or west of the polar meridian. Like KSC, Vandenberg has allowable launch azimuths that do not pass over habitable areas or involve safety, abort, separation and political considerations.
Mission requirements and payload weight penalties also are major factors in selecting a launch site.
The Earth rotates from west to east at a speed of approximately 900 nautical miles per hour (1,035 mph). A launch to the east uses the Earth's rotation somewhat as a springboard. The Earth's rotational rate also is the reason the orbiter has a cross-range capability of 1,100 nautical miles (1,265 statute miles) to provide the abort-once-around capability in polar orbit launches.
Attempting to launch and place a \Jspacecraft\j in polar orbit from KSC to avoid habitable landmass would be uneconomical because the Shuttle's payload would be reduced severely-down to approximately 17,000 pounds. A northerly launch into polar orbit of 8 to 20 degrees \Jazimuth\j would necessitate a path over a landmass; and most safety, abort, and political constraints would have to be waived. This prohibits polar orbit launches from the KSC.
NASA's latest assessment of orbiter ascent and landing weights incorporates currently approved modifications to all vehicle elements, including crew escape provisions, and assumes a maximum Space Shuttle main engine throttle setting of 104 percent. It is noted that the resumption of Space Shuttle flights initially requires more conservative flight design criteria and additional instrumentation, which reduces the following basic capabilities by approximately 1,600 pounds:
Kennedy Space Center Eastern Space and Missile Center (ESMC) satellite deploy missions. The basic cargo-lift capability for a due east (28.5 degrees) launch is 55,000 pounds to a 110-nautical-mile (126-statute-mile) orbit using OV-103 (Discovery) or OV-104 (Atlantis) to support a 4-day satellite deploy mission. This capability will be reduced approximately 100 pounds for each additional nautical mile of altitude desired by the customer.
The payload capability for the same satellite deploy mission with a 57-degree inclination is 41,000 pounds.
The performance for intermediate inclinations can be estimated by allowing 500 pounds per degree of plane change between 28.5 and 57 degrees.
If OV-102 (Columbia) is used, the cargo-lift weight capability must be decreased by approximately 8,400 pounds. This weight difference is attributed to an approximately 7,150-pound difference in inert weight, 850 pounds of orbiter experiments, 300 pounds of additional thermal protection system and 100 pounds to accommodate a fifth cryogenic liquid oxygen and liquid \Jhydrogen\j tank set for the power reactant storage and distribution system.
Vandenberg Air Force Base Western Space and Missile Center (WSMC) satellite deploy missions. Using OV-103 (Discovery) or OV-104 (Atlantis), the cargo-lift weight capability is 29,600 pounds for a 98-degree launch inclination and 110-nautical-mile (126-statute-mile) polar orbit.
Again, an increase in altitude costs approximately 100 pounds per nautical mile. NASA assumes also that the advanced solid rocket motor will replace the filament-wound solid rocket motor case previously used for western test range assessments.
The same mission at 68 degrees inclination (minimum western test range inclination based on range safety limitations) is 49,600 pounds. Performance for intermediate inclinations can be estimated by allowing 660 pounds for each degree of plane change between inclinations of 68 and 98 degrees.
Landing weight limits. All the Space Shuttle orbiters are currently limited to a total vehicle landing weight of 240,000 pounds for abort landings and 230,000 pounds for nominal end-of-mission landings. It is noted that each additional crew person beyond the five-person standard is chargeable to the cargo weight allocation and reduces the payload capability by approximately 500 pounds. (This is an increase of 450 pounds to account for the crew escape equipment.)
#
"Shuttle Background and Status",221,0,0,0
On July 26, 1972, NASA selected Rockwell's Space Transportation Systems Division in Downey, Calif., as the industrial contractor for the design, development, test and evaluation of the orbiter. The contract called for fabrication and testing of two orbiters, a full-scale structural test article, and a main propulsion test article. The award followed years of NASA and Air Force studies to define and assess the feasibility of a reusable space transportation system.
NASA previously (March 31, 1972) had selected Rockwell's Rocketdyne Division to design and develop the Space Shuttle main engines. Contracts followed to Martin Marietta for the external tank (Aug. 16, 1973) and Morton Thiokol's Wasatch Division for the solid rocket boosters (June 27, 1974).
In addition to the orbiter DDT&E contract, Rockwell's Space Transportation Systems Division was given contractual responsibility as system integrator for the overall Shuttle system.
Rockwell's Launch Operations, part of the Space Transportation Systems Division, was under contract to NASA's Kennedy Space Center for turnaround, processing, prelaunch testing, and launch and recovery operations from STS-1 through the STS-11 mission.
On Oct. 1, 1983, the Lockheed Space Operations Co. was awarded the Space Shuttle processing contract at KSC for turnaround processing, prelaunch testing, and launch and recovery operations.
The first orbiter \Jspacecraft\j, Enterprise (OV-101), was rolled out on Sept. 17, 1976. On Jan. 31, 1977, it was transported 38 miles overland from Rockwell's assembly facility at Palmdale, Calif., to NASA's Dryden Flight Research Facility at Edwards Air Force Base for the Approach and Landing Test (ALT) program.
The 9-month-long ALT program was conducted from February through November 1977 at Dryden and demonstrated the orbiter could fly in the atmosphere and land like an airplane except without power, a gliding flight.
The ALT program involved ground tests and flight tests.
The ground tests included taxi tests of the 747 shuttle carrier \Jaircraft\j (SCA) with the Enterprise mated atop the SCA to determine structural loads and responses and assess the mated capability in ground handling and control characteristics up to flight takeoff speed.
The taxi tests also validated 747 steering and braking with the orbiter attached. A ground test of orbiter systems followed the unmanned captive tests. All orbiter systems were activated as they would be in atmospheric flight. This was the final preparation for the manned captive-flight phase.
Five captive flights of the Enterprise mounted atop the SCA with the Enterprise unmanned and Enterprise systems inert were conducted to assess the structural integrity and performance-handling qualities of the mated craft.
Three manned captive flights that followed the five unmanned captive flights included an \Jastronaut\j crew aboard the orbiter operating its flight control systems while the orbiter remained perched atop the SCA. These flights were designed to exercise and evaluate all systems in the flight environment in preparation for the orbiter release (free) flights. They included flutter tests of the mated craft at low and high speed, a separation trajectory test and a dress rehearsal for the first orbiter free flight.
In the five free flights the \Jastronaut\j crew separated the \Jspacecraft\j from the SCA and maneuvered to a landing at Edwards Air Force Base. In the first four such flights the landings were on a dry lake bed; in the fifth, the landing was on Edwards' main concrete runway under conditions simulating a return from space.
The last two free flights were made without the tail cone, which is the \Jspacecraft\j's configuration during an actual landing from Earth orbit. These flights verified the orbiter's pilot-guided approach and landing capability; demonstrated the orbiter's subsonic terminal area energy management autoland approach capability; and verified the orbiter's subsonic airworthiness, integrated system operations and selected subsystems in preparation for the first manned orbital flight.
The flights demonstrated the orbiter's ability to approach and land safely with a minimum gross weight and using several center-of-gravity configurations.
For all of the captive flights and the first three free flights, the orbiter was outfitted with a tail cone covering its aft section to reduce aerodynamic drag and turbulence. The final two free flights were without the tail cone, and the three simulated Space Shuttle main engines and two orbital maneuvering system engines were exposed aerodynamically.
The final phase of the ALT program prepared the \Jspacecraft\j for four ferry flights. Fluid systems were drained and purged, the tail cone was reinstalled and elevon locks were installed.
The forward attachment strut was replaced to lower the orbiter's cant from 6 to 3 degrees. This reduces drag to the mated vehicles during the ferry flights.
After the ferry flight tests, OV-101 was returned to the NASA hangar at Dryden and modified for vertical ground vibration tests at NASA's Marshall Space Flight Center, \JHuntsville\j, Ala.
On March 13, 1978, the Enterprise was ferried atop the SCA to MSFC. At Marshall, Enterprise was mated with the external tank and SRB and subjected to a series of vertical ground vibration tests. These tested the mated configuration's critical structural dynamic response modes, which were assessed against analytical math models used to design the various element interfaces.
These were completed in March 1979. On April 10, 1979 the Enterprise was ferried to Kennedy Space Center, mated with the external tank and SRB and transported via the mobile launcher platform to Launch Complex 39-A. At Launch Complex 39-A, the Enterprise served as a practice and launch complex fit-check verification tool representing the flight vehicles.
It was ferried back to Dryden at Edwards AFB in \JCalifornia\j on Aug. 16, 1979, and then returned overland to Rockwell's Palmdale final assembly facility on Oct. 30, 1979. Certain components were refurbished for use on flight vehicles being assembled at Palmdale. The Enterprise was then returned overland to Dryden on Sept. 6, 1981.
During exhibition at the Paris, May and June 1983, Enterprise was ferried to \JFrance\j for the Air Show as well as to \JGermany\j, \JItaly\j, England and Canada before returning to Dryden.
From April to October 1984, Enterprise was ferried to Vandenberg AFB and to Mobile, Ala., where it was taken by barge to New Orleans, La., for the United States 1984 World's Fair.
In November 1984 it was transported to Vandenberg and used as a practice and fit-check verification tool. On May 24, 1985, Enterprise was ferried from Vandenberg to Dryden.
On Sept. 20, 1985, Enterprise was ferried from Dryden Flight Research Facility to KSC. On Nov. 18, 1985, Enterprise was ferried from KSC to Dulles Airport, Washington, D.C., and became the property of the Smithsonian Institution. The Enterprise was built as a test vehicle and is not equipped for space flight.
The second orbiter, Columbia (OV-102), was the first to fly into space. it was transported overland on March 8, 1979, from Palmdale to Dryden for mating atop the SCA and ferried to KSC. It arrived on March 25, 1979, to begin preparations for the first flight into space.
The structural test article, after 11 months of extensive testing at Lockheed's facility in Palmdale, was returned to Rockwell's Palmdale facility for modification to become the second orbiter available for operational missions. It was redesignated OV-099, the Challenger.
The main propulsion test article (MPTS-098) consisted of an orbiter aft fuselage, a truss arrangement that simulated the orbiter's mid-fuselage and the Shuttle main propulsion system (three Space Shuttle main engines and the external tank). This test structure is at the Stennis Space Center in Mississippi. A series of static firings was conducted from 1978 through 1981 in support of the first flight into space.
On Jan. 29, 1979, NASA contracted with Rockwell to manufacture two additional orbiters, OV-103 and OV-104 (Discovery and Atlantis), convert the structural test article to space flight configuration (Challenger) and modify Columbia from its development configuration to that required for operational flights.
NASA named the first four orbiter spacecrafts after famous exploration sailing ships. In the order they became operational, they are: Columbia (OV-102), after a sailing \Jfrigate\j launched in 1836, one of the first Navy ships to circumnavigate the globe.
Columbia also was the name of the Apollo 11 command module that carried Neil Armstrong, Michael Collins and Edward (Buzz) Aldrin on the first lunar landing mission, July 20, 1969. Columbia was delivered to Rockwell's Palmdale assembly facility for modifications on Jan. 30, 1984, and was returned to KSC on July 14, 1985, for return to flight.
Challenger (OV-099), also a Navy ship, which from 1872 to 1876 made a prolonged exploration of the Atlantic and Pacific oceans. It also was used in the Apollo program for the Apollo 17 lunar module. Challenger was delivered to DSC on July 5, 1982.
Discovery (OV-103), after two ships, the vessel in which Henry Hudson in 1610-11 attempted to search for a northwest passage between the Atlantic and Pacific oceans and instead discovered Hudson Bay and the ship in which Capt. Cook discovered the Hawaiian Islands and explored southern \JAlaska\j and western Canada.
Discovery was delivered to KSC on Nov. 9, 1983. Atlantis (OV-104), after a two-masted ketch operated for the Woods Hole Oceanographic Institute from 1930 to 1966, which traveled more than half a million miles in ocean research. Atlantis was delivered to KSC on April 3, 1985.
In April 1983, under contract to NASA, Rockwell's Space Transportation Systems Division, Downey, Calif., began the construction of structural spares for completion in 1987. The structural spares program consisted of an aft fuselage, crew compartment, forward reaction control system, lower and upper forward fuselage, mid-fuselage, wings (elevons), payload bay doors, vertical stabilizer (rudder/speed brake), body flap and one set of orbital maneuvering system/reaction control system pods.
On Sept. 12, 1985, Rockwell International's Shuttle Operations Co., Houston, \JTexas\j, was awarded the Space Transportation System operation contract at NASA's Johnson Space Center, consolidating work previously performed under 22 contracts by 16 different contractors.
On July 31, 1987, NASA awarded Rockwell's Space Transportation Systems Division, Downey, Calif., a contract to build a replacement Space Shuttle orbiter using the structural spares. The replacement orbiter will be assembled at Rockwell's Palmdale, Calif., assembly facility and is scheduled for completion in 1991. This orbiter is designated OV-105.
#
"Shuttle Mission Profile",222,0,0,0
In the launch configuration, the orbiter and two SRBs are attached to the external tank in a vertical (nose-up) position on the launch pad. Each SRB is attached at its aft skirt to the mobile launcher platform by four bolts.
Emergency exit for the flight crew on the launch pad up to 30 seconds before liftoff is by slidewire. There are seven 1,200-foot-long slidewires, each with one basket. Each basket is designed to carry three persons.
The baskets, 5 feet in diameter and 42 inches deep, are suspended beneath the slide mechanism by four cables. The slidewires carry the baskets to ground level. Upon departing the basket at ground level, the flight crew progresses to a bunker that is designed to protect it from an explosion on the launch pad.
At launch, the three Space Shuttle main engines - fed liquid \Jhydrogen\j fuel and liquid oxygen oxidizer from the external tank - are ignited first. When it has been verified that the engines are operating at the proper thrust level, a signal is sent to ignite the SRB. At the proper thrust-to-weight ratio, initiators (small explosives) at eight hold-down bolts on the SRB are fired to release the Space Shuttle for liftoff. All this takes only a few seconds.
Maximum dynamic pressure is reached early in the ascent, nominally approximately 60 seconds after liftoff. Approximately 1 minute later (2 minutes into the ascent phase), the two SRB have consumed their propellant and are jettisoned from the external tank. This is triggered by a separation signal from the orbiter.
The boosters briefly continue to ascend, while small motors fire to carry them away from the Space Shuttle. The boosters then turn and descend, and at a predetermined altitude, parachutes are deployed to decelerate them for a safe splashdown in the ocean. Splashdown occurs approximately 141 nautical miles (162 statute miles) from the launch site. The boosters are recovered and reused.
Meanwhile, the orbiter and external tank continue to ascend, using the thrust of the three Space Shuttle main engines. Approximately 8 minutes after launch and just short of orbital velocity, the three Space Shuttle engines are shut down (main engine cutoff), and the external tank is jettisoned on command from the orbiter.
The forward and aft reaction control system engines provide attitude (pitch, yaw and roll) and the translation of the orbiter away from the external tank at separation and return to attitude hold prior to the orbital maneuvering system thrusting maneuver.
The external tank continues on a ballistic trajectory and enters the atmosphere, where it disintegrates. Its projected impact is in the Indian Ocean (except for 57-degree inclinations) in the case of equatorial orbits KSC launch) and in the extreme southern Pacific Ocean in the case of a Vandenberg launch.
Normally, two thrusting maneuvers using the two OMS engines at the aft end of the orbiter are used in a two-step thrusting sequence: to complete insertion into Earth orbit and to circularize the \Jspacecraft\j's orbit. The OMS engines are also used on orbit for any major velocity changes.
In the event of a direct-insertion mission, only one OMS thrusting sequence is used.
The orbital altitude of a mission is dependent upon that mission. The nominal altitude can vary between 100 to 217 nautical miles (115 to 250 statute miles).
The forward and aft RCS thrusters (engines) provide attitude control of the orbiter as well as any minor translation maneuvers along a given axis on orbit.
At the completion of orbital operations, the orbiter is oriented in a tail first attitude by the reaction control system. The two OMS engines are commanded to slow the orbiter for deorbit.
The reaction control system turns the orbiter's nose forward for entry. The reaction control system controls the orbiter until atmospheric density is sufficient for the pitch and roll aerodynamic control surfaces to become effective.
Entry interface is considered to occur at 400,000 feet altitude approximately 4,400 nautical miles (5,063 statute miles) from the landing site and at approximately 25,000 feet per second velocity.
At 400,000 feet altitude, the orbiter is maneuvered to zero degrees roll and yaw (wings level) and at a predetermined angle of attack for entry. The angle of attack is 40 degrees. The flight control system issues the commands to roll, pitch and yaw reaction control system jets for rate \Jdamping\j.
The forward RCS engines are inhibited prior to entry interface, and the aft reaction control system engines maneuver the \Jspacecraft\j until a dynamic pressure of 10 pounds per square foot is sensed, which is when the orbiter's ailerons become effective.
The aft RCS roll engines are then deactivated. At a dynamic pressure of 20 pounds per square foot, the orbiter's elevators become active, and the aft RCS pitch engines are deactivated. The orbiter's speed brake is used below Mach 10 to induce a more positive downward elevator trim deflection. At approximately Mach 3.5, the rudder becomes activated, and the aft reaction control system yaw engines are deactivated at 45,000 feet.
Entry guidance must dissipate the tremendous amount of energy the orbiter possesses when it enters the Earth's atmosphere to assure that the orbiter does not either burn up (entry angle too steep) or skip out of the atmosphere (entry angle too shallow) and that the orbiter is properly positioned to reach the desired touchdown point.
During entry, energy is dissipated by the atmospheric drag on the orbiter's surface. Higher atmospheric drag levels enable faster energy dissipation with a steeper trajectory. Normally, the angle of attack and roll angle enable the atmospheric drag of any flight vehicle to be controlled. However, for the orbiter, angle of attack was rejected because it creates surface temperatures above the design specification. The angle of attack scheduled during entry is loaded into the orbiter computers as a function of relative velocity, leaving roll angle for energy control. Increasing the roll angle decreases the vertical component of lift, causing a higher sink rate and energy dissipation rate. Increasing the roll rate does raise the surface temperature of the orbiter, but not nearly as drastically as an equal angle of attack command.
If the orbiter is low on energy (current range-to-go much greater than nominal at current velocity), entry guidance will command lower than nominal drag levels. If the orbiter has too much energy (current range-to-go much less than nominal at the current velocity), entry guidance will command higher-than-nominal drag levels to dissipate the extra energy.
Roll angle is used to control cross range. \JAzimuth\j error is the angle between the plane containing the orbiter's position vector and the heading alignment cylinder tangency point and the plane containing the orbiter's position vector and velocity vector. When the \Jazimuth\j error exceeds a computer-loaded number, the orbiter's roll angle is reversed.
Thus, descent rate and down ranging are controlled by bank angle. The steeper the bank angle, the greater the descent rate and the greater the drag. Conversely, the minimum drag attitude is wings level. Cross range is controlled by bank reversals.
The entry thermal control phase is designed to keep the backface temperatures within the design limits. A constant heating rate is established until below 19,000 feet per second.
The equilibrium glide phase shifts the orbiter from the rapidly increasing drag levels of the temperature control phase to the constant drag level of the constant drag phase. The equilibrium glide flight is defined as flight in which the flight path angle, the angle between the local horizontal and the local velocity vector, remains constant. Equilibrium glide flight provides the maximum downrange capability. It lasts until the drag acceleration reaches 33 feet per second squared.
The constant drag phase begins at that point. The angle of attack is initially 40 degrees, but it begins to ramp down in this phase to approximately 36 degrees by the end of this phase.
In the transition phase, the angle of attack continues to ramp down, reaching the approximately 14-degree angle of attack at the entry Terminal Area Energy Management (TAEM) interface, at approximately 83,000 feet altitude, 2,500 feet per second, Mach 2.5 and 52 nautical miles (59 statute miles) from the landing runway. Control is then transferred to TAEM guidance.
During the entry phases described, the orbiter's roll commands keep the orbiter on the drag profile and control cross range.
TAEM guidance steers the orbiter to the nearest of two heading alignment cylinders, whose radii are approximately 18,000 feet and which are located \Jtangent\j to and on either side of the runway centerline on the approach end. In TAEM guidance, excess energy is dissipated with an S-turn; and the speed brake can be used to modify drag, lift-to-drag ratio and flight path angle in high-energy conditions.
This increases the ground track range as the orbiter turns away from the nearest Heading Alignment Circle (HAC) until sufficient energy is dissipated to allow a normal approach and landing guidance phase capture, which begins at 10,000 feet altitude. The orbiter also can be flown near the velocity for maximum lift over drag or wings level for the range stretch case. The \Jspacecraft\j slows to subsonic velocity at approximately 49,000 feet altitude, about 22 nautical miles (25.3 statute miles) from the landing site.
At TAEM acquisition, the orbiter is turned until it is aimed at a point \Jtangent\j to the nearest HAC and continues until it reaches way point 1. At WP-1, the TAEM heading alignment phase begins. The HAC is followed until landing runway alignment, plus or minus 20 degrees, has been achieved.
In the TAEM pre-final phase, the orbiter leaves the HAC; pitches down to acquire the steep glide slope, increases airspeed; banks to acquire the runway centerline and continues until on the runway centerline, on the outer glide slope and on airspeed. The approach and landing guidance phase begins with the completion of the TAEM pre-final phase and ends when the \Jspacecraft\j comes to a complete stop on the runway.
The approach and landing trajectory capture phase begins at the TAEM interface and continues to guidance lock-on to the steep outer glide slope. The approach and landing phase begins at about 10,000 feet altitude at an equivalent airspeed of 290, plus or minus 12, knots 6.9 nautical miles (7.9 statute miles) from touchdown.
Autoland guidance is initiated at this point to guide the orbiter to the minus 19- to 17-degree glide slope (which is over seven times that of a commercial airliner's approach) aimed at a target 0.86 nautical mile (1 statute mile) in front of the runway.
The \Jspacecraft\j's speed brake is positioned to hold the proper velocity. The descent rate in the later portion of TAEM and approach and landing is greater than 10,000 feet per minute (a rate of descent approximately 20 times higher than a commercial airliner's standard 3-degree instrument approach angle).
At 1,750 feet above ground level, a pre-flare maneuver is started to position the \Jspacecraft\j for a 1.5-degree glide slope in preparation for landing with the speed brake positioned as required. The flight crew deploys the landing gear at this point.
The final phase reduces the sink rate of the \Jspacecraft\j to less than 9 feet per second. Touchdown occurs approximately 2,500 feet past the runway threshold at a speed of 184 to 196 knots (213 to 226 mph).
#
"Shuttle Aborts",223,0,0,0
Selection of an ascent abort mode may become necessary if there is a failure that affects vehicle performance, such as the failure of a Space Shuttle main engine or an orbital maneuvering system. Other failures requiring early termination of a flight, such as a cabin leak, might require the selection of an abort mode.
There are two basic types of ascent abort modes for Space Shuttle missions: intact aborts and contingency aborts. Intact aborts are designed to provide a safe return of the orbiter to a planned landing site. Contingency aborts are designed to permit flight crew survival following more sever failures when an intact abort is not possible. A contingency abort would generally result in a ditch operation.
There are four types of intact aborts: Abort to Orbit (ATO), Abort Once Around (AOA), Transatlantic Landing (TAL) and Return to Launch Site (RTLS).
The ATO mode is designed to allow the vehicle to achieve a temporary orbit that is lower than the nominal orbit. This mode requires less performance and allows time to evaluate problems and then choose either an early deorbit maneuver or an orbital maneuvering system thrusting maneuver to raise the orbit and continue the mission.
The AOA is designed to allow the vehicle to fly once around the Earth and make a normal entry and landing. This mode generally involves two orbital maneuvering system thrusting sequences, with the second sequence being a deorbit maneuver. The entry sequence would be similar to a normal entry.
The TAL mode is designed to permit an intact landing on the other side of the Atlantic Ocean. This mode results in a ballistic trajectory, which does not require an orbital maneuvering system maneuver.
The RTLS mode involves flying downrange to dissipate propellant and then turning around under power to return directly to a landing at or near the launch site.
There is a definite order of preference for the various abort modes. The type of failure and the time of the failure determine which type of abort is selected. In cases where performance loss is the only factor, the preferred modes would be ATO, AOA, TAL and RTLS, in that order.
The mode chosen is the highest one that can be completed with the remaining vehicle performance. In the case of some support system failures, such as cabin leaks or vehicle cooling problems, the preferred mode might be the one that will end the mission most quickly. In these cases, TAL or RTLS might be preferable to AOA or ATO. A contingency abort is never chosen if another abort option exists.
The Mission Control Center-Houston is prime for calling these aborts because it has a more precise knowledge of the orbiter's position than the crew can obtain from onboard systems. Before main engine cutoff, Mission Control makes periodic calls to the crew to tell them which abort mode is (or is not) available. If ground communications are lost, the flight crew has onboard methods, such as cue cards, dedicated displays and display information, to determine the current abort region.
Which abort mode is selected depends on the cause and timing of the failure causing the abort and which mode is safest or improves mission success. If the problem is a Space Shuttle main engine failure, the flight crew and Mission Control Center select the best option available at the time a space shuttle main engine fails.
If the problem is a system failure that jeopardizes the vehicle, the fastest abort mode that results in the earliest vehicle landing is chosen. RTLS and TAL are the quickest options (35 minutes), whereas an AOA requires approximately 90 minutes. Which of these is elected depends on the time of the failure with three good Space Shuttle main engines.
The flight crew selects the abort mode by positioning an abort mode switch and depressing an abort push button.
#
"Shuttle, Return To Launch Site",224,0,0,0
The RTLS abort mode is designed to allow the return of the orbiter, crew, and payload to the launch site, Kennedy Space Center. Approximately 25 minutes after lift-off. The RTLS profile is designed to accommodate the loss of thrust from one space shuttle main engine between liftoff and approximately four minutes 20 seconds, at which time not enough main propulsion system propellant remains to return to the launch site.
An RTLS can be considered to consist of three stages-a powered stage, during which the main engines are still thrusting; an ET separation phase; and the glide phase, during which the orbiter glides to a landing at the KSC. The powered RTLS phase begins with the crew selection of the RTLS abort, which is done after SRB separation.
The crew selects the abort mode by positioning the abort rotary switch to RTLS and depressing the abort push button. The time at which the RTLS is selected depends on the reason for the abort. For example, a three-engine RTLS is selected at the last moment, approximately 3 minutes, 34 seconds into the mission; whereas an RTLS chosen due to an engine out at liftoff is selected at the earliest time, approximately two minutes 20 seconds into the mission (after SOR separation).
After RTLS is selected, the vehicle continues downrange to dissipate excess main propulsion system propellant. The goal is to leave only enough main propulsion system propellant to be able to turn the vehicle around, fly back towards KSC and achieve the proper main engine cutoff conditions so the vehicle can glide to the KSC after external tank separation.
During the downrange phase, a pitch-around maneuver is initiated (the time depends in part on the time of a main engine failure) to orient the orbiter/ external tank configuration to a heads up attitude, pointing toward the launch site. At this time, the vehicle is still moving away from the launch site, but the main engines are now thrusting to null the downrange velocity. In addition, excess orbital maneuvering system and reaction control system propellants are dumped by continuous orbital maneuvering system and reaction control system engine thrustings to improve the orbiter weight and center of gravity for the glide phase and landing.
The vehicle will reach the desired main engine cutoff point with less than 2 percent excess propellant remaining in the external tank. At main engine cutoff minus 20 seconds, a pitch-down maneuver (called powered pitch-down) takes the mated vehicle to the required external tank separation attitude and pitch rate. After main engine cutoff has been commanded, the external tank separation sequence begins, including a reaction control system translation that ensures that the orbiter does not recontact the external tank and that the orbiter has achieved the necessary pitch attitude to begin the glide phase of the RTLS.
After the reaction control system translation maneuver has been completed, the glide phase of the RTLS begins. From then on, the RTLS is handled similarly to a normal entry.
#
"Shuttle, Transatlantic Landing Abort",225,0,0,0
The TAL abort mode was developed to improve the options available when a main engine fails after the last RTLS opportunity but before the first time that an AOA can be accomplished with only two main engines or when a major orbiter system failure, for example, a large cabin pressure leak or cooling system failure, occurs after the last RTLS opportunity, making it imperative to land as quickly as possible.
In a TAL abort, the vehicle continues on a ballistic trajectory across the Atlantic Ocean to land at a predetermined runway. Landing occurs approximately 45 minutes after launch. The landing site is selected near the nominal ascent ground track of the orbiter in order to make the most efficient use of space shuttle main engine propellant.
The landing site also must have the necessary runway length, weather conditions and U.S. State Department approval. Currently, the three landing sites that have been identified for a due east launch are Moron, \JSpain\j; Banjul, The Gambia; and Ben Guerir, Morocco.
To select the TAL abort mode, the crew must place the abort rotary switch in the TAL/AOA position and depress the abort push button before main engine cutoff. (Depressing it after main engine cutoff selects the AOA abort mode.)
The TAL abort mode begins sending commands to steer the vehicle toward the plane of the landing site. It also rolls the vehicle heads up before main engine cutoff and sends commands to begin an orbital maneuvering system propellant dump (by burning the propellants through the orbital maneuvering system engines and the reaction control system engines).
This dump is necessary to increase vehicle performance (by decreasing weight), to place the center of gravity in the proper place for vehicle control, and to decrease the vehicle's landing weight. TAL is handled like a nominal entry.
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"Shuttle, Abort To Orbit",226,0,0,0
An ATO is an abort mode used to boost the orbiter to a safe orbital altitude when performance has been lost and it is impossible to reach the planned orbital altitude. If a Space Shuttle main engine fails in a region that results in a main engine cutoff under speed, the Mission Control Center will determine that an abort mode is necessary and will inform the crew. The orbital maneuvering system engines would be used to place the orbiter in a circular orbit.
#
"Shuttle, Abort Once Around",227,0,0,0
The AOA abort mode is used in cases in which vehicle performance has been lost to such an extent that either it is impossible to achieve a viable orbit or not enough Orbital Maneuvering System (OMS) propellant is available to accomplish the OMS thrusting maneuver to place the orbiter on orbit and the deorbit thrusting maneuver.
In addition, an AOA is used in cases in which a major systems problem (cabin leak, loss of cooling) makes it necessary to land quickly. In the AOA abort mode, one OMS thrusting sequence is made to adjust the post-main engine cutoff orbit so a second orbital maneuvering system thrusting sequence will result in the vehicle deorbiting and landing at the AOA landing site (White Sands, N.M.; Edwards AFB; or KSC). Thus, an AOA results in the orbiter circling the Earth once and landing approximately 90 minutes after liftoff.
After the deorbit thrusting sequence has been executed, the flight crew flies to a landing at the planned site much as it would for a nominal entry.
#
"Shuttle, Contingency Abort",228,0,0,0
Contingency aborts are caused by loss of more than one main engine or failures in other systems. Loss of one main engine while another is stuck at a low thrust setting may also necessitate a contingency abort. Such an abort would maintain orbiter integrity for in-flight crew escape if a landing cannot be achieved at a suitable landing field.
Contingency aborts due to system failures other than those involving the main engines would normally result in an intact recovery of vehicle and crew. Loss of more than one main engine may, depending on engine failure times, result in a safe runway landing. However, in most three-engine-out cases during ascent, the orbiter would have to be ditched. The in-flight crew escape system would be used before ditching the orbiter.
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"Orbiter Ground Turnaround",229,0,0,0
Spacecraft recovery operations at the nominal end-of-mission landing site are supported by approximately 160 Space Shuttle launch operations team members. Ground team members wearing self-contained atmospheric protective ensemble suits that protect them from toxic chemicals approach the \Jspacecraft\j as soon as it stops rolling.
The ground team members take sensor measurements to ensure the atmosphere in the vicinity of the \Jspacecraft\j is not explosive. In the event of propellant leaks, a wind machine truck carrying a large fan will be moved into the area to create a turbulent airflow that will break up gas concentrations and reduce the potential for an explosion.
A ground support equipment air-conditioning purge unit is attached to the right-hand orbiter T-0 umbilical so cool air can be directed through the orbiter's aft fuselage, payload bay, forward fuselage, wings, vertical stabilizer, and orbital maneuvering system/reaction control system pods to dissipate the heat of entry.
A second ground support equipment ground cooling unit is connected to the left-hand orbiter T-0 umbilical \Jspacecraft\j Freon Coolant loops to provide cooling for the flight crew and avionics during the postlanding and system checks. The \Jspacecraft\j fuel cells remain powered up at this time. The flight crew will then exit the \Jspacecraft\j, and a ground crew will power down the \Jspacecraft\j.
AT KSC, the orbiter and ground support equipment convoy move from the runway to the Orbiter Processing Facility.
If the \Jspacecraft\j lands at Edwards, the same procedures and ground support equipment are used as at the KSC after the orbiter has stopped on the runway. The orbiter and ground support equipment convoy move from the runway to the orbiter mate and demate facility at Edwards. After detailed inspection, the \Jspacecraft\j is prepared to be ferried atop the Shuttle carrier \Jaircraft\j from Edwards to KSC. For ferrying, a tail cone is installed over the aft section of the orbiter.
In the event of a landing at an alternate site, a crew of about eight team members will move to the landing site to assist the \Jastronaut\j crew in preparing the orbiter for loading aboard the Shuttle carrier \Jaircraft\j for transport back to the KSC. For landings outside the United States, personnel at the contingency landing sites will be provided minimum training on safe handling of the orbiter with emphasis on crash rescue training, how to tow the orbiter to a safe area, and prevention of propellant conflagration.
Upon its return to the Orbiter Processing Facility (OPF) at KSC, the orbiter is safed (ordnance devices safed), the payload (if any) is removed, and the orbiter payload bay is reconfigured from the previous mission for the next mission. Any required maintenance and inspections are also performed while the orbiter is in the OPF. A payload for the orbiter's next mission may be installed in the orbiter's payload bay in the OPF or may be installed in the payload bay when the orbiter is at the launch pad.
The \Jspacecraft\j is then towed to the Vehicle Assembly Building and mated to the external tank. The external tank and solid rocket boosters are stacked and mated on the mobile launcher platform while the orbiter is being refurbished. Space Shuttle orbiter connections are made and the integrated vehicle is checked and ordnance is installed.
The mobile launcher platform moves the entire space shuttle system on four crawlers to the launch pad, where connections are made and servicing and checkout activities begin. If the payload was not installed in the OPF, it will be installed at the launch pad followed by prelaunch activities.
Space Shuttle launches from Vandenberg will use the Vandenberg Launch Facility (SL6), which was built but never used for the manned orbital laboratory program. This facility was modified for Space Transportation System use.
The runway at Vandenberg was strengthened and lengthened from 8,000 feet to 12,000 feet to accommodate the orbiter returning from space.
When the orbiter lands at Vandenberg, the same procedures and ground support equipment and convoy are used as at KSC after the orbiter stops on the runway. The orbiter and ground support equipment are moved from the runway to the Orbiter Maintenance and Checkout Facility at Vandenberg. The orbiter processing procedures used at this facility are similar to those used at the OPF at the KSC.
Space Shuttle buildup at Vandenberg differs from that of the KSC in that the vehicle is integrated on the launch pad. The orbiter is towed overland from the Orbiter Maintenance and Checkout Facility at Vandenberg to launch facility SL6.
SL6 includes the launch mount, access tower, mobile service tower, launch control tower, payload preparation room, payload changeout room, solid rocket booster refurbishment facility, solid rocket booster disassembly facility, and liquid \Jhydrogen\j and liquid oxygen storage tank facilities.
The SRB start the on-the-launch-pad buildup followed by the external tank. The orbiter is then mated to the external tank on the launch pad.
The launch processing system at the launch pad is similar to the one used at KSC.
Kennedy Space Center Launch Operations has responsibility for all mating, prelaunch testing and launch control ground activities until the Space Shuttle vehicle clears the launch pad tower. Responsibility is then turned over to Mission Control Center-Houston. The Mission Control Center's responsibility includes ascent, on-orbit operations, entry, approach and landing until landing runout completion, at which time the orbiter is handed over to the postlanding operations at the landing site for turnaround and re-launch. At the launch site the SRBs and external tank are processed for launch and the SRBs are recycled for reuse.
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"Orbital Maneuvering System and Reaction Control System (Modifications)",230,0,0,0
The 64 valves operated by AC-motors in the OMS and RCS were modified to incorporate a "sniff" line for each valve to permit monitoring of \Jnitrogen\j tetroxide or monomethyl hydrazine in the electrical portion of the valves during ground operations. This new line reduces the probability of floating particles in the electrical microswitch portion of each valve, which could affect the operation of the microswitch position indicators for onboard displays and telemetry. It also reduces the probability of \Jnitrogen\j tetroxide or monomethyl hydrazine leakage into the bellows of each ac-motor-operated valve.
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"Orbiter, Fuel Cell Power Plants (Modifications)",231,0,0,0
End-cell heaters on each fuel cell power plant were deleted because of potential electrical failures and replaced with Freon coolant loop passages to maintain uniform temperature throughout the power plants. In addition, the \Jhydrogen\j pump and water separator of each fuel cell power plant were improved to minimize excessive \Jhydrogen\j gas entrained in the power plant product water. A current measurement detector was added to monitor the \Jhydrogen\j pump of each fuel cell power plant and provide an early indication of \Jhydrogen\j pump overload.
The starting and sustaining heater system for each fuel cell power plant was modified to prevent overheating and loss of heater elements. A stack inlet temperature measurement was added to each fuel cell power plant for full visibility of thermal conditions.
The product water from all three fuel cell power plants flows to a single water relief control panel. The water can be directed from the single panel to the Environmental Control and Life Support System (ECLSS) potable water tank A or to the fuel cell power plant water relief nozzle. Normally, the water is directed to water tank A. In the event of a line rupture in the vicinity of the single water relief panel, water could spray on all three water relief panel lines causing them to freeze and preventing water discharge.
The product water lines from all three fuel cell power plants were modified to incorporate a parallel (redundant) path of product water to ECLSS potable water tank B in the event of a freeze-up in the single water relief panel. If the single water relief panel freezes up, pressure would build up and discharge through the redundant paths to water tank B.
A water purity sensor (pH) was added at the common product water outlet of the water relief panel to provide a redundant measurement of water purity (a single measurement of water purity in each fuel cell power plant was provided previously). If the fuel cell power plant \JpH\j sensor failed in the past, the flight crew had to sample the potable water.
#
"Orbiter, Auxiliary Power Units (Modifications)",232,0,0,0
The APUs that have been in use to date have a limited life. Each unit was refurbished after 25 hours of operation because of cracks in the \Jturbine\j housing, degradation of the gas generator catalyst (which varied up to approximately 30 hours of operation) and operation of the gas generator valve module (which also varied up to approximately 30 hours of operation). The remaining parts of the APU were qualified for 40 hours of operation.
Improved APUs are scheduled for delivery in late 1988. A new \Jturbine\j housing increases the life of the housing to 75 hours of operation (50 missions); a new gas generator increases its life to 75 hours; a new standoff design of the gas generator valve module and fuel pump deletes the requirement for a water spray system that was required previously for each APU upon shutdown after the first OMS thrusting period or orbital checkout; and the addition of a third seal in the middle of the two existing seals for the shaft of the fuel pump/lube oil system (previously only two seals were located on the shaft, one on the fuel pump side and one on the gearbox lube oil side) reduces the probability of hydrazine leaking into the lube oil system.
The deletion of the water spray system for the gas generator valve module and fuel pump for each APU results in a weight reduction of approximately 150 pounds for each orbiter. Upon the delivery of the improved units, the life-limited APUs will be refurbished to the upgraded design.
In the even that a fuel tank valve switch in an auxiliary power unit is inadvertently left on or an electrical short occurs within the valve electrical coil, additional protection is provided to prevent overheating of the fuel isolation valves.
#
"Orbiter, Main Landing Gear (Modifications)",233,0,0,0
The following modifications were made to improve the performance of the main landing gear elements:
The thickness of the main landing gear axle was increased to provide a stiffer configuration that reduces brake-to-axle deflections and precludes brake damage experienced in previous landings. The thicker axle should also minimize tire wear.
Orifices were added to hydraulic passages in the brake's piston housing to prevent pressure surges and brake damage caused by a wobble/pump effect.
The electronic brake control boxes were modified to balance hydraulic pressure between adjacent brakes and equalize energy applications. The anti-skid circuitry previously used to reduce brake pressure to the opposite wheel if a flat tire was detected has now been removed.
The carbon-lined \Jberyllium\j stator discs in each main landing gear brake were replaced with thicker discs to increase braking energy significantly.
A long-term structural carbon brake program is in progress to replace the carbon-lined \Jberyllium\j stator discs with a carbon configuration that provides higher braking capacity by increasing maximum energy absorption.
Strain gauges were added to each nose and main landing gear wheel to monitor tire pressure before launch, deorbit and landing.
Other studies involve arresting barriers at the end of landing site runways (except lakebed runways), installing a skid on the landing gear that could preclude the potential for a second blown tire on the same gear after the first tire has blown, providing "roll on rim" for a predictable roll if both tires are lost on a single or multiple gear and adding a drag chute.
Studies of landing gear tire improvements are being conducted to determine how best to decrease tire wear observed after previous KSC landings and how to improve crosswind landing capability.
Modifications were made to the KSC Shuttle Landing Facility runway. The full 300-foot width of 3,500-foot sections at both ends of the runway were ground to smooth the runway surface texture and remove cross grooves. The modified corduroy ridges are smaller than those they replaced and run the length of the runway rather than across its width. The existing landing zone light fixtures were also modified, and the markings of the entire runway and overruns were repainted. The primary purpose of the modifications is to enhance safety by reducing tire wear during landing.
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"Orbiter, Nose Wheel Steering (Modifications)",234,0,0,0
The nose wheel steering system was modified on Columbia (OV-102) for the 61-C mission, and Discovery (OV-103) and Atlantis (OV-104) are being similarly modified before their return to flight. The modification allows a safe high-speed engagement of the nose wheel steering system and provides positive lateral directional control of the orbiter during rollout in the presence of high crosswinds and blown tires.
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"Orbiter, Thermal Protection System (Modifications)",235,0,0,0
The area aft of the reinforced carbon-carbon nose cap to the nose landing gear doors has sustained damage (tile slumping) during flight operations from impact during ascent and overheating during reentry. This area, which previously was covered with high-temperature reusable surface \Jinsulation\j tiles, will now be covered with reinforced carbon-carbon.
The low-temperature thermal protection system tiles on Columbia's midbody, payload bay doors and vertical tail were replaced with advanced Flexible Reusable Surface \JInsulation\j (FRSI) blankets.
Because of evidence of plasma flow on the lower wing trailing edge and elevon landing edge tiles (wing/elevon cove) at the outboard elevon tip and inboard elevon, the low-temperature tiles are being replaced with Fibrous Refractory Composite \JInsulation\j (FRC1-12) and High-Temperature (HRSI-22) tiles along with gap fillers on Discovery and Atlantis. On Columbia only gap fillers are installed in this area.
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"Orbiter, Wing Modification",236,0,0,0
Before the wings for Discovery and Atlantis were manufactured, a weight reduction program was instituted that resulted in a redesign of certain areas of the wing structure. An assessment of wing air loads from actual flight data indicated greater loads on the wing structure than predicted. To maintain positive margins of safety during ascent, structural modifications were incorporated into certain areas of the wings.
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"Orbiter, Mid-Fuselage Modifications",237,0,0,0
Because of additional detailed analysis of actual flight data concerning descent-stress thermal-gradient loads, torsional straps were added to tie all the lower mid-fuselage stringers in bays 1 through 11 together in a manner similar to a box section. This eliminates rotational (torsional) capabilities to provide positive margins of safety.
Also, because of the detailed analysis of actual descent flight data, room-temperature vulcanizing silicone rubber material was bonded to the lower mid-fuselage from bays 4 through 11 to act as a heat sink, distributing temperatures evenly across the bottom of the mid-fuselage, reducing thermal gradients and ensuring positive margins of safety.
#
"Orbiter, General Purpose Computers (Modifications)",238,0,0,0
New upgraded General Purpose Computers (GPC), \JIBM\j AP-101S, will replace the existing GPCs aboard the Space Shuttle orbiters in late 1988 or early 1989. The upgraded computers allow NASA to incorporate more capabilities into the orbiters and apply advanced computer technologies that were not available when the orbiter was first designed.
The new computer design began in January 1984, whereas the older design began in January 1972. The upgraded GPCs provide two-and-a-half times the existing memory capacity and up to three times the existing processor speed with minimum impact on flight software. They are half the size, weigh approximately half as much, and require less power to operate.
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"Orbiter, Inertial Measurement Units (Modifications)",239,0,0,0
The new High-Accuracy Inertial Navigation System (HAINS) will be phased in in 1988-89 to augment the present KT-70 inertial measurement units . These new Inertial Measurement Units (IMUs) will result in lower program costs over the next decade, ongoing production support, improved performance, lower failure rates and reduced size and weight.
The HAINS IMUs also contain an internal dedicated \Jmicroprocessor\j with memory for processing and storing compensation and scale factor data from the IMU manufacturer's \Jcalibration\j, thereby reducing the need for extensive initial load data for the orbiter's computers. The HAINS is both physically and functionally interchangeable with the KT-70 IMU.
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"Orbiter, Crew Escape System (Modifications)",240,0,0,0
The in-flight crew escape system is provided for use only when the orbiter is in controlled gliding flight and unable to reach a runway. This would normally lead to ditching. The crew escape system provides the flight crew with an alternative to water ditching or to landing on terrain other than a landing site. The probability of the flight crew surviving a ditching is very small.
The hardware changes required to the orbiters would enable the flight crew to equalize the pressurized crew compartment with the outside pressure via a depressurization valve opened by pyrotechnics in the crew compartment aft bulkhead that would be manually activated by a flight crew member in the middeck of the crew compartment; pyrotechnically jettison the crew ingress/ egress side hatch in the middeck of the crew compartment; and bail out from the middeck of the orbiter through the ingress/ egress side hatch opening after manually deploying the escape pole through, outside and down from the side hatch opening.
One by one, each crewmember attaches a lanyard hook assembly, which surrounds the deployed escape pole, to his parachute harness and egresses through the side hatch opening. Attached to the escape pole, the crewmember slides down the pole and off the end. The escape pole provides a trajectory that takes the crew members below the orbiter's left wing.
Changes were also made in the software of the orbiter's general purpose computers. The software changes were required for the primary avionics software system and the backup flight system for transatlantic-landing and glide-return-to-launch-site aborts. The changes provide the orbiter with an automatic-mode input by the flight crew through keyboards on the commander's and/or pilot's panel C3, which provides the orbiter with an automatic stable flight for crew bailout.
The side hatch jettison feature also could be used in a landing emergency.
The emergency egress slide provides orbiter flight crew members with a means for rapid and safe exit through the orbiter middeck ingress/egress side hatch after a normal opening of the side hatch or after jettisoning the side hatch at the nominal end-of-mission landing site or at a remote or emergency landing site.
The emergency egress slide replaces the emergency egress side hatch bar, which required the flight crewmembers to drop approximately 10.5 feet to the ground. The previous arrangement could have injured crewmembers or prevented an already-injured crewmember from evacuating and moving a safe distance from the orbiter.
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"Orbiter, 17-Inch Orbiter/External Tank Disconnects (Modifications)",242,0,0,0
Each mated pair of 17-inch disconnects contains two flapper valves: one on the orbiter side and one on the external tank side. Both valves in each disconnect pair are opened to permit propellant flow between the orbiter and the external tank. Prior to separation from the external tank, both valves in each mated pair of disconnects are commanded closed by pneumatic (helium) pressure from the main propulsion system.
The closure of both valves in each disconnect pair prevents propellant discharge from the external tank or orbiter at external tank separation. Valve closure on the orbiter side of each disconnect also prevents contamination of the orbiter main propulsion system during landing and ground operations.
Inadvertent closure of either valve in a 17-inch disconnect during main engine thrusting would stop propellant flow from the external tank to all three main engines. Catastrophic failure of the main engines and external tank feed lines would result.
To prevent inadvertent closure of the 17-inch disconnect valves during the Space Shuttle main engine thrusting period, a latch mechanism was added in each orbiter half of the disconnect. The latch mechanism provides a mechanical backup to the normal fluid-induced-open forces. The latch is mounted on a shaft in the flowstream so that it overlaps both flappers and obstructs closure for any reason.
In preparation for external tank separation, both valves in each 17-inch disconnect are commanded closed. Pneumatic pressure from the main propulsion system causes the latch actuator to rotate the shaft in each orbiter 17-inch disconnect 90 degrees, thus freeing the flapper valves to close as required for external tank separation.
A backup mechanical separation capability is provided in case a latch pneumatic actuator malfunctions. When the orbiter umbilical initially moves away from the ET umbilical, the mechanical latch disengages from the ET flapper valve and permits the orbiter disconnect flapper to toggle the latch. This action permits both flappers to close.
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"Space Shuttle Main Engine Margin Improvement Program",243,0,0,0
Improvements to the Space Shuttle Main Engines (SSMEs) for increased margin and durability began with a formal Phase II program in 1983.
Phase II focused on turbo-machinery to extend the time between high-pressure turbopump overhauls by reducing the operating temperature in the high-pressure fuel turbopump and by incorporating margin improvements to the High Pressure Fuel Turbopump (HPFT) rotor dynamics (whirl), \Jturbine\j blade and HPFT bearings. Phase II certification was completed in 1985, and all the changes have been incorporated into the SSMEs for the STS-26 mission.
In addition to the Phase II improvements, additional changes in the SSME have been incorporated to further extend the engines' margin and durability. The main changes were to the high-pressure turbo-machinery, main combustion chamber, hydraulic actuators and high-pressure \Jturbine\j discharge temperature sensors. Changes were also made in the controller software to improve engine control.
Minor high-pressure turbo-machinery design changes resulted in margin improvements to the \Jturbine\j blades, thereby extending the operating life of the turbopumps. These changes included applying surface texture to important parts of the fuel \Jturbine\j blades to improve the material properties in the pressure of \Jhydrogen\j and incorporating a damper into the high-pressure oxidizer \Jturbine\j blades to reduce vibration.
Main combustion chamber life has been increased by plating a welded outlet manifold with nickel. Margin improvements have also been made to five hydraulic actuators to preclude a loss in redundancy on the launch pad. Improvements in quality have been incorporated into the servo-component coil design along with modifications to increase margin. To address a temperature sensor in-flight anomaly, the sensor has been redesigned and extensively tested without problems.
To certify the improvements to the SSMEs and demonstrate their reliability through margin (or limit testing), an aggressive ground test program was initiated in December 1986. From December 1986 to December 1987, 151 tests and 52.363 seconds of operation (equivalent to 100 Shuttle missions) were performed. The SSMEs have exceeded 300,000 seconds total test time, the equivalent of 615 Space Shuttle missions. These hot-fire ground tests are performed at the single-engine test stands NASA's Stennis Space Center in Mississippi and at Rockwell International's Rocketdyne Division's Santa Susana Field Laboratory in \JCalifornia\j.
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"SSME Flight Program",244,0,0,0
By January 1986, there have been 25 flights (75 engine launches with three SSMEs per flight) of the SSMEs. A total of 13 engines were flown, and SSME reusability was demonstrated. One engine (serial number 2012) has been flown 10 times; 10 other engines have flown between five and nine times. Two off-nominal conditions were experienced on the launch pad and one during flight.
Two fail-safe shutdowns occurred on the launch pad during engine start but before SRB ignition. In each case, the controller detected a loss of redundancy in the hydraulic actuator system and commanded engine shutdown in keeping with the launch commit criteria.
Another loss of redundancy occurred in flight with a loss of a red-line temperature sensor and its backup. The engine was commanded to shut down, but the other two engines safely delivered the Space Shuttle to orbit. A major upgrade of these components was implemented to prevent a recurrence of these conditions and will be incorporated for STS-26.
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"Solid Rocket Motor Redesign",245,0,0,0
On June 13, 1986, President Reagan directed NASA to implement, as soon as possible, the recommendations of the "Presidential Commission on the Space Shuttle Challenger Accident." NASA developed a plan to provide a Redesigned Solid Rocket Motor (RSRM).
The primary objective of the redesign effort was to provide an SRM that is safe to fly. A secondary objective was to minimize impact on the schedule by using existing hardware, to the extent practical, without compromising safety. A joint redesign team was established that included participation from Marshall Space Flight Center, Morton Thiokol and other NASA centers as well as individuals from outside NASA.
An "SRM Redesign Project Plan" was developed to formalize the methodology for SRM redesign and requalification. The plan provided an overview of the organizational responsibilities and relationships, the design objectives, criteria and process; the verification approach and process; and a master schedule. The companion "Development and Verification Plan" defined the test program and analyses required to verify the redesign and the unchanged components of the SRM.
All aspects of the existing SRM were assessed, and design changes were required in the field joint, case-to-nozzle joint, nozzle, factory joint, propellant grain shape, ignition system and ground support equipment. No changes were made in the propellant, liner or castable inhibitor formulations. Design criteria were established for each component to ensure a safe design with an adequate margin of safety. These criteria focused on loads, environments, performance, redundancy, margins of safety and verification philosophy.
The criteria were converted into specific design requirements during the Preliminary Requirements Reviews held in July and August 1986. The design developed from these requirements was assessed at the Preliminary Design Review held in September 1986 and baselined in October 1986.
The final design was approved at the Critical Design Review held in October 1987. Manufacture of the RSRM test hardware and the first flight hardware began prior to the Preliminary Design Review (PDR) and continued in parallel with the hardware certification program. The Design Certification Review will review the analyses and test results versus the program and design requirements to certify the redesigned SRM is ready to fly.
#
"Shuttle SRM Field Joint Redesigned",246,0,0,0
The SRM (Solid Rocket Motor) field-joint metal parts, internal case \Jinsulation\j and seals were redesigned and a weather protection system was added.
In the STS 51-L design, the application of actuating pressure to the upstream face of the O-ring was essential for proper joint sealing performance because large sealing gaps were created by pressure-induced deflections, compounded by significantly reduced O-ring sealing performance at low temperature.
The major change in the motor case is the new tang capture feature to provide a positive metal-to-metal interference fit around the \Jcircumference\j of the tang and clevis ends of the mating segments. The interference fit limits the deflection between the tang and clevis O-ring sealing surfaces caused by motor pressure and structural loads. The joints are designed so that the seals will not leak under twice the expected structural deflection and rate.
The new design, with the tang capture feature, the interference fit and the use of custom shims between the outer surface of the tang and inner surface of the outer clevis leg, controls the O-ring sealing gap dimension.
The sealing gap and the O-ring seals are designed so that a positive compression (squeeze) is always on the O-rings. The minimum and maximum squeeze requirements include the effects of temperature, O-ring resiliency and compression set, and pressure. The clevis O-ring groove dimension has been increased so that the O-ring never fills more than 90 percent of the O-ring groove and pressure actuation is enhanced.
The new field joint design also includes a new O-ring in the capture feature and an additional leak check port to ensure that the primary O-ring is positioned in the proper sealing direction at ignition. This new or third O-ring also serves as a thermal barrier in case the sealed \Jinsulation\j is breached.
The field joint internal case \Jinsulation\j was modified to be sealed with a pressure-actuated flap called a J-seal, rather than with putty as in the STS 51-L configuration.
Longer field-joint-case mating pins, with a reconfigured retainer band, were added to improve the shear strength of the pins and increase the metal parts' joint margin of safety. The joint safety margins, both thermal and structural, are being demonstrated over the full ranges of ambient temperature, storage compression, grease effect, assembly stresses and other environments. External heaters with integral weather seals were incorporated to maintain the joint and O-ring temperature at a minimum of 75 F. The weather seal also prevents water intrusion into the joint.
The SRM case-to nozzle joint, which experienced several instances of O-ring erosion in flight, has been redesigned to satisfy the same requirements imposed upon the case field joint. Similar to the field joint, cast-to-nozzle joint modifications have been made in the metal parts, internal \Jinsulation\j and O-rings. Radial bolts with Stato-O-Seals were added to minimize the joint sealing gap opening.
The internal \Jinsulation\j was modified to be sealed adhesively, and third O-ring was included. The third O-ring serves as a dam or wiper in front of the primary O-ring to prevent the polysulfide adhesive from being extruded into the primary O-ring groove. It also serves as a thermal barrier in case the polysulfide adhesive is breached. The polysulfide adhesive replaces the putty used in the 51-L joint. Also, an additional leak check port was added to reduce the amount of trapped air in the joint during the nozzle installation process and to aid in the leak check procedure.
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"Shuttle Nozzle Redesigned",248,0,0,0
The internal joints of the nozzle metal parts have been redesigned to incorporate redundant and verifiable O-rings at each joint. The nozzle steel fixed housing part has been redesigned to permit the incorporation of the 100 radial bolts that attach the fixed housing to the case's aft dome. Improved bonding techniques are being used for the nozzle nose inlet, cowl/boot and aft exit cone assemblies.
The distortion of the nose inlet assembly's metal-part-to-ablative-parts bond line has been eliminated by increasing the thickness of the aluminum nose inlet housing and improving the bonding process. The tape-wrap angle of the carbon cloth fabric in the areas of the nose inlet and throat assembly parts was changed to improve the ablative \Jinsulation\j erosion tolerance. Some of these ply-angle changes were in progress prior to STS 51-L. The cowl and outer boot ring has additional structural support with increased thickness and contour changes to increase their margins of safety. Additionally, the outer boot ring ply configuration was altered.
#
"Shuttle Factory Joint Modified",249,0,0,0
Minor modifications were made in the case factory joints by increasing the \Jinsulation\j thickness and lay-up to increase the margin of safety on the internal \Jinsulation\j. Longer pins were also added, along wit a reconfigured retainer band and new weather seal to improve factory joint performance and increase the margin of safety. Additionally, the O-ring and O-ring groove size was changed to be consistent with the field joint.
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"Shuttle Solid Rocket Motor Ignition System Modified",250,0,0,0
Several minor modifications were incorporated into the ignition system. The aft end of the igniter steel case, which contains the igniter nozzle insert, was thickened to eliminate a localized weakness. The igniter internal case \Jinsulation\j was tapered to improve the manufacturing process. Finally, although vacuum putty is still being used at the joint of the igniter and case forward dome, it was changed to eliminate \Jasbestos\j as one of its constituents.
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"Shuttle Ground Support Equipment Redesigned",251,0,0,0
The ground support equipment has been redesigned to (1) minimize the case distortion during handling at the launch site; (2) improve the segment tang and clevis joint measurement system for more accurate reading of case diameters to facilitate stacking; (3) minimize the risk of O-ring damage during joint mating; and (4) improve leak testing of the igniter, case and nozzle field joints.
A Ground Support Equipment (GSE) assembly aid guides the segment tang into the clevis and rounds the two parts with each other. Other GSE modifications include transportation monitoring equipment and lifting beam.
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"Shuttle Design Analysis Summary",252,0,0,0
Improved, state-of-the-art, analyses related to structural strength, loads, stress, dynamics, fracture mechanics, gas and thermal dynamics, and material characterization and behavior were performed to aid the field joint, nozzle-to-case joint and other designs. Continuing these analyses will ensure that the design integrity and system compatibility adhere to design requirements and operational use. These analyses will be verified by tests, whose results will be correlated with pre-test predictions.
The verification program demonstrates that the RSRM meets all design and performance requirements, and that failure modes and hazards have been eliminated or controlled. The verification program encompasses the following program phases: development, certification, acceptance, preflight checkout, flight and postflight.
Redesigned SRM (Solid Rocket Motor) certification is based on formally documented results of development motor tests; qualification motor tests and other tests and analyses. The certification tests are conducted under strict control of environments, including thermal and structural loads; assembly, inspection and test procedures; and safety, reliability, maintainability and quality assurance surveillance to verify that flight hardware meets the specified performance and design requirements. The "Development and Verification Plan" stipulates the test program, which follows a rigorous sequence wherein successive tests build on the results of previous tests leading to formal certification.
The test activities include laboratory and component tests, subscale tests, full-scale simulation and full-scale motor static test firings. Laboratory and component tests are used to determine component properties and characteristics. Subscale motor firings are used to simulate gas dynamics and thermal conditions for components and subsystem design.
Full-scale hardware simulators are used to verify analytical models; determine hardware assembly characteristics; determine joint deflection characteristics; determine joint performance under short-duration hot-gas tests, including joint flaws and flight loads; and determine redesigned hardware structural characteristics.
Fourteen full-scale joint assembly demonstration vertical mate/demate tests, with eight interspersed hydro tests to simulate flight hardware refurbishment procedures, were completed early for the redesigned capture-feature hardware. Assembly loads were as expected, and the case growth was as predicted with no measurable increase after three hydro-proof tests.
Flight-configuration aft and center segments were fabricated, loaded with live propellant, and used for assembly test article stacking demonstration tests at Kennedy Space Center. These tests were pathfinder demonstrations for the assembly of flight hardware using newly developed ground support equipment.
In a long-term stack test, a full-scale casting segment, with live propellant, has been mated vertically with a J-seal \Jinsulation\j segment and is undergoing temperature \Jcycling\j. This will determine the compression set of the J-seal, aging effects and long-term propellant slumping effects.
The Structural Test Article (STA-3), consisting of flight-type forward and aft motor segments and forward and aft skirts, was subjected to extensive static and dynamic structural testing, including maximum prelaunch, liftoff and flight (maximum dynamic pressure) structural loads.
Redesigned SRM certification includes testing the actual flight configuration over the full range of operating environments and conditions. The joint environment simulator, transient pressure test article, and the nozzle joint environment simulator test programs all utilize full-scale flight design hardware and subject the RSRM design features to the maximum expected operating pressure, maximum pressure rise rate and temperature extremes during ignition tests. Additionally, the Transient Pressure Test Article (TPTA) is subjected to ignition and liftoff loads as well as maximum dynamic pressure structural loads.
Four TPTA tests have been completed to subject the redesigned case field and case-to-nozzle joints to the above-described conditions. The field and case-to-nozzle joints were temperature-conditioned to 75 F. and contained various types of flaws in the joints so that the primary and secondary O-rings could be pressure-actuated, joint rotation and O-ring performance could be evaluated and the redesigned joints could be demonstrated as fail safe.
Six of the seven Joint Environment Simulators (JES) tests have been completed. The JES test program initially used the STS 51-L configuration hardware to evaluate the joint performance with prefabricated blowholes through the putty. The JES-1 test series, which consisted of two tests, established a structural and performance \Jdatabase\j for the STS 51-L configuration with and without a replicated joint failure.
The JES-2 series, two tests, also used the STS 51-L case metal-part joint but with a bonded labyrinth and U-seal \Jinsulation\j that was an early design variation of the J-seal. Tests were conducted with and without flaws built into the U-seal joint \Jinsulation\j; neither joint showed O-ring erosion or blow-by. The JES-3 series, three tests, uses almost exact flight configuration hardware, case field-joint capture feature with interference fit and J-seal \Jinsulation\j.
Four of five nozzle JES tests have been successfully conducted. The STS 51-L hardware configuration hydro test confirmed predicted case-to-nozzle-joint deflection. The other three tests used the radially bolted RSRM configuration.
Seven full-scale, full-duration motor static tests are being conducted to verify the integrated RSRM performance. These include one \Jengineering\j test motor used to (1) provide a data base for STS 51-L-type field joints; (2) evaluate new seal material; (3) evaluate the ply-angle change in the nozzle parts,; (4) evaluate the effectiveness of \Jgraphite\j composite stiffener rings to reduce joint rotation; and (5) evaluate field-joint heaters.
There were two development motor tests and three qualification motor tests for final flight configuration and performance certification. There will be one flight Production Verification Motor that contains intentionally induced defects in the joints to demonstrate joint performance under extreme worse case conditions. The QM-7 and QM-8 motors were subjected to liftoff and maximum dynamic pressure structural loads, QM-7 was temperature-conditioned to 90 F., and QM-8 was temperature-conditioned to 40 F.
An assessment was conducted to determine the full-duration static firing test attitude necessary to certify the design changes completely. The assessment included establishing test objectives, defining and quantifying attitude-sensitive parameters, and evaluating attitude options. Both horizontal and vertical (nozzle up and down) test attitudes were assessed.
In all three options, consideration was given to testing with and without externally applied loads. This assessment determined that the conditions influencing the joint and \Jinsulation\j behavior could best be tested to design extremes in the horizontal attitude. In conjunction with the horizontal attitude for the RSRM full-scale testing, it was decided to incorporate externally applied loads.
A second horizontal test stand for certification of the RSRM was constructed at Morton Thiokol. This new stand, designated as the T-97 Large Motor Static Test Facility, is being used to simulate environmental stresses, loads and temperatures experienced during an actual Shuttle launch and ascent. The new test stand also provides redundancy for the existing stand.
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"Shuttle, Non-Destructive Evaluation",254,0,0,0
The Shuttle 51-L and Titan 34D-9 vehicle failures, both of which occurred in 1986, resulted in major reassessments of each vehicle's design, processing, inspection and operations. While the Shuttle SRM insulation/ propellant integrity was not implicated in the 51-L failure, the intent is to preclude a failure similar to that experienced by Titan.
The RSRM field joint is quite tolerant of unbonded \Jinsulation\j. It has sealed \Jinsulation\j to prevent hot combustion products from reaching the insulation-to-case bond line. The bonding processes have been improved to reduce contamination potential, and the new \Jgeometry\j of the tang capture feature inherently provides more isolation of the edge \Jinsulation\j area from contaminating agents.
A greatly enhanced Non-Destructive Evaluation program for the RSRM has been incorporated. The enhanced non-destructive testing includes ultrasonic inspection and mechanical testing of propellant and \Jinsulation\j bonded surfaces. All segments will again be X-rayed for the first flight and near-term subsequent flights.
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"Shuttle, Contingency Planning",255,0,0,0
To provide additional program confidence, both near- and long-term contingency planning was implemented. Alternative designs, which might be incorporated into the flight program at discrete decision points, include field-joint graphite-composite overwrap bands and alternative seals for the field joint and case-to-nozzle joint. Alternative designs for the nozzle include a different composite lay-up technique and a steel nose inlet housing.
Alternative designs with long-lead-time implications were also developed. These designs focus on the field joint and cast-to-nozzle joint. Since fabrication of the large steel components dictates the schedule, long-lead procurement of maximum-size steel ingots was initiated. This allowed machining of case joints to either the new baseline or to an alternative design configuration. Ingot processing continued through forging and heat treating. At that time, the final design was selected. A principal consideration in this configuration decision was the result of verification testing on the baseline configuration.
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"NASA, Independent Oversight Panel",256,0,0,0
As recommended in the "Presidential Commission Report" and at the request of the NASA administrator, the National Research Council established an Independent Oversight Panel chaired by Dr. H. Guyford Stever, who reports directly to the NASA Administrator.
Initially, the panel was given introductory briefings on the Shuttle system requirements, implementation and control, the original design and manufacturing of the SRM, Mission 51-L accident analyses and preliminary plans for the redesign. The panel has met with major SRM manufacturers and vendors, and has visited some of their facilities.
The panel frequently reviewed the RSRM design criteria, \Jengineering\j analyses and design, and certification program planning. Panel members continuously review the design and testing for safe operation, selection and specifications for material, and quality assurance and control. The panel has continued to review the design as it progresses through certification and review the manufacturing and assembly of the first flight RSRM.
Panel members have participated in major program milestones, project requirements review, and preliminary design review; they also will participate in future review. Six written reports have been provided by the panel to the NASA administrator.
In addition to the NRC, the redesign team has a design review group of 12 expert senior engineers from NASA and the aerospace industry. They have advised on major program decisions and serve as a "sounding board" for the program.
Additionally, NASA requested the four other major SRM companies -- Aerojet Strategic Propulsion Co., Atlantic Research Corp., Hercules Inc. and United Technologies Corp.'s Chemical Systems Division -- to participate in the redesign efforts by critiquing the design approach and providing experience on alternative design approaches.
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"Solid Rocket Boosters (SRBs)",257,0,0,0
The two SRBs provide the main thrust to lift the space shuttle off the pad and up to an altitude of about 150,000 feet, or 24 nautical miles (28 statute miles). In addition, the two SRBs carry the entire weight of the external tank and orbiter and transmit the weight load through their structure to the mobile launcher platform. Each booster has a thrust (sea level) of approximately 3,300,000 pounds at launch.
They are ignited after the three space shuttle main engines' thrust level is verified. The two SRBs provide 71.4 percent of the thrust at lift- off and during first-stage ascent. Seventy- five seconds after SRB separation, SRB apogee occurs at an altitude of approximately 220,000 feet, or 35 nautical miles (41 statute miles). SRB impact occurs in the ocean approximately 122 nautical miles (141 statute miles) downrange.
The SRBs are the largest solid- propellant motors ever flown and the first designed for reuse. Each is 149.16 feet long and 12.17 feet in diameter.
Each SRB weighs approximately 1,300,000 pounds at launch. The propellant for each solid rocket motor weighs approximately 1,100,000 pounds. The inert weight of each SRB is approximately 192,000 pounds.
Primary elements of each booster are the motor (including case, propellant, igniter and nozzle), structure, separation systems, operational flight instrumentation, recovery avionics, pyrotechnics, deceleration system, thrust vector control system and range safety destruct system.
Each booster is attached to the external tank at the SRB's aft frame by two lateral sway braces and a diagonal attachment. The forward end of each SRB is attached to the external tank at the forward end of the SRB's forward skirt. On the launch pad, each booster also is attached to the mobile launcher platform at the aft skirt by four bolts and nuts that are severed by small explosives at lift-off.
During the downtime following the Challenger accident, detailed structural analyses were performed on critical structural elements of the SRB. Analyses were primarily focused in areas where anomalies had been noted during postflight inspection of recovered hardware.
One of the areas was the attach ring where the SRBs are connected to the external tank. Areas of distress were noted in some of the fasteners where the ring attaches to the SRB motor case. This situation was attributed to the high loads encountered during water impact. To correct the situation and ensure higher strength margins during ascent, the attach ring was redesigned to encircle the motor case completely (360 degrees). Previously, the attach ring formed a C and encircled the motor case 270 degrees.
Additionally, special structural tests were performed on the aft skirt. During this test program, an anomaly occurred in a critical weld between the hold-down post and skin of the skirt. A redesign was implemented to add reinforcement brackets and fittings in the aft ring of the skirt.
These two modifications added approximately 450 pounds to the weight of each SRB.
The propellant mixture in each SRB motor consists of an ammonium perchlorate (oxidizer, 69.6 percent by weight), aluminum (fuel, 16 percent), iron oxide (a catalyst, 0.4 percent), a polymer (a binder that holds the mixture together, 12.04 percent), and an epoxy curing agent (1.96 percent). The propellant is an 11-point star- shaped perforation in the forward motor segment and a double- truncated- cone perforation in each of the aft segments and aft closure. This configuration provides high thrust at ignition and then reduces the thrust by approximately a third 50 seconds after lift-off to prevent overstressing the vehicle during maximum dynamic pressure.
The SRBs are used as matched pairs and each is made up of four solid rocket motor segments. The pairs are matched by loading each of the four motor segments in pairs from the same batches of propellant ingredients to minimize any thrust imbalance. The segmented-casing design assures maximum flexibility in fabrication and ease of transportation and handling. Each segment is shipped to the launch site on a heavy- duty rail car with a specially built cover.
The nozzle expansion ratio of each booster beginning with the STS-8 mission is 7-to-79. The nozzle is gimbaled for thrust vector (direction) control. Each SRB has its own redundant auxiliary power units and hydraulic pumps. The all-axis gimbaling capability is 8 degrees. Each nozzle has a carbon cloth liner that erodes and chars during firing. The nozzle is a convergent- divergent, movable design in which an aft pivot- point flexible bearing is the gimbal mechanism.
The cone- shaped aft skirt reacts the aft loads between the SRB and the mobile launcher platform. The four aft separation motors are mounted on the skirt. The aft section contains avionics, a thrust vector control system that consists of two auxiliary power units and hydraulic pumps, hydraulic systems and a nozzle extension jettison system.
The forward section of each booster contains avionics, a sequencer, forward separation motors, a nose cone separation system, drogue and main parachutes, a recovery beacon, a recovery light, a parachute camera on selected flights and a range safety system.
Each SRB has two integrated electronic assemblies, one forward and one aft. After burnout, the forward assembly initiates the release of the nose cap and frustum and turns on the recovery aids. The aft assembly, mounted in the external tank/SRB attach ring, connects with the forward assembly and the orbiter avionics systems for SRB ignition commands and nozzle thrust vector control. Each integrated electronic assembly has a multiplexer/ demultiplexer, which sends or receives more than one message, signal or unit of information on a single communication channel.
Eight booster separation motors (four in the nose frustum and four in the aft skirt) of each SRB thrust for 1.02 seconds at SRB separation from the external tank. Each solid rocket separation motor is 31.1 inches long and 12.8 inches in diameter.
Location aids are provided for each SRB, frustum/ drogue chutes and main parachutes. These include a transmitter, antenna, strobe/ converter, battery and salt water switch \Jelectronics\j. The location aids are designed for a minimum operating life of 72 hours and when refurbished are considered usable up to 20 times. The flashing light is an exception. It has an operating life of 280 hours. The battery is used only once.
The SRB nose caps and nozzle extensions are not recovered.
The recovery crew retrieves the SRBs, frustum/ drogue chutes, and main parachutes. The nozzles are plugged, the solid rocket motors are dewatered, and the SRBs are towed back to the launch site. Each booster is removed from the water, and its components are disassembled and washed with fresh and deionized water to limit salt water \Jcorrosion\j. The motor segments, igniter and nozzle are shipped back to Thiokol for refurbishment.
Each SRB incorporates a range safety system that includes a battery power source, receiver/ decoder, antennas and ordnance.
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"SRB Hold-Down Posts",258,0,0,0
Each solid rocket booster has four hold- down posts that fit into corresponding support posts on the mobile launcher platform. Hold- down bolts hold the SRB and launcher platform posts together. Each bolt has a nut at each end, but only the top nut is frangible. The top nut contains two NASA standard detonators, which are ignited at solid rocket motor ignition commands.
When the two NSDs are ignited at each hold- down, the hold- down bolt travels downward because of the release of tension in the bolt (pretensioned before launch), NSD gas pressure and gravity. The bolt is stopped by the stud deceleration stand, which contains sand. The SRB bolt is 28 inches long and is 3.5 inches in diameter. The frangible nut is captured in a blast container.
The solid rocket motor ignition commands are issued by the orbiter's computers through the master events controllers to the hold- down pyrotechnic initiator controllers on the mobile launcher platform. They provide the ignition to the hold- down NSDs. The launch processing system monitors the SRB hold- down PICs for low voltage during the last 16 seconds before launch. PIC low voltage will initiate a launch hold.
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"SRB Ignition",259,0,0,0
SRB (Solid Rocket Booster) ignition can occur only when a manual lock pin from each SRB safe and arm device has been removed. The ground crew removes the pin during prelaunch activities. At T minus five minutes, the SRB safe and arm device is rotated to the arm position. The solid rocket motor ignition commands are issued when the three SSMEs are at or above 90-percent rated thrust, no SSME fail and/or SRB ignition PIC low voltage is indicated and there are no holds from the LPS.
The solid rocket motor ignition commands are sent by the orbiter computers through the MECs to the safe and arm device NSDs in each SRB. A PIC single-channel capacitor discharge device controls the firing of each pyrotechnic device.
Three signals must be present simultaneously for the PIC to generate the pyro firing output. These signals- arm, fire 1 and fire 2-originate in the orbiter general- purpose computers and are transmitted to the MECs. The MECs reformat them to 28-volt dc signals for the PICs. The arm signal charges the PIC capacitor to 40 volts dc (minimum of 20 volts dc).
The fire 2 commands cause the redundant NSDs to fire through a thin barrier seal down a flame tunnel. This ignites a pyro booster charge, which is retained in the safe and arm device behind a perforated plate. The booster charge ignites the propellant in the igniter initiator; and combustion products of this propellant ignite the solid rocket motor initiator, which fires down the length of the solid rocket motor igniting the solid rocket motor propellant.
The GPC launch sequence also controls certain critical main propulsion system valves and monitors the engine- ready indications from the SSMEs. The MPS start commands are issued by the onboard computers at T minus 6.6 seconds (staggered start- engine three, engine two, engine one- all approximately within 0.25 of a second), and the sequence monitors the thrust buildup of each engine. All three SSMEs must reach the required 90-percent thrust within three seconds; otherwise, an orderly shutdown is commanded and safing functions are initiated.
Normal thrust buildup to the required 90-percent thrust level will result in the SSMEs being commanded to the lift- off position at T minus three seconds as well as the fire 1 command being issued to arm the SRBs. At T minus three seconds, the vehicle base bending load modes are allowed to initialize (movement of approximately 25.5 inches measured at the tip of the external tank, with movement towards the external tank).
At T minus zero, the two SRBs are ignited, under command of the four onboard computers; separation of the four explosive bolts on each SRB is initiated (each bolt is 28 inches long and 3.5 inches in diameter); the two T-0 umbilicals (one on each side of the spacecraft) are retracted; the onboard master timing unit, event timer and mission event timers are started; the three SSMEs are at 100 percent; and the ground launch sequence is terminated.
The solid rocket motor thrust profile is tailored to reduce thrust during the maximum dynamic pressure region.
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"SRB Electrical Power Distribution",260,0,0,0
Electrical power distribution in each SRB (Solid Rocket Booster) consists of orbiter- supplied main dc bus power to each SRB via SRB buses A, B and C. Orbiter main dc buses A, B and C supply main dc bus power to corresponding SRB buses A, B and C. In addition, orbiter main dc bus C supplies backup power to SRB buses A and B, and orbiter bus B supplies backup power to SRB bus C. This electrical power distribution arrangement allows all SRB buses to remain powered in the event one orbiter main bus fails.
The nominal dc voltage is 28 volts dc, with an upper limit of 32 volts dc and a lower limit of 24 volts dc.
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"SRB Hydraulic Power Units",261,0,0,0
There are two self- contained, independent HPUs on each SRB . Each HPU consists of an auxiliary power unit, fuel supply module, hydraulic pump, hydraulic reservoir and hydraulic fluid manifold assembly. The APUs are fueled by hydrazine and generate mechanical shaft power to a hydraulic pump that produces hydraulic pressure for the SRB hydraulic system.
The two separate HPUs and two hydraulic systems are located on the aft end of each SRB between the SRB nozzle and aft skirt. The HPU components are mounted on the aft skirt between the rock and tilt actuators. The two systems operate from T minus 28 seconds until SRB separation from the orbiter and external tank. The two independent hydraulic systems are connected to the rock and tilt servoactuators.
The APU controller \Jelectronics\j are located in the SRB aft integrated electronic assemblies on the aft external tank attach rings.
The APUs and their fuel systems are isolated from each other. Each fuel supply module (tank) contains 22 pounds of hydrazine. The fuel tank is pressurized with gaseous \Jnitrogen\j at 400 psi, which provides the force to expel (positive expulsion) the fuel from the tank to the fuel distribution line, maintaining a positive fuel supply to the APU throughout its operation.
The fuel isolation valve is opened at APU startup to allow fuel to flow to the APU fuel pump and control valves and then to the gas generator. The gas generator's catalytic action decomposes the fuel and creates a hot gas. It feeds the hot gas exhaust product to the APU two- stage gas \Jturbine\j. Fuel flows primarily through the startup bypass line until the APU speed is such that the fuel pump outlet pressure is greater than the bypass line's. Then all the fuel is supplied to the fuel pump.
The APU \Jturbine\j assembly provides mechanical power to the APU gearbox. The gearbox drives the APU fuel pump, hydraulic pump and lube oil pump. The APU lube oil pump lubricates the gearbox. The \Jturbine\j exhaust of each APU flows over the exterior of the gas generator, cooling it, and is then directed overboard through an exhaust duct.
When the APU speed reaches 100 percent, the APU primary control valve closes, and the APU speed is controlled by the APU controller \Jelectronics\j. If the primary control valve logic fails to the open state, the secondary control valve assumes control of the APU at 112-percent speed.
Each HPU on an SRB is connected to both servoactuators on that SRB. One HPU serves as the primary hydraulic source for the servoactuator, and the other HPU serves as the secondary \Jhydraulics\j for the servoactuator. Each servoactuator has a switching valve that allows the secondary \Jhydraulics\j to power the actuator if the primary hydraulic pressure drops below 2,050 psi. A switch contact on the switching valve will close when the valve is in the secondary position. When the valve is closed, a signal is sent to the APU controller that inhibits the 100-percent APU speed control logic and enables the 112-percent APU speed control logic. The 100-percent APU speed enables one APU/HPU to supply sufficient operating hydraulic pressure to both servoactuators of that SRB.
The APU 100-percent speed corresponds to 72,000 rpm, 110-percent to 79,200 rpm, and 112-percent to 80,640 rpm.
The hydraulic pump speed is 3,600 rpm and supplies hydraulic pressure of 3,050, plus or minus 50, psi. A high- pressure relief valve provides overpressure protection to the hydraulic system and relieves at 3,750 psi.
The APUs/HPUs and hydraulic systems are reusable for 20 missions.
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"SRB Thrust Vector Control",262,0,0,0
Each SRB has two hydraulic gimbal servoactuators: one for rock and one for tilt. The servoactuators provide the force and control to gimbal the nozzle for thrust vector control.
The space shuttle ascent thrust vector control portion of the flight control system directs the thrust of the three shuttle main engines and the two SRB nozzles to control shuttle attitude and trajectory during lift- off and ascent. Commands from the guidance system are transmitted to the ATVC drivers, which transmit signals proportional to the commands to each servoactuator of the main engines and SRBs.
Four independent flight control system channels and four ATVC channels control six main engine and four SRB ATVC drivers, with each driver controlling one hydraulic port on each main and SRB servoactuator.
Each SRB servoactuator consists of four independent, two- stage servovalves that receive signals from the drivers. Each servovalve controls one power spool in each actuator, which positions an actuator ram and the nozzle to control the direction of thrust.
The four servovalves in each actuator provide a force- summed majority voting arrangement to position the power spool. With four identical commands to the four servovalves, the actuator force-sum action prevents a single erroneous command from affecting power ram motion. If the erroneous command persists for more than a predetermined time, differential pressure sensing activates a selector valve to isolate and remove the defective servovalve hydraulic pressure, permitting the remaining channels and servovalves to control the actuator ram spool.
Failure monitors are provided for each channel to indicate which channel has been bypassed. An isolation valve on each channel provides the capability of resetting a failed or bypassed channel.
Each actuator ram is equipped with transducers for position feedback to the thrust vector control system. Within each servoactuator ram is a splashdown load relief assembly to cushion the nozzle at water splashdown and prevent damage to the nozzle flexible bearing.
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"SRB Rate Gyro Assemblies",263,0,0,0
Each SRB contains two RGAs, with each RGA containing one pitch and one yaw gyro. These provide an output proportional to angular rates about the pitch and yaw axes to the orbiter computers and guidance, navigation and control system during first- stage ascent flight in conjunction with the orbiter roll rate gyros until SRB separation. At SRB separation, a switchover is made from the SRB RGAs to the orbiter RGAs.
The SRB RGA rates pass through the orbiter flight aft multiplexers/ demultiplexers to the orbiter GPCs. The RGA rates are then mid-value- selected in redundancy management to provide SRB pitch and yaw rates to the user software. The RGAs are designed for 20 missions.
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"SRB Separation",264,0,0,0
SRB (Solid Rocket Booster) separation is initiated when the three solid rocket motor chamber pressure transducers are processed in the redundancy management middle value select and the head- end chamber pressure of both SRBs is less than or equal to 50 psi. A backup cue is the time elapsed from booster ignition.
The separation sequence is initiated, commanding the thrust vector control actuators to the null position and putting the main propulsion system into a second-stage configuration (0.8 second from sequence initialization), which ensures the thrust of each SRB is less than 100,000 pounds. Orbiter yaw attitude is held for four seconds, and SRB thrust drops to less than 60,000 pounds.
The SRBs separate from the external tank within 30 milliseconds of the ordnance firing command.
The forward attachment point consists of a ball (SRB) and socket (ET) held together by one bolt. The bolt contains one NSD pressure cartridge at each end. The forward attachment point also carries the range safety system cross-strap wiring connecting each SRB RSS and the ET RSS with each other.
The aft attachment points consist of three separate struts: upper, diagonal and lower. Each strut contains one bolt with an NSD pressure cartridge at each end. The upper strut also carries the umbilical interface between its SRB and the external tank and on to the orbiter.
There are four booster separation motors on each end of each SRB. The BSMs separate the SRBs from the external tank. The solid rocket motors in each cluster of four are ignited by firing redundant NSD pressure cartridges into redundant confined detonating fuse manifolds.
The separation commands issued from the orbiter by the SRB separation sequence initiate the redundant NSD pressure cartridge in each bolt and ignite the BSMs to effect a clean separation.
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"SRB Range Safety System",265,0,0,0
The shuttle vehicle has three RSSs. One is located in each SRB and one in the external tank. Any one or all three are capable of receiving two command messages (arm and fire) transmitted from the ground station. The RSS is used only when the shuttle vehicle violates a launch trajectory red line.
An RSS consists of two antenna couplers, command receivers/ decoders, a dual distributor, a safe and arm device with two NSDs, two confined detonating fuse manifolds, seven CDF assemblies and one linear-shaped charge.
The antenna couplers provide the proper \Jimpedance\j for radio frequency and ground support equipment commands. The command receivers are tuned to RSS command frequencies and provide the input signal to the distributors when an RSS command is sent. The command decoders use a code plug to prevent any command signal other than the proper command signal from getting into the distributors. The distributors contain the logic to supply valid destruct commands to the RSS pyrotechnics.
The NSDs provide the spark to ignite the CDF, which in turn ignites the LSC for shuttle vehicle destruction. The safe and arm device provides mechanical isolation between the NSDs and the CDF before launch and during the SRB separation sequence.
The first message, called arm, allows the onboard logic to enable a destruct and illuminates a light on the flight deck display and control panel at the commander and pilot station. The second message transmitted is the fire command.
The SRB distributors in the SRBs and the ET are cross- strapped together. Thus, if one SRB received an arm or destruct signal, the signal would also be sent to the other SRB and the ET.
Electrical power from the RSS battery in each SRB is routed to RSS system A. The recovery battery in each SRB is used to power RSS system B as well as the recovery system in the SRB. The SRB RSS is powered down during the separation sequence, and the SRB recovery system is powered up. Electrical power for the ET RSS system A and system B is independently supplied by two RSS batteries on the ET.
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"SRB Descent and Recovery",266,0,0,0
The recovery sequence begins with the operation of the high-altitude baroswitch, which triggers the functioning of the pyrotechnic nose cap thrusters. This ejects the nose cap, which deploys the pilot parachute. This occurs at 15,704 feet altitude 225 seconds after separation. The 11.5-foot-diameter conical ribbon pilot parachute provides the force to pull the lanyard activating the zero-second cutter, which cuts the loop securing the drogue retention straps.
This allows the pilot chute to pull the drogue pack from the SRB, causing the drogue suspension lines to deploy from their stored position. At full extension of the 12 95-foot suspension lines, the drogue deployment bag is stripped away from the canopy, and the 54-foot-diameter conical ribbon drogue parachute inflates to its initial reefed condition. The drogue disreefs twice after specified time delays, and it reorients/stabilizes the SRB for main chute deployment. The drogue parachute can withstand a load of 270,000 pounds and weighs approximately 1,200 pounds.
After the drogue chute has stabilized the vehicle in a tailfirst attitude, the frustum is separated from the forward skirt by a charge triggered by the low-altitude baroswitch at an altitude of 5,975 feet 248 seconds after separation. It is then pulled away from the SRB by the drogue chute. The main chutes' suspension lines are pulled out from deployment bags that remain in the frustum.
At full extension of the lines, which are 204 feet long, the three main chutes are pulled from the deployment bags and inflate to their first reefed condition. The frustum and drogue parachute continue on a separate trajectory to splashdown. After specified time delays, the main chutes' reefing lines are cut and the chutes inflate to their second reefed and full open configurations.
The main chute cluster decelerates the SRB to terminal conditions. Each of the 136-foot-diameter, 20-degree conical ribbon parachutes can withstand a load of 180,000 pounds and weighs 2,180 pounds. The nozzle extension is severed by pyrotechnic charge either at apogee or 20 seconds after low baroswitch operation.
Water impact occurs 295 seconds after separation at a velocity of 81 feet per second. The water impact range is approximately 140 miles off the eastern coast of \JFlorida\j. Because the parachutes provide for a nozzlefirst impact, air is trapped in the empty (burned out) motor casing, causing the booster to float with the forward end approximately 30 feet out of the water.
The main chutes are released from the SRB at impact using the parachute release nut ordnance system. Residual loads in the main chutes deploy the parachute attach fittings with the redundant flotation tethered to each fitting. The drogue and frustum; each main chute, with its flotation; and the SRB are buoyant.
The SRB recovery aids are the radio beacon and flashing lights, which become operable at frustum separation. The radio transponder in each SRB has a range of 8.9 nautical miles (10.35 statute miles), and the flashing light has a nighttime range of 4.9 nautical miles (5.75 statute miles).
Various parameters of SRB operation are monitored and displayed on the orbiter flight deck control and display panel and are transmitted to ground telemetry.
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"Shuttle External Tank",267,0,0,0
The external tank contains the liquid \Jhydrogen\j fuel and liquid oxygen oxidizer and supplies them under pressure to the three space shuttle main engines in the orbiter during lift-off and ascent. When the SSMEs are shut down, the ET is jettisoned, enters the Earth's atmosphere, breaks up, and impacts in a remote ocean area. It is not recovered.
The largest and heaviest (when loaded) element of the space shuttle, the ET has three major components: the forward liquid oxygen tank, an unpressurized intertank that contains most of the electrical components, and the aft liquid \Jhydrogen\j tank. The ET is 153.8 feet long and has a diameter of 27.6 feet.
Beginning with the STS-6 mission, a lightweight ET was introduced. Although future tanks may vary slightly, each will weigh approximately 66,000 pounds inert. The last heavyweight tank, flown on STS-7, weighed approximately 77,000 pounds inert. For each pound of weight reduced from the ET, the cargo-carrying capability of the space shuttle \Jspacecraft\j is increased almost one pound. The weight reduction was accomplished by eliminating portions of stringers (structural stiffeners running the length of the \Jhydrogen\j tank), using fewer stiffener rings and by modifying major frames in the \Jhydrogen\j tank.
Also, significant portions of the tank are milled differently to reduce thickness, and the weight of the ET's aft solid rocket booster attachments were reduced by using a stronger, yet lighter and less expensive \Jtitanium\j alloy. Earlier several hundred pounds were eliminated by deleting the anti-geyser line. The line paralleled the oxygen feed line and provided a circulation path for liquid oxygen to reduce accumulation of gaseous oxygen in the feed line while the oxygen tank was being filled before launch.
After propellant-loading data from ground tests and the first few space shuttle missions was assessed, the anti- \Jgeyser\j line was removed for STS-5 and subsequent missions. The total length and diameter of the ET remain unchanged.
The ET is attached to the orbiter at one forward attachment point and two aft points. In the aft attachment area, there are also umbilicals that carry fluids, gases, electrical signals and electrical power between the tank and the orbiter. Electrical signals and controls between the orbiter and the two solid rocket boosters also are routed through those umbilicals.
#
"Shuttle Liquid Oxygen Tank",268,0,0,0
The liquid oxygen tank is an aluminum monocoque structure composed of a fusion-welded assembly of preformed, chem-milled gores, panels, machined fittings and ring chords. It operates in a pressure range of 20 to 22 psig. The tank contains anti-slosh and anti-vortex provisions to minimize liquid residuals and damp fluid motion.
The tank feeds into a 17-inch- diameter feed line that conveys the liquid oxygen through the intertank, then outside the ET to the aft right-hand ET / orbiter disconnect umbilical. The 17-inch-diameter feed line permits liquid oxygen to flow at approximately 2,787 pounds per second with the SSMEs operating at 104 percent or permits a maximum flow of 17,592 gallons per minute.
The liquid oxygen tank's double-wedge nose cone reduces drag and heating, contains the vehicle's ascent air data system (for nine tanks only) and serves as a \Jlightning\j rod. The liquid oxygen tank's volume is 19,563 cubic feet. It is 331 inches in diameter, 592 inches long and weighs 12,000 pounds empty.
#
"Shuttle Intertank",269,0,0,0
The intertank is a steel / aluminum semimonocoque cylindrical structure with flanges on each end for joining the liquid oxygen and liquid \Jhydrogen\j tanks. The intertank houses ET instrumentation components and provides an umbilical plate that interfaces with the ground facility arm for purge gas supply, hazardous gas detection and \Jhydrogen\j gas boiloff during ground operations.
It consists of mechanically joined skin, stringers and machined panels of aluminum alloy. The intertank is vented during flight. The intertank contains the forward SRB-ET attach thrust beam and fittings that distribute the SRB loads to the liquid oxygen and liquid \Jhydrogen\j tanks. The intertank is 270 inches long, 331 inches in diameter and weighs 12,100 pounds.
#
"Shuttle Liquid Hydrogen Tank",270,0,0,0
The liquid \Jhydrogen\j tank is an aluminum semimonocoque structure of fusion-welded barrel sections, five major ring frames, and forward and aft ellipsoidal domes. Its operating pressure range is 32 to 34 psia. The tank contains an anti-vortex baffle and siphon outlet to transmit the liquid \Jhydrogen\j from the tank through a 17-inch line to the left aft umbilical.
The liquid \Jhydrogen\j feed line flow rate is 465 pounds per second with the SSMEs at 104 percent or a maximum flow of 47,365 gallons per minute. At the forward end of the liquid \Jhydrogen\j tank is the ET / orbiter forward attachment pod strut, and at its aft end are the two ET / orbiter aft attachment ball fittings as well as the aft SRB-ET stabilizing strut attachments. The liquid \Jhydrogen\j tank is 331 inches in diameter, 1,160 inches long, and has a volume of 53,518 cubic feet and a dry weight of 29,000 pounds.
#
"Shuttle ET Thermal Protection System",271,0,0,0
The ET thermal protection system consists of sprayed-on \Jfoam\j \Jinsulation\j and premolded ablator materials. The system also includes the use of phenolic thermal insulators to preclude air liquefaction. Thermal isolators are required for liquid \Jhydrogen\j tank attachments to preclude the liquefaction of air-exposed metallic attachments and to reduce heat flow into the liquid \Jhydrogen\j. The thermal protection system weighs 4,823 pounds.
#
"Shuttle ET Hardware",272,0,0,0
The external hardware, ET / orbiter attachment fittings, umbilical fittings, electrical and range safety system weigh 9,100 pounds.
Each propellant tank has a vent and relief valve at its forward end. This dual-function valve can be opened by ground support equipment for the vent function during prelaunch and can open during flight when the ullage (empty space) pressure of the liquid \Jhydrogen\j tank reaches 38 psig or the ullage pressure of the liquid oxygen tank reaches 25 psig.
The liquid oxygen tank contains a separate, pyrotechnically operated, propulsive tumble vent valve at its forward end. At separation, the liquid oxygen tumble vent valve is opened, providing impulse to assist in the separation maneuver and more positive control of the entry aerodynamics of the ET.
There are eight propellant-depletion sensors, four each for fuel and oxidizer. The fuel-depletion sensors are located in the bottom of the fuel tank. The oxidizer sensors are mounted in the orbiter liquid oxygen feed line manifold downstream of the feed line disconnect. During SSME thrusting, the orbiter general-purpose computers constantly compute the instantaneous mass of the vehicle due to the usage of the propellants. Normally, main engine cutoff is based on a predetermined velocity; however, if any two of the fuel or oxidizer sensors sense a dry condition, the engines will be shut down.
The locations of the liquid oxygen sensors allow the maximum amount of oxidizer to be consumed in the engines, while allowing sufficient time to shut down the engines before the oxidizer pumps cavitate (run dry). In addition, 1,100 pounds of liquid \Jhydrogen\j are loaded over and above that required by the 6-1 oxidizer / fuel engine mixture ratio. This assures that MECO from the depletion sensors is fuel-rich; oxidizer-rich engine shutdowns can cause burning and severe erosion of engine components.
Four pressure transducers located at the top of the liquid oxygen and liquid \Jhydrogen\j tanks monitor the ullage pressures.
Each of the two aft external tank umbilical plates mate with a corresponding plate on the orbiter. The plates help maintain alignment among the umbilicals. Physical strength at the umbilical plates is provided by bolting corresponding umbilical plates together. When the orbiter GPCs command external tank separation, the bolts are severed by pyrotechnic devices.
The ET has five propellant umbilical valves that interface with orbiter umbilicals: two for the liquid oxygen tank and three for the liquid \Jhydrogen\j tank. One of the liquid oxygen tank umbilical valves is for liquid oxygen, the other for gaseous oxygen. The liquid \Jhydrogen\j tank umbilical has two valves for liquid and one for gas. The intermediate-diameter liquid \Jhydrogen\j umbilical is a recirculation umbilical used only during the liquid \Jhydrogen\j chill-down sequence during prelaunch.
The ET also has two electrical umbilicals that carry electrical power from the orbiter to the tank and the two SRBs and provide information from the SRBs and ET to the orbiter.
A swing-arm-mounted cap to the fixed service structure covers the oxygen tank vent on top of the ET during the countdown and is retracted about two minutes before lift- off. The cap siphons off oxygen vapor that threatens to form large ice on the ET, thus protecting the orbiter's thermal protection system during launch.
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"Shuttle ET Range Safety System",273,0,0,0
A range safety system provides for dispersing tank propellants if necessary. It includes a battery power source, a receiver / decoder, antennas and ordnance.
Various parameters are monitored and displayed on the flight deck display and control panel and are transmitted to the ground.
The contractor for the external tank is Martin Marietta Aero space, New Orleans, La. The tank is manufactured at Michoud, La. Motorola, Inc., Scottsdale, Ariz., is the contractor for range safety receivers.
#
"Orbiter Structure",274,0,0,0
The orbiter structure is divided into nine major sections: the forward fuselage, which consists of upper and lower sections that fit clam-like around a pressurized crew compartment; wings; midfuselage; payload bay doors; aft fuselage; forward reaction control system; vertical tail; orbital maneuvering system/reaction control system pods; and body flap. The majority of the sections are constructed of conventional aluminum and protected by reusable surface \Jinsulation\j.
The forward fuselage structure is composed of 2024 aluminum alloy skin-stringer panels, frames and bulkheads.
The crew compartment is supported within the forward fuselage at four attachment points and is welded to create a pressure-tight vessel. The three-level compartment has a side hatch for normal passage and hatches in the airlock to permit extravehicular and intravehicular activities. The side hatch can be jettisoned.
The midfuselage is a 60-foot section of primary load-carrying structure between the forward and aft fuselages. It includes the wing carry-through structure and the payload bay doors. The skins consist of integral-machined aluminum panels and aluminum honeycomb sandwich panels.
The frames are constructed from a combination of aluminum panels with riveted or machined integral stiffeners and a truss structure center section. The upper half of the midfuselage consists of structural payload bay doors hinged along the side and split at the top centerline. The doors are \Jgraphite\j epoxy frames and honeycomb panel construction.
The aft fuselage includes a truss-type internal structure of diffusion-bonded elements that transfer the main engine thrust loads to the midfuselage and external tank. (In OV-105 , the truss-type internal structure is of a forging construction.) The aft fuselage's external surface is of standard construction except for the removable OMS/RCS pods, which are constructed of \Jgraphite\j epoxy skins and frames. An aluminum bulkhead shield with reusable \Jinsulation\j at the rear of the orbiter protects the rear portion of the aft fuselage.
The wing is constructed of a conventional aluminum alloy, using a corrugated spar web, truss-type ribs and riveted skin-stringer and honeycomb covers. The elevons are constructed of aluminum honeycomb and are split into two segments to minimize hinge binding and interaction with the wing.
The vertical tail, a conventional aluminum alloy structure, is a two-spar, multirib, integrally machined skin assembly. The tail is attached to the aft fuselage by bolted fittings at the two main spars. The rudder/speed brake assembly is divided into upper and lower sections, which are split longitudinally and actuated individually to serve as both rudder and speed brake.
These major structural assemblies are mated and held together by rivets and bolts. The midfuselage is joined to the forward and aft fuselage primarily by shear ties, with the midfuselage overlapping the bulkhead caps at stations Xo 582 and Xo 1307. The wing is attached to the midfuselage and aft fuselage primarily by shear ties, except in the area of the wing carry-through, where the upper panels are attached with tension bolts. The vertical tail is attached to the aft fuselage with bolts that work in both shear and tension. The body flap, which has aluminum honeycomb covers, is attached to the lower aft fuselage by four rotary actuators.
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"Orbiter Forward Fuselage",275,0,0,0
The forward fuselage consists of the upper and lower fuselages. It houses the crew compartment and supports the forward reaction control system module, nose cap, nose gear wheel well, nose gear and nose gear doors.
The forward fuselage is constructed of conventional 2024 aluminum alloy skin-stringer panels, frames and bulkheads. The panels are single curvature and stretch-formed skins with riveted stringers spaced 3 to 5 inches apart. The frames are riveted to the skin-stringer panels. The major frames are spaced 30 to 36 inches apart. The Y o 378 upper forward bulkhead is constructed of flat aluminum and formed sections riveted and bolted together; the lower is a machined section. The bulkhead provides the interface fitting for the nose section.
The nose section contains large machined beams and struts. The structure for the nose landing gear wheel well consists of two support beams, two upper closeout webs, drag-link support struts, nose landing gear strut and actuator attachment fittings, and the nose landing gear door fittings. The left and right landing gear doors are attached by hinge fittings in the nose section.
The doors are constructed of aluminum alloy honeycomb, and although the doors are the same length, the left door is wider than the right. Each door has an up-latch fitting at the forward and aft ends to lock the door closed when the gear is retracted, and each has a pressure seal in addition to a thermal barrier.
Lead ballast in the nose wheel well and on the X o 378 bulkhead provides weight and center-of-gravity control. The nose wheel well will accommodate 1,350 pounds of ballast, and the X o 378 bulkhead will accommodate a maximum of 2,660 pounds.
The forward fuselage carries the basic body-bending loads (a tendency to change the radius of a curvature of the body) and reacts nose landing gear loads.
The forward fuselage is covered with reusable \Jinsulation\j, except for the six windshields, two overhead windows and side hatch window areas around the forward RCS engines. The nose cap is also a reusable thermal protection system. It is constructed of reinforced carbon-carbon and has thermal barriers at the nose cap-structure interface.
The forward fuselage skin has structural provisions for installing antennas, deployable air data probes and the door eyelet openings for the two star trackers. Two openings are required in the upper forward fuselage for star tracker viewing. Each opening has a door for environmental control.
The forward orbiter/external tank attach fitting is at the Xo 378 bulkhead and the skin panel structure aft of the nose gear wheel well. Purge and vent control is provided by flexible boots between the forward fuselage and crew compartment around the windshield windows, overhead observation window, crew hatch window and star tracker openings. The forward fuselage is isolated from the payload bay by a flexible membrane between the forward fuselage and crew compartment at Xo 582.
Six forward outer pane windshields are installed on the forward fuselage. They are described in the section on windows. The window structural frames in the forward fuselage are five-axis machined parts.
The forward RCS module is constructed of conventional 2024 aluminum alloy skin-stringer panels and frames. The panels are composed of single-curvature and stretch-formed skins with riveted stringers. The frames are riveted to the skin-stringer panels. The forward RCS module is secured to the forward fuselage nose section and forward bulkhead of the forward fuselage with 16 fasteners, which permit the installation and removal of the module.
The components of the forward RCS are mounted and attached to the module, which will have a reusable thermal protection cover, in addition to thermal barriers installed around it and the RCS engine interfaces and the interface-attachment area to the forward fuselage.
The forward fuselage and forward RCS module are built by Rockwell's Space Transportation Systems Division, Downey, Calif.
#
"Orbiter Crew Compartment",276,0,0,0
The three-level crew compartment is constructed of 2219 aluminum alloy plate with integral stiffening stringers and internal framing welded together to create a pressure-tight vessel. The compartment has a side hatch for normal ingress and egress, a hatch into the airlock from the middeck, and a hatch from the airlock through the aft bulkhead into the payload bay for extravehicular activity and payload bay access.
Redundant pressure window panes are provided in the six forward windshields, the two overhead viewing windows, the two aft viewing windows and the side hatch windows; they are described in the window section. Approximately 300 penetrations in the pressure shell are sealed with plates and fittings.
A large removable panel in the aft bulkhead provides access to the interior of the crew compartment during initial fabrication and assembly and provides for airlock installation and removal. The compartment supports the environmental control and life support system; avionics; guidance, navigation and control equipment; inertial measurement units; displays and controls; star trackers; and crew accommodations for sleeping, waste management, seats and an optional galley.
The crew compartment is supported within the forward fuselage at only four attach points to minimize the thermal \Jconductivity\j between them. The two major attach points are located at the aft end of the crew compartment at the flight deck floor level. The vertical load reaction link is on the centerline of the forward bulkhead. The lateral load reaction is on the lower segment of the aft bulkhead.
The compartment is configured to accommodate a crew of four on the flight deck and three in the middeck. In OV-102, four can be accommodated in the middeck. The crew cabin arrangement consists of a flight deck, middeck and lower level equipment bay.
The crew compartment is pressurized to 14.7 psia, plus or minus 0.2 psia, and is maintained at an 80-percent \Jnitrogen\j and 20-percent oxygen composition by the ECLSS, which provides a shirt-sleeve environment for the flight crew. The crew compartment is designed for 16 psia.
The crew compartment's volume with the airlock in the middeck is 2,325 cubic feet. If the airlock is in the payload bay, the crew compartment's cabin volume is 2,625 cubic feet.
The flight deck is the uppermost compartment of the cabin. The commander's and pilot's work stations are positioned side by side in the forward portion of the flight deck. These stations have controls and displays for maintaining autonomous control of the vehicle throughout all mission phases. Directly behind and to the sides of the commander and pilot centerline are the mission specialist seats.
The commander's and pilot's seats have two shoulder harnesses and a lap belt for restraints. The shoulder harnesses have an \Jinertia\j reel lock/unlock feature. The unlocked position allows the shoulder harness to move. The commander and pilot can move their seats along the orbiter's Z (vertical) and X (longitudinal) axes so they can reach and see controls better during the ascent and entry phases of flight.
Seat movement for each axis is provided by a single ac motor. The total travel distance for the Z and X axes is 10 and 5 inches, respectively. Seat adjustment controls are located on the left side of the seat pan and consist of a three-position toggle switch for power bus selection and one spring-loaded, three-position toggle switch each to control horizontal and vertical seat movement.
Each mission and payload specialist's seat has two shoulder harnesses and a lap belt for restraints. The specialists' seats have controls to manually lock and unlock the tilt of the seat back. Each seat has removable seat cushions and mounting provisions for oxygen and communications connections to the CAPS. The specialists' seats are removed and stowed in the middeck on orbit. No tools are required since the legs of each seat have quick-disconnect fittings. Each seat is 25.5 inches long, 15.5 inches wide and 11 inches high when folded for stowage.
The aft flight deck has two overhead and aft viewing windows for viewing orbital operations. The aft flight deck station also contains displays and controls for executing attitude or translational maneuvers for rendezvous, stationkeeping, docking, payload deployment and retrieval, payload monitoring, remote manipulator system controls and displays, payload bay door operations and closed-circuit \Jtelevision\j operations.
The forward flight deck, which includes the center console and seats, is approximately 24 square feet. However, the side console controls and displays add approximately 3.5 square feet more. If the center console is subtracted from the 24 square feet, this would amount to approximately 5.2 square feet.
The aft flight deck is approximately 40 square feet.
Directly beneath the flight deck is the middeck. Access to the middeck is through two interdeck openings, which measure 26 by 28 inches. Normally, the right interdeck opening is closed and the left is open. A ladder attached to the left interdeck access allows easy passage in 1-g conditions. The middeck provides crew accommodations and contains three avionics equipment bays.
The two forward avionics bays utilize the complete width of the cabin and extend into the middeck 39 inches from the forward bulkhead. The aft bay extends into the middeck 39 inches from the aft bulkhead on the right side of the airlock. Just forward of the waste management system is the side hatch. The completely stripped middeck is approximately 160 square feet; the gross mobility area is approximately 100 square feet.
The side hatch in the middeck is used for normal crew entrance/exit and may be operated from within the crew cabin middeck or externally. It can be jettisoned for emergencies, as discussed in the escape system section. It is attached to the crew cabin tunnel by hinges, a \Jtorque\j tube and support fittings.
The hatch opens outwardly 90 degrees down with the orbiter horizontal or 90 degrees sideways with the orbiter vertical. It is 40 inches in diameter and has a 10-inch clear-view window in the center of the hatch. The window consists of three panes of glass. The side hatch has a pressure seal that is compressed by the side hatch latch mechanisms when the hatch is locked closed.
A thermal barrier of Inconel wire mesh spring with a ceramic fiber braided sleeve is installed between the reusable surface \Jinsulation\j tiles on the forward fuselage and the side hatch. The total weight of the side hatch is 294 pounds.
Depending on the mission requirements, bunk sleep stations and a galley can be installed in the middeck. In addition, three or four seats of the same type as the mission specialists' seats on the flight deck can be installed in the middeck. Three seats over the normal three could be installed in the middeck for rescue missions if the bunk sleep stations were removed.
The waste management system, located in the middeck, can also accommodate payloads in the pressurized crew compartment environment.
The middeck also provides a stowage volume of 140 cubic feet. Accommodations are included for dining, sleeping, maintenance, exercising and data management. On the orbiter centerline, just aft of the forward avionics equipment bay, an opening in the ceiling provides access to the inertial measurement units.
The middeck floor contains removable panels that provide access to the ECLSS equipment. The middeck equipment bay below the middeck floor houses the major components of the waste management and air revitalization systems, such as pumps, fans, \Jlithium\j \Jhydroxide\j, absorbers, heat exchangers and ducting. This compartment has space for stowing \Jlithium\j \Jhydroxide\j canisters and five separate spaces for crew equipment stowage with a volume of 29.92 cubic feet.
Modular stowage lockers are used to store the flight crew's personal gear, mission-necessary equipment, personal hygiene equipment and experiments. The modular lockers are made of sandwich panels of Kevlar/epoxy and a non-metallic core. This reduced the lockers' weight by 83 percent compared to all-aluminum lockers, a reduction of approximately 150 pounds. There are 42 identical boxes, which are 11 by 18 by 21 inches.
An airlock, located in the middeck, is composed of machined aluminum sections welded together to form a cylinder with hatch mounting flanges. The upper cylindrical section and bulkheads are constructed of aluminum honeycomb. Two semicylindrical aluminum sections are welded to the airlock's primary structure to house the ECLSS and electrical support equipment. Each semicylindrical section has three feedthrough plates for plumbing and cable routings from the orbiter to the airlock.
Normally, two extravehicular mobility units are stowed in the airlock. The EMU is an integrated space suit assembly and life support system that enables flight crew members to leave the pressurized orbiter crew cabin and work outside the cabin in space.
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"Orbiter Airlock",277,0,0,0
The airlock is normally located inside the middeck of the \Jspacecraft\j's pressurized crew cabin. It has an inside diameter of 63 inches, is 83 inches long and has two 40-inch- diameter D-shaped openings that are 36 inches across. It also has two pressure-sealing hatches and a complement of airlock support systems. The airlock's volume is 150 cubic feet.
The airlock is sized to accommodate two fully suited flight crew members simultaneously. Support functions include airlock depressurization and repressurization, extravehicular activity equipment recharge, liquid-cooled garment water cooling, EVA equipment checkout, donning and communications. The EVA gear, checkout panel and recharge stations are located on the internal walls of the airlock.
The airlock hatches are mounted on the airlock. The inner hatch is mounted on the exterior of the airlock (orbiter crew cabin middeck side) and opens into the middeck. The inner hatch isolates the airlock from the orbiter crew cabin. The outer hatch is mounted inside the airlock and opens into the airlock. The outer hatch isolates the airlock from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock to the payload bay when open.
Airlock repressurization is controllable from the orbiter crew cabin middeck and from inside the airlock. It is performed by equalizing the airlock's and cabin's pressure with equalization valves mounted on the inner hatch. The airlock is depressurized from inside the airlock by venting the airlock's pressure overboard. The two D-shaped airlock hatches open toward the primary pressure source, the orbiter crew cabin, to achieve pressure-assist sealing when closed.
Each hatch has six interconnected latches and a gearbox/actuator, a window, a hinge mechanism and hold-open device, a differential pressure gauge on each side and two equalization valves.
The 4-inch diameter window in each airlock hatch is used for crew observation from the cabin/airlock and the airlock/payload bay. The dual window panes are made of polycarbonate plastic and mounted directly to the hatch by means of bolts fastened through the panes. Each hatch window has dual pressure seals, with seal grooves located in the hatch.
Each airlock hatch has dual pressure seals to maintain pressure integrity. One seal is mounted on the airlock hatch and the other on the airlock structure. A leak check quick disconnect is installed between the hatch and the airlock pressure seals to verify hatch pressure integrity before flight.
The gearbox with latch mechanisms on each hatch allows the flight crew to open and close the hatch during transfers and EVA operations. The gearbox and the latches are mounted on the low-pressure side of each hatch; with a gearbox handle installed on both sides to permit operation from either side of the hatch.
Three of the six latches on each hatch are double-acting and have cam surfaces that force the sealing surfaces apart when the latches are opened, thereby acting as crew assist devices. The latches are interconnected with push-pull rods and an idler bell crank that is installed between the rods for pivoting the rods. Self-aligning dual rotating bearings are used on the rods for attachment to the bellcranks and the latches. The gearbox and hatch open support struts are also connected to the latching system by the same rod/bellcrank and bearing system. To latch or unlatch the hatch, the gearbox handle must be rotated 440 degrees.
The hatch actuator/gearbox is used to provide the mechanical advantage to open and close the latches. The hatch actuator lock lever requires a force of 8 to 10 pounds through an angle of 180 degrees to unlatch the actuator. A minimum rotation of 440 degrees with a maximum force of 30 pounds applied to the actuator handle is required to operate the latches to their fully unlatched positions.
The hinge mechanism for each hatch permits a minimum opening sweep into the airlock or the crew cabin middeck. The inner hatch (airlock to crew cabin) is pulled or pushed forward to the crew cabin approximately 6 inches. The hatch pivots up and to the right side. Positive locks are provided to hold the hatch in both an intermediate and a full-open position. A spring-loaded handle on the latch hold-open bracket releases the lock. Friction is also provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The outer hatch (airlock to payload bay) opens and closes to the contour of the airlock wall. The hatch is hinged to be pulled first into the airlock and then forward at the bottom and rotated down until it rests with the low-pressure (outer) side facing the airlock ceiling (middeck floor). The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke.
The hatch has a hold-open hook that snaps into place over a flange when the hatch is fully open. The hook is released by depressing the spring-loaded hook handle and pushing the hatch toward the closed position. To support and protect the hatch against the airlock ceiling, the hatch incorporates two deployable struts. The struts are connected to the hatch linkage mechanism and are deployed when the hatch linkage is rotated open. When the hatch latches are rotated closed, the struts are retracted against the hatch.
The airlock hatches can be removed in flight from the hinge mechanism using pip pins, if required.
The airlock air circulation system provides conditioned air to the airlock during non-EVA periods. The airlock revitalization system duct is attached to the outside airlock wall at launch. Upon airlock hatch opening in flight, the duct is rotated by the flight crew through the cabin/airlock hatch, installed in the airlock and held in place by a strap holder.
The duct has a removable air diffuser cap, installed on the end of the flexible duct, which can adjust the air flow from 216 pounds per hour. The duct must be rotated out of the airlock before the cabin/airlock hatch is closed for airlock depressurization. During the EVA preparation period, the duct is rotated out of the airlock and can be used for supplemental air circulation in the middeck.
To assist the crew member before and after EVA operations, the airlock incorporates handrails and foot restraints. Handrails are located alongside the avionics and ECLSS panels. Aluminum alloy handholds mounted on each side of the hatches have oval configurations 0.75 by 1.32 inches and are painted yellow. They are bonded to the airlock walls with an epoxyphenolic adhesive.
Each handrail has a clearance of 2.25 inches between the airlock wall and the handrail to allow the astronauts to grip it while wearing a pressurized glove. Foot restraints are installed on the airlock floor nearer the payload bay side. The ceiling handhold is installed nearer the cabin side of the airlock. The foot restraints can be rotated 360 degrees by releasing a spring-loaded latch and lock in every 90 degrees.
A rotation release knob on the foot restraint is designed for shirt-sleeve operation and, therefore, must be positioned before the suit is donned. The foot restraint is bolted to the floor and cannot be removed in flight. It is sized for the EMU boot. The crew member first inserts his foot under the toe bar and then rotates his heel from inboard to outboard until the heel of the boot is captured.
There are four floodlights in the airlock.
If the airlock is relocated to the payload bay from the middeck, it will function in the same manner as in the middeck. \JInsulation\j is installed on the airlock's exterior for protection from the extreme temperatures of space.
For Spacelab pressurized module missions, the airlock remains in the crew compartment middeck, and a tunnel adapter that mates with the airlock and the Spacelab tunnel is installed in the payload bay.
The airlock tunnel adapter, hatches, tunnel extension and tunnel permit the flight crew members to transfer from the \Jspacecraft\j's pressurized middeck crew compartment to Spacelab's pressurized shirt-sleeve environment.
In addition, the airlock, tunnel adapter and hatches permit the EVA flight crew members to transfer from the airlock/tunnel adapter in the space suit assembly into the payload bay without depressurizing the crew cabin and Spacelab.
The Spacelab tunnel and Spacelab are accessed via the tunnel adapter, which is located in the payload bay and is attached to the airlock at orbiter station Xo 576 and the tunnel extension at X o 660. The tunnel adapter has an inside diameter of 63 inches at its widest section and tapers in the cone area at each end to two 40-inch- diameter D-shaped openings 36 inches across. A 40-inch- diameter D-shaped opening 36 inches across is located at the top of the tunnel adapter.
Two pressure-sealing hatches are located in the tunnel adapter, one in the upper area of the tunnel adapter and one in the aft end of the tunnel adapter. The tunnel adapter is a welded structure constructed of 2219 aluminum with 2.4- by 2.4-inch exposed structural ribs on the exterior surface and external waffle skin stiffening.
The hatch located on the middeck side of the airlock is mounted on the exterior of the airlock and opens into the middeck. The hatch isolates the airlock from the crew cabin. The hatch located in the tunnel adapter's aft end isolates the tunnel adapter/airlock from the tunnel extension, tunnel and Spacelab. This hatch opens into the tunnel adapter.
The hatch located in the tunnel adapter at the upper D-shaped opening isolates the airlock/tunnel adapter from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock/tunnel adapter to the payload bay when open. This hatch opens into the tunnel adapter.
The hinge mechanism for each hatch permits a minimum opening sweep into the tunnel adapter or the \Jspacecraft\j crew cabin middeck. The airlock crew cabin hatch in the middeck is pulled/pushed forward to the middeck approximately 6 inches.
The hatch pivots up and right. Positive locks are provided to hold the latch in both an intermediate and a full-open position. A spring-loaded handle on the latch hold-open bracket releases the lock. Friction is provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The aft hatch is hinged to be pulled first into the tunnel adapter and then forward at the bottom. The top of the hatch is rotated towards the tunnel and downward until the hatch rests with the Spacelab side facing the tunnel adapter floor. The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The upper (EVA) hatch in the tunnel adapter opens and closes to the left wall of the tunnel adapter. The hatch is hinged to be pulled first into the tunnel adapter and then forward at the hinge area and rotated down until it rests against the port wall of the tunnel adapter. The linkage mechanism guides the hatch from the closed/open, open/closed position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The hatches can be removed in flight from the hinge mechanisms via pip pins, if required.
The crew compartment, bunk sleep stations (if installed), airlock and modular stowage lockers are built by Rockwell's Space Transportation Systems Division, Downey, Calif. The original crew seat contractor was AMI of \JColorado\j Springs, Colo., but later Rockwell's Space Transportation Systems Division. The Spacelab pressurized module tunnel adapter and tunnel contractor is McDonnell Douglas Astronautics, Huntington Beach, Calif.
#
"Orbiter Forward Fuselage and Cabin Windows",278,0,0,0
The orbiter windows provide visibility for entry, landing and on-orbit operations. For atmospheric flight, the flight crew needs forward, left and right viewing areas. On-orbit mission phases require visibility for rendezvous, docking and payload-handling operations.
The six windows located at the forward flight deck commander and pilot stations provide forward, left and right viewing. The two overhead windows and two payload-viewing windows at the aft station location on the flight deck provide rendezvous, docking and payload viewing. There is also a window in the middeck side hatch.
The six planeform-shaped forward windows are the thickest pieces of glass ever produced in the optical quality for see-through viewing. Each consists of three individual panes. The innermost pane is constructed of tempered aluminosilicate glass to withstand the crew compartment pressure. It is 0.625 of an inch thick. Aluminosilicate glass is a low-expansion glass that can be tempered to provide maximum mechanical strength. The exterior of this pane, called a pressure pane, is coated with a red reflector coating to reflect the infrared (heat portion) rays while transmitting the visible spectrum.
The center pane is constructed of low-expansion, fused \Jsilica\j glass because of its high optical quality and excellent thermal shock resistance. This pane is 1.3 inches thick.
The inner and outer panes are coated with a high-efficiency, anti-reflection coating to improve visible light transmission. These windows withstand a proof pressure of 8,600 psi at 240 F and 0.017 relative \Jhumidity\j.
The outer pane is made of the same material as the center pane and is 0.625 of an inch thick. The exterior is uncoated, but the interior is coated with high-efficiency, anti-reflection coating. The outer surface withstands approximately 800 F.
Each of the forward six windows' outer panes measures 42 inches diagonally, and the center and inner panes each measure 35 inches diagonally. The outer panes of the forward six windows are mounted and attached to the forward fuselage. The center and inner panes are mounted and attached to the crew compartment. Redundant seals are employed on each window. No sealing/bonding compounds are used.
The two overhead windows at the flight deck aft station are identical in construction to the six forward windows except for thickness. The inner and center panes are 0.45 of an inch thick, and the outer pane is 0.68 of an inch thick. The outer pane is attached to the forward fuselage, and the center and inner panes are attached to the crew compartment.
The two overhead windows' clear view area is 20 by 20 inches. The left-hand overhead window provides the crew members with a secondary emergency egress. The inner and center panes open into the crew cabin, and the outer pane is jettisoned up and over the top of the orbiter. This provides a secondary emergency exit area of 20 by 20 inches.
On the aft flight deck, each of the two windows for viewing the payload bay consists of only two panes of glass, which are identical to the forward windows' inner and center panes. The outer thermal panes are not installed. Each pane is 0.3 of an inch thick. The windows are 14.5 by 11 inches. Both panes are attached to the crew compartment.
The side hatch viewing window consists of three panes of glass identical to the six forward windows. The inner pane is 11.4 inches in diameter and 0.25 of an inch thick. The center pane is 11.4 inches in diameter and 0.5 of an inch thick. The outer pane is 15 inches in diameter and 0.3 of an inch thick.
During orbital operations, the large window areas of transparency expose the flight crew to sun glare; therefore, window shades and filters are provided to preclude or minimize exposure. Shades are provided for all windows, and filters are supplied for the aft and overhead viewing windows. The window shades and filters are stored in the middeck of the orbiter crew compartment. Attachment mechanisms and devices are provided for their installation at each window on the flight deck.
The forward station window shades (W-1 through W-6) are fabricated from Kevlar/epoxy glass fabric with silver and Inconel-coated Teflon tape on the outside surface and paint on the inside surface. When the shade is installed next to the inner window pane, a silicone rubber seal around the periphery deforms to prevent light leakage. The shade is held in place by the shade installation guide, the hinge plate and the Velcro keeper.
The overhead window shades (W-7 and W-8) are nearly the same as the forward shades; but the rubber seal is deleted, and the shade is sealed and held in place by a separate seal around the window opening, a hinge plate and secondary frame, and Velcro retainer. The overhead window filters are fabricated from Lexan and are used interchangeably with the shades.
The aft window shades (W-9 and W-10) are the same as the overhead window shades except that a 0.63-inch-wide strip of Nomex Velcro has been added around the perimeter of the shade. The shade is attached to the window by pressing the Velcro strip to the pile strip around the window opening. The aft window filters are the same as the overhead window filters except for the addition of the Velcro hook strip. The filters and shades are used interchangeably.
The side hatch window cover is permanently attached to the window frame and is hinged to allow opening and closing.
The contractor for the windows is Corning Glass Co., Corning, N.Y.
#
"Orbiter Wing",279,0,0,0
The wing is an aerodynamic lifting surface that provides conventional lift and control for the orbiter. The left and right wings consist of the wing glove; the intermediate section, which includes the main landing gear well; the \Jtorque\j box; the forward spar for mounting the reusable reinforced carbon-carbon leading edge structure thermal protection system; the wing/elevon interface; the elevon seal panels; and the elevons.
The wing is constructed of conventional aluminum alloy with a multirib and spar arrangement with skin-stringer-stiffened covers or honeycomb skin covers. Each wing is approximately 60 feet long at the fuselage intersection and has a maximum thickness of 5 feet.
The forward wing box is an extension of the basic wing that aerodynamically blends the wing leading edge into the midfuselage wing glove. The forward wing box is a conventional design of aluminum ribs, aluminum tubes and tubular struts. The upper and lower wing skin panels are stiffened aluminum. The leading edge spar is constructed of corrugated aluminum.
The intermediate wing section consists of the conventional aluminum multiribs and aluminum tubes. The upper and lower skin covers are constructed of aluminum honeycomb. A portion of the lower wing surface skin panel includes the main landing gear door.
The intermediate section houses the main landing gear compartment and reacts a portion of the main landing gear loads. A structural rib supports the outboard main landing gear door hinges and the main landing gear trunnion and drag link. The support for the inboard main landing gear trunnion and drag link attachment is provided by the midfuselage. The main landing gear door is conventional aluminum honeycomb.
The four major spars are constructed of corrugated aluminum to minimize thermal loads. The forward spar provides the attachment for the thermal protection system reusable reinforced carbon-carbon leading edge structure. The rear spar provides the attachment interfaces for the elevons, hinged upper seal panels, and associated hydraulic and electrical system components. The upper and lower wing skin panels are stiffened aluminum.
The elevons provide orbiter flight control during atmospheric flight. The two-piece elevons are conventional aluminum multirib and beam construction with aluminum honeycomb skins for compatibility with the acoustic environment and thermal interaction. The elevons are divided into two segments for each wing, and each segment is supported by three hinges. The elevons are attached to the flight control system hydraulic actuators at points along their forward extremities, and all hinge moments are reacted at these points. Each elevon travels 40 degrees up and 25 degrees down.
The transition area on the upper surface between the \Jtorque\j box and the movable elevon consists of a series of hinged panels that provide a closeout of the wing-to-elevon cavity. These panels are of Inconel honeycomb sandwich construction outboard of wing station Y w 312.5 and of \Jtitanium\j honeycomb sandwich construction inboard of wing station Y w 312.5.
The upper leading edge of each elevon incorporates \Jtitanium\j rub strips. The rub strips are of \Jtitanium\j honeycomb construction and are not covered with the thermal protection system reusable surface \Jinsulation\j. They provide the sealing surface area for the elevon seal panels.
The exposed areas of the wings, main landing gear doors and elevons are covered with reusable surface \Jinsulation\j thermal protection system materials except for the elevon seal panels.
Thermal seals are provided on the elevon lower cove area along with thermal spring seals on the upper rub panels. Pressure seals and thermal barriers are provided on the main landing gear doors.
The wing is attached to the fuselage with a tension bolt splice along the upper surface. A shear splice along the lower surface in the area of the fuselage carry-through completes attachment interface.
Prior to the manufacturing of the wings for Discovery (OV-103) and Atlantis (OV-104), a weight reduction program resulted in a redesign of certain areas of the wing structure. An assessment of wing air loads was made from actual flight data that indicated greater loads on the wing structure. As a result, to maintain positive margins of safety during ascent, structural modifications were incorporated into certain areas of the wings. The modifications consisted of the addition of doublers and stiffeners.
The wing, elevon and main landing gear door contractor is Grumman Corp., Bethpage, N.Y.
#
"Orbiter Midfuselage",280,0,0,0
The midfuselage structure interfaces with the forward fuselage, aft fuselage and wings. It supports the payload bay doors, hinges, tie-down fittings, forward wing glove, and various orbiter system components and forms the payload bay area.
The forward and aft ends of the midfuselage are open, with reinforced skin and longerons interfacing with the bulkheads of the forward and aft fuselages. The midfuselage is primarily an aluminum structure 60 feet long, 17 feet wide and 13 feet high. It weighs approximately 13,502 pounds.
The midfuselage skins are integrally machined by numerical control. The panels above the wing glove and the wings for the forward eight bays have longitudinal T-stringers. The five aft bays have aluminum honeycomb panels. The side skins in the shadow of the wing are also numerically control machined but have vertical stiffeners.
Twelve main-frame assemblies stabilize the midfuselage structure. The assemblies consist of vertical side elements and horizontal elements. The side elements are machined; whereas the horizontal elements are boron/aluminum tubes with bonded \Jtitanium\j end fittings, which reduced the weight by 49 percent (approximately 305 pounds).
In the upper portion of the midfuselage are the sill and door longerons. The machined sill longerons not only make up the primary body-bending elements, but also take the longitudinal loads from payloads in the payload bay. The payload bay door longerons and associated structure are attached to the 13 payload bay door hinges. These hinges provide the vertical reaction from the payload bay doors.
Five of the hinges react the payload bay door shears. The sill longeron also provides the base support for the payload bay manipulator arm (if installed) and its stowage provisions, the Ku-band rendezvous antenna, the antenna base support and its stowage provisions, and the payload bay door actuation system.
The side wall forward of the wing carry-through structure provides the inboard support for the main landing gear. The total lateral landing gear loads are reacted by the midfuselage structure.
The midfuselage also supports the two electrical wire trays that contain the wiring between the crew compartment and aft fuselage.
Plumbing and wiring in the lower portion of the midfuselage are supported by fiberglass milk stools.
The remainder of the exposed areas of the midfuselage is covered with the reusable surface \Jinsulation\j thermal protection system.
Because of additional detailed analysis of actual flight data concerning descent stress thermal \Jgradient\j loads, torsional straps were added to the lower midfuselage stringers in bays 1 through 11. The torsional straps tie all stringers together similarly to a box section, which eliminates rotational (torsional) capabilities to provide positive margins of safety.
Also, because of additional detailed analysis of actual flight data during descent, room-temperature vulcanizing silicone rubber material was bonded to the lower midfuselage from bay 4 through 12 to act as a heat sink and distribute temperatures evenly across the bottom of the midfuselage, which will reduce thermal gradients and ensure positive margins of safety.
The contractor for the midfuselage is General Dynamics Corp., Convair Aerospace Division, San Diego, Calif.
#
"Orbiter Payload Bay Doors",281,0,0,0
The payload bay doors are opened shortly after orbit is achieved to allow exposure of the environmental control and life support system radiators for heat rejection of the orbiter's systems. The payload bay doors consist of port and starboard doors hinged at each side of the midfuselage and latched mechanically at the forward and aft fuselage and at the split-top centerline.
Thermal seals on the doors provide a relatively air-tight payload compartment when the doors are closed and latched. During prelaunch and postlanding, the purge, vent and drain system permits purging of undesirable gases and maintains a positive delta pressure for venting of payloads within the payload area when the doors are closed.
The port and starboard doors are 60 feet long with a combined area of approximately 1,600 square feet. Each door is made up of five segments that are interconnected by circumferential expansion joints. Each door hinges on 13 Inconel 718 external hinges (five shear and eight idlers).
The lower half of each hinge attaches to the midfuselage sill longeron. The hinges rotate on bearings with dual rotational surfaces. There are five shear hinges and eight floating hinges. The floating hinges allow fore and aft movement of the door panels for thermal expansion.
Each door actuation system provides the mechanism to drive each door side to the open or closed position. Each mechanism consists of an electromechanical power drive unit and six rotary gear actuators, which are connected by \Jtorque\j tubes to each other and to the power drive unit. Linkages transmit \Jtorque\j from the rotary actuators to the doors.
The forward 30-foot sections of both doors incorporate radiators that can be deployed; they are hinged and latched to the door inner surface in order to reject the excess heat of the Freon-21 coolant loops from both sides of the radiator panels when the doors are open. An electromechanical actuation system on the door unlatches and deploys the radiators when open and latches and stows the radiators when closed.
The radiators may be left in the stowed position for a given flight and will only radiate the excess heat from the one side. Fixed radiator panels are installed on the forward end of the aft payload bay doors and radiate from one side only. Kitted fixed radiator panels may be installed on the aft end of the aft payload bay doors when required by a specific mission; they also will radiate from only one side.
During payload bay door closure, the aft flight deck payload bay door crewman optical alignment sight is used to check door alignment.
When the payload bay doors are closed, they are fixed at the aft fuselage bulkhead and allowed to move longitudinally at the forward fuselage. The doors also accommodate vehicle torsional loads (a force that causes a body, such as a shaft, to twist about its longitudinal axis), aerodynamic pressure loads and payload bay vent lag pressures. The payload bay is not a pressurized area.
Thermal and pressure seals are used to close the gaps at the forward and aft fuselage interface, door centerline and circumferential expansion joints.
The doors are 60 feet long. Each consists of five segments interconnected by expansion joints. The chord of each half of these curved doors is approximately 10 feet, and the doors are 15 feet in diameter.
The doors are constructed of \Jgraphite\j epoxy composite material, which reduces the weight by 23 percent over that of aluminum honeycomb sandwich. This is a reduction of approximately 900 pounds, which brings the weight of the doors down to approximately 3,264 pounds. The payload bay doors are the largest aerospace structure to be constructed from composite material.
The composite doors will withstand 163-decibel acoustic noise and a temperature range of minus 170 to plus 135 F.
The doors are made up of subassemblies consisting of \Jgraphite\j epoxy honeycomb sandwich panels, solid \Jgraphite\j epoxy laminate frames, expansion joint frames, \Jtorque\j box, seal depressor, centerline beam intercostals, gussets, end fittings and clips. There are also aluminum 2024 shear pins, \Jtitanium\j fittings, and Inconel 718 floating and shear hinges. The assembly is joined by mechanical fasteners. \JLightning\j strike protection is provided by aluminum mesh wire bonded to the outer skin.
Extravehicular activity handholds are attached in the \Jtorque\j box areas.
The payload bay doors are covered with reusable surface \Jinsulation\j.
The left door with attached systems weighs approximately 2,375 pounds and the right weighs about 2,535 pounds. The right door contains the centerline latch active mechanisms, which accounts for the weight difference. These weights do not include the radiator panel system, which adds 833 pounds per door.
When closed, the doors are latched to the forward and aft bulkheads and along the upper centerline of the doors. The latching system consists of 16 bulkhead latches (eight aft and eight forward) and 16 payload bay door centerline latches. The forward and aft bulkhead latches are in groups of four ganged latch hooks. The centerline latches are also in groups of four ganged latches. Each centerline latch gang incorporates four latches, bellcranks, push rods, levers, rollers and an electromechanical actuator.
The forward and aft bulkhead latches are arranged in groups of four ganged latches. Each group is opened or closed by an electromechanical actuator with two redundant, three-phase ac reversible motors that receive ac power from mid motor controller assemblies when commanded in the automatic predetermined sequence or by manual keyboard entries. In the automatic mode, the forward and aft bulkhead latches operate simultaneously.
The payload bay door centerline latch groups are controlled automatically in a predetermined sequence or manually by individual latch groups through keyboard entries in a manner similar to the bulkhead latch groups. The 16 centerline latches are arranged into groups of four, similar to the bulkhead latches.
Each centerline latch group consists of two ac reversible electric motors that drive a rotary shaft and bellcrank and four hooks to engage a corresponding passive roller to latch the door closed or disengage the passive roller to unlatch the door. All 16 centerline hook assemblies contain alignment rollers to eliminate payload bay door overlap due to thermal distortion. Passive shear fittings in each centerline latch group align door closure and cause the fore and aft shear loads to react once the doors are closed.
The centerline latch group ac reversible motors are automatically turned off by limit switches when the latches are opened or closed. Each motor has a brake that operates similarly to the brakes in the bulkhead motors. When both motors are operating, the nominal operating time is 20 seconds.
If only one motor is operating, the time is 40 seconds. Each mid motor controller assembly has its own timer set to twice the normal operating time to allow single-motor operation of the centerline latch group without causing a sequence fail signal PLB doors CRT message and SM alert.
The power drive unit drives a 55-foot-long \Jtorque\j shaft. The shaft turns the rotary actuators, which causes the push rod, bell crank and link to push the doors open. The same arrangement pulls the doors closed.
The payload bay door opening and closing sequence is controlled automatically through in a predetermined sequence or manually through keyboard entries. The starboard doors must be opened first and closed last due to the arrangement of the centerline latching mechanism and the structural and seal overlap.
Limit switches on each power drive unit turn the ac motors off when the doors are open or closed. Each ac motor has an associated brake that operates similarly to the bulkhead and centerline latch motors. When both motors are operating, the nominal time for payload bay door opening or closing is 63 seconds.
If only one motor is operating, the time is 126 seconds. Each MMCA has its own timer set to twice the normal operating time to allow single-motor operation of the payload bay doors without causing a sequence fail signal PLB doors CRT message and an SM alert .
Torque limiters are incorporated into the rotary actuators to avoid damaging the drive motors or mechanisms if limit switches fail to turn off an electrical drive motor or the mechanisms jam.
Two bolts on the bellcrank and the bolt connecting the link to the rotary actuator can be EVA disconnect points if the linkage fails when the doors close. The power drive unit can be disengaged manually on the ground or on orbit.
The payload bay doors open through an angle of 175.5 degrees.
Two radiator panels on each forward payload bay door can be deployed when the doors are opened on orbit and stowed when the doors are closed before entry, or they can be left in the stowed position for a given flight. Freon-21 coolant loop 1 flows through the left-hand radiator panels, and the No. 2 loop flows through the right-hand panels.
On orbit, the panels radiate excess heat collected by the Freon-21 coolant loops from heat exchangers and cold plates throughout the orbiter. Coolant flows through the radiators from aft to forward. The radiator panels mounted on the forward end of the aft payload bay doors are fixed to the bay doors.
The radiator deploy and stow operation is controlled manually from the aft flight deck panel R13. The PL bay mech (payload bay mechanisms) pwr, radiator latch and radiator control sys switches control the panels. Four indicators show the radiator latch and deploy status.
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"Orbiter Aft Fuselage",282,0,0,0
The aft fuselage consists of an outer shell, thrust structure and internal secondary structure. It is approximately 18 feet long, 22 feet wide and 20 feet high.
The aft fuselage supports and interfaces with the left-hand and right-hand aft orbital maneuvering system/reaction control system pods, the wing aft spar, midfuselage, orbiter/external tank rear attachments, space shuttle main engines, aft heat shield, body flap, vertical tail and two T-0 launch umbilical panels.
The aft fuselage provides the load path to the midfuselage main longerons, main wing spar continuity across the forward bulkhead of the aft fuselage, structural support for the body flap, and structural housing around all internal systems for protection from operational environments (pressure, thermal and acoustic) and controlled internal pressures during flight.
The forward bulkhead closes off the aft fuselage from the midfuselage and is composed of machined and beaded sheet metal aluminum segments. The upper portion of the bulkhead attaches to the front spar of the vertical tail.
The internal thrust structure supports the three SSMEs. The upper section of the thrust structure supports the upper SSME, and the lower section of the thrust structure supports the two lower SSMEs. The internal thrust structure includes the SSMEs, load reaction truss structures, engine interface fittings and the actuator support structure. It supports the SSMEs, the SSME low-pressure turbopumps and propellant lines. The two orbiter/external tank aft attach points interface at the longeron fittings.
The internal thrust structure is composed mainly of 28 machined, diffusion-bonded truss members. In diffusion bonding, \Jtitanium\j strips are bonded together under heat, pressure and time. This fuses the \Jtitanium\j strips into a single hollow, homogeneous mass that is lighter and stronger than a forged part. In looking at the cross section of a diffusion bond, one sees no weld line.
It is a homogeneous parent metal, yet composed of pieces joined by diffusion bonding. (In OV-105, the internal thrust structure is a forging.) In selected areas, the \Jtitanium\j construction is reinforced with boron/epoxy tubular struts to minimize weight and add stiffness. This reduced the weight by 21 percent, approximately 900 pounds.
The upper thrust structure of the aft fuselage is of integral-machined aluminum construction with aluminum frames except for the vertical fin support frame, which is \Jtitanium\j. The skin panels are integrally machined aluminum and attach to each side of the vertical fin to react drag and torsion loading.
The outer shell of the aft fuselage is constructed of integral-machined aluminum. Various penetrations are provided in the shell for access to installed systems. The exposed outer areas of the aft fuselage are covered with reusable thermal protection system.
The secondary structure of the aft fuselage is of conventional aluminum construction except that \Jtitanium\j and fiberglass are used for thermal isolation of equipment. The aft fuselage secondary structures consist of brackets, buildup webs, truss members, and machined fittings, as required by system loading and support constraints.
Certain system components, such as the avionics shelves, are shock-mounted to the secondary structure. The secondary structure includes support provisions for the auxiliary power units, \Jhydraulics\j, ammonia boiler, flash evaporator and electrical wire runs.
The two external tank umbilical areas interface with the orbiter's two aft external tank attach points and the external tank's liquid oxygen and \Jhydrogen\j feed lines and electrical wire runs. The umbilicals are retracted, and the umbilical areas are closed off after external tank separation by an electromechanically operated \Jberyllium\j door at each umbilical. Thermal barriers are employed at each umbilical door. The exposed area of each closed door is covered with reusable surface \Jinsulation\j.
The aft fuselage heat shield and seal provide a closeout of the orbiter aft base area. The aft heat shield consists of a base heat shield of machined aluminum. Attached to the base heat shield are domes of honeycomb construction that support flexible and sliding seal assemblies. The engine-mounted heat shield is of Inconel honeycomb construction and is removable for access to the main engine power heads. The heat shield is covered with a reusable thermal protection system except for the Inconel segments.
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"Orbiter OMS/RCS Pods",283,0,0,0
The orbital maneuvering system/reaction control system left- and right-hand pods are attached to the upper aft fuselage left and right sides. Each pod is fabricated primarily of \Jgraphite\j epoxy composite and aluminum. Each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet.
Each pod is divided into two compartments: the OMS and the RCS housings. Each pod houses all the OMS and RCS propulsion components and is attached to the aft fuselage with 11 bolts. The pod skin panels are \Jgraphite\j epoxy honeycomb sandwich. The forward and aft bulkhead aft tank support bulkhead and floor truss beam are machined aluminum 2124.
The centerline beam is 2024 aluminum sheet with \Jtitanium\j stiffeners and \Jgraphite\j epoxy frames. The OMS thrust structure is conventional 2124 aluminum construction. The cross braces are aluminum tubing, and the attach fittings at the forward and aft fittings are 2124 aluminum. The intermediate fittings are corrosion-resistant steel.
The RCS housing, which attaches to the OMS pod structure, contains the RCS thrusters and associated propellant feed lines. The RCS housing is constructed of aluminum sheet metal, including flat outer skins. The curved outer skin panels are \Jgraphite\j epoxy honeycomb sandwich. Twenty-four doors in the skins provide access to the OMS and RCS and attach points.
The two \Jgraphite\j epoxy pods per \Jspacecraft\j reduce the weight by 10 percent, approximately 450 pounds. The pods will withstand 162-decibel acoustic noise and a temperature range from minus 170 to plus 135 F.
The exposed areas of the OMS/RCS pods are covered with a reusable thermal protection system, and a pressure and thermal seal is installed at the OMS/RCS pod aft fuselage interface. Thermal barriers are installed, and they interface with the RCS thrusters and reusable thermal protection system.
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"Orbiter Body Flap",284,0,0,0
The body flap thermally shields the three SSMEs during entry and provides the orbiter with pitch control trim during its atmospheric flight after entry.
The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. The main upper and lower forward honeycomb skin panels are joined to the ribs, spars and honeycomb trailing edge with structural fasteners. The removable upper forward honeycomb skin panels complete the body flap structure.
The upper skin panels aft of the forward spar and the entire lower skin panels are mechanically attached to the ribs. The forward upper skin consists of five removable access panels attached to the ribs with quick-release fasteners. The four integral-machined aluminum actuator ribs provide the aft fuselage interface through self-aligning bearings.
Two bearings are located in each rib for attachment to the four rotary actuators located in the aft fuselage, which are controlled by the flight control system and the hydraulically actuated rotary actuators. The remaining ribs consist of eight stability ribs and two closeout ribs constructed of chemically milled aluminum webs bonded to aluminum honeycomb core.
The forward spar web is of chemically milled sheets with flanged holes and stiffened beads. The spar web is riveted to the ribs. The trailing edge includes the rear spar, which is composed of piano-hinge half-cap angles, chemically milled skins, honeycomb aluminum core, closeouts and plates. The trailing edge attaches to the upper and lower forward panels by the piano-hinge halves and hinge pins. Two moisture drain lines and one hydraulic fluid drain line penetrate the trailing edge honeycomb core for horizontal and vertical drainage.
The body flap is covered with a reusable thermal protection system and an articulating pressure and thermal seal to its forward cover area on the lower surface of the body flap to block heat and air flow from the structures.
The aft fuselage is built by Rockwell's Space Transportation Systems Division, Downey, Calif. The OMS/RCS pods are built by McDonnell Douglas, St. Louis, Mo. The body flap is built by Rockwell's Columbus, Ohio, division.
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"Orbiter Vertical Tail",285,0,0,0
The vertical tail consists of a structural fin surface, the rudder/speed brake surface, a tip and a lower trailing edge. The rudder splits into two halves to serve as a speed brake.
The vertical tail structure fin is made of aluminum. The main \Jtorque\j box is constructed of integral-machined skins and strings, ribs, and two machined spars. The fin is attached by two tension tie bolts at the root of the front spar of the vertical tail to the forward bulkhead of the aft fuselage and by eight shear bolts at the root of the vertical tail rear spar to the upper structural surface of the aft fuselage.
The rudder/speed brake control surface is made of conventional aluminum ribs and spars with aluminum honeycomb skin panels and is attached through rotating hinge parts to the vertical tail fin.
The lower trailing edge area of the fin, which houses the rudder/speed brake power drive unit, is made of aluminum honeycomb skin.
The hydraulic power drive unit/mechanical rotary actuation system drives left- and right-hand drive shafts in the same direction for rudder control of plus or minus 27 degrees. For speed brake control, the drive shafts turn in opposite directions for a maximum of 49.3 degrees each. The rotary drive actions are also combined for joint rudder/speed brake control. The hydraulic power drive unit is controlled by the orbiter flight control system.
The vertical tail structure is designed for a 163-decibel acoustic environment with a maximum temperature of 350 F.
All-Inconel honeycomb conical seals house the rotary actuators and provide a pressure and thermal seal that withstands a maximum of 1,200 F.
The split halves of the rudder panels and trailing edge contain a thermal barrier seal.
The vertical tail and rudder/speed brake are covered with a reusable thermal protection system. A thermal barrier is also employed at the interface of the vertical stabilizer and aft fuselage.
The contractor for the vertical tail and rudder/speed brake is Fairchild Republic, Farmingdale, N.Y.
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"Orbiter Passive Thermal Control",286,0,0,0
A passive thermal control system helps maintain the temperature of the orbiter \Jspacecraft\j, systems and components within their temperature limits. This system uses available orbiter heat sources and heat sinks supplemented by \Jinsulation\j blankets, thermal coatings and thermal isolation methods. Heaters are provided on components and systems in areas where passive thermal control techniques are not adequate. (The heaters are described under the various systems.)
The \Jinsulation\j blankets are of two basic types: fibrous bulk and multilayer. The bulk blankets are fibrous materials with a density of 2 pounds per cubic foot and a sewn cover of reinforced acrylic film Kapton. The cover material has 13,500 holes per square foot for venting. Acrylic film tape is used for cutouts, patching and reinforcements. Tufts throughout the blankets minimize billowing during venting.
The multilayer blankets are constructed of alternate layers of perforated acrylic film Kapton reflectors and Dacron net separators. There are 16 reflector layers in all, the two cover halves counting as two layers. Covers, tufting and acrylic film tape are similar to that used for the bulk blankets.
The contractors are Hi-Temp \JInsulation\j Inc., Camarillo, Calif. (fibrous insulation); Scheldahl, Northfield, Minn. (cover materials and inner layers); Apex Mills, Los Angeles, Calif. (separators).
#
"Orbiter Purge, Vent and Drain System",287,0,0,0
The purge, vent and drain system on the orbiter is designed to perform the following functions: provide unpressurized compartments with gas purge for thermal conditioning and prevent accumulation of hazardous gases, vent the unpressurized compartments during ascent and entry, drain trapped fluids (water and hydraulic fluid) and condition window cavities to maintain visibility.
Three purge circuits are connected by the T-0 umbilical to ground equipment before launch during the preflight countdown and postlanding phases. Purge gas (cool, dry air and gaseous nitrogen) is provided to three sets of distribution plumbing: the forward fuselage, orbital maneuvering system/reaction control system pods, wings and vertical stabilizer; the midfuselage; and the aft fuselage. The purge gas makes all the unpressurized volumes inert, maintains constant \Jhumidity\j and temperature, forces out any hazardous gases and ensures that external contaminants cannot enter.
The active vent system provides the flow area to control pressure during purge, depressurization during ascent, molecular venting in orbit and repressurization during entry.
The vent and purge system is controlled exclusively through guidance, navigation and control software. The active ports are positioned by the software on the basis of mission time or mission events during ascent, entry and aborts and by crew inputs on the CRT and keyboard in the crew compartment flight deck.
There are 18 active vents in the orbiter fuselage, nine on each side. Each vent has a door that can be positioned for a specific purpose at various phases of flight. For identification, each door is numbered, starting at the nose of the orbiter. Each compartment has a dedicated vent on the left and right side of the orbiter for redundancy.
Internal vents are used to vent compartments that have no vent doors of their own, such as the nose wheel well, the two main wheel wells and the vertical tail section. Passive vents are used to back up vent 7 of the forward wing compartment, which responds to a delta pressure to open a check valve (passive vent) during ascent to vent the wing to the midbody if vent 7 fails; or, on descent, the midfuselage pressurizes the wing if vent 7 fails at a delta pressure of 0.72 to 1 psid. The aft bulkhead (X o 1307) has 14 one-way check valves that vent the payload bay into the aft fuselage at a delta pressure of 0.004 to 0.04 psid. Vent 8 vents the OMS/RCS pods, which are joined by a duct that enables the pod to vent through the opposite side of the vehicle if vent 8 fails to open.
All vent doors are driven by an electromechanical actuator. Vent doors located near each other share common actuators and controls. Vents 1 and 2, 4 and 7, and 8 and 9 share drive mechanisms on the left and right side. The 18 doors are divided into six groups of four ac motors each and are staggered so that all 24 motors do not run at the same time. All vent doors are driven inward, and each door has a pressure seal and thermal seal. The normal opening or closing time of a door with two motors operating is five seconds.
Vent doors 1, 2, 8 and 9 have purge positions that control flow from the forward and aft volumes, respectively. Vent 6 has two purge positions and a closed position that accommodates the different purge flow rates available to the payloads and payload bay. These doors are in the purge position before launch.
Two minutes 20 seconds before launch, the launch processing system reduces the purge flow in anticipation of closing vent 6. At T minus 35 seconds, vent 6 is closed. At T minus 25 seconds, the onboard general-purpose computers are enabled and take over the sequences.
At T minus 10 seconds, the vents are configured for launch. Vents 3, 4, 5, 6 and 7 are closed to limit sound pressure levels in the payload bay. Vents 1, 2, 8 and 9 are opened. If a launch abort occurs from T minus 10 seconds to T minus zero, the vent doors reposition to the prelaunch configuration. At T minus four seconds, any vent door out of configuration and not overridden causes the onboard GPCs to call a hold.
At T plus 10 seconds, all vent doors are commanded open. At T plus 80 seconds, vent doors 8 and 9 are commanded closed to prevent hazardous gases from entering the aft fuselage; at T plus 122 seconds, vents 8 and 9 are commanded open. All vent doors remain open during the remainder of ascent and on-orbit operations.
In preparation for entry, the onboard operational sequence software (OPS 3) closes all vent doors. The doors remain closed until the velocity of the orbiter reaches 2,400 feet per second, when all vents are opened by the onboard GPCs.
At the end of the mission, after the orbiter stops on the runway, vent doors 1, 2, 6, 8 and 9 are configured to their purge positions for ground cooling.
The purge and vent ducting is now made of Kevlar/epoxy (115 pieces up to 11 inches in diameter), which replaced the fiberglass or aluminum ducts and reduced the weight of the ducts 33 percent, or approximately 200 pounds.
The window cavity conditioning system prevents moisture from entering into the windshields and the cavities of the overhead and payload-viewing windows. It also depressurizes and repressurizes these cavities during flight and supplies the purge conditioning to dry them during ground operations. The side hatch window is self-contained.
A hazardous gas detection system detects hazardous levels of explosive or toxic gases. The onboard orbiter sample lines duct the compartment gases to the ground support equipment at the T-0 right-hand umbilical panel and to the ground-based mass spectrometer for analysis at the launch pad.
#
"Orbiter In-Flight Crew Escape System",288,0,0,0
The in-flight crew escape system is provided for use only when the orbiter would be in controlled gliding flight and unable to reach a runway. This condition would normally lead to ditching. The crew escape system provides the flight crew with an alternative to water ditching or to landing on terrain other than a landing site. The probability of the flight crew surviving a ditching is very slim.
The hardware changes required to the orbiters enable the flight crew to equalize the pressurized crew compartment with the outside pressure via the depressurization valve opened by pyrotechnics in the crew compartment aft bulkhead that would be manually activated by a flight crew member in the middeck of the crew compartment; pyrotechnically jettison the crew ingress/egress side hatch manually in the middeck of the crew compartment; and bail out from the middeck through the ingress/egress side hatch opening after manually deploying the escape pole through, outside and down from the side hatch opening.
One by one, each flight crew member attaches a lanyard hook assembly, which surrounds the deployed escape pole, to his or her parachute harness and egresses through the side hatch opening. Attached to the escape pole, the crew member slides down the pole and off the end. The escape pole provides each crew member with a trajectory that takes the crew member below the orbiter's left wing.
Changes were also made in the software of the orbiter's general-purpose computers. The software changes were required for the primary avionics software system and the backup flight system for transatlantic-landing and glide-return-to-launch-site abort modes. The changes provide the orbiter with an automatic-mode input by flight crew members through keyboards at the commander's and/or pilot's panel C3, which provides the orbiter with an automatic stable flight for crew bailout. This software change, which is required to allow the flight crew commander's departure, automatically controls the orbiter's velocity and angle of attack to the desired bailout conditions.
The crew would make the escape decision at an altitude of approximately 60,000 feet and would immediately make an input to the flight control system software autopilot mode.
When the orbiter descends to an altitude of approximately 30,000 feet, its airspeed must be decreased to approximately 200 knots (230 mph). At approximately 25,000 feet, a crew member in the middeck (referred to as the jump master and seated in the forward left seat in the middeck) raises a cover on the left side of the crew compartment middeck at floor level and pulls the T-handle, which activates the pyrotechnics for the depressurization valve at the crew compartment X o 576 aft bulkhead. This equalizes the crew compartment cabin and outside pressure before the side hatch is jettisoned.
At approximately 25,000 feet, the software for the automatic autopilot mode changes the orbiter's angle of attack to approximately 15 degrees. This angle of attack must remain nearly constant for approximately three minutes until the orbiter reaches an altitude of approximately 2,000 feet.
At approximately 25,000 feet, the jump master jettisons the side hatch by pulling the hatch jettison T-handle next to the depressurization T-handle. When the T-handle is pulled, pyrotechnics separate the hatch assembly by severing the side hatch hinge, and three pyrotechnic thrusters jettison the tunnel/hatch from the orbiter at a velocity of approximately 50 feet per second.
The jump master pulls the pip pin on the escape pole and pulls the ratchet handle down, which permits the two telescoping sections of the escape pole to be deployed through the hatch opening by spring tension.
A magazine assembly located near the side hatch contains a lanyard assembly for each flight crew member. Each lanyard assembly consists of a hook attached to a Kevlar strap that surrounds the escape pole. Five roller bearings on each strap surround the pole and permit the lanyard to roll freely down the pole. Each flight crew member positions himself or herself at the hatch opening and attaches himself or herself to the escape pole via the lanyard hook assembly and jumps out the hatch opening.
Each lanyard assembly incorporates an energy absorber rated at 1,000 pounds. The Kevlar strap consists of two sections of permanent Nomex thread stitching and a section of breakaway Kevlar thread stitching. When the crew member exits the side hatch on the escape pole, the breakaway Kevlar thread stitching can break away, providing the crew member with an energy absorber. The crew member slides down the escape pole and off the end into a free-fall. The escape pole extends downward 9.8 feet from the side hatch and provides the crew member with a trajectory that will carry him or her beneath the orbiter's left wing.
It would take approximately 90 seconds for a maximum crew of eight to bail out. After the first crew member bails out from the middeck, the remaining crew members follow at approximately 12-second intervals until all are out by approximately 10,000 feet altitude.
A handhold was added in the middeck next to the side hatch to permit the crew members to position themselves through the side hatch opening for bailout.
The escape pole is constructed of aluminum and steel. The arched housing for the pole is 126.75 inches long and is attached to the middeck ceiling above the airlock hatch and at the 2 o'clock position at the side hatch for deployment during launch and entry. The escape pole telescopes from the middeck housing through the side hatch in two sections. The primary extension is 73 inches long, and the end extension is 32 inches long. The diameter of the housing is 3.5 inches. The two telescoping sections are slightly smaller in diameter. The escape pole weighs approximately 241 pounds-248 pounds with attachments.
On orbit, the escape pole's primary stowage position requires unpinning the escape pole at the starboard and port attachments, rotating the pole so it is flat against the middeck ceiling and strapping it to the ceiling. An alternate on-orbit stowage approach also requires unpinning the escape pole at the starboard and port attachments, rotating it so it is flat against the middeck ceiling and strapping it to the ceiling.
The side hatch water coolant lines for side hatch thermal conditioning were modified to accommodate the installation of the side hatch pyrotechnic separation system.
The flight crew members' seats were also modified to accommodate the seat/crew altitude protection system suit for each crew member.
The pyrotechnically operated crew compartment depressurization valve consists of two flapper valves with debris screens on the crew compartment side and payload bay side that open to depressurize the crew compartment and close when the pressure equalizes.
It is noted that the hatch jettison features could be used in a landing emergency.
The crew member's altitude protection suit includes an emergency oxygen system, pilot and drogue parachutes that are operated automatically and have manual backup, a main parachute that is operated automatically and has manual backup, a seawater activation release system, flotation devices, a life raft and survival equipment. The crew altitude protection suit and its associated equipment weigh approximately 70 pounds.
The side hatch jettison thruster contractor is OEA, \JDenver\j, Colo. The pyrotechnics contractor for the hatch tunnel, hinge and the energy transfer system lines is Explosive Technology, Fairfield, Calif. The escape pole is government-furnished equipment that is supplied by NASA's Johnson Space Center, Houston, \JTexas\j, as is the crew altitude protection suit.
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"Orbiter Emergency Egress Slide",289,0,0,0
The emergency egress slide provides the orbiter flight crew members with a rapid and safe emergency egress through the orbiter middeck ingress/egress side hatch after a normal opening of the side hatch or after jettisoning of the side hatch at the nominal end-of-mission landing site or at a remote or emergency landing site.
The emergency egress slide replaces the emergency egress side hatch bar, which required the flight crew members to drop approximately 10.5 feet to the ground. This drop could cause injury to the flight crew members and prevent an injured flight crew member from moving to a safe distance from the orbiter.
The emergency egress slide will support return- to- launch- site, transatlantic-landing, abort-once-around and normal end-of-mission landings.
The system will be activated manually by the flight crew rotating the slide from the middeck through the egress side hatch opening onto the side hatch if the hatch has not been jettisoned or through the egress side hatch opening if the hatch has been jettisoned. The flight crew pulls a lanyard to inflate the slide with a self-contained air bottle supply.
The slide allows the safe egress of the flight crew members to the ground within 60 seconds after the side hatch is fully opened or jettisoned; accommodates the egress of the flight crew members wearing the launch and entry crew altitude protection system; accommodates the egress of incapacitated crew members; withstands and remains functional in the egress environment for a minimum of six minutes after deployment; and can be released from the side hatch to permit fire truck access.
The slide is installed inside the middeck below the side hatch where it will not inhibit ingress/egress when the system is not required and not interfere with normal on-orbit operations.
The egress slide contractor is Inflatable Systems Inc., a division of OEA, \JDenver\j, Colo.
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"Orbiter Secondary Emergency Egress",290,0,0,0
The left-hand flight deck overhead window provides the flight crew with a secondary emergency egress route. The left overhead window consists of three panes of glass, an inner pane attached to the crew compartment and a center and outer pane attached to the upper forward fuselage.
When the secondary emergency egress path is utilized, pulling the T handle located forward of the flight deck center console (between the commander and pilot) activates the overhead window jettison system. When initiated, the center and outer panes are jettisoned as a unit, upward and aft. A time delay in the pyrotechnic firing circuit delays the initiation of the jettisoning of the inner pane 0.3 of a second after the center and outer panes are jettisoned.
Upon the initiation of the jettisoning of the inner window pane, it rotates downward and aft into the crew compartment aft flight deck on hinges located at the aft portion of the window frame. A capture device attenuates the opening rate and holds the window in position.
The overhead window jettison system consists primarily of expanding tube assemblies, mild detonating fuses, frangible bolts and associated initiators.
The left overhead window jettison system can be initiated from the outside of the orbiter on the right side of the forward fuselage by ground personnel.
Egress steps are mounted at the aft flight deck station (left side) to assist the flight crew up through the window.
Emergency ground descent devices are stowed on the overhead aft flight deck adjacent to the left overhead window. One device is provided for each flight crew member. The emergency ground descent device enables flight crew members to lower themselves to the ground over the side of the orbiter.
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"Orbiter Side Hatch Jettison",291,0,0,0
The middeck ingress/egress side hatch was modified to provide the capability of pyrotechnically jettisoning the side hatch for emergency egress on the ground. In addition, a crew compartment pressure equalization valve provided at the crew compartment aft bulkhead, X o 576, is also pyrotechnically activated to equalize cabin/outside pressure before the jettisoning of the side hatch.
A panel on the left side of the middeck of the crew compartment contains two T-handles. One T-handle controls the initiation of the pyrotechnic pressure equalization valves, which equalize the cabin pressure with outside pressure.
The other T-handle in the same panel in the middeck jettisons the side hatch pyrotechnically. When this T-handle is activated, pyrotechnics sever the hinges of the side hatch and three pyrotechnic tunnel/hatch thrusters are initiated, which jettisons the side hatch from the orbiter.
The side hatch jettison thruster contractor is OEA, \JDenver\j, Colo. The pyrotechnics contractor for the hatch tunnel, hinges and the energy transfer system lines is Explosive Technology, Fairfield, Calif.
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"Shuttle Food System and Dining",292,0,0,0
The middeck of the orbiter is equipped with facilities for food stowage, preparation, and dining for each crew member. The food supply is categorized as either menu food or pantry food. Menu food consists of three daily meals per crew member and provides an average energy intake of approximately 2,700 calories per crew member per day. The pantry food is a two-day contingency food supply that also contains food for snacks and beverages between meals and for individual menu changes.
It provides an average energy intake of 2,100 calories per crew member per day. The types of food include fresh, thermostabilized, rehydratable, irradiated, intermediate-moisture, and natural-form food and beverages.
If a payload is installed in the middeck in lieu of the galley, the food preparation system is limited. It consists of the water dispenser, food warmer, food trays and food system accessories.
The water dispenser provides the flight crew with ambient and chilled water for drinking and reconstituting food. The water dispenser consists of a housing assembly, rehydration station, hygiene water quick disconnect and water lines. Two flex lines 10 feet long connect the housing assembly to the ambient and chilled potable water system. Both lines have quick disconnects. A 12-foot-long flex line with a quick disconnect and water-dispensing valve supplies water for personal hygiene.
The water selector valve "amb" position provides ambient water to the rehydration station between 65 and 75 F. The "off" position prevents water from flowing to the rehydration station (it does not shut off water flow to the personal hygiene water outlet quick disconnect). The "chd" water position provides chilled water to the rehydration station between 45 and 55 F.
Depressing the hygiene water valve handle allows a constant flow of ambient water. Releasing the handle prevents water flow. The locked-open position allows a constant flow of ambient water without holding the handle.
The rehydration station is an electronic dispensing system that interfaces directly with food and beverage packages to provide rehydration capability and drinking water for flight crew members. The system dispenses 2, 3, 4 and 8 ounces of water through a replaceable needle. A spare needle is stowed at the rear of the rehydration unit and another in the in-flight maintenance middeck locker. The needles are removed and installed with a 3/8-inch open-end wrench. Depressing the "pwr" push button at the rehydration station provides power to the electronic rehydration system and an indicating light is illuminated within the switch upon activation.
Depressing the "pwr" push button again deactivates the system. The water quantity rotary switch's 2, 3, 4 and 8 positions provide 2, 3, 4 and 8 ounces of water, respectively. The needle must be inserted into the package before depressing the "fill" push button to prevent free water from being dispensed into the crew cabin environment. Depressing the "fill" push button activates the electronic filling mechanism when the water quantity selection has been made.
A light comes on within the "fill" switch during filling and goes out when filling is complete. The operation is automatically deactivated. The bypass valve provides a continuous flow of water to the food rehydration unit when the handle is depressed or lifted to the "up locked-open" position.
The rehydratable food container is inserted into the rehydration station, the water dispenser needle penetrates the rubber septum on the rehydratable container, and the specified amount of water is discharged into the container. The rehydrated food is mixed and heated, if required.
The rehydrated food container is opened by grasping the center portion of the lid liner with the fingers, piercing the liner with a knife or scissors and pulling the liner up to aspirate air. While grasping the center of the liner, the \Jastronaut\j swings container in a gentle forward and backward semicircular motion to place food contents at the bottom of the container. The inside edge of the lid liner (three sides) is cut with a knife or scissors to expose the food.
The rehydratable beverage container is inserted into the rehydration station, the water dispenser needle penetrates the rubber septum on the rehydratable container, and the specified amount of water is discharged into the container. The rehydratable beverage is mixed and heated, if required. A plastic clip is affixed to the straw in the closed position, the probe end of the straw is inserted into the container rubber septum, the straw is placed in the mouth, the clip is released, and the beverage is drunk. All straws are color-coded for each crew member.
Food trays are kept in a middeck stowage locker (or in the galley, if installed) at launch and are removed and installed in the use locations during preparations for the first meal. The tray is a clear, anodized aluminum sheet that restrains food and accessories during dining. The trays are color-coded for each crew member. Velcro on the bottom of the food trays allows them to be attached to the front of the middeck lockers (or the galley door, if installed) for food preparation or dining.
The straps will also hold the trays on the crew member's leg for dining. A cutout on the tray allows three rehydratable food packages to be secured to the tray. Another cutout with rubber strips adapts to various-sized food packages, including cans, pouches and rehydratable food packages.
Two magnetic strips hold eating utensils and two 0.75-inch-wide binder clips on the tray retain such things as condiment packets and wipes. Accessories used during food preparation and dining include condiments, gum and candy, vitamins, wet wipes, dry wipes, drinking containers, drinking straws, utensils and a re-entry kit that contains salt tablets and long straws.
Condiments include salt, pepper, taco sauce, hot pepper sauce, catsup, mayonnaise and mustard. The salt and pepper are liquids stored in small plastic squeeze bottles. The remaining condiments are packaged in individual, sealed, flexible plastic pouches. Vitamin tablets supplement dietary requirements. Wet wipes are packaged in 21 individual packets per dispenser for cleaning utensils after dining.
A light spring action retains and positions wipes for dispensing. Empty beverage containers of rigid plastic for drinking and storing water are carried in crew members' clothing and can be filled at the water dispenser. Approximately five to 10 different color-coded straws are provided for each crew member (depending on the flight's duration) for drinking beverages and water.
Additional straws are kept in the pantry beverages and various menu locker trays. Color-coded utensils include a knife, two spoons (large and small), a fork and a can opener for each crew member. They are stowed with a soft plastic holder that has a Velcro snap cover. Dry wipes are packaged in a 30-wipe container that can be attached to the crew cabin wall with Velcro for cleanup after dining.
The re-entry kit consists of one package containing eight salt tablets for each crew member and long straws (four per crew member). Two salt tablets are to be taken with 8 ounces of water or other beverage by each crew member four times before entry. The re-entry kit may be stowed in one of three locations depending on space available-with the accessories, near the last meal to be consumed on orbit or with the pantry beverages.
The food warmer is a portable heating unit that can warm a meal for at least four crew members within one hour when the galley is not flown. It is stowed in a middeck locker at launch and is removed and installed during meal preparation activity. The food is heated by thermal \Jconduction\j on a hot plate (element). The warmer is thermostatically controlled between 165 and 175 F.
The case is constructed of aluminum with an exterior envelope of 13 by 18 by 6 inches. It has latches and is lined with clear urethane \Jfoam\j \Jinsulation\j coated with room-temperature vulcanizing compound. The case has straps for handling and on-orbit installation. The exterior contains controls and displays, a power connector that interfaces with the power cable and Velcro attachment. A hinged element is sandwiched between two aluminum plates and is contained by a fiberglass frame.
The aluminum plates have spring-bungee restraints for foil-backed food packages on one side. An on/off switch provides two-phase ac power to the unit and a light indicates the warmer is operating. The power cable is 156 inches long and attaches to a middeck ac utility outlet. The cable is stowed inside the case at launch.
A maximum of 14 packages can be installed on the side of the spring bungees and eight on the other side. Rehydratable beverages should be placed on the side opposite the spring bungees, and the \Jfoam\j on the other side is additionally relieved to prevent the packages from popping out in zero gravity.
When 14 rehydratable packages are heated, no foil-backed food pouches can be heated. A maximum of six foil-backed food pouches can be heated in conjunction with 12 rehydratable packages. When foil-backed pouches are heated, only four rehydratable packages can be heated on the side of the spring bungees. The foil-backed pouches are stacked three deep. Four rehydratable packages are inserted in the outer recessed \Jfoam\j cutouts. At the bottom of the cutouts, 0.5-inch-thick uncoated \Jfoam\j absorbs moisture or spilled liquids.
The galley is a multipurpose facility that provides a centralized location for one individual to handle all food preparation activities for a meal. The galley has facilities for heating food, rehydrating food and stowing food system accessories and food trays. The galley consists of a rehydration station, oven, food trays and food system accessories.
The oven is divided into two principal compartments-a lower compartment designed for heating at least 14 rehydratable food containers inserted on tracks and an upper compartment designed to accept a variety of food packages, including the rehydratable containers. At least seven food pouches can be heated in the upper compartment and are held against the heat sink by four spring-loaded plates. The oven has a heating range of 145 to 185 F. During launch and entry, the oven door is held closed by a restraining strap, which is removed from the door by releasing the snap for on-orbit operations. An on/off switch enables and removes power to three fans.
The rehydration station dispensing system interfaces directly with food and beverage packages, providing rehydration capability and drinking water for crew members. A gauge indicates hot water temperatures of 100 to 220 F. The volume/ounces switch selects the volume of water to be dispensed by the rehydration station in 0.5-ounce increments from 0.5 of an ounce to 8 ounces.
The yellow hot push button indicator allows hot water to be dispensed when it is depressed and is illuminated when energized. When the selected volume of water has been dispensed, the push button will begin to flash on and off. The light will be extinguished when the food package is retracted, releasing the \Jhydration\j station lever arm/limit switch.
The rehydration station lever arm/limit switch serves as an interlock so water can be dispensed only when a food package is connected to the needle. The food package makes contact with the rehydration station lever, which activates the limit switch (note that the flight crew does not physically actuate the lever).
The blue cold push button indicator allows cold water to be dispensed when it is depressed and is illuminated when energized. When the selected volume of water has been dispensed, the push button will begin to flash on and off. The light will be extinguished when the food package is retracted, releasing the rehydration station lever arm/limit switch.
The galley light is located on the upper left-hand side of the galley structure surface and has a single light brightness control. Moving the knob clockwise from off applies power to the light and provides variable brightness control.
Two condiment dispensers are attached to the galley by Velcro tabs on the back of the dispensers. The dispensers are available for holding individual packets, such as catsup, taco sauce, mayonnaise and mustard, on the front panel below the oven. The dispensers are open-ended boxes designed to hold the stack of packets together so they may be individually removed as needed. A slide plate keeps the packets from becoming loose as the items are depleted.
A single dispenser for holding individual packets of wet wipes is located on the front panel below the oven and is slightly different in design than the condiment dispensers.
Dispensers for liquid salt and pepper and vitamins can be restrained by clips conveniently located below the rehydration station.
Food trays and food system accessories are the same as those used on flights without the galley.
On the upper left-hand corner of the galley behind a Teflon cloth panel is an MV3 valve that has "emer off" and "on" positions. The "on" position serves as the nominal open position of the manual shutoff valve, and "emer off" serves as a manual shutoff valve for the ambient temperature water supply to the galley.
On the upper right-hand corner of the galley behind a Teflon cloth panel are test connectors, a dc power bus A and B switch and a flush port quick-disconnect test port. The two test connectors serve as a hookup for ground support equipment. The dc power bus A switch's "on" position activates the galley oven heaters, rehydration station system and one of six water tank strip heaters; the "off" position deactivates the heaters. The dc power bus B switch's "on" position activates five of the six galley water tank strip heaters, and the "off" position deactivates them. The flush port quick disconnect serves as the galley water system GSE flush port.
On the lower left-hand side of the galley is an auxiliary port water quick disconnect that allows the crew members to obtain ambient potable water when the MV3 valve is off or hygiene water when the 12-foot flex line and water dispensing valve are attached to the quick disconnect.
Three one-hour meal periods are scheduled for each day of the mission. This hour includes actual eating time and the time required to clean up. Breakfast, lunch and dinner are scheduled as close to the usual hours as possible. Dinner is scheduled at least two to three hours before crew members begin preparations for their sleep period.
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"Shuttle Orbiter Medical System",293,0,0,0
The shuttle orbiter medical system is required to provide medical care in flight for minor illnesses and injuries. It also provides support for stabilizing severely injured or ill crew members until they are returned to Earth. The SOMS consists of two separate packages: the medications and bandage kit and the emergency medical kit. The MBK is blue and the EMK is also blue with red Velcro.
The medical kits are stowed in a modular locker in the middeck of the crew compartment. If the kits are required on orbit, they are unstowed and installed on the locker doors with Velcro.
Each kit contains pallets. The MBK pallet designators are D, E and F. The D pallet contains oral medications consisting of pills, capsules and suppositories. The E pallet contains bandage materials for covering or immobilizing body parts. The F pallet contains medications to be administered by topical application.
The EMK pallet designators are A, B, C and G. The A pallet contains medications to be administered by injection. The B pallet contains items for performing minor surgeries. The C pallet contains diagnostic/therapeutic items consisting of instruments for measuring and inspecting the body. The G pallet contains a microbiological test kit for testing for bacterial infections.
The diagnostic equipment on board and information from the flight crew will allow diagnosis and treatment of injuries and illnesses through consultation with flight surgeons in the Mission Control Center in Houston.
The operational bioinstrumentation system provides an amplified electrocardiograph analog signal from either of two designated flight crew members to the orbiter avionics system, where it is converted to digital tape and transmitted to the ground in real time or stored on tape for dump at a later time. The designated flight crew members wear the OBS during the ascent and entry phases. On-orbit use will be limited to contingency situations.
The OBS electrodes are attached to the skin with electrode paste to establish electrical contact. The electrode is composed of a plastic housing containing a non-polarizable pressed pellet. The housing is attached to the skin with double-sided adhesive tape and the pellet contacts the skin. There are three electrodes on the harness marked LC (lower chest), UC (upper chest) and G (ground).
The ECG signal conditioner is a hybrid microcircuit with variable gain (adjusted for each crew member before flight). It provides a zero- to 5-volt output and has an on/off switch within the input plug, which is actuated when the intravehicular activity biomed cable is plugged in. The unit has batteries that will not be replaced in flight.
The IVA biomed cable connects to the signal conditioner and is routed under the IVA clothing to connect to the biomed seat cable. The biomed seat cable is routed to one of the biomed input connectors located on panel A11, A15 or M062M. Rotary control switches on panel R10 provide circuits from the biomed outlets to the orbiter's network signal processor for downlink or recording. The two rotary switches on panel R10 are biomed channel 1 and channel 2 . Extravehicular activity positions provide circuits for the EVA UHF transceiver.
The electrode application kit contains components to aid in the application of electrodes. The components include wet wipes, double-sided adhesive tape, overtapes, electrode paste and a cue card illustrating electrode placement.
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"Shuttle Radiation Equipment",295,0,0,0
The harmful biological effects of radiation must be minimized through mission planning based on calculated predictions and monitoring of dosage exposures. Preflight requirements include a projection of mission radiation dosage, an assessment of the probability of solar flares during the mission and a radiation exposure history of flight crew members. In-flight requirements include the carrying of passive dosimeters by the flight crew members and, in the event of solar flares or other radiation contingencies, the readout and reporting of the active dosimeters.
There are four types of active dosimeters: pocket dosimeter high, pocket dosimeter low, pocket dosimeter FEMA and high-rate dosimeter. All four function in the same manner and contain a \Jquartz\j fiber positioned to zero by electrostatic charging before flight. The unit discharges according to the amount of radiation received; and as the unit discharges, the \Jquartz\j moves. The position of the fiber along a scale is noted visually. The PDH unit's range is zero to 100 rads. The PDF and PDL units' ranges are zero to 200 millirads and the HRD unit's range is zero to 600 rads.
The rad is a unit based on the amount of energy absorbed and is defined as any type of radiation that is deposited in the absorbing media, and radiation absorbed by man is expressed in roentgen equivalent in man, or rems. The rem is determined by multiplying rads times a qualifying factor that is a variable depending on wavelength, source, etc. For low-inclination orbits (35 degrees and lower), the qualifying factor is approximately equal to one; therefore, the rem is approximately equal to the rad. In space transportation system flights, the doses received have ranged from 0.05 to 0.07 rem, well below flight crew exposure limits.
The flight crew's passive dosimeters are squares of fine-ground photo film sandwiched between plastic separators in a light-proof package. Radiation striking the silver halide causes spots on the film, which can be analyzed after the flight. Included in the badge dosimeters are thermoluminescent dosimeter chips, which are analyzed on Earth.
Passive radiation dosimeters are placed in the crew compartment before launch by ground support personnel and removed after landing for laboratory analysis. Each flight crew member carries a passive dosimeter at all times during the mission. The remaining dosimeters are stowed in a pouch in a middeck modular locker. If a radiation contingency arises, the PDL, PDH, HRD and PDF active dosimeters will be unstowed, read, and recorded for downlink to the ground.
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"Shuttle Crew Apparel",296,0,0,0
During launch and entry, crew members wear the crew altitude protection system, which consists of a helmet; communications cap; pressure garment; anti-exposure, anti-gravity suit; gloves; and boots.
The crew wears escape equipment over the CAPS during launch and entry. It consists of an emergency oxygen system; parachute harness, parachute pack with automatic opener, pilot chute, drogue chute and main canopy; a life raft; 2 liters of emergency drinking water; flotation devices; and survival vest pockets containing a radio/beacon, signal mirror, shroud cutter, pen gun flare kit, sea dye marker, smoke flare and beacon. Manual activation of the parachute automatic opening sequences is provided, as well as manual release of the parachute main canopy.
On orbit, optional clothing and equipment include underwear, urine collection devices, eyeglasses, communications headset, emesis bag, flashlight, Swiss army knife, kneeboard, pens and pencils, stowage bags, watches and food and drink containers.
Crew clothing and equipment used during on-orbit activities include flight suits, IVA trousers, IVA jackets, IVA shirts, sleep shorts, IVA soft slippers, underwear, scissors/lanyard, pocket dosimeter and pocket food.
Crew clothing is designed for use by 90 percent of the male and female population, the 5th to 95th percentile.
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"Shuttle Sleeping Provisions",297,0,0,0
Sleeping provisions for flight crew members consist of sleeping bags, sleep restraints or rigid sleep stations. The sleeping arrangements can consist of a mix of bags and sleep restraints or rigid sleep stations on a given mission. During a mission with one shift, all crew members sleep simultaneously. If all crew members sleep simultaneously, at least one crew member will wear a communication headset to ensure reception of ground calls and orbiter caution and warning alarms.
If sleeping bags are used, they are installed on the starboard middeck wall and deployed for use on orbit.
If the rigid sleep station is used on a mission, it is installed on the starboard side of the middeck. There are two types of rigid sleep stations. One sleep station type accommodates three crew members and the other accommodates four.
If the rigid sleep station is not installed for a mission, a sleeping bag is furnished each crew member. Each sleeping bag contains a support pad with adjustable restraining straps and a reversible/removable pillow and head restraint. Apollo sleeping bags may be provided for the crew members on request. The Apollo sleeping bag is constructed of beta material and is perforated for thermal comfort.
Six adjustable straps permit the sleeping bag to be adjusted to its proper configuration. Three helical springs above the adjustable straps on one side of the bay relieve loads exerted by the crew member on the crew compartment structure. Six pip pins allow the bag to be attached to the middeck locker face in either a horizontal or vertical configuration. Two elastic adjustable straps restrain the upper and lower parts of the body in the bag. Velcro strips on the ends of both sides of the head restraint attach it to the pillow.
A double zipper arrangement permits the sleeping bag to be opened and closed from the bottom to the top of the bag. One zipper on each side of the sleeping bag allows the bag to be attached to a support pad for better rigidity.
The Apollo beta cloth sleeping bag has four adjustable straps with pip pins that are connected to any two lockers in the middeck separated by a distance equal to a four-tiered locker configuration. For torso restraint, a single two-piece strap is provided and a single zipper opens the bag. The bags are stowed in a middeck locker during launch and entry.
A sleep kit is provided for each crew member and is stowed in the crew member's clothing locker during launch and entry. Each kit contains eye covers and ear plugs for use as required during the sleep period.
The three- or four-tier rigid sleep stations contain a sleeping bag, personal stowage provisions, a light and a ventilation inlet and outlet in each of the tiers. The cotton sleeping bag is installed on the ground in each tier and held in place by six spring clips. The light in each tier is a single fluorescent fixture with a brightness control knob and an off position. The air ventilation inlet duct is an air diffuser similar to an \Jautomobile\j ventilation duct. It is adjusted by moving the vane control knob.
The air ventilation outlet duct is located in the fixed panel at each tier and is opened or closed by moving the vane control knob. The air inlet is located at the crew member's head. The outlet is at the feet. All crew members' heads are toward the airlock and their feet toward the avionics bay.
In the three-tier configuration, the upper and middle crew members face the ceiling and the lower tier crew member faces the floor. The fixed panel at the lower sleep station is removable to provide access to the cabin debris trap door for cleaning the cabin filter, to gain access to floor locker MD76C and to enter the forward portion of the lower equipment bay to clean the avionics bay fan filter.
In the four-tier configuration, the bottom tier sleep restraint hookup provision allows the crew member to position himself at a 15-degree angle, which provides more room, or in the normal horizontal position. The sleeping bag, personal stowage provisions, light and ventilation inlet and outlet are the same as in the three-tier configuration. The head and feet orientations of the crew members are also the same as in the three-tier configuration. The lowest tier is removable so access can be obtained to the cabin debris trap door to clean the cabin filter, gain access to floor locker MD76C and enter the forward portion of the lower equipment bay to clean the avionics bay fan filter.
The three-tier rigid sleep station is made of plastic honeycomb panel and weighs approximately 205 pounds. The four-tier rigid sleep station is made of metal and weighs 173 pounds.
A 24-hour period is normally divided into an eight-hour sleep period and a 16-hour wake period for each crew member. Forty-five minutes are allocated for the crew members to prepare for the sleep period and another 45 minutes when they awake to wash and get ready for the day.
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"Shuttle Personal Hygiene Provisions",298,0,0,0
To maintain good hygiene and appearance, personal hygiene and grooming provisions are furnished for both male and female flight crew members. Water is provided by the water dispensing system.
A personal hygiene kit is furnished each crew member for brushing teeth, hair care, shaving, nail care, etc. A kit is also furnished with articles essential to female hygiene and grooming.
Two washcloths and one towel per crew member per day are provided in addition to two paper tissue dispensers per crew member for each seven days. The washcloths are 12 by 12 inches and the towels 16 by 27 inches. The tissues are absorbent, multi-ply, low-linting paper. Rubber restraints with a Velcro base allow the crew members to restrain their towels and washcloths on the waste management door or middeck walls.
The personal hygiene provisions are stowed in middeck stowage lockers at launch and are removed for use on orbit.
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"Shuttle Housekeeping",299,0,0,0
In addition to time scheduled for sleep periods and meals, each crew member has housekeeping tasks that require from five to 15 minutes of his time at intervals throughout the day. These include cleaning the waste management compartment, the dining area and equipment, floors and walls (as required), the cabin air filters; trash collection and disposal; and changeout of the crew compartment carbon dioxide (lithium hydroxide) absorber canisters.
The materials and equipment available for cleaning operations are biocidal cleanser, disposable gloves, general-purpose wipes and a vacuum cleaner. The cleaning materials and vacuum are stowed in middeck lockers. The vacuum cleaner is powered by the orbiter's electrical power system.
The biocidal cleanser is a liquid detergent formulation in a container approximately 2 inches in diameter and 6 inches long. The container has a built-in bladder, dispensing valve and nozzle. The cleanser is sprayed on the surface to be cleaned and wiped off with dry general-purpose wipes. It is used for periodic cleansing of the waste collection system urinal and seat and the dining area and equipment. It is also used, as required, to clean walls and floors. Disposable plastic gloves are worn while using the biocidal cleanser.
General-purpose wipes are also used for general-purpose cleaning.
The vacuum cleaner is provided for general housekeeping and cleaning of the crew compartment air filters and Spacelab filters (on Spacelab missions). It has a normal hose, extension hose and several attachments. It is powered by the orbiter dc electrical power system.
Trash management operations include routine stowage and daily collection of wet and dry trash, such as expended wipes, tissues and food containers. Wet trash includes all items that could offgas. The equipment available for trash management includes trash bags, trash bag liners, wet trash containers and the stowable wet trash vent hose.
Three trash bags are located in the crew compartment. Each bag contains a disposable trash bag liner. Two bags are designated for dry trash and one for wet trash. At a scheduled time each day, the trash bag liner for dry trash is removed from its trash bag. The liner is closed with a strip of Velcro and stowed in an empty locker.
When more than 8 cubic feet of wet trash is expected, the trash bag liners for wet trash are removed at a scheduled time each day and placed in a wet trash container. The container is then closed with a zipper and the unit is stowed. If expansion due to offgassing is evident, the container is connected to a vent in the waste management system for overboard venting of the gas.
The wet trash container is made of airtight fabric and is closed with a seal-type slide fastener. The container has a volume of approximately 0.7 cubic foot and has an air inlet valve on one end and a quick disconnect on the other end. It is attached to the waste management vent system beneath the commode, enabling air to flow through the wet trash container and then overboard. It is attached through a 41-inch- long vent hose filter. When the container is full, it is removed and stowed in a modular locker.
An 8-cubic-foot wet trash stowage compartment is available under the middeck floor. Each day, the trash bag liners for wet trash are removed from the trash bags and stowed in the wet trash stowage compartment, which is vented overboard. If the compartment becomes full, the trash bag liners for wet trash are stowed in wet trash containers.
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"Shuttle Sighting Aids",300,0,0,0
Sighting aids include all items used to aid the flight crew within and outside the crew compartment. The sighting aids include the crewman optical alignment sight, binoculars, adjustable mirrors, spotlights and eyeglasses.
The COAS is a collimator device similar to an \Jaircraft\j gunsight. Two are installed in the crew compartment flight deck. One COAS is mounted during launch and entry over the positive X commander's forward window and on orbit is removed and mounted next to the aft flight deck overhead right negative Z window. The other COAS is mounted at the aft flight deck station for checking the alignment of the payload bay doors.
When the COAS is mounted at the commander's station, it allows the viewers to reassure themselves of proper attitude orientation during the ascent and deorbit thrusting periods. When the COAS is removed from the commander's station to the aft flight deck for on-orbit operations, it provides a backup to the orbiter star trackers for inertial measurement unit alignment. It is also used as the primary optical instrument for measuring range and rotational rates and allows the flight crew members to align the vehicles and dock.
The COAS consists of a lamp with an intensity control, a reticle, a barrel-shaped housing, a mount, a combiner assembly and a power cable. The reticle consists of a 10-degree circle, vertical and horizontal cross hairs with 1-degree marks, and an elevation scale on the right side of minus 10 degrees to 31.5 degrees.
For IMU alignments, the flight crew member at the aft flight deck station maneuvers the orbiter using the COAS at the right overhead negative Z window until the selected star is in the field of view. The crew member continues maneuvering the orbiter until the star crosses the center of the reticle.
At the instant of crossing, the crew member makes a mark, which means he depresses the "att ref" (attitude reference) push button. At the time of the mark, software stores the gimbal angles of the three IMUs. The mark can be taken again if it is felt the star was not centered as well as it could have been. When the crew member feels a good mark was taken, the software is notified to accept it. Good marks for two stars are required for an IMU alignment.
By knowing the star being sighted and the COAS location and mounting relationship in the orbiter, software can determine a line-of-sight vector from the COAS to the star in an inertial coordinate system. Line-of-sight vectors to two stars define the attitude of the orbiter in inertial space. This attitude can be compared to the attitude defined by the IMUs, and if the IMUs are in error, they can be realigned to the more correct orientation by the COAS sightings.
The COAS requires 115-volt ac power for reticle illumination. The COAS is 9.5 by 6 by 4.3 inches and weighs 2.5 pounds.
The 10-by-40 binoculars are a space-modified version of the commercial Leitz Trinovid binocular noted especially for its small size, high \Jmagnification\j, wide field of view, and rugged sealed construction. The 7-by-35 binoculars are noted for close focal distance at high \Jmagnification\j. The 14-by-40 gyrostabilized binoculars contain a gyrostabilized system that enhances target acquisition and retention.
When the crew member is subjected to ambient vibrations or hand tremor while using the gyrostabilized binoculars, the target image remains clear and stable. The gyrostabilized binoculars are electrically powered by six alkaline-type AA batteries and will operate continuously up to three hours on one battery pack.
Adjustable mirrors are installed before launch on handholds located between windows 2 and 3 for the commander and windows 4 and 5 for the pilot. During ascent and entry, the commander and pilot use the adjustable mirrors to better see controls that are in obscured areas of their vision. On orbit, the mirrors can be removed and stowed if desired. Each mirror is approximately 3 by 5 inches and weighs approximately 1 pound.
The spotlight is a high-intensity, hand-held flashlight powered by a battery pack consisting of five 1.2-volt one-half D size nickel-cadmium batteries. The spotlight produces a 20,000-candlepower output with a continuous running time of 1.5 hours. The lamp is a 6-volt \Jtungsten\j filament and cannot be replaced in flight. A spare battery pack is available on board.
For those crew members requiring them, two pairs of eyeglasses are available on board.
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"Shuttle Microcassette Recorder",301,0,0,0
The microcassette recorder is flown primarily for voice recording of data but may also be used to play prerecorded tapes. A microcassette tape has a recording time of 30 minutes per side. It is powered by two 1.5-volt AAA alkaline batteries.
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"Shuttle Photographic Equipment",302,0,0,0
Three camera systems-16mm, 35mm and 70mm-are used by the flight crew to document activities inside and outside the orbiter. All three camera systems are used to document on-orbit operations. The 16mm camera is also used during the launch and landing phases of the flight.
The 16mm camera is like a motion picture camera with independent shutter speeds and frame rates. The camera can be operated in one of three modes: pulse, cine, or time exposure. In the pulse mode, the camera operates at a continuous frame rate of two, six or 12 frames per second. In the cine mode, the camera operates at 24 frames per second. In the time exposure mode, the first switch actuation opens the shutter and the second actuation closes it. The camera uses 140-foot film magazines and has 5mm, 10mm and 18mm lenses.
The 35mm camera is a motorized, battery-operated Nikon camera with reflex viewing, through-the-lens coupled light metering and automatic film advancement. The camera has the standard manual operation and three automatic (electrically controlled) modes-single exposure, continuous and time. It uses an f/1.4 lens.
The 70mm camera system is a modified battery-powered, motor-driven, single-reflex Hasselblad camera that has 80mm and 250mm lenses and film magazines. Each magazine contains approximately 80 exposures. This camera has only one mode of operation, automatic; however, there are five automatic-type camera functions from which to select. The camera has a fixed viewfinder for through-the-lens viewing.
Interdeck light shades are provided to minimize light leakage between the flight deck and middeck during in-cabin photography. The light shade is attached with Velcro to the middeck ceiling around the interdeck access. Adjustable louvers are provided to regulate the amount of light between the flight deck and middeck.
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"Shuttle Wicket Tabs",303,0,0,0
Wicket tabs are devices that help the crew member activate controls when his vision is degraded. The tabs provide the crew member with tactile cues to the location of controls to be activated as well as a memory aid to their function, sequence of activation and other pertinent information. Wicket tabs are found on controls that are difficult to see during the ascent and entry flight phases on panels O8, C3 and R2.
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"Shuttle Reach Aid",304,0,0,0
The reach aid, sometimes known as the ''swizzle stick,'' is a short adjustable bar with a multipurpose end effector that is used to actuate controls that are out of the reach of seated crew members. The reach aid is used to push in and pull out circuit breakers and move toggle switches. It may be used during any phase of flight, but is not recommended for use during ascent because of the attenuation and switch-cueing difficulties resulting from acceleration forces.
Operation of the reach aid consists of extending it and actuating controls with the end effector. To extend the reach aid, one depresses the spring-loaded extension tab and pulls the end effector out to the desired length.
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"Shuttle Restraints and Mobility Aids",305,0,0,0
Restraints and mobility aids are provided in the orbiter to enable the flight crew to perform all tasks safely and efficiently during ingress (1-g, orbiter vertical), egress (1-g, orbiter horizontal) and orbital flight (orbiter orientation arbitrary). Restraints and mobility aids consist of foot loop restraints, the airlock foot restraint platform and the work/dining table. In-flight restraints consist of temporary stowage bags, Velcro, tape, snaps, cable restraints, clips, bungees and tethers.
Mobility aids and devices consist of handholds, footholds, handrails, ladders and the ingress-egress platform.
Foot loop restraints are cloth loops attached to the crew compartment decks by adhesive to secure crew members to the deck. Before launch, the foot loop restraints are installed on the floor areas of the aft flight deck work stations, middeck lockers, waste collection system and galley (if installed).
Spares will be stowed in the modular lockers. To install a foot restraint, the protective backing on the underside of the restraint is removed and the restraint is placed in its desired location. The foot loop restraints are easily used by placing one or both feet in the loop.
The temporary stowage bag is used to restrain, stow or transport loose equipment temporarily. It is snapped or attached with Velcro to the crew station standard Velcro and snap patterns.
Mobility aids and devices are located in the crew compartment for movement of the flight crew members during ingress, egress and orbital flight. These devices consist of handholds for ingress and egress to and from crew seats in the launch and landing configuration, handholds in the primary interdeck access opening for ingress and egress in the launch and landing configuration, a platform in the middeck for ingress and egress to and from the middeck when the orbiter is in the launch configuration, and an interdeck access ladder to enter the flight deck from the middeck in the launch configuration and go from the flight deck to the middeck in the launch and landing configuration.
The flight data file is a flight reference data file that is readily available to crew members aboard the orbiter. It consists of the onboard complement of documentation and related crew aids and includes documentation, such as procedural checklists (normal, backup and emergency procedures), malfunction procedures, crew activity plans, schematics, photographs, cue cards, star charts, Earth maps and crew notebooks; FDF stowage containers; and FDF ancillary equipment, such as tethers, clips, tape and erasers.
Four permanently mounted containers are located to the left and right side of the commander's and pilot's seats for stowing FDFs on the flight deck. The remaining FDF items are stowed in a middeck modular stowage locker.
The flight data file quantity and stowage locations are similar for all flights. The baseline stowage volume is sufficient to contain all FDF items for all orbiter configurations except the pallet-mounted payload. In this case, a larger flight data file and, consequently, additional locker space are required because all payload operations are performed in the orbiter.
FDF items are used throughout the flight-from prelaunch use of the ascent checklist through crew use of the entry checklist.
Flight data files are packaged and stowed on an individual flight basis. FDF items will be stowed in five types of stowage containers: lockers, the flight deck module, the commander's and pilot's seat-back FDF assemblies, the middeck FDF assembly and the map bag. The portable containers are stowed in a middeck modular locker for launch and entry.
If the flight carries a Spacelab module, all Spacelab books are stowed for launch in a portable container on the middeck and transferred in flight to the Spacelab. The FDF stowage is flexible and easily accessible.
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"Shuttle Crew Equipment Stowage",306,0,0,0
Crew equipment on board the orbiter is stowed in lockers with insertable trays. The trays can be adapted to accommodate a wide variety of soft goods, loose equipment and food. The lockers are interchangeable and attach to the orbiter with crew fittings. The lockers can be removed or installed in flight by the crew members. There are two sizes of trays: a half-size tray (two of which fit inside a locker) and a full-size tray. Approximately 150 cubic feet of stowage space is available, almost 95 percent of it on the middeck.
The lockers are made of either epoxy- or polyimide-coated Kevlar honeycomb material joined at the corners with aluminum channels. Inside dimensions are approximately 10 by 17 by 20 inches. The honeycomb material is approximately 0.25 of an inch thick and was chosen for its strength and light weight. The lockers contain about 2 cubic feet of space and can hold up to 60 pounds.
Dividers are used in the trays to provide a friction fit for zero-g retention. This will reduce the necessity for the straps, bags, Velcro snaps and other cumbersome attach devices previously used. Soft containers will be used in orbiter spaces too small for the fixed lockers.
The trays are packed with gear so that no item covers another type of gear. This method of packing will reduce the confusion usually associated with finding loose equipment and maintaining a record of the equipment.
Stowage areas in the orbiter crew compartment are located in the forward flight deck, the aft flight deck, the middeck, the equipment bay and the airlock module.
In the aft flight deck, stowage lockers are located below the rear payload control panels in the center of the deck. Container modules can be mounted to the right and left of the payload control station. Since these side containers are interchangeable, they may not be carried on every mission, depending on any payload-unique installed electronic gear.
In the middeck, container modules can be inserted in the forward avionics bay. Provisions for 42 containers are available in this area. In addition, there is an area to the right side of the airlock module where nine containers can be attached.
Harness stowage bags stowed in a middeck stowage locker or airlock are used on orbit to stow flight crew members' launch equipment, such as helmets, harnesses, boots and waste/trash materials.
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"Shuttle Exercise Equipment",307,0,0,0
The only exercise equipment presently being flown is a treadmill. The exact stowage location in the crew compartment middeck for launch, orbit and entry depends on the mission.
The treadmill is used with a restraint system to allow a crew member to run or jog in orbit. The treadmill kit is stowed on top of the treadmill and contains the waist belt, two shoulder straps, four extender hooks and a physiological monitor. The treadmill kit is restrained by four force cords that are used to restrain the body during exercise. The treadmill attaches to four middeck quick disconnects.
The quick disconnects contain several metal hooks that are hinged within the quick disconnect and actuated by the knurled lock ring. To release the quick disconnects, the lock push button is depressed and the knurled lock ring is pushed up, releasing the metal hooks. When the lock ring is pushed down, the metal hooks converge and capture the top of the middeck stud.
The treadmill has a speed control knob, which controls a rapid onset braking system. When the preset speed is reached, the brake engages and produces increased drag on the running track.
The physiological monitor provides heart rate, the time run and the distance run. The heart rate is determined by an ear clip, which has an infrared sensor that detects increased blood flow (pulses) in the ear lobe. Distance run is determined by connecting a mechanical sensor wire on the side of the treadmill to the physiological monitor.
The mechanical sensor detects the number of revolutions of the track and sends an electrical signal to the physiological monitor, where the distance is computed and shown on the display along with the heart rate. The monitor is stowed on the treadmill handle while the crew member runs.
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"Shuttle Sound Level Meter",308,0,0,0
The sound level meter is provided to determine on-orbit acoustical noise levels in the cabin. Depending on the requirements for each flight, the flight crew is required to take meter readings at specified crew compartment and equipment locations. The data obtained by the flight crew is logged and/or voice recorded. The meter is operated by four 1.5-volt batteries.
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"Shuttle Air Sampling System",309,0,0,0
The air sampling system consists of air bottles that are stowed in a modular locker. They are removed for sampling and restowed for entry.
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"Shuttle Main Propulsion System",310,0,0,0
The main propulsion system, assisted by the two solid rocket boosters during the initial phases of the ascent trajectory, provides the velocity increment from lift-off to a predetermined velocity increment before orbit insertion. The two SRBs are jettisoned after their fuel has been expended, but the MPS continues to thrust until the predetermined velocity is achieved. At that time, main engine cutoff is initiated.
The external tank is jettisoned, and the orbital maneuvering system is ignited to provide the final velocity increment for orbital insertion. The magnitude of the velocity increment supplied by the OMS depends on payload weight, mission trajectory and system limitations.
Coincident with the start of the OMS thrusting maneuver (which settles the MPS propellants), the remaining liquid oxygen propellant in the orbiter feed system and space shuttle main engines is dumped through the nozzles of the three SSMEs. At the same time, the remaining liquid \Jhydrogen\j propellant in the orbiter feed system and SSMEs is dumped overboard through the \Jhydrogen\j fill and drain valves for six seconds.
Then the \Jhydrogen\j inboard fill and drain valve is closed, and the \Jhydrogen\j recirculation valve is opened, continuing the dump. The \Jhydrogen\j flows through the engine \Jhydrogen\j bleed valves to the orbiter \Jhydrogen\j MPS line between the inboard and outboard \Jhydrogen\j fill and drain valves, and the remaining \Jhydrogen\j is dumped through the outboard fill and drain valve for approximately 120 seconds.
During on-orbit operations, the flight crew vacuum inerts the MPS by opening the liquid oxygen and liquid \Jhydrogen\j fill and drain valves, which allows the remaining propellants to be vented to space.
Before entry, the flight crew repressurizes the MPS propellant lines with \Jhelium\j to prevent contaminants from being drawn into the lines during entry and to maintain internal positive pressure. MPS \Jhelium\j is also used to purge the \Jspacecraft\j's aft fuselage. The last activity involving the MPS occurs at the end of the landing rollout. At that time, the \Jhelium\j remaining in onboard \Jhelium\j storage tanks is released into the MPS to provide an inert atmosphere for safety.
The MPS consists of the following major subsystems: three SSMEs, three SSME controllers, the external tank, the orbiter MPS propellant management subsystem and \Jhelium\j subsystem, four ascent thrust vector control units, and six SSME hydraulic servoactuators.
The main engines are reusable, high-performance, liquid-propellant rocket engines with variable thrust. The propellant fuel is liquid \Jhydrogen\j and the oxidizer is liquid oxygen. The propellant is carried in separate tanks in the external tank and supplied to the main engines under pressure. Each engine can be gimbaled plus or minus 10.5 degrees in the yaw axis and plus or minus 10.5 degrees in the pitch axis for thrust vector control by hydraulically powered gimbal actuators.
The main engines can be throttled over a range of 65 to 109 percent of their rated power level in 1-percent increments. A value of 100 percent corresponds to a thrust level of 375,000 pounds at sea level and 470,000 pounds in a vacuum. A value of 104 percent corresponds to 393,800 pounds at sea level and 488,800 pounds in a vacuum; 109 percent corresponds to 417,300 pounds at sea level and 513,250 pounds in a vacuum.
At sea level, the engine throttling range is reduced due to flow separation in the nozzle, prohibiting operation of the engine at its 65-percent throttle setting, referred to as minimum power level. All three main engines receive the same throttle command at the same time. Normally, these come automatically from the orbiter general-purpose computers through the engine controllers. During certain contingency situations, manual control of engine throttling is possible through the speed brake/thrust controller handle. The throttling ability reduces vehicle loads during maximum aerodynamic pressure and limits vehicle acceleration to 3 g's maximum during boost.
Each engine is designed for 7.5 hours of operation over a life span of 55 starts. Throughout the throttling range, the ratio of the liquid oxygen-liquid \Jhydrogen\j mixture is 6-to-1. Each nozzle area ratio is 77.5-to-1. The engines are 14 feet long and 7.5 feet in diameter at the nozzle exit.
The SSME controllers are digital, computer system, electronic packages mounted on the SSMEs. They operate in conjunction with engine sensors, valve actuators and spark igniters to provide a self-contained system for monitoring engine control, checkout and status. Each controller is attached to the forward end of the SSME.
Engine data and status collected by each controller are transmitted to the engine interface unit, which is mounted in the orbiter. There is one EIU for each main engine. The EIU transmits commands from the orbiter GPCs to the main engine controller. When engine data and status are received by the EIU, the data are held in a buffer until the EIU receives a request for data from the computers.
Three orbiter hydraulic systems provide hydraulic pressure to position the SSME servoactuators for thrust vector control during the ascent phase of the mission in addition to performing other functions in the main propulsion system. The three orbiter auxiliary power units provide mechanical shaft power through a gear train to drive the hydraulic pumps that provide hydraulic pressure to their respective hydraulic systems.
The ascent thrust vector control units receive commands from the orbiter GPCs and send commands to the engine gimbal actuators. The units are \Jelectronics\j packages (four in all) mounted in the orbiter's aft fuselage avionics bays. Hydraulic isolation commands are directed to engine gimbal actuators that indicate faulty servovalve position. In conjunction with this, a servovalve isolation signal is transmitted to the computers.
The SSME hydraulic servoactuators are used to gimbal the main engine. There are two actuators per engine, one for pitch motion and one for yaw motion. They convert electrical commands received from the orbiter GPCs and position servovalves, which direct hydraulic pressure to a piston that converts the pressure into a mechanical force that is used to gimbal the SSMEs. The hydraulic pressure status of each servovalve is transmitted to the ATVC units.
The orbiter MPS propellant management subsystem consists of the manifolds, distribution lines and valves by which the liquid propellants pass from the external tank to the main engines and the gaseous propellants pass from the main engines to the external tank. The SSMEs' gaseous propellants are used to pressurize the external tank. All the valves in the propellant management subsystem are under direct control of the orbiter GPCs and are either electrically or pneumatically actuated.
The orbiter MPS \Jhelium\j subsystem consists of a series of \Jhelium\j supply tanks and regulators, check valves, distribution lines and control valves. The subsystem supplies the \Jhelium\j used within the engine to purge the high-pressure oxidizer turbopump intermediate seal and preburner oxidizer domes and to actuate valves during emergency pneumatic shutdown. The balance of the \Jhelium\j is used to actuate all the pneumatically operated valves within the propellant management subsystem and to pressurize the propellant lines before re-entry.
#
"Orbiter Main Propulsion System Helium Subsystem",311,0,0,0
The MPS \Jhelium\j subsystem consists of seven 4.7-cubic-foot \Jhelium\j supply tanks; three 17.3-cubic-foot \Jhelium\j supply tanks; and associated regulators, check valves, distribution lines and control valves. Four of the 4.7-cubic-foot \Jhelium\j supply tanks are located in the aft fuselage, and the other three are located below the payload bay liner in the midfuselage in the area originally reserved for the cryogenic storage tanks of the power reactant storage and distribution system. The three 17.3-cubic-foot \Jhelium\j supply tanks are also located below the payload bay liner in the midfuselage.
The tanks are composite structures consisting of a \Jtitanium\j liner with a fiberglass structural overwrap. The large tanks are 40.3 inches in diameter and have a dry weight of 272 pounds. The smaller tanks are 26 inches in diameter and have a dry weight of 73 pounds. The tanks are serviced before lift-off to a pressure of 4,500 psi.
Each of the larger supply tanks is plumbed to two of the smaller supply tanks (one in the midbody, the other in the aft body), forming three sets of three tanks for the engine \Jhelium\j pneumatic supply system. Each set of tanks normally provides \Jhelium\j to only one engine and is commonly referred to as left, center, or right engine \Jhelium\j, depending on the engine serviced. Each set normally provides \Jhelium\j to its designated engine for in-flight purges and provides pressure for actuating engine valves during emergency pneumatic shutdown.
The remaining 4.7-cubic-foot \Jhelium\j tank is referred to as the pneumatic \Jhelium\j supply tank. It normally provides pressure to actuate all of the pneumatically operated valves in the propellant management subsystem.
There are eight \Jhelium\j supply tank isolation valves grouped in pairs. One pair of valves is connected to each engine \Jhelium\j supply tank cluster, and one pair is connected to the pneumatic supply tank. In the engine \Jhelium\j supply tank system, each pair of isolation valves is connected in parallel, with each valve in the pair controlling \Jhelium\j flow through one leg of a dual-redundant \Jhelium\j supply circuit. Each \Jhelium\j supply circuit contains two check valves, a filter, an isolation valve, a regulator and a relief valve.
The two isolation valves connected to the pneumatic supply tanks are also connected in parallel; however, the rest of the pneumatic supply system consists of a filter, the two isolation valves, a regulator, a relief valve and a single check valve. Each engine \Jhelium\j supply isolation valve can be individually controlled by its He isolation A left , ctr , right open , GPC , close and He isolation B left , ctr , right , open , GPC, close switches on panel R2. The two pneumatic \Jhelium\j supply isolation valves are controlled by a single pneumatic He isol , open, GPC, close switch on panel R2.
All of the valves in the \Jhelium\j subsystem (with the exception of the supply tank isolation valves) are spring loaded to one position and electrically actuated to the other position. The supply tank isolation valves are spring loaded to the closed position and pneumatically actuated to the open position. Valve position is controlled via electrical signals from either the onboard GPCs or manually by the flight crew. All of the valves can be controlled automatically by the GPCs, and the flight crew can control some of the valves.
The \Jhelium\j source pressure of the pneumatic, left, center and right supply systems can be monitored on the \Jhelium\j , pneu , l (left), c (center), r (right) meters on panel F7 by positioning the tank, reg (regulator) switch below the meters to tank . In addition, the regulated pressure of the pneumatic, left, center and right systems can be monitored on the same meters by placing the switch to reg .
Each of the four \Jhelium\j supply systems operates independently until after main engine cutoff. Each engine \Jhelium\j supply has two interconnect (crossover) valves associated with it, and each valve in the pair of interconnect valves is connected in series with a check valve. The check valves allow \Jhelium\j to flow through the interconnect valves in one direction only.
One check valve associated with one interconnect valve controls \Jhelium\j flow in one direction, and the other interconnect valve and its associated check valve permit \Jhelium\j flow in the opposite direction. The in interconnect valve controls \Jhelium\j flow into the associated engine \Jhelium\j distribution system from the pneumatic \Jhelium\j supply tank. The out interconnect valve controls \Jhelium\j flow out of the associated engine \Jhelium\j supply system to the pneumatic distribution system.
Each pair of interconnect valves is controlled by a single switch on panel R2. Each He interconnect , left , ctr , right switch has three positions- in open/out close , GPC , and in close/out open. With the switch in the in open/out close position, the in interconnect valve is open and the out interconnect valve is closed. The in close/out open position does the reverse. With the switch in GPC, the out interconnect valve opens automatically at the beginning of the liquid oxygen dump and closes automatically at the end of the liquid \Jhydrogen\j dump.
In a return-to-launch-site abort, the GPC position will cause the in interconnect valve to open automatically at MECO and close automatically 20 seconds later. If an engine was shut down before MECO, its in interconnect valve will remain closed at MECO. At any other time, placing the switch in GPC results in both interconnect valves being closed.
An additional interconnect valve between the left engine \Jhelium\j supply and pneumatic \Jhelium\j supply would be used if the pneumatic \Jhelium\j supply regulator failed. This crossover valve would be opened and the pneumatic \Jhelium\j supply tank isolation valves would be closed, allowing the left engine \Jhelium\j supply system to supply \Jhelium\j to the pneumatic \Jhelium\j supply. The crossover \Jhelium\j valve is controlled by its own three-position switch on panel R2. The pneumatics l (left) eng He xovr (crossover) switch positions are open, GPC and close. The GPC position allows the valve to be controlled by the flight software loaded in the GPCs.
Manifold pressurization valves located downstream of the pneumatic \Jhelium\j pressure regulator are used to control the flow of \Jhelium\j to propellant manifolds during a nominal propellant dump and manifold repressurization. There are four of these valves grouped in pairs. One pair controls \Jhelium\j pressure to the liquid oxygen propellant manifolds, and the other pair controls \Jhelium\j pressure to the liquid \Jhydrogen\j propellant manifold.
The liquid \Jhydrogen\j RTLS dump pressurization valves located downstream of the pneumatic \Jhelium\j pressure regulator are used to control the pressurization of the liquid \Jhydrogen\j propellant manifolds during an RTLS liquid \Jhydrogen\j dump. There are two of these valves connected in series.
Unlike the liquid \Jhydrogen\j manifold pressurization valves, the liquid \Jhydrogen\j RTLS dump pressurization valves cannot be controlled by flight deck switches. During an RTLS abort, these valves are opened and closed automatically by GPC commands. An additional difference between the nominal and the RTLS liquid \Jhydrogen\j dumps is in the routing of the \Jhelium\j and the place where it enters the liquid \Jhydrogen\j feed line manifold.
For the nominal liquid \Jhydrogen\j dump, \Jhelium\j passes through the liquid \Jhydrogen\j manifold pressurization valves and enters the feed line manifold in the vicinity of the liquid \Jhydrogen\j feed line disconnect valve. For the liquid \Jhydrogen\j RTLS dump, \Jhelium\j passes through the RTLS liquid \Jhydrogen\j dump pressurization valves and enters the feed line manifold in the vicinity of the liquid \Jhydrogen\j inboard fill and drain valve on the inboard side. There is no RTLS liquid oxygen dump pressurization valve since the liquid oxygen manifold is not pressurized during the RTLS liquid oxygen dump.
Each engine \Jhelium\j supply tank has two pressure regulators operating in parallel. Each regulator controls pressure in one leg of a dual-redundant \Jhelium\j supply circuit and is capable of providing all of the \Jhelium\j needed by the main engines.
The pressure regulator for the pneumatic \Jhelium\j supply system is not redundant and is set to provide outlet pressure between 715 to 770 psig. Downstream of the regulator are two more regulators: the liquid \Jhydrogen\j manifold pressure regulator and the liquid oxygen manifold pressure regulator. These regulators are used only during MPS propellant dumps and manifold pressurization. Both regulators are set to provide outlet pressure between 20 to 25 psig. Flow through the regulators is controlled by the appropriate set of two normally closed manifold pressurization valves.
Downstream of each pressure regulator, with the exception of the two manifold repressurization regulators, is a relief valve. The valve protects the downstream \Jhelium\j distribution lines from overpressurization if the associated regulator fails fully open. The two relief valves in each engine \Jhelium\j supply are set to relieve at 785 to 850 psig and reseat at 785 psig. The relief valve in the pneumatic \Jhelium\j supply circuit also relieves at 785 to 850 psig and reseats at 785 psig.
There is one pneumatic control assembly on each of the three space shuttle main engines. The PCA is essentially a manifold pressurized by one of the engine \Jhelium\j supply systems and contains \Jsolenoid\j valves to control and direct pressure to perform various essential functions. The valves are energized by discrete on/off commands from the output \Jelectronics\j of the associated SSME controller. Functions controlled by the PCA include the high-pressure oxidizer turbopump intermediate seal cavity and preburner oxidizer dome purge, pogo system postcharge and pneumatic shutdown.
#
"Shuttle Main Propulsion System Propellant Management Subsystem",312,0,0,0
Within the orbiter aft fuselage, liquid \Jhydrogen\j and liquid oxygen pass through the manifolds, distribution lines and valves of the propellant management subsystem.
During prelaunch activities, this subsystem is used to control the loading of liquid oxygen and liquid \Jhydrogen\j in the external tank. During SSME thrusting periods, propellants from the external tank flow into this subsystem and to the three SSMEs. The subsystem also provides a path that allows gases tapped from the three SSMEs to flow back to the external tank through two gas umbilicals to maintain pressure in the external tank's liquid oxygen and liquid \Jhydrogen\j tanks. After MECO, this subsystem controls MPS dumps, vacuum inerting and MPS repressurization for entry.
All the valves in the MPS are either electrically or pneumatically operated. Pneumatic valves are used where large loads are encountered, such as in the control of liquid propellant flows. Electrical valves are used for lighter loads, such as in the control of gaseous propellant flows.
The pneumatically actuated valves are divided into two types: those that require pneumatic pressure to open and close the valve (type 1) and those that are spring loaded to one position and require pneumatic pressure to move to the other position (type 2).
Each type 1 valve actuator is equipped with two electrically actuated \Jsolenoid\j valves. Each \Jsolenoid\j valve controls \Jhelium\j pressure to an ''open'' or ''close'' port on the actuator. Energizing the \Jsolenoid\j valve on the open port allows \Jhelium\j pressure to open the pneumatic valve.
Energizing the \Jsolenoid\j on the close port allows \Jhelium\j pressure to close the pneumatic valve. Removing power from a \Jsolenoid\j valve removes \Jhelium\j pressure from the corresponding port of the pneumatic actuator and allows the \Jhelium\j pressure trapped in that side of the actuator to vent overboard.
Removing power from both solenoids allows the pneumatic valve to remain in the last commanded position. This type of valve is used for the liquid oxygen and liquid \Jhydrogen\j feed line 17-inch umbilical disconnect valves (two), the liquid oxygen and liquid \Jhydrogen\j prevalves (six), the three liquid \Jhydrogen\j and liquid oxygen inboard and outboard fill and drain valves (four), and the liquid \Jhydrogen\j 4-inch recirculation disconnect valves.
Each type 2 valve is a single electrically actuated \Jsolenoid\j valve that controls \Jhelium\j pressure to either an open or a close port on the actuator. Removing power from the \Jsolenoid\j valve removes \Jhelium\j pressure from the corresponding port of the pneumatic actuator and allows \Jhelium\j pressure trapped in that side of the actuator to vent overboard. Spring force takes over and drives the valve to the opposite position.
If the spring force drives the valve to the open position, the valve is referred to as a normally open valve. If the spring force drives the valve to a closed position, the valve is referred to as a normally closed valve.
This type of valve is used for the liquid \Jhydrogen\j RTLS inboard dump valve (NC), the liquid \Jhydrogen\j RTLS outboard dump valve (NC), the liquid \Jhydrogen\j feed line relief shutoff valve (NO), the liquid oxygen feed line relief shutoff valve (NO), the three liquid \Jhydrogen\j engine recirculation valves (NC), the two liquid oxygen pogo recirculation valves (NO), the liquid \Jhydrogen\j topping valve (NC), the liquid \Jhydrogen\j high-point bleed valve (NC), and the liquid oxygen overboard bleed valve (NO).
The electrically actuated \Jsolenoid\j valves are spring loaded to one position and move to the other position when electrical power is applied. These valves also are referred to as either normally open or normally closed, based on their position in the de-energized state. Electrically actuated \Jsolenoid\j valves are the gaseous \Jhydrogen\j pressurization line vent valve (NC), the three gaseous \Jhydrogen\j pressurization flow control valves (NO) and the three gaseous oxygen pressurization flow control valves (NO).
There are two 17-inch-diameter MPS propellant feed line manifolds in the orbiter aft fuselage, one for liquid oxygen and one for liquid \Jhydrogen\j. Each manifold has an outboard and inboard fill and drain valve in series that interface with the respective port (left) and starboard (right) T-0 umbilical. The port T-0 umbilical is for liquid \Jhydrogen\j; the starboard, for liquid oxygen. In addition, each manifold connects the orbiter to the external tank in the lower aft fuselage through a port 17-inch liquid \Jhydrogen\j disconnect valve umbilical and a starboard 17-inch liquid oxygen disconnect valve umbilical.
There are three outlets in both the liquid oxygen and liquid \Jhydrogen\j 17-inch manifolds between the orbiter-external tank 17-inch umbilical disconnect valves and the inboard fill and drain valve. The outlets in the manifolds provide liquid oxygen and liquid \Jhydrogen\j to each SSME in 12-inch-diameter feed lines.
Each 17-inch liquid \Jhydrogen\j and liquid oxygen manifold has a 1-inch-diameter line that is routed to a feed line relief isolation valve and feed line relief valve in the respective liquid \Jhydrogen\j and liquid oxygen system. The LO 2 and LH 2 feed line rlf (relief) isol (isolation) switches on panel R4 have open , GPC and close positions. When a feed line relief isolation valve is opened, the corresponding manifold can relieve excessive pressure overboard through its relief valve.
The liquid \Jhydrogen\j feed line manifold has another outlet directed to the two liquid \Jhydrogen\j RTLS dump valves in series. Both valves are controlled by the MPS prplt dump LH 2 valve switch on panel R2, which has backup LH 2 vlv open , GPC , close positions. When opened, these valves enable the liquid \Jhydrogen\j dump during RTLS aborts or provide a backup to the normal liquid \Jhydrogen\j dump after a nominal main engine cutoff. In an RTLS abort dump, liquid \Jhydrogen\j is dumped overboard through a port at the outer aft fuselage's left side between the orbital maneuvering system/reaction control system pod and the upper surface of the wing.
The MPS propellant management subsystem also contains two 2-inch-diameter manifolds, one for gaseous oxygen and one for gaseous \Jhydrogen\j. Each manifold individually permits ground support equipment servicing with \Jhelium\j through the respective T-0 umbilical and provides initial pressurization of the external tank's liquid oxygen and liquid \Jhydrogen\j orbiter/external tank disconnect umbilicals. Self-sealing quick disconnects are provided at the T-0 umbilical and the orbiter/external tank umbilical.
Six 0.63-inch-diameter pressurization lines, three for gaseous oxygen and three for gaseous \Jhydrogen\j, are used after SSME start to pressurize the external tank's liquid oxygen and liquid \Jhydrogen\j tanks.
In each SSME, a small portion of liquid oxygen is diverted into the engine's oxidizer heat exchanger, and the heat generated by the engine's high-pressure oxidizer turbopump converts the liquid oxygen into gaseous oxygen and directs it through a check valve to two orifices and a flow control valve for each engine.
During SSME thrusting periods, liquid oxygen tank pressure is maintained between 20 and 22 psig by the orifices in the two lines and the action of the flow control valve from each SSME. The flow control valve is controlled by one of three liquid oxygen pressure transducers. When tank pressure decreases below 20 psig, the valve opens. If the tank pressure is greater than 24 psig, it is relieved through the liquid oxygen tank's vent and relief valve.
In each SSME, gaseous \Jhydrogen\j from the low-pressure fuel turbopump is directed through two check valves to two orifices and a flow control valve for each engine. During the main engine thrusting period, the liquid \Jhydrogen\j tank's pressure is maintained between 32 and 34 psia by the orifices and the action of the flow control valve from each SSME.
The flow control valve is controlled by one of three liquid \Jhydrogen\j pressure transducers. When tank pressure decreases below 32 psia, the valve opens; and when tank pressure increases to 33 psia, the valve closes. If the tank pressure is greater than 35 psia, the pressure is relieved through the liquid \Jhydrogen\j tank's vent and relief valve. If the pressure falls below 32 psia, the LH 2 ullage press switch on panel R2 is positioned from auto to open , which will cause all three flow control valves to go to full open and remain in the full-open position.
The single gaseous \Jhydrogen\j manifold repressurization line connects to the \Jhydrogen\j line vent valve, which is controlled by the H 2 press line vent switch on panel R4. This valve is normally closed, and the switch is positioned to open when vacuum inerting the gaseous \Jhydrogen\j pressurization lines after MECO and the liquid \Jhydrogen\j dump. The gnd position allows the launch processing system to control the valve during ground operations.
#
"Shuttle MPS External Tank",313,0,0,0
The external tank is attached to the orbiter at one forward and two aft attach points. At the two aft attach points are the two external tank/orbiter umbilicals for the fluid, gas, signal and electrical power connections between the orbiter and the external tank. Each external tank umbilical plate mates with a corresponding umbilical plate on the orbiter. The umbilical plates help maintain alignment of the various connecting components. The corresponding umbilical plates are bolted together; and when external tank separation is commanded, the bolts are severed by pyrotechnics.
At the forward end of each external tank propellant tank is a vent and relief valve that can be opened by GSE-supplied \Jhelium\j before launch for venting or by excessive tank pressure for relief. The vent function is available only before launch; after lift-off only the relief function is operable.
The liquid oxygen tank relieves at an ullage pressure of 25 psig, while the liquid \Jhydrogen\j tank relieves at an ullage pressure of 38 psi. The flight crew has no control over the position of the vent and relief valves before launch or during ascent. Normal range of the tank ullage pressure of the liquid \Jhydrogen\j tank during ascent is 32 to 39 psia.
During prelaunch activities, the liquid \Jhydrogen\j tank is pressurized to 44.1 psi to meet the start requirement of the main engine LPFT. The liquid oxygen and liquid \Jhydrogen\j tanks' ullage pressures are monitored on the panel F7 eng manf LO2 and LH2 meters as well as on a \Jcathode\j ray tube display.
In addition to the vent and relief valve, the liquid oxygen tank has a tumble vent valve that is opened during the external tank separation sequence. The thrust force provided by opening the valve imparts an angular velocity to the external tank to assist in the separation maneuver and provide more positive control of the external tank's re-entry aerodynamics.
There are eight propellant depletion sensors. Four of them sense fuel depletion and four sense oxidizer depletion. The oxidizer depletion sensors are mounted in the external tank's liquid oxygen feed line manifold downstream of the tank. The fuel depletion sensors are located in the liquid \Jhydrogen\j tank. During prelaunch activities, the launch processing system tests each propellant depletion sensor.
If any are found to be in a failed condition, the LPS sets a flag in the computer's SSME operational sequence, sequence logic that will instruct the computer to ignore the output of the failed sensor or sensors. During main engine thrusting, the computer constantly computes the instantaneous mass of the vehicle, which constantly decreases due to propellant usage from the external tank.
When the computed vehicle mass matches a predetermined initialized-loaded value, the computer arms the propellant depletion sensors. After this time, if any two of the good fuel depletion sensors (those not flagged before launch) or any two of the good oxidizer depletion sensors indicate a dry condition, the computers command main engine cutoff. This type of MECO is a backup to the nominal MECO, which is based on vehicle velocity.
The oxidizer sensors sense propellant depletion before the fuel sensors to ensure that all depletion cutoffs are fuel-rich since an oxidizer-rich cutoff can cause burning and severe erosion of engine components. To ensure that the oxidizer sensors sense depletion first, a plus 700-pound bias is included in the amount of liquid \Jhydrogen\j loaded in the external tank.
This amount is in excess of that dictated by the 6-1 ratio of oxidizer to fuel. The position of the oxidizer propellant depletion sensors allows the maximum amount of oxidizer to be consumed in the engines and allows sufficient time to cut off the engines before the oxidizer turbopumps cavitate (run dry).
Four ullage pressure transducers are located at the top end of each propellant tank (liquid oxygen and liquid hydrogen). One of the four is considered a spare and is normally off-line. Before launch, GSE normally checks out the four transducers; and if one of the three active transducers is determined to be bad, it can be taken off-line and the output of the spare \Jtransducer\j selected.
The flight crew can also perform this operation after lift-off via the computer keyboard; however, because of the time involved from lift-off to MECO, this would probably be impractical. The three active ullage pressure sensors provide outputs for CRT display and control of ullage pressure within their particular propellant tanks. For CRT display, computer processing selects the middle value output of the three transducers and displays this single value. For ullage pressure control, all three outputs are used.
The external tank/orbiter aft umbilicals have five propellant disconnects: two for the liquid oxygen tank and three for the liquid \Jhydrogen\j tank. One of the liquid oxygen propellant umbilicals carries liquid oxygen and the other carries gaseous oxygen.
The liquid \Jhydrogen\j tank has two disconnects that carry liquid \Jhydrogen\j and one that carries gaseous \Jhydrogen\j. The external tank liquid \Jhydrogen\j recirculation disconnect is the smaller of the two disconnects that carry liquid \Jhydrogen\j and is used only during the liquid \Jhydrogen\j chill-down sequence before launch.
In addition, the external tank/orbiter umbilicals contain two electrical umbilicals, each made of many smaller electrical cables. These cables carry electrical power from the orbiter to the external tank and the two solid rocket boosters and bring telemetry back to the orbiter from the SRBs and external tank.
The operational instrumentation telemetry that comes back from the SRBs is conditioned, digitized and multiplexed in the SRBs themselves. The external tank OI measurements that return to the orbiter are raw \Jtransducer\j outputs and must be processed within the orbiter telemetry system.
The external tank's liquid oxygen tank is serviced at the launch pad before prelaunch from ground support equipment through the starboard T-0 umbilical of the orbiter, the MPS outboard and inboard fill and drain valves, the MPS 17-inch liquid oxygen line, and the orbiter/external tank 17-inch umbilical disconnect valves.
Once the liquid oxygen is loaded and ready for main engine ignition, the liquid oxygen tank's vent and relief valve is closed, and the tank is pressurized to 21 psig by GSE-supplied \Jhelium\j. During SSME thrusting, liquid oxygen flows out of the external tank through the orbiter/external tank umbilical into the orbiter MPS and to each SSME. Pressurization in the tank is maintained by gaseous oxygen tapped from the three main engines and supplied to the liquid oxygen tank through the orbiter/external tank gaseous oxygen umbilical.
The external tank's liquid \Jhydrogen\j tank is serviced before launch from GSE at the launch pad similarly to the liquid oxygen tank but through the port T-0 umbilical and port orbiter/external tank umbilical. When the liquid \Jhydrogen\j is loaded and ready for main engine ignition, the liquid \Jhydrogen\j tank's vent and relief valve is closed, and the tank is pressurized to 42.5 psia by GSE-supplied \Jhelium\j.
Approximately 45 minutes after loading starts, three electrically powered liquid \Jhydrogen\j pumps in the orbiter begin to circulate the liquid \Jhydrogen\j in the external tank through the three SSMEs and back to the external tank through a special recirculation umbilical. This recirculation chills down the liquid \Jhydrogen\j lines between the external tank and the high-pressure fuel turbopump in the SSMEs so that the path is free of any gaseous \Jhydrogen\j bubbles and is at the proper temperature for engine start.
Recirculation ends approximately six seconds before engine start. During engine thrusting, liquid \Jhydrogen\j flows from the external tank and through the orbiter/external tank liquid \Jhydrogen\j umbilical into the orbiter MPS and to the main engines. Tank pressurization is maintained by gaseous \Jhydrogen\j tapped from the three SSMEs and supplied to the liquid \Jhydrogen\j tank through the orbiter/external tank gaseous \Jhydrogen\j umbilical.
#
"Space Shuttle Main Engines",314,0,0,0
Oxidizer from the external tank enters the orbiter at the orbiter/external tank umbilical disconnect and then the orbiter's main propulsion system liquid oxygen feed line. There it branches out into three parallel paths, one to each engine. In each branch, a liquid oxygen prevalve must be opened to permit flow to the low-pressure oxidizer turbopump.
The LPOT is an axial-flow pump driven by a six-stage \Jturbine\j powered by liquid oxygen. It boosts the liquid oxygen's pressure from 100 psia to 422 psia. The flow from the LPOT is supplied to the high-pressure oxidizer turbopump. During engine operation, the pressure boost permits the HPOT to operate at high speeds without cavitating. The LPOT operates at approximately 5,150 rpm. The LPOT, which is approximately 18 by 18 inches, is connected to the vehicle propellant ducting and supported in a fixed position by the orbiter structure.
The HPOT consists of two single-stage centrifugal pumps (a main pump and a preburner pump) mounted on a common shaft and driven by a two-stage, hot-gas \Jturbine\j. The main pump boosts the liquid oxygen's pressure from 422 psia to 4,300 psia while operating at approximately 28,120 rpm.
The HPOT discharge flow splits into several paths, one of which is routed to drive the LPOT \Jturbine\j. Another path is routed to and through the main oxidizer valve and enters into the main combustion chamber. Another small flow path is tapped off and sent to the oxidizer heat exchanger.
The liquid oxygen flows through an anti-flood valve that prevents it from entering the heat exchanger until sufficient heat is present to convert the liquid oxygen to gas. The heat exchanger utilizes the heat contained in the discharge gases from the HPOT \Jturbine\j to convert the liquid oxygen to gas.
The gas is sent to a manifold and is then routed to the external tank to pressurize the liquid oxygen tank. Another path enters the HPOT second-stage preburner pump to boost the liquid oxygen's pressure from 4,300 psia to 7,420 psia. It passes through the oxidizer preburner oxidizer valve into the oxidizer preburner and through the fuel preburner oxidizer valve into the fuel preburner. The HPOT is approximately 24 by 36 inches. It is attached by flanges to the hot-gas manifold.
Fuel enters the orbiter at the liquid \Jhydrogen\j feed line disconnect valve, then flows into the orbiter gaseous \Jhydrogen\j feed line manifold and branches out into three parallel paths to each engine. In each liquid \Jhydrogen\j branch, a prevalve permits liquid \Jhydrogen\j to flow to the low-pressure fuel turbopump when the prevalve is open.
The LPFT is an axial-flow pump driven by a two-stage \Jturbine\j powered by gaseous \Jhydrogen\j. It boosts the pressure of the liquid \Jhydrogen\j from 30 psia to 276 psia and supplies it to the high-pressure fuel turbopump. During engine operation, the pressure boost provided by the LPFT permits the HPFT to operate at high speeds without cavitating. The LPFT operates at approximately 16,185 rpm. The LPFT is approximately 18 by 24 inches. It is connected to the vehicle propellant ducting and is supported in a fixed position by the orbiter structure 180 degrees from the LPOT.
The HPFT is a three-stage centrifugal pump driven by a two-stage, hot-gas \Jturbine\j. It boosts the pressure of the liquid \Jhydrogen\j from 276 psia to 6,515 psia. The HPFT operates at approximately 35,360 rpm. The discharge flow from the turbopump is routed to and through the main valve and then splits into three flow paths. One path is through the jacket of the main combustion chamber, where the \Jhydrogen\j is used to cool the chamber walls. It is then routed from the main combustion chamber to the LPFT, where it is used to drive the LPFT \Jturbine\j. A small portion of the flow from the LPFT is then directed to a common manifold from all three engines to form a single path to the external tank to maintain liquid \Jhydrogen\j tank pressurization.
The remaining \Jhydrogen\j passes between the inner and outer walls to cool the hot-gas manifold and is discharged into the main combustion chamber. The second \Jhydrogen\j flow path from the main fuel valve is through the engine nozzle (to cool the nozzle). It then joins the third flow path from the chamber coolant valve. The combined flow is then directed to the fuel and oxidizer preburners. The HPFT is approximately 22 by 44 inches. It is attached by flanges to the hot-gas manifold.
The oxidizer and fuel preburners are welded to the hot-gas manifold. The fuel and oxidizer enter the preburners and are mixed so that efficient combustion can occur. The augmented spark igniter is a small combination chamber located in the center of the injector of each preburner.
The two dual-redundant spark igniters, which are activated by the engine controller, are used during the engine start sequence to initiate combustion in each preburner. They are turned off after approximately three seconds because the combustion process is then self-sustaining. The preburners produce the fuel-rich hot gas that passes through the turbines to generate the power to operate the high-pressure turbopumps. The oxidizer preburner's outflow drives a \Jturbine\j that is connected to the HPOT and the oxidizer preburner pump. The fuel preburner's outflow drives a \Jturbine\j that is connected to the HPFT.
The HPOT \Jturbine\j and HPOT pumps are mounted on a common shaft. Mixing of the fuel-rich hot gas in the \Jturbine\j section and the liquid oxygen in the main pump could create a hazard. To prevent this, the two sections are separated by a cavity that is continuously purged by the MPS engine \Jhelium\j supply during engine operation. Two seals minimize leakage into the cavity. One seal is located between the \Jturbine\j section and the cavity, and the other is between the pump section and cavity. Loss of \Jhelium\j pressure in this cavity results in an automatic engine shutdown.
The speed of the HPOT and HPFT turbines depends on the position of the corresponding oxidizer and fuel preburner oxidizer valves. These valves are positioned by the engine controller, which uses them to throttle the flow of liquid oxygen to the preburners and, thus, control engine thrust.
The oxidizer and fuel preburner oxidizer valves increase or decrease the liquid oxygen flow, thus increasing or decreasing preburner chamber pressure, HPOT and HPFT \Jturbine\j speed, and liquid oxygen and gaseous \Jhydrogen\j flow into the main combustion chamber, which increases or decreases engine thrust, thus throttling the engine. The oxidizer and fuel preburner valves operate together to throttle the engine and maintain a constant 6-1 propellant mixture ratio.
The main oxidizer valve and the main fuel valve control the flow of liquid oxygen and liquid \Jhydrogen\j into the engine and are controlled by each engine controller. When an engine is operating, the main valves are fully open.
A coolant control valve is mounted on the combustion chamber coolant bypass duct of each engine. The engine controller regulates the amount of gaseous \Jhydrogen\j allowed to bypass the nozzle coolant loop, thus controlling its temperature. The chamber coolant valve is 100 percent open before engine start. During engine operation, it will be 100 percent open for throttle settings of 100 to 109 percent for minimum cooling. For throttle settings between 65 to 100 percent, its position will range from 66.4 to 100 percent open for maximum cooling.
Each engine main combustion chamber receives fuel-rich hot gas from a hot-gas manifold cooling circuit. The gaseous \Jhydrogen\j and liquid oxygen enter the chamber at the injector, which mixes the propellants. A small augmented spark igniter chamber is located in the center of the injector.
The dual-redundant igniter is used during the engine start sequence to initiate combustion. The igniters are turned off after approximately three seconds because the combustion process is self-sustaining. The main injector and dome assembly is welded to the hot-gas manifold. The main combustion chamber also is bolted to the hot-gas manifold.
The inner surface of each combustion chamber, as well as the inner surface of each nozzle, is cooled by gaseous \Jhydrogen\j flowing through coolant passages. The nozzle assembly is a bell-shaped extension bolted to the main combustion chamber.
The nozzle is 113 inches long, and the outside diameter of the exit is 94 inches. A support ring welded to the forward end of the nozzle is the engine attach point to the orbiter-supplied heat shield. Thermal protection for the nozzles is necessary because of the exposure that portions of the nozzles experience during the launch, ascent, on-orbit and entry phases of a mission. The \Jinsulation\j consists of four layers of metallic batting covered with a metallic foil and screening.
The five propellant valves on each engine (oxidizer preburner oxidizer, fuel preburner oxidizer, main oxidizer, main fuel, and chamber coolant) are hydraulically actuated and controlled by electrical signals from the engine controller. They can be fully closed by using the MPS engine \Jhelium\j supply system as a backup actuation system.
The low-pressure oxygen and low-pressure fuel turbopumps are mounted 180 degrees apart on the orbiter's aft fuselage thrust structure. The lines from the low-pressure turbopumps to the high-pressure turbopumps contain flexible bellows that enable the low-pressure turbopumps to remain stationary while the rest of the engine is gimbaled for thrust vector control. The liquid \Jhydrogen\j line from the LPFT to the HPFT is insulated to prevent the formation of liquid air.
The main oxidizer valve and fuel bleed valve are used after shutdown. The main oxidizer valve is opened during a propellant dump to allow residual liquid oxygen to be dumped overboard through the engine, and the fuel bleed valve is opened to allow residual liquid \Jhydrogen\j to be dumped through the liquid \Jhydrogen\j fill and drain valves overboard. After the dump is completed, the valves close and remain closed for the remainder of the mission.
The gimbal bearing is bolted to the main injector and dome assembly and is the thrust interface between the engine and orbiter. The bearing assembly is approximately 11.3 by 14 inches.
Overall, a space shuttle main engine weighs approximately 7,000 pounds.
#
"Shuttle Pogo Suppression System",315,0,0,0
A pogo suppression system prevents the transmission of low-frequency flow oscillations into the high-pressure turbopump and, ultimately, prevents main combustion chamber pressure (engine thrust) \Joscillation\j. Flow oscillations transmitted from the space shuttle vehicle are suppressed by a partially filled gas accumulator, which is attached by flanges to the high-pressure oxidizer turbopump's inlet duct.
The system consists of a 0.6-cubic-foot accumulator with an internal standpipe, \Jhelium\j precharge valve package, gaseous oxygen supply valve package and two recirculation isolation valves (one located on the orbiter).
During engine start, the accumulator is charged with \Jhelium\j 2.4 seconds after the start command to provide pogo protection until the engine heat exchanger is operational and gaseous oxygen is available.
The accumulator is partially chilled by liquid oxygen during the engine chill-down operation. It fills to the overflow standpipe line inlet level, which is sufficient to preclude gas ingestion at engine start.
During engine operation, the accumulator is charged with a continuous gaseous oxygen flow maintained at a rate governed by the engine operation point.
The liquid level in the accumulator is controlled by the overflow standpipe line in the accumulator, which is orificed to regulate the gaseous oxygen overflow over the engine's operating power level. The system is sized to provide sufficient replenishment of gaseous oxygen at the minimum flow rate and to permit sufficient gaseous oxygen overflow at the maximum decreasing pressure transient in the low-pressure oxidizer turbopump discharge duct. At all other conditions, excess gaseous and liquid oxygen are recirculated to the low-pressure oxidizer turbopump inlet through the engine oxidizer bleed duct. The pogo accumulator is charged (pressurized) at engine shutdown to provide a positive pressure at the HPOT inlet, which prevents HPOT overspeed in the zero-gravity environment.
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"Space Shuttle Main Engine Controllers",316,0,0,0
The controller is an \Jelectronics\j package mounted on each SSME. It contains two digital computers and the associated \Jelectronics\j to control all main engine components and operations. The controller is attached to the main combustion chamber by shock-mounted fittings.
Each controller operates in conjunction with engine sensors, valves, actuators and spark igniters to provide a self-contained system for engine control, checkout and monitoring. The controller provides engine flight readiness verification; engine start and shutdown sequencing; closed-loop thrust and propellant mixture ratio control; sensor excitation; valve actuator and spark igniter control signals; engine performance limit monitoring; onboard engine checkout, response to vehicle commands and transmission of engine status; and performance and maintenance data.
Each engine controller receives engine commands transmitted by the orbiter's general-purpose computers through its own engine interface unit. The engine controller provides its own commands to the main engine components. Engine data are sent to the engine controller, where they are stored in a vehicle data table in the controller's computer memory.
Data on the controller's status compiled by the engine controller's computer are also added to the vehicle data table. The vehicle data table is periodically output by the controller to the EIU for transmission to the orbiter's GPCs.
The engine interface unit is a specialized multiplexer/demultiplexer that interfaces with the GPCs and with the engine controller. When engine commands are received by the EIU, the data are held in a buffer until the EIU receives a request for data from the GPCs. The EIU then sends data to each GPC. Each EIU is dedicated to one space shuttle main engine and communicates only with the engine controller that controls its SSME. The EIUs have no interface with each other.
The controller provides responsive control of engine thrust and propellant mixture ratio throughout the digital computer in the controller, updating the instructions to the engine control elements 50 times per second (every 20 milliseconds). Engine reliability is enhanced by a dual-redundant system that allows normal operation after the first failure and a fail-safe shutdown after a second failure. High-reliability electronic parts are used throughout the controller.
The digital computer is programmable, allowing engine control equations and constants to be modified by changing the stored program (software). The controller is packaged in a sealed, pressurized chassis and is cooled by convection heat transfer through pin fins as part of the main chassis. The \Jelectronics\j are distributed on functional modules with special thermal and vibration protection.
The controller is divided into five subsystems: input \Jelectronics\j, output \Jelectronics\j, computer interface \Jelectronics\j, digital computer and power supply \Jelectronics\j. Each subsystem is duplicated to provide dual-redundant capability.
The input \Jelectronics\j receive data from all engine sensors, condition the signals and convert them to digital values for processing by the digital computer. Engine control sensors are dual-redundant, and maintenance data sensors are non-redundant.
The output \Jelectronics\j convert computer digital control commands into voltages suitable for powering the engine spark igniters, the off/on valves and the engine propellant valve actuators.
The computer interface \Jelectronics\j control the flow of data within the controller, data input to the computer and computer output commands to the output \Jelectronics\j. They also provide the controller interface with the vehicle engine \Jelectronics\j interface unit for receiving engine commands that are triple-redundant channels from the vehicle and for transmitting engine status and data through dual-redundant channels to the vehicle. The computer interface \Jelectronics\j include the watchdog timers that determine which channel of the dual-redundant \Jmechanization\j is in control.
The digital computer is an internally stored, general-purpose computer that provides the computational capability necessary for all engine control functions. The memory has a program storage capacity of 16,384 data and instruction words (17-bit words; 16 bits for program use, one bit for parity).
The power supply \Jelectronics\j convert the 115-volt, three-phase, 400-hertz vehicle ac power to the individual power supply voltage levels required by the engine control system and monitor the level of power supply channel operation to ensure it is within satisfactory limits.
Each orbiter GPC, operating in a redundant set, issues engine commands to the engine interface units for transmission to their corresponding engine controllers. Each orbiter GPC has SSME subsystem operating program applications software residing in it. Engine commands are output over the engine's assigned flight-critical data bus (a total of four GPCs outputting over four FC data buses). Therefore, each EIU will receive four commands. The nominal ascent configuration has GPCs 1, 2, 3 and 4 outputting on FC data buses 5, 6, 7 and 8, respectively. Each FC data bus is connected to one multiplexer interface adapter in each EIU.
The EIU checks the received engine commands for transmission errors. If there are none, the EIU passes the validated engine commands on to the controller interface assemblies, which output the validated engine commands to the engine controller. An engine command that does not pass validation is not sent to the controller interface assembly. Instead, it is dead-ended in the EIU's multiplexer interface adapter. Commands that come through MIAs 1 and 2 are sent to CIAs 1 and 2, respectively.
Commands that come to MIAs 3 and 4 pass through a \JCIA\j 3 data-select logic. This logic outputs the command that arrives at the logic first, from either MIA 3 or 4. The other command is dead-ended in the \JCIA\j 3 select logic. The selected command is output through \JCIA\j 3. In this manner, the EIU reduces the four commands sent to the EIU to three commands output by the EIU.
The engine controller vehicle interface \Jelectronics\j receive the three engine commands output by its EIU, check for transmission errors (hardware validation), and send controller hardware-validated engine commands to the controller A and B \Jelectronics\j. Normally, channel A \Jelectronics\j are in control, with channel B \Jelectronics\j active, but not in control. If channel A fails, channel B will assume control.
If channel B subsequently fails, the engine controller will shut down the engine pneumatically. If two or three commands pass voting, the engine controller will issue its own commands to accomplish the function commanded by the orbiter GPCs. If command voting fails and two or all three commands fail, the engine controller will maintain the last command that passed voting.
The engine controller provides all the main engine data to the GPCs. Sensors in the engine supply pressures, temperatures, flow rates, turbopump speeds, valve position and engine servovalve actuator positions to the engine controller. The engine controller assembles these data into a vehicle data table and adds status data of its own to the vehicle data table.
The vehicle data tables output channels A and B to the vehicle interface \Jelectronics\j for transmission to the EIUs. The vehicle interface \Jelectronics\j output over both data paths. The data paths are called primary and secondary. The channel A vehicle data table is normally sent over both primary and secondary control (channel A has failed); then the vehicle interface \Jelectronics\j output the channel B vehicle data table over both the primary and secondary data paths.
The vehicle data table is sent by the controller to the EIU. There are only two data paths versus three command paths between the engine controller and the EIU. The data path that interfaces with \JCIA\j 1 is called primary data. The path that interfaces with \JCIA\j 2 is called secondary data. Primary and secondary data are held in buffers until the GPCs send a data request command to the EIUs. The GPCs request both primary and secondary data. Primary data is output only through MIA 1 on each EIU. Secondary data is output only through MIA 4 on each EIU.
During prelaunch, the orbiter's computers look at both primary and secondary data. Loss of either primary or secondary data will result in data path failure and either an engine ignition inhibit or a launch pad shutdown of all three main engines.
At T minus zero, the orbiter GPCs request both primary and secondary data from each EIU. For no failures, only primary data are looked at. If there is a loss of primary data (which can occur between the engine controller channel A \Jelectronics\j and the SSME SOP), the secondary data are looked at.
There are two primary written engine controller computer software programs: the flight operational program and the test operational program. The flight operational program is an on-line, real-time, process-control program that processes inputs from engine sensors; controls the operation of the engine servovalves, actuators, solenoids and spark igniters; accepts and processes vehicle commands; provides and transmits data to the vehicle; and provides checkout and monitoring capabilities.
The test operational program supports engine testing. Functionally, it is similar to the flight operational program but differs with respect to implementation. The computer software programs are modular and are defined as computer program components, which consist of a data base organized into tables and 15 computer program components.
During application of the computer program components, the programs perform data processing for failure detection and status to the vehicle. As system operation progresses through an operating phase, different combinations of control functions are operative at different times. These combinations within a phase are defined as operating modes.
The checkout phase initiates active control monitoring or checkout. The standby mode in this phase is a waiting mode of controller operation while active control sequence operations are in process. Monitoring functions that do not affect engine hardware status are continually active during the mode. Such functions include processing of vehicle commands, status update and controller self-test.
During checkout, data and instructions can be loaded into the engine controller's computer memory. This permits updating of the software program and data as necessary to proceed with engine-firing operations or checkout operations. Also in this phase, component checkout, consisting of checkout or engine leak tests, is performed on an individual engine system component.
The start preparation phase consists of system purges and propellant conditioning, which are performed in preparation for engine start. The purge sequence 1 mode is the first purge sequence, including oxidizer system and intermediate seal purge operation. The purge sequence 2 mode is the second purge sequence, including fuel system purge operation and the continuation of purges initiated during purge sequence 1.
The purge sequence 3 mode includes propellant recirculation (bleed valve operation). The purge sequence 4 mode includes fuel system purge and the indication engine is ready to enter the start phase. The engine-ready mode occurs when proper engine thermal conditions for start have been attained and other criteria for start have been satisfied, including a continuation of the purge sequence 4 mode.
The start phase covers operations involved with starting or firing the engines, beginning with scheduled open-loop operation of propellant valves. The start initiation mode includes all functions before ignition confirmed and the closing of the thrust control loop. The thrust buildup mode detects ignition by monitoring main combustion chamber pressure and verifying that closed-loop thrust buildup sequencing is in progress.
The main stage phase is automatically entered upon successful completion of the start phase. The normal control mode has initiated mixture ratio control, and thrust control is operating normally. In case of a malfunction, the electrical lock mode will be activated. In that mode, engine propellant valves are electrically held in a fixed configuration, and all control loop communications are suspended. There is also the hydraulic lockup mode, in which all fail-safe valves are deactivated to hydraulically hold the propellant valves in a fixed configuration and all control loop functions are suspended.
The shutdown phase covers operations to reduce main combustion chamber pressure and drive all valves closed to effect full engine shutdown. Throttling to minimum power level is the portion of the shutdown in progress at a programmed shutdown thrust reference level above the MPL.
The valve schedule throttling mode is the stage in the shutdown sequence at which the programmed thrust reference has decreased below the MPL. Propellant valves closed is the stage in the shutdown sequence after all liquid propellant valves have been closed, the shutdown purge has been activated, and verification sequences are in progress. The fail-safe pneumatic mode is when the fail-safe pneumatic shutdown is used.
The post-shutdown phase represents the state of the SSME and engine controller at the completion of engine firing. The standby mode is a waiting mode of controller operations whose functions are identical to those of standby during checkout. It is the normal mode that is entered after completion of the shutdown phase. The terminate sequence mode terminates a purge sequence by a command from the vehicle. All propellant valves are closed, and all \Jsolenoid\j and \Jtorque\j motor valves are de-energized.
Each controller utilizes ac power provided by the MPS engine power left, ctr, right switches on panel R2.
Each controller has internal electrical heaters that provide environmental temperature control and are powered by main bus power through a remote power controller. The RPC is controlled by the main propulsion system engine cntrl htr left, ctr, right switches on panel R4. The heaters are not normally used until after main engine cutoff and are only turned on if environmental control is required during the mission.
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"Shuttle Malfunction Detection",317,0,0,0
There are three separate means of detecting malfunctions within the main propulsion system: the engine controllers, the caution and warning system and the GPCs.
The engine controller, through its network of sensors, has access to numerous engine operating parameters. A group of these parameters has been designated critical operating parameters, and special limits defined for these parameters are hard-wired and limit sensed within the caution and warning system. If a violation of any limit is detected, the caution and warning system will illuminate the red MPS caution and warning light on panel F7.
The light will be illuminated by an MPS engine liquid oxygen manifold pressure above 249 psia; an MPS engine liquid \Jhydrogen\j manifold pressure below 28 psia or above 60 psia; an MPS center, left or right \Jhelium\j pressure below 1,150 psia; an MPS center, left or right \Jhelium\j regulated pressure above 820 psia; or an MPS left, center or right \Jhelium\j delta pressure/delta time above 29 psia.
Note that the flight crew can monitor the MPS press \Jhelium\j pneu, l, c, r meter on panel F7 when the switch is placed in the tank or reg position. The MPS press eng manf LO 2 , LH2 meter can also be monitored on panel F7. A number of the conditions will require crew action.
For example, an MPS engine liquid \Jhydrogen\j manifold pressure below the minimum setting will require the flight crew to pressurize the external liquid \Jhydrogen\j tank by setting the LH2 ullage press switch on panel R2 to open , and a low \Jhelium\j pressure may require the flight crew to interconnect the pneumatic \Jhelium\j tank and the engine \Jhelium\j tanks using the MPS He interconnect valve switches on panel R2 for the engine \Jhelium\j system that is affected.
The engine controller also has a self-test feature that allows it to detect certain malfunctions involving its own sensors and control devices. For each of the three engines, a yellow main engine status left, ctr, right light (lower half) on panel F7 will be illuminated when the corresponding engine \Jhelium\j pressure is below 1,150 psia or regulated \Jhelium\j pressure is above 820 psia.
The lower half of the main engine status left, ctr, right light on panel F7 may also be illuminated by the SSME SOP (GPC- detected malfunctions). The yellow light may be illuminated due to an electronic hold, hydraulic lockup, loss of two or more command channels or command reject between the GPC and the SSME controller, or loss of both data channels from the SSME controller to the GPC of the corresponding engine.
In an electronic hold for the affected SSME, loss of data from both pairs of the four fuel flow rate sensors and the four chamber pressure sensors will result in the propellant valve actuators being maintained electronically in the positions existing at the time the second sensor failed. (To fail both sensors in a pair, it is only necessary to fail one sensor.) In the case of either the hydraulic lockup or an electronic hold, all engine-throttling capability for the affected engine is lost; thus, subsequent throttling commands to that engine will not change the thrust level.
The red upper half of the main engine status left, ctr, right light on panel F7 will be illuminated if the corresponding engine's high-pressure oxidizer \Jturbine\j's discharge temperature is above 1,760 degrees R, the main combustion chamber's pressure is below 1,000 psia, the high-pressure oxidizer turbopump's intermediate seal purge pressure is below 170 or above 650 psia, or the high-pressure oxidizer turbopump's secondary seal purge pressure is below 5 or above 85 psia.
Because of the rapidity with which it is possible to exceed these limits, the engine controller has been programmed to sense the limits and automatically cut off the engine if the limits are exceeded. Although a shutdown as a result of violating operating limits is normally automatic, the flight crew can, if necessary, inhibit an automatic shutdown through the use of the main engine limit shut dn switch on panel C3. The switch has three positions: enable, auto and inhibit.
The enable position allows only the first engine that violates operating limits to be shut down automatically. If either of the two remaining engines subsequently violates operating limits, it would be inhibited from automatically shutting down. The inhibit position inhibits all automatic shutdowns. The main engine shutdown left, ctr, right push buttons on panel C3 have spring-loaded covers (guards). When the guard is raised and the push button is depressed, the corresponding engine shuts down immediately.
The backup caution and warning processing of the orbiter GPCs can detect certain specified out-of-limit or fault conditions of the MPS. The backup C/W alarm light on panel F7 is illuminated, a fault message appears on all CRT displays, and an audio alarm sounds if the MPS engine liquid oxygen manifold pressure is zero or above 29 psia; the MPS engine liquid \Jhydrogen\j manifold pressure is below 30 or above 46 psia; the MPS left, center or right \Jhelium\j pressure is below 1,150 psia; or the MPS regulated left, center or right \Jhelium\j pressure is below 680 or above 820 psia. This is identical to the parameter limit sensed by the caution and warning system; thus, the MPS red light on panel F7 will also be illuminated.
The SM alert indicator on panel F7 is illuminated, a fault message appears on all CRT displays, and an audio alarm is sounded when MPS malfunctions/conditions are detected by the SSME SOP or special systems-monitoring processing.
The first four conditions are detected by the SSME SOP and are identical to those that illuminate the yellow lower light of the respective main engine status light on panel F7 due to electronic hold, hydraulic lockup, loss of two or more command channels or command reject between the GPC and the SSME controller, or loss of both data channels from the SSME controller to the orbiter GPC.
The last four conditions are special systems-monitoring processing and illuminate the SM alert light on panel F7, sound an audio alarm and provide a fault message on all CRTs because of an external tank liquid \Jhydrogen\j ullage pressure below 30 psia or above 46 psia or an external tank liquid oxygen ullage pressure of zero or above 29 psia. (Note that the main engine status lights on panel F7 will not be illuminated.)
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"Orbiter Hydraulic Systems",318,0,0,0
The three orbiter hydraulic systems supply hydraulic pressure to the main propulsion system for providing thrust vector control and actuating engine valves on each SSME.
The three hydraulic supply systems are distributed to the MPS TVC valves. These valves are controlled by \Jhydraulics\j MPS/TVC 1, 2, 3 switches on panel R4. A valve is opened by positioning its respective switch to open. The talkback indicator above each switch indicates op or cl for open and close.
When the three MPS TVC hydraulic isolation valves are opened, hydraulic pressure actuates the engine main fuel valve, the main oxidizer valve, the fuel preburner oxidizer valve, the oxidizer preburner oxidizer valve and the chamber coolant valve. All hydraulically actuated engine valves on an engine receive hydraulic pressure from the same hydraulic system.
The left engine valves are actuated by hydraulic system 2, the center engine valves are actuated by hydraulic system 1, and the right engine valves are actuated by hydraulic system 3. Each engine valve actuator is controlled by dual-redundant signals: channel A/engine servovalve 1 and channel B/engine servovalve 2 from that engine controller \Jelectronics\j.
As a backup, all of the hydraulically actuated engine valves on an engine are supplied with \Jhelium\j pressure from the \Jhelium\j subsystem left, center and right engine \Jhelium\j tank supply system. In the event of a hydraulic lockup in an engine, \Jhelium\j pressure is used to actuate the engine's propellant valves to their fully closed position when the engine is shut down.
Hydraulic lockup is a condition in which all of the propellant valves on an engine are hydraulically locked in a fixed position. This is a built-in protective response of the MPS propellant valve actuator/control circuit. It takes effect any time low hydraulic pressure or loss of control of one or more propellant valve actuators renders closed-loop control of engine thrust or propellant mixture ratio impossible.
Hydraulic lockup allows an engine to continue to thrust in a safe manner under conditions that normally would require that the engine be shut down; however, the affected engine will continue to operate at approximately the throttle level in effect at the time hydraulic lockup occurred. Once an engine is in a hydraulic lockup, any subsequent shutoff commands, whether nominal or premature, will cause a pneumatic \Jhelium\j shutdown.
Hydraulic lockup does not affect the capability of the engine controller to monitor critical operating parameters or issue an automatic shutdown if an operating limit is out of tolerance; however, the engine shutdown would be accomplished pneumatically.
The three MPS thrust vector control valves must also be opened to supply hydraulic pressure to the six main engine TVC actuators. There are two servoactuators per SSME: one for yaw and one for pitch. Each actuator is fastened to the orbiter thrust structure and to the powerhead of one of the three SSMEs. The two actuators per engine provide attitude control and trajectory shaping by gimbaling the SSMEs in conjunction with the solid rocket boosters during first-stage ascent and without the SRBs during second-stage ascent.
Each SSME servoactuator receives hydraulic pressure from two of the three orbiter hydraulic systems; one system is the primary system and the other is a standby system. Each servoactuator has its own hydraulic switching valve. The switching valve receives hydraulic pressure from two of the three orbiter hydraulic systems and provides a single source to the actuator. Normally, the primary hydraulic supply is directed to the actuator; however, if the primary system were to fail and lose hydraulic pressure, the switching valve would automatically switch over to the standby system, and the actuator would continue to function on the standby system.
The left engine's pitch actuator utilizes hydraulic system 2 as the primary and hydraulic system 1 as the standby. The engine's yaw actuator utilizes hydraulic system 1 as the primary and hydraulic system 2 as the standby. The center engine's pitch actuator utilizes hydraulic system 1 as the primary and hydraulic system 3 as the standby, and the yaw actuator utilizes hydraulic system 3 as the primary and hydraulic system 1 as the standby. The right engine's pitch actuator utilizes hydraulic system 3 as the primary and hydraulic system 2 as the standby. Its yaw actuator utilizes hydraulic system 2 as the primary and hydraulic system 3 as the standby.
The hydraulic systems are distributed among the actuators and engine valves to equalize the hydraulic work load among the three systems.
The hydraulic MPS/TVC isol vlv sys1, sys2, sys3 switches on panel R4 are positioned to close during on-orbit operations to protect against hydraulic leaks downstream of the isolation valves. In addition, there is no requirement to gimbal the main engines from the stow position. During on-orbit operations when the MPS TVC valves are closed, the hydraulic pressure supply and return lines within each MPS TVC component are interconnected to enable hydraulic fluid to circulate for thermal conditioning.
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"Shuttle MPS Thrust Vector Control",319,0,0,0
The space shuttle ascent thrust vector control portion of the flight control system directs the thrust of the three main engines and two solid rocket boosters to control attitude and trajectory during lift-off and first-stage ascent and the main engines alone during second-stage ascent.
Ascent thrust vector control is provided by avionics hardware packages that supply gimbal commands and fault detection for each hydraulic gimbal actuator. The MPS ATVC packages are located in the three aft avionics bays in the orbiter aft fuselage and are cooled by cold plates and the Freon-21 system. The associated flight aft multiplexers/demultiplexers are also located in the aft avionics bays.
The MPS TVC command flow starts in the general-purpose computers, in which the flight control system generates the TVC position commands, and terminates at the SSME servoactuators, where the actuators gimbal the SSMEs in response to the commands. All the MPS TVC position commands generated by the flight control system are issued to the MPS TVC command subsystem operating program, which processes and disburses them to their corresponding flight aft MDMs.
The flight aft MDMs separate these linear discrete commands and disburse them to ATVC channels, which generate equivalent command analog voltages for each command issued. These voltages are, in turn, sent to the servoactuators, commanding the SSME hydraulic actuators to extend or retract, thus gimbaling the main engines to which they are fastened.
Six MPS TVC actuators respond to the command voltages issued by four ATVC channels. Each ATVC channel has six MPS drivers and four SRB drivers. Each actuator receives four identical command voltages from four different MPS drivers, each located in different ATVC channels.
Each main engine servoactuator consists of four independent, two-stage servovalves, which receive signals from the drivers. Each servovalve controls one power spool in each actuator, which positions an actuator ram and the engine to control thrust direction.
The four servovalves in each actuator provide a force-summed majority voting arrangement to position the power spool. With four identical commands to the four servovalves, the actuator's force-sum action prevents a single erroneous command from affecting power ram motion.
If the erroneous command persists for more than a predetermined time, differential pressure sensing activates an isolation driver, which energizes an isolation valve that isolates the defective servovalve and removes hydraulic pressure, permitting the remaining channels and servovalves to control the actuator ram spool provided the FCS channel 1, 2, 3, 4 switch on panel C3 is in the auto position. A second failure would isolate the defective servovalve and remove hydraulic pressure in the same manner as the first failure, leaving only two channels remaining.
Failure monitors are provided for each channel on the CRT and backup caution and warning light to indicate which channel has been bypassed for the MPS and/or SRB. If the FCS channel 1, 2, 3, or 4 switch on panel C3 is positioned to off, that ATVC channel is isolated from its servovalve on all MPS and SRB actuators. The override position of the FCS channel 1, 2, 3, 4 switch inhibits the isolation valve driver from energizing the isolation valve for its respective channel and provides the capability of resetting a failed or bypassed channel.
The ATVC 1, 2, 3, 4 power switch is located on panel O17. The on position enables the ATVC channel selected; off disables the channel.
Each actuator ram is equipped with transducers for position feedback to the TVC system.
The SSME servoactuators change each main engine's thrust vector direction as needed during the flight sequence. The three pitch actuators gimbal the engine up or down a maximum of 10 degrees 30 minutes from the installed null position. The three yaw actuators gimbal the engine left or right a maximum of 8 degrees 30 minutes from the installed position.
The installed null position for the left and right main engines is 10 degrees up from the X axis in a negative Z direction and 3 degrees 30 minutes outboard from an engine centerline parallel to the X axis. The center engine's installed null position is 16 degrees above the X axis for pitch and on the X axis for yaw. When any engine is installed in the null position, the other engines cannot collide with it.
The minimum gimbal rate is 10 degrees per second; the maximum rate is 20 degrees per second.
There are three actuator sizes for the main engines. The piston area of the one upper pitch actuator is 24.8 square inches, its stroke is 10.8 inches, it has a peak flow of 50 gallons per minute, and it weighs 265 pounds. The piston area of the two lower pitch actuators is 20 square inches, their stroke is 10.8 inches, their peak flow is 45 gallons per minute, and they weigh 245 pounds. All three yaw actuators have a piston area of 20 square inches, a stroke of 8.8 inches and a peak flow of 45 gallons per minute and weigh 240 pounds.
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"Orbiter/External Tank Separation System",320,0,0,0
The orbiter/external tank separation system consists of the oxygen and \Jhydrogen\j umbilical disconnects located at the lower left and right aft fuselage, one forward structural attach point just aft of the nose landing gear doors and two structural attach points located in the orbiter/external tank umbilical disconnect cavities. An umbilical retraction system retracts the orbiter umbilicals within the orbiter aft fuselage, and umbilical doors close over each of the umbilical cavities after separation.
The 17-inch liquid oxygen and liquid \Jhydrogen\j disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three space shuttle main engines. The respective 17-inch disconnects also provide the capability for external tank fill and drain of oxygen and \Jhydrogen\j through the orbiter main propulsion system and the T-0 umbilicals.
The liquid \Jhydrogen\j interface between the orbiter and the ground storage tank is provided by a T-0 umbilical located on the left side of the aft fuselage. The liquid oxygen interface between the orbiter and the ground storage tank is provided by a T-0 umbilical on the right side of the aft fuselage.
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"Orbiter 17-Inch Disconnect",321,0,0,0
Each mated pair of 17-inch disconnects contains two flapper valves, one on the orbiter side of the interface and one on the external tank side of the interface. Both valves in each disconnect pair are opened to permit propellant flow between the orbiter and the external tank. Before the separation of the external tank, both valves in each mated pair of disconnects are commanded closed by pneumatic (helium) pressure from the main propulsion system. The closure of both valves in each disconnect pair prevents propellant discharge from the external tank or orbiter at separation. Valve closure on the orbiter side of each disconnect also prevents contamination of the orbiter main propulsion system during landing and ground operations.
Inadvertent closure of either valve in a 17-inch disconnect during space shuttle main engine thrusting would stop propellant flow from the external tank to all three main engines. Catastrophic failure of the main engines and external tank feed lines would result.
To prevent inadvertent closure of the 17-inch disconnect valves during the main engine thrusting, a latch mechanism was added in the orbiter half of the disconnects. The latch mechanism provides a mechanical backup to the normal fluid-induced-open forces. The latch is mounted on a shaft in the flowstream so it overlaps both flappers and obstructs closure for any reason.
In preparation for external tank separation, both valves in each 17-inch disconnect are commanded closed. Pneumatic (helium) pressure from the main propulsion system causes the latch actuator to rotate the latch shaft in each orbiter 17-inch disconnect 90 degrees, thus freeing the flapper valves to close as required for external tank separation.
If the latch pneumatic actuator malfunctions, a backup mechanical separation capability is provided. When the orbiter umbilical initially moves away from the external tank umbilical, the mechanical latch disengages from the external tank flapper valve and permits the orbiter disconnect flapper to toggle the latch. This action permits both flappers to close.
During ground mating of the external tank to the orbiter, the latch engagement mechanism in each 17-inch disconnect provides a go/no-go verification that flapper angle rigging is within stability limits. Misrigged flappers will prevent full engagement of latch. The angle of each flapper in each disconnect is still carefully rigged within specific tolerances to assure basic stability independently of the latch safety feature.
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"Orbiter External Tank Separation System",322,0,0,0
The external tank is separated from the orbiter at three structural attach points. Separation from the orbiter occurs before orbit insertion and is automatically controlled by the orbiter's general-purpose computers. External tank separation can be manually initiated by the flight crew using the same jettison circuits as the automatic sequence.
Separation is controlled by the ET separation "auto", "man" switch on panel C3 and the "sep" push button on panel C3. In the auto position, the onboard GPCs initiate separation. To manually initiate separation, the ET separation switch is positioned to "man" and the "sep" push button is depressed.
The forward structural attachment consists of a shear bolt unit mounted in a spherical bearing. The bolt separates at a break area when two pressure cartridges are initiated. The pressure from one or both cartridges drives one of a pair of pistons to shear the bolt, with the second piston acting as a hole plugger to fill the cavity left by the sheared bolt. A centering mechanism rotates the unit from the displacement position to a centered position, aligning the bearing flush with the adjacent thermal protection system mold line.
The aft structural attachment consists of two special bolts and pyrotechnically actuated frangible nuts that attach the external tank strut hemisphere to the orbiter's left- and right-side cavities. At separation the frangible nuts are split by a booster cartridge initiated by a \Jdetonator\j cartridge. The attach bolts are driven by the separation forces and a spring into a cavity in the tank strut. The frangible nut, cartridge fragments and hot gases are contained within a cover assembly, and a hole plugger isolates the fragments in the container.
The aft separation involves right and left umbilical assemblies. Each assembly contains three dual-detonator frangible nut and bolt combinations that hold the orbiter and external tank umbilical plates together during mated flight. Each bolt has a retraction spring that, after release of the nut, retracts the bolt to the external tank side of the interface. On the orbiter side, each frangible nut and its detonators are enclosed in a debris container that captures nut fragments and hot gases generated by the operation of the detonators, either of which will fracture the nut.
The right aft umbilical assembly consists of an electrical disconnect, the gaseous oxygen 2-inch pressurization disconnect used for pressurization of the external tank's oxygen tank and the 17-inch liquid oxygen disconnect.
The left aft umbilical assembly consists of an electrical disconnect plate, the gaseous \Jhydrogen\j 2-inch pressurization disconnect used for pressurization of the external tank's \Jhydrogen\j tank, the 4-inch recirculation disconnect used during prelaunch to precondition the main engine and the 17-inch liquid \Jhydrogen\j disconnect.
After release of the three frangible nuts and bolts at each aft umbilical, three lateral support arms at each orbiter umbilical plate hold the plates in the lateral position when the external tank separates from the umbilical plates. Each 17-inch disconnect has been commanded closed. The orbiter umbilical plates are retracted inside the orbiter aft fuselage approximately 2.5 inches by three hydraulic actuators and locked to permit closure of the umbilical doors in the bottom of the aft fuselage.
Hydraulic system 1 source pressure is supplied to one actuator at each umbilical, hydraulic system 2 source pressure is supplied to the second actuator at each umbilical, and hydraulic system 3 source pressure is supplied to a third actuator at each umbilical.
The retraction of each umbilical disconnects the external tank and orbiter electrical umbilical in the first 0.5 of an inch of travel and releases any fluids trapped between the 17-inch disconnect flappers.
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"Orbiter Umbilical Doors",323,0,0,0
An electromechanical actuation system on each umbilical door closes the left and right umbilical cavities after the external tank is jettisoned and the umbilical plates retracted inside the orbiter's aft fuselage. Each umbilical door is approximately 50 inches square.
The doors are held in the full-open position by two centerline latches, one forward and one aft. They are opened before the mating of the orbiter to the external tank in the Vehicle Assembly Building.
The orbiter umbilical doors normally are controlled by the flight crew with switches on panel R2. In return-to-launch-site aborts, the doors are controlled automatically. The ET umbilical door mode switch on panel R2 positioned to GPC enables automatic control of the doors. The GPC/man position enables manual flight crew control of the doors.
The ET umbilical door centerline latch switch on panel R2 positioned to gnd permits ground control of the door centerline latches during ground turnaround operations. The stow position, enables flight crew manual control of the door centerline latches. The talkback indicator above the switch indicates "sto" when the door centerline latches are stowed, which permits closure of the doors, and barberpole when the latches are latched or the doors are in transit.
The ET umbilical door left and right latch, off, release switches on panel R2 are used by the flight crew to unlatch the corresponding centerline latches during normal operations. Positioning the respective switch to release provides electrical power to redundant ac reversible motors which operate an electromechanical actuator for each centerline latch that causes the latch to rotate and retract the latch blade flush with the reusable thermal protection system mold line.
It takes approximately six seconds for the latches to complete their motion. The talkback indicator above the respective switch indicates rel when the corresponding latches are released. The latch position of each switch is used during ground turnaround operations to latch the respective door open, and the talkback indicator indicates lat when the latches have latched the doors in the open position. The talkback indicators indicate barberpole when the latches are in transit. The off position of the switches removes power from the motors, which stops the latches.
The ET umbilical door left and right , open , off , latch switches on panel R2 normally are used by the flight crew to close the umbilical doors. Positioning the switches to close provides electrical power to redundant ac reversible motors, which position the doors closed through a system of bellcranks and push rods. It takes approximately 24 seconds for the doors to close; and when they are within 2 inches of the closed position, ready-to-latch indicators activate the door uplatch system.
Three uplatch hooks for each door engage three corresponding rollers near the outboard edge of the door and lock the door in preparation for entry. The motors are automatically turned off. The talkback indicator above the respective switch indicates cl when two of the three ready-to-latch switches for that door have sensed door closure. The open position of the switches is used during ground turnaround operations to open the doors. The talkback indicator indicates op when the doors are open and barberpole when they are in transit. The off position removes power from the motors, which stops the doors' movement.
The ET umbilical door switch on panel R2 positioned to GPC provides a backup method of releasing the centerline latches and closing the umbilical doors through guidance, navigation and control software through \Jcathode\j ray tube display item entry during an RTLS abort. The operation of the centerline latches and closing of the umbilical doors are completely automated after external tank separation when the ET umbilical door switch on panel R2 is positioned to GPC .
Two seconds after external tank separation, the centerline latches release the doors and the latches are stowed. The ET umbilical door centerline latch talkback indicator indicates "sto" when the centerline latches complete their motion eight seconds after external tank separation. The left and right umbilical doors are closed, and the ET umbilical door left and right talkback indicates cl 32 seconds after separation. The left and right umbilical door latches latch the doors closed, and the ET umbilical door left and right talkback indicates lat 38 seconds after separation.
Each umbilical door is covered with reusable thermal protection system in addition to an aerothermal barrier that required approximately 6 psi to compress to seal the door with adjacent thermal protection system tiles.
A closeout curtain is installed at each of the orbiter/external tank umbilicals. After external tank separation, the residual liquid oxygen in the main propulsion system is dumped through the three space shuttle main engines and the residual liquid \Jhydrogen\j is dumped overboard. The umbilical curtain prevents hazardous gases (gaseous oxygen and hydrogen) from entering the orbiter aft fuselage through the umbilical openings before the umbilical doors are closed.
The curtain also acts as a seal during the ascent phase of the mission to permit the aft fuselage to vent through the orbiter purge and vent system, thereby protecting the orbiter aft bulkhead at station Xo 1307. The curtain is designed to operate in range of minus 200 F to plus 250 F. The umbilical doors are opened when the orbiter has stopped at the end of landing rollout.
Various parameters are monitored and displayed on the flight deck control panel and CRT and transmitted by telemetry.
Contractors for the separation system include Hoover Electric, Los Angeles, Calif. (external tank umbilical centerline latch and actuator; umbilical door actuator and umbilical door latch actuator); U.S. Bearing, Chatsworth, Calif. (external tank/orbiter spherical bearing); Bertea Corp., Irvine, Calif. (umbilical retractor actuator); Space Ordnance Systems Division, Trans Technology Corp., Saugus, Calif. (orbiter/external tank separation bolt/cartridge \Jdetonator\j assembly, 0.75-inch frangible nut orbiter/external tank umbilical separation and 2.5-inch frangible nut/pyro components in orbiter/external tank aft attach separation system).
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"Orbital Maneuvering System (OMS)",324,0,0,0
The orbital maneuvering system provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around and can provide up to 1,000 pounds of propellant to the aft reaction control system. The OMS is housed in two independent pods located on each side of the orbiter's aft fuselage.
The pods also house the aft RCS and are referred to as the OMS/RCS pods. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers. The two pods provide redundancy for the OMS. The vehicle velocity required for orbital adjustments is approximately 2 feet per second for each nautical mile of altitude change.
The ascent profile of a mission determines if one or two OMS thrusting periods are used and the interactions of the RCS. After main engine cutoff, the RCS thrusters in the forward and aft RCS pods are used to provide attitude hold until external tank separation. At ET separation, the RCS provides a minus (negative) Z translation maneuver of about minus 4 feet per second to maneuver the orbiter away from the ET. Upon completion of the translation, the RCS provides orbiter attitude hold until time to maneuver to the OMS-1 thrusting attitude.
The targeting data for the OMS-1 thrusting period is selected before launch; however, the target data in the onboard general-purpose computers can be modified by the flight crew via the \Jcathode\j ray tube keyboard, if necessary, before the OMS thrusting period.
During the first OMS thrusting period, both OMS engines are used to raise the orbiter to a predetermined elliptical orbit. During the thrusting period, vehicle attitude is maintained by gimbaling (swiveling) the OMS engines. The RCS will not normally come into operation during an OMS thrusting period. If, during an OMS thrusting period, the OMS gimbal rate or gimbal limits are exceeded, RCS attitude control is required. If only one OMS engine is used during an OMS thrusting period, RCS roll control is required.
During the OMS-1 thrusting period, the liquid oxygen and liquid \Jhydrogen\j trapped in the main propulsion system ducts are dumped. The liquid oxygen is dumped out through the space shuttle main engines' combustion chambers and the liquid \Jhydrogen\j is dumped through the starboard (right) side T-0 umbilical overboard fill and drain. This velocity was precomputed in conjunction with the OMS-1 thrusting period.
Upon completion of the OMS-1 thrusting period, the RCS is used to null any residual velocities, if required. The flight crew uses the rotational hand controller and/or translational hand controller to command the applicable RCS thrusters to null the residual velocities. The RCS then provides attitude hold until time to maneuver to the OMS-2 thrusting attitude.
The second OMS thrusting period using both OMS engines occurs near the apogee of the orbit established by the OMS-1 thrusting period and is used to circularize the predetermined orbit for that mission. The targeting data for the OMS-2 thrusting period is selected before launch; however, the target data in the onboard GPCs can be modified by the flight crew via the CRT keyboard, if necessary, before the OMS thrusting period.
Upon completion of the OMS-2 thrusting period, the RCS is used to null any residual velocities, if required, in the same manner as during OMS-1. The RCS is then used to provide attitude hold and minor translation maneuvers as required for on-orbit operations. The flight crew can select primary or vernier RCS thrusters for attitude control on orbit. Normally, the vernier RCS thrusters are selected for on-orbit attitude hold.
If the ascent profile for a mission uses a single OMS thrusting maneuver, it is referred to as direct insertion. In a direct-insertion ascent profile, the OMS-1 thrusting period after main engine cutoff is eliminated and is replaced with a 5-feet- per-second RCS translation maneuver to facilitate the main propulsion system dump. The RCS provides attitude hold after the translation maneuver. The OMS-2 thrusting period is then used to achieve orbit insertion. The direct-insertion ascent profile allows the MPS to provide more energy to orbit insertion and permits easier use of onboard software.
Additional OMS thrusting periods using both or one OMS engine are performed on orbit according to the mission's requirements to modify the orbit for rendezvous, payload deployment or transfer to another orbit.
The two OMS engines are used to deorbit. Target data for the deorbit maneuver is computed by the ground and loaded in the onboard GPCs via uplink. This data is also voiced to the flight crew for verification of loaded values. After verification of the deorbit data, the flight crew initiates an OMS gimbal test on the CRT keyboard unit.
Before the deorbit thrusting period, the flight crew maneuvers the \Jspacecraft\j to the desired deorbit thrusting attitude using the rotational hand controller and RCS thrusters. Upon completion of the OMS thrusting period, the RCS is used to null any residual velocities, if required. The \Jspacecraft\j is then maneuvered to the proper entry interface attitude using the RCS. The remaining propellants aboard the forward RCS are dumped by burning the propellants through the forward RCS thrusters before the entry interface if it is necessary to control the orbiter's center of gravity.
The aft RCS plus X jets can be used to complete any planned OMS thrusting period in the event of an OMS engine failure. In this case, the OMS-to-aft-RCS interconnect would feed OMS propellants to the aft RCS.
From entry interface at 400,000 feet, the orbiter is controlled in roll, pitch and yaw with the aft RCS thrusters. The orbiter's ailerons become effective at a dynamic pressure of 10 pounds per square foot, and the aft RCS roll jets are deactivated. At a dynamic pressure of 20 pounds per square foot, the orbiter's elevons become effective, and the aft RCS pitch jets are deactivated. The rudder is activated at Mach 3.5, and the aft RCS yaw jets are deactivated at Mach 1 and approximately 45,000 feet.
The OMS in each pod consists of a high-pressure gaseous \Jhelium\j storage tank, \Jhelium\j isolation valves, dual pressure regulation systems, vapor isolation valves for only the oxidizer regulated \Jhelium\j pressure path, quad check valves, a fuel tank, an oxidizer tank, a propellant distribution system consisting of tank isolation valves, crossfeed valves, and an OMS engine. Each OMS engine also has a gaseous \Jnitrogen\j storage tank, gaseous \Jnitrogen\j pressure isolation valve, gaseous \Jnitrogen\j accumulator, bipropellant \Jsolenoid\j control valves and actuators that control bipropellant ball valves, and purge valves.
In each of the OMS pods, gaseous \Jhelium\j pressure is supplied to \Jhelium\j isolation valves and dual pressure regulators, which supply regulated \Jhelium\j pressure to the fuel and oxidizer tanks. The fuel is monomethyl hydrazine and the oxidizer is \Jnitrogen\j tetroxide. The propellants are Earth-storable liquids at normal temperatures.
They are pressure-fed to the propellant distribution system through tank isolation valves to the OMS engines. The OMS engine propellant ball valves are positioned by the gaseous \Jnitrogen\j system and control the flow of propellants into the engine. The fuel is directed first through the engine combustion chamber walls and provides regenerative cooling of the chamber walls; it then flows into the engine injector. The oxidizer goes directly to the engine injector. The propellants are sprayed into the combustion chamber, where they atomize and ignite upon contact with each other (hypergolic), producing a hot gas and, thus, thrust.
The gaseous \Jnitrogen\j system is also used after the OMS engines are shut down to purge residual fuel from the injector and combustion chamber, permitting safe restarting of the engines. The nozzle extension of each OMS engine is radiation-cooled and is constructed of columbium alloy.
Each OMS engine produces 6,000 pounds of thrust. The oxidizer-to-fuel ratio is 1.65-to-1. The expansion ratio of the nozzle exit to the throat is 55-to-1. The chamber pressure of the engine is 125 psia. The dry weight of each engine is 260 pounds.
Each OMS engine can be reused for 100 missions and is capable of 1,000 starts and 15 hours of cumulative firing. The minimum duration of an OMS engine firing is two seconds. The OMS may be utilized to provide thrust above 70,000 feet. For vehicle velocity changes of between 3 and 6 feet per second, normally only one OMS engine is used.
Each engine has two electromechanical gimbal actuators, which control the OMS engine thrust direction in pitch and yaw (thrust vector control). The OMS engines can be used singularly by directing the thrust vector through the orbiter center of gravity or together by directing the thrust vector of each engine parallel to the other.
During a two-OMS-engine thrusting period, the RCS will come into operation only if the OMS gimbal rate or gimbal limits are exceeded and should not normally come into operation during the OMS thrust period. However, during a one-OMS-engine thrusting period, roll RCS control is required. The pitch and yaw actuators are identical except for the stroke length and contain redundant electrical channels (active and standby), which couple to a common mechanical drive assembly.
The OMS/RCS pods are designed to be reused for up to 100 missions with only minor repair, refurbishment and maintenance. The pods are removable to facilitate orbiter turnaround, if required.
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"OMS Helium Pressurization",325,0,0,0
Each pod pressurization system consists of a \Jhelium\j tank, two \Jhelium\j isolation valves, two dual pressure regulator assemblies, parallel vapor isolation valves on the regulated \Jhelium\j pressure to the oxidizer tank only, dual series-parallel check valve assemblies and pressure relief valves.
The \Jhelium\j storage tank in each pod has a \Jtitanium\j liner with a fiberglass structural overwrap. This increases safety and decreases the weight of the tank 32 percent over that of conventional tanks. The \Jhelium\j tank is 40.2 inches in diameter and has a volume of 17.03 cubic feet minimum. Its dry weight is 272 pounds. The \Jhelium\j tank's operating pressure range is 4,800 to 460 psia with a maximum operating limit of 4,875 psia at 200 F.
A pressure sensor downstream of each \Jhelium\j tank in each pod monitors the \Jhelium\j source pressure and transmits it to the N 2 , He , kit He switch on panel F7. When the switch is in the He position, the \Jhelium\j pressure of the left and right OMS is displayed on the OMS press left, right meters. This pressure also is transmitted to the CRT and displayed.
The two \Jhelium\j pressure isolation valves in each pod permit \Jhelium\j source pressure to the propellant tanks or isolate the \Jhelium\j from the propellant tanks. The parallel paths in each pod assure \Jhelium\j flow to the propellant tanks of that pod. The \Jhelium\j valves are continuous-duty, solenoid-operated. They are energized open and spring loaded closed.
The OMS He press/vapor isol switches on panel O8 permit automatic or manual control of the valves. With the switches in the GPC position, the valves are automatically controlled by the general-purpose computer during an engine thrusting sequence. The valves are controlled manually by placing the switches to open or close.
The pressure regulators reduce the \Jhelium\j source pressure to the desired working pressure. Pressure is regulated by assemblies downstream of each \Jhelium\j pressure isolation valve. Each assembly contains primary and secondary regulators in series and a flow limiter. Normally, the primary regulator is the controlling regulator. The secondary regulator is normally open during a dynamic flow condition.
It will not become the controlling regulator until the primary regulator allows a higher pressure than normal. All regulator assemblies are in reference to a bellows assembly that is vented to ambient. The primary regulator outlet pressure at normal flow is 252 to 262 psig and 247 psig minimum at high abort flow, with lockup at 266 psig maximum.
The secondary regulator outlet pressure at normal flow is 259 to 269 psig and 254 psig minimum at high abort flow, with lockup at 273 psig maximum. The flow limiter restricts the flow to a maximum of 1,040 standard cubic feet per minute and to a minimum of 304 standard cubic feet per minute.
The vapor isolation valves in the oxidizer pressurization line to the oxidizer tank prevent oxidizer vapor from migrating upstream and over into the fuel system. These are low-pressure, two-position, two-way, solenoid-operated valves that are energized open and spring loaded closed.
They can be commanded manually or automatically by the positioning of the He press/vapor isol switches on panel O8. When either of the A or B switches is in the open position, both vapor isolation valves are energized open; and when both switches are in the close position, both vapor isolation valves are closed. When the switches are in the GPC position, the GPC opens and closes the valves automatically.
The check valve assembly in each parallel path contains four independent check valves connected in a series-parallel configuration to provide a positive checking action against a reverse flow of propellant liquid or vapor, and the parallel path permits redundant paths of \Jhelium\j to be directed to the propellant tanks. Filters are incorporated into the inlet of each check valve assembly.
Two pressure sensors in the \Jhelium\j pressurization line upstream of the fuel and oxidizer tanks monitor the regulated tank pressure and transmit it to the RCS/OMS press rotary switch on panel O3. When the switch is in the OMS prplnt position, the left and right fuel and oxidizer pressure is displayed. If the tank pressure is lower than 234 psia or above 284 psia, the left or right OMS red caution and warning light on panel F7 will be illuminated. These pressures also are transmitted to the CRT and displayed.
The relief valves in each pressurization path limit excessive pressure in the propellant tanks. Each pressure relief valve also contains a burst diaphragm and filter. If excessive pressure is caused by \Jhelium\j or propellant vapor, the diaphragm will rupture and the relief valve will open and vent the excessive pressure overboard. The filter prevents particulates from the non-fragmentation-type diaphragm from entering the relief valve seat.
The relief valve will close and reset after the pressure has returned to the operating level. The burst diaphragm is used to provide a more positive seal of \Jhelium\j and propellant vapors than the relief valve. The diaphragm ruptures between 303 and 313 psig. The relief valve opens at a minimum of 286 psig and a maximum of 313 psig. The relief valve's minimum reseat pressure is 280 psig. The maximum flow capacity of the relief valve at 60 F and 313 psig is 520 cubic feet per minute.
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"OMS Propellant Storage and Distribution",326,0,0,0
The propellant storage and distribution system consists of one fuel tank and one oxidizer tank in each pod. It also contains propellant feed lines, interconnect lines, isolation valves and crossfeed valves.
The OMS propellant tanks of both pods enable the orbiter to reach a 1,000-foot- per-second velocity change with a 65,000-pound payload in the payload bay. An OMS pod crossfeed line allows the propellants in the pods to be used to operate either OMS engine.
The propellant is contained in domed cylindrical \Jtitanium\j tanks within each pod. Each propellant tank is 96.38 inches long with a diameter of 49.1 inches and a volume of 89.89 cubic feet unpressurized. The dry weight of each tank is 250 pounds. The propellant tanks are pressurized by the \Jhelium\j system.
Each tank contains a propellant acquisition and retention assembly in the aft end and is divided into forward and aft compartments. The propellant acquisition and retention assembly is located in the aft compartment and consists of an intermediate bulkhead with communication screen and an acquisition system.
The propellant in the tank is directed from the forward compartment through the intermediate bulkhead through the communication screen into the aft compartment during OMS velocity maneuvers. The communication screen retains propellant in the aft compartment during zero-gravity conditions.
The acquisition assembly consists of four stub galleries and a collector manifold. The stub galleries acquire wall-bound propellant at OMS start and during RCS velocity maneuvers to prevent gas ingestion. The stub galleries have screens that allow propellant flow and prevent gas ingestion.
The collector manifold is connected to the stub galleries and also contains a gas arrestor screen to further prevent gas ingestion, which permits OMS engine ignition without the need of a propellant-settling maneuver employing RCS thrusters. The propellant tank's nominal operating pressure is 250 psi, with a maximum operating pressure limit of 313 psia.
A capacitance gauging system in each OMS propellant tank measures the propellant in the tank. The system consists of a forward and aft probe and a totalizer. The forward and aft fuel probes use fuel (which is a conductor) as one plate of the capacitor and a glass tube that is metallized on the inside as the other.
The forward and aft oxidizer probes use two concentric nickel tubes as the capacitor plates and oxidizer as the dielectric. (Helium is also a dielectric, but has a different dielectric constant than the oxidizer.) The aft probes in each tank contain a resistive temperature-sensing element to correct variations in fluid density. The fluid in the area of the communication screens cannot be measured.
The totalizer receives OMS valve operation information and inputs from the forward and aft probes in each tank and outputs total and aft quantities and a low level quantity. The inputs from the OMS valves allow control logic in the totalizer to determine when an OMS engine is thrusting and which tanks are being used. The totalizer begins an engine flow rate/time \Jintegration\j process at the start of the OMS thrusting period, which reduces the indicated amount of propellants by a preset estimated rate for the first 14.8 seconds.
After 14.8 seconds of OMS thrusting, which settles the propellant surface, the probe capacitance gauging system outputs are enabled, which permits the quantity of propellant remaining to be displayed. The totalizer outputs are displayed on the OMS/RCS prplnt qty meters on panel O3 when the rotary switch is positioned to the OMS fuel or oxid positions.
When the wet or dry analog comparator indicates the forward probe is dry, the ungaugeable propellant in the region of the intermediate bulkhead is added to the aft probe output quantity, decreasing the total quantity at a preset rate for 98.15 seconds, and updates from the aft probes are inhibited. After 98.15 seconds of thrusting, the aft probe output inhibit is removed, and the aft probe updates the total quantity. When the quantity decreases to 5 percent, the low-level signal is output.
There are 64 ac -motor-operated valve actuators in the OMS/RCS \Jnitrogen\j tetroxide and monomethyl hydrazine propellant systems. Each valve actuator was modified to incorporate a 0.25-inch-diameter stainless steel sniff line from the actuator to the mold line of the orbiter. The sniff line permits the monitoring of \Jnitrogen\j tetroxide or monomethyl hydrazine in the electrical portion of each valve actuator during ground operations.
There are sniff lines in the 12 ac -motor-operated valve actuators in the forward RCS and in the 44 actuators in the aft left and aft right RCS. The remaining 0.25-inch-diameter sniff lines are in the eight OMS tank isolation and crossfeed ac-motor-operated valve actuators in the left and right orbital maneuvering systems. The 44 aft left and right RCS sniff lines and the eight OMS left and right sniff lines are routed to the respective left and right OMS/RCS pod Y web access servicing panels.
During ground operations, an interscan can be connected to the sniff ports to check for the presence of \Jnitrogen\j tetroxide or monomethyl hydrazine in the electrical portion of the ac-motor-operated valve actuators.
An electrical microswitch in each of the ac-motor-operated valve actuators signals the respective valves' position (open or closed) to the onboard flight crew displays and controls as well as telemetry. An extensive improvement program was implemented to reduce the probability of floating particulates in the electrical microswitch portion of each ac-motor-operated valve actuator. Particulates could affect the operation of the microswitch in each valve and, thus, the position indication of the valves to the onboard displays and controls and telemetry.
Each OMS engine receives pressure-fed propellants at its bipropellant valve assembly. The bipropellant ball valve assembly is controlled by its gaseous \Jnitrogen\j system. The \Jnitrogen\j system consists of a storage tank, engine pressure isolation valve, regulator, relief valve, check valve, accumulator, engine purge valves, bipropellant \Jsolenoid\j control valves and actuators that control the bipropellant ball valves.
A gaseous \Jnitrogen\j spherical storage tank is mounted next to the combustion chamber to supply pressure to its engine pressure isolation valve. The tank contains enough \Jnitrogen\j to operate the ball valves and purge the engine 10 times. Nominal tank capacity is 60 cubic inches. The maximum tank operating pressure is 3,000 psi, with a proof pressure of 6,000 psig.
Each tank's pressure is monitored by two pressure sensors. One sensor transmits the tank pressure to the N 2 , He, kit He switch on panel F7. When the switch is positioned to N 2 , tank pressure is displayed on the OMS press N 2 tank left, right meters on panel F7. The other sensor transmits pressure to telemetry.
A dual-coil, solenoid-operated engine pressure isolation valve is located in each gaseous \Jnitrogen\j system. The valve is energized open and spring-loaded closed. The engine pressure isolation valve permits gaseous \Jnitrogen\j flow from the tank to the regulator, accumulator, the bipropellant ball valve control valves and purge valves 1 and 2 when energized open and isolates the \Jnitrogen\j tank from the gaseous \Jnitrogen\j supply system when closed. The engine pressure isolation valves in each system are controlled by the OMS eng left, right switches on panel C3. When the OMS eng left switch is placed in the arm press position, the left OMS engine pod's pressure isolation valve is energized open.
When the OMS eng right switch is placed in the arm press position, the right OMS engine pod's pressure isolation valve is energized open. The gaseous \Jnitrogen\j engine pressure isolation valve, when energized open, allows gaseous \Jnitrogen\j supply pressure to be directed into a regulator, through a check valve, an in-line accumulator and to a pair of engine bipropellant control valves. The engine bipropellant control valves are controlled by the OMS thrust on/off commands from the GPCs.
A single-stage regulator is installed in each gaseous \Jnitrogen\j pneumatic control system between the gaseous \Jnitrogen\j engine pressure isolation valve and the engine bipropellant control valves. The regulator reduces the gaseous \Jnitrogen\j service pressure to a desired working pressure of 315 to 360 psig.
A pressure relief valve downstream of the gaseous \Jnitrogen\j regulator limits the pressure to the engine bipropellant control valves and actuators if a gaseous \Jnitrogen\j regulator malfunctions. The relief valve relieves between 450 and 500 psig and resets at 400 psig minimum.
A pressure sensor downstream of the regulator monitors the regulated pressure and transmits it to the CRT display and to telemetry.
The check valve located downstream of the gaseous \Jnitrogen\j regulator will close if gaseous \Jnitrogen\j pressure is lost on the upstream side of the check valve and will isolate the remaining gaseous \Jnitrogen\j pressure on the downstream side of the check valve.
The 19-cubic- inch gaseous \Jnitrogen\j accumulator downstream of the check valve and upstream of the bipropellant control valves provides enough pressure to operate the engine bipropellant control valves one time with the engine pressure isolation valve closed or in the event of loss of pressure on the upstream side of the check valve.
Two solenoid-operated, three-way, two-position bipropellant control valves on each OMS engine control the bipropellant control valve actuators and bipropellant ball valves. Control valve 1 controls the No. 1 actuator and the fuel and oxidizer ball valves. Control valve 2 controls the No. 2 actuator and two ball valves, one fuel and oxidizer ball valve in series to the No. 1 system. Each control valve contains two \Jsolenoid\j coils, either of which, when energized, opens the control valve.
The right OMS engine gaseous \Jnitrogen\j \Jsolenoid\j control valves 1 and 2 are energized open by computer commands if the right OMS eng switch on panel C3 is in the arm or arm/press position and the right OMS eng vlv switch on panel O16 is on; the valves are de-energized normally when thrust off is commanded or if the right OMS eng switch is positioned to off . The left OMS engine gaseous \Jnitrogen\j \Jsolenoid\j control valves 1 and 2 are controlled in the same manner, but through the left OMS eng switch on panel C3 and the left OMS eng vlv switch on panel O14.
When the gaseous \Jnitrogen\j \Jsolenoid\j control valves are energized open, pressure is directed into the two actuators in each engine. The \Jnitrogen\j acts against the piston in each actuator, overcoming the spring force on the opposite side of the actuators. Each actuator has a rack-and-pinion gear; and the linear motion of the actuator connecting arm is converted into rotary motion, which drives two ball valves, one fuel and one oxidizer, to the open position. Each pair of ball valves opens simultaneously. Fuel and oxidizer are then directed to the combustion chamber of the engine, where the propellants atomize and ignite upon contact. The hypergolic propellants produce a hot gas, thus thrust.
The chamber pressure of each engine is monitored by a pressure sensor and is transmitted to the OMS press left and right Pc (chamber pressure) meter on panel F7.
When the computer commands thrust off or an engine's OMS eng switch on panel C3 or eng vlv switch on panel O14/O16 is positioned off, the \Jsolenoid\j control valves are de-energized, removing gaseous \Jnitrogen\j pressure from the actuators; and the gaseous \Jnitrogen\j pressure in the actuators is vented overboard through the \Jsolenoid\j control valve. The spring in the actuator forces the actuator's piston to move in the opposite direction, and the actuator drives the fuel and oxidizer ball valves closed simultaneously. The series-redundant arrangement of ball valves ensures engine thrusting is terminated.
Each actuator incorporates a linear position \Jtransducer\j, which supplies ball valve position to a CRT.
Check valves are installed in the vent port outlet of each gaseous \Jnitrogen\j \Jsolenoid\j control valve on the spring pressure side of each actuator to protect the seal of these components from atmospheric contamination.
Each engine has two gaseous \Jnitrogen\j purge valves in series. These valves are solenoid-operated open and spring-loaded closed. They are normally energized open after each thrusting period by the GPCs unless inhibited by a crew entry on the maneuver CRT display. The two purge valves of an engine are energized open 0.36 second after OMS engine thrust off has been commanded and permit gaseous \Jnitrogen\j to flow through the valves and check valve into the fuel line downstream of the ball valves and out through the combustion chamber and engine injector to space for two seconds. This purges the residual fuel from the combustion chamber and injector of the engine, permitting safe engine restart. The purge valves are then de-energized and spring-loaded closed. When the purge is completed, the gaseous \Jnitrogen\j tank pressure isolation valve is closed by placing the respective OMS eng switch (panel C3) to off. The check valve downstream of the purge valves prevents fuel from flowing to the engine purge valves during engine thrusting.
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"OMS Engine Thrust Chamber Assembly",328,0,0,0
When the fuel reaches the thrust chamber, it is directed through 102 coolant channels in the combustion chamber wall, providing regenerative cooling to the combustion chamber walls, and then to the injector of the engine. The oxidizer is routed directly to the injector. The platelet injector assembly consists of a stack of plates, each with an etched pattern that provides proper distribution and propellant injection velocity vector. The stack is diffusion-bonded and welded to the body of the injector. The fuel and oxidizer orifices are positioned so that the propellants will impinge and atomize, causing the fuel and oxidizer to ignite because of hypergolic reaction.
The contoured nozzle extension is bolted to the aft flange of the combustion chamber. The nozzle extension is made of a columbium alloy and is radiantly cooled.
The nominal flow rate of oxidizer and fuel to each engine is 11.93 pounds per second and 7.23 pounds per second, respectively, producing 6,000 pounds of thrust at a vacuum specific impulse of 313 seconds.
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"OMS Thrusting Sequence",329,0,0,0
The OMS thrusting sequence commands the OMS engines on or off and commands the engine purge function. The flight crew can select, via item entry on the maneuver display, a one- or two-engine thrusting maneuver and can inhibit the OMS engine purge.
The sequence determines which engines are selected and then provides the necessary computer commands to open the appropriate \Jhelium\j vapor isolation valves and the engine gaseous \Jnitrogen\j \Jsolenoid\j control valves and sets an engine-on indicator. The sequence will monitor the OMS engine fail flags and, if one or both engines have failed, issue the appropriate OMS cutoff commands as soon as the crew has confirmed the failure by placing the OMS eng switch in the off position. This will then terminate the appropriate engine's control valve commands.
In a normal OMS thrusting period, when the OMS cutoff flag is true, the sequence terminates commands to the \Jhelium\j pressurization, \Jhelium\j vapor isolation valves and two gaseous \Jnitrogen\j engine control valves. If the engine purge sequence is not inhibited, the sequence will check for the left and right engine arm press signals and after 0.36 second open the engine gaseous \Jnitrogen\j purge valves for two seconds for the engines that have the arm press signals present.
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"OMS Engine Thrust Vector Control System",330,0,0,0
The engine TVC system consists of a gimbal ring assembly, two gimbal actuator assemblies and two gimbal actuator controllers. The engine gimbal ring assembly and gimbal actuator assemblies provide OMS TVC by gimbaling the engines in pitch and yaw. Each engine has a pitch actuator and a yaw actuator. Each actuator is extended or retracted by one of a pair of dual-redundant electric motors and is actuated by general-purpose computer control signals.
The gimbal ring assembly contains two mounting pads to attach the engine to the gimbal ring and two pads to attach the gimbal ring to the orbiter. The ring transmits engine thrust to the pod and orbiter.
The pitch and yaw gimbal actuator assembly for each OMS engine provides the force to gimbal the engines. Each actuator contains a primary and secondary motor and drive gears. The primary and secondary drive systems are isolated and are not operated concurrently. Each actuator consists of two redundant brushless dc motors and gear trains, a single jackscrew and nut-tube assembly and redundant linear position feedback transducers.
A GPC position command signal from the primary electronic controller energizes the primary dc motor, which is coupled with a reduction gear and a no-back device. The output from the primary power train drives the jackscrew of the drive assembly, causing the nut-tube to translate (with the secondary power train at idle), which causes angular engine movement.
If the primary power train is inoperative, a GPC position command from the secondary electronic controller energizes the secondary dc motor, providing linear travel by applying \Jtorque\j to the nut-tube through the spline that extends along the nut-tube for the stroke length of the unit. Rotation of the nut-tube about the stationary jackscrew causes the nut-tube to move along the screw. A no-back device in each drive system prevents backdriving of the standby system.
The electrical interface, power and electronic control elements for active and standby control channels are assembled in separate enclosures designated the active actuator controller and standby actuator controller. These are mounted on the OMS/RCS pod structure. The active and standby actuator controllers are electrically and mechanically interchangeable.
The gimbal assembly provides control angles of plus or minus 6 degrees in pitch and plus or minus 7 degrees in yaw with clearance provided for an additional 1 degree for snubbing and tolerances. The engine null position is with the engine nozzles up 15 degrees 49 seconds (as projected in the orbiter XZ plane) and outboard 6 degrees 30 seconds (measured in the 15-degree 49-second plane).
The thrust vector control command subsystem operating program processes and outputs pitch and yaw OMS engine actuator commands and the actuator power selection discretes. The OMS TVC command SOP is active during operational sequences, orbit insertion (OMS-1 and OMS-2), orbit coast, deorbit, deorbit coast and return-to-launch-site abort.
The OMS TVC feedback SOP monitors the primary and secondary actuator selection discretes from the maneuver display and performs compensation on the selected pitch and yaw actuator feedback data. This data is output to the OMS actuator fault detection and identification and to the maneuver display. The OMS TVC feedback SOP is active during orbit insertion (OMS-1 and OMS-2), orbit coast, deorbit maneuver and deorbit maneuver coast. The present OMS gimbal positions can be monitored on the maneuver CRT display when this SOP is active and the primary or secondary actuator motors are selected.
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"OMS Thermal Control",331,0,0,0
OMS thermal control is achieved by \Jinsulation\j on the interior surface of the pods that enclose the OMS hardware components and the use of strip heaters. Wrap around heaters and \Jinsulation\j condition the crossfeed lines. The heaters prevent propellant from freezing in the tanks and lines. The heater system is divided into two areas: the OMS/RCS pods and the aft fuselage crossfeed and bleed lines. Each heater system has two redundant heater systems, A and B, and is controlled by the RCS/OMS heaters switches on panel A14.
Each OMS/RCS pod is divided into eight heater areas. Each of the heater areas in the pods contains an A and B element, and each element has a thermostat that controls the temperature from 55 to 75 F. These heater elements are controlled by the left pod and right pod switches on panel A14. Sensors located throughout the pods supply temperature information to the propellant thermal CRT display and telemetry.
The crossfeed line thermal control in the aft fuselage is divided into 11 heater areas. Each area is heated in parallel by heater systems A and B, and each area has a control thermostat to maintain temperature at 55 F minimum to 75 F maximum. Each circuit also has an overtemperature thermostat to protect against a failed-on heater switch. These heater elements are controlled by the respective crsfd lines switch on panel A14. Temperature sensors near the control thermostats on the crossfeed and bleed lines supply temperature information on the propellant thermal CRT display and telemetry.
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"OMS-RCS Interconnect",332,0,0,0
An interconnect between the OMS crossfeed line and the aft RCS manifolds provides the capability to operate the aft RCS using 1,000 pounds per pod of OMS propellant for orbital maneuvers. The aft RCS may use OMS propellant from either OMS pod in orbit.
The orbital interconnect sequence is available during orbit operations and on-orbit checkout.
The left OMS \Jhelium\j pressure vapor isolation valve A will be commanded open when the left OMS tank (ullage) pressure decays to 236 psig, and the open commands will be terminated 30 seconds later. If the left OMS tank (ullage) pressure remains below 236 psia, the sequence will set an OMS/RCS valve miscompare flag and will set a Class 3 alarm and a CRT fault message. The sequence also will enable the OMS-to-RCS gauging sequence at the same time.
The flight crew can terminate the sequence and inhibit the OMS-to-RCS gauging sequence by use of the OMS press ena-off item entry on the RCS SPEC display. The valves can then be reconfigured to their normal position on panels O7 and O8. The OMS-to-aft-RCS interconnect sequence is not available in the backup flight control system.
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"OMS-To-RCS Gauging Sequence",333,0,0,0
The OMS-to-aft-RCS propellant quantities are calculated by burn time \Jintegration\j. Once each cycle, the accumulated aft RCS thruster cycles are used to compute the OMS propellant used since the initiation of gauging. The number of RCS thruster cycles is provided by the RCS command subsystem operating program to account for minimum-impulse firing of the RCS thrusters. The gauging sequence is initiated by item entry of the OMS right or OMS left interconnect on the RCS SPEC CRT display and is terminated by the return to normal item entry.
The gauging sequence maintains a cumulative total of left and right OMS propellant used during OMS-to-aft-RCS interconnects and displays the cumulative totals as percentage of left and right OMS propellant on the RCS SPEC display. The flight crew will be alerted by a Class 3 alarm and a fault message when the total quantity used from either OMS pod exceeds 1,000 pounds or 8.37 percent.
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"OMS Abort Control Sequences",334,0,0,0
The abort control sequence is the software that manages, among other items, the OMS and aft RCS configuration and thrusting periods during ascent aborts to improve performance or to consume OMS and aft RCS propellants for orbiter center-of-gravity control.
Premission-determined parameters are provided for the OMS and aft RCS thrusting periods during aborts since the propellant loading and orbiter center of gravity vary with each mission.
The premission-determined parameters for the abort-to-orbit thrusting period are modified during flight, based on the vehicle velocity at abort initiation. The premission-determined parameters for abort once around are grouped with different values for early or late AOA. The return-to-launch-site parameters are contained in a single table.
The abort control sequence is available in OPS 1 and 6 and is initiated at SRB separation if selected before then or at the time of selection if after SRB separation.
ATO and AOA Aborts. The OMS and aft RCS begin thrusting as soon as an ATO or AOA is initiated with one main engine out.
For some aborts, an OMS-to-aft-RCS interconnect is not desired. A parallel aft RCS plus X thrusting period using aft RCS propellant and the four aft RCS plus X thrusters will be performed during the OMS-1 thrusting period to achieve the desired orbit.
If a plus X aft RCS thrusting period is required before main engine cutoff, the abort control sequence will command the four aft plus X RCS jets on if vehicle acceleration is greater than 0.8 g and will monitor the RCS cutoff time to terminate the thrusting period. If an RCS propellant dump (burn) is required before MECO and vehicle acceleration is greater than 1.8 g, the abort control sequence will command an eight-aft-RCS-jet null thrust and monitor the RCS cutoff time to terminate the thrusting period.
In other abort cases, an OMS-to-aft-RCS interconnect is desired. This thrusting is performed with the OMS and four aft RCS plus X thrusters to consume OMS propellant for orbiter center-of-gravity control. More aft RCS jets can be commanded if needed to increase OMS propellant usage. For example, for an OMS propellant dump (burn), 14 aft RCS null jets can be commanded to thrust to improve orbiter center-of-gravity location.
If the amount of OMS propellant used before MECO leaves less than 28 percent of OMS propellants, a 15-second aft RCS ullage thrust is performed after MECO to provide a positive OMS propellant feed to start the OMS-1 thrusting period.
The OMS-to-aft-RCS interconnect sequence provides for an automatic interconnect of the OMS propellant to the aft RCS when required and reconfigures the propellant feed from the OMS and aft RCS tanks to their normal state after the thrusting periods have ended. The interconnect sequence is initiated by the abort control sequence.
In order to establish a known configuration of the valves, the interconnect sequence terminates the GPC commands to the following valves if they have not been terminated before honoring a request from the abort control sequence: left and right OMS crossfeed A and B valves, aft RCS crossfeed valves and aft RCS tank isolation valves.
A request from the abort control sequence for an OMS-to-aft-RCS interconnect will sequentially configure the OMS/RCS valves as follows: close the left and right aft RCS propellant tank isolation valves, open the left and right OMS crossfeed A and B valves, and open the left and right aft RCS crossfeed valves. The OMS-to-aft-RCS interconnect complete flag is then set to true.
When the abort control sequence requests a return to normal configuration, all affected OMS/RCS propellant valve commands are removed to establish a known condition; and the interconnect sequence will then sequentially configure the valves as follows: close aft RCS crossfeed valves, close left and right OMS crossfeed valves and open aft RCS propellant tank isolation valves. The OMS-to-aft-RCS reconfiguration complete flag is then set to false, and the sequence is terminated.
Return-to-Launch-Site Abort. An RTLS abort requires the dumping of OMS propellant by burning the OMS propellant through both OMS engines and through the 24 aft RCS thrusters to improve abort performance and to achieve an acceptable entry orbiter vehicle weight and center-of-gravity location. The thrusting period is premission-determined and depends on the OMS propellant load.
The OMS engines start the thrusting sequence; and after the OMS-to-aft-RCS interconnect is complete, the aft RCS thrusters are commanded on. The OMS engines and RCS thrusters then continue their burn for a predetermined period. The interconnect sequence is the same for ATO and AOA aborts. The OMS and aft RCS will begin thrusting at SRB staging if the abort is initiated during the first stage of flight or immediately upon abort initiation during second stage.
Contingency Abort. A contingency abort is selected automatically at the loss of a second main engine or manually by the flight crew using an item entry on the RTLS TRAJ or RTLS TRANS CRT displays. For the contingency aborts, the OMS-to-aft-RCS interconnect is performed in a modified manner to allow continuous flow of propellants to the aft RCS jets for vehicle control and to allow contingency rapid dump (burning) of OMS and RCS propellants. The abort control sequence tracks the total time the OMS and aft RCS are on to determine the amount of propellants used.
The request for an interconnect will cause the interconnect sequence to configure the valves sequentially as follows: open the aft RCS crossfeed valves, open the left OMS crossfeed valves A, open the right OMS crossfeed valves B, close the left and right aft RCS tank isolation valves, open the left OMS crossfeed valves B and open the right OMS crossfeed valves A. The OMS-to-aft-RCS interconnect complete flag will then be set to true.
If the rapid dump is selected before MECO, the OMS-to-aft-RCS interconnect occurs, and both OMS engines and the 24 aft RCS jets are commanded to thrust until the desired amount of propellant has been consumed. The rapid dump will be interrupted during external tank separation if the thrusting period is not completed before MECO; otherwise, the thrusting period terminates when thrusting time equals zero or if the normal acceleration exceeds a threshold value.
Upon completion of the thrusting period, the OMS-to-aft-RCS configuration flag will be set to false, and the sequence will be terminated. A return-to-normal-configuration request by the abort control sequence will cause the interconnect sequence to configure the valves sequentially as follows: open aft RCS propellant tank isolation valves, close the aft RCS crossfeed valves, and close the left and right OMS crossfeed A and B valves. The OMS-to-aft-RCS interconnect complete flag will be set to false, and the sequence will be terminated.
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"OMS Engine Fault Detection and Identification",335,0,0,0
The OMS engine FDI function detects and identifies off-nominal performance of the OMS engine, such as off-failures during OMS thrusting periods, on-failures after or before a thrusting period and high or low engine chamber pressures.
Redundancy management software performs OMS engine FDI. It is assumed that the flight crew arms only the OMS engine to be used; the OMS engine not armed cannot be used for thrusting. FDI will be initialized at SRB ignition and terminated after the OMS-1 thrusting period or, in the case of an RTLS abort, at the transition from RTLS entry to the RTLS landing sequence program. The FDI also will be initiated before each OMS burn and will be terminated after the OMS thrusting period is complete.
The OMS engine FDI uses both a velocity comparison and a chamber pressure comparison method to determine a failed-on or failed-off engine. The velocity comparison is used only after MECO since the OMS thrust is small compared to main propulsion thrust before MECO.
The measured velocity increment is compared to a predetermined one-engine and two-engine acceleration threshold value by the redundancy management software to determine the number of engines actually firing. This information, along with the assumption that an armed engine is to be used, allows the software to determine if the engine has low thrust or has shut down prematurely.
The chamber pressure comparison test compares a predetermined threshold chamber pressure level to the measured chamber pressure to determine a failed engine (on, off or low thrust).
The engine-on command and the chamber pressure are used before MECO to determine a failed engine. The velocity indication and the chamber pressure indication are used after MECO to determine a failed engine. If the engine fails the chamber pressure test but passes the velocity test after MECO, the engine will be considered failed.
Such a failure would illuminate the red right OMS or left OMS caution and warning light on panel F7 and the master alarm and produce a fault message. In addition, if an engine fails the chamber pressure and velocity tests, a down arrow is displayed on the maneuver CRT next to the failed engine.
When the flight crew disarms a failed engine by turning the arm/press switch on panel C3 to off , a signal is sent to the OMS thrusting sequence to shut down the engine and to signal guidance to reconfigure. Guidance reconfigures and downmodes from two OMS engines, to one OMS engine, to four plus X RCS jets.
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"OMS Gimbal Actuator FDI",336,0,0,0
The OMS gimbal actuator FDI detects and identifies off-nominal performance of the pitch and yaw gimbal actuators of the OMS engines.
The OMS gimbal actuator FDI is divided into two processes. The first determines if the actuators should move from their present position. If the actuators must move, the second part determines how much they should move and whether the desired movement has occurred.
The first part checks the actuators' gimbal deflection error (which is the difference between the commanded new position and the actuators' last known position) and determines whether the actuators should extend or retract or if they are being driven against a stop. If the actuators are in the desired position or being driven against a stop, the first part of the process will be repeated. If the first part determines that the actuator should move, the second part of the actuator FDI process is performed.
The second part of the actuator FDI process checks the present position of each actuator against its last known position to determine whether the actuators have moved more than a threshold amount. If the actuators have not moved more than this amount, an actuator failure is incremented by one.
Each time an actuator fails this test, the failure is again incremented by one. When the actuator failure counter reaches an I-loaded value of four, the actuator is declared failed and a fault message is output. The actuator failure counter is reset to zero any time the actuator passes the threshold test.
The first and second parts of the actuator FDI process continue to perform in this manner. The actuator FDI process can detect full-off gimbal failures and full-on failures indirectly. The full-on failure determines that the gimbal has extended or retracted too far and commands reverse motion. If no motion occurs, the actuator will be declared failed. The flight crew's response to a failed actuator is to select the secondary actuator \Jelectronics\j by item entry on the maneuver CRT display.
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"Shuttle Electrical Power System",337,0,0,0
The EPS consists of three subsystems: power reactant storage and distribution, fuel cell power plants (electrical power generation) and electrical power distribution and control.
The PRSD subsystem stores the reactants (cryogenic \Jhydrogen\j and oxygen) and supplies them to the three fuel cell power plants, which generate all the electrical power for the vehicle during all mission phases. In addition, cryogenic oxygen is supplied to the environmental control and life support system for crew cabin pressurization. The \Jhydrogen\j and oxygen are stored in their respective storage tanks at cryogenic temperatures and supercritical pressures. The storage temperature of liquid oxygen is minus 285 F and minus 420 F for liquid \Jhydrogen\j.
The three fuel cell power plants, through a chemical reaction, generate all of the 28-volt direct-current electrical power for the vehicle from launch through landing rollout. Before launch, electrical power is provided by ground power supplies and the onboard fuel cell power plants until T minus three minutes and 30 seconds.
Each fuel cell power plant consists of a power section, where the chemical reaction occurs, and a compact accessory section attached to the power section, which controls and monitors the power section's performance. The three fuel cell power plants are individually coupled to the reactant (hydrogen and oxygen) distribution subsystem, the heat rejection subsystem, the potable water storage subsystem and the EPDC subsystem.
The fuel cell power plants generate heat and water as by-products of electrical power generation. The excess heat is directed to fuel cell heat exchangers, where the excess heat is rejected to Freon coolant loops. The water is directed to the potable water storage subsystem.
The EPDC subsystem distributes the 28 volts dc generated by each of the three fuel cell power plants to a three-bus system that distributes dc power to the forward, mid-, and aft sections of the orbiter for equipment in those areas. The three main dc buses-MNA, MNB and MNC-are the prime sources of power for the vehicle's dc loads.
Each of the three dc main buses supplies power to three solid-state (static), single-phase inverters, which constitute one three-phase alternating-current bus; thus, the nine inverters convert dc power to 115-volt, 400-hertz ac power for distribution to three ac buses-AC1, AC2 and AC3-for the vehicle's ac loads.
The EPDC subsystem controls and distributes electrical power (ac and dc) to the orbiter subsystems, the solid rocket boosters, the external tank and payloads. Power is controlled and distributed by assemblies. Each assembly is a housing for electrical components, such as remote switching devices, buses, resistors, diodes and fuses. Each assembly usually contains a power bus or buses and remote switching devices for distributing bus power to subsystems located in its area.
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"Shuttle Power Reactant Storage and Distribution",338,0,0,0
Cryogenic \Jhydrogen\j and oxygen are stored in a supercritical condition in double-walled, thermally insulated spherical tanks with a vacuum annulus between the inner pressure vessel and outer shell of the tank. Each tank has heaters to add energy to the reactants during depletion to control pressure. Each tank is capable of measuring quantity remaining.
The tanks are grouped in sets consisting of one \Jhydrogen\j and one oxygen tank. The number of tank sets installed depends on the specific mission requirement. Up to five tank sets can be installed. The five tank sets are all installed in the midfuselage under the payload bay liner.
The oxygen tanks are identical and consist of inner pressure vessels of Inconel 718 and outer shells of aluminum 2219. The inner vessel is 33.43 inches in diameter and the outer shell is 36.8 inches in diameter. Each tank has a volume of 11.2 cubic feet and stores 781 pounds of oxygen. The dry weight of each tank is 201 pounds. The initial temperature of the stored oxygen is minus 285 F. Maximum fill time is 45 minutes.
The \Jhydrogen\j tanks also are identical. Both the inner pressure vessel and the outer shell are constructed of aluminum 2219. The inner vessel's diameter is 41.51 inches and the outer shell's is 45.5 inches. The volume of each tank is 21.39 cubic feet, and each stores 92 pounds of \Jhydrogen\j. Each tank weighs 216 pounds dry. The initial storage temperature is minus 420 F. Maximum fill time is 45 minutes.
The inner pressure vessels are kept supercold by minimizing conductive, convective and radiant heat transfer. Twelve low-conductive supports suspend the inner vessel within the outer shell. Radiant heat transfer is reduced by a shield between the inner vessel and outer shell (hydrogen tanks only), and convective heat transfer is minimized by maintaining a vacuum between the vessel and shell. A vacuum ion pump maintains the required vacuum level and is also used as a vacuum gauge to determine the vacuum's integrity.
Dual-mode heater operation is available for pairs of oxygen and \Jhydrogen\j tanks. If the heaters of both tanks 1 and 2 or tanks 3 and 4 are placed in the automatic mode, the tank heater logic is interconnected. In this case, the heater controllers of both tanks must send a low pressure signal to the heater logic before the heaters will turn on. Once the heaters are on, a high pressure signal from either tank will turn off the heaters in both tanks.
In the manual mode, the flight crew controls the heaters by using the on/off positions for each heater switch on panel R1 or A11. High or low pressure in each tank is shown on the CRT display or the gauges on panel O2. The specific tank is selected by setting the rotary switch on panel O2.
Before lift-off, the oxygen and \Jhydrogen\j tank 1 and 2 heater switches are set on auto. After SRB separation, all the \Jhydrogen\j and oxygen tank 1 and 2 heater switches are positioned to auto, and the tank 3 and 4 heaters remain off.
On orbit, the tank 3 and 4 heater switches are positioned to auto. Because the tank 3 and 4 heater controller pressure limits are higher than those of tanks 1 and 2, tanks 3 and 4 supply the reactants to the fuel cells. For entry, the tank 3 and 4 heater switches are set to off, and tanks 1 and 2 supply the reactants to the fuel cells.
The cryo oxygen htr assy temp meter on panel O2, in conjunction with the rotary switch tk1 1-2, tk2 1-2, tk3 1-2, tk4 1-2, selects one of the two heaters in each tank and permits the temperature of the heater element to be displayed. The range of the display is from minus 425 F to plus 475 F. The temperature sensor in each heater also is hard-wired directly to the yellow O 2 heater temp caution and warning light on panel F7.
This light is illuminated if the temperature is at or above 349 F. A signal also is sent to the computers, where software checks the limit; and if the temperature is at or above 349 F, the backup C/W alarm light on panel F7 is illuminated. This signal also is transmitted to the CRT and telemetry.
Each tank set (one \Jhydrogen\j and one oxygen tank) has a hydrogen/oxygen control box that contains the electrical logic for the \Jhydrogen\j and oxygen heaters and controllers. The control box is located on cold plates in the midbody under the payload bay envelope.
The reactants from the tanks flow through two relief valve/filter package modules and valve modules and then to the fuel cells through a common manifold. Oxygen is supplied to the manifold from the tank at a pressure of 815 to 881 psia, and \Jhydrogen\j is supplied at a pressure of 200 to 243 psia. The pressure of the reactants will be essentially the same at the fuel cell interface as it is in the tanks since only a small decrease in pressure occurs in the manifolds.
The relief valve/filter package module contains the tank relief valve and a 12-micron filter. The filter removes contaminants that could affect the performance of components within the power reactant storage and distribution subsystem and fuel cells. The valve relieves excessive pressure that builds up in the tank, and a manifold valve relieves pressure in the manifold lines. The oxygen tank relief valve relieves at 1,005 psia, and the \Jhydrogen\j tank relief valve relieves at 310 psia.
The reactants flow from the relief valve/filter packages through four reactant valve modules: two \Jhydrogen\j (hydrogen valve modules 1 and 2) and two oxygen (oxygen modules 1 and 2). Each valve module contains a check valve for each cryogenic tank line to prevent the reactants from flowing from one tank to another tank in the event of a tank leak. This prevents a total loss of reactants. The oxygen valve modules also contain the environmental control and life support system atmosphere pressure control system 1 and 2 oxygen supply. Each module also contains a manifold valve and fuel cell reactant valves.
Each fuel cell reactant valve consists of two valves-one for \Jhydrogen\j and one for oxygen. The valves are controlled by the fuel cell 1, 2, 3 reac open/close switches on panel R1. When the switch is positioned to open, the \Jhydrogen\j and oxygen reactant valves for that fuel cell are opened, and reactants are allowed to flow from the manifold into the fuel cell.
When the switch is positioned to close, the \Jhydrogen\j and oxygen reactant valves for that fuel cell are closed, isolating the reactants from the fuel cell and rendering that fuel cell inoperative. Each fuel cell reac switch on panel R1 also has a talkback indicator. The corresponding talkback indicator indicates op when both valves are open and cl when either valve is closed.
Because it is critical to have reactants available to the fuel cells, the red fuel cell reac light on panel F7 is illuminated when any fuel cell reactant valve is closed, a caution/warning tone is sounded, and the computers sense the closed valve, which causes the backup C/W alarm light on panel F7 to be illuminated, an SM alert to occur, and the data to be displayed on the CRT. This alerts the flight crew that the fuel cell will be inoperative within approximately 20 seconds for a \Jhydrogen\j valve closure and 130 seconds for an oxygen valve closure.
The manifold relief valves are a built-in safety device in the event a manifold valve and fuel cell reactant valves are closed because of a malfunction. The reactants trapped in the manifold lines would be warmed up by the internal heat of the orbiter and overpressurize. The manifold relief valve will open at 290 psi for \Jhydrogen\j and 975 psi for oxygen to relieve pressure and allow the trapped reactants to flow back to their tanks.
Two pressure sensors located in the respective \Jhydrogen\j and oxygen valve modules transmit data to the CRT. This data is also sent to the systems management computer, where its lower limit is checked; and if the respective \Jhydrogen\j and oxygen manifold pressures are below 150 psia and 200 psia, respectively, an SM alert will occur.
If cryogenic tank set 5 is added to an orbiter, the displays and controls associated with controlling the tank set will be added to panel A15.
During prelaunch operations, the onboard fuel cell reactants (oxygen and hydrogen) are supplied by ground support equipment to assure a full load of onboard reactants before lift-off. At T minus two minutes 35 seconds, the GSE filling operation is terminated. The GSE supply pressure is 300 to 320 psia for \Jhydrogen\j and 1,000 to 1,020 psia for oxygen, which is higher than the onboard PRSD pressures. The GSE supply valves close automatically to transfer to onboard reactants.
#
"Shuttle Fuel Cell Power Plants",339,0,0,0
Each of the three fuel cell power plants is reusable and restartable. The fuel cells are located under the payload bay area in the forward portion of the orbiter's midfuselage.
The three fuel cells operate as independent electrical power sources, each supplying its own isolated, simultaneously operating 28-volt dc bus. The fuel cell consists of a power section, where the chemical reaction occurs, and an accessory section that controls and monitors the power section's performance. The power section, where \Jhydrogen\j and oxygen are transformed into electrical power, water and heat, consists of 96 cells contained in three substacks.
Manifolds run the length of these substacks and distribute \Jhydrogen\j, oxygen and coolant to the cells. The cells contain \Jelectrolyte\j consisting of \Jpotassium\j \Jhydroxide\j and water, an oxygen electrode (cathode) and a \Jhydrogen\j electrode (anode).
The accessory section monitors the reactant flow, removes waste heat and water from the chemical reaction and controls the temperature of the stack. The accessory section consists of the \Jhydrogen\j and oxygen flow system, the coolant loop and the electrical control unit.
Oxygen is routed to the fuel cell's oxygen electrode, where it reacts with the water and returning electrons to produce hydroxyl ions. The hydroxyl ions then migrate to the \Jhydrogen\j electrode, where they enter into the \Jhydrogen\j reaction. \JHydrogen\j is routed to the fuel cell's \Jhydrogen\j electrode, where it reacts with the hydroxyl ions from the \Jelectrolyte\j.
This electrochemical reaction produces electrons (electrical power), water and heat. The electrons are routed through the orbiter's EPDC subsystem to perform electrical work. The oxygen and \Jhydrogen\j are reacted (consumed) in proportion to the orbiter's electrical power demand.
Excess water vapor is removed by an internal circulating \Jhydrogen\j system. \JHydrogen\j and water vapor from the reaction exits the cell stack, is mixed with replenishing \Jhydrogen\j from the storage and distribution system, and enters a condenser, where waste heat from the \Jhydrogen\j and water vapor is transferred to the fuel cell coolant system.
The resultant temperature decrease condenses some of the water vapor to water droplets. A centrifugal water separator extracts the liquid water and pressure-feeds it to potable tanks in the lower deck of the pressurized crew cabin. Water from the potable water storage tanks can be used for crew consumption and cooling the Freon-21 coolant loops. The remaining circulating \Jhydrogen\j is directed back to the fuel cell stack.
The fuel cell coolant system circulates a liquid fluorinated hydrocarbon and transfers the waste heat from the cell stack through the fuel cell heat exchanger of the fuel cell power plant to the Freon-21 coolant loop system in the midfuselage. Internal control of the circulating fluid maintains the cell stack at a normal operating temperature of approximately 200 F.
When the reactants enter the fuel cells, they flow through a preheater (where they are warmed from a cryogenic temperature to 40 F or greater); a 6-micron filter; and a two-stage, integrated dual gas regulator module. The first stage of the regulator reduces the pressure of the \Jhydrogen\j and oxygen to 135 to 150 psia.
The second stage reduces the oxygen pressure to a range of 62 to 65 psia and maintains the \Jhydrogen\j pressure at 4.5 to 6 psia differential below the oxygen pressure. The regulated oxygen lines are connected to the accumulator, which maintains an equalized pressure between the oxygen and the fuel cell coolant. If the oxygen's and \Jhydrogen\j's pressure decreases, the coolant's pressure is also decreased to prevent a large differential pressure inside the stack that could deform the cell stack structural elements.
Upon leaving the dual gas regulator module, the incoming \Jhydrogen\j mixes with the hydrogen-water vapor exhaust from the fuel cell stack. This saturated gas mixture is routed through a condenser, where the temperature of the mixture is reduced, condensing a portion of the water vapor to form liquid water droplets. The liquid water is then separated from the hydrogen-water mixture by the \Jhydrogen\j pump/water separator.
The \Jhydrogen\j pump circulates the \Jhydrogen\j gas back to the fuel cell stack, where some of the \Jhydrogen\j is consumed in the reaction. The remainder flows through the fuel cell stack, removing the product water vapor formed at the \Jhydrogen\j electrode. The hydrogen-water vapor mixture then combines with the regulated \Jhydrogen\j from the dual gas generator module, and the loop begins again.
The oxygen from the dual gas regulator module flows directly through two ports into a closed-end manifold in the fuel cell stack, achieving optimum oxygen distribution in the cells. All oxygen that flows into the stack is consumed, except during purge operations.
Reactant consumption is directly related to the electrical current produced: if there are no internal or external loads on the fuel cell, no reactants will be used. Because of this direct proportion, leaks may be detected by comparing reactant consumption and current produced. An appreciable amount of excess reactants used indicates a probable leak.
Water and electricity are the products of the chemical reaction of oxygen and \Jhydrogen\j that takes place in the fuel cells. The water must be removed or the cells will become saturated with water, decreasing reaction efficiency. With an operating load of about 7 kilowatts, it takes only a few minutes to flood the fuel cell with produced water, thus effectively halting power generation. \JHydrogen\j is pumped through the stack, reacting with oxygen and picking up and removing water vapor on the way. After being condensed, the liquid water is separated from the \Jhydrogen\j by the \Jhydrogen\j pump/water separator and discharged from the fuel cell to be stored in the ECLSS potable water storage tanks.
If the water tanks are full or there is line blockage, the water relief valves open at 45 psia to allow the water to vent overboard through the water relief line and nozzle. Check valves prevent water tanks from discharging through an open relief valve. An alternate water delivery path is also available to deliver water to the ECLSS tanks if the primary path is lost.
For redundancy, there are two thermostatically activated heaters wrapped around the discharge and relief lines to prevent blockage caused by the formation of ice in the lines. Two switches on panel R12, fuel cell H\D2\dO line htr and H\D2\dO relief htr , provide the flight crew with the capability to select either auto A or auto B for the fuel cell water discharge line heaters and the water relief line and vent heaters, respectively.
Thermostatically controlled heaters will maintain the water line temperature above 53 F, when required. The normal temperature of product water is approximately 140 to 150 F. The thermostatically controlled heaters maintain the water relief valve's temperature when in use between 70 to 100 F. Temperature sensors located on the fuel cell water discharge line, relief valve, relief line and vent nozzle are displayed on the CRT.
If the \Jpotassium\j \Jhydroxide\j \Jelectrolyte\j in the fuel cell migrates into the product water, a \JpH\j sensor located downstream of the \Jhydrogen\j pump/water separator will sense the presence of the \Jelectrolyte\j, and the crew will be alerted by an SM alert and display on the CRT.
During normal fuel cell operation, the reactants are present in a closed-loop system and are 100 percent consumed in the production of electricity. Any inert gases or other contaminants will accumulate in and around the porous electrodes in the cells and reduce the reaction efficiency and electrical load support capability. Purging, therefore, is required at least twice daily to cleanse the cells.
When a purge is initiated by opening the purge valves, the oxygen and \Jhydrogen\j systems become open-loop systems; and increased flows allow the reactants to circulate through the stack, pick up the contaminants and blow them out overboard through the purge lines and vents. Electrical power is produced throughout the purge sequence, although no more than 10 kilowatts should be required from a fuel cell being purged because of the increased reactant flow and preheater limitations.
Fuel cell purge can be activated automatically or manually by the use of fuel cell switches on panel R12. In the automatic mode, the fuel cell purge heater switch is positioned to GPC . The purge line heaters are turned on to heat the purge lines to ensure that the reactants will not freeze in the lines. The \Jhydrogen\j reactant is the more likely to freeze because it is saturated with water vapor.
Depending on the orbit trajectory and vehicle orientation, the heaters may require 27 minutes to heat the lines to the required temperatures. The fuel cell current is checked to ensure a load of less than 350 amps, due to limitations on the \Jhydrogen\j and oxygen preheaters in the fuel cells.
As the current output of the fuel cell increases, the reactant flow rates increase, and the preheaters raise the temperature of the reactants to a minimum of minus 40 F in order to prevent the seals in the dual gas regulator from freezing.
The purge lines from all three fuel cells are manifolded together downstream of their purge valves and associated check valves. The line leading to the purge outlet is sized to permit unrestricted flow from only one fuel cell at a time. If purging of more than one cell at a time is attempted, pressure could build in the purge outlet line and cause a decrease in the flow rate through the individual cells, which would result in an inefficient purge.
When the fuel cell purge valves 1, 2 and 3 switches are positioned to GPC, the fuel cell GPC purge seq switch is positioned to start and must be held until the GPC purge seq talkback indicator indicates gray (in approximately three seconds). The automatic purge sequence will not begin if the indicator indicates barberpole. The GPC turns the purge line heaters on and monitors the temperature.
The one oxygen line temperature sensor must register at least 69 F and the two \Jhydrogen\j line temperature sensors 79 and 40 F, respectively, and be verified by the GPC before the purge sequence begins. If the temperatures are not up to minimum after 27 minutes, the GPC will issue an SM alert and display the data on the CRT.
When the proper temperatures have been attained, the GPC will open for two minutes and then close the \Jhydrogen\j and oxygen purge valves for fuel cells 1, 2 and 3 in that order. Thirty minutes after the fuel cell 3 purge valves have been closed (to ensure that the purge lines have been totally evacuated), the GPC will turn off the purge line heaters. This provides sufficient time and heat to bake out any remaining water vapor. If the heaters are turned off before 30 minutes have elapsed, water vapor left in the lines may freeze.
In order to cool the fuel cell stack during its operations, distribute heat during fuel cell starting, and warm the cryogenic reactants entering the stack, the fuel cell circulates a coolant-fluorinated hydrocarbon-throughout the fuel cell. The fuel cell coolant loop and its interface with the ECLSS Freon-21 coolant loops are identical in fuel cells 1, 2 and 3.
Where the coolant enters the fuel cell, the temperature of the F-40 coolant returning from the ECLSS Freon-21 coolant loops is sensed before it passes through a 75-micron filter. After the filter, two temperature-controlled mixing valves allow some of the hot coolant to mix with the cool returning coolant to prevent the condenser exit control valve from oscillating. The condenser exit control valve adjusts the flow of the coolant through the condenser to maintain the hydrogen-water vapor exiting the condenser at a temperature between 148 and 153 F.
The stack inlet control valve maintains the temperature of the coolant entering the stack between 177 and 187 F. The accumulator is the interface with the oxygen cryogenic reactant to maintain an equalized pressure between the oxygen and the coolant (the oxygen and \Jhydrogen\j pressures are controlled at the dual gas regulator) to preclude a high-pressure differential in the stack. The pressure in the coolant loop is sensed before the coolant enters the stack.
The coolant is circulated through the fuel cell stack to absorb the waste heat from the hydrogen/oxygen reaction occurring in the individual cells. The coolant pump utilizes three-phase ac power to circulate the coolant through the loop.
The coolant (that which is not made to bypass) exits the fuel cells to the fuel cell heat exchanger, where it transfers its excess heat to be dissipated through the ECLSS Freon-21 coolant loop systems in the midfuselage.
In addition to thermal conditioning by means of the coolant loop, the fuel cell has internal startup and sustaining heaters. The 2,400-watt startup heater is used only during startup to warm the fuel cell to its operational level. The 1,100-watt sustaining heaters normally are used during low power periods to maintain the fuel cells at their operational temperature.
Two 160-watt end-cell electrical heaters on each fuel cell power plant were used to maintain a uniform temperature throughout the fuel cell power section. As an operational improvement, the end-cell electrical heaters on each fuel cell power plant were deleted due to potential electrical failures and were replaced by fuel cell power plant coolant (F-40) passages. This permits waste heat from each fuel cell power plant to be used to maintain a uniform temperature profile for each fuel cell power plant.
The \Jhydrogen\j pump and water separator of each fuel cell power plant were also improved. To minimize excessive \Jhydrogen\j gas entrained in each fuel cell power plant's product water, modifications were made to the water pickup (pitot) system. The centrifugal force of high-velocity water flowing around the pitot tube's bends separates the \Jhydrogen\j gas and water. Pitot pressure then expels the \Jhydrogen\j gas into the \Jhydrogen\j pump's inlet housing though a bleed orifice.
A current measurement detection system was added to monitor the \Jhydrogen\j pump load for each fuel cell power plant. Excessive load could indicate improper water removal, which could lead to flooding of the fuel cell power plant and eventually render that power plant inoperative.
The start/sustaining heater system for each fuel cell power plant was also modified. The modification was required specifically for fuel cell power plant No. 1, mounted on the port, or left, side. The No. 1 fuel cell power plant start/sustaining heater system added heat to that fuel cell power plant's F-40 coolant loop system during the startup of the power plant. Because of its orientation, any entrained gas in the coolant could enter the heater and become trapped at the heater elements.
This would result in overheating of the heater elements, which could vaporize the F-40 coolant, causing heater failure and extensive damage to the fuel cell power plant. The F-40 coolant loop flow system within the start/sustaining heater of each fuel cell power plant was modified to prevent a gas bubble from developing or being trapped at the heater elements, preventing the loss of the start/sustaining heater.
A stack inlet temperature measurement was added to each fuel cell power plant. The temperature measurement was added to the in-flight system to provide full visibility of the thermal conditions of each fuel cell power plant (similar to the existing stack exit and condenser exit temperatures of each fuel cell power plant).
The product water from all three fuel cell power plants flows to a single water relief control panel. The water can be directed from the single panel to the ECLSS potable water tank A or to the fuel cell power plant water relief nozzle. Normally, the water is directed to water tank A. In the event of a line rupture in the vicinity of the single water relief panel, water could spray on all three water relief panel lines, causing them to freeze and prevent fuel cell power plant water discharge.
The product water lines from all three fuel cell power plants were modified to incorporate a parallel (redundant) path of product water to ECLSS potable water tank B in the event of a freeze-up of the single water relief panel. In the event of the single water relief panel freeze-up, pressure would build up and relieve through the redundant paths to water tank B. Temperature sensors and a pressure sensor installed on each of the redundant water line paths transmit data via telemetry for ground monitoring.
A water purity sensor (pH) was added at the common product water outlet of the water relief panel to provide a redundant measurement of water purity. A single measurement of water purity in each fuel cell power plant was provided previously. If the fuel cell power plant \JpH\j sensor failed, the flight crew was required to sample the potable water.
The electrical control unit located in each fuel cell power plant is the brain of the power plants. The ECU contains the start up logic, heater thermostats, and 30-second timer and interfaces with the controls and displays for fuel cell startup, operation and shutdown. The ECU controls the supply of ac power to the coolant pump, \Jhydrogen\j pump/water separator, the \JpH\j sensor, and the dc power supplied to the flow control bypass valve (open only during startup) and the internal startup and sustaining heaters. The ECU also controls the status of the fuel cell 1, 2, 3 ready for load and coolant pump P talkback indicators on panel R1.
The nine fuel cell circuit breakers that connect the three-phase ac power to the three fuel cells are located on panel L4, and the fuel cell ECU receives its power from an essential bus through the FC cntlr switch on panel O14.
The fuel cell start/stop switch on panel R1 for each fuel cell is used to initiate the start sequence or stop the fuel cell operation. When this switch is held in its momentary start position, the ECU connects the three-phase ac power to the coolant pump and \Jhydrogen\j pump/water separator (allowing the coolant and the hydrogen-water vapor to circulate through these loops) and connects the dc power to the internal startup and sustaining heaters and the flow control bypass valve. The switch must be held in the start position until the coolant pump P talkback shows gray in approximately three to four seconds, which indicates that the coolant pump is functioning properly by creating a differential pressure across the pump. When the coolant pump P talkback indicates barberpole, it indicates the coolant pump is not running.
The ready for load talkback for each fuel cell will show gray after the 30-second timer times out and the stack-out temperature is above 187 F (which can be monitored on panel O2 in conjunction with the 1, 2, 3 switch located beneath the fuel cell stack out temp meter). This indicates that the fuel cell is up to the proper operating temperature and is ready for loads to be attached to it.
It should not take longer than 25 minutes for the fuel cell to warm up and become fully operational, the actual time depending on the fuel cell's initial temperature. The ready for load indicator remains gray until the fuel cell start/stop switch for each fuel cell is placed to stop, the FC cntlr switch is placed to off , or the essential bus power is lost to the ECU.
The startup heater enable/inhibit switch on panel R12 for each fuel cell provides the crew control of the off/on status of the startup heaters during fuel cell startup. The inhibit position allows the startup heaters to remain off and would be used only when immediate power is required from a shutdown fuel cell.
Fuel cell 1, 2 or 3 dc voltage and current (amps) can be monitored on the dc volts and dc amps meters on panel F9, using the fuel cell volts/amp rotary switch to select a specific fuel cell.
The fuel cells will be on when the crew boards the vehicle, and the vehicle is powered by the fuel cells and load sharing with the ground support equipment power supplies. Just before lift-off (T minus three minutes and 30 seconds), the GSE is powered off and the fuel cells take over all of the vehicle's electrical loads. Indication of the switchover can be noted on the CRT display and the dc amps meter.
The fuel cell current will increase to approximately 220 amps; the oxygen and \Jhydrogen\j flow will increase to approximately 4 and 0.6 pound per hour, respectively; and the fuel cell stack temperature will increase slightly.
Fuel cell standby consists of removing the electrical loads but continuing operation of the fuel cell pumps, controls, instrumentation and valves, with the electrical power being supplied by the remaining fuel cells. A small amount of reactants is used to generate power for the fuel cell internal heaters.
Fuel cell shutdown, after standby, consists of stopping the coolant pump and \Jhydrogen\j pump/water separator by positioning that fuel cell start/stop switch on panel R1 to the stop position. If the temperature in the fuel cell compartment beneath the payload bay is lower than 40 F, the fuel cell should be left in standby instead of being shut down to prevent it from freezing.
Each fuel cell power plant is 14 inches high, 15 inches wide and 40 inches long and weighs 255 pounds.
#
"Shuttle Environmental Control and Life Support System",340,0,0,0
The ECLSS consists of an air revitalization system, water coolant loop systems, atmosphere revitalization pressure control system, active thermal control system, supply water and waste water system, waste collection system and airlock support system. These systems interact to provide a habitable environment for the flight crew in the crew compartment in addition to cooling or heating various orbiter systems or components.
The ARS controls relative \Jhumidity\j between 30 and 75 percent, maintains carbon dioxide and carbon monoxide at non-toxic levels, controls temperature and ventilation in the crew compartment, and provides cooling to various flight deck and middeck electronic avionics and the crew compartment.
The ARS consists of water coolant loops, cabin air loops and pressure control. Cabin air is ducted to the crew compartment cabin heat exchanger, where the cabin air is cooled by the WCLs; therefore, cabin air cools the crew cabin, flight crew and crew compartment electronic avionics. The water coolant loop system collects heat from the crew compartment cabin heat exchanger and heat from some of the electronic units in the crew compartment and transfers it to the water coolant/Freon-21 coolant loop heat exchanger of the ATCS.
The ATCS provides orbiter heat rejection during all phases of the mission. It consists of two Freon-21 coolant loops, cold plate networks for cooling electronic avionics units, liquid/liquid heat exchangers for cooling various orbiter systems, and four heat sink systems for rejecting excess heat outside the orbiter-ground support equipment heat exchanger, flash evaporators, radiator panels and ammonia boilers.
The Freon-21 coolant loops transport excess heat from the fuel cell power plant heat exchangers, payload heat exchangers and midbody and aft avionics electronic units; heat the hydraulic systems; and deliver that heat to the heat sinks. During checkout, prelaunch and postlanding ground operations, the GSE heat exchanger in the orbiter's Freon-21 coolant loops rejects excess heat from the orbiter through ground systems cooling.
Approximately 125 seconds after lift-off, the flash evaporator system is activated and provides orbiter heat rejection of the Freon-21 coolant loops via water boiling. When the orbiter is on orbit and the payload bay doors are opened, radiator panels on the underside of the doors are exposed to space and provide heat rejection. If combinations of heat loads and orbiter attitude exceed the capacity of the radiator panels during on-orbit operations, the flash evaporator can be activated to meet the heat rejection requirements.
At the conclusion of orbital operations, the payload bay doors are closed, rendering the radiator panels inoperative for heat rejection; and the flash evaporator is again brought into operation through deorbit and entry until atmospheric pressure buildup no longer permits the boiling water to provide adequate cooling at approximately 100,000 feet altitude. At this point the ammonia boilers reject heat from the Freon-21 coolant loops by evaporating ammonia through the remainder of entry, landing and postlanding until ground cooling is connected to the GSE heat exchanger.
The ARPCS controls crew compartment cabin pressure at 14.7 psia, plus or minus 0.2 psia, with an average of 80-percent \Jnitrogen\j and 20-percent oxygen mixture. Oxygen partial pressure is maintained between 2.95 psia and 3.45 psia, with sufficient \Jnitrogen\j pressure of 11.5 psia added to achieve the cabin total pressure of 14.7 psia, plus or minus 0.2 psia.
The pressurization control system receives oxygen from two power reactant storage and distribution cryogenic oxygen systems in the midfuselage of the orbiter. Gaseous \Jnitrogen\j is supplied from two \Jnitrogen\j systems consisting of two \Jnitrogen\j tanks for each system located in the midfuselage of the orbiter. An optional mission kit consists of an emergency gaseous oxygen tank, and the system can be located in the midfuselage of the orbiter. The gaseous \Jnitrogen\j system is also used to pressurize the potable and waste water tanks located below the crew compartment middeck floor.
Potable water produced by the three fuel cell power plants is directed and stored in potable water tanks for flight crew consumption and personal hygiene. The potable water system is the supply to the flash evaporator system when it is used to cool the Freon-21 coolant loops. A waste water tank is also located below the crew compartment middeck floor to collect waste water from the crew cabin heat exchanger and flight crew waste water. Solid waste remains in the waste management system in the crew compartment middeck until the orbiter is serviced during ground turnaround operations.
The orbiter crew compartment provides a life-sustaining environment for a flight crew of eight. The crew cabin volume with the airlock inside the middeck is 2,325 cubic feet. For extravehicular activity requirements, only the airlock is depressurized and repressurized. If the airlock is located outside of the middeck in the payload bay, the crew cabin volume would be 2,625 cubic feet.
The cabin is pressurized to 14.7 psia, plus or minus 0.2 psia, and maintained at an average 80-percent \Jnitrogen\j and 20-percent oxygen mixture by the air revitalization system. Oxygen partial pressure is maintained between 2.95 and 3.45 psi, with sufficient \Jnitrogen\j pressure of 11.5 psia added to achieve the cabin total pressure of 14.7 psia, plus or minus 0.2 psia.
The pressurization system consists of two oxygen systems and two gaseous \Jnitrogen\j systems. The two oxygen systems are supplied by the PRSD oxygen system, which is the same source that supplies oxygen to the orbiter fuel cell power plants. The PRSD cryogenic supercritical oxygen storage system is controlled by electrical heaters within the tanks and supplies the oxygen to the ECLSS pressurization control system at a pressure of 835 to 852 psia in a gaseous state.
The gaseous \Jnitrogen\j supply system consists of two systems with two gaseous \Jnitrogen\j tanks for each system. The \Jnitrogen\j storage tanks are serviced to a nominal pressure of 2,964 psia at 80 F. If the auxiliary gaseous oxygen supply tank is installed, it is serviced to 2,440 psia at 80 F and stores 67.6 pounds of gaseous oxygen to provide high flow along with gaseous \Jnitrogen\j.
It would maintain the crew cabin at 8 psi with oxygen partial pressure at 2 psia. For normal on-orbit operations one oxygen and \Jnitrogen\j supply system is used. For launch and entry both oxygen and \Jnitrogen\j supply systems are used in addition to repressurization of the airlock.
The heart of the cabin pressurization is the nitrogen/oxygen control and supply panels, the PPO 2 sensor, and crew cabin positive and negative pressure relief valves. The nitrogen/oxygen control panel selects and regulates primary (system 1) or secondary (system 2) oxygen and \Jnitrogen\j. The primary and secondary nitrogen/oxygen supply panels are located in the lower forward portion of the midfuselage. The primary and secondary oxygen supply systems have a crossover capability, as do the primary and secondary \Jnitrogen\j supply systems. If installed, the auxiliary oxygen supply system is also controlled by the supply panel.
The oxygen and \Jnitrogen\j supply systems provide the makeup cabin oxygen gas consumed by the flight crew and \Jnitrogen\j for pressurizing the potable and waste water tanks and repressurizing the airlock. An average of 1.76 pounds of oxygen is used per flight crew member per day. Up to 7.7 pounds of \Jnitrogen\j and 9 pounds of oxygen are expected to be used per day for normal loss of crew cabin gas to space and metabolic usage. The potable and waste water tanks are pressurized to 17 psia.
The \Jnitrogen\j regulators in the primary and secondary supply system reduce the pressure to 200 psi. Each \Jnitrogen\j regulator is a two-stage regulator with the second stage functioning as a relief valve. The second stage relieves pressure overboard at 245 psi.
The regulated pressure of each \Jnitrogen\j system is directed to the \Jnitrogen\j manual crossover valve, the water tank regulator inlet valve and the oxygen and \Jnitrogen\j controller valve in each system.
The \Jnitrogen\j crossover manual valve connects both regulated \Jnitrogen\j systems when the valve is open and isolates the \Jnitrogen\j supply systems from each other when closed. A check valve between the \Jnitrogen\j regulator and \Jnitrogen\j crossover valve in each nitrogen-regulated supply line prevents flow from one \Jnitrogen\j source supply pressure to the other if the \Jnitrogen\j crossover valve is open.
The partial pressure of oxygen in the flight crew cabin can be controlled automatically by one of two oxygen and \Jnitrogen\j controllers. Two PPO2 sensors are located under the crew cabin flight deck mission support console. The PPO 2 A and B sensors provide inputs to the PPO2 control systems 1 and 2 controller and switches, respectively.
The oxygen systems 1 and 2 and \Jnitrogen\j systems 1 and 2 flows are monitored and sent to the O 2 /N 2 flow rotary switch on panel O1. The rotary switch permits system 1 oxygen or \Jnitrogen\j or system 2 oxygen or \Jnitrogen\j flow to be monitored on the flow meter on panel O1 in pounds per hour.
PPO2 sensors A and B monitor the oxygen partial pressure and transmit the signal to the PPO 2 sensor select switch on panel O1. When the switch is positioned to sensor A, oxygen partial pressure from sensor A is monitored on the PPO 2 meter on panel O1 in psia. If the switch is set on sensor B, oxygen partial pressure from sensor B is monitored. The cabin pressure sensor transmits directly to the cabin press meter on panel O1 and is monitored in psia.
The red cabin atm caution and warning light on panel F7 is illuminated for any of the following monitored parameters:
- Cabin pressure below 14.0 psia or above 15.4 psia.
- PPO2 below 2.8 psia or above 3.6 psia.
- Oxygen flow rate above 5 pounds per hour.
- Nitrogen flow rate above 5 pounds per hour.
A klaxon will sound in the crew cabin and the master alarm push button light indicators will be illuminated if the change in pressure versus change in time decreases at a rate of 0.05 psi per minute or greater. The normal cabin dP/dT is zero psi per minute, plus or minus 0.01 psi, for all normal operations.
The temperature and pressure of the primary and secondary \Jnitrogen\j and emergency oxygen tanks are monitored and transmitted to the systems management computer. This information is used to compute oxygen and \Jnitrogen\j quantities.
The two cabin relief valves are in parallel to provide overpressurization protection of the crew module cabin above 16 psid. Each cabin relief valve is controlled by its corresponding switch on panel L2. The cabin relief A switch controls cabin relief A, and the cabin relief B switch controls cabin relief B.
When the switch is positioned to enable, the corresponding motor-operated valve allows the cabin pressure to a corresponding positive pressure relief valve that relieves at 16 psid and reseats at 15.5 psid. The relief valve maximum flow capability is 150 pounds per hour. A talkback indicator above the respective switch indicates barberpole when the motor-operated valve is in transit and op when the motor-operated valve is open. When the switch is positioned to close , the corresponding motor-operated valve isolates cabin pressure from the relief valve, and the talkback indicator indicates cl .
The crew module cabin vent isolation valve and cabin vent valve are in series to vent the crew cabin to ambient pressure. Approximately one hour and 30 minutes before lift-off, the crew module cabin is pressurized to approximately 16.7 psi for leak checks of the crew cabin.
The cabin vent isolation valve is controlled by the cabin vent , vent isol switch on panel L2, and the cabin vent valve is controlled by the cabin vent, vent switch on panel L2. Each switch is positioned to open to control its respective motor-operated valve.
When both valves are open, the cabin pressure is vented into the midfuselage. The maximum flow capability through the valves at 0.2 psid is 900 pounds per hour. A talkback indicator above each switch indicates the position of the respective valve-barberpole when the valve is in transit and op when it is open.
If the crew cabin pressure is lower than the pressure outside the crew cabin, two negative pressure relief valves in parallel will open at 0.2 to 0.7 psid, permitting flow of ambient pressure into the crew cabin. The maximum flow rate at 0.5 psid is zero to 654 pounds per hour.
The manual water tank \Jnitrogen\j regulator inlet valve in each nitrogen-regulated supply system permits \Jnitrogen\j to flow to its corresponding regulator and water tank \Jnitrogen\j regulator isolation manual valve. The inlet and isolation manual valves are on panel M010W.
The water regulator in each \Jnitrogen\j system reduces the 200-psi supply pressure to 16 psi. Each regulator is a two-stage regulator with the second stage relieving pressure into the crew cabin at a differential pressure of 20 psi. The \Jnitrogen\j pressurization system for the potable and waste water tanks is discussed later in this section.
#
"Shuttle Cabin Air Revitalization",342,0,0,0
There are five independent air loops in the cabin: the cabin itself, three avionics bays and inertial measurement units. The cabin pressure atmosphere is circulated by the air revitalization system. The air circulated through the flight crew cabin picks up heat, moisture, odor, carbon dioxide and debris with additional heat from electronic units in the crew cabin. The cabin air is drawn through the cabin loop and through a 300-micron filter by one of two cabin fans located downstream of the filter.
Each cabin fan is controlled by its respective cabin fan A and B switch on panel L1. Normally, only one fan is used at a time. The cabin fans are located under the middeck floor.
The cabin air from the cabin fan is ducted to the two \Jlithium\j \Jhydroxide\j canisters, where carbon dioxide is removed and activated charcoal removes odors and trace contaminants. An orifice in the duct directs a specific amount of cabin air through each \Jlithium\j \Jhydroxide\j canister.
The canisters are also located under the middeck floor. They are changed alternately every 12 hours through an access door in the floor. For a flight crew of seven, the \Jlithium\j \Jhydroxide\j canisters are changed alternately every 11 hours. Replacement canisters are stored under the middeck floor between the cabin heat exchanger and water tanks.
Cabin air is then directed to the crew cabin heat exchanger located under the middeck floor and cooled by the water coolant loops. \JHumidity\j condensation that forms in the slurper of the cabin heat exchanger is removed by a fan separator that draws air and water from the cabin heat exchanger. The moist air is drawn from the slurper into the \Jhumidity\j separator fan, where centrifugal force separates the water from the air.
The fan separator removes up to approximately 4 pounds of water per hour. The water is routed to the waste water tank, and the air is ducted through the exhaust for return to the cabin. There are two fan separators controlled individually by \Jhumidity\j sep A and B switches on panel L1. The \Jhumidity\j sep A switch controls \Jhumidity\j separator fan A, and the B switch controls \Jhumidity\j separator fan B.
Normally, only one fan separator is used at a time. The relative \Jhumidity\j in the crew cabin is maintained between 30 and 65 percent in this manner. A small portion of the revitalized and conditioned air from the cabin heat exchanger is ducted to the carbon monoxide removal unit, which converts carbon monoxide to carbon dioxide.
Based on the crew cabin volume of 2,300 cubic feet and 330 cubic feet of air per minute, one volume crew cabin air change occurs in approximately seven minutes, and approximately 8.5 air changes occur in one hour.
The cabin heat exchanger outlet temperature is transmitted to the cabin hx out av bay rotary switch on panel O1. When the switch is positioned to cabin hx out , the temperature can be monitored on the panel O1 air temp meter. The cabin heat exchanger outlet temperature provides an input to the yellow av bay/cabin air caution and warning light on panel F7. The C/W light is illuminated if the cabin heat exchanger outlet temperature is above 65 F or if the cabin fan delta pressure is 2.8 inches of water or above 7.1 inches of water.
The air from the cabin heat exchanger and the bypassed air come together in the supply duct and are exhausted into the crew cabin through consoles and middeck and various station duct outlets into the crew cabin.
If cabin temperature controllers 1 and 2 or the cabin temp cool/warm rotary switch is unable to control the single bypass valve, the flight crew can position the single bypass valve actuator drive arm to the desired position and pin the bypass valve in place at panel MD44F.
The full cool position at panel MD44F establishes the maximum cabin air flow rate to the cabin heat exchanger, the 2/3 cool position establishes a flow rate that provides approximately two-thirds of the maximum cooling capability, the 1/3 cool position establishes a flow rate that provides approximately one-third of the maximum cooling capability, and the full heat position establishes the minimum cabin air flow rate to the cabin heat exchanger.
The cabin air is also used to cool the three avionics equipment bays and some of the electronic avionics units in the avionics bays in addition to the three IMUs. Each of the three avionics equipment bays in the middeck has a closeout cover to minimize air interchange and thermal gradients between the avionics bay and crew cabin; however, the closeout cover for each avionics equipment bay is not airtight.
The electronic avionics units in each avionics bay meet outgassing and flammability requirements to minimize toxicity levels. Each of the three avionics equipment bays has identical air-cooled systems. Two fans per avionics equipment bay are controlled by individual avionics bay fan A and B switches on panel L1. Normally, only one fan is used at a time.
When the A or B switch for an avionics bay is positioned to on, the fan draws cabin air from the floor of the avionics bay, through the applicable air-cooled avionics units, through connectors at the back of the applicable air-cooled units, to the cabin fan inlet, through a 300-micron filter and to the cabin fan. The cabin fan outlet directs the air through that avionics bay heat exchanger.
The water coolant loops flow through the heat exchanger to cool the fan outlet air, and the cooled air is returned to the avionics bay. A check valve in the outlet of the fan that is not operating prevents a reverse flow through that fan. The air outlet from the fan in each avionics bay is monitored and transmitted to the cabin hx out av bay 1,2,3 rotary switch on panel O1.
When the rotary switch is positioned to av bay 1, 2 or 3 , that avionics bay's fan outlet temperature can be displayed on the air temp meter on panel O1. The air outlet temperature of each avionics bay also provides an input to the yellow av bay/cabin air C/W light on panel F7. This light is illuminated if any of the avionics bay outlet temperatures are above 135 F. The off position of the A or B switch removes power from that avionics bay fan.
The three IMUs are cooled by one of three fans drawing cabin air through a 300-micron filter and across the three IMUs. The fan outlet air flows through the IMU heat exchanger and is cooled by the water coolant loops before returning to the crew cabin. Each IMU fan is controlled by the IMU fan A, B, C switches on panel L1. The on position turns the corresponding fan on and the off position turns it off. Normally, one fan is sufficient because one fan cools all three IMUs. A check valve is installed on the outlet of each fan to prevent a reverse air flow through the fans not operating.
If the payload bay contains the Spacelab pressurized module, a kit is installed to provide ducting for the flow of cabin air from the middeck through the airlock and tunnel to the module. The \Jhumidity\j separators, cabin fans, cabin heat exchanger, avionics bay heat exchangers, IMU heat exchanger, waste water tank, \Jlithium\j \Jhydroxide\j filters, carbon monoxide unit, and waste and potable water tanks are located beneath the middeck crew compartment floor.
#
"Shuttle Water Coolant Loop System",343,0,0,0
The WCLS provides thermal conditioning of the crew cabin by collecting heat at the cabin-air-to-water-coolant-loop heat exchanger and transfers heat to the water coolant loops. The water coolant loops transfer the heat at the water and Freon-21 coolant loop interchanger. The WCLS also provides thermal conditioning for the three avionics bays by an air-to-water heat exchanger in each avionics bay, the cabin heat exchanger, liquid-cooled garment and water chiller, which transfers heat to the water coolant loops. The IMU air-to-water heat exchanger also transfers heat to the water coolant loops.
#
"Shuttle Supply and Waste Water",344,0,0,0
The supply and waste water systems provide water for the flash evaporator, crew consumption and hygiene. The supply water system stores water generated by the fuel cell power plants, and the waste water system stores waste from the crew cabin \Jhumidity\j separator and from the flight crew. There are four supply water tanks and one waste water tank located beneath the crew compartment middeck floor.
Each of the four potable water tanks has a usable capacity of 165 pounds, is 35.5 inches in length and 15.5 inches in diameter, and weighs 39.5 pounds dry.
The waste water tank's usable capacity is 165 pounds. It is 35.5 inches in length and 15.5 inches in diameter and weighs 39.5 pounds dry.
The three fuel cell power plants generate a maximum of 25 pounds of potable water per hour. The product water from all three fuel cell power plants flows to a single water relief control panel. The water can be directed to potable water tank A or to the fuel cell power plant water relief nozzle.
Normally, the water is directed to water tank A. If a line ruptured in the vicinity of the single water relief panel, water could spray on all three water relief panel lines, causing them to freeze and preventing the fuel cell power plants from discharging water, which would cause flooding of the fuel cell power plants.
The product water lines from all three fuel cell power plants were modified to incorporate a parallel (redundant) path of product water to potable water tank B in the event of a freeze-up of the single water relief panel. In the event of a water freeze-up, pressure would build up and relieve through the redundant paths to potable water tank B. Temperature sensors are installed on each of the redundant paths in addition to a pressure sensor that is transmitted to telemetry.
A water purity sensor (pH) was added at the common product water outlet of the water relief panel. It provides a redundant measurement of water purity-a single measurement of water purity in each fuel cell power plant was provided previously. If the single fuel cell power plant \JpH\j sensor failed, the flight crew was required to sample the potable water.
The hydrogen-enriched water from the fuel cell power plants that flows through the single water relief panel to potable tank A passes through two \Jhydrogen\j separators, where 85 percent of the excess \Jhydrogen\j is removed. The \Jhydrogen\j separators consist of a matrix of silver \Jpalladium\j tubes, which have an affinity for \Jhydrogen\j. The \Jhydrogen\j is dumped overboard through a vacuum vent.
Water passing through the \Jhydrogen\j separators can be stored in all four potable water tanks. The four potable water tanks are identified as tanks A, B, C and D. The water entering tank A passes through a microbial filter that adds approximately one-half parts per million \Jiodine\j to the water.
The water stored in tank A is normally used for flight crew consumption but could also be used for flash evaporator cooling. The water from the microbial check valve is also directed to a galley supply valve. If the water tank A inlet valve is closed or tank A is full, water is directed to tank B through a 1.5-psid check valve where it branches off to tank B. If the tank B inlet valve is closed or tank B is full of water, the water is directed through another 1.5-psid check valve to the inlets to tanks C and D.
Each potable water tank has an inlet and outlet valve that can be opened or closed selectively to use water; however, the tank A outlet valve normally remains closed since the water has been treated by passage through the microbial filter for flight crew consumption.
The controls and displays for the potable water tank supply system are located on panels R12 and ML31C. Potable water tanks A, B and C are controlled from panel R12, and tank D is controlled from panel ML31C.
When the supply H\D2\dO inlet tk A, B or C switch on panel R12 is positioned to open , the inlet valve for the tank permits water into that tank. A talkback indicator next to the corresponding switch on panel R12 indicates op when the corresponding valve is open, barberpole when the valve is in transit and cl when that valve is closed. When the switch is positioned to close, the water inlet to that tank is isolated from the inlet water supply. The supply H\D2\dO tk inlet D switch and talkback indicator are located on panel ML31C and operate in the same manner as the switches and talkbacks for tanks A, B and C.
Each potable water and waste water tank is pressurized with gaseous \Jnitrogen\j from the crew compartment \Jnitrogen\j supply system. \JNitrogen\j supply systems 1 and 2 can be used individually to pressurize the tanks with \Jnitrogen\j at 16 psig. \JNitrogen\j supply system 1 is controlled by the water tank \Jnitrogen\j regulator inlet and water tank \Jnitrogen\j isolation 1 manual valves from panel M010W.
Nitrogen supply system 2 is controlled by the water tank \Jnitrogen\j regulator inlet and water tank \Jnitrogen\j isolation 2 manual valves from panel M010W. The regulator in each \Jnitrogen\j supply system controls the \Jnitrogen\j pressure to the tanks at 16 psig, and a relief valve in each \Jnitrogen\j supply system will relieve into the crew cabin if the \Jnitrogen\j supply increases to 18.5 psig, plus or minus 1.5 psig, to protect the tanks from overpressurization.
For only tank A, inlet \Jnitrogen\j pressure is controlled by pressure and vent manual valves on panel ML26C. When the tank A isolation valve is closed, the tank A vent valve is opened to the crew cabin atmosphere. For launch the tank A isolation valve is closed, which lowers tank A pressure so the fuel cell power plants' water head pressure is lower to help prevent flooding of the fuel cell power plants during ascent.
On orbit the tank A isolation valve is opened, and the tank A vent to the cabin is closed, allowing \Jnitrogen\j supply pressure to tank A and inhibiting cabin atmosphere to tank A. \JNitrogen\j supply pressure is supplied to tanks B, C and D to support flash evaporator operation.
If neither \Jnitrogen\j supply system 1 nor 2 can be used to pressurize the water tanks, the H\D2\dO alternate press switch on panel L1 can be positioned to open, which would reference the water tank pressurization system to the crew compartment ambient pressure. Normally, this switch is positioned to close to isolate the cabin pressurization system from the water tank pressurization system.
The supply H\D2\dO outlet tk A, B or C switch on panel R12 positioned to open permits water from the corresponding tank to flow from the tank into the water outlet manifold by the tank \Jnitrogen\j pressurization system. A talkback indicator next to the switch would indicate op when that valve is open, barberpole when it is in transit and cl when it is closed.
The close position of each switch isolates that water tank from the water outlet manifold. The supply H\D2\dO tk inlet D switch and talkback indicator are located on panel ML31C and operate in the same manner as the tank A, B and C switches and talkback indicators on panel R12.
If the potable water tank A or B outlet valve is opened, water from the corresponding tank is directed to the water outlet manifold. The tank A and B water is then available to the extravehicular mobility unit fill in the airlock, to the flash evaporator water supply system A and to the water dump.
As stated previously, the tank A outlet valve is normally closed to prevent contamination of the water in tank A. Thus, tank B would supply water to flash evaporator water supply system A and to the EMU fill in the airlock. If it is necessary to provide space for storing water in tank A and/or B, tank A and/or B water can be dumped overboard.
Tank C or D is normally saved full of water for contingency purposes. If the tank C or D outlet valve is opened, water from either tank is directed to the water outlet manifold. The water is then available to the flash evaporator water supply system B.
A crossover valve installed in the water outlet manifold is controlled by the supply H\D2\dO crossover vlv switch on panel R12. When the switch is positioned to open , the crossover valve opens and allows tank A or B to supply flash evaporator water supply systems A and B, the EMU fill in the airlock and water dump. It would also allow tank C or D to supply flash evaporator water supply systems A and B, the EMU fill in the airlock and water dump.
A talkback indicator next to the switch indicates op when the crossover valve is opened, barberpole when the valve is in transit and cl when the valve is closed. The close position isolates the water manifold between the tank A and B outlets and the tank C and D outlets.
Water from supply system A is routed directly to the flash evaporator. Water from system B is routed to an isolation valve in the system. The valve is controlled by the supply H\D2\dO B supply isol vlv switch on panel R12. When the switch is positioned to open , water from supply system B is directed to the flash evaporator. A talkback indicator next to the switch indicates op when the valve is opened, barberpole when it is in transit and cl when the valve is closed. The close position isolates water supply system B from the flash evaporator.
Potable water from tank A or B can be dumped overboard, if necessary, through a dump isolation valve and a dump valve. Potable water from tank C or D can also be dumped overboard, if necessary, through the crossover valve and through the dump isolation valve and dump valve. The overboard dump isolation valve is located in the crew cabin, and the dump valve is located in the midfuselage.
The dump isolation valve is controlled by the supply H\D2\dO dump isol vlv switch on panel R12; the dump valve is controlled by supply H\D2\dO dump vlv switch on panel R12. The supply H\D2\dO dump valve enable/noz htr switch on panel R12 must be positioned to on to supply electrical power to the supply H\D2\dO dump vlv switch.
When each switch is positioned to open , the corresponding valve is opened, which allows potable water to be dumped overboard. A talkback indicator next to each switch indicates op when the corresponding valve is open, barberpole when it is in transit and cl when it is closed. Closing either valve inhibits the dumping of potable water. At the completion of the dump, each switch is positioned to close to close the corresponding valve.
A contingency crosstie valve, in the supply water overboard dump line between the dump isolation valve and dump valve, permits the joining of the waste water system through a flexible hose to the supply water system for emergency dumping of waste water through the supply water dump or the use of waste water for the flash evaporators.
The potable supply water dump nozzle employs a heater to prevent freezing of the supply water dump nozzle at the midfuselage. The dump nozzle heater is powered when the supply H\D2\dO dump vlv enable/noz htr switch on panel R12 is positioned to on . When the switch is positioned to off, it removes electrical power from the nozzle heater as well as the supply H\D2\dO dump vlv switch, which causes the dump valve to close.
The potable supply water line upstream of the water dump nozzle has electrical heaters on the line to prevent supply water from freezing. The A and B heaters on the line are thermostatically controlled and are powered by the H\D2\dO line heater A and B circuit breakers on panel M186B. (These circuit breakers also provide power to thermostatically controlled heaters on the waste water line and the waste collection system vacuum vent line.)
The potable supply water feed lines to the flash evaporators are approximately 100 feet long. To prevent the water in the lines from freezing, redundant heaters are installed along the length of the water lines. The heaters are controlled by the flash evap feed-line heater A supply and B supply switches on panel L2.
When a switch is positioned to 1, it enables the thermostatically controlled heaters on the corresponding supply line to automatically control the temperature on that line. When a switch is positioned to 2 , it enables another thermostatically controlled heater system on the corresponding supply line. The off position of each switch inhibits heater operation on the corresponding supply line.
The galley supply valve in the supply water line from the microbial filter permits or isolates the supply water from the microbial filter to the middeck ECLSS supply water bay. When the switch is positioned to open , supply water is routed through parallel paths; one path flows through the ARS water coolant loop water chiller to cool the supply water, and the other path bypasses the water chiller with ambient water. A talkback indicator next to the switch indicates op when the valve is open, barberpole when the valve is in transit and cl when the valve is closed. The close position of the switch isolates the potable supply water from the middeck ECLSS supply water panel.
If a payload is installed in the middeck in lieu of the galley, the chilled water and ambient water are connected to an Apollo water dispenser to dispense ambient and chilled water for drinking and food reconstitution. The temperature of the chilled water is within the range of 43 to 55 F, and the ambient water's temperature is between 65 to 95 F. A personal hygiene dispenser is also provided with ambient water.
If the galley is installed in the middeck of the crew cabin, the water supply is directed to the galley. A hot-water heater heats water to 155 to 165 F. Chilled water is also provided at 45 to 55 F.
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"Shuttle Waste Collection System",345,0,0,0
The waste collection system is an integrated, multifunctional system used primarily to collect and process biological wastes from crew members in a zero-gravity environment. The WCS is located in the middeck of the orbiter crew compartment in a 29-inch -wide area immediately aft of the crew ingress and egress side hatch. The commode is 27 by 27 by 29 inches and is used like a standard toilet.
The system collects, stores and dries fecal wastes and associated tissues; processes urine and transfers it to the waste water tank; processes EMU condensate water from the airlock and transfers it to the waste water tank if an extravehicular activity is required on a mission; provides an interface for venting trash container gases overboard; provides an interface for dumping ARS waste water overboard in a contingency situation; and transfers ARS waste water to the waste water tank.
A door on the waste management compartment and two privacy curtains attached to the inside of the compartment door provide privacy for crew members. One curtain is attached to the top of the door and interfaces with the edge of the interdeck access, and the other is attached to the door and interfaces with the galley, if installed.
The door also serves as an ingress platform during prelaunch (vertical) operations since the flight crew must enter the flight deck over the waste management compartment. The door has a friction hinge and must be open to gain access to the waste management compartment.
The WCS consists of a commode, urinal, fan separators, odor and bacteria filter, vacuum vent quick disconnect and waste collection system controls. The commode contains a single multilayer \Jhydrophobic\j porous bag liner for collecting and storing solid waste.
When the commode is in use, it is pressurized, and transport air flow is provided by the fan separator. When the commode is not in use, it is depressurized for solid waste drying and deactivation. The urinal is essentially a funnel attached to a hose and provides the capability to collect and transport liquid waste to the waste water tank. The fan separator provides transport air flow for the liquid. The fan separators separate the waste liquid from the air flow.
The liquid is drawn off to the waste water tank, and the air returns to the crew cabin through the odor and bacteria filter. The filter removes odors and bacteria from the air that returns to the cabin. The vacuum quick disconnect is used to vent liquid directly overboard from equipment connected to the quick disconnect through the vacuum line.
The urinal can accommodate both males and females. The urinal assembly is a flexible hose with attachable funnels for males or females. It can be used in a standing position or can be attached to the commode by a pivoting mounting bracket for use in a sitting position.
All waste collection system gases are ducted from the fan separator into the odor and bacteria filter and then mixed with cabin air. The filter can be removed for in-flight replacement.
The system employs various restraints and adjustments to enable the user to achieve the proper body positioning to urinate or defecate in a zero-gravity environment. Two foot restraints are provided. A toe bar is located at the commode base and is used to urinate standing.
It consists of two flexible cylindrical pads on a shaft that can be adjusted to various heights by releasing two locking levers that are turned 90 degrees counterclockwise. The crew member is restrained by slipping the feet under the toe bar restraint. A footrest restrains the feet of a crew member sitting on the commode. It consists of an adjustable platform with detachable Velcro straps for securing the feet.
The Velcro straps are wrapped crosswise over each foot and secured around the back. The footrest can be adjusted to various angles and heights. Two locking handles pulled outward adjust the angle; two other locking levers adjust the height of the footrest.
Two body restraints are provided for use when crew members are seated on the commode. The primary restraint is a thigh bar that the crew member lifts up out of the detent position, rotates over the thigh and releases. It exerts a preloaded force on each thigh of approximately 10 pounds. The second restraint is a backup method. It consists of four Velcro fabric thigh straps with a spring hook on one end.
Two of the straps are attached to the top front commode surface mating attach points, and the other two are installed on a bracket with five holes on the upper sides of the commode, below and outboard of the thigh bars. The crew member is secured in position by wrapping two straps over each thigh and attaching the mating Velcro surfaces.
Handholds are used for positioning or stabilizing the crew member while using the WCS and form an integral part of the top cover of the waste management collection system assembly.
The controls on the waste collection system are the vacuum valve, fan separator select switch, mode switch, fan separator bypass switches and commode control handle. The system uses dc power to control the fan separators and ac power for fan separator operations. The mode switch and the commode control handle are mechanically interlocked to prevent undesirable system configurations. The remaining controls operate independently. The fan separator bypass switches allow the crew member to manually override a fan separator limit switch failure.
For launch and entry the vacuum valve is closed. During on-orbit operations when the WCS is not in use, the vacuum valve is opened. This exposes the commode (overboard) via the vacuum vent system, and any solid wastes in the commode are dried. The \Jhydrophobic\j bag liner in the commode allows gas from the commode to vent overboard, but does not allow the passage of free liquid.
In the urine collection mode, the vacuum valve remains in open . The fan sep select switch is positioned to 1 or 2 . When positioned to 1, main bus A dc power is supplied to the mode switch; and when positioned to 2 , MNB dc power is supplied to the mode switch. The mode switch positioned to WCS/EMU energizes a relay for a fan separator (dependent on fan sep 1 or 2 position).
The active fan separator pulls cabin air flow through the urinal at 8 cubic feet per minute and ballast cabin air through the wet-trash storage modules at 30 cubic feet per minute. The ballast air mixes with the urine transport air flow in the fan separator. The fan separator is designed to operate at 38 cubic feet per minute and thus requires the 30-cubic- feet-per-minute ballast air flow. Liquid check valves at the waste water outlet from each fan separator provide a back pressure for proper separator operation and prevent backflow through the non-operating separator.
The liquid and air mixture from the urinal line enters the fan separator axially and is carried to a rotating chamber. The mixture first contacts a rotating impact separator that throws the liquid to the outer walls of the rotating fluid reservoir. Centrifugal force separates the liquid and draws it into a stationary pitot tube in a reservoir and directs the liquid to the waste water tank. Air is drawn out of the rotating chamber by a blower that passes the air through the odor and bacteria filter, where it mixes with cabin air and re-enters the crew cabin.
In the EMU and airlock water collection mode, a guard is rotated over the mode switch to preclude WCS use or deactivation during the EMU and airlock water collection mode. A urinal protective screen cap is installed on the urinal because it cannot be used during the EMU dump because of possible separator flooding. The EMU dump will be used only if an EVA is required on a mission. The EMU waste water is dumped through waste water valves in the airlock. Other than these requirements, EMU dump is the same as the urine collection mode.
In the urine and feces collection mode, the commode control handle is pulled up, and the commode is pressurized with cabin air through the debris screen and flow restrictor in approximately 20 seconds. (Note that if the mode switch is positioned to off , the handle cannot be pulled up because of a mechanical interlock.) The commode control handle is positioned to push fwd after 20 seconds (it cannot be pushed forward until after 20 seconds because of the delta pressure across the sliding gate valve, and it cannot be pushed forward unless the mode switch is positioned all the way to the WCS/EMU position).
When the commode control handle is pushed forward, the sliding gate valve on the commode is opened. The commode outlet control valve and ballast air control valves are positioned to connect the commode to the fan separator, and the commode pressurization valve is closed. The WCS is used like a normal toilet. The commode seat is made of a contoured, compliant, semisoft material that provides proper positioning of the user and is sealed to minimize air leakage. Feces enter the commode through the 4-inch-diameter seat opening and are drawn in by cabin air flowing through holes under the seat at 30 cubic feet per minute.
Fecal matter and tissues are deposited on the porous bag liner, and the air is drawn through the \Jhydrophobic\j material to the fan separator. The \Jhydrophobic\j liner material prevents free liquid and bacteria from leaving the collector. (Toilet tissue is the only paper item permitted to be disposed of in the commode.) Urine is processed as in the urine collection mode. The off/down position closes the sliding gate valve and depressurizes the commode for deactivation and solid waste drying.
If the handle is left partially up, it would cause loss of cabin air through the vacuum vent. After usage, the WCS should be cleaned with wet wipes, if required, to maintain an odorless and sanitary environment. The seat can be lifted for cleaning, and the WCS should be cleaned once a day with a biocidal cleanser. The urinal should also be cleaned and flushed with water once a day.
The vacuum vent quick disconnect provides the capability to dump ARS waste water overboard through the vacuum vent by connecting a water transfer hose to the vacuum vent quick disconnect and the waste water crosstie quick disconnect.
If fan separator 1 is inoperative or fails to achieve proper operational speed (which can be verified by a reduced noise level or lack of air flow), the fan sep switch is positioned from 1 to 2, and fan separator 2 will operate in the same manner as 1.
The fan sep 1 bypass and fan sep 2 bypass switches permit the crew members to manually override a fan separator limit switch failure either in the fan sep or mode switches. The bypass switches are located on the waste collector and are lever-locked. When either switch is positioned to on, dc power is applied to the corresponding relay, energizing it and providing ac power to activate the corresponding fan separator.
Both separator bypass switches should not be on at the same time. Before the fan sep bypass switch is activated, the fan sep select switch should be positioned in the corresponding fan separator position to preset the fan separator inlet valve, and the mode switch should be positioned to WCS/EMU to preset the urine collection valve.
A vacuum vent isolation valve is located in the vacuum vent line from the waste collector to the overboard vacuum line. It is controlled by the vacuum vent control switch on panel ML31C. This switch receives electrical power from the vacuum vent bus select switch on panel ML31C when the bus select switch is positioned to MNA or MNB.
When the control switch is positioned to open , the vacuum vent isolation valve is opened, allowing the vacuum vent line to be open to vacuum. A talkback indicator next to the control switch indicates op when the valve is open, barberpole when the valve is in transit and cl when the valve is closed. The off position of the control switch closes the valve.
Thermostatically controlled heaters are installed on the vacuum vent line. Electrical power for the A and B heaters are from the respective H\D2\dO line htr A and B circuit breakers on panel ML86B. (These circuit breakers also supply electrical power to supply water dump line A and B heaters and waste water line A and B heaters.)
Heaters are also installed on the vacuum vent nozzle and are controlled by the vacuum vent noz htr switch on panel ML31C. Electrical power is supplied to the vacuum vent nozzle heaters when the switch is positioned to on. The off position removes electrical power from the vacuum vent nozzle heaters.
If both fan separators in the waste collection system fail, feces are collected by the Apollo fecal bag. To dispose of the Apollo fecal bag, the waste collection system is configured as in the urine and feces collection mode and the bag is stowed in the commode.
If both fan separators in the waste collection system fail and it is not possible to dump urine overboard, urine may be collected using a contingency urine collection device.
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"Shuttle Waste Water Tank",346,0,0,0
A single waste water tank receives waste water from the ARS \Jhumidity\j separator and the waste collection system. The tank is located beneath the crew compartment middeck floor next to the potable water tanks.
The waste water tank holds 165 pounds, is 35.5 inches long and 15.5 inches in diameter, and weighs 39.5 pounds dry. It is pressurized by gaseous \Jnitrogen\j from the same source as the potable water tanks.
Waste water is directed to the waste water tank 1 inlet valve, which is controlled by the waste H\D2\dO tank 1 switch on panel ML31C. When the switch is positioned to open , waste water is directed to the waste water tank. A talkback indicator next to the switch indicates op when the valve is open, barberpole when the valve is in transit and cl when the valve is closed.
The switch positioned to off closes the waste water tank inlet, isolating the waste water tank from the waste water collection system. When the valve is open, waste water from the tank can also be directed to the waste water dump for overboard dumping.
The waste water dump isolation valve and waste water dump valve in the waste water dump line allow waste water to be dumped overboard through the waste water dump. The waste H\D2\dO dump isol valve switch on panel ML31C positioned to open allows waste water to be directed to the waste water dump valve. A talkback indicator above the waste H\D2\dO dump isol valve switch indicates op when the valve is open, barberpole when the valve is in transit and cl when the valve is closed.
In order for waste water to be dumped overboard, the waste water dump valve must be opened. It is controlled by the waste H\D2\dO dump valve enable/nozzle heater switch and the waste H\D2\dO dump valve switch on panel ML31C. When the waste H\D2\dO dump valve enable/nozzle heater switch is positioned to on , electrical power is supplied to the waste water dump heaters and the waste H\D2\dO dump valve switch.
When the waste H\D2\dO dump valve switch is positioned to open, the dump valve allows waste water to be dumped overboard. A talkback indicator above the dump switch indicates op when the valve is open, barberpole when the valve is in transit and cl when the valve is closed. If waste water is dumped overboard, the dump isolation valve switch is positioned to cl upon completion of the dump.
The waste H\D2\dO dump valve is positioned to cl , and the waste H\D2\dO dump valve enable/nozzle heater switch is set to off . (If the heater switch is positioned to off before the dump valve switch is positioned to close, the dump valve will remain open.) The heaters at the waste water dump prevent waste water from freezing at the overboard dump.
The waste water dump line, upstream of the waste dump nozzle, has electrical heaters on the line to prevent waste water from freezing. The A and B heaters are powered by the H\D2\dO line htr A and B circuit breakers on panel ML86B and are thermostatically controlled. (These circuit breakers also provide power to thermostatically controlled heaters on the supply water line and waste collection system vacuum vent line.)
The contingency crosstie quick disconnect in the waste water overboard dump line between the dump isolation valve and dump valve permits waste water to be joined with the supply water system through a flexible hose for emergency dumping of supply water through the waste water dump or using waste water for the flash evaporators.
The waste water tank 1 drain valve controls the draining of the waste water tank during ground operations through the ground support equipment flush and drain. When the waste H\D2\dO drain tank 1 valve switch on panel ML31C is positioned to open , the valve permits the draining and flushing of the waste water tank. The drain line is capped during flight. A talkback indicator above the switch indicates op when the valve is open, barberpole when the valve is in transit and cl when the valve is closed.
Personal hygiene accommodations for flight crew members include personal hygiene kits, pressure-packed personal hygiene agents and towel tissue dispensers. The personal hygiene kits have provisions for tooth brushing, hair care, shaving, nail care, etc. The pressure-packaged personal hygiene agents are for cleaning the hands, face and body.
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"Shuttle Airlock Support",347,0,0,0
An airlock is located in the crew cabin middeck. The airlock and airlock hatches permit EVA flight crew members to transfer from the middeck crew compartment into the payload bay in EMUs without depressurizing the orbiter crew cabin.
Normally, two EMUs are stowed in the airlock. The EMU is an integrated space suit assembly and life support system that enables flight crew members to leave the pressurized orbiter crew cabin and work in space.
The airlock has an inside diameter of 63 inches, is 83 inches long and has two 40-inch- diameter D-shaped openings that are 36 inches across, plus two pressure sealing hatches and a complement of airlock support systems. The airlock's volume is 150 cubic feet.
The airlock is sized to accommodate two fully suited flight crew members simultaneously. The airlock support provides airlock depressurization and repressurization, EVA equipment recharge, liquid-cooled garment water cooling, EVA equipment checkout, donning and communications. All EVA gear, checkout panel and recharge stations are located against the internal walls of the airlock.
The airlock hatches are mounted on the airlock. The inner hatch is mounted on the exterior of the airlock (orbiter crew cabin middeck side) and opens into the middeck. The inner hatch isolates the airlock from the orbiter crew cabin. The outer hatch is mounted in the interior of the airlock and opens into the airlock. The outer hatch isolates the airlock from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock to the payload bay when open.
Airlock repressurization is controllable from the orbiter crew cabin middeck and inside the airlock. It is performed by equalizing the airlock and cabin pressure with airlock-hatch-mounted equalization valves on the inner hatch. Depressurization of the airlock is controlled from inside the airlock. The airlock is depressurized by venting the airlock pressure overboard. The two D-shaped airlock hatches are installed to open toward the primary pressure source, the orbiter crew cabin, to achieve pressure-assist sealing when closed.
Each hatch has six interconnected latches with gearbox and actuator, a window, a hinge mechanism and hold-open device, a differential pressure gauge on each side and two equalization valves.
The window in each airlock hatch is 4 inches in diameter. The window is used for crew observation from the cabin and airlock and the airlock and payload bay. The dual window panes are made of polycarbonate plastic and are mounted directly to the hatch using bolts fastened through the panes. Each hatch window has dual pressure seals with seal grooves located in the hatch.
Each airlock hatch has dual pressure seals to maintain the airlock's pressure integrity. One seal is mounted on the airlock hatch and the other on the airlock structure. A leak check quick disconnect is installed between the hatch and the airlock pressure seals to verify hatch pressure integrity before flight.
The gearbox with latch mechanisms on each hatch allows the flight crew to open or close the hatch during transfers and EVA operations. The gearbox and the latches are mounted on the low-pressure side of each hatch, and a gearbox handle is installed on both sides to permit operation from either side of the hatch.
Three of the six latches on each hatch are double-acting. They have cam surfaces that force the sealing surfaces apart when the latches are opened, thereby acting as crew assist devices. The latches are interconnected, with push-pull rods and an idler bell crank installed between the rods for pivoting the rods.
Self-aligning dual rotating bearings are used on the rods to attach the bellcranks and the latches. The gearbox and hatch's open support struts are also connected to the latching system, using the same rod and bellcrank and bearing system. To latch or unlatch the hatch, a rotation of 440 degrees on the gearbox handle is required.
The hatch actuator and gearbox are used to provide the mechanical advantage to open and close the latches. The hatch actuator lock lever requires a force of 8 to 10 pounds through an angle of 180 degrees to unlatch the actuator. A minimum rotation of 440 degrees with a maximum force of 30 pounds applied to the actuator handle is required to operate the latches to their fully unlatched positions.
The hinge mechanism for each hatch permits a minimum opening sweep into the airlock or the crew cabin middeck. The inner hatch (airlock to crew cabin) is pulled and pushed forward into the crew cabin approximately 6 inches. The hatch pivots up and to the right side. Positive locks are provided to hold the hatch in both an intermediate and a full-open position. To release the lock, a spring-loaded handle is provided on the latch hold-open bracket. Friction is also provided in the linkage to prevent the hatch from moving if released during any part of the swing.
The outer hatch (in airlock to payload bay) opens and closes to the contour of the airlock wall. The hatch is hinged to be pulled first into the airlock and then pulled forward at the bottom and rotated down until it rests with the low-pressure (outer) side facing the airlock ceiling (middeck floor). The linkage mechanism guides the hatch from the close/open, open/close position with friction restraint throughout the stroke.
The hatch has a hold-open hook that snaps into place over a flange when the hatch is fully open. The hook is released by depressing the spring-loaded hook handle and pushing the hatch toward the closed position. To support and protect the hatch against the airlock ceiling, the hatch incorporates two deployable struts. The struts are connected to the hatch linkage mechanism and are deployed when the hatch linkage mechanism is rotated open. When the hatch latches are rotated closed, the struts are retracted against the hatch.
The airlock hatches can be removed in flight from the hinge mechanism via pip pins, if required.
An air circulation system provides conditioned air to the airlock during non-EVA operation periods. The airlock revitalization system duct is attached to the outside airlock wall at launch. When the airlock hatch is opened in flight, the duct is rotated by the flight crew through the cabin and airlock hatch and installed in the airlock. It is held in place by a strap holder.
The duct has a removable air diffuser cap on the end of the flexible duct that can adjust the air flow from zero to 216 pounds per hour. The duct must be rotated out of the airlock before the cabin and airlock hatch is closed for airlock depressurization. During the EVA preparation period, the duct is rotated out of the airlock and can be used as supplemental air circulation in the middeck.
To assist the crew member in pre- and post-EVA operations, the airlock incorporates handrails and foot restraints. Handrails are located alongside the avionics and environmental control and life support system panels. Oval aluminum alloy handholds 0.75 by 1.32 inches are mounted in the airlock. They are painted yellow. The handrails are bonded to the airlock walls with an epoxyphenolic adhesive.
Each handrail has a clearance of 2.25 inches from the airlock wall to allow gripping in a pressurized glove. Foot restraints are installed on the airlock floor nearer the payload bay side. A ceiling handhold installed nearer the cabin side of the airlock was removed to make room to stow a third EMU. The foot restraints can be rotated 360 degrees by releasing a spring-loaded latch and lock every 90 degrees.
A rotation release knob on the foot restraint is designed for shirt-sleeve operation; therefore, it must be positioned before the suit is donned. The foot restraint is bolted to the floor and cannot be removed in flight. It is sized for the EMU boot. The crew member first inserts his foot under the toe bar and then rotates his heel from inboard to outboard until the heel of the boot is captured.
There are four floodlights in the airlock. The lights are controlled by switches in the airlock on panel AW18A. Lights 1, 3 and 4 are controlled by a corresponding on/off switch on panel AW18A. Light 2 can be controlled by an on/off switch on panels AW18A and M013Q, allowing illumination of the airlock prior to entry. Lights 1, 3 and 4 are powered by main buses A, B and C, respectively, and light 2 is powered by essential bus 1 BC. The circuit breakers are on panel ML86B.
The airlock provides stowage for two EMUs, two service and cooling umbilicals and miscellaneous support equipment. Both EMUs are mounted on the airlock walls by means of an airlock adapter plate.
The prime contractor to NASA for the space suit and life support system is United Technologies' Hamilton Standard Division in Windsor Locks, Conn. Hamilton Standard is program systems manager, designer and builder of the space suit and life support system. Hamilton Standard's major subcontractor is ILC Dover of Frederica, Del., which fabricates the space suit.
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"Shuttle Extravehicular Mobility Unit (Space Suit)",348,0,0,0
The EMUs provide the necessities for life support, such as oxygen, carbon dioxide removal, a pressurized enclosure, temperature control and meteoroid protection during EVA.
The EMU space suit comes in various sizes so that flight crew members can pick their suits before launch. Components are designed to fit men and women from the 5th to the 95th percentiles of body size.
The self-contained life support system contains seven hours of expendables, such as oxygen, a battery for electrical power, water for cooling, \Jlithium\j \Jhydroxide\j for carbon dioxide removal and a 30-minute emergency life support system during an EVA.
The airlock adapter plate in the airlock also provides a fixed position for the EMUs to assist the crew member during donning, doffing, checkout and servicing. The EMU weighs approximately 225 pounds, and its overall storage envelope is 26 by 28 by 40 inches. For launch and entry, the lower torso restraint, a cloth bag attached to the airlock adapter plate with straps, is used to hold the lower torso and arms securely in place.
The EMU is pressurized to 4 psid. It is designed for a 15-year life with cleaning and drying between flights. The EMU consists of a hard upper torso, lower torso assembly, gloves, helmet and visor assembly, communications carrier assembly, liquid cooling and ventilation garment, urine collection device and operational bioinstrumentation system.
The upper torso, including arms, is that portion of the pressure suit above the waist, excluding the gloves and helmet. It provides the structural mounting for most of the EMU-helmet, arms, lower torso, portable life support system, display and control module and electrical harness. The arm assembly contains the shoulder joint and upper arm bearings that permit shoulder mobility as well as the elbow joint and wrist bearing.
The PLSS is made of fiberglass and provides a mounting for other EMU components. It includes oxygen bottles; water storage tanks; a fan, separator and pump motor assembly; a sublimator; a contaminant control cartridge; various regulators, valves and sensors; communications; bioinstrumentation; and a \Jmicroprocessor\j module.
The secondary oxygen pack attaches to the bottom of the PLSS. The PLSS expendables include 1.2 pounds of oxygen pressurized to 850 psia in the primary bottles, 2.6 pounds of oxygen at 6,000 psia in the secondary pack, 10 pounds of water for cooling in three bladders and \Jlithium\j \Jhydroxide\j in the contaminant control cartridge.
The primary oxygen system and water bladders provide enough of these expendables for seven hours inside the EMU, including 15 minutes for checkout, six hours of EVA, 15 minutes for EMU doffing and 30 minutes for reserve. The SOP will supply oxygen and maintain suit pressure for 30 minutes in the event of a failure in the primary system or depletion of the primary oxygen system.
The lower torso assembly is that portion of the EMU below the waist, including boots. It consists of pants and hip, knee and ankle joints. The lower torso comes in various sizes and connects to the hard upper torso by a waist ring. It is composed of several layers, beginning with a pressure bladder of urethane-coated nylon, a restraining layer made of Dacron, an outer thermal garment made of neoprene-coated nylon, four layers of aluminized Mylar and a surface layer of Gortex and Nomex. The foot section consists of specialized socks that contain return air ports. The EVA crew members' feet are fitted with boot inserts that fit into the boots.
The gloves contain the wrist connection, wrist joint and \Jinsulation\j padding for palms and fingers. They connect to the arms and are available in 15 sizes.
The helmet is a clear polycarbonate bubble with a neck disconnect and ventilation pad that provides pressurization for the head. An assembly that goes over the helmet contains visors that are manually adjusted to shield the EVA crew members' eyes from micrometeoroids and from ultraviolet and infrared radiation from the sun. Two EVA lights are attached on each side of the helmet. A TV camera can also be attached to the helmet.
A cap, known as the Snoopy cap, is worn under the EMU helmet. It fits over the crew member's head and is held in place by a chin guard. It contains a \Jmicrophone\j and headphones for two-way communications and receiving caution and warning tones.
The liquid cooling and ventilation garment worn by the EVA crew member under the pressure suit has sewn-in tubes. It provides circulation of cooling water and pickup of vent flow at the extremities. It is a mesh one-piece suit made of spandex and has a zipper in the front for entry. It has 300 feet of plastic tubing that carries cooling water at a rate of 240 pounds per hour. It is controlled by a valve on the display and control module.
Ducting along the garment's arms and legs directs oxygen and carbon dioxide from the suit to the life support system for purification and recirculation. The garment weighs 6.5 pounds and provides cooling to maintain desired body temperature and physical activity that nominally generates 1,000 Btu per hour and can generate up to 2,000 Btu per hour, which is considered extremely vigorous.
The urine collection device collects urine. It can store approximately 1 quart of urine. It consists of adapter tubing, a storage bag and disconnect hardware for emptying after an EVA to the orbiter waste water tank.
The bioinstrumentation system monitors the EVA crew member's heart rate (electrocardiogram) during an EVA.
An in-suit drink bag stores approximately 0.5 of a quart of drinking water in the upper torso. A tube from the upper hard torso to the helmet permits the EVA crew member to drink water while suited.
The life support system consists of the portable life support system, display and control module, contaminant control cartridge, battery, secondary oxygen pack, and EVA communicator and EMU antenna. The PLSS is also referred to as a backpack. The PLSS normally provides the EVA crew member with oxygen for breathing, ventilation and pressurization and water for cooling.
The contaminant cartridge consists of \Jlithium\j \Jhydroxide\j, charcoal and filters to remove carbon dioxide, odors, particulates and other contaminants from the ventilation circuit. It is replaceable upon completion of an EVA.
A silver-zinc battery provides all electrical power used by the EMU and life support system. It is stored dry, filled, sealed and charged before flight. It is rechargeable upon completion of an EVA and is rated at 17 volts dc.
The SOP provides oxygen for breathing, ventilation, pressurization and cooling in the event of a PLSS malfunction. It is mounted at the base of the PLSS and contains a 30-minute oxygen supply, a valve and a regulator assembly.
The EVA communicator and EMU antenna provide EVA communications via its transceiver and antenna between the suited crew member and the orbiter. In addition, the crew member's electrocardiogram is telemetered through the communicator to the orbiter. It is a separate subassembly that attaches to the upper portion of the life support system at the back of the hard upper torso. The controls are located on the display and control module mounted at the front of the upper torso.
The radios for space walk communications have two single-UHF-channel transmitters, three single-channel receivers and a switching mechanism. In addition, telemetry equipment is included so that ground personnel can monitor the \Jastronaut\j's heart beat. These backpack radios have a low-profile antenna, a 1-foot-long rectangular block fitted to the top of the packs. The radios weigh 8.7 pounds and are 12 inches long, 4.3 inches high and 3.5 inches wide.
The EMU electrical harness provides biomedical instrumentation and communications connections to the PLSS. The harness connects the communications carrier assembly and the biomedical instrumentation subsystem to the hard upper torso, where internal connections are routed to the EVA communicator. The cable routes signals from the electrocardiogram sensors, which are attached to the crew member, through the bioinstrumentation system to the EVA communicator. It also routes caution and warning signals and communications from the communicator to the crew member's headset.
The DCM is an integrated assembly that attaches directly to the front of the hard upper torso. The module contains a series of mechanical and electrical controls, a \Jmicroprocessor\j, and an alphanumeric LED display easily seen by a crew member wearing the space suit. It contains the displays and controls associated with the operation of the EMU.
The function of the display and control module is to enable the crew member to control the PLSS and the secondary oxygen pack. It also indicates the status of the PLSS, the suit and, the manned maneuvering unit (when it is attached) visibly and audibly.
The mechanical controls consist of a suit purge valve, the liquid cooling and ventilation garment cooling valve, and the oxygen actuator control, which has four positions: off , iv (which turns primary oxygen on to a 0.5-psid suit pressure setting), press (which turns primary oxygen on to a 4.1-psid suit pressure setting), and ev (which leaves primary oxygen on the 4.1-psid setting and turns the secondary oxygen pack on).
The electrical controls include a voice communications mode switch, dual volume controls, push-to-talk switches, a power mode switch, feedwater and C/W switches and the LED display brightness control. The displays on the module are a 12-digit LED display, a built-in test equipment indicator and an analog suit pressure gauge.
The display and control module is connected to the hard upper torso and to the PLSS by both internal and external hookups. A multiple-function connector links the display module to the service and cooling umbilical, thus enabling the use of the display module controls during suit checkout inside the airlock station.
The display module interacts with a \Jmicroprocessor\j in the PLSS that contains a program that enables the crew member to cycle the display through a series of systems checks and thereby determine the condition of a variety of components. The \Jmicroprocessor\j monitors oxygen pressure and calculates the time remaining at the crew member's present use rate. It signals an alarm at high oxygen use in the primary oxygen tanks. It also monitors water pressure and temperature in the cooling garment.
The carbon dioxide level is monitored and an alarm is signaled when it reaches high concentrations in the suit. The \Jmicroprocessor\j monitors the power consumed and signals at high current-drain rates and also when an estimated 30 minutes of battery power is left. All the warnings are displayed on the LED display.
The display module also has a fiber-optic cable that is used when the MMU is connected to the EMU. The fiber-optic cable connects the display unit to the MMU. A fiber-optic cable is more reliable and more convenient and safer to use than an electrical connector for extravehicular applications.
The MMU is mounted on the back of the portable life support system. When the MMU is connected, the display module also provides a cycled readout of propellant pressures, temperatures, and battery condition (in the MMU) and an audible thruster cue. The C/W system warns of low propellant, low battery, and failed components.
Oxygen from the system enters the suit at the helmet and flows from behind the head down through the suit. Oxygen and carbon dioxide are removed from the suit through the liquid cooling and ventilation garment at ports near the crew member's wrists and feet.
Return air goes first through the contaminant control cartridge, where activated charcoal and \Jlithium\j \Jhydroxide\j beds remove carbon dioxide, odors and dust. From there the return air goes through a water separator, where moisture from exhalation and the \Jlithium\j \Jhydroxide\j and carbon dioxide reaction is removed. The oxygen then goes through the fan, which maintains air flow at 6 cubic feet per minute.
It is then routed through the sublimator, where it is cooled to 85 F, and then passes through a vent and flow detector and back to the suit. Oxygen for the air system is fed from the primary oxygen containers through regulators that maintain suit pressure at 4.1 psid.
The system is protected from suit overpressure, primary oxygen supply depletion or mechanical failure by regulators, sensors and the secondary oxygen pack. The secondary oxygen pack can maintain suit pressure at 3.45 psid. A purge valve on the display and control module allows a crew member to completely replace system oxygen in the suit if, for instance, the carbon dioxide level rises too high too quickly.
The cooling water system takes the warm water from the cooling garment and divides it into two loops. One loop goes to the sublimator, where the water in that loop is cooled and sent back to the cooling control valve. The other loop goes directly back to the cooling control valve, where the loops are recombined and full flow goes back to the cooling garment. Thus, the cooling garment has a constant flow of cooling water at a temperature set by the crew member using the cooling control valve.
During the process, the full flow from the cooling garment goes through a gas separator, where gas is removed from the loop, and then through a pump that maintains a flow of 260 pounds per hour. Another side loop circulates 20 pounds per hour through the contaminant control cartridge to cool the \Jlithium\j \Jhydroxide\j canister since the \Jlithium\j \Jhydroxide\j and carbon dioxide reaction produces heat and needs to be kept cool for an efficient reaction.
Since the system is a closed-loop design, water from the water separator is fed back to the water system, and air from the gas trap is fed back to the oxygen system. Water from the water tanks is also fed, through regulators, into the cooling system.
However, the primary purpose of the water tanks is to feed water to the sublimator. The sublimator works on the principle of sublimation, that is, the process by which a solid turns directly into a vapor, bypassing the liquid phase. In this case, ice is formed on the sublimator evaporator sieve and is allowed to vaporize to space, removing heat with it. Air and cooling water are passed through fins in the sublimator, which extracts heat from each system.
The PLSS sensors detect system air flow, air pressure, water flow, water pressure, differential water pressure (between the circulating system and the water tanks), water temperature and carbon dioxide content in the return air. In addition, there are a number of crew-selectable valves, including a purge valve, a cooling control valve (infinitely variable), oxygen supply and a direct-reading air pressure gauge. The sensors supply information to the display and control module, where a \Jmicroprocessor\j maintains an automatic watch over system integrity.
Normally, the day before an EVA, the orbiter crew compartment cabin pressure is allowed to decrease from 14.7 psia to 12.5 psia through metabolic usage. One hour before depressurizing the crew compartment from 12.5 psia to 10.2 psia, the EVA crew member and prebreathes 100-percent oxygen for 45 minutes. There are two PEAPs in the airlock.
The crew compartment is then depressurized from 12.5 psia to 10.2 psia and remains at this pressure until after the EVA is completed. This is necessary to remove \Jnitrogen\j from the EVA crew member's blood before the EVA crew member works in the pure oxygen environment of the EMU. Without the prebreathing, bends can occur.
When an individual fails to reduce \Jnitrogen\j levels in the blood before working in a pressure condition, it can result in \Jnitrogen\j coming out of solution in the form of bubbles in the bloodstream. This condition results in pain in the body joints, possibly because of restricted blood flow to connective tissues or because of the extra pressure caused by bubbles in the blood in the joint area.
In preparation for an EVA, the crew member dons the liquid cooling and ventilation garment first, enters the airlock and dons the lower torso assembly. The crew member then squats under the hard upper torso mounted on the airlock adapter plate and slides up into the upper torso. The upper and lower torsos are connected with a waist ring. The gloves and helmet are then put on, and the EMU is disconnected from the AAP.
The orbiter provides electrical power, oxygen, liquid cooling and ventilation garment cooling and water to the EMUs in the airlock via the service and cooling umbilical for EVA preparation and after EVA operations.
The service and cooling umbilical contains communication lines, electrical power, water, water drain line and oxygen recharge lines. The umbilical permits the EVA crew member to check out the suit in the airlock without using the EMU supply of water, oxygen and battery power.
The SCU is launched with the orbiter end fittings permanently connected to the appropriate ECLSS panels in the airlock and the EMU connected to the airlock adapter plate stowage connector. It allows all supplies (oxygen, water, electrical and communication) to be transported from the airlock control panels to the EMU before and after EVA without using the EMU expendable supplies of water, oxygen and battery power that are scheduled for use in the EVA.
The SCU also provides EMU recharge. The SCU umbilical is disconnected just before the crew member leaves the airlock on an EVA and is reconnected when he returns to the airlock. Each SCU is 144 inches long, 3.5 inches in diameter and weighs 20 pounds. Actual usable length after attachment to the control panel is approximately 7 feet.
\B\IFor more information click on \b\i\JShuttle Extravehicular Mobility Unit (continued)\j
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"Shuttle Extravehicular Mobility Unit (continued)",349,0,0,0
The airlock has two display and control panels. The airlock control panels are basically split to provide either ECLSS or avionics operations. The ECLSS panel provides the interface for the SCU waste and potable water, liquid cooling and ventilation garment cooling water, EMU hardline communications, EMU power and oxygen supply. The avionics panel includes the airlock lighting, airlock audio system and EMU power and battery recharge controls.
Airlock communications are provided with the orbiter audio system at airlock panel AW82D, where connectors for the headset interface units and the EMUs are located at airlock panel AW18D, the airlock audio terminal. The HIUs are inserted in the crew member communications carrier unit connectors on airlock panel AW82D.
During EVA, the EVA communicator is part of the same UHF system that is used for air-to-air and air-to-ground voice communications between the orbiter and landing site control tower. The EVA communicator provides full duplex (simultaneous transmission and reception) communications between the orbiter and the EVA crew members. It also supplies continuous data reception of electrocardiogram signals from each crew member by the orbiter and processing by the orbiter and relay of electrocardiogram signals to the ground. The UHF airlock antenna in the forward portion of the payload bay provides the UHF EVA capability.
Cooling for flight crew members before and after the EVA is provided by the liquid-cooled garment circulation system via the SCU and LCG supply and return connections on panel AW82B. These connections are routed to the orbiter LCG heat exchanger, which transfers the collected heat to the orbiter Freon-21 coolant loops.
With the suit connected to the SCU, oxygen at 900 psia, plus or minus 500 psia, is supplied through airlock panel AW82B from the orbiter's oxygen system when the oxygen valve is in the open position on the airlock panel. This provides the suited crew member with breathing oxygen and prevents depletion of the PLSS oxygen tanks before the EVA. Before the crew member seals the helmet, an oxygen purge adapter hose is connected to the airlock panel to flush \Jnitrogen\j out of the suit.
When the SCU is disconnected, the PLSS provides oxygen for the suit. When the EVA is completed and the SCU is reconnected, the orbiter's oxygen supply begins recharging the PLSS, assuming that the oxygen valve on panel AW82B is open. Full oxygen recharge takes approximately one hour (allowing for thermal expansion during recharge), and the tank pressure is monitored on the EMU display and control panel as well as on the airlock oxygen pressure readout.
The EMU water supply and waste valves are opened during the EVA preparation by switches on panel AW82D. This provides the EMU, via the SCU, access to the orbiter's potable water and waste water systems. The support provided to the EMU PLSS is further controlled by the EMU display and control panel.
Potable water (supplied from the orbiter at 16 psi, plus or minus 0.5 psi; 100 to 300 pounds per hour; and 40 to 100 F) is allowed to flow to the feedwater reservoir in the EMU that provides pressure, which would top off any tank not completely filled. Waste water condensate developed in the PLSS is allowed to flow to the orbiter waste water system via the SCU whenever the regulator connected at the bacteria filters (airlock end of the SCU) detects upstream pressure in excess of 16 psi, plus or minus 0.5 psi.
When the SCU is disconnected from the EMU, the PLSS assumes its functions. When the SCU is reconnected to the EMU upon completion of the EVA, it performs the same functions it did before the EVA except that the water supply is allowed to continue until the PLSS water tanks are filled, which takes approximately 30 minutes.
In preparation for the EVA, the airlock hatch to the orbiter crew cabin is closed and depressurization of the airlock begins.
Airlock depressurization is accomplished in two stages by a three-position valve located on the ECLSS panel AW82A in the airlock. The airlock depressurization valve is covered with a pressure and dust cap. Before the cap is removed from the valve, it is necessary to vent the area between the cap and valve by pushing the vent valve on the cap.
In flight the pressure and dust cap is stored next to the valve. The airlock depressurization valve is connected to a 2-inch-inside-diameter stainless steel overboard vacuum line. The airlock depressurization valve controls the rate of depressurization by varying the valve's diameter. Closing the valve prevents any air flow from escaping to the overboard vent system.
When the crew members have completed the 40-minute prebreathe in the EMUs, the airlock is depressurized from 10.2 psia to 5 psia by moving the airlock depressurization valve to the 5 position, which opens the depressurization valve and allows the pressure in the airlock to decrease at a controlled rate.
The airlock depressurization valve must be closed to maintain 5 psia. During depressurization, pressure can be monitored by the delta pressure gauge on either airlock hatch. A delta pressure gauge is installed on each side of both airlock hatches.
At this time, the flight crew performs an EMU suit leak check, electrical power is transferred from the umbilicals to the EMU batteries, the umbilicals are disconnected, and the suit oxygen packs are brought on-line.
The second stage of airlock depressurization is accomplished by positioning the airlock depressurization valve to 0 , which increases the valve's diameter and allows the pressure in the airlock to decrease from 5 psia to zero psia. The suit sublimators are activated for cooling, EMU system checks are performed, and the airlock and payload bay hatch can be opened. The hatch is capable of opening against a 0.2 psia differential maximum.
Hardware provisions are installed in the orbiter payload bay for use by the crew member during the EVA.
Handrails and tether points are located on the payload bulkheads, forward bulkhead station Xo 576 and aft bulkhead station Xo 1307 along the sill longeron on both sides of the bay to provide translation and stabilization capability for EVA crew members and facilitate movement in the payload bay. The handrails are designed to withstand a load of 200 pounds, or 280 pounds maximum, in any direction. Tether attach points are designed to sustain a load of 574 pounds, 804 pounds maximum, in any direction.
The handrails have a cross section of 1.32 inches by 0.75 of an inch. They are made of aluminum alloy tubing and are painted yellow. The end braces and side struts of the handrails are constructed of \Jtitanium\j. An aluminum alloy end support standoff functions as the terminal of the handrail. Each end support standoff incorporates a 1-inch- diameter tether point.
A 25-foot safety tether is attached to each crew member at all times during an EVA.
The tether consists of a reel case with an integral D-ring, a reel with a light takeup spring, a cable and a locking hook. The safety tether hook is locked onto the slidewire before launch, and the cable is routed and clipped along the left and right handrails to a position just above the airlock and payload bay hatch. After opening the airlock hatch but before leaving the airlock, the crew member attaches a waist tether to the D-ring of the safety tether to be used.
The other end of the waist tether is hooked to a ring on the EMU waist bearing. The crew member may select either the left or the right safety tether. With the selector on the tether in the locked position, the cable will not retract or reel out. Moving the selector to the unlocked position allows the cable to reel out and the retract feature to take up slack. The cable is designed for a maximum load of 878 pounds. The routing of the tethers follows the handrails, which allows the crew member to deploy and restow his tether during translation.
The two slidewires, approximately 46.3 feet long, are located in the longeron sill area on each side of the payload bay. They start approximately 9.3 feet aft of the forward bulkhead and extend approximately 46.3 feet down the payload bay. The slidewires withstand a tether load of 574 pounds with a safety factor of 1.4 or 804 pounds maximum.
EVA support equipment may consist of a small work station, tool caddies and equipment tethers. The work station contains a universal attachment tether for crew member restraint and a carrying location for the tool caddies. The caddies hold the tools and provide tethers for them when they are not in use.
A cargo bay stowage assembly installed in the orbiter payload bay contains miscellaneous tools for use in the payload bay during an EVA. The CBSA is approximately 42 inches wide, 24 inches deep and 36 inches high. The CBSA weighs 573 pounds.
The airlock and cabin hatch has two pressure equalization valves that can be operated from both sides of the hatch to repressurize the airlock volume. Each valve has three positions- closed , norm (normal) and emerg (emergency)-and is protected by a debris pressure cap on the intake (high-pressure) side of the valve. The pressure cap on the outer hatch must be vented for removal.
The caps are tethered to the valves and also have small Velcro spots that allow them to be stored temporarily on the hatch. The exit side of the valve contains an air diffuser to provide uniform flow out of the valve.
Through the use of the equalization valves, the airlock is initially pressurized to 5 psia, and the space suit is connected to the umbilical in the airlock and electrical power is transferred back to umbilical power. After the airlock is pressurized to the 10.2-psia cabin pressure, the EVA crew members remove and recharge their EMUs. Shortly thereafter, the crew compartment cabin is pressurized from 10.2 psia to 14.7 psia.
The orbiter can accommodate three six-hour EVAs by two crew members per flight at no weight or volume cost to the payload. Two of the EVAs are for payload support; the third is reserved for orbiter contingency. Additional EVAs can be considered with consumables charged to payloads.
When fitted with a tunnel adapter, hatches, tunnel extension and tunnel, the middeck airlock permits flight crew members to transfer from the orbiter middeck into the Spacelab pressurized modules, where they can work in a pressurized shirt-sleeve environment. The airlock, tunnel adapter and hatches also permit EVA flight crew members to transfer into the payload bay from the tunnel adapter in the space suit assembly without depressurizing the \Jspacecraft\j crew cabin and Spacelab.
The tunnel adapter is located in the payload bay and is attached to the airlock at orbiter station X o 576 and to the tunnel extension at Xo 660, thus attaching it to the Spacelab tunnel and Spacelab. The tunnel adapter has an inside diameter of 63 inches at the widest section and tapers in the cone area at each end to two 40-inch-diameter D-shaped openings, 36 inches across.
An identical D-shaped opening is located at the top of the tunnel adapter. Two pressure-sealing hatches are located in the tunnel adapter: one at the upper area of the tunnel adapter and one at the aft end of the tunnel adapter. The tunnel adapter is constructed of 2219 aluminum and is a welded structure with 2.4- by 2.4-inch exposed structural ribs on the exterior surface and external waffle skin stiffening.
The hatch located in the aft end isolates the tunnel adapter and airlock from the extension tunnel and Spacelab. This hatch opens into the tunnel adapter. The hatch located in the tunnel adapter at the upper D-shaped opening isolates the airlock and tunnel adapter from the unpressurized payload bay when closed and permits EVA crew members to exit the airlock and tunnel adapter to the payload bay when open. This hatch opens into the tunnel adapter.
The two hatches in the tunnel adapter are installed to open toward the primary pressure source and the orbiter crew cabin to achieve pressure-assist sealing when closed.
Each hatch has six interconnected latches (with the exception of the aft hatch, which has 17) with a gearbox and actuator, window, hinge mechanism and hold-open device, differential pressure gauge on each side and two equalization valves.
The window in each hatch is 4 inches in diameter. The window is used for crew observation from the cabin and airlock, tunnel adapter to tunnel, and tunnel adapter to payload bay. The dual window panes are made of polycarbonate plastic and are mounted directly to the hatch using bolts fastened through the panes. Each hatch window has dual pressure seals with seal grooves located in the hatch.
Each hatch has dual pressure seals to maintain pressure integrity. One seal is mounted on the hatch and the other on the structure. Leak check quick disconnects are installed between the hatches and the pressure seals to verify the hatches' pressure integrity before flight.
The gearbox with latch mechanisms on each hatch allows the flight crew to open or close the hatch during transfers and EVA operation. The gearbox and the latches are mounted on the low-pressure side of each hatch and a gearbox handle is installed on both sides to permit operation from either side of the hatch.
The aft hatch is hinged to be first pulled into the tunnel adapter and then pulled forward at the bottom. The top of the hatch is rotated toward the tunnel and downward until the hatch rests with the Spacelab side facing the tunnel adapter's floor. The linkage mechanism guides the hatch from the close/open, open/close position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro.
The upper (EVA) hatch in the tunnel adapter opens and closes to the left wall of the tunnel adapter. The hatch is hinged to be pulled first into the tunnel adapter and then pulled forward at the hinge area and rotated down until it rests against the left wall of the tunnel adapter. The linkage mechanism guides the hatch from the close/open, open/close position with friction restraint throughout the stroke. The hatch is held in the open position by straps and Velcro. The hatches can be removed in flight from the hinge mechanism via pip pins, if required.
When the airlock hatch is opened on orbit, a duct is connected to the cabin air system to provide conditioned air to the airlock, tunnel adapter and tunnel during non-EVA-operation periods. The duct must be disconnected before the airlock hatch is closed for entry.
For an EVA during a mission with pressurized Spacelab modules, all hatches are closed and depressurization of the airlock tunnel adapter begins. The prebreathe requirements, lowering of cabin pressure to 10.2 psia and space suit preparations, etc., remain the same as for an EVA from the airlock. The difference is that the EVA crew member enters the payload bay from the upper tunnel adapter hatch.
Hardware is installed in the orbiter payload bay and in the tunnel adapter, tunnel and Spacelab for use by the crew member during the EVA.
When EVA is completed, the crew member enters the upper tunnel adapter hatch and closes it. The airlock and tunnel adapter are repressurized in the same manner as when the airlock alone is used.
The crew altitude protection system is worn by the flight crew members during launch and entry. Oxygen from the orbiter's environmental control and life support system is supplied to the crew members through the CAPS. A skullcap is worn under the helmet.
A microphone and headphones are worn over the skullcap for two-way communications and receiving caution and warning tones. The CAPS also incorporates an anti-exposure, anti-gravity suit to prevent pooling of body fluids and to aid in maintaining the circulation of blood during entry. Pooling of body fluids can occur when high ''g'' loads are imposed on the body, particularly after crew members have had more than 12 hours of zero-gravity activity. The anti-gravity portion of the CAPS is pressurized from the ECLSS oxygen system.
A radioisotope thermoelectric generator cooling and gaseous \Jnitrogen\j purge system is installed in Discovery OV-103) and Atlantis OV-104) to support those payloads with RTGs or gaseous \Jnitrogen\j purging requirements. The water coolant system lines are located in the aft fuselage and payload bay of the orbiters. The water system flows water through the existing ground support equipment heat exchanger in the aft fuselage. The heat exchanger cools the water from the RTGs and directs it back through the RTGs. An accumulator located in the aft fuselage next to the GSE heat exchanger acts as a reservoir to absorb excess water and is pressurized with gaseous \Jnitrogen\j. The water preflight servicing umbilical is located at X o 835.05.
The gaseous \Jnitrogen\j purge system connects at the existing nitrogen/oxygen panel at the X o 576 bulkhead. The gaseous \Jnitrogen\j is directed from the nitrogen/oxygen panel to the existing payload interface panel in the payload bay and provides gaseous \Jnitrogen\j purging as required by the payload.
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"Shuttle Landing Gear System",352,0,0,0
The landing gear system on the orbiter is a conventional \Jaircraft\j tricycle configuration consisting of a nose landing gear and a left and right main landing gear. Each landing gear includes a shock strut with two wheel and tire assemblies. Each main landing gear wheel is equipped with a brake assembly with anti-skid protection. The nose landing gear is steerable. The nose landing gear is located in the lower forward fuselage, and the main landing gear are located in the lower left and right wing area adjacent to the midfuselage.
The nose landing gear is retracted forward and up into the lower forward fuselage and is enclosed by two doors. The main landing gear are also retracted forward and up into the left and right lower wing area, and each is enclosed with a single door. The nose and main landing gear can be retracted only during ground operations.
For retraction, each gear is hydraulically rotated forward and up during ground operations until it engages an uplock hook for each gear in its respective wheel well. The uplock hook locks onto a roller on each strut. Mechanical linkage driven by each landing gear mechanically closes the respective landing gear doors. All three landing gear doors have high-temperature reusable surface \Jinsulation\j thermal protection system tiles bonded to their outer surface with thermal barriers to protect and prevent the landing gear and wheel well from the high-temperature thermal loads encountered during the shuttle's entry into the atmosphere.
For deployment of the landing gear, the uplock hook for each gear is activated by the flight crew initiating a gear-down command. The uplock hook is hydraulically unlocked by hydraulic system 1 pressure applied to release it from the roller on the strut to allow the gear, assisted by springs and hydraulic actuators, to rotate down and aft.
Mechanical linkage released by each gear actuates the respective doors to the open position. The landing gear reach the full-down and extended position within 10 seconds and are locked in the down position by spring-loaded downlock bungees. If hydraulic system 1 pressure is not available to release the uplock hook, a pyrotechnic initiator at each landing gear uplock hook automatically releases the uplock hook on each gear one second after the flight crew has commanded gear down.
The landing gears are deployed only after the \Jspacecraft\j has an indicated airspeed of less than 300 knots (345 mph) and an altitude of approximately 250 feet.
The shock strut of each landing gear is the primary source of shock attenuation at landing. The struts have air/oil shock absorbers to control the rate of compression extension and prevent damage to the vehicle by controlling load application rates and peak values.
Each main landing gear wheel contains an electrohydraulic disc brake assembly with anti-skid control. The main landing gear brakes are controlled by the commander or pilot applying toe pressure to the rudder pedals; electrical signals produced by rudder pedal toe pressure control hydraulic servovalves at each wheel and allow hydraulic system pressure to perform braking.
Main landing gear brakes cannot be applied until weight on the main gear has been sensed. The anti-skid system monitors wheel velocity and controls brake \Jtorque\j to prevent wheel lock and tire skidding. The braking/anti-skid system is redundant in that it utilizes system 1 and 2 hydraulic pressure as the active system with system 3 as standby and also utilizes all three main dc electrical systems.
The nose landing gear contains a hydraulic steering actuator that is electrohydraulically steerable through the use of the onboard general-purpose computers, the commander's or pilot's rudder pedals in conjunction with the orbiter flight control system in the control stick steering mode, or through the use of the commander's or pilot's rudder pedals in the direct mode.
If hydraulic system 1 is inoperative, nose wheel steering changes to caster mode, and the commander or pilot would then apply toe pressure to the brake pedals to apply hydraulic pressure to the left and right main gear brakes as required for directional control using differential braking.
Each landing gear shock strut assembly is constructed of high-strength, stress- and corrosion-resistant steel alloys, aluminum alloys, stainless steel and aluminum bronze. Cadmium and \Jchromium\j plating and urethane paint are applied to the strut surfaces for space flight protection. The shock strut is a pneudraulic shock absorber containing gaseous \Jnitrogen\j and hydraulic fluid. Because the shock strut is subjected to zero-g conditions during space flight, a floating piston separates the gaseous \Jnitrogen\j from the hydraulic fluid to maintain absorption integrity.
Landing gear wheels are made in two halves from forged aluminum and are primed and painted with two coats of urethane paint.
The LG hyd isol vlv 1, 2 and 3 switches on panel R4 control the corresponding landing gear isolation valve in hydraulic systems 1, 2 and 3. When the LG hyd isol vlv 1 switch on panel R4 is positioned to close , hydraulic system 1 is isolated from the nose and main landing gear deployment uplock hook actuators and strut actuators, nose wheel steering actuator and main landing gear brake control valves. A talkback indicator next to the switch would indicate cl.
The landing gear isolation valves will not close or open unless the pressure in that system is at least 100 psi. When the LG hyd isol vlv 1 switch is positioned to open , it allows hydraulic system 1 source pressure to the main landing brake control valves and to the normally closed extend valve.
The normally closed extend valve is not energized until a gear-down command is initiated by the commander or pilot on panel F6 or panel F8. The talkback indicator would indicate op . The LG hyd isol vlv 1 switch is left in the close position during the mission to prevent inadvertent gear deployment.
The LG hyd isol vlv 2 and 3 switches on panel R4 positioned to close isolate the corresponding hydraulic system from only the main landing gear brake control valves. The adjacent talkback indicator would indicate cl . When switches 2 and 3 are positioned to open, the corresponding hydraulic system source pressure is available to the main landing gear brake control valves. The corresponding talkback indicator would indicate op .
Thus, only hydraulic system 1 is used to deploy the nose and main landing gear and for nose wheel steering. When the nose- and main-landing-gear-down command is initiated by the commander or pilot on panel F6 or F8, hydraulic system 1 pressure is directed to the nose and main landing gear uplock hook actuators and strut actuators (provided that the LG hyd isol vlv 1 switch is in the open position) to actuate the mechanical uplock hook for each landing gear and allow the gear to be deployed and also provide hydraulic system 1 pressure to the nose wheel steering actuator.
The main landing gear brake control valves receive hydraulic system 1 source pressure when the LG hyd isol vlv 1 switch is positioned to open. If hydraulic system 1 is unavailable, a pyrotechnic actuator attached to the nose and main landing gear uplock actuator would deploy the landing gear automatically one second after the gear-down command, actuate the mechanical uplock hook for each landing gear and allow the gear to be deployed. Because powered nose wheel steering would not be functional, directional control for steering would be accomplished by differential braking to caster the nose wheel.
The GPC position of the LG hyd isol vlv 1, 2 and 3 switches on panel R2 permits the onboard computer to automatically control the valves in conjunction with computer control of the corresponding hydraulic system circulation pump. The LG hyd isol vlv 2 and 3 switches provide fluid circulation to only the main landing gear brake system, which dead-ends at the brake control valves. The LG hyd isol vlv 1 switch is left closed to prevent inadvertent gear deployment.
The normally open hydraulic system 1 redundant shutoff valve is a backup to the retract/circulation valve to prevent hydraulic pressure from being directed to the retract side of the nose and main landing gear uplock hook actuators and strut actuators if the retract/circulation valve fails to open during nose and main landing gear deployment.
The normally closed hydraulic system 1 dump valve is energized open to allow hydraulic system 1 fluid to return from the nose and main landing gear areas when deployment of the landing gear is commanded by the flight crew.
The activation/deactivation limits of the hydraulic fluid circulation systems can be changed during the mission by the flight crew or the Mission Control Center-Houston. The program also includes a timer to limit the maximum time a circulation pump will run and a priority system that automatically monitors hydraulic bootstrap pressure to allow all three circulation pumps to be on at the same time. The software timers allow this software to be used in contingency situations for ''time-controlled'' circulation pump operations in order to periodically boost an accumulator that is losing hydraulic fluid through a leaking priority valve or unloader valve.
During entry, if required, LG hyd isol vlv 1, 2 and 3 are positioned to GPC . At 19,000 feet per second, the landing gear isolation valve automatic opening sequence begins under GN&C software control. If the landing gear isolation valve is not opened automatically, the flight crew will be requested by the Mission Control Center to open the valve by positioning the applicable LG hyd isol vlv to open.
Landing gear isolation valve 2 is automatically opened six minutes and 37 seconds later, and this is followed by the automatic opening of landing gear isolation valve 1 when orbiter velocity is at 800 feet per second or less. Landing gear isolation valve 3 is automatically opened at ground speed enable. Landing gear isolation valve 1 is next to last to ensure that an inadvertent gear deployment would occur as late (low airspeed) as possible.
Note that the hydraulic system 1 retract/circulation valve would be automatically closed when the landing gear system is armed for deployment.
The commander and pilot have a landing gear deployment arm and dn (down) guarded push button switch/light indicators and landing left, nose and right indicators. The commander's controls and indicators are on panel F6, and the pilot's controls and indicators are on panel F8. The dn push button, when depressed, energizes the hydraulic system 1 normally closed extend valve, permitting hydraulic system 1 source pressure for gear deployment and nose wheel steering.
The proximity switches on the nose and main landing gear doors and struts provide electrical signals to control the landing gear nose, left and right indicators on panels F6 and F8. The output signals of the landing gear and door uplock switches drive the landing gear up position indicators and the backup pyrotechnic release system. The output signals of the landing gear downlock switches drive the landing gear dn position indicators. The landing gear indicators are barberpole when the gear is deploying (or retracting).
The left and right main landing gear weight-on-wheels switches produce output signals to the guidance, navigation and control software to reconfigure the flight control system for landing.
The two weight-on-nose-gear signals run to the main landing gear brake/skid control boxes to prevent the main landing gear brakes from being applied until the nose gear is in contact with the runway and also to the GN&C software, which computes a nose wheel steering enable signal. This enable signal is then sent to the NWS control box to prevent NWS until the nose gear is in contact with the runway.
The six group 1 switches are signal conditioned by the landing gear proximity sensor \Jelectronics\j box 1, located in avionics bay 1. The six group 2 switches are signal conditioned by the landing gear proximity sensor \Jelectronics\j box 2, located in avionics bay 2.
Landing gear deployment is initiated when the commander or pilot depresses the guarded arm push button switch/light indicator and then the guarded dn push button switch/light indicator at least 15 seconds before predicted touchdown and at a speed no greater than 300 knots (345 mph).
Depressing the arm push button switch/light indicator energizes latching relays that close the hydraulic system 1 landing gear retract/circulation valve and the normally open redundant shutoff valve to the retract/circulation valve. It also arms the nose and main landing gear pyrotechnic initiator controllers and illuminates the yellow light in the arm push button switch/light indicator.
The dn push button switch/light indicator is then depressed. This energizes latching relays that open the hydraulic system 1 landing gear normally closed extend control valve, permitting the fluid in hydraulic system 1 to flow to the landing gear uplock and strut actuators and nose wheel steering. The relays also open the normally closed dump valve, allowing the landing gear retract line fluid to flow in to the hydraulic system 1 return line. The green light in the dn push button switch/light indicator is illuminated.
Hydraulic system 1 source pressure is routed to the nose and main landing gear uplock actuators, which releases the nose and main landing gear and door uplock hooks. As the uplock hooks are released, the gear begins its deployment and mechanical linkage attached to the doors and fuselage is powered by landing gear strut camming action, during gear extension, which opens the landing gear doors.
There are two landing gear doors for the nose gear and one for each main gear. The landing gear free falls into the extended position, assisted by the strut actuators and airstream in the deployment. The hydraulic strut actuator incorporates a hydraulic fluid flow through orifice (snubber) to control the rate of landing gear extension and thereby prevent damage to the gear's downlock linkages.
If hydraulic system 1 fails to release the landing gear within one second after the dn push button is depressed, the nose and left and right main landing gear uplock sensors (proximity switches) will provide inputs to the pyro initiator controllers for initiation of the redundant NASA standard detonators (nose, left and right main landing gear pyrotechnic backup release system).
They release the same uplock hooks as the hydraulic system. The nose landing gear, in addition, has a PIC and redundant NSDs that initiate a pyrotechnic power thruster two seconds after the dn push button is depressed to assist gear deployment.
The landing gear drag brace overcenter lock and spring-loaded bungee lock the nose and main landing gear in the down position.
The ldg gr/arm/dn reset switch positioned to reset on panel A12 unlatches the relays that were latched during landing gear deployment by the landing gear arm and dn push button light/switch indicators. This is primarily a ground function, which will be performed only during landing gear deactivation.
The reset position also will extinguish the yellow light in the arm push button switch/light indicator and the green light in the dn push button switch/light indicator. In addition, the hydraulic system 1 landing gear dump valve is closed, the extend control valve is closed, the retract/circulation valve is opened only if the switch is in the open position, and its redundant shutoff valve is opened (de-energized) and de-energizes the landing gear PIC circuits.
The nose landing gear tires are 32 by 8.8 inches and will withstand a burst pressure of not less than 3.2 times the normal \Jinflation\j pressure of 300 psi. The \Jinflation\j agent is gaseous \Jnitrogen\j. The maximum allowable load per nose landing gear tire is approximately 45,000 pounds and rated at 224 knots (258 mph) landing speed.
The nose landing gear shock strut has a 22-inch stroke. The maximum allowable derotation rate is approximately 9.4 degrees per second or 11 feet per second, vertical sink rate.
The main landing gear tires are 44.5 by 16 and 21 inches. The normal \Jinflation\j pressure is 315 psi, and the \Jinflation\j agent is gaseous \Jnitrogen\j. The maximum allowable load per main landing gear tire is 123,000 pounds. If the orbiter touches down with a 60/40 percent load distribution on a strut's two tires, with one tire supporting the maximum load, then the other tire can support a load of only 82,410 pounds. Therefore, the maximum tire load on a strut is 205,410 pounds with a 60/40 percent tire load distribution. The tires are rated at 225 knots (258 mph).
The main landing gear shock strut stroke is 16 inches. The allowable main gear sink rate for a 212,000-pound orbiter is 9.6 feet per second; for a 240,000-pound orbiter, it is 6 feet per second. With a 20-knot (23-mph) crosswind, the maximum allowable gear sink rate for a 212,000-pound orbiter is 6 feet per second; for a 240,000-pound orbiter, it is approximately 5 feet per second.
The landing gear tires have a life of one landing.
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"Shuttle Main Landing Gear Brakes",353,0,0,0
Each of the orbiter's four main landing gear wheels has electrohydraulic disc brakes and an anti-skid system.
Each main landing gear wheel has a disc brake assembly consisting of nine discs, four rotors, three stators, a backplate and a pressure plate. The carbon-lined \Jberyllium\j rotors are splined to the inside of the wheel and rotate with the wheel. The carbon-lined \Jberyllium\j stators are splined to the outside of the axle assembly and do not rotate with the wheel.
Each of the four main landing gear wheel brake assemblies is supplied with pressure from two different hydraulic systems. Each brake hydraulic piston housing has two separate brake supply chambers. One chamber receives hydraulic source pressure from hydraulic system 1 and the other from hydraulic system 2. There are eight hydraulic pistons in each brake assembly.
Four are manifolded together from hydraulic system 1 in a brake chamber. The remaining four pistons are manifolded together from hydraulic system 2. When the brakes are applied, the eight hydraulic pistons press the discs together, providing brake \Jtorque\j.
In the event of the loss of hydraulic system 1 or 2 source pressure, switching valves provide automatic switching to the standby hydraulic system 3 when the active hydraulic system source pressure drops below approximately 1,000 psi. If hydraulic system 1 is unavailable, it has no effect on the braking system because standby system 3 would automatically replace system 1.
Loss of hydraulic system 2 or both 1 and 2 would also have no effect on the braking system because system 3 would automatically switch to replace system 2 or 1 and 2. Loss of hydraulic systems 1 and 3 would cause the loss of half of the braking power on each wheel and additional braking distance would be required. Loss of hydraulic systems 2 and 3 would also cause the loss of half of the braking power on each wheel, requiring additional braking distance.
As in the landing gear deployment, the landing gear isolation valve in hydraulic systems 1, 2 and 3 must be open to allow the applicable hydraulic source pressure to the main landing gear brakes.
The brakes MN A, MN B and MN C switches are located on the flight deck display and control panels O14, O15 and O16 and allow electrical power to brake/anti-skid control boxes A and B. The antiskid switch located on panel L2 provides electrical power for enabling the anti-skid portion of the braking system boxes A and B.
The brakes MN A, MN B and MN C switches are positioned to on to supply electrical power to brake boxes A and B and to off to remove electrical power. The antiskid switch is positioned to on to enable the anti-skid system and to off to disable the system.
When weight is sensed on the main landing gear, the brake/anti-skid boxes A and B are enabled, permitting the main landing gear brakes to become operational.
The main landing gear brakes controlled by the commander's or pilot's brake pedals are located on the rudder pedal assemblies at the commander's and pilot's stations. The pedals' positions are adjustable by a handle. The braking commands are accomplished by the commander or pilot initiating toe pressure on the top of the rudder pedal assembly.
Each brake pedal (left and right) has four linear variable differential transducers. The left pedal \Jtransducer\j unit will output four separate braking signals through the brake/skid control boxes for braking control of the two left main wheels. The right pedal \Jtransducer\j unit does likewise for the two right main wheels. When toe pressure is applied to the brake pedal, the transducers transmit electrical signals of zero to 5 volts dc to the brake/anti-skid control boxes.
If both right pedals are moved, the pedal with the greatest toe pressure becomes the controlling pedal through electronic OR circuits. The electrical signal is proportional to the toe pressure. The electrical output energizes the main landing gear brake coils proportionally to brake pedal deflection, allowing the desired hydraulic pressure to be directed to the main landing gear brakes for braking action. The brake system bungee at each brake pedal provides the braking artificial feel to the crew member.
Each of the three hydraulic systems' source pressure of 3,000 psi is reduced by a regulator in each of the brake hydraulic systems to 1,500 psi.
The anti-skid portion of the brake system provides optimum braking by preventing tire skid or wheel lock and subsequent tire damage.
Each main landing gear wheel has two speed sensors that supply wheel rotational velocity information to the skid control circuits in the brake/skid control boxes. The velocity of each wheel is continuously compared to the average wheel velocity of all four wheels. Whenever the wheel velocity of one wheel is below 30 percent of the average velocity of the four wheels, skid control removes brake pressure from the slow wheel until the velocity of that wheel increases to an acceptable range.
The brake system contains eight brake/skid control valves that receive signals from the brake/skid control boxes. Each valve controls the hydraulic brake pressure to one of the brake chambers. The brake/skid control valves contain a brake coil and a skid coil. The brake coil allows hydraulic pressure to enter the brake chambers. The skid coil, when energized by the skid control circuit, provides reverse polarity to the brake coil, preventing the brake coil from allowing brake pressure to the brake chamber.
Anti-skid control is automatically disabled below 9 to 14 knots (11 to 17 mph) to prevent loss of braking for maneuvering and/or coming to a complete stop.
The anti-skid system control circuits contain fault detection logic. The antiskid yellow caution and warning light located on the flight deck display and control panel F3 will be illuminated if the anti-skid fault detection circuit detects an open or short in a wheel speed sensor, open or short in a anti-skid control valve servocoil or a failure in an anti-skid control circuit. A failure of these items will only deactivate the failed circuit, not total anti-skid control. If the brake power switches are on and the antiskid switch is off , the antiskid caution and warning light will be illuminated.
Insulation and electrical heaters are installed on the portions of the hydraulic systems that are not adequately thermally conditioned by the individual hydraulic circulation pump system because of stagnant hydraulic fluid areas.
Redundant electrical heaters are installed on the main landing hydraulic flexible lines located on the back side of each main landing gear strut between the brake module and brakes.
These heaters are required because the hydraulic fluid systems are dead-ended and fluid cannot be circulated with the circulation pumps. In addition, on OV-103 and OV-104, the hydraulic system 1 lines to the nose landing gear are located in a tunnel between the crew compartment and forward fuselage.
The passive thermal control systems on OV-103 and OV-104 are attached to the crew compartment, which leaves the hydraulic system 1 lines to the nose landing gear exposed to environmental temperatures, thus requiring electrical heaters on the lines in the tunnel. Since the passive thermal control system on OV-102 is attached to the inner portion of the forward fuselage rather than the crew compartment, no heaters are required on the hydraulic system 1 lines to the nose landing gear on OV-102.
The \Jhydraulics\j brake heater A, B, and C switches on panel R4 enable the heater circuits. On OV-103 and OV-104, \Jhydraulics\j brake heater switches A, B and C provide electrical power from the corresponding main buses A, B and C to the redundant heaters on the main landing gear flexible lines and the hydraulic system 1 lines in the tunnel between the crew compartment and forward fuselage leading to the nose landing gear. Thermostats on each electrical A, B and C system cycle the heaters automatically off or on.
The \Jhydraulics\j brake heater A, B and C switches on panel R4 enable the heater circuits on only the main landing gear hydraulic flexible lines on OV-102.
Because problems were encountered with the main landing gear braking system in the majority of the first 24 landings, an improvement program has been implemented for the main landing gear and braking system in addition to a long-term improvement program for the main landing gear brakes.
Main landing gear axle stiffness has been increased to reduce brake-to-axle deflections to preclude brake damage, which occurred in previous landings. This should also minimize tire wear. With the increased axle thickness, existing axle/bearing and axle/sensor interfaces are maintained. All main landing gear axles will be changed before the three orbiters return to flight.
Six orifices were added to the hydraulic passages in the brake hydraulic piston housing to restrict circular fluid flow within the chambers in order to stop the whirl phenomenon, which has been identified as the cause of brake damage.
The electronic brake control boxes were modified to provide hydraulic pressure balancing between adjacent brakes in order to equalize energy applications. This results in higher efficiency and allows full capability of adjacent brakes. The anti-skid circuitry that reduced brake pressure to the opposite wheel if a flat tire was detected was removed.
The previous, thinner, carbon-lined \Jberyllium\j stator discs are being replaced in two positions with thicker discs to provide a significant increase in braking energy capability. The additional material added to the stators improves heat capacity, with resulting lower temperatures, and provides the stators with greater strength.
Note that the main landing gear brakes, which were exposed to two 14-million- foot-pound wear-in cycles added before installation on the orbiter, reduced damage to the brakes during landing. The thicker stator discs will provide approximately 65 million foot pounds of energy absorption, which is a significant increase over the thinner stator discs.
A long-term structural carbon brake program is in progress to provide higher braking capability by increasing maximum energy absorption capability to 82 million foot pounds and to reduce refurbishment costs. These new brakes will consist of a five-rotor, disc-type carbon configuration for each main landing gear wheel brake.
The goal is to demonstrate that the carbon heat sink brake design will have the capability of providing a one-time stop of 100 million foot pounds. The go-ahead for the carbon brake design was given in January 1986, with delivery scheduled in late April 1988.
Upon the return to flight of the space shuttle, end-of-mission landings are planned for Edwards Air Force Base in \JCalifornia\j until the performance of the landing gear system is fully understood and a higher confidence in the weather prediction capability is established at the Kennedy Space Center shuttle landing facility runway area in \JFlorida\j.
Strain gauges have been added to each nose and main landing gear wheel in order to monitor tire pressure and provide the status of tire pressures during launch pad stay, launch, orbit, deorbit and landing to the flight crew and Mission Control Center in Houston.
A landing gear tire improvement and runway-surface study is in progress at NASA's Langley Research Center to determine how best to decrease tire wear experienced during previous Kennedy Space Center landings and improve crosswind landing capability. Six new tire designs will be evaluated at the Langley Research Center with various yaw and tilt angles at speeds up to 253 mph. Additional tests at Langley Research Center are to provide the ability to mathematically model tire side-force characteristics.
Modifications were made to the Kennedy Space Center shuttle landing facility runway. The full 300-foot width of 3,500-foot sections at both ends of the runway were ground to smooth the runway surface and remove cross grooves. The corduroy ridges are smaller than those they replace and run the length of the runway rather than across its width.
The existing landing zone light fixtures were also modified, and the markings of the entire runway and overruns were repainted. The primary purpose of the modifications is to enhance safety by reducing tire wear during landing.
The current strength of the orbiter landing system is a maximum of 240,000 pounds. Evaluations for landing loads of 256,000 pounds, associated with abort landings, are to be completed by the spring of 1988.
Other studies in progress are arrest barriers at landing sites (except lake bed runways) to provide safe stops in the event of main landing gear brake failure or unforeseen wet runway conditions. The barrier net study will determine whether the barrier can safely stop a 256,000-pound orbiter traveling at 100 knots (115 mph) at the end of runway.
Also under study are (1) the installation of a skid on the landing gear that could preclude the potential for a second blown tire on a gear on which one tire has blown, (2) a rim that would provide a predictable roll in the event of the loss of both tires on a single or multiple gear and (3) the addition of a drag chute.
#
"Shuttle Nose Wheel Steering",354,0,0,0
The orbiter nose wheel is steerable after nose wheel touchdown at landing. The nose wheel is electrohydraulically steerable through the use of the general-purpose computers and the commander's or pilot's rudder pedals, in conjunction with the orbiter flight control system in the control stick steering mode. In addition, the nose wheel may also be steerable through the use of the commander's or pilot's rudder pedals in the direct mode.
Nose wheel steering is advantageous to allow positive lateral directional control of the orbiter during the rollout phase of a mission in the presence of high crosswinds and blown tires. Recent modifications of the nose wheel steering system have been incorporated to allow a safe high-speed engage of the nose wheel steering system.
The first in the series of changes was to redefine the NWS switch on panel L2. The forward position of the switch now activates the direct mode of nose wheel steering, the center position activates the GPC mode of nose wheel steering, and the aft switch off position deactivates all nose wheel steering. A flexible handle extension was also added to the switch handle.
The commander selects and activates the GPC mode while performing the pre-entry checklist of cockpit switches while his eyes and attention are inside the cockpit. No other flight crew action is required in order to engage active nose wheel steering at the time of nose wheel touchdown. Other added features provide assurance that the nose wheel will be positioned straight ahead at the moment of nose wheel touchdown, and the final enable signals in the GPC mode are sequenced at nose wheel touchdown and begin active nose wheel steering at that time.
During derotation and nose wheel touchdown, the flight crew is looking out the windshield and at the head-up display. If the commander chooses to select the direct mode of nose wheel steering or deactivate all nose wheel steering by selecting off without visual verification of the actual switch position, he can sweep his left hand over the switch (forward motion for direct or aft motion for off ) and be certain of the position of the NWS switch without looking at the switch.
Only hydraulic system 1 supplies hydraulic system supply pressure to the nose wheel hydraulic actuator for steering in either the GPC or direct mode. If hydraulic system 1 supply pressure is unavailable, the commander or pilot can apply the left and right main landing brakes by applying toe pressure on the rudder pedals differentially, which allows directional control by differential braking. The nose wheel steering actuator limits the motion of the nose wheel to plus or minus 10 degrees and also prevents nose wheel shimmey.
When the NWS switch is set to either GPC or direct , the nose wheel steering system fault detection logic detects (1) the interruption of hydraulic system 1 supply pressure; (2) an open or short in the nose wheel steering servovalve circuitry; (3) an open or short in the position feedback, rate position error; (4) an open or short in the command \Jtransducer\j; and (5) broken linkage or loss of electrical power. At this time, the NWS fail C/W light on panel F3 is illuminated, and the nose wheel steering reverts automatically to the caster mode.
General-Purpose Computer Mode. The NWS switch on panel L2, positioned to GPC, enables the nose wheel steering \Jsolenoid\j control valves, which supplies hydraulic system 1 pressure to the nose wheel steering servovalves. In addition to the GPC mode selection of the NWS switch, the flight control system roll/yaw CSS push button switch/light indicator on panel F2 or F4 must be depressed to enable GPC mode steering. When either push button switch/light indicator on panel F2 or F4 is depressed, a white light is illuminated within the push button.
When the commander or pilot positions the rudder pedals in the GPC roll/yaw CSS mode, the rudder pedals command position is appropriately scaled within the GPC's software and transmitted to a summing network, along with accelerometer inputs from within the flight control system. The accelerometer inputs are utilized to prevent any sudden lateral deviation of the orbiter's velocity vector direction. From this summing network, a nose wheel steering command is sent to a comparison network as well as to the steering servo system.
The three new steering position transducers in the unit added on the nose wheel strut receive redundant electrical excitation from the new steering position amplifier (added to the middeck ceiling), which receives redundant electrical power from data display unit 2 on the flight deck.
Each of the three new transducers transmits nose wheel position feedback to a redundancy management mid-value-select software, which then transmits a nose wheel position signal to the comparison network. The orbiter nose wheel commanded and actual positions are compared for position error and for rates to reduce any error.
Absence of an error condition will enable nose wheel steering after weight on the nose gear is sensed in the software. Weight on the nose gear requires weight on all three landing gear and a nose-down orbiter attitude. The enable signal permits hydraulic system 1 pressure to be applied to the nose wheel steering actuator. If hydraulic pressure is below 1,350 psi, the actuator remains in the shimmy damp mode, and a failure is annunciated to the NWS fail C/W yellow light on panel F3.
If hydraulic system 1 pressure is above 1,350 psi, the actuator is configured to the GPC CSS mode to position the nose wheel utilizing the commander's or pilot's rudder pedals.
If more than one of the three new position \Jtransducer\j feedback signals are lost, the comparison logic switches hydraulic system 1 supply pressure off to the nose wheel steering actuator, and pressure sensing immediately inhibits nose wheel steering in the GPC CSS mode automatically.
Nose wheel steering will then revert to a nose wheel caster mode and the NWS fail yellow C/W light will be illuminated on panel F3. This redundancy in the GPC CSS mode provides immediate protection against any undesired lateral nose wheel response during nose wheel steering.
If nose wheel steering has failed, the NWS switch on panel L2 positioned to off will extinguish the NWS fail C/W yellow light on panel F3, unless the GPC CSS mode comparison circuits caused the light to be illuminated. When the NWS switch is positioned to off, hydraulic system 1 supply pressure to the nose wheel steering system is removed, allowing the nose wheel to caster.
Direct Mode. When the NWS switch on panel L2 is positioned to direct, the nose wheel steering \Jsolenoid\j control is enabled after weight is sensed on the nose gear. This supplies hydraulic system 1 supply pressure to the nose wheel steering servo system.
When the commander or pilot positions the rudder pedals in the direct mode, the rudder pedals command position is shaped to provide a less sensitive rudder pedal command in the midrange and provides steering commands directly to the nose wheel steering servoactuator, bypassing the GPC software altogether.
The two additional steering position transducer/amplifier boxes are from Honeywell, the NWS box was modified by Sterer, and the DDU 2 modification was made by Collins.
A new additional nose wheel steering system design is expected to be available in late 1988 or early 1989.
#
"Shuttle Caution and Warning System",355,0,0,0
The primary caution and warning system is designed to warn the crew of conditions that may adversely affect orbiter operations. The system consists of hardware and \Jelectronics\j that provide the crew with both visual and aural cues when a system exceeds predefined operating limits.
The primary system's visual cues consist of four master alarm lights, a 40-light array on panel F7 and a 120-light array on panel R13. The aural cue is sent to the communications system for distribution to flight crew headsets or speaker boxes.
The C/W system interfaces with the auxiliary power units, data processing system, environmental control and life support system, electrical power system, flight control system, guidance and navigation, \Jhydraulics\j, main propulsion system, reaction control system, orbital maneuvering system and payloads. The audio alarms are classified as emergency (class 1), C/W (class 2) and alert (class 3).
The emergency alarms consist of a siren (activated by the smoke detection system) and a klaxon (activated by the delta pressure/delta time sensor that recognizes a rapid loss of cabin pressure), and they are annunciated by hardware. The siren's frequency varies from 666 to 1,470 hertz and returns at a five-second-per-cycle rate. The klaxon is a 2,500-hertz signal with an on/off cycle of 2.1 milliseconds on and 1.6 milliseconds off, mixed with a 270-hertz signal with a cycle of 215 milliseconds on and 70 milliseconds off.
The class 2 alarm is activated by the primary (hardware) system, the backup (software) system or both. The C/W tone is an alternating 375 hertz and 1,000 hertz at 2.5 hertz. The alternating C/W alarm tone is generated when the hardware system detects an out-of-limit condition on any of the 120 parameters it monitors or when the software (backup) system detects a parameter that is out of limits.
Both guidance, navigation and control and systems management software sense out-of-limit conditions. These software systems also serve some less critical parameters and annunciate the systems management alert tone. The SM alert tone is a steady tone of 512 hertz of predefined duration generated in the C/W \Jelectronics\j when activated by inputs from the onboard computers.
Visual cues for the flight crew consist of four red master alarm push button light indicators on panels F2, F4, A7 and M052J; the 40-light (red or yellow) C/W light array on panel F7; the 120 parameter status lights on panel R13; the blue SM alert light on panel F7; the red backup C/W light on panel F7; fault messages on \Jcathode\j ray tubes; and status characters on CRTs.
Inputs enter the C/W logic circuitry from the onboard computers through multiplexers/demultiplexers to activate alarm tones and the backup C/W alarm. Some of these are used to turn the backup C/W light on panel F7 on and off. One additional signal resets the master alarm lights and tones.
The primary C/W system has three modes of operation: ascent, normal and acknowledge. These modes are controlled by the caution/warning asc , norm, ack switch on panel C3. The normal mode is discussed first.
The C/W ascent mode is the same as the normal mode, except that the commander's red master alarm push button light indicator will not be illuminated.
The C/W acknowledge mode is also the same as the normal mode, except that the 40 annunciator lights on panel F7 will not be illuminated unless one of the red master alarm push button light indicators on panel F2 for the commander or panel F4 for the pilot is depressed.
Each of the 120 status C/W red parameter lights on panel R13 receives an input from a specific parameter. A primary C/W parameter matrix cue card identifies the 120 input channels and correlates them to the panel F7 C/W annunciator light matrix. If an out-of-limit condition exists on a specific parameter that is set on panel R13, it illuminates the corresponding light on panel F7. If the caution/warning param status switch on panel R13 is held in the tripped position when an out-of-limit parameter light on panel F7 is illuminated, the corresponding light on panel R13 will also be illuminated.
The three caution/warning parameter select thumbwheels on panel R13 provide signals to the C/W \Jelectronics\j unit and define the specific parameter for enabling and inhibiting the parameter and setting and reading the parameter's limits.
The caution/warning limit set switch grouping on panel R13 is used to change limits or to read a parameter's limits. The three value thumbwheels provide the signals to the C/W unit, defining the voltage value setting of a parameter's upper or lower limit, X.XX.
The caution/warning limit set limit upper switch on panel R13 provides a signal to the C/W \Jelectronics\j unit, which modes the \Jelectronics\j to set or read the upper limit of a parameter specified by the settings on the value thumbwheels for that parameter; and the caution/warning limit func switch is cycled to set or read the upper limit of that parameter. The caution/warning limit lower switch on panel R13 functions in the same manner as the limit upper switch, except for the lower limit for a parameter.
The caution/warning limit set func set switch position on panel R13 provides a signal to the C/W \Jelectronics\j unit, which sets the value specified by the limit set value thumbwheels into the parameter as specified by the parameter select thumbwheels and limit set limit switch.
The limit set func read switch position on panel R13 provides a signal to the C/W \Jelectronics\j unit, which illuminates the lights under the status limit volts X.XX columns on panel R13, that correspond to the voltage parameter limit specified by the parameter select thumbwheels and the limit set limit switch. The value read corresponds to the parameter's full-scale range on a scale of zero to 5 volts dc. The limit sec func switch center position disables the set and read functions.
The caution/warning param enable switch position on panel R13 provides a signal to the C/W \Jelectronics\j unit to enable the parameter indicated on the parameter select thumbwheels, which allows the parameter to trigger the primary C/W alarm when out of limits. The inhibited position operates the same as enable , except it inhibits the parameter from triggering the primary C/W alarm. The center position of the switch disables the enable and inhibit functions.
The caution/warning param status tripped switch position on panel R13 provides a signal to the C/W \Jelectronics\j unit, which illuminates the C/W status lights on panel R13 that correspond to the parameters that are presently out of limits, including those that are inhibited. The inhibited position illuminates those C/W lights on panel R13 that have been inhibited. The center position disables the tripped and inhibited functions.
The caution/warning memory read switch position on panel R13 provides a signal to the C/W \Jelectronics\j unit, which illuminates the C/W status lights on panel R13 that correspond to the parameters that have been out of limits since the last positioning of this switch or the caution/warning memory switch on panel C3 to clear .
The clear position on panel R13 or panel C3 provides a signal to the C/W \Jelectronics\j unit that clears from the memory any parameters that are presently within limits, but any parameters that are out of limits during this action remain in memory. The center position of the switch on panel R13 or panel C3 disables the clear and read functions.
The caution/warning tone volume A switch on panel R13, when adjusted clockwise, increases the system A siren, klaxon, C/W, and SM tone generator output signals to the audio central control unit. The B switch functions the same as the A switch for system B tone generators.
The caution/warning lamp test switch on panel R13, when positioned to left, provides a signal to the C/W \Jelectronics\j unit, which illuminates the left five columns of the C/W status matrix lights on panel R13. The right position functions the same as the left , except for the right five columns of lights.
The backup C/W system is part of the systems management fault detection and \Jannunciation\j, GN&C and backup flight system software programs. All backup C/W alarms are class 2. Only the 69 backup C/W alarms that are produced by FDA have limits that can be changed and displayed in \Jengineering\j units accessed through the SM table maintenance specialist function display (SPEC 60).
The remaining backup C/W alarms that are produced by the guidance and navigation program are accessed through general-purpose computer read/write procedures. A backup C/W out-of-tolerance condition will trigger a master alarm light, illuminate the red backup C/W alarm light on panel F7, and display a message on the fault message line and fault summary page on the SM CRT.
The SM alert program is another portion of the SM program and operates like the backup C/W system. It is designed to inform the flight crew of a situation leading up to a C/W or one that may require additional procedures. When an SM alert parameter exceeds its limits, the blue SM alert light on panel F7 is illuminated, a discrete is sent to the primary C/W system to turn on the SM tone, and the software displays a fault message on the fault message line and fault summary page on the SM CRT.
Annunciator lights provide visual indications of the status of the vehicle and payload systems. The annunciator lights are classified as emergency, warning, caution and advisory. Emergency and warning annunciators are red; cautions are yellow; and advisory may be white (status), green (normal configuration), yellow (alternate configuration) or blue (special applications).
Annunciator lighting is provided by incandescent lamps that illuminate the lens area of the annunciators. Most annunciators are driven by an annunciator control assembly that controls the illumination of the lights during a normal or test input and the brightness level. The C/W status lights and GPC status lights have separate electronic units for lighting control.
There are three different lens configurations for push button indicator and indicator lights. One configuration has illuminated nomenclature in the appropriate color on an opaque black background, and the nomenclature cannot be seen until it is illuminated.
Another configuration has non-illuminated white nomenclature on an opaque black background and a bar that illuminates in the appropriate color; this nomenclature is always visible. The third configuration has a bar that is illuminated on an opaque black background and no nomenclature on the lens, but the nomenclature is available as part of the panel.
#
"Orbiter Lighting System",356,0,0,0
The orbiter lighting system provides both interior and exterior lighting. The interior lighting provides illumination for display and control visibility and general flight station and crew equipment operations. Exterior lighting provides illumination for payload bay door operations, extravehicular activity, remote manipulator system operations, and stationkeeping and docking. Interior lighting consists of floodlights, panel lights, instrument lights, numeric lights and annunciator lights. Annunciator lighting is discussed with the caution and warning system. Exterior lighting consists of floodlights and spotlights.
Interior floodlights provide general illumination throughout the crew cabin and allow the flight crew to function within the flight deck, middeck, airlock and tunnel adapter (if installed). Both fluorescent and incandescent lamps are used. Emergency lighting is provided by selected fixtures that are powered via a separate power input from an essential bus.
Dual fluorescent lamp fixtures provide lighting for the mission station and payload station. The mission station lighting is controlled by an on/off switch with a rotary control switch to control brightness on panel R10. The payload station lighting is controlled by an on/off switch with a rotary control switch to control brightness on panel L9.
A single fluorescent lamp fixture is employed on each side of the commander's and pilot's forward flight deck glareshield, the commander's and pilot's side consoles and the orbit station. The commander's glareshield light is controlled by the bright, var , off switch and a dim, brt rotary control on panel O6. The dim, brt rotary control operates in conjunction with the var position. The pilot's glareshield light functions the same as the commander's, except the control is on panel O8.
The commander's side console light is controlled by an integral off/variable/on control switch. The pilot's side console light is also controlled by an integral off/variable/on control switch. The orbit station light is controlled by an on/off switch and dim, brt rotary control on panel A6.
The galley and middeck sleep station bunks (if installed) use the same floodlights as the commander's and pilot's flight deck consoles and are also controlled individually.
The airlock floodlights are similar to those at the commander's and pilot's flight deck side consoles, except they are controlled by switches on panels AW18A and M013Q.
If the tunnel adapter is installed for a Spacelab mission, the floodlights are also similar to those at the commander's and pilot's side consoles. Tunnel adapter lights 2, 3 and 4 are controlled by individual on/off switches on the tunnel adapter panel. The remaining tunnel adapter light is controlled by the tunnel adapter 1 on/off switch on panel M013Q and the on/off 1 switch on the tunnel adapter panel. The emergency floodlights are controlled by on/off switches on either panel C3 or ML18F.
\BPanel Lighting\b
Many flight deck instrument panels have integral lighting that illuminates the panel nomenclature and markings on the displays and controls. This illumination aids the flight crew in locating displays and controls while operating the orbiter. Panel lighting is transmitted from behind a panel overlay through the panel nomenclature, making it appear white-lighted. It is also transmitted to the edges of the displays and controls for general illumination. The lighting source consists of small incandescent, grain-of-wheat lamps mounted between the metal panel face and the plastic panel overlay. The overlay has a layer of white paint and a layer of gray paint on the top surface. The panel nomenclature is formed by etching the letters and symbols into the gray paint, leaving the white layer underneath. The panel, mission station and orbit station lighting is controlled by off, var, brt rotary controls on panels O6, O8, R10 and A6.
\BInstrument Lighting\b
The flight deck instruments have integral lighting that enables the flight crew to read the displayed data. Lighting is provided by incandescent lamps located behind the face of the instruments. Prisms are used to distribute the light evenly over the face. Instrument, lighting panel, and orbit station lighting are controlled by off, var, brt rotary controls on panels O6, O8 and A6.
\BNumeric Lighting\b
Six indicators on the flight deck use illuminated numeric (digital) readouts to display data. The illumination is provided by a single incandescent lamp in each segment of a digit. Seven segments are required to generate the numbers zero through nine. Each numeric indicator has a red light to indicate failures in the indicator and will be illuminated when any lamp in the indicator fails. The six numeric (digital) indicators are event time (panels F7 and A4), mission time (panels O3 and A4), RCS/OMS prplt qty (panel O3) and rri (rendezvous radar) (panel A2). The panel lighting and numeric orbit station lighting are controlled by off, var, brt rotary controls on panels O8 and A6.
\BExterior Floodlights\b
The exterior floodlights improve visibility for the flight crew during payload bay door operations, EVA operations, RMS operations, and stationkeeping and docking. The payload bay floodlights are controlled by switches on panel A7. The RMS floodlight is also controlled on panel A7.
The payload bay floodlights are metal halide lamps that are gas discharge arc tubes similar to mercury vapor lamps. Two different fixtures are used with the lamps: one fixture mounts the floodlight on the payload bay forward bulkhead and the other fixture mounts the floodlight within the payload bay. The RMS floodlight uses an incandescent lamp. It is located near the RMS end effector.
#
"Shuttle Smoke Detection and Fire Suppression",357,0,0,0
Smoke detection and fire suppression capabilities are provided in the crew cabin avionics bays, the crew cabin and the Spacelab pressurized module. \JIonization\j detection elements, which sense levels of smoke concentrations or rate of concentration change, trigger alarms and provide information on smoke concentration levels to the performance-monitoring CRT system and displays on flight deck display panel L1.
The \Jionization\j detection system is divided into two groups: group A and group B. Group A \Jionization\j detection elements are located in the environmental control and life support system cabin fan plenum outlet beneath the crew cabin middeck floor and in the left return air duct on the crew cabin flight deck; and one element is located in each of the three avionics bays (bays 1, 2 and 3A). Group B \Jionization\j detection elements are located in the right return air duct on the crew cabin flight deck and in avionics bays 1, 2 and 3A. On Spacelab missions, \Jionization\j detection elements are located in the pressurized Spacelab module.
If an \Jionization\j detection element senses a smoke concentration of 2,200, plus or minus 200, micrograms per cubic meter or a rate of smoke increase of 22 micrograms per cubic meter per second for eight consecutive counts in 20 seconds, a trip signal illuminates the applicable red smoke detection A or B light on panel L1, activates the C/W master alarm red lights and sounds the siren in the crew cabin. The normal reading on the CRT for the smoke detection elements is 0.3 to 0.4. A reading on the CRT of 2.2, plus or minus 0.2, corresponds to 2,200, plus or minus 200, micrograms per cubic meter.
Fire suppression in the crew cabin avionics bays is provided by one Freon-1301 (bromotrifluoromethane) extinguisher bottle in each of the three avionics bays. Each bottle contains 3.74 to 3.8 pounds of Freon-1301 in a pressure vessel that is 8 inches long and 4.25 inches in diameter.
To activate the applicable Freon-1301 bottle in an avionics bay, the corresponding fire suppression av bay switch on panel L1 is positioned to arm, and the corresponding agent disch push button light indicator on panel L1 is depressed for at least two seconds. The agent disch push button light indicator activates the corresponding pyro initiator controller, which initiates a pyrotechnic valve on the bottle to discharge the Freon-1301 into the avionics bay. When the Freon-1301 bottle is fully discharged, the push button light indicator white light will be illuminated. The white light will also be illuminated if the pressure in an avionics bottle decays before use.
The red smoke detection cabin light on panel L1 is illuminated by the smoke detection \Jionization\j element in the ECLSS cabin fan plenum, the red l flt deck fire detection light is illuminated by the crew cabin left flight deck return air duct smoke \Jionization\j element, the red r flt deck fire detection light is illuminated by the crew cabin right flight deck return air duct smoke \Jionization\j element, and the red payload fire detection light is illuminated by the smoke detection \Jionization\j elements in the Spacelab pressurized module if it is a Spacelab mission.
The applicable smoke detection \Jionization\j element trips at the same levels as the avionics bay elements, illuminates the applicable red smoke detection A or B light on panel L1, activates the C/W master alarm red lights and sounds the siren in the crew cabin.
Three portable hand-held fire extinguishers are available in the crew cabin. Two are located in the crew cabin middeck and one is on the flight deck. Each fire extinguisher nozzle is tapered to fit fire hole ports in the display and control panels. The extinguishing agent is Halon-1301 (monobromotrifluoromethane). Halon-1301 minimizes the major hazards of a conflagration: smoke; heat; oxygen depletion; and formation of pyrolysis products, such as carbon monoxide. The fire extinguishers are 13 inches long. The portable fire extinguishers can also be used as a backup for extinguishers in the avionics bays.
Various parameters of the smoke detection system and remote fire extinguishing agent system are provided to telemetry.
The smoke detection circuit test switch on panel L1 tests the smoke detection system, lights and alarm circuitry. When the switch is positioned to A or B, electrical power is applied to the ACA channels controlling the agent disch lights, and the white lights are illuminated. After approximately a 20-second delay, the smoke detection A or B lights are illuminated and the siren is triggered.
The smoke detection sensor switch on panel L1 resets a tripped smoke detection element.
The contractors involved with the smoke detection and fire suppression system are \JBrunswick\j Celesco, Costa Mesa, Calif. (smoke detectors and remote control fire extinguishing agent); J.L. Products, Gardena, Calif. (arming fire push button); and Metalcraft Inc., Baltimore, Md. (portable fire extinguishers).
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"Shuttle Payload Deployment and Retrieval System",358,0,0,0
The payload deployment and retrieval system includes the electromechanical arm that maneuvers a payload from the payload bay of the space shuttle orbiter to its deployment position and then releases it. It can also grapple a free-flying payload, maneuver it to the payload bay of the orbiter and berth it in the orbiter. This arm is referred to as the remote manipulator system.
The RMS is installed in the payload bay of the orbiter for those missions requiring it. Some payloads carried aboard the orbiter for deployment do not require the RMS.
The RMS is capable of deploying or retrieving payloads weighing up to 65,000 pounds. The RMS can also retrieve, repair and deploy satellites; provide a mobile extension ladder for extravehicular activity crew members for work stations or foot restraints; and be used as an inspection aid to allow the flight crew members to view the orbiter's or payload's surfaces through a \Jtelevision\j camera on the RMS.
The PDRS was built via an international agreement between the National Research Council of Canada and NASA. Spar Aerospace Ltd., a Canadian company, designed, developed, tested and built the RMS. CAE \JElectronics\j Ltd. in Montreal provides electronic interfaces, servoamplifiers and power conditioners. Dilworth, Secord, Meagher and Assoc. Ltd. in Toronto is responsible for the RMS end effector. Rockwell International's Space Transportation Systems Division designed, developed, tested and built the systems used to attach the RMS to the payload bay of the orbiter.
The basic RMS configuration consists of a manipulator arm; an RMS display and control panel, including rotational and translational hand controllers at the orbiter aft flight deck flight crew station; and a manipulator controller interface unit that interfaces with the orbiter computer. Normally, only one RMS is installed on the left longeron of the orbiter payload bay.
The RMS could be installed on the right side, but the orbiter Ku-band antenna would have to be removed to accommodate the RMS there. Two arms could be installed in the payload bay if the orbiter Ku-band antenna were removed, but only one arm could be operated at a time because only a single software package (computer program) and a single set of display and control panel hardware are provided at the flight deck aft control station. Electrical wiring is in the flight deck aft station for both arms.
One flight crew member operates the RMS from the aft flight deck control station, and a second flight crew member usually assists with \Jtelevision\j camera operations. This allows the RMS operator to view RMS operations through the aft flight deck payload and overhead windows and through the closed-circuit \Jtelevision\j monitors at the aft flight deck station.
The RMS arm is 50 feet 3 inches long and 15 inches in diameter and has six degrees of freedom. It weighs 905 pounds, and the total system weighs 994 pounds.
The RMS has six joints that correspond roughly to the joints of the human arm, with shoulder yaw and pitch joints; an elbow pitch joint; and wrist pitch, yaw and roll joints. The end effector is the unit at the end of the wrist that actually grabs, or grapples, the payload. The two lightweight boom segments are called the upper and lower arms. The upper boom connects the shoulder and elbow joints, and the lower boom connects the elbow and wrist joints. The RMS arm attaches to the orbiter payload bay longeron at the shoulder manipulator positioning mechanism. Power and data connections are located at the shoulder MPM.
The RMS can operate with standard or special-purpose end effectors. The standard end effector can grapple a payload, keep it rigidly attached as long as required and then release it. Special-purpose end effectors are designed by payload developers and installed instead of the standard end effector during ground turnaround. An optional payload electrical connector can receive electrical power through a connector located in the standard end effector.
The booms are made of \Jgraphite\j epoxy. They are 13 inches in diameter by 17 feet and 20 feet, respectively, in length and are attached by metallic joints. The composite in one arm weighs 93 pounds. The joint and electronic housings are made of aluminum alloy.
A shoulder brace relieves launch loads on the shoulder pitch gear train of the RMS. On orbit, the brace is released to allow RMS operations. It cannot be relatched on orbit, but it is not required that it be relatched for entry or landing loads. A plunger is extended between two pieces of tapered metal, pushing the ends of the pieces outward, wedging the ends of the receptacle on the outer casing of the shoulder yaw joint, and engaging the shoulder brace.
Shoulder brace release is controlled by the lever-locked shoulder brace release starboard, port switch on panel A8U. Normally, the RMS arm is installed on the orbiter's port, or left, side. Positioning the switch to port releases the port shoulder brace, which withdraws the plunger by an electrical linear actuator. This allows the tapered metal pieces to relax and move toward each other, which permits the brace to slide out of the shoulder yaw outer casing, unlatching the brace.
The switch must be held until the shoulder brace release talkback indicator on panel A8U indicates gray, which usually takes six to nine seconds. A microswitch at the end of the plunger's travel controls the talkback indicator. A barberpole indication shows that the shoulder brace is still latched. The shoulder brace release switch and talkback indicator cannot receive electrical power until the RMS select switch and RMS power switch on panel A8L are positioned for electrical power.
The RMS select switch on panel A8L selects the arm to be used, or it indicates off if no arm is to be used. The switch essentially tells the manipulator controller interface unit and display and control panel A8L which arm is being powered and operated. The status of the switch is also displayed on the aft flight control station CRT display.
The RMS power switch on panel A8L has a guard over it and connects orbiter main dc bus and AC1 phase A power to the RMS, MCIU and displays and controls on panels A8L and A8U when positioned to primary . Main dc bus and AC2 phase A power are connected to the RMS and some displays and controls on panels A8L and A8U when the switch is positioned to backup . The status of the switch is also displayed on the aft flight deck CRT. The off position removes all power from the selected arm. Panel A8L has a set of switches for the starboard RMS on the left side and an identical set for the port RMS on the right side of the panel that controls the selected RMS heaters, retention system and positioning mechanism. Normally, the port RMS set of switches is used.
The RMS is rotated 31.36 degrees toward the payload so the payload bay doors can be closed and is rotated 31.36 degrees away from the payload bay when the payload bay doors are opened. The manipulator positioning mechanism rolls the arm from its stowed position (toward the payload bay) to its operating position (away from the payload bay). The MPM consists of four pedestal joints joined by a \Jtorque\j tube and operated as a single assembly. The shoulder of the RMS arm attaches to the orbiter payload bay longeron (the sill of the bay) at the forwardmost MPM pedestal, and the aft three MPM pedestals contain latches that secure the arm along the orbiter payload bay longeron.
An RMS talkback indicator above the RMS deploy, off, stow switch on panel A8L indicates sto when the arm is stowed, barberpole when the arm is in transit and dep when the arm is deployed. The talkback indicator is controlled by four microswitches on each of the four pedestals. These microswitches can also be monitored by telemetry.
A manipulator retention latch is located in each of the three aft MPM pedestals. It locks a corresponding striker bar on the arm, locking the arm to the MPM.
When the RMS retention latches, release, off, latch switch on panel A8L is positioned to release , each MRL has two redundant ac motors that drive the MRL open; and two microswitches, one for each motor on each MRL, remove electrical power from that ac motor when it reaches its limit of travel. With both ac motors operating, it takes approximately eight seconds to fully open that latch.
If only one ac motor is in operation, it takes approximately 18 seconds for the latch to fully open. The talkback indicator above the release, off, latch switch on panel A8L indicates barberpole when the MRLs are in transit and rel when they are released. There are two release microswitches on each of the MRLs that control the talkback indicator. These microswitches can be monitored by the flight crew on the aft flight deck CRT.
The RMS arm is now available for operation.
When the RMS arm is properly aligned and resting on the MPM pedestals for latching to the MPMs, the three RMS striker bars are in the MRL ready-to-latch envelope. Two microswitches in each MRL control the corresponding aft, mid, or fwd ready for latch talkback indicators on panel A8L.
When the talkback indicators show gray, the corresponding MRL is positioned to latch the corresponding arm striker bar. If a talkback indicator shows barberpole, the MRL is not correctly aligned or not in position to be able to latch down the arm. As a result, the flight crew must reposition the arm until the talkback indicators show gray.
These microswitches can be monitored by the flight crew at the aft flight deck CRT. When the flight crew sees three gray ready for latch talkbacks, it positions the retention latches switch to latch. The two ac motors in each MRL drive the MRL closed; and two microswitches, one for each motor, remove electrical power from that ac motor when it reaches its limit of travel. The operating time for both motors or one motor would be the same as for release.
The talkback indicator above the release, off, latch switch indicates barberpole when the MRLs are in transit and lat when they are latched, thus holding the arm in the MRLs.
The RMS has both passive and active thermal control systems. The passive system consists of multilayer \Jinsulation\j blankets and thermal coatings that reflect solar energy away from the arm and aid in controlling the temperature of the hardware. The blankets are attached to the arm structure and to each other with Velcro. Exposed areas around the moving parts are painted with a special white paint.
To maintain the arm's temperature within predetermined operating limits, an active system, which consists of 26 heaters on the arm, supplies 520 watts of power at 28 volts dc. There are two redundant heater systems: one powered from the orbiter's main A dc bus and the other from the main B dc bus. Only one system is required for proper thermal control.
The heaters in each system are concentrated at the arm's joint and end effector to heat the \Jelectronics\j and ac motor modules. The heaters are enabled by the heater auto, off guarded switch on panel A8L. When the switch is positioned to auto, the heaters are thermostatically controlled by 12 thermistors located along the arm. The heaters are automatically turned on at 14 F and off at 43 F.
The orbiter's CCTV aids the flight crew in monitoring PDRS operations. The arm has provisions on the wrist joint for a CCTV camera that can be zoomed, a viewing light on the wrist joint and a CCTV with pan and tilt capability on the elbow of the arm. In addition, four CCTV cameras in the payload bay can be panned, tilted and zoomed.
Keel cameras may be provided, depending on the mission payload. The two CCTV monitors at the aft flight deck station can each display any two of the CCTV camera views simultaneously with split screen capability. This shows two views on the same monitor, which allows crew members to work with four different views at once. Crew members can also view payload operations through the aft flight station overhead and aft (payload) viewing windows.
The RMS can only be operated in a zero-gravity environment, since the arm dc motors are unable to move the arm's weight under the influence of Earth's gravity. Each of the six joints has an extensive range of motion, allowing the arm to reach across the payload bay, over the crew compartment or to areas on the undersurface of the orbiter. Arm joint travel limits are annunciated to the flight crew arm operator before the actual mechanical hardstop for a joint is reached.
Each joint of the arm is driven electromechanically. Each joint has one dc motor and associated ABE. Each dc motor turns a gear train, which produces joint motion. A tachometer on the output side of each motor measures motor rate. Also, on the output side of the gear train is an optical encoder that measures the actual joint angle and feeds it back to the software. There are two optical commutators on the input side of each motor: one commutator electronically interfaces with the primary motor drive amplifier; and one electronically interfaces with the backup drive amplifier, which is the only redundancy in each joint motor.
The arm has a number of operating modes. Some of these modes are computer-assisted, moving the joints simultaneously as required to put the end point (the point of resolution, such as the tip of the end effector) in the desired location. Other modes move one joint at a time; e.g., single mode uses software assistance and direct and backup hard-wired command paths that bypass the computers.
When the arm is used in the computer-assisted mode, the command from the flight crew operator is converted by the computer to a set of motor speeds (one for each joint) that move the arm to the desired configuration. The software scales down the set of commands so that no joint exceeds the maximum allowable joint rate.
This is called rate limiting, with the maximum joint rates dependent on the payload being flown and chosen so the arm can be stopped in 2 feet. The software also checks that the POR can be stopped within 2 feet. This is called POR rate limiting. For example, the tip of the unloaded arm cannot be moved more than 2 feet per second, and a 32,000-pound payload cannot be moved more than 0.2 foot per second.
The motor drive amplifier for each joint (total of six) can receive either a hard-wired direct drive input when the arm is not in the computer-assisted mode of operation or receive the error signal from the tachometer feedback loop. When the MDA receives its signal from the feedback loop, it gets additional input, called the current limit command, from the computer. This input controls the maximum \Jtorque\j of the motor; thus, arm loads are maintained within the defined limits while operating. The current limit can only be changed with a computer memory read/write procedure.
One backup drive amplifier for the entire arm is located in the shoulder \Jelectronics\j compartment. When the arm is operated in the backup mode, the drive unit goes to the motor via the BDA and bypasses the feedback loop.
Normal braking is accomplished by each joint motor deceleration; however, each joint has a mechanical friction-type brake, and all six brakes are operated by a single switch on panel A8U. When the brakes on, off switch is positioned to on, brakes in all joints of the arm are applied.
The brakes are applied only after all joints are brought to rest or for an emergency and should be left on whenever the arm is unattended. The switch positioned to off removes the brakes from all joints of the arm. The talkback indicator above the brakes switch on panel A8U indicates when the brakes are on or off.
The standard end effector can be considered the hand of the RMS. It is a hollow canlike device attached to the wrist roll joint at the end of the arm. Payloads to be captured by the standard end effector must be equipped with a grapple fixture. To capture a payload, the flight crew operator aligns the end effector over the grapple fixture probe to capture it. The end effector snare consists of three cables that have one end attached to a fixed ring and one attached to a rotating ring.
The end effector extend talkback indicator on panel A8U indicates gray when the end effector snare assembly is fully extended toward the opening of the canister. Barberpole indicates the end effector snares are somewhere between rigidize and derigidize.
\B\IFor more information click on \b\i\JShuttle Payload Deployment and Retrieval System (continued)\j
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"Shuttle Payload Deployment and Retrieval System (continued)",359,0,0,0
The payload is rigidized by drawing the snare assembly inside the end effector using a jackscrew, pulling the payload tightly against the face of the end effector and rigidizing the arm/payload assembly. During this process, current limit commands are sent to each joint motor to limp the arm, allowing the arm to move and compensate for misalignment errors. Wrist roll can still be commanded with a limp arm. The end effector rigid talkback indicator on panel A8U indicates gray when the end effector snare assembly is fully withdrawn in the end effector canister and the payload is rigidized. Barberpole indicates the end effector and payload are not rigidized.
There is one dual-end motor that produces all the motion in the end effector. The end effector \Jelectronics\j unit processes the end effector commands to produce the appropriate motor, clutch and brake commands from the displays and controls.
The rigidize sequence can be accomplished automatically or manually. The mode is selected by the flight crew operator with the end effector mode auto, off, man switch on panel A8U. Positioning the switch to auto causes the rigidize sequence to proceed automatically. If the switch is positioned to man , the end effector man contr switch on panel A8U must be positioned to rigid .
To release a payload, the snare mechanism moves outward until there is no force pulling the payload against the end effector, which is called derigidizing. If the end effector mode auto, off, man switch is in auto, lifting the RMS RHC switch guard and depressing the top half of the rocker switch commands a release. Derigidization automatically occurs, the snares of the end effector rotate open, and the payload grapple fixture is released. If the switch is positioned to man , the end effector man contr switch on panel A8U must be positioned to derigid .
The end effector rigid talkback indicator indicates barberpole when the end effector is no longer rigidized. The end effector derigid talkback indicator indicates gray when the end effector is derigidized. The end effector extend talkback indicator indicates gray when the end effector is fully extended. The end effector open talkback indicator indicates gray when the snares are fully open, releasing the payload, and the end effector close talkback indicator indicates barberpole.
The end effector is also equipped with a backup release capability that is controlled by the lever-locked payload release, off switch on panel A8U. The switch is only powered when the RMS power switch on panel A8L is positioned to backup . When the snares are closed, the snares wind up a spring device in the end effector. When the RMS power switch is in backup and the payload release switch is positioned to on, the snares are opened by the energy stored in the spring, releasing the payload. The end effector does not derigidize before releasing the payload. The payload release switch positioned to off de-energizes the circuit that opened the snares.
There are two types of automatic modes that can be used to position the RMS arm: preprogrammed and operator-commanded. The RMS software may be placed in the auto mode by positioning the mode rotary switch on panel A8U to auto 1, auto 2, auto 3, auto 4 or opr cmd and depressing the enter push button indicator on panel A8U. RMS joint rate commands are computed to drive the arm from its present position to a given point. (Point refers to a position and attitude of the point of resolution relative to the orbiter.) The RMS joint rates are computed so that the desired position and attitude are reached at the same time.
The operator-commanded automatic mode moves the end effector from its present position and orientation to a new one defined by the operator via the keyboard and RMS CRT display. After the data are keyed in, the operator must do a command check to verify that there is a set of joint angles that will put the arm at the desired point, but this command check does not verify the trajectory the arm must travel to get to that point (a straight line).
If the point is valid, good appears on the CRT; if not, fail appears. The mode is then entered by selecting opr cmd on the rotary mode switch, positioning the brakes switch to off , and depressing the enter push button indicator. The white ready light on panel A8U then is illuminated.
To start the arm moving to the desired point, the auto seq switch on panel A8U is positioned momentarily to proceed. The ready light is extinguished, and the white in prog light on panel A8U is illuminated. The arm will move in a straight line to the desired position and orientation, the in prog light will be extinguished, and the arm will then enter the hold mode. The RMS operator can stop and start the sequence through the auto seq proceed, stop switch on panel A8U.
The preprogrammed auto sequences operate in a manner similar to the operator-commanded sequences. Instead of the RMS operator entering the data on the computer via the keyboard and CRT display, the RMS arm is maneuvered according to sets programmed before the flight, called sequences. Up to 200 points may be preprogrammed into as many as 20 sequences. A given sequence is assigned via the CRT into auto 1, auto 2, auto 3 or auto 4. The mode is determined by then selecting auto 1, auto 2, auto 3 or auto 4 on the rotary mode switch and depressing the enter push button indicator. Each sequence is an ordered set of points to which the arm will move.
The preprogrammed sequences also consist of pause and fly-by. Pauses may be preprogrammed into the arm trajectory at any point that will cause the arm to come to rest. In order for the arm to proceed with the automatic sequence, the auto seq proceed, stop switch is positioned to proceed . (The operator can stop the arm at any place in the auto sequence by positioning the auto seq switch to stop .)
When the last point in the sequence is reached, the computer will terminate the movement of the arm and enter a position hold mode. The speed of the end effector between points in a sequence is governed by the individual joint rate limits set in the RMS software. In the fly-by sequence, the arm does not stop at a fly-by point; it continues to the next point in the sequence.
The single-joint drive control mode enables the operator to move the arm on a joint-by-joint basis with full computer support, thereby enabling full use of joint drive characteristics on a joint-by-joint basis. The operator places the rotary mode switch in the single position, depresses the enter push button indicator, and operates the arm by driving one joint at a time with the joint rotary switch on panel A8U and the single/direct drive switch on panel A8U.
In this mode, actuation of the single/direct drive switch removes the brakes from the joint that is selected by the joint rotary switch. The computer sends rate commands to the selected joint while holding position on the other joints. The single-joint drive mode is used to stow and unstow the arm and drive it out of joint travel limits.
Direct-drive control is a contingency mode. The direct mode is selected by positioning the rotary mode switch to direct and the brakes switch to on; the individual joints are driven with the rotary joint switch and the single/direct drive + or - switch. In the direct mode, the brakes remain on those joints not being driven, and the drive commands are hard-wired to the selected joint.
Direct drive bypasses the manipulator control interface unit, computer and data buses to send a direct command to the motor drive amplifier. The direct-drive mode is used when the MCIU or computer has a problem that necessitates arm control by the direct-drive mode to maneuver the loaded arm to a safe payload release position or to maneuver the unloaded arm to the storage position. Since this is a contingency mode, full joint performance characteristics are not available. Computer-supported displays may or may not be available, depending on the fault that necessitated the use of direct drive.
Backup drive control is a contingency mode to be used when the prime channel drive modes are not available. The backup is a degraded joint-by-joint drive system. The RMS software is in a suspend mode when backup is selected. The backup mode is selected by positioning the RMS power switch on panel A8L to backup . The arm is controlled using the backup control rotary switch on panel A8U and the single/direct drive + or - switch located below the rotary switch. The brakes are on the joints not being driven. The motors are driven bypassing the servoloop system. Since the MCIU has no power, there are no data from the arm.
Four RMS manually augmented modes are used to grapple a payload and maneuver it into or out of the orbiter payload retention fittings. The four manually augmented modes require the RMS operator to use the RMS translational hand controller and rotational hand controller with the computer to augment operations. The RMS takes up 32 percent of the fifth central processor unit for RMS operation and 30 percent for the manually augmented modes. The four manually augmented modes are controlled by the mode rotary switch on panel A8U. The modes are orbiter unloaded, end effector, orbiter loaded or payload. The coordinate system to which the motion is referred and point of resolution differ for each of these modes.
The THC and RHC located at the aft flight deck station are used exclusively for RMS operations. The THC is located between the two aft viewing windows. The RHC is located on the left side of the aft flight station below the CCTV monitors. The THC and RHC have only one output channel per axis. Both RMS hand controllers are proportional, which means that the command supplied is linearly proportional to the deflection of the controller.
The manually augmented mode enables the RMS operator to direct the end effector of the arm using two three-degree-of-freedom RMS hand controllers to control the end effector translation and rotation rates. The control algorithms process the hand controller signals into a rate for each joint.
When a manually augmented mode is selected, rate commands from the RMS THC result in motions at the tip of the end effector that are parallel to the orbiter-referenced coordinate frame and compatible with the up/down, left/right, in/out direction of the THC. Commands from the RMS RHC result in rotation at the tip of the end effector, which is also about the orbiter-referenced coordinate frame.
The manually augmented end effector mode maintains compatibility at all times among rate commands at the THC and RHC and the instantaneous orientation of the end effector. The end effector mode is used for grappling operations in conjunction with the RMS wrist-mounted CCTV camera, which is oriented with the end effector coordinates and rolls with the end effector.
The CCTV scene presented on the \Jtelevision\j monitor has viewing axes that are oriented with the end effector's coordinate frame. This results in compatible motion among the rate commands applied at the hand controllers and movement of the background image presented on the \Jtelevision\j monitor. Up/down, left/right and in/out motions of the THC result in the same direction of motion of the end effector as seen on the \Jtelevision\j monitor, except that the background in the scene will move in the opposite direction. Therefore, the operator must remember to use a fly-to control strategy and apply commands to the THC and RHC that are toward the target area in the \Jtelevision\j scene.
The manually augmented orbiter-loaded mode enables the operator to translate and rotate a payload about the orbiter axis with the point of resolution of the resolved rate \Jalgorithm\j at a predetermined point within the payload, normally the center of \Jgeometry\j. This allows for pure rotations of the payload for berthing operations. The manually augmented payload mode is analogous to the manually augmented end effector mode.
Safing and braking are the two methods available for bringing the arm to rest. Safing can be accomplished by positioning the safing switch on panel A8U to safe , which brings the arm to rest using the servocontrol loops. When the safing switch is positioned to auto , safing is initiated by the MCIU when certain critical built-in test equipment failures are detected. The cancel position of the safing switch removes the safing state. The safing talkback indicator indicates gray when safing is not in progress and barberpole when safing is in progress.
The RMS has a built-in test capability to detect and display critical failures. It monitors the arm-based \Jelectronics\j, displays and controls, and the MCIU software checks in the computer monitor computations. Failures are displayed on panel A8U and on the CRT and are also available for downlinking through orbiter telemetry.
All of the major systems of the ABE are monitored by BITE. The MCIU checks the integrity of the communications link among itself and ABE, displays and controls, and the orbiter computer. It also monitors end effector functions, thermistor circuit operation and its own internal consistency. The computer checks include an overall check of each joint's behavior through the consistency check, encoder data validity and end effector behavior as well as the proximity of the arm to reach limits, softstops and singularities.
The white auto 1, auto 2, auto 3, auto 4, opr cmd, test, orb unl, end eff, orb ld, payload, single and direct lights on panel A8U indicate the current RMS operating mode.
The software stop talkback indicator on panel A8U indicates gray when a stop has been commanded by the computer. Barberpole indicates a software stop has occurred, at least one joint has reached its limit of travel, and the computer has commanded arm motion to cease.
The rate meter on panel A8U reads in feet per second. Act indicates the translational speed, and cmd indicates the computer-commanded speed.
The rate min talkback indicator on panel A8U indicates on when the RHC vernier speed has been selected. Off indicates that the RHC coarse speed has been selected.
The rate hold talkback indicator on panel A8U indicates on when rate hold has been commanded and implemented by the computer. Off indicates that the rate hold function is not in effect. The rate hold function is engaged or disengaged by the rate hold button on the RHC.
If the manipulator arm cannot be restowed for any reason, it will be jettisoned so the payload bay doors can be closed. There are four separation points: one at the shoulder and one at each of the three retention latches. Each separation point is individually released. The switches for jettisoning the right or left RMS are located on panel A14.
An RMS jett deadface switch is located on panel A14. When the switch is positioned to deadface , the \Jelectronics\j of the three RMS retention latches are deadfaced. The safe position removes power from the deadface circuits and the ground reset circuits. The gnd reset position resets relays in the retention latches if the RMS was jettisoned. The relays are reset on the ground.
The pyro port or starboard RMS arm switches on panel A14 control the corresponding arm jettison functions. The jett arm, safe, \Jguillotine\j switch on panel A14 positioned to safe opens the corresponding RMS arm circuitry to disable the \Jguillotine\j and jettison operations. Positioning the switch to \Jguillotine\j closes the circuits, which arm the corresponding \Jguillotine\j circuits. The arm position enables power to the RMS jett switch.
The pyro port or starboard RMS jett switches on panel A14 control the corresponding arm shoulder jettison. The jett arm, safe, \Jguillotine\j switch on panel A14 positioned to safe opens the jettison logic circuitry for the corresponding arm.
Positioning the switch to \Jguillotine\j allows power to the \Jguillotine\j logic circuitry, guillotining the corresponding arm shoulder wire bundle (the corresponding RMS arm switch must be in the \Jguillotine\j position). The wire bundle is severed by a redundant pyro-operated \Jguillotine\j.
Positioning the switch to jett allows power to the jettison logic circuitry, jettisoning the corresponding arm shoulder (the corresponding RMS arm switch must be in the jett position). The separation system has redundant pyro-operated pressure cartridges to force a retractor down and pulls four overcenter tie-down hooks back, which releases the arm at the shoulder joint support.
The pyro starboard or port retention latches, fwd, mid, aft switches on panel A14 control the corresponding arm retention latches. Positioning the switches individually to the safe position opens the jettison logic circuitry for the corresponding retention latch. Positioning the switches to \Jguillotine\j individually allows power to \Jguillotine\j the corresponding retention latch.
Positioning the switches to jett individually allows power to the jettison circuitry, jettisoning the corresponding latch. The separation of the retention latches operates in a similar manner as the jettisoning of the shoulder joint. The separation system imparts a minimum impulse velocity on the RMS arm.
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"Shuttle Payload Retention Mechanisms",360,0,0,0
Non-deployable payloads are retained by passive retention devices, and deployable payloads are secured by motor-driven, active retention devices. Payloads are secured in the orbiter payload bay with the payload retention system or are equipped with their own unique retention systems.
The orbiter payload retention system provides three-axis support for up to five payloads per flight. The payload retention mechanisms secure the payloads during all mission phases and allow installation and removal of the payloads when the orbiter is either horizontal or vertical.
The longeron bridge fittings are attached to the payload bay frame at the longeron level and at the side of the bay. Keel bridge fittings are attached to the payload bay frame at the bottom of the payload bay.
The payload trunnions are the portion of the payload that interface with the orbiter retention system. The trunnions that interface with the longeron are 3.25 inches in diameter and 7 or 8.75 inches long, depending on their position in the payload bay. The keel trunnions are 3 inches in diameter and vary in length from 4 to 11.5 inches, depending on where they fit in the payload bay.
The orbiter and payload attachments are the trunnion/bearing/journal type. The longeron and keel attach fittings have a split, self-aligning bearing for non-release-type payloads in which the hinged half is bolted closed. For on-orbit deployment and retrieval payloads, the hinged half fitting releases or secures the payload by latches that are driven by dual-redundant electric motors.
Payload guides and scuff plates assist in deploying and berthing payloads in the payload bay. The payload is constrained in the X direction by guides and in the Y direction by scuff plates and guides. The guides are mounted to the inboard side of the payload latches and interface with the payload trunnions and scuff plates. The scuff plates are attached to the payload trunnions and interface with the payload guides.
The guides are V-shaped with one part of the V being 2 inches taller than the other part. Parts are available to make either the forward or aft guide taller.
This difference enables the operator monitoring the berthing or deployment operations through the aft bulkhead CCTV cameras to better determine when the payload trunnion has entered the guide. The top of the taller portion of the guide is 24 inches above the centerline of the payload trunnion when it is all the way down in the guide. The top of the guide has a 9-inch opening. These guides are mounted to the 8-inch guides that are a part of the longeron payload retention latches.
A set of payload active retention latches may consist of as many as five latches per payload. Three payloads can be accommodated with active latches. Each of the active latches is controlled by dual-redundant ac electric motors that release or latch the active retention latch. The active retention latches are controlled from panel A6U.
When the payload retention logic power system 1 switch on panel A6U is positioned to on, it provides main bus power to the rotary payload select switch on panel A6U. The system 2 switch, when positioned to on, provides MNB bus power to the rotary payload select switch for payloads 1, 2 and 3.
Positioning the payload select rotary switch on panel A6U to 1 provides power-on logic for the dual actuator motors of up to five latches for one payload and the talkback indications associated with up to five latches for the payload. Position 2 of the payload select switch provides power-on logic for the dual actuator motors of up to five latches for the second payload and the talkback indications associated with up to five latches for the payload. Position 3 provides power-on logic for the dual actuator motors of up to five latches for the third payload and the talkback indications associated with up to five latches.
The payload retention latches 1, 2, 3, 4 and 5 switches on panel A6U are enabled by the payload select rotary switch. Positioning the payload select switch to 1 enables up to five retention latches for payload 1, and each of the five retention latches for payload 1 would be controlled by the individual 1, 2, 3, 4 and 5 release, off, latch switches. Positioning the payload select switch to 2 or 3 has the same effect for payloads 2 and 3.
Positioning a payload retention latches switch to release provides ac power to the dual electric motors associated with the retention latch of the selected payload, driving the retention latch open. The operating time of the latch with both motors operating is 30 seconds; with only one motor operating it is 60 seconds. The talkback indicator immediately above a retention latches switch indicates rel when the latch is fully open. There are two microswitches for the rel talkback indication; however, only one is required to control the talkback indicator. The payload retention latches ready for latch talkback indicator for a retention latches switch is barberpole when the payload latch is set in the release position. There are two microswitches for the ready-for-latch talkback indication; however, only one is required to control the talkback indicator.
Positioning a payload retention latches switch to latch provides ac power to the dual electric motor associated with the latch of the payload selected, driving the retention latch closed. The operating time of one or both motors is the same as for releasing a payload. A barberpole talkback indicator immediately above each retention latches switch indicates that latch is ready to latch.
The indicator shows lat when the latch is closed. There are two microswitches for the lat indication; however, only one is required to control the talkback indicator. The payload retention latches ready for latch talkback indicator for a retention latches switch is gray when the payload latch is ready to latch.
Positioning the payload select rotary switch to monitor inhibits the logic circuits of all payload actuator latch sets and inhibits the talkback indicators but provides power for payload latch telemetry.
The keel active latch centers the payload in the yaw direction in the payload bay; therefore, the keel latch must be closed before the longeron latches are closed. The keel latch can float plus or minus 2.75 inches in the X direction.
#
"Space Flight Tracking and Data Network",361,0,0,0
The Networks Division of Goddard Space Flight Center, Greenbelt, Md., is responsible for operating, maintaining and controlling the space flight tracking and data network, which consists of the space network and ground network for providing tracking, data acquisition and associated support. The network is operated through NASA contracts and interagency and international agreements that provide staffing and logistic support for space missions.
The Networks Division also operates the Network Control Center and NASA Ground Terminal. The division is responsible for testing, \Jcalibration\j and configuring network resources to ensure network support capability before each mission. It coordinates, schedules and directs all network activity and provides the necessary interface among GSFC elements and other agencies, centers and networks.
The STDN, controlled by the NCC at Goddard, is composed of the White Sands Ground Terminal and NASA Ground Terminal in White Sands, N.M.; the NASA Communications Network, Flight Dynamics Facility and Simulation Operations Center at GSFC; and the ground network.
These elements are linked by voice and data communication services provided by Nascom. The prime operational communications data are formatted into 4,800-bit blocks and transmitted on the Nascom wide-band data and message switching system. Other communications are transmitted by teletype and facsimile facilities.
The Tracking and Data Relay Satellite system will consist of two Tracking and Data Relay satellites in geosynchronous orbit (130 degrees apart in longitude), an on-orbit spare, and a ground terminal facility (located at White Sands). The TDRS can transmit and receive data and track a user \Jspacecraft\j in a low Earth orbit for a minimum of 85 percent of its orbit. TDRSS telecommunication services to and from the user's control and data processing facilities operate in a real-time, bent-pipe mode.
The White Sands Ground Terminal contains the ground terminal communications relay equipment for the command, telemetry, tracking and control equipment of the TDRSS. The NASA Ground Terminal is collocated with WSGT. The NGT is managed and operated by the Networks Division and, in combination with Nascom, is NASA's physical and electrical interface with the TDRSS. The NGT provides the interfaces with the common carrier, monitors the quality of the service from the TDRSS, and remotes data quality to the NCC.
Goddard's ground tracking stations for various communications are located throughout the world:
- Ascension Island (ACN)-S-band and ultrahigh frequency air-to-ground.
- Bermuda (BDA)-S-band, C-band and UHF air-to- ground.
- Guam (GWM)-S-band and UHF air-to- ground.
- Kauai, Hawaii (HAW)-S-band and UHF air-to- ground.
- Merritt Island, Fla. (MIL)-S-band and UHF air-to- ground.
- Santiago, \JChile\j (AGO)- S-band.
- Ponce de Leon, Fla. (PDL)- S-band.
- Canberra, \JAustralia\j (CAN)- S-band.
- Dakar, Senegal-UHF air-to- ground.
- Wallops, Va. (WFF)- C-band.
Also supporting the STDN are several instrumented United States Air Force \Jaircraft\j, referred to as advanced range instrumentation \Jaircraft\j, that are situated upon request at various locations around the world where ground stations cannot support space shuttle missions.
The various antennas at each STDN site accomplish a specific task, usually in a specific frequency band. Functioning like giant electronic magnifying glasses, the larger antennas absorb radiated electronic signals transmitted by \Jspacecraft\j in a radio form called telemetry.
This communication network is composed of \Jtelephone\j, microwave, radio, submarine cables and communication satellites. These various systems link the data flow through 11 countries of the free world with 15 foreign and domestic carriers and provide the required information between tracking sites and Johnson Space Center (Houston, Texas) and Goddard control centers. Special wide-band and video circuitry is used as needed. GSFC has the largest wide-band system in existence.
Included in the equipment of the worldwide STDN are numerous computers located at the different stations that control tracking antennas, handle commands and process data for transmission to the JSC and GSFC control centers. Shuttle data from all the tracking stations are funneled into the main switching computers at GSFC and rerouted to JSC without delay by domestic communications satellites. Commands generated at JSC are transmitted to the main switching computers at GSFC and switched to the proper tracking station for transmission to the space shuttle.
If NASA's JSC Mission Control Center should be impaired for an extended period of time, an emergency control center would be established at NASA's ground terminal at White Sands and manned by NASA JSC personnel.
A station conferencing and monitoring arrangement allows various traffic managers to hold conferences with as many as 220 different voice terminals throughout the United States and abroad with talking and listening capability at the touch of a few buttons. The system is redundant, which accounts for its mission support reliability record of 99.6 percent. All space shuttle voice traffic is routed through this arrangement at GSFC.
Communication satellites electronically connect the Earth stations and permit transmission of 10 to 20 times more data. Ground terminals for domestic communications satellites are situated at JSC; \JKauai\j, Hawaii; Goldstone, Calif.; Kennedy Space Center. \JFlorida\j; NASA's Dryden Flight Research Facility, \JCalifornia\j; GSFC, Greenbelt, Md.; and White Sands, N.M.
The tracking station at Ponce de Leon Inlet, Fla. (near New \JSmyrna\j Beach), provides support during powered flight because of attenuation problems from the solid rocket booster motor plume.
The existing worldwide ground stations provide coverage for approximately 20 percent of a satellite's or \Jspacecraft\j's orbit, limited to brief periods when the satellite or \Jspacecraft\j is within the line of sight of a given tracking station.
A new era in space communication began with the STS-6 mission in April 1983, when the first Tracking and Data Relay Satellite was deployed. TDRS-A was the first of three identical satellites planned for the system. The TDRS system was developed after studies in the early 1970s showed that a telecommunication satellite system could support the projected scientific and application mission requirements better than ground stations and also could halt the spiraling cost of upgrading and operating a worldwide network of tracking and communication ground stations.
#
"Satellite System, Tracking and Data Relay",362,0,0,0
When fully operational, the TDRSS will provide continuous global coverage of Earth-orbiting satellites at altitudes from 750 miles to about 3,100 miles. At lower altitudes, there will be brief periods when satellites or \Jspacecraft\j over the Indian Ocean near the equator are out of view. The TDRSS will be able to handle up to 300 million bits of information per second. Because eight bits of information make one word, this capability is equivalent to processing 300 14-volume sets of encyclopedias every minute.
The fully operational TDRSS network will consist of three satellites in geosynchronous orbits. The first, positioned at 41 degrees west longitude, is TDRS-East (TDRS-A). The next satellite, TDRS-West, will be carried into Earth orbit aboard the space shuttle and deployed and positioned at 171 degrees west longitude. The remaining TDRS will be positioned above a central station just west of South America at 62 degrees west longitude as a backup.
The satellites are positioned in geosynchronous orbits above the equator at an altitude of 22,300 statute miles. At this altitude, because the speed of the satellite is the same as the rotational speed of Earth, it remains fixed in orbit over one location. The eventual positioning of two TDRSs will be 130 degrees apart instead of the usual 180-degree spacing. This 130-degree spacing will reduce the ground station requirements to one station instead of the two stations required for 180-degree spacing.
The TDRS system serves as a radio data relay, carrying voice, \Jtelevision\j, and analog and digital data signals. It offers three frequency band services: S-band, C-band and high-capacity Ku-band. The C-band transponders operate at 4 to 6 GHz and the Ku-band transponders operate at 12 to 14 GHz.
The highly automated TDRSS network ground station, located at the White Sands Ground Terminal, is owned and managed by Contel.
TDRSS also provides communication and tracking services for low Earth-orbiting satellites. It measures two-way range and Doppler for up to nine user satellites and one-way and Doppler for up to 10 user satellites simultaneously. These measurements are relayed to the Flight Dynamics Facility at GSFC from the WSGT.
Six TDRSs will be built by TRW's Defense and Space Systems Group, Redondo Beach, Calif. Contel owns and operates the satellites and the White Sands Ground Terminal, which was built jointly by the team of TRW, Harris Corporation and Spacecom. Electronic hardware was jointly supplied by TRW and Harris's Government Communications Division, Melbourne, Fla. TRW integrated and tested the ground station, developed software for the TDRS system and integrated the hardware with the ground station and satellites.
The ground station is located at a longitude with a clear line of sight to the TDRSs and very little rain, because rain can interfere with the Ku-band uplink and downlink channels. It is one of the largest and most complex communication terminals ever built.
The most prominent features of the ground station are three 60-foot Ku-band dish antennas used to transmit and receive user traffic. Several other antennas are used for S-band and Ku-band communications. NASA developed sophisticated operational control facilities at GSFC and next to the WSGT to schedule TDRSS support of each user and to distribute the user's data from White Sands to the user.
Automatic data processing equipment at the WSGT aids in satellite tracking measurements, control and communications. Equipment in the TDRS and the ground station collects system status data for transmission, along with user \Jspacecraft\j data, to NASA. The ground station software and computer component, with more than 900,000 machine language instructions, will eventually control three geosynchronous TDRSs and the 300 racks of ground station electronic equipment.
Many command and control functions ordinarily found in the space segment of a system are performed by the ground station, such as the formation and control of the receive beam of the TDRS multiple-access phased-array antenna and the control and tracking functions of the TDRS single-access antennas.
Data acquired by the satellites are relayed to the ground terminal facilities at White Sands. White Sands sends the raw data directly by domestic communications satellite to NASA control centers at JSC (for space shuttle operations) and GSFC, which schedules TDRSS operations and controls a large number of satellites. To increase system reliability and availability, no signal processing is done aboard the TDRSs; instead, they act as repeaters, relaying signals to and from the ground station or to and from satellites or \Jspacecraft\j. No user signal processing is done aboard the TDRSs.
A second TDRS ground terminal is being built at White Sands approximately 3 miles north of the initial ground station. The $18.5-million facility will back up the existing facility and meet the growing communication needs of the 1990s.
When the TDRSS is fully operational, ground stations of the worldwide STDN will be closed or consolidated, resulting in savings in personnel and operating and maintenance costs. However, the Merritt Island, Fla.; Ponce de Leon, Fla.; and \JBermuda\j ground stations will remain open to support the launch of the space transportation system and the landing of the space shuttle at the Kennedy Space Center in \JFlorida\j.
Deep-space probes and Earth-orbiting satellites above approximately 3,100 miles will use the three ground stations of the deep-space network, operated for NASA by the Jet Propulsion Laboratory, Pasadena, Calif. The deep-space network stations are in Goldstone, Calif.; Madrid, \JSpain\j; and \JCanberra\j, \JAustralia\j.
During the lift-off and ascent phase of a space shuttle mission launched from the Kennedy Space Center. the space shuttle S-band system is used in a high-data-rate mode to transmit and receive through the Merritt Island, Ponce de Leon and \JBermuda\j STDN tracking stations. When the shuttle leaves the line-of-sight tracking station at \JBermuda\j, its S-band system transmits and receives through the TDRSS. (There are two communication systems used in communicating between the space shuttle and the ground. One is referred to as the S-band system; the other, the Ku-band, or K-band, system.)
To date, the TDRSs are the largest privately owned telecommunication satellites ever built. Each satellite weighs nearly 5,000 pounds in orbit. The TDRSs will be deployed from the space shuttle at an altitude of approximately 160 nautical miles, and inertial upper stage boosters will propel them to geosynchronous orbit.
The TDRS single-access parabolic antennas deploy after the satellite separates from the IUS. After the TDRS acquires the sun and Earth, its sensors provide attitude and velocity control to achieve the final geostationary position.
Three-axis stabilization aboard the TDRS maintains attitude control. Body-fixed momentum wheels in a vee configuration combine with body-fixed antennas pointing constantly at Earth, while the satellite's solar arrays track the sun. Monopropellant hydrazine thrusters are used for TDRS positioning and north-south, east-west stationkeeping.
The antenna module houses four antennas. For single-access services, each TDRS has two dual-feed S-band / Ku-band deployable parabolic antennas. They are 16 feet in diameter, unfurl like a giant umbrella when deployed, and are attached on two axes that can move horizontally or vertically (gimbal) to focus the beam on satellites or \Jspacecraft\j below. Their primary function is to relay communications to and from user satellites or \Jspacecraft\j. The high-bit-rate service made possible by these antennas is available to users on a time-shared basis. Each antenna simultaneously supports two user satellites or \Jspacecraft\j (one on S-band and one on Ku-band) if both users are within the antenna's bandwidth.
The antenna's primary reflector surface is a gold-clad molybdenum wire mesh, woven like cloth on the same type of machine used to make material for women's hosiery. When deployed, the antenna's 203 square feet of mesh are stretched tautly on 16 sup porting tubular ribs by fine threadlike \Jquartz\j cords. The antenna looks like a glittering metallic spiderweb. The entire antenna structure, including the ribs, reflector surface, a dual-frequency antenna feed and the deployment mechanisms needed to fold and unfold the structure like a parasol, weighs approximately 50 pounds.
For multiple-access service, the multielement S-band phased array of 30 helix antennas on each satellite is mounted on the satellite's body. The multiple-access forward link (between the TDRS and the user satellite or spacecraft) transmits command data to the user satellite or \Jspacecraft\j, and the return link sends the signal outputs separately from the array elements to the WSGT's parallel processors. Signals from each helix antenna are received at the same frequency, frequency-division-multiplexed into a single composite signal and transmitted to the ground. In the ground equipment, the signal is demultiplexed and distributed to 20 sets of beam-forming equipment that discriminates among the 30 signals to select the signals of individual users. The multiple-access system uses 12 of the 30 helix antennas on each TDRS to form a transmit beam.
A 6.6-foot parabolic reflector is the space-to-ground-link antenna that communicates all data and tracking information to and from the ground terminal on Ku-band. The omni telemetry, tracking and communication antenna is used to control TDRS while it is in transfer orbit to geosynchronous altitude.
The solar arrays on each satellite, when deployed, span more than 57 feet from tip to tip. The two single-access, high-gain parabolic antennas, when deployed, measure 16 feet in diameter and span 42 feet from tip to tip.
Each TDRS is composed of three distinct modules: the equipment module, the communication payload module and the antenna module. The modular structure reduces the cost of individual design and construction.
The equipment module housing the subsystems that operate the satellite and the communication service is located in the lower hexagon of the satellite. The attitude control subsystem stabilizes the satellite so that the antennas are properly oriented toward the Earth and the solar panels are facing toward the sun. The electrical power subsystem consists of two solar panels that provide approximately 1,850 watts of power for 10 years. Nickel-cadmium rechargeable batteries supply full power when the satellite is in the shadow of the Earth. The thermal control subsystem consists of surface coatings and controlled electric heaters. The solar sail compensates for the effects of solar winds against the asymmetrical body of the TDRS.
The communication payload module on each satellite contains electronic equipment and associated antennas required for linking the user \Jspacecraft\j or satellite with the ground terminal. The receivers and transmitters are mounted in compartments on the back of the single-access antennas to reduce complexity and possible circuit losses.
TDRS-A and its IUS were carried aboard the space shuttle Challenger on the April 1983 STS-6 mission. After it was deployed on April 4, 1983, and first-stage boost of the IUS solid rocket motor was completed, the second-stage IUS motor malfunctioned and TDRS-A was left in an egg-shaped orbit 13,579 by 21,980 statute miles-far short of the planned 22,300-mile geosynchronous altitude. Also, TDRS-A was spinning out of control at a rate of 30 revolutions per minute until the Contel/TRW flight control team recovered control and stabilized it.
Later Contel, TRW and NASA TDRS program officials devised a procedure for using the small (1-pound) hydrazine-fueled reaction control system thrusters on TDRS-A to raise its orbit. The thrusting, which began on June 6, 1983, required 39 maneuvers to raise TDRS-A to geosynchronous orbit. The maneuvers consumed approximately 900 pounds of the satellite's propellant, leaving approximately 500 pounds of hydrazine for the 10-year on-orbit operations.
During the maneuvers, overheating caused the loss of one of the redundant banks of 12 thrusters and one thruster in the other bank. The flight control team developed procedures to control TDRS-A properly in spite of the thruster failures.
TDRS-A was turned on for testing on July 6, 1983. Tests proceeded without incident until October 1983, when one of the Ku-band single-access-link diplexers failed. Shortly afterward, one of the Ku-band traveling-wave-tube amplifiers on the same single-access antenna failed, and the forward link service was lost. On November 19, 1983, one of the Ku-band TWT amplifiers serving the other single-access antenna failed. TDRS-A testing was completed in December 1984. Although the satellite can provide only one Ku-band single-access forward link, it is still functioning.
TDRS-B, C and D are identical to TDRS-A except for modifications to correct the malfunctions that occurred in TDRS-A and a modification of the C-band antenna feeds. The C-band minor modification was made to improve coverage for providing government point-to-point communications. TDRS-B was lost on the 51-L mission.
The mission plan for TDRS-C is similar to that originally planned for TDRS-A. Backup project operations control centers have been added at TRW and at the TDRS Launch/Deployment Control Center in White Sands. These facilities will improve the reliability of control operations and the simultaneous control of TDRS-A in mission support and of TDRS-C during launch and deployment operations.
TDRS-C and its IUS are to be deployed from the space shuttle orbiter. Approximately 60 minutes later, the IUS first-stage solid rocket motor is scheduled to ignite. This will be followed by five maneuvers to allow monitoring of TDRS-C telemetry.
After the IUS second-stage thrusting is completed, the TDRSS mission team at White Sands will command deployment of the TDRS-C solar arrays, the space-ground link antenna and the C-band antenna while the TDRS is still attached to the IUS. Upon separation of the IUS from TDRS-C, the 16-foot-diameter single-access antennas will be deployed, unfurled and oriented toward Earth. Nominal deployment will place TDRS-C at 178 degrees west longitude.
Testing of TDRS-C will be initiated; and after initial checkout, TDRS-C will drift westward to its operational location at 171 degrees west longitude, southwest of Hawaii, where it will be referred to as TDRS-West. Operational testing will continue to verify the full-system capability with two operating satellites. On completion of this testing, about three to five months after the launch of TDRS-C, the TDRSS, for the first time, will provide its full-coverage capability in support of NASA space missions.
TDRS-D, identical to TDRS-C, will take the place of TDRS-A at 41 degrees west longitude above the equator, over the northeast corner of \JBrazil\j, and will be referred to as TDRS-East. TDRS-A will then be relocated, probably 79 degrees west longitude above the equator, over central South America, and will be maintained as an on-orbit spare.
These three satellites will make up the space segment of the TDRS system. The on-orbit spare, available for use if one of the operational satellites malfunctions, will augment system capabilities during peak periods. The two remaining satellites will be available as flight-ready spares.
The failure of TDRS-A's Ku-band forward link prohibits the operation of the text and graphics system that it is desired be placed on board all space shuttle orbiters. TAGS is a high-resolution facsimile system that scans text or graphic material and converts the analog scan data into serial digital data. It provides on-orbit capability to transmit text material, maps, schematics and photographs to the \Jspacecraft\j through a two-way Ku-band link through the TDRSS. This is basically a hard-copy machine that operates by telemetry.
Until there is a dual TDRS capability, a teleprinter must be used on orbit to receive and reproduce text only (such as procedures, weather data and crew activity plan updates or changes) from the Mission Control Center. The teleprinter uses S-band and is not dependent on the TDRSS Ku-band.
When the space shuttle orbiter is on orbit and its payload bay doors are opened, the space shuttle orbiter Ku-band antenna, stowed on the right side of the forward portion of the payload bay, is deployed. One drawback of the Ku-band system is its narrow pencil beam, which makes it difficult for the TDRS antennas to lock on to the signal. Because the S-band system has a larger beamwidth, the orbiter uses it first to lock the Ku-band antenna into position. Once this has occurred, the Ku-band signal is turned on.
The Ku-band system provides a much higher gain signal with a smaller antenna than the S-band system. The orbiter's Ku-band antenna is gimbaled so that it can acquire the TDRS. Upon communication acquisition, if the TDRS is not detected within the first 8 degrees of spiral conical scan, the search is automatically expanded to 20 degrees. The entire TDRS search requires approximately three minutes. The scanning stops when an increase in the received signal is sensed. The orbiter Ku-band system and antenna then transmits and receives through the TDRS in view.
At times, the orbiter may block its Ku-band antenna's view to the TDRS because of attitude requirements or certain payloads that cannot withstand Ku-band radiation from the main beam of the orbiter's antenna. The main beam of the Ku-band antenna produces 340 volts per meter, which decreases in distance from the antenna-e.g., 200 volts per meter 65 feet away from the antenna. A program can be instituted in the orbiter's Ku-band antenna control system to limit the \Jazimuth\j and elevation angle, which inhibits direction of the beam toward areas of certain onboard payloads. This area is referred to as an obscuration zone. In other cases, such as deployment of a satellite from the orbiter payload bay, the Ku-band system is turned off temporarily.
When the orbital mission is completed, the orbiter's payload bay doors must be closed for entry; therefore, its Ku-band antenna must be stowed. If the antenna cannot be stowed, provisions are incorporated to jettison the assembly from the \Jspacecraft\j so that the payload bay doors can be closed for entry. The orbiter can then transmit and receive through the S-band system, the TDRS in view and the TDRS system. After the communications blackout during entry, the space shuttle again operates in S-band through the TDRS system in the low- or high-data-rate mode as long as it can view the TDRS until it reaches the S-band landing site ground station.
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"Shuttle Avionics Systems",363,0,0,0
The space shuttle avionics system controls, or assists in controlling, most of the shuttle systems. Its functions include automatic determination of the vehicle's status and operational readiness; implementation sequencing and control for the solid rocket boosters and external tank during launch and ascent; performance monitoring; digital data processing; communications and tracking; payload and system management; guidance, navigation and control; and electrical power distribution for the orbiter, external tank and solid rocket boosters.
Automatic vehicle flight control can be used for every phase of the mission except docking, which is a manual operation performed by the flight crew. Manual control-referred to as the control stick steering mode-also is available at all times as a flight crew option.
The avionics equipment is arranged to facilitate checkout, access and replacement with minimal disturbance to other systems. Almost all electrical and electronic equipment is installed in three areas of the orbiter: the flight deck, the three avionics equipment bays in the middeck of the orbiter crew compartment and the three avionics equipment bays in the orbiter aft fuselage. The flight deck of the orbiter crew compartment is the center of avionics activity, both in flight and on the ground. Before launch, the orbiter avionics system is linked to ground support equipment through umbilical connections.
The space shuttle avionics system consists of more than 300 major electronic black boxes located throughout the vehicle, connected by more than 300 miles of electrical wiring. There are approximately 120,400 wire segments and 6,491 connectors in the vehicle. The wiring and connectors weigh approximately 7,000 pounds, wiring alone weighing approximately 4,600 pounds. Total weight of the black boxes, wiring and connectors is approximately 17,116 pounds.
The black boxes are connected to a set of five general-purpose computers through common party lines called data buses. The black boxes offer dual or triple redundancy for every function.
The avionics are designed to withstand multiple failures through redundant hardware and software (computer programs) managed by the complex of five computers; this arrangement is called a fail-operational/fail-safe capability. Fail-operational performance means that, after one failure in a system, redundancy management allows the vehicle to continue on its mission. Fail-safe means that after a second failure, the vehicle still is capable of returning to a landing site safely.
#
"Shuttle Data Processing System",364,0,0,0
The space shuttle vehicle relies on computerized control and monitoring for successful performance. The data processing system, through the use of various hardware components and its self-contained computer programming (software), provides the vehicle with this monitoring and control.
The DPS hardware consists of five general-purpose computers for computation and control, two magnetic tape mass memory units for large-volume bulk storage, a time-shared computer data bus network consisting of serial digital data buses (essentially party lines) to accommodate the data traffic between the GPCs and space shuttle vehicle systems, 19 orbiter and four solid rocket booster multiplexers/demultiplexers to convert and format data from the various vehicle systems, three space shuttle main engine interface units to command the SSMEs, four multifunction CRT display systems used by the flight crew to monitor and control the vehicle and payload systems, two data bus isolation amplifiers to interface with the ground support equipment/launch processing system and the solid rocket boosters, two master events controllers, and a master timing unit.
The software stored in and executed by the GPCs is the most sophisticated and complex set of programs ever developed for aero space use. The programs are written to accommodate almost every aspect of space shuttle operations, including orbiter checkout at Rockwell's Palmdale, Calif., assembly facility; space shuttle vehicle prelaunch and final countdown for launch; turnaround activities at the Kennedy Space Center and eventually Vandenberg Air Force Base; control and monitoring during launch ascent, on-orbit activities, entry and landing; and aborts or other contingency mission phases. A multicomputer mode is used for the critical phases of the mission, such as launch, ascent, entry, landing and aborts.
Some of the DPS functions are as follows: support the guidance, navigation and control of the vehicle, including calculations of trajectories, SSME thrusting data and vehicle attitude control data; process vehicle data for the flight crew and for transmission to the ground and allow ground control of some vehicle systems via transmitted commands; check data transmission errors and crew control input errors; support \Jannunciation\j of vehicle system failures and out-of-tolerance system conditions; support payloads with flight crew/software interface for activation, deployment, deactivation and retrieval; process rendezvous, tracking and data transmissions between payloads and the ground; and monitor and control vehicle subsystems.
\BSoftware\b
DPS software is divided into two major groups, system software and applications software. The two software program groups are combined to form a memory configuration for a specific mission phase. The software programs are written in HAL/S (high-order assembly language/shuttle) especially developed for real-time space flight applications.
The system software is the GPC operating software that controls the interfaces among the computers and the rest of the DPS. It is loaded into the computer when it is first initialized. It always resides in the GPC main memory and is common to all memory configurations. The system software controls the GPC input and output, loads new memory configurations, keeps time, monitors discretes into the GPCs and performs many other functions required for the DPS to operate. The system software has nothing to do with orbiter systems or systems management software.
The system software consists of three sets of programs: the flight computer operating program (the executive) that controls the processors, monitors key system parameters, allocates computer resources, provides for orderly program interrupts for higher priority activities and updates computer memory; the user interface programs that provide instructions for processing flight crew commands or requests; and the system control program that initializes each GPC and arranges for multi-GPC operation during flight-critical phases. The system software program tells the general-purpose computers how to perform and how to communicate with other equipment.
One of the system software responsibilities is to manage the GPC input and output operations, which includes assigning computers as commanders and listeners on the data buses and exercising the logic involved in sending commands to these data buses at specified rates and upon request from the applications software.
The applications software contains (1) specific software programs for vehicle guidance, navigation and control required for launch, ascent to orbit, maneuvering in orbit, entry and landing on a runway; (2) systems management programs with instructions for loading memories in the space shuttle main engine computers and for checking the vehicle instrumentation system, aiding in vehicle subsystem checkout, ascertaining that flight crew displays and controls perform properly and updating inertial measurement unit state vectors; (3) payload processing programs with instructions for controlling and monitoring orbiter payload systems that can be revised depending on the nature of the payload; and (4) vehicle checkout programs needed to handle data management, performance monitoring, special processing, and display and control processing.
The applications software performs the actual duties required to fly and operate the vehicle. To conserve main memory, the applications software is divided into three major functions: guidance, navigation and control; systems management; and payload. Each GPC operates in one major function at a time, and usually more than one computer is in the GN&C major function simultaneously for redundancy.
The highest level of the applications software is the operational sequence required to perform part of a mission phase. Each OPS is a set of unique software that must be loaded separately into a GPC from the mass memory units. Therefore, all the software residing in a GPC at any time consists of system software and an OPS. An OPS can be further subdivided into groups called major modes, each representing a portion of the OPS mission phase.
During the transition from one OPS to another, the flight crew requests a new set of applications software to be loaded in from the MMU. Every OPS transition is initiated by the flight crew. An exception is GN&C OPS 1, which is divided into six major modes and contains the OPS 6 return-to-launch-site abort, since there would not be time to load in new software for an RTLS. When an OPS transition is requested, the redundant OPS overlay contains all major modes of that sequence.
Each major mode has with it an associated CRT display, called an OPS display, that provides the flight crew with information concerning the current portion of the mission phase and allows flight crew interaction. There are three levels of CRT displays. Certain portions of each OPS display can be manipulated by flight crew keyboard input (or ground link) to view and modify system parameters and enter data. The specialist function of the OPS software is a block of displays associated with one or more operational sequences and enabled by the flight crew to monitor and modify system parameters through keyboard entries. The display function of the OPS software is a block of displays associated with one OPS or more. These displays are for parameter monitoring only (no modification capability) and are called from the keyboard.
The principal software used to operate the vehicle during a mission is the primary avionics software system. It contains all the programming needed to fly the vehicle through all phases of the mission and manage all vehicle and payload systems.
Since the ascent and entry phases of flight are so critical, four of the five GPCs are loaded with the same PASS software and perform all GN&C functions simultaneously and redundantly. As a safety measure, the fifth GPC contains a different set of software, programmed by a company different from the PASS developer, designed to take control of the vehicle if a generic error in the PASS software or other multiple errors should cause a loss of vehicle control. This software is called the backup flight system. In the less dynamic phases of on-orbit operations, the BFS is not required.
GPCs running together in the same GN&C OPS are part of a redundant set performing identical tasks from the same inputs and producing identical outputs. Therefore, any data bus assigned to a commanding GN&C GPC is heard by all members of the redundant set (except the instrumentation buses because each GPC has only one dedicated bus connected to it).
These transmissions include all CRT inputs and mass memory transactions, as well as flight-critical data. Thus, if one or more GPCs in the redundant set fail, the remaining computers can continue operating in GN&C. Each GPC performs about 325,000 operations per second during critical phases.
Each computer in a redundant set operates in synchronized steps and cross-checks results of processing about 440 times per second. Synchronization refers to the software scheme used to ensure simultaneous intercomputer communications of necessary GPC status information among the primary avionics computers.
If a GPC operating in a redundant set fails to meet two redundant synchronization codes in a row, the remaining computers will vote it out of the redundant set. Or if a GPC has a problem with its multiplexer interface adapter receiver during two successive reads of response data and does not receive any data while the other members of the redundant set do not receive the data, they in turn will vote the GPC out of the set. A failed GPC is halted as soon as possible.
GPC failure votes are annunciated in a number of ways. The GPC status matrix on panel O1 is a 5-by-5 matrix of lights. For example, if GPC 2 sends out a failure vote against GPC 3, the second white light in the third column is illuminated.
The yellow diagonal lights from upper left to lower right are self-failure votes. Whenever a GPC receives two or more failure votes from other GPCs, it illuminates its own yellow light and resets any failure votes that it made against other GPCs (any white lights in its row are extinguished). Any time a yellow matrix light is illuminated, the GPC red caution and warning light on panel F7 is illuminated, in addition to master alarm illumination, and a GPC fault message is displayed on the CRT.
\BGPCs\b
Five identical general-purpose computers aboard the orbiter control space shuttle vehicle systems. Each GPC is composed of two separate units, a central processor unit and an input/output processor. All five GPCs are \JIBM\j AP-101 computers. Each CPU and IOP contains a memory area for storing software and data. These memory areas are collectively referred to as the GPC's main memory.
The central processor controls access to GPC main memory for data storage and software execution and executes instructions to control vehicle systems and manipulate data. In other words, the CPU is the ''number cruncher'' that computes and controls computer functions.
The IOP formats and transmits commands to the vehicle systems, receives and validates response data from the vehicle systems and maintains the status of interfaces with the CPU and the other GPCs.
The IOP of each computer has 24 independent processors, each of which controls 24 data buses used to transmit serial digital data between the GPCs and vehicle systems, and secondary channels between the telemetry system and units that collect instrumentation data. The 24 data buses are connected to each IOP by multiplexer interface adapters that receive, convert and validate the serial data in response to discrete signals calling for available data to be transmitted or received from vehicle hardware.
During the receive mode, the multiplexer interface adapter validates the received data (notifying the IOP control logic when an error is detected) and reformats the data. During the receive mode, its transmitter is inhibited unless that particular GPC is in command of that data bus.
During the transmit mode, a multiplexer interface adapter transmits and receives 28-bit command/data words over the computer data buses. When transmitting, the MIA adds the appropriate parity and synchronization code bits to the data, reformats the data, and sends the information out over the data bus. In this mode, the MIA's receiver and transmitters are enabled.
The first three bits of the 28-bit word provide synchronization and indicate whether the information is a command or data. The next five bits identify the destination or source of the information. For command words, 19 bits identify the data transfer or operations to be performed; for data words, 16 of the 19 bits contain the data and three bits define the word validity. The last bit of each word is for an odd parity error test.
The main memory of each GPC is non-volatile (the software is retained when power is interrupted). The memory capacity of each CPU is 81,920 words, and the memory capacity of each IOP is 24,576 words; thus, the CPU and IOP constitute a total of 106,496 words.
The hardware controls for the GPCs are located on panel O6. Each computer reads the position of its corresponding output , initial program load and mode switches from discrete input lines that go directly to the GPC. Each GPC also has an output and mode talkback indicator on panel O6 that are driven from GPC output discretes.
Each GPC power on , off switch is a guarded switch. Positioning a switch to on provides the computer with triply redundant power (not through a discrete) by three essential buses-ESS1BC, 2AC and 3AB-which run through the GPC power switch. The essential bus power is transferred to remote power controllers, which permits main bus power from the three main buses (MNA, MNB and MNC) to power the GPC. There are three RPCs for the IOP and three for the CPU; thus, any GPC will function normally, even if two main or essential buses are lost.
Each computer uses over 600 watts of power. GPCs 1 and 4 are located in forward middeck avionics bay 1, GPCs 2 and 5 are located in forward middeck avionics bay 2, and GPC 3 is located in aft middeck avionics bay 3. The GPCs receive forced-air cooling from an avionics bay fan. There are two fans in each avionics bay but only one is powered at a time. If both fans in an avionics bay fail, the computers will overheat and could not be relied on to operate properly for more than 20 minutes if the initial condition is warm.
Each GPC output switch is a guarded switch with backup , normal and terminate positions. The output switch provides a hardware override to the GPC that precludes that GPC from outputting (transmitting) on the flight-critical buses. The switches for the primary avionics GN&C GPCs are positioned to normal , which permits them to output (transmit). The backup flight system GPC switch is positioned to backup, which precludes it from outputting until it is engaged. The switch for a GPC designated on orbit to be a systems management computer is positioned to terminate since the GPC is not to command anything on the flight-critical buses.
The output talkback indicator above each output switch on panel O6 indicates gray if that GPC output is enabled and barberpole if it is not.
Each GPC receives run , stby , or halt discrete inputs from its mode switch on panel O6, which determines whether that GPC can process software. The mode switch is lever-locked in the run position. The halt position for a GPC initiates a hardware-controlled state in which no software can be executed. A GPC that fails to synchronize with others is moded to halt as soon as possible to prevent the failed computer from outputting erroneous commands. The mode talkback indicator above the mode switch for that GPC indicates barberpole when that computer is in halt.
\B\IFor more information click on \b\i\JShuttle Data Processing System (continued)\j
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"Shuttle Data Processing System (continued)",365,0,0,0
In standby, a GPC is also in a state in which no software can be executed but is in a software-controlled state. The stby discrete allows an orderly startup or shutdown of processing. It is necessary, as a matter of procedure, for a GPC that is shifting from run to halt to be temporarily (more than one second) in the standby mode before going to halt since the standby mode allows for an orderly software cleanup and allows a GPC to be correctly initialized without an initial program load. If a GPC is moded from run to halt without pausing in standby, it may not perform its functions correctly upon being remoded to run. There is no stby indication on the mode talkback indicator above the mode switch; however, it would indicate barberpole in the transition from run to standby and run from standby to halt.
The run position permits a GPC to support its normal processing of all active software and assigned vehicle operations. Whenever a computer is moded from standby or halt to run, it initializes itself to a state in which only system software is processed (called OPS 0). If a GPC is in another OPS before being moded out of run and the initial program has not been loaded since, that software still resides in main memory; but it will not begin processing until that OPS is recalled by flight crew keyboard entry. The mode talkback indicator always reads run when that GPC switch is in run and the computer has not failed.
Placing the backup flight system GPC in standby does not stop BFS software processing or preclude BFS engagement; it only prevents the BFS from commanding.
The IPL push button indicator for a GPC on panel O6 activates the initial program load command discrete input when depressed. When the input is received, that GPC initiates an IPL from whichever mass memory unit is specified by the IPL source , MMU 1 , MMU 2 , off switch on panel O6. The talkback indicator above the mode switch for that GPC indicates IPL .
During non-critical flight periods in orbit, only one or two GPCs are used for GN&C tasks and another for systems management and payload operations.
A GPC on orbit can also be ''freeze-dried;'' that is, it can be loaded with the software for a particular memory configuration and then moded to standby. It can then be moded to halt and powered off. Since the GPCs have non-volatile memory, the software is retained. Before an OPS transition to the loaded memory configuration, the freeze-dried GPC can be moded back to run and the appropriate OPS requested.
A simplex GPC is one in run and not a member of the redundant set, such as the BFS GPC. Systems management and payload major functions are always in a simplex GPC.
A failed GPC can be hardware-initiated, stand-alone-memory-dumped by switching the powered computer to terminate and halt and then selecting the number of the failed GPC on the GPC memory dump rotary switch on panel M042F in the crew compartment middeck. Then the GPC is moded to standby to start the dump, which takes three minutes.
Each CPU is 7.62 inches high, 10.2 inches wide and 19.55 inches long; it weighs 57 pounds. The IOPs are the same size and weight as the CPUs.
The new upgraded general-purpose computers, AP-101S from \JIBM\j, will replace the existing GPCs, AP-101B, aboard the space shuttle orbiters in mid-1990.
The upgraded GPCs allow NASA to incorporate more capabilities into the space shuttle orbiters and apply more advanced computer technologies than were available when the orbiter was first designed. The new design began in January 1984, whereas the older GPC design began in January 1972.
The upgraded computers provide 2.5 times the existing memory capacity and up to three times the existing processor speed with minimum impact on flight software. The upgraded GPCs are half the size and approximately half the weight of the old GPCs, and they require less power to operate.
The upgraded GPCs consist of a central processor unit and an input/output processor in one avionics box instead of the two separate CPU and IOP avionics boxes of the old GPCs. The upgraded GPC can perform more than 1 million benchmark tests per second in comparison to the older GPC's 400,000 operations per second. The upgraded GPCs have a semiconductor memory of 256,000 32-bit words; the older GPCs have a core memory of up to 104,000 32-bit words.
The upgraded GPCs have volatile memory, but each GPC contains a battery pack to preserve the software when the GPC is powered off.
The initial predicted reliability of the upgraded GPCs is 6,000 hours mean time between failures, with a projected growth to 10,000 hours mean time between failures. The mean time between failures for the older GPCs is 5,200 hours-more than five times better than the original reliability estimate of 1,000 hours.
The AP-101S avionics box is 19.55 inches long, 7.62 inches high and 10.2 inches wide, the same as one of the two previous GPC avionics boxes. Each of the five upgraded GPCs aboard the orbiter weighs 64 pounds, in comparison to 114 pounds for the two units of the older GPCs. This change reduces the weight of the orbiter's avionics by approximately 300 pounds and frees a volume of approximately 4.35 cubic feet in the orbiter avionics bays. The older GPCs require 650 watts of electrical power versus 550 watts for the upgraded units.
Thorough testing, documentation and \Jintegration\j, including minor modifications to flight software, were performed by \JIBM\j and NASA's Shuttle Avionics Integration Laboratory in NASA's Avionics \JEngineering\j Laboratory at the Johnson Space Center.
\BMaster Events Controllers\b
The two master events controllers under GPC control send signals to arm and safe pyrotechnics and command and fire pyrotechnics during the solid rocket booster/external tank separation process and the orbiter/external tank separation process. The MEC contractor is Rockwell International, Autonetics Group, Anaheim, Calif.
\BBackup Flight Control\b
Even though the four primary avionics software system GPCs control all GN&C functions during the critical phases of the mission, there is always a possibility that a generic failure could cause loss of vehicle control. Thus, the fifth GPC is loaded with different software created by a different company than the PASS developer.
This different software is the backup flight system. To take over control of the vehicle, the BFS monitors the PASS GPCs to keep track of the current state of the vehicle. If required, the BFS can take over control of the vehicle upon the press of a button. The BFS also performs the systems management functions during ascent and entry because the PASS GPCs are operating in GN&C. BFS software is always loaded into GPC 5 before flight, but any of the five GPCs could be made the BFS GPC if necessary.
The BFS interface programs, events and applications controllers, and GN&C are provided by the Charles Stark Draper Laboratory Inc., Cambridge, Mass. The remainder of the software, as well as the \Jintegration\j of the total backup flight control system, is provided by Intermetrics and Rockwell International. The GN&C software is written in HAL/S by Intermetrics of \JBoston\j, Mass.
Since the BFS is intended to be used only in a contingency, its programming is much simpler than that of the PASS. Only the software necessary to complete ascent or entry safely, maintain vehicle control in orbit and perform systems management functions during ascent and entry is included. Thus, all the software used by the BFS can fit into one GPC and never needs to access mass memory. For added protection, the BFS software is loaded into the MMUs in case of a BFS GPC failure.
The BFS, like PASS, consists of system software and applications software. System software in the BFS performs basically the same functions as it does in PASS. These functions include time management, PASS/BFS interface, multifunction CRT display system, input/output, uplink/downlink and engage/disengage control. The system software is always operating when the BFS GPC is not in halt.
Applications software in the BFS has different major functions, GN&C and systems management; but all of its applications software resides in main memory at one time, and the BFS can process software in both major functions simultaneously. The GN&C functions of the BFS, designed as a backup capability, support the ascent phase beginning at major mode 102 and the deorbit/entry phase beginning at major mode 301. In addition, the various ascent abort modes are supported by the BFS.
The BFS provides only limited support for on-orbit operations through major modes 106 or 301. Because the BFS is designed to monitor everything the PASS does during ascent and entry, it has the same major modes as the PASS in OPS 1, 3 and 6.
The BFS systems management contains software to support the ascent and entry phases of the mission. Whenever the BFS GPC is in the run or standby mode, it runs continuously; however, the BFS does not control the payload buses in standby. The systems management major function in the BFS is not associated with any operational sequence.
The BFS also monitors some inputs to PASS CRTs and updates its own GN&C parameters accordingly. When the BFS GPC is tracking the PASS GPCs, it cannot command over the FC buses but may listen to FC inputs through the listen mode.
The BFS GPC controls its own instrumentation/PCMMU data bus. The BFS GPC intercomputer communication data bus is not used to transmit status or data to the other GPCs; and the MMU data buses are not used except during initial program load and MMU assignment, which use the same IPL source switch used for PASS IPL.
A major difference between the PASS and BFS is that the BFS can be shifted into OPS 1 or 3 at any time, even in the middle of ascent or entry.
The BFC lights on panels F2 and F4 remain unlighted as long as PASS is in control and the BFS is tracking. The lights flash if the BFS loses track of the PASS and stands alone. The flight crew must then decide whether to engage the BFS or try to initiate BFS tracking again by a reset. When BFS is engaged and in control of the flight-critical buses, the BFC lights are illuminated and stay on until the BFC is disengaged.
Since the BFS does not operate in a redundant set, its discrete inputs and outputs, which are fail votes from and against other GPCs, are not enabled; thus, the GPC matrix status light on panel O1 for the BFS GPC does not function as it does in PASS. The BFS can illuminate its own light on the GPC matrix status panel if the watchdog timer in the BFS GPC times out or if the BFS GPC does not complete its cyclic processing.
To engage the BFS, which is considered a last resort to save the vehicle, the crew presses a BFS engage momentary push button located on the commander's or pilot's rotational hand controller. As long as the RHC is powered and the BFS GPC output switch is in backup on panel O6, depressing the engage push button on the RHC engages the BFS and causes PASS to relinquish control during ascent or entry.
There are three contacts in each engage push button, and all three contacts must be made to engage the BFS. The signals from the RHC are sent to the backup flight controller, which handles the engagement logic.
When the BFS is engaged, the BFC lights on panels F2 and F4 are illuminated; the BFS output talkback indicator on panel O6 turns gray; all PASS GPC output and mode talkback indicators on panel O6 display a barberpole; the BFS controls the CRTs selected by the BFS CRT select switch on panel C3; big X and poll fail appear on the remaining CRTs; and all four GPC status matrix indicators for PASS GPCs are illuminated on panel O1.
When the BFS is disengaged and the BFC CRT switch on panel O3 is positioned to on, the BFS commands the first CRT indicated by the BFC CRT select switch. The BFC CRT select switch positions on panel C3 are 1 + 2 , 2 + 3 and 3+1. When the BFS is engaged, it assumes control of the second CRT as well.
If the BFS is engaged during ascent, the PASS GPCs can be recovered on orbit to continue a normal mission. This procedure takes about two hours, since the PASS inertial measurement unit reference must be re-established. To disengage the BFS after all PASS GPCs have been hardware-dumped and software-loaded, the PASS GPCs must be taken to GN&C OPS 3. Positioning the BFC disengage momentary switch on panel F6 to the up position disengages the BFS. The switch sends a signal to the BFC that resets the engage discretes to the GPCs. The BFS then releases control of the flight-critical buses as well as the payload buses if it is in standby, and the PASS GPCs assume command.
Indications of the PASS engagement and BFS disengagement are as follows: BFC lights on panels F2 and F4 are out, BFS output talkback indicator on panel O6 displays a barberpole, PASS output talkback indicators on panel O6 are gray and BFS release/PASS control appears on the CRT. After disengagement, the PASS and BFS GPCs return to their normal pre-engaged state.
If the BFS is engaged, there is no manual thrust vector control or manual throttling capability during first- and second-stage ascent. If the BFS is engaged during entry, the speed brake is positioned using the speed brake/thrust controller and the body flap is positioned manually. The BFC system also augments the control stick steering mode of maneuvering the vehicle with the commander's rotational hand controller.
The software of the BFC system is processed only for the commander's attitude director indicator, horizontal situation indicator and RHC. The BFC system supplies attitude errors on the CRT trajectory display, whereas PASS supplies attitude errors to the ADIs; however, when the BFC system is engaged, the errors on the CRT are blanked.
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"Shuttle Guidance, Navigation and Control",366,0,0,0
Guidance, navigation and control software command the GN&C system to effect vehicle control and to provide the sensor and controller data needed to compute these commands. The process involves three steps: guidance equipment and software first compute the orbiter location required to satisfy mission requirements, navigation then tracks the vehicle's actual location, and flight control then transports the orbiter to the required location.
A redundant set of four orbiter general-purpose computers forms the primary avionics software system; a fifth GPC is used as the backup flight system.
The GPCs interface with the various systems through the orbiter's flight forward and flight aft multiplexers/demultiplexers. The data buses serve as a conduit for signals going to and from the various sensors that provide velocity and attitude information as well as for signals traveling to and from the orbiter propulsion systems, orbiter aerodynamic control surfaces, and displays and controls.
The GN&C system consists of two operational modes: auto and manual (control stick steering). In the automatic mode, the primary avionics software system essentially allows the GPCs to fly the vehicle; the flight crew simply selects the various operational sequences. The flight crew may control the vehicle in the control stick steering mode using hand controls, such as the rotational hand controller, translational hand controller, speed brake/thrust controller and rudder pedals. The translational hand controller is available only for the commander, but both the commander and pilot have a rotational hand controller.
In the control stick steering mode, flight crew commands must still pass through and be issued by the GPCs. There are no direct mechanical links between the flight crew and the orbiter's various propulsion systems or aerodynamic surfaces; the orbiter is an entirely digitally controlled, fly-by-wire vehicle.
During launch and ascent, most of the GN&C commands are directed to gimbal the three space shuttle main engines and solid rocket boosters to maintain thrust vector control through the vehicle's center of gravity at a time when the amount of consumables is changing rapidly.
In addition, the GN&C controls SSME throttling for maximum aerodynamic loading of the vehicle during ascent-referred to as max q-and to maintain an acceleration of no greater than 3 g's during the ascent phase. To circularize the orbit and perform on-orbit and deorbit maneuvers, the GN&C commands the orbital maneuvering system engines. At external tank separation, on orbit and during portions of entry, GN&C controls commands to the reaction control system. In atmospheric flight, GN&C controls the orbiter aerodynamic flight control surfaces.
Functions of GN&C software include flight control, guidance, navigation, hardware data processing and flight crew display. Specific function tasks and their associated GN&C hardware vary with each mission phase.
Vehicle control is maintained and in-flight trajectory changes are made during powered flight by firing and gimbaling engines. During atmospheric flight, these functions are performed by deflecting aerosurfaces. Flight control computes and issues the engine fire and gimbal commands and aerosurface deflection commands.
Flight control includes attitude processing, steering, thrust vector control and digital autopilots. Flight control receives vehicle dynamics commands (attitudes, rates and accelerations) from guidance software or flight crew controllers and processes them for conversion to effector commands (engine fire, gimbal or aerosurface). Flight control output commands are based on errors for stability augmentation. The errors are the difference between the commanded attitude, aerosurface position, body rate or body acceleration and the actual attitude, position, rate or acceleration.
Actual attitude is derived from inertial measurement unit angles, aerosurface position is provided by feedback transducers in the aerosurface servoamplifiers, body rates are sensed by rate gyro assemblies, and accelerations are sensed by accelerometer assemblies. In atmospheric flight, flight control adjusts control sensitivity based on air data parameters derived from local pressures sensed by air data probes and performs turn coordination using body attitude angles derived from IMU angles. Thus, GN&C hardware required to support flight control is a function of the mission phase.
The guidance steering commands used by the flight control software are augmented by the guidance software or are manually commanded by the hand controller or speed brake/thrust controller. When flight control software uses the steering commands computed by guidance software, it is termed automatic guidance; when the flight crew is controlling the vehicle by hand, it is called control stick steering.
The commands computed by guidance are those required to get from the current state (position and velocity) to a desired state (specified by target conditions, attitude, airspeed and runway centerline). The steering commands consist of translational and rotational angles, rates and accelerations. Guidance receives the current state from navigation software. The desired state or targets are part of the initialized software load and some may be changed manually in flight.
The navigation system maintains an accurate estimate of vehicle position and velocity, referred to as a state vector. From position, attitude and velocity, other parameters (acceleration, angle of attack) are calculated for use in guidance and for display to the crew. The current state vector is mathematically determined from the previous state vector by integrating the equations of motion using vehicle acceleration as sensed by the IMUs and/or computed from gravity and drag models.
The alignment of the IMU and, hence, the accuracy of the resulting state vector deteriorate as a function of time. Celestial navigation instruments (star trackers and crewman optical alignment sight) are used to maintain IMU alignment in orbit. For entry, the accuracy of the IMU-derived state vector is, however, insufficient for either guidance or the flight crew to bring the \Jspacecraft\j to a pinpoint landing. Therefore, data from other navigation sensors-air data system, tactical air navigation, microwave scan beam landing system and radar altimeter-is blended into the state vector at different phases of entry to provide the necessary accuracy.
The three IMUs maintain an inertial reference and provide velocity changes until the microwave scan beam landing system is acquired. Navigation-derived air data are needed during entry as inputs to guidance, flight control and flight crew dedicated displays. Such data are provided by tactical air navigation, which supplies range and bearing measurements beginning at 160,000 feet; the air data system provides information at about Mach 3. Tactical air navigation is used until the microwave scan beam landing system is acquired or an altitude of 1,500 feet is reached if MSBLS is not available.
During rendezvous and proximity operations, the onboard navigation system maintains the state vectors of both the orbiter and target vehicle. During close operations (separation of less than 15 miles), these two state vectors must be very accurate in order to maintain an accurate relative state vector. Rendezvous radar measurements (range and range rate) are used for a separation of about 15 miles to 100 feet to provide the necessary relative state vector accuracy. When two vehicles are separated by less than 100 feet, the flight crew relies primarily on visual monitoring (aft and overhead windows and closed-circuit television).
In summary, GN&C hardware sensors used by navigation include IMUs, star trackers, the crewman optical alignment sight, tactical air navigation, air data system, microwave scan beam landing system, radar \Jaltimeter\j and rendezvous radar. The GN&C hardware sensors used by the flight control system are accelerometer assemblies, orbiter rate gyro assemblies, solid rocket booster rate gyro assemblies, controllers and aerosurface servoamplifiers.
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"Shuttle Flight Control System Hardware",367,0,0,0
Each piece of GN&C hardware is hard-wired to one of eight flight-critical multiplexers/demultiplexers, which are connected to each of the five GPCs by data buses. Each GPC is assigned to command on one or more data buses; assignments can be changed.
A sensor, controller or flight control effector cannot be assigned or rerouted to another MDM during flight, but the MDM it is wired to can be assigned to a different GPC. Each multiple unit of each type of GN&C hardware is hard-wired to a different MDM. For example, there are four accelerometer assemblies on board the orbiter. AA 1 is wired to forward flight MDM 1 and is part of string 1, AA 2 is wired to FF MDM 2 on string 2, AA 3 is wired to FF MDM 3 on string 3, and AA 4 is wired to FF MDM 4 on string 4.
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"Shuttle Navigation Aids",368,0,0,0
Navigation aids on board the orbiter include the three inertial measurement units, three tactical air navigation units, two air data probe assemblies, three microwave scan beam landing systems and two radar altimeters.
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"Shuttle Inertial Measurement Units",369,0,0,0
The IMUs consist of an all-attitude, four-gimbal, inertially stabilized platform. They provide inertial attitude and velocity data to the GN&C software functions. Navigation software uses the processed IMU velocity and attitude data to propagate the orbiter state vector. Guidance uses the attitude data, along with state vector from the navigation software, to develop steering commands for flight control. Flight control uses the IMU attitude data to convert the steering commands into control surface, engine gimbal (thrust vector control) and reaction control system thruster fire commands.
Although flight could be accomplished with only one, three IMUs are installed on the orbiter for redundancy. The IMUs are mounted on the navigation base, which is located inside the crew compartment flight deck forward of the flight deck control and display panels. The navigation base mounting platform is pitched down 10.6 degrees from the orbiter's plus X body axis. The navigation base provides a platform for the IMUs that can be repeatedly mounted with great accuracy, enabling the definition of transformations that relate IMU reference frame measurements to any other reference frame.
The IMU consists of a platform isolated from vehicle rotations by four gimbals. Since the platform does not rotate with the vehicle, its orientation remains fixed, or inertial, in space. The gimbal order from outermost to innermost is outer roll, pitch, inner roll and \Jazimuth\j. The platform is attached to the \Jazimuth\j gimbal. The inner roll gimbal is a redundant gimbal used to provide an all-attitude IMU while preventing the possibility of gimbal-lock (a condition that can occur with a three-gimbal system and cause the inertial platform to lose its reference). The outer roll gimbal is driven from error signals generated from disturbances to the inner roll gimbal. Thus, the inner roll gimbal will remain at its null position, orthogonal to the pitch gimbal.
The inertial sensors consist of two gyros, each with two degrees of freedom, that provide platform stabilization. The gyros are used to maintain the platform's inertial orientation by sensing rotations of the platform caused by vehicle-rotation-induced friction at the gimbal pivot points. The gyros output a signal that is proportional to the motion and is used by the gimbal \Jelectronics\j to drive the appropriate gimbals to null the gyro outputs.
Thus, the platform remains essentially undisturbed, maintaining its inertial orientation while the gimbals respond to vehicle motion. One gyro-called the vertical gyro-is oriented so its input axes are aligned with the X and Y platform axes; its input axes provide IMU platform roll and pitch stabilization. The second gyro is oriented so that one input axis lies along the platform's Z axis and the other lies in the X-Y plane.
This gyro-the \Jazimuth\j gyro-provides platform yaw stabilization with the Z input axis, while the second input axis is used as a platform rate detector for built-in test equipment. Each gyro contains a two-axis pick-off that senses deflection of the rotating wheel. The gyro also contains a pair of two-axis torquers that provide compensation torquing for gyro drift and a means to reposition the platform.
The spin axis of a gyro is its axis of rotation. The inertial stability of the spin axis is a basic property of gyroscopes and is used in stabilization loops, which consist of the gyro pick-off, gimbals and gimbal torquers. When the vehicle is rotated, the platform also tends to rotate due to friction at the gimbal pivot points. Since the gyro casing is rigidly mounted to the platform, it will also rotate.
The gyro resists this rotation tendency to remain inertial, but the resistance is overcome by friction. This rotation is detected by the pick-offs as a deflection of the rotating gyro wheel. A signal proportional to this deflection is sent to the gimbal \Jelectronics\j, which routes the signals to the appropriate torquers, which in turn rotate their gimbals to null the pick-off point. When the output is nulled, the loop is closed.
Two accelerometers in each IMU measure linear vehicle accelerations in the IMU inertial reference frame; one measures the acceleration along the platform's X and Y axes, the other along the Z axis. The accelerometer is basically a force rebalance-type instrument.
When the accelerometer experiences an acceleration along its input axes, it causes a \Jpendulum\j mass displacement. This displacement is measured by a pick-off device, which generates an electrical signal that is proportional to the sensed acceleration. This signal is amplified and returned to a torquer within the accelerometer, which attempts to reposition the proof mass to its null (no output) position.
The velocity data measured by the IMU are the primary sources that propagate the orbiter state vector during ascent and entry. On orbit, a sophisticated drag model is substituted for IMU velocity information, except during large vehicle accelerations. During large on-orbit accelerations, IMU velocity data are used in navigation calculations.
Platform attitude can be reoriented by two methods: slewing or pulse torquing. Slewing rotates the platform at a high rate (72 degrees per minute), while pulse torquing rotates it very slowly (0.417 degree per minute). Platform reorientation relies on another property of gyroscopes: precession. If a force is applied to a spinning \Jgyroscope\j, the induced motion is 90 degrees from the input force.
In each IMU, a two-axis torquer is located along the input axes of both gyros. Commands are sent to the torquers from the GPC to apply a force along the input axes. The result is a deflection of the gyro spin axis that is detected and nulled by the stabilization loops. Since the gyro spin axis is forced to point in a new direction, the platform has to rotate to null the gyro outputs.
The three IMUs have skewed orientations-their axes are not coaligned and are not aligned with the vehicle axes. This is done for two reasons. First, gimbaled platforms have problems at certain orientations. This skewing ensures that no more than one IMU will have an orientation problem for a given attitude. Skew allows resolution of a single-axis failure on one IMU by multiple axes on another IMU since the possibility of multiple-axis failure is more remote. Second, skewing is also used by redundancy management to determine which IMUs have failures.
The IMU platform is capable of remaining inertial for vehicle rotations of up to 35 degrees per second and angular accelerations of 35 degrees per second squared. Each IMU interfaces with the five onboard GPCs through a different flight forward multiplexer/demultiplexer of the data bus network. Under GPC control, each IMU is capable of orienting its platform to any attitude, determining platform alignment relative to a reference and providing velocity and attitude data for flight operations.
Very precise thermal control must be maintained in order to meet IMU performance requirements. The IMU thermal control system consists of an internal heater system and a forced-air cooling system. The internal heater system is completely automatic and is powered on when power is initially applied to the IMU. It continues to operate until the IMU is powered down.
The forced-air cooling consists of three fans that serve all three IMUs. Only one fan is necessary to provide adequate air flow. The IMU fan pulls cabin air through the casing of each IMU and cools it in an IMU heat exchanger before returning it to the cabin. Each IMU fan is controlled by an individual on/off switch located on panel L1.
Each IMU is supplied with redundant 28-volt dc power through separate remote power controllers when control bus power is applied to the RPCs by the IMU power switch. The IMU 1 , 2 and 3 on/off power switches are located on panels O14, O15 and O16, respectively. Loss of one control bus or one main bus will not cause the loss of an IMU.
Each IMU has two modes of operation: a warm-up/standby mode and an operate mode. When the respective IMU switch is positioned to on , that IMU is powered and enters the warm-up/standby mode, which applies power only to the heater circuits. It takes approximately 30 minutes for the IMU to reach its operating range, at which time the IMU enters a standby mode, when it can be moded to the operate mode by flight crew command in GN&C OPS 2, 3 or 9.
To mode the IMU to operate, the controlling GPC sends the operate discrete to the IMU through the forward flight multiplexer/demultiplexer. The IMU, upon receiving this command, initiates its run-up sequence.
The run-up sequence first cages the IMU-a process of reorienting the IMU gimbals and then mechanically locking them into place so that the gyros may begin to spin. When the IMU is caged, its platform orientation will be known when it becomes inertial. The caged orientation is defined as the point at which all resolver outputs are zero. This causes the IMU platform to lie parallel to the navigation base plane with its coordinate axes lying parallel to the navigation base's coordinate axes.
Once the IMU gimbals are caged, the gyros begin to spin and power is applied to the remaining IMU components. When the gyros have reached the correct spin rate, the stabilization loops are powered, and the IMU becomes inertial. At this time, the IMU returns an in operate mode discrete to the GPC, indicating that the run-up sequence is complete. This process requires approximately 38 seconds.
The IMUs are in operate by the time the flight crew enters the vehicle before launch and remain in that state for the duration of the flight unless powered down to minimize power consumption. While in the operate mode, the IMU maintains its inertial orientation and is used for calibrations and preflight, flight and on-orbit alignments.
Before preflight, the IMUs are taken through three levels of \Jcalibration\j to correct for hardware inaccuracies: factory \Jcalibration\j, hangar \Jcalibration\j and preflight \Jcalibration\j. Sixty-one IMU parameters are developed during this extensive \Jcalibration\j period. These parameters are stored in the orbiter GPC mass memory units and are used in the software to compensate for hardware inaccuracies.
At T minus two hours during the launch countdown, the IMU \Jcalibration\j is complete and the IMUs are ready for the preflight alignment. At T minus one hour and one minute, the Launch Control Center initiates this alignment by a display \Jelectronics\j unit equivalent. (A DEU equivalent is simply a ground command that looks to the GPCs like a crew keyboard input.)
Preflight alignment requires 48 minutes to complete and consists of two different operations: gyrocompass alignment and velocity/tilt initialization.
In the gyrocompass alignment, each IMU is oriented so that the desired relative skew is achieved when the platforms are at their alignment orientation. During this phase, the IMUs are placed in two orientations relative to the north-west-up coordinate system. These two orientations differ only in a 90-degree rotation about the up axis.
Data are collected for 90 seconds by the accelerometers to remove any misalignment resulting from the reorientation. The accelerometers are used here because their accuracy is much better than that of the resolvers and the acceleration due to Earth rotation is definitely known. Therefore, any unexpected acceleration is due to IMU misalignment. Once this misalignment is nulled, the platform is torqued about the north axis to compensate for Earth rotation.
Data are then collected for 10 minutes to measure platform drifts. This sequence of data collecting is repeated at the second orientation. The relative attitude errors for each IMU pair are also computed, first with resolver data and then with accelerometer data. The two values are subtracted and transformed into body coordinates.
A factory-calibrated relative resolver error term is then subtracted, and a reasonableness test is performed to check the relative alignment between each IMU pair to assure a good preflight alignment. The velocity/tilt initialization mode is then entered, during which the drifts experienced while waiting for the OPS 1 transition are estimated.
The compensation developed by these drifts is applied to the gyros from the OPS 1 transition to T minus 12 minutes and is also used to compute the current platform to the mean of 1950 reference stable member matrix at the OPS 1 transition. In addition, a level-axis test is performed on each platform three times a second; failure to pass this test requires the alignment to be repeated.
At T minus 22 seconds, a one-shot data transfer from the primary avionics software system to the backup flight system is commanded by display \Jelectronics\j unit equivalent. IMU compensation data computed by the PASS GPCs in OPS 9 are sent to the BFS GPC at this time so that it will have the same data for controlling the IMUs if it is engaged.
At the OPS 1 transition, the IMUs enter the ''tuned inertial'' drift compensation mode. It is tuned because a compensation factor computed in the velocity/tilt is applied to the IMU gyros. At T minus 12 seconds, this compensation is removed and the IMUs enter the ''free inertial'' mode. The IMUs are now flight ready, and all functions, both hardware and software, remain the same throughout the flight.
During ascent, the IMUs provide accelerometer and resolver data to the GN&C software to propagate the state vector, determine attitude and display flight parameters.
During the orbital flight phase, the IMUs provide GN&C software with attitude and accelerometer data.
On-orbit alignments are necessary to correct platform misalignment caused by uncompensated gyro drift.
During entry, IMU operation differs only in the manner in which accelerometer data are used by navigation.
The IMU can be safely powered off from either the warm-up/standby mode or the operate mode. If an IMU is moded to standby, an internal timer inhibits moding operation for three minutes to allow the gyros to spin to a stop so that the proper sequencing to the operate mode can occur.
The IMU software scheme is designed to select the best data for GPC use and to detect system failures. This scheme is referred to as redundancy management.
In the event of an IMU failure, the IMU red caution and warning light on panel F7 will be illuminated. If temperatures are out of limits or if built-in test equipment detects a failure, a fault message and SM alert will be annunciated.
The accuracy of the IMU deteriorates with time. If the errors are known, they can be physically or mathematically corrected. Software based on preflight calibrations is used to compensate for most of the inaccuracy. The star trackers and crewman optical alignment sight are used to determine additional inaccuracies.
The IMU subsystem operating program processes the data from the IMUs and converts it to a usable form for other users. The following computations are performed in the IMU SOP: conversion of velocities to the mean of 1950 coordinates; conversion of resolver outputs to gimbal angles; computation of accelerations for displays; performance of additional software built-in test equipment checks; support of selection, control and monitoring of IMU submodes of the operate mode; and computation of torquing commands based on the misalignment determined by the star trackers, crewman optical alignment sight or another IMU. Misalignments are due to gyro drifts.
Each KT-70 IMU is 10.28 inches high, 11.5 inches wide and 22 inches long and weighs 58 pounds.
A new high-accuracy inertial navigation system will be phased in to augment the present KT-70 IMU during 1988-89. The HAINS will provide spares support for the inertial navigation system and will eventually phase out the KT-70 IMU design. Benefits of the HAINS include lower program costs over the next decade, ongoing production support, improved performance, lower failure rates, and reduced size and weight. The HAINS is 9.24 inches high, 8.49 inches wide and 22 inches long.
The unit weighs 43.5 pounds. The HAINS also contains an internal dedicated \Jmicroprocessor\j with memory for processing and storing compensation and scale factor data from the vendor's calibrations, thereby reducing the need for extensive initial load data for the orbiter GPCs. The HAINS is both physically and functionally interchangeable with the KT-70 IMU.
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"Shuttle Star Trackers",370,0,0,0
The star tracker system is part of the orbiter's navigation system. Its two units are located just forward and to the left of the commander's plus X window in a well outside the pressurized crew compartment-an extension of the navigation base on which the IMUs are mounted. The star trackers are slightly inclined off the vehicle's negative Y and negative Z axes, for which they are named. The star trackers are used to align the IMUs on board the orbiter as well as to track targets and provide line-of-sight vectors for rendezvous calculations.
Alignment of the IMUs is required approximately every 12 hours to correct IMU drift, within one to two hours before major on-orbit thrusting duration or after a crewman optical alignment sight IMU alignment. IMU alignment is accomplished by using the star trackers to measure the line-of-sight vector to at least two stars. With this information, the GPC calculates the orientation between these stars and the orbiter to define the orbiter's attitude. A comparison of this attitude with the attitude measured by the IMU provides the correction factor necessary to null the IMU error.
The GPC memory contains inertial information for 50 stars chosen for their brightness and their ability to provide complete sky coverage.
The star trackers are oriented so that the optical axis of the negative Z star tracker is pointed approximately along the negative Z axis of the orbiter and the optical axis of the negative Y star tracker is pointed approximately along the negative Y axis of the orbiter. Since the navigation base provides the mount for the IMUs and star trackers, the star tracker line of sight is referenced to the navigation base and the orbiter coordinate system; thus, the GPC knows where the star tracker is pointed and its orientation with respect to the IMUs.
Each star tracker has a door to protect it during ascent and entry. The doors are opened on orbit to permit use of the star trackers.
To enable the star tracker doors to open, the star tracker power minus Y and minus Z switches on panel O6 must be positioned to on. The star tracker door control sys 1 and sys 2 switches on panel O6 control one three-phase ac motor on each door. Positioning the sys 1 switch to open controls motor control logic and drives the minus Y and minus Z star tracker door by electromechanical actuators to the open position.
Limit switches stop the motors when the doors are open and control a talkback indicator above the sys 1 switch. The talkback indicator indicates when the doors are open. Setting the sys 2 switch to open controls a redundant ac motor and electromechanical actuators to open the minus Y and minus Z star tracker door, and limit switches stop the motors when the doors are open and control the talkback indicator above the sys 2 switch in the same manner as for system 1.
The difference between the inertial attitudes defined by the star tracker and the IMU is processed by software and results in IMU torquing angles. If the IMU gimbals are physically torqued or the matrix defining its orientation is recomputed, the effects of the IMU gyro drift are removed and the IMU is restored to its inertial attitude.
If the IMU alignment is in error by more than 1.4 degrees, the star tracker is unable to acquire and track stars. In this case, the crewman optical alignment sight must be used to realign the IMUs to within 1.4 degrees; the star trackers can then be used to realign the IMUs more precisely.
The star tracker cannot be used if the IMU alignment error is greater than 1.4 degrees because the angles the star tracker is given for searching are based on current knowledge of the orbiter attitude, which is based on IMU gimbal angles. If that attitude is greatly in error, the star tracker may acquire and track the wrong star.
In addition to aligning the IMUs, the star trackers can be used to provide angular data from the orbiter to a target. This capability can be used during rendezvous or proximity operations with a target satellite.
The star tracker includes a light shade assembly and an \Jelectronics\j assembly mounted on top of the navigation base. The light shade assembly defines the tracker field of view (10 degrees square). Its shutter mechanism may be opened manually by the crew using an entry on the \Jcathode\j ray tube display, or it can be opened and closed automatically by a bright object sensor or target suppress software.
The bright object sensor reacts before a bright object, such as the sun or moon, can damage the star tracker (the sensor has a larger field of view than the star tracker shutter). The target suppress software reacts to a broad light source (such as the sunlit Earth), which may not trip the bright object sensor but could produce overall illumination large enough to cause photo currents larger than desired.
The \Jelectronics\j assembly contains an image dissector tube mounted on the underside of the navigation base. The star tracker itself does not move-the field of view is scanned electronically. The star tracker may be commanded to scan the entire field of view or a smaller offset field of view (1 degree square) about a point defined by horizontal and vertical offsets.
An object is tracked when the proper intensity and the correct location are sensed. Star tracker outputs are the horizontal and vertical position within the field of view of the object being tracked and its intensity.
There is no redundancy management for the star tracker assemblies; they operate independently, and either can do the whole task. They can be operated either separately or concurrently.
The star tracker subsystem operating program supports the modes that are commanded manually: self-test, star track, target track, break track and term/idle. Self-test consists of software and hardware tests. In the star track mode, the star tracker does an offset scan search for the star, acquires it and tracks it.
The star may be selected by the flight crew or GPC; in either case, field-of-view and \Joccultation\j checks are made. Target track is the same as star track, but the flight crew must specify the target and its threshold. Break track forces the star tracker to stop tracking a star and to perform a search scan from the current location to track the next star it acquires. In the term/idle mode, the star tracker continues its operation, but all star tracker software processing ceases.
In addition, the star tracker SOP maintains the star table. When a star tracker has acquired and tracked a star and the data has passed software checks, the star identification, time tag and line-of-sight vector are stored. The identification and time elapsed since time tag are displayed in the star table.
When two or three stars are in the table, the angular difference between their line-of-sight vectors is displayed. The difference between the star tracker and star catalog angular differences is displayed as an error. The star tracker SOP selects line-of-sight vectors of two stars in the star table for IMU alignment and outputs an align ena discrete.
The software selects the star pair whose angular difference is closest to 90 degrees or the pair whose elapsed time of entry into the table is less than 60 minutes. The flight crew may manually override the SOP selection or clear the table if desired.
The SOP also determines and displays star tracker status.
The crewman optical alignment sight is used if inertial measurement unit alignment is in error by more than 1.4 degrees, rendering the star tracker unable to acquire and track stars. The COAS must be used to realign the IMUs to within 1.4 degrees. The star trackers can then be used to realign the IMUs more precisely.
The COAS is mounted at the commander's station so the crew can check for proper attitude orientation during ascent and deorbit thrusting periods. For on-orbit operations, the COAS at the commander's station is removed and installed next to the aft flight deck overhead right minus Z window.
The COAS is an optical device with a reticle projected on a combining glass that is focused on infinity. The reticle consists of 10-degree-wide vertical and horizontal cross hairs with 1-degree marks and an elevation scale on the right side of minus 10 to 31.5 degrees. A light bulb with variable brightness illuminates the reticle. The COAS requires 115-volt ac power for reticle illumination. The COAS is 9.5 by 6 by 4.3 inches and weighs 2.5 pounds.
After mounting the COAS at the aft flight station, the flight crew member must manually maneuver the orbiter until the selected star is in the field of view. The crew member maneuvers the orbiter so that the star crosses the center of the reticle. At the instant of the crossing, the crew member makes a mark by depressing the most convenient att ref push button; the three att ref push buttons are located on panels F6, F8 and A6.
At the time of the mark, software stores the gimbal angles of the three IMUs. This process can be repeated if the accuracy of the star's centering is in doubt. When the crew member feels a good mark has been taken, the software is notified to accept it. Good marks for two stars are required for an IMU alignment. The separation between the two stars should be between 60 and 120 degrees.
By knowing the star being sighted and the COAS's location and mounting relationship in the orbiter, software can determine a line-of-sight vector from the COAS to the star in an inertial coordinate system. Line-of-sight vectors to two stars define the attitude of the orbiter in inertial space. This attitude can be compared to the attitude defined by the IMUs and can be realigned to the more correct orientation by the COAS sightings if the IMUs are in error.
The COAS's mounting relative to the navigation base on which the IMUs are mounted is calibrated before launch. The constants are stored in software, and COAS line-of-sight vectors are based on known relationships between the COAS line of sight and the navigation base.
COAS can also be used to visually track targets during proximity operations or to visually verify tracking of the correct star by the minus Z star tracker.
COAS data processing is accomplished in the star tracker SOP. This SOP accepts and stores crew inputs on COAS location, star identification or \Jcalibration\j mode; accepts marks; computes and stores the line-of-sight vectors; enables IMU alignment when two marks have been accepted; and computes, updates and provides display data.
#
"Shuttle Air Data System",372,0,0,0
Two air data probes are located on the left and right sides of the orbiter's forward lower fuselage. During the ascent, on-orbit, deorbit and initial entry heat load environment phases, the probes are stowed inside the forward lower fuselage. The air data probe (except for the probe itself) is covered by thermal protection system tiles while in the stowed position. At approximately Mach 3, the air data probes are deployed.
The air data system provides information on the movement of the orbiter in the air mass (flight environment).
The air data system senses air pressures related to \Jspacecraft\j movement through the atmosphere to update navigation state vector in altitude; provide guidance in calculating steering and speed brake commands; and update flight control law computations and provide display data for the commander's and pilot's alpha Mach indicators, altitude/vertical velocity indicators and CRTs.
The AMIs display essential flight parameters relative to the \Jspacecraft\j's travel in the air mass, such as angle of attack (alpha), acceleration, Mach/velocity and knots equivalent airspeed. The altitude/vertical velocity indicators display such essential flight parameters as radar altitude, barometric altitude, altitude rate and altitude acceleration.
Each probe is independently deployed by an actuator consisting of two ac motors connected to rotary electromechanical actuator and limit switches. Each probe is controlled by its air data probe switch on panel C3. To deploy the air data probes, the left and right switches are positioned to deploy. The redundant motors for each probe drive the probe to the deployed position.
When the probe is fully deployed, limit switches remove electrical power from the motors. Deployment time is 15 seconds for two-motor operation and 30 seconds for single-motor operation. The deploy position deploys the probe without electrical heaters. The deploy/heat position also deploys the air data probes with heaters powered.
The air data probe stow left and right switches on panel C3 are used during ground turnaround operations to stow the respective probe. Positioning the respective switch to enable and positioning the corresponding air data probe switch to stow stows the corresponding air data probe. The air data probe stow inhibit position opens the ac motor circuits, disables the stow and protects microswitches.
Each air data probe has four pressure-port sensors and two temperature sensors. The pressures sensed are static pressure, total pressure, angle-of-attack upper pressure and angle-of-attack lower pressure. The four pressures are sensed at ports on each probe: static pressure at the side, total pressure at the front and angle-of-attack lower near the bottom front.
The probe-sensed pressures are connected by a set of pneumatic lines to two air data \Jtransducer\j assemblies. The two temperature sensors are installed on each probe and are wired to an ADTA. The pressures sensed by the left probe are connected by pneumatic tubing to ADTAs 1 and 3. Those sensed by the right probe are connected to ADTAs 2 and 4. Temperatures and sensed pressure from the probes are sent to the same ADTAs.
Within each ADTA, the pressure signals are directed to four transducers, and the temperature signal is directed to a bridge. The pressure \Jtransducer\j analogs are converted to digital data by digital-processor-controlled counters. The temperature signal is converted by an analog-to-digital converter. The digital processor corrects errors, linearizes the pressure data and converts the temperature bridge data to temperatures in degrees \Jcentigrade\j.
These data are sent to the digital output device, which converts the signals into serial digital format, and then to the onboard computers to update the navigation state vector. The data are also sent to the commander's and pilot's altitude/vertical velocity indicators, alpha Mach indicators and CRT.
The ADTA SOP uses ADTA data to compute angle of attack, Mach number (M), equivalent airspeed (EAS), true airspeed (TAS), dynamic pressure (q), barometric altitude (h) and altitude rate ( . h ).
The altitude/vertical velocity indicators and alpha Mach indicators are located on panels F6 and F8 for the commander and pilot, respectively. The information to be displayed on the commander's AVVI and AMI is controlled by the commander's air data switch on panel F6, and the pilot's AVVI and AMI data are controlled by the pilot's air data switch on panel F8.
When the commander's or pilot's switch is positioned to nav , the AVVIs and AMIs receive information from the navigation attitude processor. When the air data probes are deployed, the commander's and pilot's switches can be positioned to left or right to receive information from the corresponding air data probe.
The AVVIs display altitude deceleration ( alt accel ) in feet per second squared, rate in feet per second, navigation/air data system (nav/ADS) in feet and radar ( rdr ) altitude in feet. The AVVI's altitude acceleration indicator remains on the navigation attitude processor from the IMUs. In addition, the radar altimeters on the AVVIs will not receive information until the orbiter reaches an altitude of 5,000 feet.
The AMIs display angle of attack ( alpha ) in degrees, acceleration (accel) in feet per second squared, Mach number or velocity (M/vel) in feet per second and equivalent airspeed ( EAS ) in knots. The AMI's acceleration indicator remains on the navigation attitude processor from the IMUs.
All but the alpha indicators (a moving drum) and the altitude acceleration indicators (a moving pointer displayed against a fixed line) are moving tapes behind fixed lines. The AMI's angle-of-attack indicator reads from minus 18 to plus 60 degrees, the acceleration indicator from minus 50 to plus 100 feet per second squared, the Mach/velocity indicator from Mach zero to 4 and 4,000 to 27,000 feet per second, and equivalent airspeed from zero to 500 knots.
The AVVIs read altitude acceleration from minus 13.3 to plus 13.3 feet per second squared, altitude rate from minus 2,940 to plus 2,940 feet per second, altitude from minus 1,100 to plus 400,000 feet and then changes scale to plus 40 to plus 165 nautical miles (barometric altitude), and radar altitude from zero to plus 9,000 feet.
Failure warning flags are provided for all four scales on the AVVIs and AMIs. The flags appear in the event of a malfunction in the indicator or in received data. In the event of power failure, all four flags appear.
The four computers compare the pressure readings from the four ADTAs for error. If all the pressure readings compare within a specified value, one set of pressure readings from each probe is summed, averaged and sent to the software. If one or more pressure signals of a set of probe pressure readings fail, the failed set's data flow from that ADTA to the averager is interrupted, and the software will receive data from the other ADTA of that probe.
If both probe sets fail, the software operates on data from the two ADTAs connected to the other probe. The best total temperature from all four ADTAs is sent to the software. A fault detection will illuminate the air data red caution and warning light on panel F7, the backup caution and warning alarm light, and the master alarm and will also sound the audible tone and generate a fault message on the CRT. A communication fault will illuminate the SM alert light.
The four ADTAs are located in the orbiter crew compartment middeck forward avionics bays and are convection cooled. Each is 4.87 inches high, 21.25 inches long and 4.37 inches wide and weighs 19.2 pounds.
The air data probe sensor contractor is Rosemount Inc., Eden Prairie, Minn. The air data \Jtransducer\j assembly contractor is AirResearch Manufacturing Co., Garrett Corp., Torrance, Calif. The contractor for the air data probe deploy system is Ellanef, Corona, N.Y.
The three onboard microwave scan beam landing systems are airborne Ku-band receiver/transmitter navigation and landing aids with decoding and computational capabilities. The MSBLS units determine slant range, \Jazimuth\j and elevation to the ground stations alongside the landing runway.
MSBLS is used during terminal area energy management, the approach and landing flight phases and return-to-launch-site aborts. When the channel (specific frequency) associated with the target runway approach is selected, the orbiter's MSBLS units receive elevation from the glide slope ground portion and \Jazimuth\j and slant range from the azimuth/distance-measuring equipment ground station.
The orbiter is equipped with three independent MSBLS sets, each consisting of a Ku-band receiver/transmitter and decoder. Data computation capabilities determine elevation angle, \Jazimuth\j angle and orbiter range with respect to the MSBLS ground station.
The MSBLS provides highly accurate three-dimensional navigation position information to the orbiter to compute state vector components for steering commands that maintain the orbiter on its proper flight trajectory. The three orbiter Ku-band antennas are located on the upper forward fuselage nose. The three MSBLS and decoder assemblies are located in the crew compartment middeck avionics bays and are convection cooled.
The ground portion of the MSBLS consists of two shelters: an elevation shelter and an azimuth/distance-measuring equipment shelter. The elevation shelter is located near the projected touchdown point, with the azimuth/DME shelter located near the far end of the runway. Both ends of the runway are instrumented to enable landing in either direction.
The MSBLS ground station signals are acquired when the orbiter is close to the landing site and has turned on its final leg. This usually occurs on or near the heading alignment cylinder, about 8 to 12 nautical miles (9 to 13 statute miles) from touchdown at an altitude of approximately 18,000 feet.
Final tracking occurs at the terminal area energy management ''autoland'' interface at approximately 10,000 feet altitude and 8 nautical miles (9 statute miles) from the azimuth/DME station.
The MSBLS angle and range data are used to compute steering commands until the orbiter is over the runway approach threshold, at an altitude of approximately 100 feet. If the autoland system is used, it may be overridden by the commander or pilot at any time using the control stick steering mode.
The commander's and pilot's horizontal situation indicators display the orbiter's position with respect to the runway. Elevation and \Jazimuth\j are shown relative to a GPC-derived glide slope on a glide slope indicator; course deviation needles and range are displayed on a mileage indicator. When the orbiter is over the runway threshold, the radar \Jaltimeter\j is used to provide elevation (pitch) guidance. Azimuth/DME data are used during the landing rollout.
The three orbiter MSBLS sets operate on a common channel during the landing phase. The MSBLS ground station transmits a DME solicit pulse. The onboard MSBLS receiver responds with a DME interrogation pulse. The ground equipment responds by transmitting a return pulse. A decoder in the onboard MSBLS decodes the pulses to determine range, \Jazimuth\j and elevation.
Range is a function of the elapsed time between interrogation pulse transmission and signal return. \JAzimuth\j pulses are returned in pairs. The spacing between the two pulses in a pair identifies the pair as \Jazimuth\j and indicates which side of the runway the orbiter is on; spacing between pulse pairs defines the angular position from runway centerline. The spacing between the two pulses in a pair identifies the pair as elevation, and the spacing between pulse pairs defines the angular position of the orbiter above the runway.
The elevation beam is 1.3 to 29 degrees high and 25 degrees to the left and right of the runway. The azimuth/DME beam is zero to 23 degrees high and 13.5 degrees to the left and right of the runway.
Each RF assembly routes range, \Jazimuth\j and elevation information in RF form to its decoder assembly, which processes the information and converts it to digital words for transmission to the onboard GN&C via the multiplexers/demultiplexers for the GPCs.
Elevation, \Jazimuth\j and range data from the MSBLS are used by the GN&C system from the time of acquisition until the runway approach threshold is reached. After that point, the \Jazimuth\j and range data are used to control rollout. Altitude data are provided separately by the orbiter's radar \Jaltimeter\j.
Since the azimuth/DME shelters are at the far ends of the runway, the MSBLS can provide useful data until the orbiter is stopped. \JAzimuth\j data give position in relation to the runway centerline, while the DME gives the distance from the orbiter to the end of the runway.
Each MSBLS has an on/off power switch on panel O8 and on the channel (frequency) selection thumbwheel on panel O8. Positioning the MLS 1, 2 and 3 switch provides power to the corresponding microwave scan beam landing system. MSBLS 1 receives power from main bus A, MSBLS 2 from main bus B and MSBLS 3 from main bus C. Positioning the channel 1 , 2 and 3 thumbwheels selects the frequency (channel) for the ground station at the selected runway for the corresponding MSBLS.
Redundancy management mid-value-selects \Jazimuth\j and elevation angles for processing navigation data. The three MSBLS sets are compared to identify any significant differences among them.
When data from all three MSBLS sets are valid, redundancy management selects middle values of three ranges, azimuths and elevations. In the event that only two MSBLS sets are valid, the two ranges, azimuths and elevations are averaged. If only one MSBLS set is valid, its range, \Jazimuth\j and elevation are passed for display. When a fault is detected, the SM alert light is illuminated, and a CRT fault message is shown.
Each MSBLS decoder assembly is 8.25 inches high, 5 inches wide and 16.16 inches long and weighs 17.5 pounds. The RF assembly is 7 inches high, 3.5 inches wide and 10.25 inches long and weighs 6 pounds.
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"Shuttle Radar Altimeter",374,0,0,0
The two radar altimeters on board the orbiter measure absolute altitude from the orbiter to the nearest terrain within the beamwidth of the orbiter's antennas.
The RAs constitute a low-altitude terrain-tracking and altitude-sensing system based on the precise time it takes an electromagnetic energy pulse to travel from the orbiter to the nearest object on the ground below and return during altitude rate changes of as much as 2,000 feet per second. This enables tracking of mountain or cliff sides ahead or alongside the orbiter if these obstacles are nearer than the ground below and warns of rapid changes in absolute altitude.
The two independent RAs consist of a transmitter and receiver antenna. The systems can operate simultaneously without affecting each other. The four C-band antennas are located on the lower forward fuselage. The two receiver/transmitters are located in the middeck forward avionics bays and are convection cooled.
Each RA transmits a C-band (4,300 MHz modulated at 8.5 kHz) pulse through its transmitting antenna. The signal is reflected by the nearest terrain, and the leading edge of the return radar echo is locked on by the RA through its receiving antenna. The altitude outputs by the RA are analog voltages that are proportional to the elapsed time required for the ground pulse to return, which is a function of height or distance to the nearest terrain. The range output of the RA is from zero to 5,000 feet. The RA will not lock on if the orbiter has large pitch or roll angles.
The onboard GPCs process the data for the autoland mode and touchdown guidance after the orbiter has crossed the runway threshold from an altitude of 100 feet down to touchdown. If the autoland mode is not used, the GPCs process the data for display on the commander's and pilot's altitude/vertical velocity meters from 5,000 feet.
The commander and pilot can select RA 1 or 2 for display on their respective AVVI. The commander's radar al tm 1 and 2 switch is located on panel F7, and the pilot's switch is located on panel F8. The radar \Jaltimeter\j on/off 1 and 2 power switches are on panel O8. Positioning radar \Jaltimeter\j 1 to on provides electrical power to RA 1 from main bus A; positioning radar \Jaltimeter\j 2 to on provides electrical power to RA 2 from main bus B.
The display scale on the commander's and pilot's AVVI raw data recorder indicators ranges from 5,000 to zero feet. Altitude is displayed on a moving tape. Above 9,000 feet, the scale will be pegged. At 1,500 feet, the raw data recorder indicator changes scale. The RA off flag will appear if there is a loss of power, loss of lock, data good-bad or after three communications faults.
Because there are only two radar altimeters on board the orbiter, the altitude data from the two units are averaged in redundancy management when the radar \Jaltimeter\j is used for the autoland mode.
Each radar \Jaltimeter\j receiver/transmitter measures 3.13 inches high, 7.41 inches long and 3.83 inches wide and weighs 4.5 pounds.
The radar \Jaltimeter\j contractor is Honeywell Inc., \JMinneapolis\j, Minn.
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"Shuttle Accelerometer Assemblies",375,0,0,0
There are four accelerometer assemblies aboard the orbiter, each containing two identical single-axis accelerometers, one of which senses vehicle acceleration along the lateral (left and right) vehicle Y axis while the other senses vehicle acceleration along the vertical (normal, yaw and pitch) Z axis.
The AAs provide feedback to the flight control system concerning acceleration errors, which are used to augment stability during first-stage ascent, aborts and entry; elevon load relief during first-stage ascent; and computation of steering errors for display on the commander's and pilot's attitude director indicators during terminal area energy management and approach and landing phases.
The lateral acceleration readings enable the flight control system to null side forces during both ascent and entry. The normal acceleration readings indicate the need to relieve the load on the wings during ascent. During entry, the normal acceleration measurements cue guidance at the proper time to begin ranging. During the latter stages of entry, these measurements provide feedback for guidance to control sink rate. In contrast, the accelerometers within the IMUs measure three accelerations used in navigation to calculate state vector changes.
Each accelerometer consists of a \Jpendulum\j suspended so that its base is in a permanent magnetic field between two torquer magnets. A lamp is beamed through an opening in one of the torquer magnets; photodiodes are located on both sides of the other torquer magnet. When acceleration deflects the \Jpendulum\j toward one photodiode, the resulting light imbalance on the two photodiodes causes a differential voltage, which increases the magnetic field on one of the torquer magnets to return the \Jpendulum\j to an offset position.
The magnitude of the current that is required to accomplish this is proportional to the acceleration. The polarity of the differential voltage depends on the direction of the \Jpendulum\j's movement, which is opposite to the direction of acceleration. The only difference between the lateral and normal accelerometers is the position in which they are mounted within the assembly.
When the acceleration is removed, the \Jpendulum\j returns to the null position. The maximum output for a lateral accelerometer is plus or minus 1 g; for a normal accelerometer, the maximum output is plus or minus 4 g.
The accelerations transmitted to the forward MDMs are voltages proportional to the sensed acceleration. These accelerations are multiplexed and sent to the GPCs, where an accelerometer assembly subsystem operating program converts the eight accelerometer output voltages to gravitational units. This data is also sent to the CRTs and attitude director indicator error needles during entry.
The accelerometer assemblies provide fail-operational redundancy during both ascent and entry. The four AAs employ a quad mid value software scheme to select the best data for redundancy management and failure detection.
Accelerometer 1 is powered from main bus A through the accel 1 circuit breaker on panel O14. Accelerometer 2 is powered from main bus B through the accel 2 circuit breaker on panel O15. Accelerometer 3 is controlled by the accel 3 on/off switch on panel O16. When the switch is positioned to on, power from control buses controls remote power controllers, which supplies main bus A and main bus C to accelerometer 3. The accel 4 on/off switch on panel O15 operates similarly, except that accelerometer 4 receives power from main bus B and main bus C. The accelerometers are turned off once on orbit and on again before entry.
An RGA/accel red caution and warning light on panel F7 will be illuminated if an accelerometer fails.
The four AAs are located in crew compartment middeck forward avionics bays 1 and 2. The AAs are convection cooled and require a five-minute warm-up period.
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"Orbiter Rate Gyro Assemblies",376,0,0,0
The orbiter rate gyro assemblies are used by the flight control system during ascent, entry and aborts as feedbacks to final rate errors that are used to augment stability and for display on the commander's and pilot's attitude director indicator rate needles on panels F6 and F8. The four orbiter RGAs are referred to as RGAs 1, 2, 3 and 4.
The RGAs sense roll rates (about the X axis), pitch rates (about the Y axis) and yaw rates (about the Z axis). These rates are used by the flight control system to augment stability during both ascent and entry.
Each RGA contains three identical single-degree-of-freedom rate gyros so that each gyro senses rotation about one of the vehicle axes. Thus, each RGA contains one gyro-sensing roll rate (about the X axis), one gyro-sensing pitch rate (about the Y axis) and one gyro-sensing yaw rate (about the Z axis).
Each gyro has three axes. A motor forces the gyro to rotate about its spin axis. When the vehicle rotates about the gyro input axis, a \Jtorque\j results in a rotation about the output axis. An electrical voltage proportional to the angular deflection about the output axis-representing vehicle rate about the input axis-is generated and transmitted through the flight aft MDMs to the GPCs and RGA SOP.
This same voltage is used within the RGA to generate a counteracting \Jtorque\j that prevents excessive gimbal movement about the output axis. The maximum output for roll rate gyros is plus or minus 40 degrees per second; for the pitch and yaw gyros, the maximum output is plus or minus 20 degrees per second.
The RGA SOP converts the voltage rate into units of degrees per second.
The RGA 1, 2, 3 and 4 on/off power switches are located on panels O14, O15, O16 and O15, respectively. The redundant power supplies for RGAs 1 and 4 prevent the loss of more than one rate gyro assembly if main bus power is lost.
The RGAs remain off on orbit except during flight control system checkout to conserve power.
The RGAs afford fail-operational redundancy during both ascent and entry. A quad mid value software scheme selects the best data for use in redundancy management and failure detection.
The RGA/accel red caution and warning light on panel F7 will be illuminated to inform the flight crew of an RGA failure.
The RGAs are located on the aft bulkhead below the floor of the payload bay. They are mounted on cold plates for cooling by the Freon-21 coolant loops. The RGAs require a five-minute warm-up time.
The solid rocket booster RGAs are used exclusively during first-stage ascent as feedback to find rate errors from lift-off to two to three seconds before SRB separation. There are three RGAs on each SRB, each containing two identical single-degree-of-freedom rate gyros for sensing rates in the vehicle pitch and yaw axes similar in function to the orbiter RGAs. The maximum outputs for the SRB RGAs are 10 degrees per second.
The SRB RGAs sense pitch and yaw rates, but not roll rates, during the first stage of ascent. Because the SRBs are more rigid than the orbiter body, these rates are less vulnerable to errors created by structural bending. They are thus particularly useful in thrust vector control.
The three RGAs in each SRB are mounted on the forward ring within the forward skirt near the SRB-external tank attach point.
The SRB RGA SOP converts the 12 voltages representing a rate into units of degrees per second. These rates are used by the flight control system during first-stage ascent as feedback to identify rate errors, which are used for stability augmentation. The pitch and yaw axes and a combination of rate, attitude and acceleration signals are blended to provide a common signal to the space shuttle main engines and SRB thrust vector control during first stage. In the roll axis, rate and attitude are summed to provide a common signal to the SSMEs and SRB thrust vector control.
Each of the SRB RGAs is hard-wired to a flight aft MDM to the GPCs through flight-critical buses 5, 6 and 7. In the GPCs, the SOP applies the rate compensation equation to each of the left or right pitch and yaw rates. The compensated rate signals are sent to redundancy management, where the mid value software scheme selects the best data for use and failure detection.
The SRB RGAs are commanded to null and switched out of the flight control system two to three seconds before SRB separation; SRB yaw and pitch rate data are then replaced with orbiter pitch and yaw RGA data.
The RGA/accel red caution and warning light on panel F7 will be illuminated to inform the flight crew of an RGA failure.
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"Shuttle Rotational Hand Controller",378,0,0,0
There are three rotational hand controllers on the orbiter crew compartment flight deck: one at the commander's station, one at the pilot's station and one at the aft flight deck station. Each RHC controls vehicle rotation about three axes: roll, pitch and yaw. During ascent, the commander's and pilot's RHCs may be used to gimbal the SSMEs and SRBs.
For insertion and deorbit, the commander's and pilot's RHCs may be used to gimbal the orbital maneuvering system engines and to command thrusting of the reaction control system engines. On orbit, the commander's, pilot's and aft flight station RHCs may be used to command RCS engine thrusting. During entry, the commander's and pilot's RHCs may be used to command RCS engine thrusting during the early portion of entry and may be used to position the orbiter elevons in roll and pitch axes in the latter portion of entry.
Human factors dictate that an RHC deflection produce a rotation in the same direction as the flight crew member's line of sight. The aft flight station RHC is used only on orbit. An aft sense -Z switch on panel A6 selects the line-of-sight reference about the minus Z axis (overhead windows), and the -X position selects the line-of-sight reference about the minus X axis (aft windows) in order for aft RHC commands to be correctly transformed to give the desired orbiter movement.
Several switches are located on the RHC. A backup flight system (BFS) mode button on the commander's and pilot's RHCs engage the BFS when depressed. The commander's, pilot's, and aft flight station RHCs have a two-contact trim switch that can be pushed forward or aft to add a trim rate to the RHC pitch command; pushing it left or right adds a roll trim rate. The aft RHC's trim switch is inactive. The communications switch on each RHC is a push-to-talk switch that enables voice transmission when the switch is depressed.
Each RHC contains nine transducers: three redundant transducers sense pitch deflection, three sense roll deflection and three sense yaw deflection. The transducers produce an electrical signal proportional to the deflection of the RHC. The three transducers are called channels 1, 2 and 3; the channel selected by redundancy management provides the command. Each channel is powered by a separate power supply in its associated display driver unit. Each controller is triply redundant; thus, it takes only one good signal from a controller for the controller to operate.
Each RHC has an initial dead band of 0.25 of a degree in all three axes. To move the RHC beyond the dead band, an additional force is required. When the amount of deflection reaches a certain level, called the softstop, a step increase in the force required for further deflection occurs. When a software detent position is exceeded, that RHC assumes control.
The softstop occurs at 19.5 degrees in the roll and pitch axes and at 9.5 to 10.5 degrees in the yaw axis. To reach the softstop in the roll axes, 40.95 inch-pounds of static \Jtorque\j deflection are required; 38.2 inch-pounds are needed in pitch and 7 inch-pounds in yaw.
The mechanical hardstop that can be obtained in an axis is 24.3 degrees in the roll and pitch axis and 14.3 degrees in the yaw axis.
Software normally flows from the RHCs to the flight control system through redundancy management and a SOP before it is passed to the aerojet digital autopilot.
In a nominal mission, the flight crew has manual control of the RHC during every major mode except terminal countdown. When an RHC deflection exceeds the detent in an axis, the RHC SOP generates a discrete signal that converts the RHC from the automatic mode to control stick steering, or hot stick.
However, during ascent when the ascent digital autopilot is active, a CSS pitch and/or roll/yaw mode push button light indicator on panel F2 or F4 must be depressed in order for manual inputs to be implemented into the flight control system from the commander's or pilot's RHC. When a CSS pitch or roll/yaw push button light indicator is depressed on panels F2 or F4, the white light for the push button indicator will be illuminated and that axis will be downmoded from automatic to CSS.
When the flight crew commands three-axis motion using the RHC, the GPCs process the RHC and motion sensor commands; and the flight control system interprets the RHC motions (fore and aft, right and left, clockwise and counterclockwise) as rate commands in pitch, roll and yaw and then processes the flight control law (equations) to enhance control response and stability. If conflicting commands are given, no commands result.
During orbital flight, any one of the three stations can input three-axis control commands to the flight control system. During entry and landing, the commander and pilot have two-axis (roll and pitch only) capability. Roll, pitch and yaw aerosurface deflection trim is controlled by the panel trim switches, while roll and pitch vehicle rate trim is controlled with the trim switches on the RHC. For a return-to-launch-site abort, both the commander's and pilot's RHC have three-axis capability during major mode 601 and roll and pitch during major modes 602 and 603.
The commander's RHC is powered when the flt cntlr (controller) on/off switch on panel F7 is positioned to on . The pilot's RHC is powered when the flt cntlr on/off switch on panel F8 is positioned to on. The aft RHC is powered when the flt cntlr on/off switch on panel A6 is positioned to on .
If a malfunction occurs in the commander's or pilot's RHC, the red RHC caution and warning light on panel F7 is illuminated.
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"Shuttle Translational Hand Controller",379,0,0,0
There are two translational hand controllers: one at the commander's station and one at the aft flight deck station. The commander's THC is active during orbit insertion, on orbit and during deorbit. The aft flight deck station THC is active only on orbit. The THCs are used for manual control of translation along the longitudinal (X), lateral (Y) and vertical (Z) vehicle axes using the RCS.
Each THC contains six three-contact switches, one in the plus and minus directions for each axis. Moving the THC to the right commands translation along the plus Y axis and closes three switch contacts (referred to as channels 1, 2 and 3). Redundancy management then selects the channel and provides the command.
An aft sense switch on panel A6 selects the line-of-sight reference along the minus X or minus Z axis of the orbiter for the aft THC. The aft sense switch must be in the -X position for aft windows and -Z for the overhead windows in order for the aft THC commands to be correctly transformed to give the desired orbiter movement.
The normal displacement of a THC is 0.5 of an inch from the center null position in both directions along each of the three THC axes. A force of 2 pounds is required to deflect either THC 0.5 of an inch in all axes.
The redundant signals from the forward and aft THC pass through a redundant management process and a SOP before being passed to the flight control system. If both the forward and aft THCs generate conflicting translation commands, the output translation command is given.
In what is referred to as the transition digital autopilot mode, the commander's THC is active and totally independent of the flight control orbital digital autopilot ( DAP ) push buttons on panel C3 or A6 or the position or status of the RHC. Whenever the commander's THC is out of detent plus or minus X, Y or Z, translation acceleration commands are sent directly to the RCS jet selection logic for continuous RCS thrusting periods. Rotational commands may be sent simultaneously with translation commands within the limits of the RCS jet selection logic; if both plus X and minus Z translations are commanded simultaneously, plus X translation is given priority.
The commander's THC is powered when the flt cntlr on/off switch on panel F7 is positioned to on . The aft THC is powered when the flt cntlr on/off switch on panel A6 is positioned to on .
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"Shuttle Control Stick Steering Push Button Light Indicators",380,0,0,0
The pitch auto, CSS and roll/yaw auto , CSS push button light indicators are located on panel F2 for the commander and on panel F4 for the pilot. Each push button light indicator is triply redundant. During entry, depressing a CSS push button light indicator will mode flight control to augmented manual in the corresponding axis, illuminate both CSS lights and extinguish both auto lights for that axis. Depressing a CSS push button light indicator to auto will return flight control in that axis to auto, extinguish both CSS lights and illuminate both auto lights.
During ascent, depressing any of the four CSS push button light indicators will mode flight control to augmented manual in all axes, illuminate all four CSS push button light indicators and extinguish all four auto lights. Depressing any of the four push button light indicators to auto will mode flight control to automatic, illuminate all four auto push button light indicators and extinguish the CSS push button light indicators.
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"Shuttle Rudder Pedals",381,0,0,0
There are two pairs of rudder pedals: one each for the commander and pilot. The commander's and pilot's rudder pedals are mechanically linked so that movement on one side moves the other side. When a pedal is depressed, it moves a mechanical input arm in a rudder pedal \Jtransducer\j assembly. Each RPTA contains three transducers-channels 1, 2 and 3-and generates an electrical signal proportional to the rudder pedal deflection. An artificial feel is provided in the rudder pedal assemblies.
The rudder pedals command orbiter rotation about the yaw axis by positioning the rudder during atmospheric flight. In atmospheric flight, flight control software performs automatic turn coordination; thus the rudder pedals are not used until the wings are level before touchdown.
The RPTA SOP converts the selected left and right commands from volts to degrees; selects the largest of the left and right commands for output to flight control software after applying a dead band; and if redundancy management declares an RPTA bad, sets that RPTA to zero.
The rudder pedals can be adjusted 3.25 inches forward or aft from the neutral position in 0.81-inch increments (nine positions). The breakout force is 10 pounds. A pedal force of 70 pounds is required to depress a pedal to its maximum forward or aft position.
The rudder pedals provide two additional functions unrelated to software after touchdown. Rudder pedal deflections provide nose wheel steering, and depressing the upper portion of the pedals by applying toe pressure provides braking. Differential braking may be used for nose wheel steering.
The commander's RPTA is powered when the flt cntlr on/off switch on panel F7 is positioned to on . The pilot's RPTA is powered when the flt cntlr on/off switch on panel F8 is positioned to on .
The RPTA contractor is Honeywell Inc., Clearwater, Fla.
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"Shuttle Speed Brake/Thrust Controller",382,0,0,0
There are two speed brake/thrust controllers: one on the left side of the flight deck forward on panel L2 for the commander and one on the right side of the center console on panel C3 for the pilot. The SBTCs serve two distinct functions: during ascent, the pilot's speed brake/thrust controller may be used to vary the thrust level of the three SSMEs. During entry, the commander's or pilot's speed brake/thrust controller may be used to control aerodynamic drag (hence airspeed) by opening or closing the speed brake.
Depressing a takeover switch (with three contacts) on each SBTC switches to manual control of the SSME thrust level setting (pilot's only) or speed brake position. Each SBTC contains three transducers-channels 1, 2 and 3, which produce a voltage proportional to the deflection. Redundancy management selects the output.
In the case of the SSME thrust-level setting, the top half of both spd bk/throt push button light indicators on panels F2 and F4 will be illuminated, indicating auto . Only the pilot's SBTC can be enabled for manual throttle control. The pilot depresses the takeover push button on the SBTC, causing the general-purpose computer throttle command to be frozen at its current value.
While depressing the takeover button, the pilot moves the SBTC to match the frozen GPC command. Manual control is established when the match is within 4 percent. When the match is achieved, the spd bk/throt man push button light indicator on panel F4 will be illuminated and the auto light extinguished. The takeover push button is then released. If the takeover push button is released before a match is achieved, the system reverts to GPC auto commands.
Under manual throttle command, depressing either or both push button light indicators on panel F2 or F4 will cause the system to revert to the GPC auto commands, extinguishing the pilot's man light and illuminating the auto lights on panels F2 and F4. Transferring back to auto during a return-to-launch-site abort leaves the throttle at the last-commanded manual setting until 3 g's, and the vehicle is held at 3 g's.
If the speed brake mode is in automatic and the commander or pilot wishes to control the speed brake manually, momentarily depressing the takeover push button grants control of the SBTC to the crew member who depressed the switch. The speed brake is driven to the position currently commanded by the SBTC.
The spd bk/throt man push button light indicator on panel F2 will be illuminated if the commander takes control, extinguishing the auto light. If the pilot takes control, the spd bk/throt man push button light indicator on panel F4 will be illuminated, and the commander's light will be extinguished. To place the speed brake under software control, either or both spd bk/throt push button indicators on panel F2 or F4 can be depressed, and the auto lights on panels F2 and F4 will be illuminated.
At the forward setting, the SSME thrust level is the greatest, and the speed brake is closed. Rotating the SBTC back decreases the SSME thrust level or opens the speed brake.
The SBTC SOP converts the selected SSME throttle command to a setting in percent and the selected speed brake command from volts to degrees. In addition, the SBTC SOP selects the speed brake command from the SBTC whose takeover button was depressed last. If both takeover buttons are depressed simultaneously, the commander's SBTC is given control. If redundancy management declares an SBTC bad, the command is frozen.
The commander's SBTC is powered by the flt cntlr on/off switch on panel F7 when positioned to on . The pilot's SBTC is powered when the flt cntlr on/off switch on panel F8 is positioned to on .
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"Shuttle Body Flap Switch",383,0,0,0
There are two body flap switches: one for the commander on panel L2 and one for the pilot on panel C3. Each switch is a lever-locked switch spring loaded to the center position. The body flap switches provide manual control for positioning the body flap for SSME thermal protection and for reducing elevon deflections during the entry phase.
The body flap can be switched from its automatic mode to its manual mode by moving either switch from the auto/off position to the up or down position. These are momentary switch positions; when released, the switch returns to auto/off . The white body flap man (lower half) of the push button light indicator on panel F2 or F4 is illuminated, indicating manual control of the body flap.
To regain automatic control, the body flap push button light indicator on panel F2 or F4 is depressed, extinguishing the man white light and illuminating the auto white light. The push button indicator can also be depressed to man for manual body flap control. The push button light indicator is triply redundant.
The up and down positions of each switch have two power supplies from a control bus.
If the commander and pilot generate conflicting commands, a body-flap-up command will be output to flight control because up has priority.
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"Shuttle Aerosurface Servoamplifiers",384,0,0,0
Aerosurface servoamplifiers are electronic devices that receive aerosurface commands during atmospheric flight from the flight control system software and electrically position hydraulic valves in aerosurface actuators, causing aerosurface deflections.
Each aerosurface is driven by a hydraulic actuator controlled by a redundant set of electrically driven valves (ports). There are four of these valves for each aerosurface actuator, except the body flap, which has only three. These valves are controlled by the selected ASAs.
There are four ASAs located in aft avionics bays 4, 5 and 6. Each ASA commands one valve for each aerosurface, except the body flap. ASA 4 does not command the body flap.
In addition to the command channels from the ASAs to the control valves, there are data feedback channels to the ASAs from the aerosurface actuators. Each aerosurface has four associated position feedback transducers that are summed with the position command to provide a servoloop closure for one of the four independent servoloops associated with the elevons, rudder and speed brake.
The body flap utilizes only three servoloops. The path from an ASA to its associated servovalve in the actuators and from the aerosurface feedback transducers to an ASA is called a flight control channel; there are, thus, four flight control channels, except for the body flap.
Each of the four elevons located on the trailing edges has an associated servoactuator that positions it. Each servoactuator is supplied with hydraulic pressure from the three orbiter hydraulic systems. A switching valve is used to control the hydraulic system that becomes the source of hydraulic pressure for that servoactuator.
The valve allows a primary source of pressure (P1) to be supplied to that servoactuator. If the primary hydraulic pressure drops to around 1,200 to 1,500 psig, the switching valve allows the first standby hydraulic pressure (P2) to supply that servoactuator. If the first standby hydraulic pressure drops to around 1,200 to 1,500 psig, the secondary standby hydraulic source pressure (P3) is then supplied to that servoactuator.
The yellow hyd press caution and warning light will be illuminated on panel F7 if the hydraulic pressure of system 1, 2 or 3 is below 2,400 psi and will also illuminate the red backup caution and warning alarm light on panel F7.
Each elevon servoactuator receives command signals from each of the four ASAs. Each actuator is composed of four two-stage servovalves that drive a modulating piston. Each piston is summed on a common mechanical shaft, creating a force to position a power spool that controls the flow of hydraulic fluid to the actuator power ram, controlling the direction of ram movement, thus driving the elevon to the desired position.
When the desired position is reached, the power spool positions the mechanical shaft to block the hydraulic pressure to the hydraulically operated ram, locking the ram at that position. If a problem develops within a servovalve or it is commanded to a position different than the positions of the other three within an actuator, secondary delta pressure should begin to rise to 2,200 psi.
Once the secondary delta pressure is at or above 2,200 psia for more than 120 seconds, the corresponding ASA sends an isolation command to the servovalve, opening the isolation valve, bypassing the hydraulic pressure to the servovalve, and causing its commanded pressure to the power spool to drop to zero, effectively removing it from operation.
The pressure differential is sensed by a primary linear differential pressure \Jtransducer\j across the modulating piston when the respective FCS channel switch on panel C3 is in auto . This automatic function prevents excessive transient motion to that aerosurface, which could result in loss of the orbiter due to slow manual redundancy.
The FCS channel yellow caution and warning light on panel F7 will be illuminated to inform the flight crew of a failed channel. A red FCS saturation caution and warning light on panel F7 will be illuminated if one of the four elevons is set at more than plus 15 degrees or less than minus 20 degrees.
There are four FCS channel switches on panel C3- FCS channels 1, 2, 3 and 4; each has an override, auto and off position. The switch for a channel controls the channel for the elevons, rudder/speed brake and body flap, except channel 4, which has no body flap commands. When an FCS channel switch is in auto and that channel was bypassed, it can be reset by positioning the applicable switch to override . When an FCS channel switch is positioned to off , that channel is bypassed.
In each elevon servoactuator ram, there are four linear ram position transducers and four linear ram secondary differential pressure transducers. The ram linear transducers provide position feedback to the corresponding servoloop in the ASA, which is then summed with the position command to close the servoloop.
This feedback is then summed with the elevon ram linear secondary differential pressures to develop an electrohydraulic valve drive current that is proportional to the error signal in order to position the ram. The maximum elevon deflection rate is 20 degrees per second.
During ascent, the elevons are deflected to reduce wing loads caused by rapid acceleration through the lower atmosphere. In this scheme, the inboard and outboard elevons are deflected together. By the time the vehicle reaches approximately Mach 2.5, the elevons have reached a null position, where they remain. This is accomplished by the initialized-loaded program.
The rudder/speed brake, which consists of upper and lower panels, is located on the trailing edge of the orbiter's vertical stabilizer. One servoactuator positions the panels together to act as a rudder; another opens the panels at the rudder's flared end so it functions as a speed brake.
The rudder and speed brake servoactuator receives four command signals from the four ASAs. Each servoactuator is composed of four two-stage servovalves that function like those of the elevons. The exception is that the rudder's power spool controls the flow of hydraulic fluid to the rudder's three reversible hydraulic motors and the power spool for the speed brake controls the flow of hydraulic fluid in the speed brake's three hydraulic reversible motors.
Each rudder and speed brake hydraulic motor receives hydraulic pressure from only one of the orbiter's hydraulic systems. Each hydraulic motor has a hydraulic brake. When the motor is supplied with hydraulic pressure, the motor's brake is released. When the hydraulic pressure is blocked to that hydraulic motor, the hydraulic brake is applied, holding that motor and the corresponding aerosurface at that position.
The three hydraulic motors provide output to the rudder differential gearbox, which is connected to a mixer gearbox that drives rotary shafts. These rotary shafts drive four rotary actuators, which position the rudder panels.
The three speed brake hydraulic motors provide power output to the speed brake differential gearbox, which is connected to the same mixer gearbox as that of the rudder. This gearbox drives rotary shafts, which drive the same four rotary actuators involved with the rudder. Within each of the four rotary actuators, planetary gears blend the rudder positioning with the opening of the rudder flared ends.
There are four rotary position transducers on the rudder differential gearbox output and one differential linear position \Jtransducer\j in each rudder servoactuator. The rotary position transducers provide position feedback to the corresponding servoloop in the ASA. The feedback is summed with the linear differential pressures that develop the electrohydraulic valve drive current in proportion to the error signal in order to position the rudder.
There are also four rotary position transducers on the speed brake differential gearbox output and one differential linear pressure \Jtransducer\j in each speed brake servoactuator.
The rotary position transducers provide position feedback to the corresponding servoloop in the ASA, which is summed with the position command to close the servoloop. These are then summed with the linear differential pressures that develop the electrohydraulic valve drive current in proportion to the error signal to position the speed brake.
If a problem occurs in one of the four rudder or speed brake servoactuator channels, the corresponding linear differential pressure \Jtransducer\j will cause the corresponding ASA to signal a \Jsolenoid\j isolation valve to remove the pressure from the failed channel and bypass it if that FCS channel switch is in auto . The FCS channel switches' override and off positions and the FCS channel caution and warning light function the same as for the elevons.
The hyd press light indicates a hydraulic failure. The rudder deflection rate is a maximum of 14 degrees per second. The speed brake deflection rate is approximately 10 degrees per second. If two of the three hydraulic motors fail in the rudder or speed brake, about half the design speed output will result from the corresponding gearbox due to its velocity summary nature.
Three servoactuators at the lower aft end of the fuselage are used to position the body flap; each is supplied with hydraulic pressure from an orbiter hydraulic system and has a solenoid-operated enable valve controlled by one of the three ASAs (the fourth ASA is not used for the body flap commands). Each solenoid-operated enable valve supplies hydraulic pressure from one orbiter hydraulic system to a corresponding solenoid-operated pilot valve, which is, in turn, controlled by one of the three ASAs.
When the individual pilot valve receives a command signal from its corresponding ASA, it positions a common mechanical shaft in the control valve, allowing hydraulic pressure to be supplied to the hydraulic motors (normally one pilot valve is enabled and moves the other two). The hydraulic motors are reversible, allowing the body flap to be positioned up or down. The hydraulic brake associated with each hydraulic motor releases the hydraulic motor for rotation.
When the desired body flap position is reached, the control valves block the hydraulic pressure to the hydraulic motor and apply the hydraulic brake, holding that hydraulic motor at that position. Each hydraulic motor provides the power output to a differential gearbox, which drives a rotary shaft, and four rotary actuators, which position the body flap. The rotary position \Jtransducer\j associated with each rotary actuator provides position feedback to the ASAs; the fourth ASA is used to provide position feedback to the flight control system software.
If the FCS channel switches are in auto , the ASAs will isolate a body flap channel through the solenoid-operated enable valve if the corresponding solenoid-operated pilot valve malfunctions or the control valve associated with the pilot valve does not provide the proper response and allows the hydraulic pressure fluid to recirculate. The FCS channel switches and FCS channel caution and warning light function the same as for the elevons. If the hydraulic system associated with the hydraulic motor fails, the remaining two hydraulic motors will position the body flap, and the hyd press caution and warning light will be illuminated. The body flap deflection rate is approximately 4.5 degrees per second.
Each ASA is hard-wired to a flight MDM. Flight control commands originate from guidance software or from controllers. These inputs go to the flight control software, where they are augmented and then routed to the ASAs.
There are several subsystem operating programs associated with the ASA commands and data. The SOPs convert elevon, rudder and speed brake commands from flight control software from degrees to millivolts; set commands to body flap valves based on an enable command from body flap redundancy management and up/down commands from flight control; convert position feedback to degrees for the elevons, rudder, speed brake and body flap; compute elevator position from elevon position feedbacks; calculate body flap and speed brake deflections as percentages; calculate elevon and rudder positions for display on the surface position indicator; monitor the FCS channel switches and if any are positioned to override, set the override command for that ASA; monitor hydraulic system pressures for failures; and rate-limit aileron and elevator commands according to the number of failures.
Each ASA is mounted on a cold plate and cooled by the Freon-21 coolant loops. Each is 20 inches long, 6.4 inches high and 9.12 inches wide and weighs 30.2 pounds.
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"Shuttle Digital Autopilot",385,0,0,0
The digital autopilot is the heart of flight control software. It is composed of several software modules that interpret maneuver requests; compare them to what the vehicle is doing; and generate commands for the appropriate effectors, as needed, to satisfy the requests. There are different DAPS for different flight phases and various modes and submodes within each.
At main engine cutoff, the transition digital autopilot becomes active, sending attitude hold commands to the reaction control system. External tank separation is automatically commanded 18 seconds after main engine cutoff, and the transition DAP immediately sets commands to fire the orbiter's minus Z RCS jets, causing the orbiter to translate in the minus Z direction. When a rate of negative 4 feet per second is reached, RCS fire commands are removed. The transition DAP is used from MECO until transition to OPS 2 (on orbit).
In the transition DAP mode, the external tank separation module initialized by the external tank separation sequencer compares Z-delta velocities from the DAP attitude processor with an initialized-loaded desired Z-delta velocity. Before this value is reached, the transition DAP and steering processor send commands to the RCS jet selection logic, which uses a table lookup technique for the primary RCS jets and commands 10 RCS jets to fire in the plus Z direction. When the desired Z-delta velocity is reached, the translation command is set to zero. Rotation commands are permitted during the external tank separation sequence.
The RCS jet selection module receives both RCS rotation and translation commands and outputs 75 discretes to the primary RCS jets to turn the 38 jets on or off by applying the table lookup technique.
The RCS reaction jet driver forward and aft assemblies provide the turn-on/turn-off jet selection logic signals to the RCS jets. There is also a driver redundancy management program that permits only ''good'' RCS jets to be turned on. The RCS jet yellow caution and warning light indicates a failed-on, failed-off or leaking RCS jet.
The 19 RCS jets thrusting in the plus or minus Z direction provide Z translation and roll or pitch rotation control and are considered independent of other axes. The 12 RCS jets thrusting in the plus or minus Y direction provide Y translation and yaw rotation only. The seven RCS jets thrusting in the plus or minus X direction provide X translation only.
Insertion flight control is accomplished using the transition DAP. The transition DAP uses commands from guidance for automatic maneuvers to orbital maneuvering system burn attitude using RCS jets. During OMS-1 and OMS-2 (or only OMS-1 in a direct insertion), the transition DAP uses the OMS engines and RCS jets, as required. The transition DAP also receives commands from the flight crew through the commander's THC and the commander's or pilot's RHC.
The transition DAP then takes these commands and converts them into appropriate RCS commands. The transition DAP monitors the resultant attitude and attitude rates and sends the necessary commands to achieve the targeted attitude and attitude rate within premission-specified dead bands.
The transition DAP reconfiguration logic controls the moding, sequencing and initialization of the control law modules and sets gains, dead bands and rate limits. The steering processor is the interface between the guidance or manual steering commands and the transition DAP. The steering processor generates commands to the RCS processor, which generates the RCS jet command required to produce the commanded \Jspacecraft\j translation and rotation using attitude and rotational rate signals or translation or rotation acceleration commands.
The flight crew interfaces with the transition DAP through the forward RHCs and THC and indirectly through entries to the OMS mnv exec CRT displays and the DAP panel push button light indicators on panel C3.
In the transition DAP, the commander's THC is active and totally independent of the DAP push button light indicators or RHC position or status. Whenever the commander's THC is out of detent plus or minus X, Y or Z, translation acceleration commands are sent directly to the RCS jet selection logic for continuous RCS jet firing. Rotational commands may be sent simultaneously with translation commands within the limits of the RCS jet selection logic; if both plus X and minus Z translations are commanded simultaneously, plus X translation receives priority.
For rotations, the flight crew can select either automatic or manual control through the use of the DAP panel push button light indicators or by moving the RHC. In manual, the capability exists to rotate in any axis in a pulse mode, in which each RHC deflection results in a single burst of jet fire, or in a discrete rate mode, in which RHC deflection results in a specified rate being commanded in that axis for the entire time the RHC is deflected.
It is also possible to go to a free drift mode, in which no RCS jets are fired, or to an attitude hold mode, in which the DAP sends commands to maintain the current attitude with null rates within premission-specified dead bands. Also, if the RHC is deflected beyond a certain point, continuous RCS jet firings will result. In translation, movement of the THC results in continuous jet firings.
The orbital maneuvering system processor generates OMS engine gimbal actuator thrust vector control commands to produce the desired spacecraft/engine relationship for the commanded thrust direction.
For the OMS thrusting period, the orbital state (position and vector) is produced by navigation incorporating inertial measurement unit delta velocities during powered and coasting flight. This state is sent to guidance, which uses target inputs through the CRT to compute thrust direction commands and commanded attitude for flight control and thrusting parameters for CRT display.
Flight control converts the commands into OMS engine gimbal angles (thrust vector control) for an automatic thrusting period. OMS thrust vector control for normal two-engine thrusting is entered by depressing the orbital DAP auto push button light indicator with both RHCs within software detents.
OMS manual thrust vector control for both OMS engines is entered by depressing the orbital man DAP push button light indicator or by moving the commander's or pilot's RHC out of detent; the flight crew supplies the rate commands to the TVC system instead of guidance.
The manual RHC rotation requests are proportional to RHC deflections and are converted into gimbal angles. OMS thrust in either case is applied through the \Jspacecraft\j's center of gravity.
Before OMS ignition, the \Jspacecraft\j is maneuvered to the OMS ignition attitude by using the RHC and RCS jets, reducing transient fuel losses. Normally, the first OMS thrusting period raises the orbiter's low elliptical orbit after the external tank is jettisoned, and the second OMS thrusting places the \Jspacecraft\j into the circular orbit designated for that mission. For orbital maneuvers that use the OMS, any delta velocities greater than 6 feet per second use the two OMS engines. Some missions use a direct insertion, requiring only one OMS thrusting period.
Automatic thrust vector control for one OMS engine is identical to that for two, except that the RCS processor is responsible for roll control. Single-OMS-engine thrust is also through the \Jspacecraft\j's center of gravity, except when pitch or yaw rate commands are non-zero. If the left or right OMS engine fails, an OMS TVC red light on panel F7 will be illuminated.
Since an OMS cutoff is based on time rather than velocity, a velocity residual may exist following the cutoff. The residual is zeroed by the RCS through the THC.
The orbital flight control software includes an RCS DAP, an OMS TVC DAP and an attitude processor module to calculate vehicle attitude as well as logic to govern the selection of a DAP. The attitudes calculated by the attitude processor are displayed on the attitude display indicator along with another crew display, universal pointing, which is available in major mode 201 (orbit coast). The vehicle attitude is used by the DAP to determine attitude and rate errors.
The only time the RCS DAP is not used in OPS 2 is during an OMS burn. This DAP controls vehicle attitudes and rates through the use of RCS jet fire commands. Either the larger primary jets or the less powerful vernier jets are used for rotational maneuvers, depending on whether norm or vern is selected on the panel C3 orbital DAP panel. The choice of primary or vernier thrusters depends on fuel consumption considerations and how quickly the vehicle needs to be maneuvered to satisfy a mission objective.
The rotation rates and dead bands, translation rate and certain other DAP options can be changed by the flight crew during the orbit phase using the DAP CRT display. The flight crew can load the DAP with these options in two ways: one option set may be accessed by depressing the DAP A push button on the orbital DAP panel, the other by depressing the DAP B push button.
For convenience, each planned DAP configuration is given a number and is referred to by that number and the DAP used to access it. Typically, the DAP A configurations will have larger dead bands and higher rates than the DAP B configurations. The wide dead bands are used to minimize fuel usage, while the tight dead bands allow greater precision in executing maneuvers or in holding attitude.
The RCS DAP can operate in both an automatic and a manual rotation mode, depending on whether the flight crew selects the auto or man push button light indicators on the orbital DAP panel. The manual mode is also accessed when the RHC is moved out of its detent (neutral) position. In both the automatic and manual modes, the rotation rate is controlled by the selection of DAP A or B and the information loaded in the DAP config display.
In addition, in automatic, the DAP determines the required attitude to be achieved from universal pointing and then computes the RCS jet fire commands necessary to achieve these requirements within the current set of dead bands. In the manual rotation mode, the RCS DAP converts flight crew inputs with any of the three RHCs to RCS jet fire commands, depending on whether pulse, disc rate or accel is selected on the orbital DAP panel.
Simply, when pulse is selected, a single burst of jet fire is produced with each RHC deflection. The resultant rotational rate is specified on the DAP config display. When disc rate is selected, jet firings continue to be made as long as the RHC is out of detent in order to maintain the rotational rate specified on the DAP config display. When accel is selected, continuous jet firings are made as long as the RHC is out of detent.
Another manual RCS DAP mode, local vertical/local horizontal, is used to maintain the current attitude with respect to the rotating LVLH reference frame. It is selected through the LVLH push button on the orbital DAP panel.
The RCS DAP has only a manual translation capability, which is executed through the forward or aft THC. Only the primary RCS jets are used. Deflections of the THC result in RCS jet firings based on the transition DAP mode push button light indicator selected on the orbital DAP panel. Pulse results in a single burst of jet fire. Norm results in continuous jet firings with a specified subset of the available jets. High results in all up-firing jets firing continuously in a Z translation.
And low enables a special technique that accomplishes a Z translation using the forward- and aft-firing RCS jets in order to not fire directly toward a target (avoiding plume impingement and contamination of a target payload).
The OMS thrust vector control DAP is available when an OMS burn is executed in major mode 202 (maneuver execute) through the orbit mnvr exec display. The TVC DAP uses the guidance-generated velocity requirements and converts these into the appropriate OMS gimbal commands to achieve this target, assuming auto is selected on the orbital DAP panel.
It generates the OMS fire commands; the OMS shutdown commands; and, if necessary due to OMS engine failure, required RCS commands to maintain attitude control. If manual is selected, the TVC DAP uses inputs from the RHC to control attitude during the burn.
As with the transition DAP, there are many subtleties in the operation of the orbital DAP.
There are 24 orbital DAP push button light indicators on panels C3 and A6. Assuming no electrical, computer bus (MDM) or hardware failures that could affect the operation of the push button light indicators, inputs made to one panel will be reflected in the configuration of the other panel. All of the push button light indicators are active in the orbital DAP, but only a subset of these are operational in the transition DAP or when the backup flight system is engaged.
None of the push button light indicators are operational during ascent or entry. As with other aft flight deck controls, aft panel A6 push button light indicators are only operational while the vehicle is on orbit.
The orbital DAP select, control, RCS jets, and manual mode translation and rotation push button light indicators are illuminated by flight control when that mode is implemented in the flight control system in the transition or orbital DAP.
Automatic rotation commands are supplied by the universal pointing processor. The universal pointing processor, through the operational sequence display, provides three-axis automatic maneuver, tracking local vertical/local horizontal about any body vector, rotating about any body vector at the DAP discrete rate, and stopping any of these options and commanding attitude hold. The parameters of these maneuvers are displayed in current attitude, required attitude, attitude error and body rates. Either total or DAP attitude errors may be selected for display on the attitude display indicator error needles.
The automatic maneuver option is used to calculate a commanded vehicle attitude and angular rate or to hold a vehicle attitude. The desired inertial commanded and rotation attitude is input into the operational sequence display in pitch, yaw and roll. When the maneuver option is selected, universal pointing sends the required attitude increment and body rate to flight control, and flight control performs the maneuver when the DAP is in automatic.
The automatic rotation option calculates a rotation about a desired body axis. This option is used for passive thermal control, also known as barbecue. Pitch and yaw body components of the desired rotational axis are first input. The orbiter is maneuvered automatically or manually so that the rotational axis is oriented properly in inertial space. When the rotational option is selected, universal pointing will calculate the required body attitude and send it to flight control. Flight control performs the maneuver if the DAP is in automatic.
The automatic stop/attitude hold option cancels universal pointing processing of the automatic maneuver, rotation or LVLH options. When the stop/attitude hold option is selected, universal pointing will cancel the processing of the maneuver options and send the current attitude to flight control. When flight control is in automatic, attitude hold will be initiated about the current attitude.
The nine rotation push button light indicators on panel C3 or A6 are meaningless when auto is selected. The transition or orbital DAP can be switched to manual by depressing the man push button light indicator or by positioning the RHC out of detent while operating in the automatic mode. In the manual mode, the man light is illuminated and the auto light is extinguished. When the manual mode is selected, the nine rotation roll, pitch and yaw push button light indicators determine the kind of control the RHC will provide.
There are three RCS rotation submodes available in the orbital DAP: automatic, manual or LVLH. LVLH is not available in the transition DAP; thus the choice in transition DAP is automatic or manual. Within each of these submodes are submodes that depend on conditions, such as the DAP push button light indicator configuration and the RHC state. Manual RCS translation modes are independent.
Depressing the RCS jets norm push button enables the primary RCS jets for rotational and translational firings and disables the vernier RCS jets. The RCS jets norm light is illuminated and the RCS jets vern light is extinguished.
Depressing the RCS jets vern push button enables the vernier RCS jets for rotational (not translational) RCS firings and disables the primary RCS jets. The RCS jets vern light is illuminated and the RCS jets norm light is extinguished.
The manual mode rotation, roll, pitch and yaw push button light indicators are used on orbit and during transition DAP operations. The RCS rotation is selected on an axis-by-axis basis. For example, pitch could be in discrete rate, yaw in acceleration and roll in pulse.
Depressing the rotation disc rate push button for an axis causes the appropriate primary or vernier RCS jets to fire to attain a predetermined rotational rate in that axis while the RHC is out of detent. When the RHC is returned to detent, the rate is nulled and attitude hold is re-established. When depressed, the disc rate push button light indicator is illuminated and the accel or pulse indicators are extinguished for that axis. Rotation units are in degrees per second.
Depressing the rotation accel push button for an axis causes the primary or vernier jets to fire when the RHC is out of detent, producing a moment with the same sense as the RHC deflection in that axis. The jets remain on as long as the RHC is out of detent and shut off when the RHC is returned to detent, allowing attitude to drift freely.
When depressed, the accel push button light indicator is illuminated and the disc rate or pulse push button light indicators are extinguished for that axis. The accel push button light indicator is not functional in the transition DAP, but the same effect can be achieved by taking the RHC beyond the softstop in any mode.
Depressing the rotation pulse push button for an axis causes the primary or vernier jets to fire for preset increments in response to each deflection of the RHC in that axis. No further firing occurs until the RHC is returned to detent and is again deflected, allowing attitude to drift. When depressed, the pulse push button light indicator is illuminated and the disc rate or accel push button light indicators are extinguished for that axis. The units for rotation pulse are in degrees per second, and the resulting vehicle rate will be a product of the pulse size and number of pulses.
Depressing the LVLH push button light indicator places the DAP in a manual mode. With rotation in discrete rate for all three axes and the RHC in detent, LVLH hold will be maintained and the LVLH push button light indicator will remain on.
The manual mode translation push button light indicators are used on orbit, and mixed modes are permitted on an axis-by-axis basis. Except for the case of the automatic minus Z external tank separation firing, all RCS translations must be performed manually. In transition DAP, none of the translation push button light indicators will be illuminated.
Depressing the translation high push button light indicator causes all nine up-firing primary plus Z jets to be fired as long as the THC is held out of detent in that axis. For a minus Z translation, the high Z mode functions identically to the normal Z and low Z. The Z high push button light indicators are illuminated and the low Z , Z norm and Z pulse push button light indicators are extinguished.
Depressing the low Z push button inhibits all up-firing jets in order to prevent plume damage to payloads or injury to extra-vehicular activity crew members. If a plus Z translation is requested, plus X and minus X jets are fired simultaneously, producing a downward translation because the X jets are oriented in such a way that they have small plus Z thrust components.
Minus Z uses the six down-firing jets with the selected high or norm Z. The jets continue to fire as long as the THC is out of detent. The low Z push button light indicators are illuminated and the Z high push button light indicator is extinguished.
Depressing the norm push button light indicator for an axis causes the appropriate primary jets in that axis to be fired for as long as the THC is out of detent. This mode is used for most RCS translations and propellant dumps. The norm push button light indicators are illuminated and the pulse and Z high indicators are extinguished.
Depressing the pulse push button light indicator for an axis causes the appropriate primary jets to fire for a preset increment in response to each deflection of the THC. The firing duration is a function of pulse size. No further firing occurs until the THC is returned to detent and is again deflected. The pulse push button light indicators are illuminated and the norm or Z high indicators are extinguished.
Following insertion, a \Jspacecraft\j's orbit is essentially fixed, although effects, such as venting and atmospheric drag, can cause orbital perturbations. OMS or RCS thrusting periods can be used to correct or modify the orbit, as required, for mission operations. The direction and magnitude of the thrusting period, as well as the time of application, determine the resulting shape of the orbit.
A posigrade thrusting period increases the speed at the point of application and will raise every point of the orbit except the thrusting point. A retrograde thrusting period decreases the speed at the point of application and will lower every point of the orbit except the thrusting point.
An out-of-plane thrusting period alters the inclination of the \Jspacecraft\j's orbital plane. It does not change the vehicle's period of orbit or height above the Earth.
A radial thrusting period is one in which the thrust is applied in a direction perpendicular to the \Jspacecraft\j's velocity vector and in the vehicle's orbital plane. With the vehicle in a circular orbit, a radial thrusting period would be applied along the radius vector either toward or away from the center of the Earth.
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"Shuttle Component Locations",387,0,0,0
The various black boxes of the avionics systems are located in the crew compartment flight deck, middeck avionics bays and the aft avionics bays.
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"Shuttle, Dedicated Display Systems",388,0,0,0
The dedicated displays provide the flight crew with information required to fly the vehicle manually or to monitor automatic flight control system performance. The data on the dedicated displays may be generated by the navigation or the flight control system software or more directly by one of the navigation sensors. The dedicated displays are located in front of the commander's and pilot's seats and on the aft flight deck panel by the aft-facing windows.
The dedicated displays are the attitude director indicators on panels F6, F8 and A1; horizontal situation indicators on panels F6 and F8; alpha Mach indicators on panels F6 and F8; altitude/vertical velocity indicators on panels F6 and F8; surface position indicator on panel F7; reaction control system activity lights on panel F6; g-meter on panel F7; and head-up display on the glare-shield in front of the commander's and pilot's seats.
Not all of the dedicated displays are available in every operational sequence or major mode. Their availability is related to the requirements of each flight phase.
The display driver unit is an electronic mechanism that connects the general-purpose computers and the primary flight displays. The DDU receives data signals from the computers and decodes them to drive the dedicated displays. The unit also provides dc and ac power for the ADIs and the rotational and translational hand controllers. It contains logic for setting flags on the dedicated instruments for such items as data dropouts and failure to synchronize. The orbiter contains three DDUs: one at the commander's station, one at the pilot's station and one at the aft station.
There are three display driver units. One interfaces with the ADI, HSI, AVVI and AMI displays on panel F6 at the commander's station, and the second interfaces with the same instruments on panel F8 at the pilot's station. The third unit interfaces with the ADI at the aft flight station.
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"Shuttle, Attitude Director Indicator",389,0,0,0
The commander's and pilot's ADIs are supported throughout the mission, while the aft ADI is active only during orbital operations. They give the crew attitude information as well as attitude rate and attitude errors, which can be read from the position of the pointers and needles. Each ADI has a set of switches by which the crew can select the mode or scale of the readout. The commander's switches are located on panel F6, the pilot's on panel F8 and the aft switches on panel A6.
The orbiter's attitude is displayed to the flight crew by an enclosed ball (sometimes called the eight ball) that is gimbaled to represent three degrees of freedom. The ball, covered with numbers indicating angle measurements (a zero is added as the last digit of each), moves in response to software-generated commands to depict the current orbiter attitude in terms of pitch, yaw and roll.
The ADI attitude select switch determines the unit's frame of reference: inrtl (inertial), LVLH (local vertical/local horizontal), and ref (reference). The inrtl position allows the flight crew to view the orbiter's attitude with respect to the inertial reference frame, useful in locating stars. The LVLH position shows the orbiter's attitude from an orbiter-centered rotating reference frame with respect to Earth. The ref position is primarily used to see the orbiter's attitude with respect to an inertial reference frame defined when the flight crew last depressed the att ref push button.
It is useful when the crew flies back to a previous attitude or monitors a maneuvering system thrusting period for attitude excursions. The two forward switches are active during ascent, orbital and transition flight phases but have no effect during entry, the latter part of a return to launch site or phases when the backup flight system is driving the ADIs. The aft switch, like the aft ADI, is operational only in orbit.
Each attitude director indicator has a set of three rate pointers that provide a continuous readout of vehicle body rotational rates. Roll, pitch and yaw rates are displayed on the top, right and bottom pointers, respectively. The center mark on the graduated scale next to the pointers shows zero rates, while the rest of the marks indicate positive or negative rates.
The adi rate switch for each indicator unit determines the magnitude of full-scale deflection. When this switch is positioned to high (the coarsest setting), the pointer at the end of the scale represents a rotation rate of 10 degrees per second. When the switch is positioned to med, a full-range deflection represents 5 degrees per second. In the low position (the finest setting), a pointer at either end of the scale is read at a rate of 1 degree per second. These pointers are ''fly to'' in the sense that the rotational hand controller must be moved in the same direction as the pointer to null a rate.
ADI rate readings are independent of the selected attitude reference. During ascent, the selected rates come directly from the solid rocket booster or orbiter rate gyros to the ADI processor for display on the rate pointers. During entry, only the pitch rate follows the direct route to the ADI display. The selected roll and yaw rates first flow through flight control, where they are processed and output to the ADI as stability roll and yaw rates. (This transformation is necessary because, in aerodynamic flight, control is achieved about stability axes, which in the cases of roll and yaw differ from body axes.)
Three needles on each attitude director indicator display vehicle attitude errors. These needles extend in front of the ADI ball, with roll, pitch and yaw arranged just as the rate pointers are. Like the rate indicators, each error needle has a background scale with graduation marks that allow the flight crew to read the magnitude of the attitude error. The errors are displayed with respect to the body-axis coordinate system and, thus, are independent of the selected reference frame of the attitude display.
The ADI error needles are driven by flight control outputs that show the difference between the required and current vehicle attitude. These needles are also ''fly to,'' meaning that the flight crew must maneuver in the direction of the needle to null the needle. For example, if the pitch error needle points down, the flight crew must manually pitch down to null the pitch attitude error. The amount of needle deflection indicating the number of degrees of attitude error depends upon the adi error switch for each ADI. In the high position, the error needles represent 10 degrees, med represents 5 degrees and low represents 1 degree.
At the aft flight station on panel A6, the aft sense switch allows the flight crew to use the aft ADI, RHC and translational hand controller in a minus X or minus Z control axis sense. These two options of the aft ADI and hand controllers correspond to the visual data out of the aft viewing (negative X) or overhead viewing (negative Z) windows.
Each ADI has a single flag labeled off on the left side of the display whenever any attitude drive signal is invalid. There are no flags for the rate and error needles; these indicators are driven out of view when they are invalid.
The horizontal situation indicator for the commander and pilot displays a pictorial view of the vehicle's position with respect to various navigation points and shows a visual perspective of certain guidance, navigation and control parameters, such as directions, distances and course/glide path deviation. The flight crew uses this information to control or monitor vehicle performance. The HSIs are active during the entry and landing and ascent/RTLS phases.
Each HSI provides an independent source to compare with ascent and entry guidance, a means of assessing the health of individual navigation aids during entry and information needed by the flight crew to fly manual ascent, RTLS and entry.
Each HSI displays magnetic heading (compass card), selected course, runway magnetic course, course deviation, glide slope deviation, primary and secondary bearing, primary and secondary range, and flags to indicate validity.
Each HSI consists of a case-enclosed compass card measuring zero to 360 degrees. At the center of the compass card is an \Jaircraft\j symbol, fixed with respect to the case and about which the compass card rotates.
The magnetic heading (the angle between magnetic north and vehicle direction measured clockwise from magnetic north) is displayed by the compass card and read under the lubber line located at the top of the indicator dial. (A lubber line is a fixed line on a compass aligned to the longitudinal axis of the craft.) The compass card is positioned at zero degrees (north) when the heading input is zero. When the heading point is increased, the compass card rotates counterclockwise.
The course pointer is driven with respect to the HSI case rather than the compass card. Therefore, a course input (from the DDU) of zero positions the pointer at the top lubber line, regardless of compass card position. To position the course pointer correctly with respect to the compass card scale, the software must subtract the vehicle magnetic heading from the runway \Jazimuth\j angle (corrected to magnetic north).
As this subtraction is done continuously, the course pointer appears to rotate with the compass card, remaining at the same scale position. An increase in the angle defining runway course results in a clockwise rotation of the course pointer.
Course deviation is an angular measurement of vehicle displacement from the extended runway centerline. On the HSI, course deviation is represented by the deflection of the deviation bar from the course pointer line. Full scale on the course deviation scale is plus or minus 10 degrees in terminal area energy management and plus or minus 2.5 degrees during approach and landing.
The course deviation indicator is driven to zero during entry. When the course deviation input is zero, the deviation bar is aligned with the end of the course pointer. With the pointer in the top half of the compass card, an increase in course deviation to the left (right) causes the bar to deflect the right (left). Therefore, the course deviation indicator is a fly-to indicator for flying the vehicle to the extended runway centerline. Software processing also ensures that the CDI remains fly to, even when the orbiter is heading away from the runway.
In the TAEM example, at a range of 9 nautical miles (10 statute miles), the CDI would read about 7.5 degrees, with the extended runway centerline to the right of the orbiter. In course deviation \Jgeometry\j, if the orbiter is to the left of the runway, it must fly right (or if the orbiter is to the right of the runway, it must fly left) to reach the extended runway centerline. The corresponding course deviation bar would deflect to the right (or to the left in the latter case). The reference point at the end of the runway is the microwave landing system station. The sense of the CDI deflection is a function of vehicle position rather than vehicle heading.
Glide slope deviation, the distance of the vehicle above or below the desired glide slope, is indicated by the deflection of the glide slope pointer on the right side of the HSI. An increase in glide slope deviation above (below) the desired slope deflects the pointer downward (upward); the pointer is a fly-to indicator. In the HSI example, the pointer shows the vehicle to be below the desired glide slope by about 4,000 feet (in TAEM, each dot represents 2,500 feet).
The ''desired glide slope'' is actually only a conceptual term in HSI processing. At any instant, glide slope deviation is really the difference between the orbiter altitude and a reference altitude computed in the same fashion as the guidance reference altitude. Also included in the reference altitude equation are factors for a ''heavy orbiter'' and for high winds.
The GSI computation is not made during entry or below 1,500 feet during approach and landing; therefore, the pointer is stowed and the GSI flag is displayed during those intervals.
The primary and secondary bearing pointers display bearings relative to the compass card. These bearings are angles between the direction to true or magnetic north and to various reference points as viewed from the orbiter. For the bearing pointers to be valid, the compass card must be positioned in accordance with vehicle heading input data.
When the bearing inputs are zero, the pointers are at the top lubber line, regardless of compass card position. Like the course pointer, the bearing pointer drive commands are developed by subtracting the vehicle heading from the calculated bearing values. This allows the pointers to be driven with respect to the HSI case but still be at the correct index point on the compass card scale. When the bearing inputs are increased, the pointers rotate clockwise about the compass card. The pointer does not reverse when it passes through 360 degrees in either direction.
For example, if the primary bearing is 190 degrees and the secondary bearing is 245 degrees, the bearing reciprocals are always 180 degrees from (opposite) the pointers. The definition of primary and secondary bearing varies with the flight regime.
The HSI is capable of displaying two four-digit values in the upper left and right side of its face. These numbers are called primary and secondary range, respectively. Each display ranges from zero to 3,999 nautical miles (4,602 statute miles). While their meaning depends on the flight regime, both numbers represent range in nautical miles from the vehicle to various points relative to the primary and secondary runways. In the HSI example, the primary range is 9 nautical miles (10 statute miles); the barberpole in the secondary range slot is an invalid data indication.
The HSI has four flags- off, brg (bearing), GS (glide slope) and CDI-and two barberpole indications that can respond to separate DDU commands, identifying invalid data. Off indicates that the entire HSI display is invalid because of insufficient power. Brg indicates invalid course, primary bearing, and/or secondary bearing data. GS indicates invalid glide slope deviation. CDI indicates invalid course deviation data. Barberpole in the range slots indicates invalid primary or secondary range data.
When the HSI source switch is in nav , the entire HSI display is driven by navigation-derived data from the orbiter state vector. This makes the HSI display dependent on the same sources as the navigation software (IMU, selected air data, selected navigational aids), but the display is independent of guidance targeting parameters. As stated previously, when the TACAN/nav/MLS switch is in the nav position, the source 1, 2, 3 switch is not processed.
The TACAN or MLS position of the source switch should be used only when TACAN or MLS data are available. TACAN data can be acquired in Earth orbit but would be unavailable during blackout; therefore, TACAN is generally not selected until acquisition after blackout. MLS has a range of 20 nautical miles (23 statute miles) and is normally selected after the orbiter is on the heading alignment cylinder.
The glide slope deviation pointer is stowed when the entry mode is selected and the flag is displayed. The GSI in TAEM indicates deviation from guidance reference attitude in plus or minus 5,000 feet. The GSI in approach indicates guidance reference altitude for approach and landing in plus or minus 1,000 feet; it is not computed below 1,500 feet and the flag deploys.
In the entry mode, the compass card heading indicates the magnetic heading of the vehicle's relative velocity vector. In TAEM and approach, the compass card indicates magnetic heading of the body X axis.
In the entry mode, the course deviation indicator is a valid software zero with no flag. In TAEM, the CDI indicates the deviation from the extended runway centerline, plus or minus 10 degrees. In approach, the CDI indicates the deviation from the extended runway centerline, plus or minus 2.5 degrees.
In the entry mode, the primary bearing indicates the spherical bearing to way point 1 for the nominal entry point at the primary landing runway. The secondary bearing indicates the spherical bearing to WP-1 for the NEP to the secondary landing runway.
In TAEM, the primary bearing indicates the bearing to WP-1 on selected HAC for the primary runway. The secondary bearing indicates the bearing to the center of the selected HAC for the primary runway. In approach, the primary and secondary bearings indicate the bearing to WP-2 at the primary runway.
In the entry mode, the primary range indicates the spherical surface range to WP-2 on the primary runway via WP-1 for NEP. The secondary range indicates the spherical surface range to WP-2 on the secondary runway via WP-1 for NEP. In TAEM, the primary range indicates the horizontal distance to WP-2 on the primary runway via WP-1. The secondary range indicates the horizontal distance to the center of the selected HAC for the primary runway. In approach, the primary and secondary ranges indicate the horizontal distance to WP-2 on the primary runway.
During ascent major modes 102 and 103 (first and second stage) and RTLS, the horizontal situation indicator provides information about the target insertion orbit. The compass card displays heading with respect to TIO, and north on the compass card points along the TIO plane. The heading of the body plus X axis with respect to the target insertion orbit is read at the lubber line.
The course pointer provides the heading of the Earth-relative velocity vector with respect to the TIO plane. The CDI deflection indicates the estimated sideslip angle, the angle between the body X axis and the relative velocity vector.
The primary bearing pointer during major modes 102 and 103 is fixed on the compass card at a predetermined value to provide a turnaround heading in the event of an RTLS abort. During RTLS major mode 601, the pointer indicates the heading to the landing site runway. The secondary bearing provides the heading of the inertial velocity vector with respect to the TIO plane.
The horizontal situation CRT display allows the flight crew to configure the software for nominal winds or high head winds. The software item entry determines the distance from the runway threshold to the intersection of the glide slope with the runway centerline. The high-wind entry pushes the intercept point close to the threshold. The distance selected is factored into the computation of reference altitude from which the GSI is derived.
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"Shuttle, Alpha Mach Indicator",391,0,0,0
The two alpha Mach indicators are located next to the attitude director indicators on panels F6 and F8. The AMIs consists of four tape meters displaying angle of attack ( alpha ), vehicle acceleration (accel), vehicle velocity ( M/vel ) and equivalent airspeed ( EAS ). The two units are driven independently but can have the same data source.
Alpha displays vehicle angle of attack, defined as the angle between the vehicle plus X axis and the wind-relative velocity vector (negative wind vector). Alpha is displayed by a combination moving scale and moving pointer. For angles between minus 4 degrees and plus 28 degrees, the scale remains stationary and the pointer moves to the correct reading.
For angles less than minus 4 degrees or greater than plus 28 degrees, the pointer stops (at minus 4 or plus 28 degrees) and the scale moves so that the correct reading is adjacent to the pointer. The alpha tape ranges from minus 18 to plus 60 degrees with no scale changes. The negative scale numbers (below zero) have no minus signs; the actual tape has black markings on a white background on the negative side and white markings on a black background on the positive side.
The accel scale displays vehicle drag acceleration, which is the deceleration along the flight path. This is a moving tape upon which acceleration is read at the fixed lubber line. The tape range is minus 50 to plus 100 with a scale change at zero feet per second squared. Minus signs are assumed on the accel scale also; the negative region has a black background and the positive side has a white background.
The M/vel scale displays Mach number or relative velocity. Mach number is the ratio of vehicle airspeed to the speed of sound in the same medium. Relative velocity in this case is the vehicle airspeed. The actual parameter displayed is always Mach number; the tape is simply rescaled above Mach 4 to read relative velocity in thousands of feet per second (above 2,000 feet per second, Mach number = V REL /1,000). The M/vel scale is a moving tape from which Mach/velocity is read at the fixed lubber line. The scale ranges from zero to 27 with a scale change at Mach 4.
The EAS scale is used to display equivalent airspeed. On the moving-tape scale, equivalent airspeed is read at the fixed lubber line. The tape range is zero to 500 knots, and scaling is 1 inch per 10 knots.
Each scale on the AMI displays an off flag if the indicator malfunctions, invalid data are received at the DDU or a power failure occurs (all flags appear).
The air data source select switch on panel F6 for the commander and panel F8 for the pilot determines the source of data for the AMI and altitude/vertical velocity indicator. The nav position of the air data switch ensures that the alpha , Mach and EAS on the AMI are the same parameters sent to guidance, flight control, navigation and other software users; accel comes from navigation software.
The left, right position of the air data switch selects predetermined data from the left or right air data probe assembly after deployment of the left and right air data probes at Mach 3 for alpha, M/vel and EAS display. Accel is always derived from navigation software during entry. It is driven to zero during terminal area energy management and approach and landing.
The altitude/vertical velocity indicators are located on panel F6 for the commander and panel F8 for the pilot. These indicators display vertical acceleration ( alt accel ), vertical velocity ( alt rate ), barometric altitude ( alt ) and radar altitude ( rdr alt ).
The alt accel indicator, which displays altitude acceleration of the vehicle, is read at the intersection of the moving pointer and the fixed scale. The scale range is minus 13.3 to 13.3 feet per second squared, and the scaling is 6.67 feet per second squared per inch. Software limits acceleration values to plus or minus 12.75 feet per second squared.
The alt rate scale displays vehicle altitude rate, which is read at the intersection of the moving tape and the fixed lubber line. The scale range is minus 2,940 to plus 2,940 feet per second with scale changes at minus 740 feet per second and plus 740 feet per second. The negative and positive regions are color-reversed: negative numbers are white on a black background and positive numbers are black on white.
The alt scale, a moving tape read against a fixed lubber line, displays the altitude of the vehicle above the runway (barometric altitude). The scale range is minus 1,100 feet to plus 165 nautical miles (189 statute miles), with scale changes at minus 100, zero, 500 feet and plus 100,000 feet. The scale is in feet from minus 1,100 to plus 400,000 and in nautical miles from plus 40 to plus 165 (46 to 189 statute miles). Feet and nautical miles overlap from plus 40 to plus 61 nautical miles (46 to 70 statute miles).
The rdr alt scale is a moving tape read against a fixed lubber line. It displays radar altitude (corrected to wheels) during major mode 305, below 9,000 feet (normally not locked in until below 5,000 feet; prior to radar \Jaltimeter\j lock-on, the meter is ''parked'' at 5,000 feet). The scale ranges from zero to 9,000 feet with a scale change at 1,500 feet. Each scale on the AVVI displays an off flag in the event of indicator malfunction, invalid data received at the DDU or power failure (all flags appear).
With the air data source switch in the nav position, the alt accel, alt rate, and alt scales are navigation-derived. The rdr alt indicator is controlled by the radar altm switch on panel F6 for the commander and panel F8 for the pilot. Radar altm positioned to 1 selects radar \Jaltimeter\j 1; 2 selects radar \Jaltimeter\j 2.
The air data switch is positioned to left or right to select the right or left air data probe, respectively, after air data probe deployment at Mach 3. The alt and alt rate scales receive information from the selected air data probe. Alt accel receives navigation data. The rdr alt scale receives data from the radar alt select switch.
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"Shuttle, Surface Position Indicator",393,0,0,0
The surface position indicator is a single display on panel F7 that is active during entry and during the entry portion of RTLS. The SPI displays the actual and commanded positions of the elevons, body flap, rudder, aileron and speed brake.
The four elevon position indicators show the elevon positions in the order of appearance as viewed from behind the vehicle (from left to right: left outboard, left inboard, right inboard, right outboard). The scales all range from plus 20 to minus 35 degrees, which are also the software limits to the elevon commands. The pointers are driven by four separate signals and can read different values, but normally the left pair is identical and the right pair is identical. Positive elevon is below the null line and negative is above.
The body flap scale reads body flap positions from zero to 100 percent of software-allowed travel. Zero percent corresponds to full up (minus 11.7 degrees); 100 percent corresponds to full down (plus 22.5 degrees). The small pointer at 35 percent is fixed and shows the trail position.
Rudder position is displayed as if viewed from the rear of the vehicle. Deflection to the left of center represents left rudder. The scale is plus 30 degrees (left) to minus 30 degrees (right), but software limits the rudder command to plus or minus 27.1 degrees.
The aileron display measures the effective aileron function of the elevons in combination. Aileron position equals the average of the left and right elevon divided by two. Deflection of the pointer to the right of center indicates a roll-right configuration (left elevons down, right elevons up) and vice versa. The scale is minus 5 to plus 5 degrees, with minus 5 at the left side. The aileron command can exceed plus or minus 5 degrees (maximum plus or minus 10 degrees), in which case the meter saturates at plus or minus 5 degrees.
The speed brake position indicator indicates the actual position on the upper scale and commanded position on the lower scale. The position ranges zero to 100 percent; zero percent is fully closed and 100 percent is fully open, which corresponds to 98 degrees with respect to the hinge lines. The small point at 25 percent is fixed and represents the point at which the speed brake surfaces and the remainder of the tail form a smooth wedge.
The speed brake command is scaled identically to position and has the same travel limits. It always represents the speed brake auto guidance command. The off flag is set only for internal meter problems or during OPS 8 display checkout.
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"Shuttle, Flight Control System Push Button Light Indicators",394,0,0,0
These indicators are located on panel F2 for the commander and panel F4 for the pilot. The flight control system's push button light indicators transmit flight crew moding requests to the digital autopilot in the flight control software and reflect selection by illuminating the effective DAP state.
The push button light indicators are used to command and reflect the status of the pitch control mode. The pitch and roll/yaw indicators transmit moding requests to the digital autopilot and indicate the effective state of the pitch, roll and yaw DAP channels by lighting.
Auto indicates that control is automatic and no crew inputs are required. CSS is control stick steering; crew inputs are required but are smoothed by the DAP (stability augmentation, turn coordination).
The spd brk/throt (speed brake/throttle) push button light indicator has two separate lights, auto and man (manual), to indicate that the DAP speed brake channel is in the automatic or manual mode. The push button light indicator transmits only the auto request.
The body flap push button light indicator also has separate auto and man lights, indicating the state of the body flap channel. Like the spd brk/throt push button light indicator, the body flap indicator transmits only the auto request.
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"Shuttle, Reaction Control System Command Lights",395,0,0,0
The reaction control system command lights on panel F6 are active during the entry and RTLS flight phases. Their primary function is to indicate RCS jet commands by axis and direction; secondary functions are to indicate when more than two yaw jets are commanded and when the elevon drive rate is saturated.
During major modes 301 through 304, up until the roll jets are no longer commanded (dynamic pressure exceeds 10 pounds per square foot), the roll l and r lights indicate that left or right roll commands have been issued by the DAP. The minimum light-on duration is extended so that the light can be seen even during minimum-impulse firings. When dynamic pressure is greater than or equal to 10 pounds per square foot, the roll lights are quiescent until 50 pounds per square foot, after which time both lights are illuminated whenever more than two yaw jets are commanded on.
The pitch u and d lights indicate up and down pitch jet commands until dynamic pressure equals 20 pounds per square foot, after which the pitch jets are no longer used. When dynamic pressure is 50 pounds per square foot or more, the pitch lights, like the roll lights, assume a new function: both light whenever the elevon surface drive rate exceeds 20 degrees per second (10 degrees per second if only one hydraulic system is left).
The yaw l and r lights function as yaw jet command indicators throughout entry until the yaw jets are disabled at Mach 1. The yaw lights have no other functions.
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"Shuttle, G-Meter",396,0,0,0
The g-meter is a self-contained accelerometer and display unit mounted on panel F7. It senses linear acceleration along the Z axis (normal) of the vehicle. A mass weight in the unit is supported vertically by two guide rods and is constrained by a constant-rate helical spring. The inertial force of the mass is proportional to the inertial force of the vehicle and, hence, to the input acceleration, under conditions of constant acceleration.
Displacement of the mass is translated to pointer displacement through a rack-and-pinion gear train whose output is linear with input acceleration. The display indicates acceleration from minus 2 g's to plus 4 g's. The g-meter requires no power and has no software interface. Like all the dedicated displays, it has an external variable incandescent lamp.
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"Shuttle, Head-Up Display",397,0,0,0
The head-up display is an optical miniprocessor that cues the commander and/or pilot during the final phase of entry and particularly in the final approach to the runway. With minimal movement of their eyes from the forward windows (head up) to the dedicated display instruments (head down), the commander and pilot can read data from HUDs located in front of them on their respective glareshields. The HUD displays the same data presented on several other instruments, including the ADI, SPI, AMI and AVVI.
The HUD allows out-of-the-window viewing by superimposing flight commands and information on a transparent combiner in the window's field of view. The baseline orbiter, like most commercial \Jaircraft\j, presents conventional electromechanical display on a panel beneath the glareshield, which necessitates that the flight crew look down for information and then up to see out the window. During critical flight phases, particularly approach and landing, this is not an easy task. In the orbiter, with its unique vehicle dynamics and approach trajectories, this situation is even more difficult.
Since the orbiter is intended to be in service for several years, the addition of a HUD was considered appropriate. Most recent military \Jaircraft\j include HUD systems, as do several European airliners. Additionally, since the display portion of some existing HUD systems could be easily installed in the orbiter, the HUD system requirements for the orbiter were patterned after existing hardware to minimize development costs.
While the display portion of the orbiter system could be similar to existing HUD systems, the drive \Jelectronics\j could not. Since the orbiter avionics systems are digital and minimal impact on the orbiter was paramount, the HUD drive \Jelectronics\j were designed to receive data from the orbiter data buses. Most existing HUD drive \Jelectronics\j use analog data or a combination analog/digital interface. In the orbiter system, the HUD drive \Jelectronics\j utilize, to the maximum extent possible, the same data that drive the existing electromechanical display devices.
The orbiter display device, designed by Kaiser \JElectronics\j of San Jose, Calif., uses a CRT to create the image, which is then projected through a series of lenses onto a combining glass (a system very similar to one they developed and produce for the Cobra jet aircraft). Certain orbiter design requirements, such as vertical viewing angles, brightness and unique mounting, dictated some changes from the Cobra configuration.
A HUD power on/off switch located on the left side of panel F3 provides and terminates electrical power to the commander's HUD. The same switch is also located on the right side of panel F3 for the pilot's HUD.
Each HUD is a single-string system but connected to two data buses for redundancy. It is an electronic/optical device with two sets of combiner glasses located above the glareshield in the direct line of sight of the commander and the pilot. Essential flight information for vehicle guidance and control during approach and landing is projected on the combiner glasses and collimated at infini
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"Shuttle, Pre-Launch Operations",398,0,0,0
After the Space Shuttle has been rolled out to the launch pad on the Mobile Launcher Platform (MLP), all pre-launch activities are controlled from the Launch Control Center (LCC).
After the Shuttle is in place on the launch pad support columns, and the Rotating Service Structure (RSS) is placed around it, power for the vehicle is activated. The MLP and the Shuttle are then electronically and mechanically mated with support launch pad facilities and ground support equipment. An extensive series of validation checks verify that the numerous interfaces are functioning properly.
Meanwhile, in parallel with pre-launch pad activities, cargo operations get underway in the RSS's Payload Changeout Room.
Vertically integrated payloads are delivered to the launch pad before the Shuttle is rolled out. They are stored in the Payload Changeout Room until the Shuttle is ready for cargo loading. Once the RSS is in place around the orbiter, the payload bay doors are opened and the cargo is installed. Final cargo and payload bay closeouts are completed in the Payload Changeout Room and the payload bay doors are closed for flight.
Initial Shuttle propellant loading involves pumping hypergolic propellants into the orbiter's aft and forward Orbital Maneuvering System and Reaction Control System storage tanks, the orbiter's hydraulic Auxiliary Power Units, and SRB hydraulic power units. These are hazardous operations, and while they are underway work on the launch pad is suspended.
Since these propellants are hypergolic -- that is they ignite on contact with one another--oxidizer and fuel loading operations are carried out serially, never in parallel.
Finally, dewar tanks on the Fixed Service Structure (FSS), are filled with liquid oxygen and liquid \Jhydrogen\j, which will be loaded into the orbiter's Power Reactant and Storage Distribution (PRSD) tanks during the launch countdown.
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"Shuttle, Final Pre-launch Activities",400,0,0,0
Before the formal Space Shuttle launch countdown starts, the vehicle is powered down while pyrotechnic devices -- various ordinance components -- are installed or hooked up. The extravehicular Mobility Units (EMUs) -- space suits -- are stored On Board along with other items of flight crew equipment.
When closeouts of the Space Shuttle and the launch pad are completed, all is in readiness for the countdown to get underway.
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"Shuttle, Launch Control Center",401,0,0,0
While the VAB can be considered the heart of LC-39, the Launch Control Center (LCC) can easily be called its brain.
The LCC is a 4-story building connected to the east side of the VAB by an elevated, enclosed bridge. It houses four firing rooms that are used to conduct NASA and classified military launches of the Space Shuttle. Each firing room is equipped with the Launch Processing System (LPS) which monitors and controls most Shuttle assembly, checkout and launch operations. Physically, the LCC is 77 ft. high, 378 ft. long and 181 ft. wide.
Thanks to the LPS, the countdown for the Space Shuttle takes only about 40 hours, compared with the 80 plus hours usually needed for a Saturn/Apollo countdown. Moreover, the LPS calls for only about 90 people to work in the firing room during launch operations -- compared with about 450 needed for earlier manned missions.
From the outside, the LCC is virtually unchanged from its original Apollo-era configuration, except that a fourth floor office has been added to the southwest and northwest corners corner of the building.
The interior of the LCC has undergone extensive modifications to meet the needs of the Space Shuttle era.
Physically, the LCC is constructed as follows: the first floor is used for administrative activities and houses the building's utilities systems control room; the second floor is occupied by the Control Data Subsystem; the four firing rooms occupy practically all of the third floor, and the fourth floor, as mentioned, earlier is used for offices.
During the Shuttle Orbital Flight Test program and the early operational missions, Firing Room l was the only fully-equipped control facility available for vehicle checkout and launch. However, as the Shuttle launch rate increased during the first half of the 1980s, the other three firing rooms were activated. Although NASA operates the firing rooms, the Department of Defense uses Firing Rooms 3 and 4 to support its classified, Shuttle-dedicated missions. Additionally, Firing Room 4 serves as an \Jengineering\j analysis and support facility for launch and checkout operations.
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"Shuttle, Launch Countdown",402,0,0,0
As experience was gained by launch crews during the early years of the Space Shuttle program, the launch countdown was refined and streamlined to the point where the average countdown now takes a little more than 40 hours. This was not the case early in the program, when countdowns of 80 hours or more were not uncommon.
The following is a narrative description of the major events of a typical countdown for the Space Shuttle. The time of liftoff is predicated on what is called the launch window -- that point in time when the Shuttle must be launched in order to meet specific mission objectives such as the deployment of \Jspacecraft\j at a predetermined time and location in space.
\BLaunch Minus 3 Days.\b
The countdown gets underway with the traditional call to stations by the NASA Test Director. This verifies that the launch team is in place and ready to proceed.
The first item of business is to checkout the backup flight system and the software stored in the mass memory units and display systems. Backup flight system software is then loaded into the Shuttle's fifth general purpose computer (GPC's).
Flight crew equipment stowage begins. Final inspection of the orbiter's middeck and flight decks are made, and removal of work crew module platforms begin. Loading preparations for the external tank get underway, and the Shuttle main engines are readied for tanking. Servicing of fuel cell storage tanks also starts. Final vehicle and facility closeouts are made.
\BLaunch Minus 2 Days.\b
The launch pad is cleared of all personnel while liquid oxygen and \Jhydrogen\j are loaded into the Shuttle fuel cell storage tanks. Upon completion, the launch pad area is reopened and the closeout crew continues its prelaunch preparations.
The orbiter's flight control, navigation and communications systems are activated. Switches located on the flight and mid- decks are checked and, if required, mission specialist seats are installed. Preparations also are made for rollback of the Rotating Service Structure (RSS).
At launch minus ll hours a planned countdown hold -- called a built-in hold -- begins and can last for up to 26 hours, 16 minute depending on the type of payload, tests required and other factors. This time is used, if needed, to perform tasks in the countdown that may not have been completed earlier.
\BLaunch Minus 1 Day.\b
Countdown is resumed after the built-in hold period has elapsed. The RSS is rolled back and remaining items of crew equipment are installed. Cockpit switch positions are verified, and oxygen samples are taken in the crew area. The fuel cells are activated following a fuel cell flow through purge. Communications with the Johnson Space Center.s Mission Control Center (MCC) are established.
Finally, the launch pad is again cleared of all personnel while conditioned air that has been blowing through the payload bay and other orbiter cavities is switched to inert gaseous \Jnitrogen\j in preparation for filling the external tank with its super-cold propellants.
\BLaunch Day\b
Filling the external tank with liquid oxygen and \Jhydrogen\j gets underway. Communications checks are made with elements of the Air Force's Eastern Space and Missile Center. Gimbal profile checks of the Orbital Maneuvering System (OMS) engines are made. Preflight \Jcalibration\j of the Inertial Measurement Units (IMU) is made, and tracking antennas at the nearby Merritt Island Tracking Station are aligned for liftoff.
At launch minus 5 hours, 20 minutes -- T minus 5 hours, 20 minutes -- a 2-hour built-in hold occurs. During this hold, an ice inspection team goes to the launch pad to inspect the external tank's \Jinsulation\j to insure that there is no dangerous accumulation of ice on the tank caused by the super-cold liquids. Meanwhile, the closeout crew is preparing for the arrival of the flight crew.
Meanwhile, the flight crew, in their quarters at the Operations and Checkout (O&C) Building, eat a meal and receive a weather briefing. After suiting up, they leave the O&C Building at about T minus 2 hours, 30 minutes for the launch pad -- the countdown having resumed at T minus 3 hours.
Upon arriving at the white room at the end of the orbiter access arm, the crew, assisted by white room personnel, enter the orbiter. Once on board they conduct air-to-ground communications checks with the LCC and MCC.
Meanwhile, the orbiter hatch is closed and hatch seal and cabin leak checks are made. The IMU preflight alignment is made and closed-loop tests with Range Safety are completed. The white room is then evacuated and the closeout crew proceeds from the launch pad to a fallback area. At this time, primary ascent guidance data is transferred to the backup flight system.
At T minus 20 minutes a planned 10-minute hold begins. When the countdown is resumed on-board computers are commanded to their launch configuration and fuel cell thermal conditioning begins. Orbiter cabin vent valves are closed and the backup flight system transitions into its launch configuration.
At T minus 9 minutes another planned 10-minute hold occurs. Just prior to resuming the countdown, the NASA Test Director gets the "go for launch" verification from the launch team. At this point, the Ground Launch Sequencer (GLS) is turned on and the terminal countdown starts. All countdown functions are now automatically controlled by the GLS computer located in the Firing Room Integration Console.
At T minus 7 minutes, 30 seconds, the orbiter access arm is retracted. Should an emergency occur requiring crew evacuation from the orbiter, the arm can be extended either manually or automatically in about 15 seconds.
At T minus 5 minutes, 15 seconds the MCC transmits a command that activates the orbiter's operational instrumentation recorders. These recorders store information relating to ascent, on-orbit and descent performance during the mission. These data are analyzed after landing.
At T minus 5 minutes, the crew activates the Auxiliary Power Units (APU) to provide pressure to the Shuttle's three hydraulic systems which move the main engine nozzles and the aero-aerosurfaces. Also at this point, the firing circuit for SRB ignition and the range safety destruct system devices are mechanically enabled by a motor-driven switch called the safe and arm device.
At about T minus 4 minutes, 55 seconds, the liquid oxygen vent on the external tank is closed. It had been open to allow the super-cold liquid oxygen to boil off, thus preventing over pressurization while the tank remained near its full level. Now, with the vent closed, preparations are made to bring the tank to its flight pressure. This occurs at T minus 2 minutes, 55 seconds.
At T minus 4 minutes the final \Jhelium\j purge of the Shuttle's three main engines is initiated in preparation for engine start. Five seconds later, the orbiter's elevons, speed brakes and rudder are moved through a pre-programmed series of maneuvers to position them for launch. This is called the aerosurface profile.
At T minus 3 minutes, 30 seconds, the ground power transition takes place and the Shuttle's fuel cells transition to internal power. Up to this point, ground power had augmented the fuel cells. Then, 5 seconds later, the main engine nozzles are gimballed through a pre-programmed series of maneuvers to confirm their readiness.
At T minus 2 minutes, 50 seconds, the external tank oxygen vent hood -- known as the beanie cap -- is raised and retracted. It had been in place during tanking operations to prevent ice buildup on the oxygen vents. Fifteen seconds later, at T minus 2 minutes, 35 seconds, the piping of gaseous oxygen and \Jhydrogen\j to the fuel cells from ground tanks is terminated and the fuel cells begin to use the on board reactants.
At T minus 1 minute, 57 seconds, the external tank's liquid \Jhydrogen\j is brought to flight pressure by closing the boil off vent, as was done earlier with the liquid oxygen vent. However, during the \Jhydrogen\j boil off of, the gas is piped out to an area adjacent to the launch pad where it is burned off.
At T minus 31 seconds, the Shuttle's on-board computers start their terminal launch sequence. Any problem after this point will require calling a "hold" and the countdown recycled to T minus 20 minutes. However, if all goes well, only one further ground command is needed for launch. This is the "go for main engine start," which comes at the T-minus-10-second point. Meanwhile, the Ground Launch Sequencer (GLS) continues to monitor more than several hundred launch commit functions and is able automatically to call a "hold" or "cutoff" if a problem occurs.
At T minus 28 seconds the SRB booster hydraulic power units are activated by a command from the GLS. The units provide hydraulic power for SRB nozzle gimballing. At T minus 16 seconds, the nozzles are commanded to carry out a pre-programmed series of maneuvers to confirm they are ready for liftoff. At the same time -- T minus 16 seconds -- the sound suppression system is turned on and water begins to pour onto the deck of the MLP and pad areas to protect the Shuttle from acoustical damage at liftoff.
At T minus ll seconds, the SRB range safety destruct system is activated.
At T minus 10 seconds, the "go for main engine start" command is issued by the GLS. (The GLS retains the capability to command main engine stop until just before the SRBs are ignited.) At this time flares are ignited under the main engines to burn away any residual gaseous \Jhydrogen\j that may have collected in the vicinity of the main engine nozzles. A half second later, the flight computers order the opening of valves which allow the liquid \Jhydrogen\j and oxygen to flow into the engine's turbopumps.
At T minus 6.6 seconds, the three main engines are ignited at intervals of 120 milliseconds. The engines throttle up to 90 percent thrust in 3 seconds. At T minus 3 seconds, if the engines are at the required 90 percent, SRB ignition sequence starts. All of these split-second events are monitored by the Shuttle's four primary flight computers.
At T minus zero, the holddown explosive bolts and the T-O umbilical explosive bolts are blown by command from the on-board computers and the SRBs ignite. The Shuttle is now committed to launch. The mission elapsed time is reset to zero and the mission event timer starts. The Shuttle lifts off the pad and clears the tower at about T plus 7 seconds. Mission control is handed over to JSC after the tower is cleared.
#
"Shuttle, Mission Control Center",403,0,0,0
Space Shuttle flights are controlled through the Mission Control Center (MCC) at Johnson Space Center, Houston, \JTexas\j. It has been central control for more than 60 NASA manned space flights since becoming operational in June 1965, for the Gemini 4 mission.
Located in a square, windowless, 3-story building, designated Building 30, the MCC has two Flight Control Rooms (FCRs--the \Jacronym\j is pronounced "fickers") from which Shuttle missions are managed. These rooms are functionally identical. One located on the second floor is used for NASA-controlled missions, the other, on the third floor, is dedicated primarily to Department of Defense missions. However, either FCR can be used for mission control. They also can be used simultaneously to control separate flights if required.
The MCC takes over mission control functions when the Space Shuttle clears the service tower at the Kennedy Space Center.s Launch Complex 39. Shuttle systems data, voice communications and \Jtelevision\j are relayed almost instantaneously to MCC through the NASA Ground and Space Networks, the latter using the orbiting Tracking and Data Relay Satellites. The MCC retains its mission control function until the end of a mission, when the orbiter lands and rolls to a stop. At that point the Kennedy Space Center again assumes control.
In the event MCC becomes inoperative because of a hurricane or other disaster, backup mission control capability would shift to the NASA Ground Terminal at JSC's White Sands Test Facility near Las Cruces, NIMI. This emergency control center is a stripped-down version of MCC, with minimal equipment and instrumentation to allow controllers to support a mission to its conclusion.
In the FCRs, teams of up to 30 flight controllers sit at consoles directing and monitoring all aspects of the flight 24 hours a day, 7 days a week. Each team is headed by a flight director and normally works an 8-hour shift.
Over the years, the flight controller teams have become specialized. One team, for example, becomes responsible for ascent-to-orbit and return-from-orbit, two others for in-space operations and a fourth planning next-day mission activities. Augmenting the FRC teams are groups of engineers, flight controllers and technicians who monitor and analyze flight data from adjacent staff support areas.
There are normally 16 major flight control consoles operating in an FCR during a Space Shuttle mission. Each console is identified by title or a "call sign" which is used when communicating with other controllers or the \Jastronaut\j flight crew.
These mission command and control positions, their individual initials, call signs and responsibilities include:
FLIGHT DIRECTOR (FD), with the call sign "Flight," is the leader of the flight control team. The Flight Director is responsible for mission and payload operations and decisions relating to safety and flight conduct.
SPACECRAFT COMMUNICATOR (CAPCOM), with the familiar call sign "CAPCOM" -- Capsule Communicator -- is the primary communicator between MCC and the Shuttle crew. The \Jacronym\j dates from the Mercury program when the Mercury \Jspacecraft\j was called a capsule.
FLIGHT DYNAMICS OFFICER (FDO), call sign "Fido," plans orbiter maneuvers and follows the Shuttle's flight trajectory along with the Guidance Officer.
GUIDANCE OFFICER (GDO), call sign "Guidance," is responsible for monitoring the orbiter navigation and guidance computer software.
DATA PROCESSING SYSTEMS ENGINEER (DPS), keeps track of the orbiter's data processing systems, including the five on-board general purpose computers, the flight-critical and launch data lines, the malfunction display system, mass memories and systems software.
FLIGHT SURGEON (Surgeon) monitors crew activities and is for the medical operations flight control team, providing medical consultations with the crew, as required, and keeping the Flight Director informed on the state of the crew's health.
BOOSTER SYSTEMS ENGINEER (Booster) is responsible for monitoring and evaluating the main engine, solid rocket booster and external tank performance before launch and during the ascent phases of a mission.
PROPULSION SYSTEMS ENGINEER (PROP) monitors and evaluates performance of the reaction control and orbital maneuvering systems during all flight phases and is charged with management of propellants and other consumables for various orbiter maneuvers.
GUIDANCE NAVIGATION AND CONTROL SYSTEMS ENGINEER (GNC) is charged with monitoring all Shuttle guidance, navigation and control systems. GNC also keeps the Flight Director and crew notified of possible abort situations and keeps the crew informed of any guidance problems.
ELECTRICAL, ENVIRONMENTAL AND CONSUMABLES SYSTEMS ENGINEER (EECOM) is responsible for monitoring the cryogenic supplies available for the fuel cells, avionics and cabin cooling systems, as well as electrical distribution, cabin pressure and orbiter lighting systems.
INSTRUMENTATION AND COMMUNICATIONS SYSTEMS ENGINEER (INCO) is charged with planning and monitoring in-flight communications and instrumentation systems.
GROUND CONTROL (GC) is responsible for maintenance and operation of MCC hardware, software and support facilities. GC also coordinates tracking and data activities with the Goddard Space Flight Center (GSFC), Greenbelt, Md.
FLIGHT ACTIVITIES OFFICER (FAO), plans and supports crew activities, checklists, procedures and schedules.
PAYLOADS OFFICER (Payloads) is in charge of coordinating the ground and on-board system interfaces between the flight control team and the payload user. The Payloads Officer also monitors Spacelab and upper stage systems and their interfaces with payloads.
MAINTENANCE, MECHANICAL ARM AND CREW SYSTEMS ENGINEER (MMACS), call sign "Max," monitors operation of the remote manipulator arm and the orbiter's structural and mechanical systems. Max also observes crew hardware and in-flight equipment maintenance.
PUBLIC AFFAIRS OFFICER (PAO) provides mission commentary, augments and explains air-to-ground conversations and flight control operations for the news media and public.
During Spacelab missions another flight control position is needed. This is the Command and Data Management Systems Officer (CDMS), who is primarily responsible for data processing of the Spacelab's two main computers. To support Spacelab missions the EECOM and the DPS both work closely with the CDMS since the missions involve monitoring additional displays involving almost 300 items and coordinating their activities with the Marshall Space Flight Center's Payload Operations Control Center (POCC).
One of the most unusual support facilities of the FCRs is the display/control system. It consists of a series of projection screens displays on the front wall which show the orbiter's "realtime" location, live \Jtelevision\j pictures of crew activities, Earth views and extravehicular activities. Other displays include mission elapsed time as well as time remaining before a maneuver or other major mission event.
Many of the decisions or recommendations made by flight controllers are based on information shown on the display/control system displays. Telemetry data is processed instantaneously for display allowing controllers to keep current on the status of Shuttle systems.
Consoles in the FCR, the adjacent multipurpose support room, and POCCs have one or more TV screens and switches to allow the controllers to view data displays on a number of different channels. It is possible to call up data of special interest simply by changing channels. Also, an extensive library of reference data is available to display static data, while digital-to-television display generators can provide dynamic, or constantly changing data.
Eventually, it is planned that the Apollo-era consoles will be superceded by modern state-of-the-art work stations that will provide more capability to monitor and analyze vast amounts of data. Moreover, instead of driving the consoles with a single main computer, each console will eventually have its own smaller computer which will be able to monitor a specific system and be linked into a network capable of sharing the data.
While the FCRs are the nerve center for MCC operations, there are other behind-the-scene work areas that are vital to successful Shuttle operations. These include the Network Interface Processor (NIP), and the Data Computation Complex (DCC) both of which are located on the first floor of the MCC building.
The NIP, as its name implies, processes incoming digital data and distributes it in realtime to the FCR and support room displays. This system also handles digital command signals to the orbiter permitting ground controllers to keep on-board guidance computers current.
The DCC processes incoming tracking and telemetry data and compares what is happening with what should be happening. Normally, it will display information only if a problem occurs. It also will decide what maneuvers should be made to correct the problem. The DCC also predicts where the orbiter will be at a specific point in flight, and it aids ground tracking stations to point their antennas in the right direction.
The DCC uses five primary computers each of which can support the FCR. During critical phases of a mission, one of the five computers is designated a "dynamic standby," processing data concurrently in case the prime computer fails. The DCC computers also are used to develop computer programs for future Shuttle missions.
Operating in conjunction with the FCRs are facilities known as Payload Operations Control Centers (POCCs) where principal investigators and commercial users can monitor and control payloads being carried on board the Shuttle. One of the most extensive POCCs is located at the Marshall Space Flight Center, \JHuntsville\j, Ala., where Spacelab missions will be coordinated with MCC.
It is the command post, communications hub and data relay station for the principal investigators, mission managers and support teams. Here decisions on payload operations are made, coordinated with the MCC Flight Director, and sent to the Spacelab or Shuttle.
The POCC at Goddard Space Flight Center, controls free-flying \Jspacecraft\j that are deployed, retrieved or serviced by the Shuttle. Planetary mission \Jspacecraft\j are controlled from the POCC at NASA's Jet Propulsion Laboratory, Pasadena, Calif. Finally, private sector payload operators and foreign governments maintain their own POCCs at various locations for control of \Jspacecraft\j systems under their control.
#
"Shuttle, Marshall Payload Operations Control Center",404,0,0,0
The Payload Operations Control Center (POCC) operated by the NASA's Marshall Space Flight Center (MSFC), \JHuntsville\j, Ala., is the largest and most diverse of the various POCCs associated with the Space Shuttle program. Since its functions in many respects parallel those of other POCCs operated by private industry, the academic community and government agencies, a description of what it does, how it operates and who operates it will serve as an overview of this type of control center.
The Marshall POCC -- like all POCCs -- is a facility designed to monitor, coordinate, and control on-orbit operation of a Shuttle payload, particularly Spacelab. During non-mission periods it also is used for crew training and simulated space operations. It is, in effect, a command post for payload activities, just as the JSC Mission Control Center (MCC) is a command post for the flight and operation of the Space Shuttle.
Both control centers work closely in coordinating mission activities. In fact, the Marshall POCC originally was housed in Building 30 at JSC, adjacent to the MCC. It has since been moved to Building 4663 at Marshall and is an important element of the Hunstville Operations Support Center (HOSC), which augments the MCC by monitoring Shuttle propulsion systems.
The Marshall POCC Capabilities Document states that the "POCC provides physical space, communications, and data system capabilities to enable user access to payload data (digital, video, and analog), command uplink, and coordination of activities internal and external to the POCC."
Members of the Marshall mission management team and principal investigators and research teams work in the POCC or in adjacent facilities around-the-clock controlling and directing payload experiment operations. Using the extensive POCC facilities they are able to communicate directly with mission crews and direct experiment activities from the ground. They also can operate experiments and support equipment on board the Shuttle and manage payload resources.
The POCC operations concept requires a team consisting of the Payload Mission Manager (PMM) directing the POCC cadre which has overall responsibility for managing and controlling POCC operations. Its scientific counterpart, the investigator's operations team, is the group that conducts, monitors and controls the experiments carried on the Shuttle, primarily those related to Spacelab.
Generally, POCC operations are carried out by a management/scientific team of 10 key individuals, headed by the Payload Operations Director (POD), who is a senior member of the PMM's cadre. The POD is charged with managing the day-to-day mission operations and directing the payload operations team and the science crew.
Other POCC key personnel include:
MISSION SCIENTIST (MSCI) who represents scientists who have experiments on a specific flight and serves as the interface between the PMM and the POD in matters relating to mission science operations and accomplishments.
CREW INTERFACE COORDINATOR (CIC), who coordinates communications between the POCC and the payload crew.
ALTERNATE PAYLOAD SPECIALIST (APS) is a trained payload specialist not assigned to flight duty who aids the payload operations team and the payload crew in solving problems, troubleshooting and modifying crew procedures, if necessary, and who advises the MSCI on the possible impact of any problem areas.
PAYLOAD ACTIVITY PLANNER (PAP), who directs mission replanning activities, as required, and coordinates mission timeline changes with POCC personnel.
MASS MEMORY UNIT MANAGER (MUM) who sends experiment command uplinks to the flight crew based on data received from the POCC operations team.
OPERATIONS CONTROLLER (OC), who coordinates activities of the payload operations team to insure the efficient accomplishment of activities supporting real-time execution of the mission timeline.
PAYLOAD COMMAND COORDINATOR (PAYCOM), who configures the POCC for ground command operation and controls the flow of experiment commands from the POCC to the flight crew.
DATA MANAGEMENT COORDINATOR (DMC), who is responsible for maintaining and coordinating the flow of payload experiment data to and within the POCC the DMC also assesses the impact of proposed changes to the experiment timeline and payload data requirements that affect the payload downlink data.
PUBLIC AFFAIRS OFFICER (PAO), who provides mission commentary on payload activities and serves as the primary source of information on mission progress to the news media and public.
#
"Space Tracking and Data Acquisition",405,0,0,0
Responsibility for Space Shuttle tracking and data acquisition is charged to the Goddard Space Flight Center, Greenbelt, Md. This involves integrating and coordinating all of the worldwide NASA and Department of Defense tracking facilities needed to support Space Shuttle missions.
These facilities include the Goddard-operated Ground Network (GN) and Space Network (SN); the Deep Space Network (DSN) managed for NASA by the Jet Propulsion Laboratory (JPL), Pasadena, Calif.; the Ames Dryden Flight Research Facility, (ADFRC) Edwards, Calif.; and extensive Department of Defense tracking systems at the Eastern and Western Space and Missile Centers, as well as the Air Force Satellite Control Network's (AFSCN) remote tracking stations.
#
"Shuttle Support Ground Network",406,0,0,0
The Ground Network (GN) is a worldwide network of tracking stations and data-gathering facilities which support Space Shuttle missions and also maintain communications with low Earth-orbiting \Jspacecraft\j. Station management is provided from the Network Control Center at Goddard. Basically, commands are sent to orbiting \Jspacecraft\j from the GN stations and, in return, scientific data are transmitted to the stations.
The system consists of 12 stations, including three DSN facilities. GN stations are located at \JAscension\j Island, a British Crown Colony in the south Atlantic Ocean; Santiago, \JChile\j; \JBermuda\j; \JDakar\j, \JSenegal\j, on the West Coast of Africa; \JGuam\j; Hawaii; Merritt Island, Fla.; Ponce de Leon, Fla.; and the Wallops Flight Facility on Virginia's Eastern Shore. The DSN tracking stations are located at \JCanberra\j, \JAustralia\j; Goldstone, Calif.; and Madrid, \JSpain\j.
The GN stations are equipped with a wide variety of tracking and data-gathering antennas, ranging in size from 14 to 85 feet in diameter. Each is designed to perform a specific task, normally in a designated frequency band, gathering radiated electronic signals (telemetry) transmitted from \Jspacecraft\j.
The communications hub for the GN is the Goddard-operated NASA Communications Center (NASCOM). It consists of more than 2 million miles of electronic circuitry linking the tracking stations and the MCC at the Johnson Space Center. NASCOM has six major switching centers to insure the prompt flow of data. In addition to Goddard and JSC, the other switching centers are located at JPL, KSC, \JCanberra\j and Madrid.
The system includes \Jtelephone\j, microwave, radio, submarine cable and geosynchronous communications satellites in ll countries. It includes communications facilities operated by 15 different domestic and foreign carriers. The system also has a wide-band and video capability. In fact, Goddard's wide-band system is the largest in the world.
A voice communications system called Station Conferencing and Monitoring Arrangement (SCAMA) can conference link up hundreds of the 220 different voice channels throughout the United States and abroad with instant talk/listen capability. With its built-in redundancy, SCAMA has realized a mission support reliability record of 99.6 percent. The majority of Space Shuttle voice traffic is routed through Goddard to the MCC.
As would be expected, computers play an important role in GN operations. They are used to program tracking antenna pointing angles, send commands to orbiting \Jspacecraft\j and process data which is sent to the JSC and Goddard control centers.
Shuttle data is sent from the tracking network to the main switching computers at GSFC. These are UNISYS 1160 computers which reformat and transmit the information to JSC almost instantaneously at a rate of l.5 million bits per second, via domestic communications satellites.
#
"Space Network",407,0,0,0
Augmenting the GN and eventually replacing it, is a unique tracking network called the Space Network (SN). The uniqueness of this network is that instead of tracking the Shuttle and other Earth-orbiting \Jspacecraft\j from a world-wide network of ground stations, its main element is an in-orbit series of satellites called the Tracking and Data Relay Satellite System (TDRSS), designed to gather tracking and data information from geosynchronous orbit and relay it to a single ground terminal located at White Sands, N.M.
The first \Jspacecraft\j in the TDRS system, TDRS-1, was deployed from the Space Shuttle Challenger on April 4, 1983. Although problems were encountered in establishing its geosynchronous orbit at 41 degrees west longitude (over the northeast corner of Brazil), TDRS-l proved the feasibility of the tracking station-in-space concept when it became operational later in the year.
Ultimately, the SN will consist of three TDRS \Jspacecraft\j in orbit, one of which will be a backup or spare to be available for use if one of the operational \Jspacecraft\j fails. Each satellite in the TDRS system is designed to operate for 10-years.
Following its planned deployment from the Space Shuttle Discovery scheduled for the STS-26 mission, TDRS-2 will be tested and then positioned in a geosynchronous orbit southwest of Hawaii at 171 degrees west longitude, about 130 degrees from TDRS-1. With these two \Jspacecraft\j and the White Sands Ground Terminal (and eventually a backup terminal) operational, the SN will be able to provide almost full-time communications and tracking of the Space Shuttle, as well as for up to 24 other Earth-orbiting \Jspacecraft\j simultaneously. The global network of ground stations can provide only about 20 percent of that coverage. Eventually some of the current ground stations will be closed when the SN becomes fully operational.
After data acquired by the TDRS \Jspacecraft\j are relayed to the White Sands Ground Terminal, they are sent directly by domestic communications satellite to NASA control centers at JSC for Space Shuttle operations, and to Goddard which schedules TDRSS operations including those of many unmanned satellites.
The TDRS are among the largest and most advanced communications satellites ever developed. They weigh almost 5,000 lb. and measure 57 ft. across at their solar panels. They operate in the S-band and Ku-band frequencies and their complex \Jelectronics\j systems can handle up to 300 million bits of information each second from a single user \Jspacecraft\j. Among the distinguishing features of the \Jspacecraft\j are their two huge, wing-like solar panels which provide l,850 watts of electric power and their two 16-ft. diameter high-gain parabolic antennas which resemble large umbrellas. These antennas weigh about 50 lb. each.
The communications capability of the TDRSS covers a wide spectrum that includes voice, \Jtelevision\j, analog and digital signals. No signal processing is done in orbit. Instead, the raw data flows directly to the ground terminal. During Space Shuttle missions, mission data and commands pass almost continuously back and forth between the orbiter and the MCC at JSC.
Like the TDRS, the White Sands ground terminal is one of the most advanced in existence. Its most prominent features include three 60-ft.-diameter Ku-band antennas which receive and transmit data. A number of smaller antennas are used for S-band and other Ku-band communications.
Ground was broken in September 1987, for a second back-up ground terminal at White Sands to accommodate increased future mission support required from the TDRSS.
The TDRSS segment of the Space Network, including the ground terminal, is owned and operated for NASA by CONTEL Federal Systems Sector, \JAtlanta\j, Ga. The \Jspacecraft\j are built the TRW Federal Systems Division, Space and Technology Group, Redondo Beach, Calif. TRW also provides software support for the White Sands facility. The TDRS parabolic antennas are built by the Harris Corp's Government Communications Systems Division, Melbourne, Fla. Harris also provides ground antennas, radio frequency equipment and other ground terminal equipment.
#
"Shuttle Flight Operations",408,0,0,0
The Space Shuttle, as it thunders away from the launch pad with its main engines and solid rocket boosters (SRB) at full power, is an unforgettable sight. It reaches the point of maximum dynamic pressure (max Q) -- when dynamic pressures on the Shuttle are greatest -- about 1 minute after liftoff, at an altitude of 33,600 ft.
At this point the main engines are "throttled down," to about 75 percent, thus keeping the dynamic pressures on the vehicle's surface to about 580 lb. per square foot. After passing through the max Q region, the main engines are throttled up to full power. This early ascent phase is often referred to as "first stage" flight.
Little more than 2 minutes into the flight, the SRBs, their fuel expended, are jettisoned from the orbiter. The Shuttle is at an altitude of about 30 miles and traveling at a speed of 2,890 miles an hour. The spent SRB casings continue to gain altitude briefly before they begin falling back to Earth.
When the spent casings have descended to an altitude of about 17,000 ft., the parachute deployment sequence starts, slowing them for a safe splashdown in the ocean. This occurs about 5 minutes after launch. The boosters are retrieved, returned to a processing facility for refurbishment and eventual reused.
Meanwhile, the "second stage" phase of the flight is underway with the main engines propelling the vehicle ever higher on its ascent trajectory. At about 8 minutes into the flight, at an altitude of about 60 miles, main engine cut-off (MECO) occurs. The Shuttle is now traveling at a speed of 16,697 mph.
After MECO, the orbiter and the external tank are moving along a trajectory that, if not corrected, would result in the vehicle entering the atmosphere about halfway around the world from the launch site. However, a brief firing of the orbiter's two Orbital Maneuvering System (OMS) thrusters changes the trajectory and orbit is achieved. This takes place just after the external tank has been jettisoned and while the orbiter is flying "upside down" in relation to Earth.
The separated external tank continues on a ballistic trajectory and enters the Earth's atmosphere to break up over a remote area of the Indian Ocean. Meanwhile, an additional firing of the OMS thrusters places the orbiter into its planned orbit, which can range from 115 to 600 miles above the Earth.
There are two ways in which orbit can be accomplished. These are the conventional OMS insertion method called "standard" and the direct insertion method.
The OMS insertion method involves a brief burn of the OMS engines shortly after MECO, placing the orbiter into an elliptical orbit. A second OMS burn is initiated when the orbiter reaches apogee in its elliptical orbit. This brings the orbiter into a near circular orbit. If required during a mission, the orbit can be raised or lowered by additional firings of the OMS thrusters.
The direct insertion technique uses the main engines to achieve the desired orbital apogee, or high point, thus saving OMS propellant. Only one OMS burn is required to circularize the orbit, and the remaining OMS fuel can then be used for frequent changes in the operational orbit, as called for in the flight plan.
The first direct insertion orbit was accomplished during the STS 41-C mission in April 1984, when the Challenger was placed in a 288-mile-high circular orbit where its flight crew was able to successfully capture, repair and redeploy a free-flying \Jspacecraft\j, the Solar Maximum satellite (Solar Max) -- an important "first" for the Space Shuttle program.
#
"Shuttle Launch Abort Modes",409,0,0,0
During the ascent phase of a Space Shuttle flight, if a situation occurs that puts the mission in jeopardy -- the loss, for example, of one or more of the main engines or the OMS thrusters -- the mission may have to be aborted. During the ascent phase, there are two basic Shuttle abort modes: intact aborts and contingency aborts. NASA has attempted to anticipate all possible emergency situations that could occur, and mission plans are prepared accordingly.
Intact aborts -- there are four different types -- permit the safe return of the orbiter and its crew to a pre-planned landing site.
When an intact abort is not possible, the contingency abort option becomes necessary. This crucial abort mode is designed to permit crew survival following a severe systems failure in which the vehicle is lost. Generally, if a contingency abort becomes necessary, the damaged vehicle would fall toward the ocean and the crew would exercise escape options that were developed in the aftermath of the Challenger accident. The four intact abort modes are:
Return to Launch Site (RTLS)
TLA (TAL)
Abort Once Around (AOA)
Abort to Orbit (ATO)
Since an intact abort could result in an emergency landing, before each flight, potential contingency landing sites are designated and weather conditions at these locations are monitored closely before a launch. Space Shuttle flight rules include provisions for minimum acceptable weather conditions at these potential landing sites in the event of intact abort is necessary.
In an abort situation, the type and time of the failure determines which abort mode is possible. There is a definite order of preference for an abort. In cases where performance loss is the only factor, the preferred modes would be ATO, AOA, TAL or RTLS, in that order. The mode selected normally would be the highest preferred one that can be completed with the remaining vehicle performance.
In the case of an extreme system failure -- the loss of cabin pressure or orbiter cooling systems -- the preferred mode would be the one that would terminate the mission as quickly as possible. This means that the TAL or RTLS modes would be more preferable than other modes.
An ascent abort during powered flight can be initiated by turning a rotary switch on a panel in the orbiter cockpit. The switch is accessible to both the commander and the pilot. Normally, flight rules call for the abort mode selection to be made by the commander upon instructions from the Mission Control Center. Once the abort mode is selected, the on board computers automatically initiate abort action for that particular abort.
A description of the intact abort modes follows.
RETURN TO LAUNCH SITE (RTLS)
The RTLS abort is a critical and complex one that becomes necessary if a main engine failure occurs after liftoff and before the point where a TAL or AOA is possible. RTLS cannot be initiated until the SRBs have completed their normal burn and have been jettisoned. Meanwhile, the orbiter with the external tank still attached continues on its downrange trajectory with the remaining operational main engines, the two OMS and four aft RCS thrusters firing until the remaining main engine propellent equals the amount needed to reverse the direction of flight and return for a landing.
A "pitch-around" maneuver of about 5 degrees per second is then performed to place the orbiter and the external tank in an attitude pointing back toward the launch site. OMS fuel is dumped to adjust the orbiter's center of gravity.
When altitude, attitude, flight path angle, heading, weight, and velocity/range conditions combine for external tank jettisoning, MECO is commanded, and the external tank separates and falls into the ocean. After this, the orbiter should glide to a landing at the launch site landing facility. From the foregoing, it can be appreciated why RTLS is the least preferred intact abort mode.
TRANS-ATLANTIC ABORT LANDING (TAL)
The TAL abort mode is designed to permit an intact landing after the Shuttle has flown a ballistic trajectory across the Atlantic Ocean and lands at a designated landing site in Africa or \JSpain\j. This abort mode was developed for the first Shuttle launch in April 1981, and has since evolved from a crew-initiated manual procedure to an automatic abort mode.
The TAL capability provides an abort option between the last RTLS opportunity up to the point in ascent known as the "single-engine press to MECO" capability --meaning that the orbiter has sufficient velocity to achieve main engine cutoff and abort to orbit, even if two main engines are shut down. TAL also can be selected if other system failures occur after the last RTLS opportunity. The TAL abort mode does not require any OMS maneuvers.
Landing sites for a TAL vary from flight to flight, depending on the launch \Jazimuth\j. For the first three Space Shuttle missions, the trajectory inclination was about 28 degrees which made the U.S. Air Force bases at Zaragoza and Moron in \JSpain\j, the most ideal landing sites for TAL. Later Shuttle missions called for air fields at \JDakar\j, \JSenegal\j, and \JCasablanca\j, Morocco, as TAL-option landing sites.
In March 1988, NASA announced that in addition to the TAL sites in \JSpain\j, that two new African contingency landing sites had been selected for future Shuttle missions: a site near Ben Guerir, Morocco, about 40 miles north of Marrakesh with a 14,000-foot runway; and at Banjul, the capital of the west African nation of The Gambia, which has an international airfield with an ll,800-foot runway.
ABORT ONCE AROUND (AOA)
This abort mode becomes available about 2 minutes after SRB separation, up to the point just before an abort to orbit is possible. AOA normally would be called for because of a main engine failure. This abort mode allows the Shuttle to fly once around the Earth and make a normal entry and landing at Edwards AFB, Calif., or White Sands Space Harbor, near Las Cruces, N.M. An AOA abort usually would require two OMS burns, the second burn being a deorbit maneuver.
There are two different AOA entry trajectories. These are the so-called normal AOA and the shallow. The entry trajectory for the normal AOA, is similar to a normal end-of-mission landing. The shallow AOA, on the other hand, results in a flatter entry trajectory, which is less desirable but uses less propellant for the OMS burn. The shallow trajectory also is less desirable because it exposes the orbiter to a longer period of atmospheric entry heating and to less predictable aerodynamic drag forces.
ABORT TO ORBIT (ATO)
The ATO mode is the most benign of the various abort modes. ATO allows the orbiter to achieve a temporary orbit that is lower than the planned. ATO is usually necessary because of a main engine failure. It places fewer performance demands on the orbiter. It also gives ground controllers and the flight crew time to evaluate the problem. Depending on the seriousness of the situation, one ATO option is to make an early deorbit and landing. If there are no major problems, other than the main engine one, an OMS maneuver is made to raise the orbit and the mission is continued as planned.
The first Space Shuttle program ATO occurred on July 29, 1985, following the STS 51-F Challenger launch, when one of the main engines was shut down early by computer command because of a failed temperature sensor. Within 10 seconds of the shutdown, Mission Control declared an ATO situation, and although a lower than planned orbit was attained, the 7-day mission carrying Spacelab-2 was successfully completed.
#
"Shuttle, On-Orbit Operations",410,0,0,0
Space Shuttle flights are controlled by Mission Control Center (MCC) -- usually referred to as "Houston" in air to ground conversations.
During a flight, Shuttle crews and ground controllers work from a common set of guidelines and planned events called the Flight Data File. The Flight Data File includes the crew activity plan, payload handbooks and other documents which are put together during the elaborate flight planning process.
Each mission includes the provision for at least two crew members to be trained for extravehicular activity (EVA). EVA is an operational requirement when satellite repair or equipment testing is called for on a mission. However, during any mission, the two crew members must be ready to perform a contingency EVA if, for example, the payload bay doors fail to close properly and must be closed manually, or equipment must be jettisoned from the payload bay.
The first Space Shuttle program contingency EVA occurred in April 1985, during STS 51-D, a Discovery mission, following deployment of the SYNCOM IV-3 (Leasat 3) communications satellite Leasats' sequencer lever failed and initiation of the antenna deployment and spin-up and perigee kick motor start sequences did not take place.
The flight was extended 2 days to give mission specialists Jeffrey Hoffman and David Griggs an opportunity to try to activate the lever during EVA operations which involved using the RMS. The effort was not successful, but was accomplished on a later mission.
Each Shuttle mission carries two complete pressurized spacesuits called Extra Vehicular Mobility Units (EMU) and backpacks called Primary Life Support Systems (PLSS). These units, along with necessary tools and equipment, are stored in the airlock off the middeck area of the orbiter, ready for use if needed.
As already mentioned, for each mission, two crew members are trained and certified to perform EVAs, if necessary. For those missions in which planned EVAs are called for, the two astronauts receive realistic training for their specific tasks in the Weightless Environment Training Facility at Johnson, with its full-scale model of the orbiter payload bay.
#
"Shuttle, Maneuvering In Orbit",411,0,0,0
Once the Shuttle orbiter goes into orbit, it is operating in the element for which it was designed: the near gravity-free vacuum of space. However, to maintain proper orbital attitude and to perform a variety of maneuvers, an extensive array of large and small rocket thrusters are used -- 46 in all. Each of these thrusters, despite their varying sizes, burn a mixture of \Jnitrogen\j tetroxide and monoethylhydrazine, an efficient but toxic combination of fuels which ignite on contact with each other.
The largest of the 46 control rockets are the two Orbital Maneuvering System (OMS) thrusters which are located in twin pods at the aft end of the orbiter, between the vertical stabilizer and just above the three main engines. Each of the two OMS engines can generate 6,000 lb. of thrust. They can cause a more than l,000 foot-per-second change in velocity of a fully loaded orbiter. This velocity change is called Delta V.
A second and smaller group of thrusters make up the Reaction Control System (RCS) of which there are two types: the primaries and the verniers. Each orbiter has 38 primary trusters, 14 in the forward nose area and 12 on each OMS pod. Each primary thruster can generate 870 lb. of thrust. The smallest of the RCS thrusters, the verniers, are designed to provide what is called "fine tuning" of the orbiter's attitude. There are two vernier thrusters on the forward end of the orbiter and four aft, each generates 24 pounds of thrust.
#
"Year in Space 1994",412,0,0,0
\J1994: Space Launches\j
\JSpace Year Review 1994: China\j
\JSpace Year Review 1994: Clementine\j
\JSpace Year Review 1994: Commercial Space Activities\j
\JSpace Year Review 1994: Compton Gamma Ray Observatory\j
\JSpace Year Review 1994: Department of Defense Space Activities\j
\JSpace Year Review 1994: European Space Activities\j
\JSpace Year Review 1994: Galileo\j
\JSpace Year Review 1994: Hubble Space Telescope (HST)\j
\JSpace Year Review 1994: India\j
\JSpace Year Review 1994: International Space Station\j
\JSpace Year Review 1994: Japan\j
\JSpace Year Review 1994: Magellan\j
\JSpace Year Review 1994: Mission to Planet Earth (MPE)\j
\JSpace Year Review 1994: Russian Space Activities\j
\JSpace Year Review 1994: Solar System Exploration\j
\JSpace Year Review 1994: United States Space Activity\j
\JSpace Year Review 1994: Summary\j
#
"1994: Space Launches",413,0,0,0
(Refer to Table)
#
"Space Year Review 1994: China",414,0,0,0
The People's Republic of China, too, joined the line-up of major space transportation suppliers when its new heavy-lift Chang Zheng (Long March) 3A performed successfully on its maiden flight on February 8, using a CZ-2 with a stretched first stage and a new large cryogenic third stage, carrying a dummy mass and a Chinese scientific satellite.
The first commercial mission of a Long March (since December 1992 when a launch failure destroyed \JAustralia\j's Opts B) followed on July 21 with the launch of Apstar-1, owned by a Chinese-controlled international consortium. A third flight took place on November 30 with the launch of the 2100-kg DFH-3 telecommunications satellite.
#
"Space Year Review 1994: Clementine",415,0,0,0
Launched by the Department of Defense on a Titan II in late January, the small (235 kg) \Jspacecraft\j Clementine, built by the Naval Research Laboratory with BMDO-support in 22 months for only $55 million, provided the first images from the Moon since the Apollo days.
Using four lightweight cameras (UV/visible, short- and long-wavelength infrared, and high-resolution) and a lidar laser transmitter designed for ballistic missile defense programs, Clementine successfully completed its primary mission of mapping 100 percent of the Moon in 11 spectral bands at spatial resolutions of 200-300 meters, complete with \Jaltimeter\j data, by which the previously unsuspected depth of the Aitken basin on the far side was discovered.
Clementine's second mission objective, to provide similar imaging during a fly-by of the asteroid Geographos, was not accomplished because of premature mission termination due to an attitude control anomaly.
#
"Space Year Review 1994: Commercial Space Activities",416,0,0,0
With the DOD launch of the 24th operational GPS on March 9, commercial use of the new high-precision locator system took an upward swing with truckers, packaging companies, car manufacturers, and other users looking into new ways of applying space-based navigation technologies to their needs.
U.S. commercial companies launched eight expendable launch vehicles (ELVs) in 1994 (General Dynamics/Martin Marietta Atlas: 5; McDonnell Douglas Delta II: 3). On the international ELV market, these two types are now encountering three competitors: Europe's Ariane, operated by Arianespace (Evry/France), \JRussia\j's Proton, offered through a U.S.-Russian joint venture called Lockheed-Krunichev-Energia (LKE), and the People's Republic of China's Long March versions, sold by the China Great Wall Industry Corp.
Even with five options, and despite the approximately 115 commercial telecommunications satellites already in geosynchronous orbit, international demands for additional future transponders (from 404 in 1994 to 880 in 2000 for the Asia-Pacific region alone [today's comsats carry from 16-48 transponders each] ) will be difficult to satisfy quickly by the world's ELV supply in coming years.
Of particular commercial interest today are satellite communications for mobile applications: by mid-1994, over 20 major "mobile" comsat systems were proposed, some already in advanced planning stages: Inmarsat P, AMSC/TMI, OmniTracs, Odyssey, Ellipso, and Iridium.
The latter, proposed and designed by Motorola, represents a major new mobile space borne system consisting of 66 satellites, 11 each in six separate orbits, to provide cellular phone service. Funded by Motorola and investors, Iridium Inc. in 1994 raised another $734 million from overseas partners, bringing its reported total funding to over $1.5 billion (estimated cost for the entire system: $4 billion).
#
"Space Year Review 1994: Compton Gamma Ray Observatory",417,0,0,0
The Gamma Ray Observatory Compton, launched April 5, 1991, continued its nearly flawless operations. One of its instruments discovered an unusually bright X-ray source, one of the three brightest in the sky, later named X-ray Nova Scorpii or GRO J1655-40, in the southern \Jconstellation\j Scorpio.
The discovery led to further observations by radio telescopes that showed ejections of matter at velocities close to the speed of light. X-ray novae may be caused by matter spilling from a normal star onto and sucked up by a black hole.
Throughout the year, Compton detected new gamma ray bursters at random locations in the sky, challenging our understanding of the high-energy sky, notably about the \Jisotropic\j distribution of these mysterious sources.
#
"Space Year Review 1994: Department of Defense Space Activities",418,0,0,0
The first (of four) MILSTAR communications satellite, designed for military communications with highly reliable "jam-proof" low-data-rate capabilities during an all-out nuclear war, was launched in February on the first Titan IV/Centaur. The 10,000-lb, \Jspacecraft\j represented the heaviest payload ever carried into geosynchronous orbit by any U.S. expendable launch vehicle.
After more than 20 years of building on its advanced navigational satellite system, the DOD completed its NAVSTAR GPS \Jconstellation\j with the launch of the 24th operational GPS satellite on a Delta II on March 9.
Also in March, the Air Force launched the first in its series of Space Test Experiment Platform (STEP) satellites on the first of the new Orbital Sciences Corp. medium-size launch vehicle \JTaurus\j, a four-stage solid-propellant rocket incorporating three Hercules Orion motors and a Peacekeeper Stage I.
For the second launch of DOD's Miniature Sensor Technology Integration (MSTI) satellite in May, NASA used the 118th and final vehicle of its highly successful Scout series of small launchers (MSTI failed during orbital checkout).
The DC-X, an innovative vertical launch-and-lander demonstrator built by McDonnell Douglas for DOD/BMDO as an early experiment toward eventual reusable single-stage-to-orbit vehicles, completed its fourth successful flight in June, reaching an altitude of 850 m (2800 ft) in a flight lasting 136 sec.
The fifth flight, launched June 27, had to be terminated prematurely because of damage to the aeroshell caused by an explosion in the ground support equipment. Further research efforts were turned over to NASA which was designated the lead agency in the reusable launch vehicle technology area.
Late in December, with the last-but-one Atlas E remaining in its inventory (future Atlas vehicles will come from Martin Marietta Co.), USAF launched the NOAA-14 weather forecasting satellite. The new \Jspacecraft\j joined the NOAA-11 and -12 satellites in polar orbit, where they have been providing data for weather prognosis and atmospheric research. It fills the void left by the failure of NOAA-13 shortly after launch in mid-1993.
#
"Space Year Review 1994: European Space Activities",419,0,0,0
Increasing economic constraints on the program of the European Space Agency (ESA) gave rise to growing concerns regarding Europe's future ability to maintain its commitment to the international space station with its Columbus Orbital Facility segment.
The participation of \JESA\j \Jastronaut\j Jean-Franτois Clervoy in shuttle mission STS-66 in November, accompanying the CRISTA-SPAS free-flyer satellite, signalled to many the end, for the next several years, of direct European engagement in human space flight (except for \JItaly\j's Dr. Umberto Guidoni signing on as payload specialist on the shuttle mission STS-75 in early 1996 to oversee the reflight of the Tethered Satellite System, a "bilateral", i.e., non-ESA, project of \JItaly\j with NASA).
On the commercial front, Arianespace continued its French-based operation as the world's first commercial space transportation company (set up in 1980), even though two of its eight launches in 1994 were failures (losing three satellites), while the payload of a third launch, AT&T's Telstar 402, died shortly after reaching orbit on September 8.
Successfully orbited payloads were Intelsat 7 on June 17, PAS-2 (PanAmSat) and BS-3N (Japan) on July 8, Brasilsat B1 and Turksat IB on August 10, Solidaridad 2 (Mexico) and Thaicom 2 (Thailand) on October 7, and Astra 1D (Europe) on October 31.
#
"Space Year Review 1994: Galileo",420,0,0,0
The deep-space probe \JGalileo\j continued to operate normally (though constrained to transmit its data via low-gain antenna), remaining on schedule to reach Jupiter on December 7, 1995, when its probe will descend into the Jovian atmosphere. \JGalileo\j will then begin two years of observations of the planet, its moons and its magnetosphere.
Unlike all near-Earth observatories, \JGalileo\j had a direct view of the Shoemaker-Levy 9 impact sites on Jupiter, and its solid state imaging system was able to take the only pictures of the impacts at the time when they happened. Data transmitted to NASA's Jet Propulsion Laboratory in early March revealed that \JGalileo\j had discovered a natural satellite of the asteroid Ida during its flyby in August 1993.
The tiny moon, about one mile in diameter, was subsequently named Dactyl by the International Astronomical Union (IAU) -- after the Dactyli, a group of tiny beings in ancient Greek \Jmythology\j residing on Mount Ida where the infant Zeus was hidden and raised by the nymph Ida. Also named were prominent surface features on the asteroid Gaspra, visited by \JGalileo\j on October 29, 1991.
#
"Space Year Review 1994: Hubble Space Telescope (HST)",421,0,0,0
After NASA announced in mid-January that the HST servicing mission aboard the space shuttle Endeavour in December 1993 had been a complete success, the orbiting observatory again turned its attention to the cosmos, rewriting \Jastronomy\j textbooks with virtually every new observation. Results that touched on some of the most fundamental astronomical questions of the 20th Century, such as the existence of black holes and the age of the universe, included:
1.Compelling evidence for a massive black hole in the center of a giant elliptical galaxy 50 million light years away.
2.Observations of great pancake-shaped disks of dust -- raw material for planet formation -- swirling around at least half of the stars in the Orion Nebula, the strongest proof yet that the process which may form planets is common in the universe.
3.Detection of primordial \Jhelium\j, confirming a critical prediction of the Big Bang theory -- that the chemical element \Jhelium\j should have been widespread in the early universe.
4.Discovery of a new \Jquasar\j -- not billions of light years away like its known cousins, but a mere 600 million, hidden behind a dark band of dust around the nucleus of the galaxy \JCygnus\j A.
5.Surprising results in the determination of the age and size of the universe which showed it to be between 8 and 12 billion years old, far younger than previous estimates of up to 20 billion years and apparently conflicting with the age of some nearby stars known to be much older.
6.Evidence ruling out a leading explanation for "dark matter", thought to make up over 90 percent of the mass of the universe, and indicating that any invisible matter, required to explain the slowing expansion of the cosmos, probably consists of exotic invisible and unmeasurable subatomic particles or other unknown material or forces beyond our experience.
7.Identification of primeval galaxies forming less than one billion years after the presumed creation of space and time in the Big Bang as well as views of the earliest moments of the universe revealing a "cosmic zoo" of bizarre fragmentary objects in a remote cluster that are the likely ancestors of our Milky Way Galaxy
8.Observations of the impacts of \JComet\j P/Shoemaker-Levy 9 with Jupiter.
9.Images of the planets Uranus and Saturn revealing several of Uranus' 11 concentric rings of the planet, five of its inner moons, and bright clouds and a high altitude haze above its south pole, as well as an Earth-size storm in the atmosphere of Saturn.
#
"Space Year Review 1994: India",422,0,0,0
India launched its first fully successful medium-lift Augmented Space Launch Vehicle (ASLV) from Sriharikota in Anohra Pradesch province on May 4.
This was the 4th ASLV launch; the first two (on March 24, 1987 and July 13, 1988) both failed and the third attempt in 1992 resulted in early mission termination because of insufficient fourth stage spin-up.
The ASLV carried the 113 kg SROSS C2 satellite (Stretched Rohini Satellite Series).
Also successful was India's first launch, on October 15, of its new Polar Satellite Launch Vehicle (PSLV), designed - with solid-propellant first stage with six strap-on boosters, a liquid-propelled second stage and a solid third stage - to carry sunsynchronous satellites into orbits over Earth's poles.
#
"Space Year Review 1994: International Space Station",423,0,0,0
1994 saw continued progress of the International Space Station program, which produced almost 25,000 pounds of flight-qualified hardware during the year. Among the major developments were a series of formal agreements bringing \JRussia\j into the multinational partnership building the space station.
The year also saw completion of a crucial review of the new space station \Jarchitecture\j, the culmination of months of intensive work following President Bill Clinton's order in February 1993 to substantially reduce the cost and time required to build the orbital laboratory. An amendment that would have terminated the program was defeated by the U.S. House of Representatives by a 123-vote margin in June.
Particular milestones achieved for the space station during the year began with the signing of documents in early February marking the end of the transition from the old Space Station Freedom program to a redesigned project with a leaner management team, a smaller price tag and a quicker development schedule.
Design, development and \Jintegration\j of the program was concentrated under a single prime contract with Boeing Defense and Space Systems Group. In March, NASA managers, the international partners and the contractor community conducted a systems design review involving a comprehensive look at the requirements, configuration and the maturity of the station's technical definition.
In June, NASA and the Russian Space Agency RKA signed documents which put U.S.-Russian space cooperation on a firm basis and underpinned Russian participation in the program. Two packages of space station hardware consisting of 45 solar energy panel modules each were shipped to \JRussia\j, and NASA took delivery from the Energia Production Facility in Kaliningrad of the \Jspacecraft\j docking mechanism that will enable the space shuttle Atlantis to join up with the orbiting Russian Mir space station in 1995.
#
"Space Year Review 1994: Japan",424,0,0,0
Japan "came of age" in space in 1994 with two faultless launches of its powerful new H-2 heavy-lift launch vehicle, the first on February 4, the second on August 28. The launches took place at the new Yoshinobu complex on Tanegashima Island.
Unlike its predecessors, the N-series and the larger H-1 vehicle (which were based on versions of the U.S. Delta rocket), the H-2 was designed and developed entirely by Japanese technology. Its first, or core, stage is powered by a cryogenic (liquid hydrogen/liquid oxygen) rocket engine quite similar to the U.S. space shuttle main engine, called the LE-7, and assisted by a pair of solid rocket boosters strapped to its sides.
The upper stage, taken from the H-1, is also cryogenic, making the H-2 the world's only expendable launch vehicle with both liquid stages powered by LOX/LH2.
Payloads on the maiden flight were an Orbital Reentry Experiment Vehicle (OREX) to test ceramic heat-shield tiles and GPS navigation for the planned HOPE spaceplane and a vehicle evaluation instrument package, totalling 3265 kg.
Geosynchronous insertion of Toshiba's ETS-6 (Engineering Test Satellite) on the second flight failed due to a problem with its kickstage. With a price tag of $150-200 million, the H-2 is currently not competitive with the roughly $60-$80 million for an Atlas 2 or Ariane 4, but rigorous cost reduction efforts will clearly be one of \JJapan\j's top priorities in future space flight activity.
#
"Space Year Review 1994: Magellan",425,0,0,0
NASA's Magellan \Jspacecraft\j, launched in 1989 and having mapped 98 percent of the planet Venus with its synthetic aperture radar since September 1990, entered its last phase by conducting a unique "windmill" experiment designed to return data about the upper atmosphere of Venus and the behavior of the \Jspacecraft\j entering it.
Its wing-like solar arrays were turned in opposite directions, like windmill sails, to encounter pressure from molecules in the upper atmospheric regions, and the measured \Jtorque\j needed to prevent the \Jspacecraft\j from spinning on its axis provided information on aerodynamics and gas-surface interactions around Venus for future mission designs involving aerobraking maneuvers such as planned for the Mars Global Surveyor.
Later, on October 12 (6:02am EDT) radio contact with Magellan was lost when the \Jspacecraft\j started its spiralling descent, burning up in the Venusian atmosphere, it is believed, within two days.
#
"Space Year Review 1994: Mission to Planet Earth (MPE)",426,0,0,0
Besides the MPE missions listed above as part of the space shuttle program, NASA -- on behalf of the National Oceanic and Atmospheric Administration (NOAA) -- launched an Atlas-Centaur with GOES-8 on April 13, the first in a series of next-generation weather satellites. Capable of much longer and more precise atmospheric measurements than its predecessors, the NOAA \Jspacecraft\j enables weather forecasters to more closely track severe storms over land and sea.
Another environmental mission came to an end, when NASA's Nimbus-7 satellite was retired in October after more than 15 years of operation. The \Jspacecraft\j, equipped with a variety of instruments to study the Earth's atmosphere, was the precursor to the current Upper Atmosphere Research Satellite (UARS) and ATLAS missions.
Its most visible success had been the first Total Ozone Mapping Spectrometer (TOMS) instrument, which provided scientists with their first full view of the Antarctic ozone hole and politicians with sufficient evidence for international treaties banning the use of ozone-depleting chemicals (Montreal Protocol).
#
"Space Year Review 1994: Russian Space Activities",427,0,0,0
Even with severe political and economic problems still pervading most of the former Soviet Union, \JRussia\j again succeeded in launching more \Jspacecraft\j in 1994 than all other nations and entities combined: 49 vs. 44.
With space flight rating that high as a military/civil-governmental/scientific tool, showcase of high-technology and object of national pride, the newly entered space station partnership with the United States, initiating cooperative operations in space, must be regarded as a particular significant political accomplishment on both sides.
Cosmonaut Sergei Krikalev's presence on board the shuttle Discovery on mission STS-60 opened the door to more joint activities in 1995 and coming years, reinforced by his colleague Vladimir Titov's joining the crew for the flight of STS-63/Discovery in February 1995.
Space Station Mir
By the end of 1994, the Mir space station had been in orbit for 3237 days, commencing in February 1986. During that time, it circled Earth approximately 50,670 times in an orbit 245 miles high and 51.65 degrees tilted against the equator.
Counting from its last brief period of nonoccupancy (September 1989) to the end of 1994, Mir has been inhabited continuously for 1943 days. Since its inception, it has been visited 17 times by two- to three-person crews.
In 1994, on January 10, Soyuz TM-18 (launched 1/8) arrived with three cosmonauts, one of which (Valery Polyakov) was still in space at end-1994. The other two returned to Earth after a six-months stay, shortly after the arrival, on July 1, by Soyuz TM-19 with two new occupants. After Soyuz TM-20 followed on October 4 with three more cosmonauts, including the German Ulf Merbold for \JESA\j, the TM-19 crew returned to Earth after four months on November 4, accompanied by Merbold (31 days).
At the end of 1994, three occupants remained in Mir, one the third Russian woman in space (Elena Kondakova).
In other space flight areas, \JRussia\j continued its numerous launches of military, scientific, and telecommunications satellites, among them the new large data relay satellite "Luch", required by upcoming joint U.S./Russian operations on Mir, on a Proton rocket on December 16. This heavy-lifter alone was used in a total of 12 launches in 1994.
#
"Space Year Review 1994: Solar System Exploration",428,0,0,0
For space sciences and \Jastronomy\j, several missions provided spectacular findings that captured headlines worldwide. A particularly seminal event for astronomers was the impact of at least 21 fragments of the \Jcomet\j P/Shoemaker-Levy 9 with Jupiter in July.
Not only had the impact been predicted accurately -- to the day -- after the discovery of the \Jcomet\j in 1993 but it also afforded an unprecedented world-wide campaign to observe this event from ground-based and space-based observatories and thereby acquire a maximum amount of data about the composition of comets and the Jovian atmosphere.
#
"Space Year Review 1994: Ulysses",429,0,0,0
The NASA/ESA-spacecraft Ulysses -- the first probe to explore the Sun's environment at high latitudes -- completed the first phase of its primary mission of exploring the complex forces at work in the polar regions of the Sun when it overflew the Sun's southern pole on November 5, continuing its sweeping path on which it will also pass over the north pole in June 1995.
Surprising findings included the very high velocity of about 750 kilometers per second (470 miles per second) of the solar wind flowing from the polar region, nearly double its speed at lower latitudes, the lack of clear evidence of a magnetic pole, and the unexpectedly low intensity of cosmic rays in the polar region.
#
"Space Year Review 1994: United States Space Activity",430,0,0,0
Centerpiece of the United States space program, the Space Shuttle continued to provide highly successful access to space for humans at a reduced cost to the nation. At the same time, dramatic reshaping and redesign efforts in the space station program yielded funding reductions around $6 billion for the period from Fiscal 1994 to 1998.
#
"Space Year Review 1994: Summary",431,0,0,0
1994 was the year of the 25th anniversary of the first Moon landing, by Apollo 11 on July 20, 1969 (as well as of the second, Apollo 12). It also saw a number of significant events which redefined the world's space programs for the remainder of this century.
Dominating the news from the National Aeronautics and Space Administration (NASA) in 1994 were the work of the newly repaired Hubble Space \JTelescope\j and the emerging international Space Station program's acquisition of \JRussia\j as a new major partner and of the Boeing Company as prime contractor.
Both for the United States and other countries involved in space flight, increasingly tight economic conditions reduced the number of new program starts and made the need for international consolidation and cooperation more evident than ever before.
Despite manpower and budgetary reductions, 1994 was one of the busiest years in the USA's quest to better understand the mechanisms that drive the climate and \Jecology\j of Earth, and how human activity is affecting the environment, since the inception of NASA's space-borne Mission to \JPlanet\j Earth program in 1992.
In Solar System Exploration, spectacular news were made by the Hubble Space \JTelescope\j and the Gamma Ray Observatory in Earth orbit, Magellan at Venus, Ulysses near the Sun, Clementine at the Moon, \JGalileo\j in deep-space en-route to Jupiter (which made headlines when it collided with fragments of \JComet\j P/Shoemaker-Levy 9), and Voyager 1 and 2 on their way into interstellar space, still taking data on energy fields and particles in researching the heliosphere, the boundary of the Sun's influence.
#
"Year in Space 1995",432,0,0,0
\J1995: Space Launches\j
\JSpace Flight 1995: United States Space Activity\j
\JSpace Flight 1995: Asian Space Activities\j
\JSpace Flight 1995: United States Commercial Space Activities\j
\JSpace Flight 1995: United States Department of Defense Space Activity\j
\JSpace Flight 1995: European Space Activities\j
\JSpace Flight 1995: United States International Space Station Activities\j
\JSpace Flight 1995: Russian Space Activities\j
\JSpace Flight 1995: United States Space Sciences and Astronomy Activities\j
\JSpace Flight 1995: Summary\j
#
"1995: Space Launches",433,0,0,0
(Refer to Table)
#
"Space Flight 1995: United States Space Activity",434,0,0,0
The four-vehicle fleet of the United States space shuttle continued its operation of carrying people and payloads to and from Earth orbit for science, technology and operational research. On the ground, the International Space Station program, after passing critical budgetary hearings in the U.S.
Congress, moved forward briskly with the finalization of the government/contractor development team, successful cooperation negotiations with the participating international partners, and accomplishment of critical review milestones.
Space shuttle. During the year, NASA successfully completed seven space shuttle missions into Earth orbit, three of which went to Mir, bringing the total number of shuttle flights since inception in 1981 to 73 and setting a new flight-duration record for shuttles.
As in the previous year, the shuttles carried 42 astronauts into space (including 10 women), with foreign crew members from \JRussia\j and Canada. In March, NASA welcomed the new \Jastronaut\j class of 1995 to Johnson Space Center, to begin a year of familiarization training. The 10 pilot and 9 mission specialist candidates, including 5 women, will be joined by two international mission specialists (Canada and Japan).
STS 63. Discovery flew its 20th mission from February 3 to 11 with the second Russian \Jcosmonaut\j, Vladimir G. Titov, and the first woman shuttle pilot, Eileen M. Collins. As a precursor and dress rehearsal for the subsequent docking missions, the \Jspacecraft\j performed a rendezvous with and fly-around of the space station Mir.
While no link-up was attempted yet, new flight techniques as well as the coordination between the mission control teams at Houston and Moscow required for such missions were validated and the two 100-ton systems closed to a distance of 36 ft. (11 m).
Later, after the three-hour joint flight, Titov used the shuttle's robot arm to release the free- flying \Jastronomy\j satellite SPARTAN-204 to make far-ultraviolet (UV) spectroscopic observations of the interstellar medium; it was retrieved two days later. The mission also included a 4h 39m spacewalk (EVA) by Mission Specialists Bernard Harris and Michael Foale to evaluate spacesuit modifications and demonstrate large object handling techniques. With Eileen Collins and Mission Specialist Janice Voss in Discovery and \JCosmonaut\j Elena Kondakova aboard Mir, three women were in space simultaneously for the second time (first time: STS 40).
STS 67. Setting a new shuttle mission record of 16 days, 15 hours in space as well as a distance mark of 6.9 million miles, Endeavour, from March 2 to 18, carried the astronomical payload ASTRO on its second mission. Its three unique telescopes for UV \Jastronomy\j mapped areas of space still largely uncharted in that spectrum, collected UV spectra of about 300 celestial objects including Jupiter and the Moon, and measured intergalactic gas abundances. Other experiments in weightlessness concerned large structures control dynamics and materials processing.
STS 71. Originally planned for launch after STS-70, which was delayed for a month, Atlantis was launched to Mir on June 27 as the 100th human flight in the United States space program and the first Phase I docking mission, 20 years after the historic joint Apollo/Soyuz Test Project (ASTP) between the United States and the then-Soviet Union.
Aboard the shuttle were the replacement crew (Mir-19) Anatoliy Solovyov and Nikolai Budarin. Steered by commander Robert "Hoot" Gibson, the orbiter docked to Mir's Kristall module on 29 June, forming the largest human-made structure in space to date. After hatch opening and exchange of greetings, joint scientific (mostly medical) investigations were carried out for 5 days, with a record number of 10 individuals aboard a single space vehicle.
Atlantis undocked on July 4 after unloading water and other supplies for Mir and taking on equipment no longer needed, and returned to Earth on July 7 with the previous Mir-18 crew Vladimir Dezhurov, Gennadiy Strekalov and NASA \Jastronaut\j Dr. Norman Thagard who had flown to Mir on March 14 in Soyuz TM-21 as a guest \Jcosmonaut\j.
STS 70. Delayed from its June liftoff date by repair of minor damage to the cork \Jinsulation\j of its external fuel tank caused by a woodpecker, Discovery flew from July 13 to 22 to launch the seventh Tracking and Data Relay Satellite (TDRS) since 1983 into geosynchronous orbit, and to conduct research with a commercial protein growth facility, a bioreactor demonstrator and various biological experiments in coordination with the National Institutes of Health.
It was also the first flight of the new Block I Space Shuttle main engine with increased stability and safety. The time between its launch and the landing of the previous mission - only six days - was the shortest interval in the shuttle program.
STS 69. Endeavour at first had to be rolled back to the Kennedy Space Center's Vehicle Assembly Building on August 1 because of Hurricane Erin and then required some repair on its solid rocket boosters. It was launched on September 7, carrying the SPARTAN-201 freeflyer satellite with instruments to observe the solar wind and the sun's outer atmosphere, the Wake Shield Facility (WSF) on its second flight to experiment with the production of advanced, thin-film semiconductor materials, and various other payloads.
Mission specialists James Voss and Michael Gernhardt conducted a 6 hour, 46 minute spacewalk, the 30th of the shuttle program, to test construction tools, an arm-sleeve computer and suit modifications for space station assembly. Return to Earth was on September 18.
STS 73. Flown from October 20 to November 5 after six delays (Hurricane Opal, a \Jhydrogen\j leak, mechanical problems, and weather conditions), Columbia on its 18th flight carried the second United States \JMicrogravity\j Laboratory with science and technology experiment from government institutes, universities and industry in areas such as fluid physics, materials science, biotechnology, combustion science, and commercial processing technologies. During the mission, students from a number of high schools interacted with the seven-member crew, discussing and comparing onboard \Jmicrogravity\j experiments with similar ground-based experiments.
STS 74. Performing its 15th mission from November 12 to 20, Atlantis was steered by commander Kenneth Cameron and pilot James Halsell to a perfect second docking with Mir during its 44th orbit (on 11/15, 1:27am EST), carrying a Russian-built 9000 lb. (4000 kg) docking module, two new solar arrays, 992 lb. (450 kg) fresh water and 1140 lb. (517 kg) equipment and food to the station.
After three days of joint operations by the eight occupants from four of the five ISS partner countries -- USA, \JRussia\j, Canada (shuttle mission specialist Chris Hadfield) and Europe (Germany's EUROMIR-cosmonaut Thomas Reiter) -- Atlantis undocked on 11/18 (2:16am EST), taking with it 816 lb. (370 kg) of science samples, data and equipment from previous Mir investigations, leaving the docking module attached for future shuttle visits.
#
"Space Flight 1995: Asian Space Activities",435,0,0,0
In 1995, space activities continued in \JJapan\j, the People's Republic of China, and South Korea.
Japan. After last year's successful introduction of the powerful H-2 heavy-lift launch vehicle with two flights, \JJapan\j's NASDA scored a third success from Tanegashima on March 17 launching twin payloads. The first, the SFU (Space Flyer Unit) experiment carrier, was deployed in Earth orbit at about 206 mi. (330 km) altitude (from where it was retrieved by the US-shuttle Endeavour on January 13, 1996).
The liquid oxygen- liquid \Jhydrogen\j LE-5A second stage then ignited again and injected the second payload, the Geostationary Meteorological Satellite 5, later named "Himawari-5" (Sunflower-5), in a geostationary orbit at 22,500 mi. (36,000 km) altitude.
On January 15, \JJapan\j's space agency ISAS launched a four-stage Mu-3S-II from Kagoshima, carrying the German Experiment Reentry Space System (EXPRESS), a Russian-built reentry vehicle. It was due to land in \JAustralia\j after a five-day flight, but the thrust vector control system on the second stage malfunctioned during ascent, and the mission failed. Given up for lost in the Pacific ocean, EXPRESS came safely down on its orange parachute 2.5 orbits later near a partially inhabited area of \JGhana\j, West Africa.
Japan Satellite Systems' communications satellite JCSAT 3 was launched by a US-Atlas on August 29. On the same day, the satellite N-STAR of Nippon \JTelephone\j and Telegraph (NTT) of \JJapan\j was carried into space by a European Ariane.
China. After its successful launches of the Long March (Chang Zheng) 2E in 1994, the People's Republic of China suffered a serious setback on January 25 with the explosion of an LM-2E approximately 50 sec. after launch, that killed six persons near the Xichang Satellite Launch Center and injured 27 others by falling debris.
A second LM-2E was launched successfully on November 28, carrying the Lockheed Martin comsat Asiasat 2 into geostationary orbit to provide services for the Hongkong-based Asia Satellite Telecommunications Co. Ltd. On December 28, a third LM-2E carried the TV-broadcast satellite EchoStar 1 successfully into orbit.
South Korea. South Korea got its first communications satellite with the (only partially successful) launch of Mugunghwa (Hibiscus) on a Delta 2 on August 5. The country also entered into negotiations with RKA for future participation in \JRussia\j's \Jcosmonaut\j training program.
#
"Space Flight 1995: United States Commercial Space Activities",436,0,0,0
To assist in early space shuttle missions of the International Space Station program, NASA in August contracted with Spacehab Inc. to lease a pressurized in- shuttle payload module, Spacehab, for ferrying cargo to Mir.
Currently booked for four flights of the double-length version, the company is confident of future new business for similar services. In a first small step toward the eventual "privatization" of the shuttle, NASA in August announced plans to select a single contractor to operate the space shuttle fleet, replacing 85 separate shuttle contracts.
In October, NASA decided to negotiate with United Space Alliance, a joint venture formed by the two firms Rockwell and Lockheed Martin for this purpose. Also in the commercial space sector in 1995, Boeing entered an agreement with the Russian firm Rocket Space Corp. (RSC) Energia to provide a turn-key service for placing and operating commercial and government payloads on the outside of the Mir space station.
In 1994, U.S. aerospace industry revenues had totalled $113 billion, with a sizable portion coming from foreign sales (trade surplus: $26 billion), and employed about 800,000 highly skilled workers. However, recent political and economic events, e.g., lower defense budgets, have changed the picture: U.S. companies were forced to restructure and downsize their operations, overseas markets are growing more vigorously than the U.S. market, and an increased number of government- supported manufacturers abroad add to the business challenge.
U.S. companies intending to be major supplier of space launchers in coming years are Lockheed Martin, which resulted from a merger of Lockheed and Martin-Marietta, and McDonnell Douglas. In 1995, the former launched 11 Atlas-2AS rockets, one refurbished Atlas-E (ICBM) and 4 Titan-4's, the latter company three Delta-2 vehicles
To replace the current Atlas rockets, Lockheed Martin is developing the Atlas-2AR, possibly using Russian liquid-fueled engines, as a next-generation medium/intermediate-class vehicle intended as the core of a new family of expendable launchers.
A smaller rocket, the Lockheed Launch Vehicle (LLV) is being developed as low-cost booster for commercial payloads between 1-4 tons to low-Earth orbit. The LLV's first launch, however, failed when the vehicle pitched out of control shortly after liftoff on August 15. McDonnell Douglas announced the development of the Delta-3, a derivative of the smaller but highly reliable Delta-2 rocket (93 launches, 98.9% successful over the past 17 years) using a larger payload fairing and a cryogenic upper stage. Its payload capability to geostationary transfer would be 8380 lb. (3800 kg), similar to the current Atlas-2 and Europe's Ariane-4.
Additional newcomers to the launch services market are two small companies, EER Systems and Orbital Sciences Corp. (OSC) The first launch of the former Conestoga-1620 rocket ended in failure on October 23, destroying its \JMeteor\j payload, a commercial \Jmicrogravity\j recoverable capsule. OSC's Pegasus-XL, launched from a Lockheed L-1011 airplane, succeeded on April 3 to launch three Orbcomm satellites, but its third flight on June 22 with a U.S. Air Force Space Test Experiment Platform (STEP) satellite as payload, was a failure, the second in the program.
OSC's Orbcomm, comprising 26 satellites for two-way communications using handsets, is one of a number of fiercely competing space-based global communications systems in early stages of build-up. The large Iridium system, using 66 satellites to provide worldwide digitally-switched point-to- point communications with multiple networks -- terrestrial cellular, public \Jtelephone\j, and space-based satellites -- is in full-scale development by Motorola and its partners, the first \Jspacecraft\j being prepared for launch in late 1996.
Loral/Qualcomm's Globalstar will employ 48 satellites on low- Earth orbits, while TRW continues development of its Odyssey system, an Atlas-launched, 12-satellite network for voice, data, paging and messaging to mobile subscribers beginning in 1999. ICO Global Communications is planning a commercial spin- off of Inmarsat to provide mobile communications over 10 satellites.
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"Space Flight 1995: United States Department of Defense Space Activity",437,0,0,0
U.S. military space organizations continued their post-Cold War "cultural change" move to make space a routine part of military operations across all service lines. Joint initiatives are aimed at bringing launch and satellite operations increasingly on a level where they can be of maximum use in directly supporting military forces in the field in a rapidly changing world.
At Cape Canaveral, continuing its build-up of spacebased resources, the Air Force launched a classified payload on a Titan-4/Centaur on May 14. Originally set for April 11, the liftoff was delayed twice by equipment and facility problems. A second Titan-4/Centaur with a classified satellite took off on July 10. On July 31, a DSCS-3 (Defense Satellite Communications System) \Jspacecraft\j -- the fifth in a \Jconstellation\j of upgraded military relay stations -- was boosted into orbit by a commercial Lockheed-Martin Atlas-2/Centaur, the 16th successful launching in a series for an Atlas. During a ground test-firing of a Titan-4 second stage on July 31, a nozzle extension made of composite material failed, leading to launch delays of several months.
The next Titan-4/Centaur launched the second MILSTAR DFS communications satellites on November 6. The 10,000-lb. (4500-kg) \Jspacecraft\j will work in concert with MILSTAR DFS-1, deployed last year, to provide jam-proof, secure communications reaching from the Persian Gulf to the western Pacific Ocean. Four upgraded MILSTARs with higher data relay rates will be launched between 1999 and 2002 to provide a global military communications network.
A fourth Titan-4 took off on December 5, carrying a classified reconnaissance satellite. It was the first launch from Vandenberg AFB, Calif., since a Titan-4 failed after liftoff in August 1993, and the fifth from there since first launch in March 1991. Another ten Titan-4's have been launched from Cape Canaveral to date.
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"Space Flight 1995: European Space Activities",438,0,0,0
In a major decision by the research/technology ministers of its membership countries meeting in \JToulouse\j on October 20, the European Space Agency reached final agreement regarding Europe's participation in the ISS project.
Affirming prior commitments, \JESA\j's contribution comprises the Columbus Orbital Facility (COF) module docked to the station, and the Ariane 5-launched Automated Transfer Vehicle (ATV). COF is being developed by \JGermany\j's Daimler-Benz Aerospace (DASA) and built in \JItaly\j. With about 41½ percent, \JGermany\j has assumed the largest share in the total cost.
In cooperation with \JRussia\j's RKA, on September 3 \JESA\j undertook the EUROMIR 95 mission -- the second flight by an \JESA\j \Jastronaut\j aboard Mir. German \Jcosmonaut\j Thomas Reiter was launched as a member of the Mir-20 crew (see Soyuz TM-22) to study living and working conditions in space with an extensive program of 47 experiments in life sciences, astrophysics, materials science and technology. Reiter, still on board at year's end, became the first guest \Jcosmonaut\j given the rank of "flight engineer".
Public deficits and rapidly declining military budgets began to force European aerospace industry into a deep crisis of restructuring in order to remain competitive. According to recent European trade analyst reports, an estimated 100,000 aerospace jobs have been lost by the 15-member-state European Union in the last 10 years.
In \JGermany\j, aerospace industry initiated major revisions to original strategic plans for growth, with a workforce down to 68,000 from 95,000 in 1991. Similar trends and moves are evident in \JFrance\j, Holland, \JItaly\j and \JBelgium\j. The downward trends, however, did not appear to affect the launch services market. \JESA\j and Arianespace in 1995 conducted brisk business at Kourou/French \JGuyana\j with the launch of 11 Ariane rockets carrying a variety of payloads, such as TV satellites, Europe's first military reconnaissance satellite \JHelios\j 1A, the giant infrared space \Jtelescope\j ISO and the remote sensing/environmental satellite ERS-2.
As compared to \JESA\j's activities, national space programs in Europe were forced by economics to remain at an almost negligible level. The German EXPRESS reentry capsule, launched on a Japanese rocket on January 15 (see Japan), did not succeed due to launch vehicle failure. In continuing its cooperation with \JRussia\j of prior years, \JGermany\j's space agency DARA entered an agreement with RKA for launching one of two German cosmonauts already in training in \JRussia\j to Mir toward the end of 1996.
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"Space Flight 1995: United States International Space Station Activities",439,0,0,0
International Space Station. In 1995, the development of the International Space Station (ISS), begun in 1994 after the redesign of the former Freedom concept and formal signings of the agreement between NASA and the Russian space agency RKA (Rossiyskoe Kosmicheskoe Agentstvo), successfully passed critical Congressional budget hearings and made significant progress, remaining on schedule and within costs.
A $5.63 billion design and development contract with the prime contractor, Boeing Co., was definitized and signed after 16 months of tough negotiations. Under the cost-plus-incentive-fee and award-fee contract extending through June 2003, Boeing is responsible for \Jintegration\j and verification of the ISS system, as well as for design, analysis, manufacture, verification and delivery of the U.S. on-orbit segments of the station. Assembly is to begin in November 1997 with the Proton-launch of the U.S.-purchased Russian "FGB" tug (Funktsionalya-gruzovod blok, Functional Cargo Block).
In March, ISS program management in the first of a series of Incremental Design Reviews successfully provided a comprehensive assessment of the design and technical feasibility for the first six U.S. and the first five Russian ISS assembly flights as well as a forward-planning review of all assembly flights.
In April, a major design review of the first construction element, the FGB, certified readiness to proceed with the manufacture of this Russian-built propulsion, guidance and control module. Fabrication of the structure of NASA's first pressurized module, called Node 1, was completed in June, while Node 2, to be used as structural test article before its launch in 1999, was delivered in April.
In September, Boeing also completed the main structure of the 28 ft. (8.53 m) long, 14 ft. (4.26 m) wide U.S. laboratory module, weighing about 6000 lb. (2700 kg). Qualification testing of the "alpha joint" for the space station's rotating solar arrays was begun, and construction of the first flight unit has started. Overall, U.S. contractors in 1995 delivered nearly 80,000 lb. (36,300 kg) of hardware including solar array panels, mast, truss segments, rack structures, hatch assemblies and various mockups.
Phase I of the ISS development, the joint Shuttle-Mir program, proceeded on schedule to meet its objectives of providing operations experience, ISS risk mitigation, technology demonstrations, and early science opportunities.
Major milestones during 1995 were three visits to Mir by U.S. space shuttles, including two link-ups (see Space shuttle), ferrying of cosmonauts and supplies by Atlantis and the first participation by a U.S. \Jastronaut\j, Dr. Norman Thagard, as a member of a Russian station crew, starting in March with the launch of Soyuz TM-21 from Baikonur, Kazakstan and ending in July with his return aboard Atlantis (see Soyuz TM-21).
Dr. Thagard's stay in space yielded the first long-duration medical data on an American \Jastronaut\j since the Skylab program in 1973. With the crews of TM-21 (3), Mir (3), and STS-67/ Endeavour (7) in orbit, the total count of humans simultaneously in space in March reached the new record of 13.
Development programs continued in other ISS partner countries as well. In Canada, the Mobile Service System (MSS) which will provide external station robotics, progressed after resolution of federal budget limitations in 1994.
Japan remained on schedule in developing the Japanese Experiment Module (JEM), and in Europe the European Space Agency (ESA) in October received final approval by the governments of the nine countries involved in ISS to proceed with the development of a pressurized laboratory called Columbus Orbital Facility (COF) and the Automated Transfer Vehicle (ATV) for supplying logistics and reboosting the station (see European Space Activities).
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"Space Flight 1995: Russian Space Activities",440,0,0,0
Represented by its space agency RKA, \JRussia\j in 1995 continued its robust space operations at a brisk pace, even if at a level considerable lower than in previous years. Again, it led the world in number of \Jspacecraft\j launches: 32 successes in a total of 74 (43%).
The new partnership with the United States gained major substance in a number of crewed flight activities and began to play a major role in shaping the Russian space program. After \Jcosmonaut\j Sergei Krikalev's participation in shuttle mission STS 60 in 1994, three more cosmonauts flew on the U.S. shuttles Discovery (STS 63) and Atlantis (STS 71) in 1995 (see Space shuttle).
Space station Mir. By the end of 1995, Mir had been in operation for 3602 days, commencing in February 1986. In that time span, it circled Earth approximately 56,380 times at 246 mi. (393 km) altitude in an orbit inclined 51.65 degrees to the Equator. Counting from its last brief period of nonoccupancy (September 1989) to the end of 1995, Mir has been inhabited continuously for 2308 days.
Since its inception, it has been visited 24 times, incl. twice by a U.S. shuttle, by two- to three-person crews. To resupply the occupants during 1995, the space station was visited by five automated Progress M cargo ships, bringing the total of Progress and Progress M ships launched to Mir and the two preceding space stations Salyut-7 and Salyut-6 to 73, with no failure.
Soyuz TM-21. Launched on March 14, TM-21 (spacecraft no. 70) carried the Mir-18 crew Vladimir Dezhurov, Gennadiy Strekalov and U.S. \JAstronaut\j Dr. Norman Thagard. Docking occurred on 3/16. Thagard became the first American launched in a Soyuz and during his stay on Mir established a new U.S. record in space of 115d 9h 44m. (previous mark: 84d, by Skylab 3 in 1973/74).
The Mir-18 crew returned to Earth on the shuttle Atlantis on July 7. On September 11, TM-21 served as return vehicle for the Mir-19 crew Anatoly Solovyev and Nikolai Budarin who had arrived with the Atlantis and stayed on Mir for 75d 11h.
Soyuz TM-20. In space since its launch on October 3, 1994, TM-20 returned to Earth on March 22, landing 13.7 mi. (22 km) northeast of Arkalyk, Kazakstan, and bringing back the Mir-17 crew Alexander Viktorenko, Elena Kondakova, and Valeriy Polyakov.
Kondakova established a new women's space endurance record of 169 days (previous mark: 14d, by Chiaki Mukai, Japan), while Polyakov, a 52-year old physician who had been on Mir since January 1994, set a new overall record of 437½ days. Combined with an earlier stay on Mir, he has now logged a total of 607 days in space and circled the Earth nearly 10,000 times.
Spektr. The 20-ton uncrewed Spektr module was launched on May 20 on a Proton rocket and docked to Mir on June 1, joining the similar Kvant-2 and Kristall blocks. Spektr, based on a military TKS \Jspacecraft\j first tested in 1977, carries an array of remote sensing instruments, a small manipulator arm, and a small science airlock.
Soyuz TM-22. The Mir-20 crew Yuri Gidzenko, Sergei Avdeyev and Thomas Reiter was launched on September 3 and docked to Mir two days later. Reiter, a German sponsored by \JESA\j, became the first Western European to perform a spacewalk (EVA), on October 20 for 5¼h.
During their stay, the crew was informed that their mission was to be extended through February 29, 1996. In a second EVA of 37 min., the 50th conducted from Mir, Gidzenko and Avdeyev prepared the remaining free port on the station's central transition node for next year's link-up of the Priroda module by attaching a docking cone.
In other space-flight areas, \JRussia\j launched about 24 military, scientific, and telecommunications satellites, among them the second Gals direct-broadcast TV satellite, and the new large data-relay satellite Luch-1, a more powerful version of the earlier Luch (Altair).
The proven heavy-lift carrier Proton (to date: 191 flights, with over 96% reliability for the last five years) was used in seven launches, three carrying geosynchronous payloads, another three launching nine GLONASS satellites in highly elliptical 12-hour orbits (completing the GLONASS system), and one used for Spektr.
GLONASS (Globalnaya Navigatsionnaya Sputnikovaya Sistema), the Russian satellite navigation system, currently has 25 satellites in orbital planes of 11,937 mi. (19,100 km) altitude, 11h 15m period and 64.8 degree inclination. Originally developed for military use, it is available to civil users since March, like its U.S. counterpart GPS but superior to the latter's 100-meter resolution by offering 50-70-m accuracy.
Commercial space activities. In its early stages of entering commercial space markets, Russian rocketry suffered a setback by a failure of a new "Start" booster on March 28 from Plesetsk Cosmodrome on its first commercial mission, carrying three \Jspacecraft\j from \JRussia\j, Israel and Mexico.
To market its launch services, the Russian firm Khrunichev, builder of the Proton, entered a joint venture with Lockheed Martin and the Russian company RSC Energia. In 1993, the resulting LKEI and Lockheed Martin Commercial Launch Services had formed International Launch Services (ILS) to market the Proton.
By mid-1995, ILS had already firmly booked about $1 billion worth of Proton launches. A new heavy booster, "Angara", larger than the Proton, is under development at Khrunichev for the early 21st century. Other high-quality space technology offered to Western companies include rocket propulsion systems like the NK-33 from \JSamara\j State Scientific and Production Enterprise, and the RD-180 from NPO Energomash.
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"Space Flight 1995: United States Space Sciences and Astronomy Activities",441,0,0,0
A plethora of important and, in part, revolutionary discoveries in space from several automated or remotely-directed missions this year enriched our knowledge about the cosmos.
Hubble Space \JTelescope\j. A string of celestial surprises from the perfectly functioning Hubble Space \JTelescope\j (HST) stunned space researchers by shaking the foundations of prior concepts about cosmological phenomena and challenging some of the most accepted and treasured theories about the form, buildup, structure, age, evolution and future of the universe.
The country's imagination particularly was captured by HST images of about 100 glittering new stars being born in the Eagle Nebula (M16), 7,000 light-years (ly) away in the \Jconstellation\j Serpens. Numerous front pages as well as network and local news programs carried the images, and they were rebroadcast several times by CNN. Many people called in to talk about their feelings and what they believed they saw in the pictures.
Other findings by HST included: Detection of critical stars called Cepheid variables in the remote Virgo cluster of galaxies, allowing better determination of the age of the universe; compelling evidence for the existence of supermassive black holes in three galaxies: a 2.4-billion-solar mass black hole in the core of the elliptical galaxy M87, a 40-million- solar-mass black hole in the spiral galaxy NGC 4258, and an extremely puzzling black hole 100 million ly away in the direction of the \Jconstellation\j Virgo.
Fueled from an 800 ly- wide spiral-shaped disk of dust around it, it is offset by 20 ly from the center of the host galaxy NGC 4261. Having presumably once been at the center, something must have pulled it outward -- or else it is "self-propelled" by the rocket-like reaction from plasma jets expelled by temperatures of tens of millions of degrees.
Hubble photographed distant galaxies through the galaxy cluster Abell 2218 which acts as a spectacular "gravitational lens", and detected mature, i.e. fully-developed, spiral galaxies that existed already when the universe was only 2 billion years old, bizarre, never-before- seen light structures in distant radio galaxies, blue dwarf galaxies, at least two (probably even four) new moons of Saturn, as well as signs of fresh volcanic activity on Jupiter's moon Io and of oxygen on its moons Europa and Ganymede.
Ulysses. The NASA-ESA solar-polar explorer Ulysses, having overflown the Sun's southern pole in November 1994, this year in June began its pass over the north pole and concluded its main mission successfully in September, while already on its return-journey out to the orbit of Jupiter, 500 million mi. (800 million km) away, to arrive there in April 1998. It will then return in September 2000 for more polar overflights.
The German-built probe for the first time detected periodic oscillations or wave motions in interplanetary space originating from deep within the Sun and took the first "snapshot" of the spiral structure of the Sun's magnetic field extending past the orbit of Venus toward Earth's orbit.
Some of its findings concerning the global differences in solar wind speed at different latitudes (up to twice as fast at high southern latitudes than near the equator, i.e., 500 mi./s [800 km/s] vs. 250 mi./s [400 km/s]), and the spiral magnetic field could mean an upset of the current solar model of physicists, necessitating its revision. By end-1995, Ulysses had travelled a total distance since its launch in October 1990 of about 2 billion mi. (3.26 billion km).
Galileo. After taking a circuitous route lasting six years and involving one gravity-assist swing-by of Venus and two of Earth, the Jupiter probe \JGalileo\j finally reached its target after traveling 2.3 billion mi. (3.7 billion km). Throughout 1995, it had continued its long string of observations of targets of opportunity begun after its launch on October 18, 1989.
On July 13, three explosive bolts connecting the entry probe to the main ship were detonated, sending the probe spinning to Jupiter. On July 27, \JGalileo\j fired its German- built main propulsion system for the first time, changing its course to pass 133,000 mi. (213,000 km) above Jupiter's clouds.
On December 7, after close approaches to the moons Europa (19,325 mi., 30921 km) and Io (552 mi., 884 km), \JGalileo\j reached its closest point to Jupiter (perijove) of 134,000 mi. (214,500 km). At 6:10pm EST, it began to receive and store 57 minutes of radio transmissions from the probe when it entered the atmosphere at latitude 6.5 deg N, longitude 4.4 deg W in the North-Equatorial Band. The data were subsequently relayed to Earth. Then, igniting its rocket engine on time at 8:20pm and firing it for 49 min., \JGalileo\j established itself in a perfect orbit around Jupiter and began its two-year mission of scientific studies of the Jovian system.
GOES 9. The second in a series of advanced weather satellites, GOES 9 (Geostationary Operational Environmental Satellite) was launched by NASA on May 23 on an Atlas-2AS/Centaur rocket. After several months of testing, the new powerful observer was handed over to the National Oceanic and Atmospheric Admninistration (NOAA).
SOHO. The joint NASA/ESA solar observatory SOHO (Solar and Heliospheric Observatory) was launched on December 2 on an Atlas/Centaur. The 4100 lbs (1860 kg) probe, manufactured by Matra Marconi Space of \JToulouse\j, \JFrance\j, is the most sophisticated solar space observatory ever built and one that promises to revolutionize solar physics when science operations begin in mid-1996, after the \Jspacecraft\j has reached its location about 1 million mi. (1.6 million km) from Earth on the Earth-Sun line where it will use its propulsion system to orbit the so-called L1 Langrangian point.
XTE. NASA's X-Ray Timing Explorer (XTE) satellite, the largest X-ray \Jtelescope\j orbited to date, was launched on December 30 on an upgraded two-stage Delta 7920-10 after six prior launch attempts, five of them caused by high upper-atmospheric winds. The 6700 lbs (3040 kg) \Jspacecraft\j, measuring 6 x 6 x 18 ft (1.8 x 1.8 x 5.5 m), was placed in a 360 mi. (576 km) orbit inclined 23 deg. to the Equator.
Built by the Goddard Space Flight Center in Maryland, the XTE is equipped with three sophisticated "telescope" instruments, the Proportional Counter Array (PCA), the High-Energy X-Ray Timing Experiment (HEXTE), and the All-Sky Monitor (ASM), to study dense objects such as white dwarf and neutron stars, binary systems, X-ray novae, black holes, active galactic nuclei and quasars throughout the universe. XTE will provide accurate timing and measurement of X-ray sources in the sky and can detect emissions as brief as 10-100 microseconds.
Voyager 1 and 2. Both Voyager deep-space probes were healthy by end-1995, continuing their departure from the solar system.
As they travel farther and farther from the Sun, they are returning data to characterize the environment of the outer solar system, searching for the heliopause -- the boundary representing the outer limit of the Sun's magnetic field and outward flow of the solar wind. Voyager 1, cruising at 39,260 mph (17.45 km/s), is currently 5.78 billion mi. (9.24 billion km) from Earth, having travelled 6.84 billion mi. (10.95 billion km) since its launch in September 1977.
Voyager 2, at a distance of 4.37 billion mi. (7 billion km), is departing our system at 36,160 mph (16.07 km/s), having travelled a distance of 6.48 billion mi. (10.37 billion km) since its launch in August 1977. Both \Jspacecraft\j are expected to operate and send back valuable data until at least 2015.
Pioneer. On September 30, NASA ended operations of Pioneer 11, one of the most durable and productive space missions in history, after nearly 22 years of exploration out to the farthest reaches of the solar system. Now far beyond the orbit of Pluto and over 4 billion mi. (6.5 billion km) from Earth, the probe is heading out into interstellar space. At that distance, faint signals from it travelling at the speed of light take over six hours to reach Earth.
Because the \Jspacecraft\j's power is too low to operate its instruments and transmit data, communications with it have been reduced from about 8-10 hours a day to about 2 hours every 2 to 4 weeks. Late in 1996, its transmitter is expected to fall silent altogether, and Pioneer 11, launched in April 1973, will become a "ghost ship" travelling through the Milky Way galaxy, passing near the star Lambda Aquila in almost 4 million years.
Its sister ship, Pioneer 10, heading in the opposite direction, continues to return scientific data and may have enough power to last until 1999. It was launched in March 1972, and, at almost 6 billion mi. (9.5 billion km), is the most distant object built by humans.
Another Pioneer, Pioneer 6, launched on December 16, 1965, into a solar orbit, was contacted by NASA/Ames in December. In the three decades gone by, the veteran \Jspacecraft\j has circled the sun more than 35 times and travelled a distance of 18 billion miles (29 billion km). It responded to the call from 31 million miles (50 million km) away, transmitting at its speed of 16 bits per second - torturously slow today, but considered lightning-fast in its days.
#
"Space Flight 1995: Summary",442,0,0,0
Space programs worldwide exhibited a number of significant developments even while continuing to adjust to severe financial restrictions, which had already become evident in 1994. As a consequence, developments increasingly tend to favor international consolidations and cooperative ventures between nations engaged in space scientifically, technologically or commercially. Thus, despite shifts in economic priorities and commensurate budgetary reductions, 1995 became one of the more active of recent years in future-oriented space flight developments.
Significant activities, for example, included the progress made in the development of the International Space Station (ISS) initiated last year, the bustling events around the operation of the Russian space station Mir, surprising discoveries in the universe by the Hubble Space \JTelescope\j continuing its cosmic research, important findings from automated deep-space explorers like Ulysses and the Jupiter-probe \JGalileo\j, and the growing attention by commercial firms to the potentials of space.
Commercial interest in worldwide space service markets has markedly increased this year. The advent of digital technology in telecommunications is a major force behind dozens of new advanced satellites being prepared to begin operations in orbit in the second half of this decade, with most of the geosynchronous orbit growth expected to come from outside the U.S., primarily Europe and Asia.
A new generation of commercial launchers for boosting them is lining up, introducing fierce competition not only in the many new voice, data and video network markets but also in a worldwide contest to fill launch vehicle payload manifests. According to one consulting firm, the global space services market may be expected to generate between $95-115 billion from 1995 through 2004.
For the United States' space shuttle and \JRussia\j's earth-circling platform, the year featured the launching of Phase I of the three-phased development process of the international megaproject ISS and, thereby, the end of decades of strict isolation and competition between both countries' space flight activities dating back to the early days of the Cold War.
Until actual construction begins in November 1997, in Phase I the shuttle Atlantis is to link up with the over-nine-year-old Mir in a total of seven missions for the purpose of conducting joint on-board research in the interest of risk mitigation, gathering practical experience in joint operations activities and testing assembly procedures for the two subsequent construction phases. Two of these link-ups, STS-71 and STS-74, already took place in 1995.
A total of nine crewed missions from the two major space-faring countries carried 48 humans into space, including ten women, bringing the total number of humans launched into space since 1958 to 667, incl. 58 women (counting repeaters), resp. 347 individuals (29 women).
#
"Year in Space 1996",443,0,0,0
\J1996: Space Launches\j
\JSpace Flight 1996: European Space Activities\j
\JSpace Flight 1996: Other Countries' Space Activities\j
\JSpace Flight 1996: Russian Space Activities\j
\JSpace Flight 1996: U.S. Commercial Space Activities \j
\JSpace Flight 1996: Department of Defense Space Activities\j
\JSpace Flight 1996: United States International Space Station Activities\j
\JSpace Flight (U.S.) 1996: Space Sciences and Astronomy\j
\JSpace Flight (U.S.) 1996: Hubble Space Telescope\j
\JSpace Flight (U.S.) 1996: Galileo\j
\JSpace Flight (U.S.) 1996: NEAR (Near Earth Asteroid Rendezvous)\j
\JSpace Flight (U.S.) 1996: POLAR\j
\JSpace Flight (U.S.) 1996: Compton Gamma Ray Observatory\j
\JSpace Flight (U.S.) 1996: Mars Exploration\j
\JSpace Flight (U.S.) 1996: Ulysses\j
\JSpace Flight 1996: Summary\j
#
"1996: Space Launches",444,0,0,0
(Refer to Table)
#
"Space Flight 1996: European Space Activities",445,0,0,0
Aerospace job losses in European countries continued abated, particularly in \JGermany\j where the leading company, Daimler Benz Aerospace AG (DASA), having already cut 26,000 jobs since 1990, in February announced additional reductions of 18,000 in the next three years. The German space sector itself comprises only 4300 engineers today, down from 7000 in 1990.
While Europe's aerospace industry, thus, generally remained in a state of crisis, the launch services market showed no decrease from previous years. With \JFrance\j's reliable Ariane 4 family of expendable launch vehicles (to date 92 flights with 7 failures = 92.4% reliability), the commercial operator Arianespace in 1996 racked up a total of 10 launches at the Kourou/French \JGuyana\j spaceport, carrying a variety of commercial satellites for customers such as Panamasat, Measat, Intelsat, M-Sat, Palapa, Arabsat, Turksat, Italsat, and others.
Europe's space activities, however, suffered a severe heavy setback when its new heavy lifter Ariane 5 was lost shortly after its first test launch, designated 501. As it turned out, the carrier rocket, using engines ten times more powerful as those of the successful Ariane 4 series, had been steered by guidance software designed for the Ariane 4 that was not adequately modified for the much larger and heavier Ariane 5.
Also destroyed in the explosion over \JKourou\j was the highly sophisticated payload CLUSTER, a set of four complex German-built satellites to study plasma streams in the solar wind and their interaction with the Earth's magnetosphere, a $150 million investment. The total loss of the 501 mission amounted to roughly $400 million.
In 1995, the European Space Agency (ESA) had extended its cooperative flight on \JRussia\j's Mir, EUROMIR 95, from the originally scheduled 135 days to nearly 180 days. On February 29, the German ESA-cosmonaut Thomas Reiter returned from the mission with Soyuz TM-22, accompanied by Yuri Gidzenko and Sergei Avdeev.
At 180 days, his flight became the longest mission flown by a West-European, and Reiter also was the first West-European to perform a spacewalk, followed by a second EVA.
In the science area, the Infrared Space Observatory (ISO), Europe's "Hubble Telescope", launched in 1995 by Ariane, in February reported the first infrared detection of cold molecular \Jhydrogen\j in deep space (the invisible "cold dark matter" astronomers have been seeking to complete astronomical models).
Other discoveries by ISO included the first galactic water vapor found in a deep space object,- in the dying star NGC 7027; hundreds of previously invisible galaxies; a galactic collision between the twin Antennae Galaxies, 60 million light-years from Earth; the birth of a star in the spiral arms and nucleus of the Whirlpool Galaxy, 20 million light-years away; and the death of a star, by ejection of layers of material, in NGC 6543, about 3000 light-years from Earth
#
"Space Flight 1996: Other Countries' Space Activities",446,0,0,0
On March 21, India successfully launched its new four-stage PSLV (polar satellite launch vehicle) carrying the 2033 -lb. (922-kg) 1-ton Indian Earth resources satellite IRS-P3. The flight was the second success in a row for the 145-ft. (44 m) booster, which develops more than 1 million lb. of thrust.
The Indian Space Research Organization (ISRO) satellite was placed into a 500- mi (800-km) sun-synchronous polar orbit, carrying a wide-field scanner and a German electro-optical scanner for ocean data, as well as a payload for X-ray \Jastronomy\j.
Earlier, India announced that the three cameras of its IRS-1C satellite launched on December 28, 1995, were successfully switched on and were functioning as expected. IRS-1C imagery is being marketed outside of India by the U.S. company Eosat. This brings the number of Indian-built Earth observing satellites in the last nine years to five, making India the operator of one of the world's most-active civilian remote-sensing programs.
In South America, \JChile\j, \JBrazil\j and Argentina continued to press ahead with efforts to expand their space launch and operations capability by the turn of the century. \JChile\j's plans to become the fourth space-faring nation in Latin America -- Argentina, \JBrazil\j and Mexico already operate satellites -- had suffered a setback on August 31, 1995, when its satellite Fasat-Alfa was lost after its launch on an Ukranian Tsiklon-3.
Argentina launched its Microsat on August 29 on a Russian Molniya- M rocket, which also carried the Czech Republic payload Magion 5, while Mexico's UNAMSat was carried into space on September 5 on a Russian Kosmos- 3M.
Other commercial international payloads were \JMalaysia\j's Measat 1 and 2, the country s first \Jtelevision\j satellites, on Ariane 44L launchers, \JIndonesia\j's Palapa C-1 and C-2 on a U.S. Atlas IIAS and an Ariane 44L, respectively, Canada's communications satellite M-Sat 1 on an Ariane 42P, \JItaly\j's \Jastronomy\j satellite BeppoSAX on an Atlas I, Israel's AMOS payload hitching a ride on an Ariane 44L, two communications satellites from Saudi \JArabia\j and one from Turkey.
#
"Space Flight 1996: Russian Space Activities",447,0,0,0
Despite continuing economic problems and a severe cash shortage for its space program, \JRussia\j in 1996 continued its space operations at a brisk pace, even if at a considerably reduced and less robust level than in the preceding years.
The new partnership with the U.S. continued to gain substance, and particularly the income from other countries availing themselves of (and paying for) the services of the space station Mir, including the United States use of it as Phase 1 of the ISS program (see International Space Station), brought much needed currency to keep \JRussia\j s space operations running. However, without increased funding in 1997, \JRussia\j will have great difficulty to maintain even a minimal space program.
Space station Mir. By the end of 1996, \JRussia\j s seventh space station Mir had been in operation for 3968 days. Its ten-year anniversary was observed on February 20, 1996. Until the end of 1996, it circled Earth approximately 62,094 times at altitudes of 238-245 mi. (380-393 km) in an orbit inclined 51.65 degrees to the equator.
Counting from its last brief period of nonoccupancy (September 1989) to the end of 1996, Mir has been inhabited continuously for 2674 days. Since its inception, it has been visited 28 times, incl. four times by a U.S. space shuttle, and by two- to three person Soyuz crews (shuttle visits: 5-7 persons). Between 1990 and the end of 1996, Mir has played host to a total of 12 (paying) guest cosmonauts from foreign countries, with more to come through 1999.
To resupply the occupants, the space station was visited in 1996 by three automated Progress M cargo ships, bringing the total of Progress and Progress M ships launched to Mir and the two preceding space stations, Salyut-7 and Salyut-6, to 76, with no failure. In 1996, Mir became increasingly dependent on the space shuttle for its logistics, due to severe setbacks in \JRussia\j s capability to finance its Progress resupply missions, particularly the Soyuz booster and its expensive Synthin kerosene fuel.
By the end of the year, the need for freight carried on board the shuttle, including critical water, air, and food supplies, had risen to about 40% of all logistics required by the space station, according to Mir officials.
In November, US-astronaut John Blaha on Mir completed the first Bioreactor experiment run to produce a three-dimensional cartilage and became the first human to harvest wheat grown in space, in an on-board "greenhouse" called Svet, built by \JSlovakia\j and sent to the station in 1990. According to space scientists, this represented the first time that an important agricultural crop and primary candidate for a future plant-based life support system successfully completed an entire life- cycle in the space environment.
Priroda. The 20-ton uncrewed Priroda module was launched on a Proton rocket on April 3. Docking to Mir occurred on the first attempt on 4/26 despite the failure of two of the module s batteries on 4/25. On 4/27, Priroda was rotated from an axial to a radial docking port. The module, built by the Khrunichev factory, carried a set of remote sensing instruments and some \Jmicrogravity\j experiments from \JRussia\j and the U.S. It also served as a cargo carrier for various equipment from these countries.
The addition of the module boosted the station s total mass (excluding transport vehicles) to about 242,500 lb. (110 metric tons) and its useful volume to nearly 14,000 cu.ft. (390 m3), finally exceeding the 12,500 cu.ft. (350 m3) available in NASA s 176,000 lb. (80-ton) Skylab in 1973/74.
Soyuz TM-23. The Mir-21 crew launched on February 21 and docked to the Mir station on 2/23, joining three crewmen of Mir-20 , who returned in TM-22 on 2/29. In the first month, onboard activities focused on national research, followed by the U.S. science program NASA-2 when U.S. \Jastronaut\j Shannon Lucid joined the crew on 3/23 (see STS-76).
On 5/21 and 24, the two Yuris (as cosmonauts Onufrienko and Usachev became known worldwide) performed their 2nd and 3rd spacewalk to install and unfurl the Mir Cooperative Solar Array brought up by Atlantis (STS-74) and a 4th on 5/30 to install a European earth resources camera on the Priroda module. On 6/13, they installed a truss structure called Rapana on Kvant-1, taking the place of the similar Strela. Later, they worked on the French research program Cassiopeia when Mir-22 arrived with French female guest \Jcosmonaut\j Andre-Deshays. Onufrienko and Usachev returned with her on 9/2.
Soyuz TM-24. The Mir-22 crew, including French/CNES researcher-cosmonaut Claudie Andre-Deshays, was launched at 9:18am EDT from Baikonur, after the primary crew, Gennadiy Manakov and Pavel Vinogradov, was grounded because of a cardiac irregularity of Manakov, and replaced by the back-up crew.
The three cosmonauts joined the Mir-21 crew on board Mir on 8/19. Prior to their arrival, on 8/1, the uncrewed cargo ship Progress M-31 was undocked and deorbited over the Pacific; Progress M-32 was launched on 7/31 and arrived on 8/2.
It was undocked on 8/18 to free up a docking port, and linked up again with Mir on 9/3. On 9/7, Shannon Lucid began setting new duration records for women in space when she passed the mark of 169 days set by Elena Kondakova in 1994/95 (Mir-17).
In other space-flight areas, \JRussia\j launched about 29 military, scientific, and telecommunications satellites on eight Proton launchers (incl. two failures), four Soyuz-U carriers (incl. two failures), four Kosmos-3M, three Molniya-M, one Zenit 2, one Tsiklon 3, and one Tsiklon 2 rockets.
In a severe setback involving more than 20 participating nations, \JRussia\j s attempt at launching the large, complex Mars exploration probe Mars-96 failed catastrophically shortly after its launch on November 16, when the fourth stage of its Proton launch vehicle malfunctioned. The 13,200 lb. (6 t) \Jspacecraft\j plunged back into Earth s atmosphere and reportedly impacted in the South Pacific Ocean near \JChile\j.
Commercial space activities. Still in its early stages of entering commercial space markets, \JRussia\j opened a new chapter in its cooperation with the U.S. when it launched, on April 8, a Proton rocket carrying the U.S.-built communications satellite Astra-1F into space.
The European satellite, part of an orbital network being developed by Luxembourg- based SES Co., was the first of 20 satellite launches planned by 2000 under a $1 billion agreement between \JRussia\j s Proton builder, Khrunichev and the joint venture International Launch Services (ILS), with Lockheed Martin of the U.S. and the Russian company RKK Energia, manager of Mir, among else.
RKK Energia initiated its own venture into becoming a satellite services provider by forming a joint venture, on February 9, with Gascom of Moscow and Loral Corp. of New York, with the objective to launch two 2976 lb. (1350 kg) Yamal satellites on a single Proton in 1997 for \JRussia\j s giant, soon-to-be- privatized natural-gas producer Gasprom.
On the second commercial launch for ILS, a Proton carried the communications satellite Inmarsat III F-2 into space. Operated by the International Maritime Satellite Organization (Inmarsat), the geostationary platform serves the Atlantic Ocean region, providing mobile communications, primarily for ships and \Jaircraft\j, at L-band frequencies. Inmarsat III F-1 satellite is operating over the Indian Ocean. (A second Inmarsat III, F3, was launched by a commercial Atlas IIA on December 18 from Cape Canaveral).
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"Space Flight 1996: U.S. Commercial Space Activities",448,0,0,0
A review of the commercial space sector shows that in the next five years more than a dozen aerospace companies, telecommunications firms and outsiders like Microsoft want to develop world-spanning or regional communication networks using satellites.
It would clearly be the most expensive space flight program since the Apollo Program of the Sixties, involving more than $25 billion in investments. With them, developers hope to realize the dream of unhindered mobile \Jtelephone\j and facsimile traffic as well as data transfer from any place on Earth to any other.
Based on the gigantic gap existing between \Jtelephone\j connections in the industrial countries (400 million sets for 800 million people) and the Third World (200 million sets for 5 billion people), investors see a services market of $100 billion in the next 10 years alone.
China alone, for example, with a population of 1.5 billion, at the beginning of 1996 had approximately 31 million telephones and 48 million switched access lines, primarily in highly populated areas. \JIndonesia\j has almost 58,000 villages with populations over 2000 that have no \Jtelephone\j service access.
Approximately 535,000 villages in India and 121,000 of the 151,000 villages in Africa have virtually no telephones. Mobile satellite systems would enable these countries to expand telecommunications quickly and economically into areas where extensive infrastructures do not now exist.
In the lead of mobile satellite network projects are Big LEO systems like Globalstar (Loral/QUALCOMM) with 56 satellites, Iridium (Motorola) with 66 satellites currently planned, Odyssey (TRW/Teleglobe) with 12 satellites, Aries/ECCO (Constellation Communications) with 46 satellites and ICO (Inmarsat), collectively representing more than $12 billion in investments. Particularly ambitious are the visions of Microsoft s Bill Gates with Teledesic, comprising up to 840 satellites. To date, licenses have been granted to Iridium, Odyssey, and Globalstar.
According to the annual review of the commercial launch market by the Arianespace launch consortium, the average mass of a communications satellite (excluding satellite constellations for mobile services), now about 5,000 lb. (2,400 kg), will increase to 7,100 lb. (3,200 kg) within four years, and then level off.
Not counting satellites of less than 2,200 lb. (1,000 kg), between 200 and 240 satellites will be launched in the next eight years, compared to 158 in the past eight years, and of these, 90% will be telecommunications satellites.
As a consequence of the boom in satellite companies, the launch services market is heating up, according to assessments by the U.S. Transportation Department which put the number of medium and heavy geosynchronous orbit (GEO) communications \Jspacecraft\j through 2010 at more than 460. This level has been estimated to amount to about $50-60 billion in commercially funded medium-to-heavy satellites and booster development around the world.
In 1996. commercial launch attempts of expendable space carriers totalled 22. There was only one failure: the sixth flight of Pegasus XL, a finned rocket launched from a Lockheed L-1011 Stargazer \Jaircraft\j as first stage, flying out of Vandenberg AFB.
In particular, McDonnell Douglas successfully launched ten Delta vehicles, followed by Lockheed Martin with six Atlas IIA launches plus one Atlas I, as well as four Titan 4 s for the DoD. Including the failed flight on November 4, Pegasus XL of Orbital Sciences Corp. (OSC) took to the air five times in 1996.
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"Space Flight 1996: Department of Defense Space Activities",449,0,0,0
U.S. military space organizations continued their post-Cold War "cultural change" move to make space a routine part of military operations across all service lines. With increasing commercialization of space and other nations' rapidly growing know-how and capabilities in noncommercial space systems, military space leaders are envisioning and concerned about an increasing potential of harmful uses of space by adversaries. For the Defense Department (DoD), this requires new military space resources to be integrated into the more traditional intelligence and defense doctrines in the most efficient and economic manner.
Military launches from Cape Canaveral and Vandenberg included nine intelligence and communications satellites on four Titan 4 launchers, three Navstar GPS-IIA satellites on Delta II, the Navy s UHF Follow-on F7 comsat on an Atlas II, and two technology development payloads, MSX and MSTI-3, on a Delta and a Pegasus air-launched carrier, respectively. Of these, MSX (Midcourse Space Experiment) tested infrared sensors for space- and ground- based missile defense systems, while MSTI-3 conducted similar tests for infrared, visible, and ultraviolet signatures of ballistic missile launches.
In its $1.4 billion program to reduce the operational cost of space transportation by one-half, the U.S. Air Force selected Lockheed Martin and McDonnell Douglas as finalists to build a new family of Evolved Expendable Launch Vehicles (EELV) replacing today s medium-lift Delta and Atlas boosters and heavy-lift Titans.
While Lockheed Martin s design builds on development of its advanced Atlas 2AR booster, powered by a Russian RD-180 engine from Pratt & Whitney and utilizing a new, storable-propellant upper stage, the Agena 2000, McDonnell Douglas proposed Delta 4 will use a new Rocketdyne cryogenic engine known as the RS-68, as well as a cryogenic upper stage being developed for the new Delta 3 ELV.
On December 3, the DoD announced the potential discovery of frozen water deep inside a crater at the south pole of the Moon. Located in the so-called Aitken Basin - twice the size of Puerto Rico and 7.5 mi. (12 km) deep, the super-cold ice formation was inferred by the radar signal returns of the small DoD Ballistic Missile Defense Organization s (BMDO) Clementine \Jspacecraft\j launched to the Moon in 1994 by the Pentagon and NASA. Scientists remained skeptical, and there was general agreement that much more study is needed to confirm the assertion.
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"Space Flight 1996: United States International Space Station Activities",450,0,0,0
In 1996, the International Space Station (ISS) program remained within budgetary guidelines and on schedule for first element launch (FEL) in November 1997. In spite of developmental challenges, cost and schedule constraints and the complexity associated with the program s multinational interrelationships (15 nations), the ISS made marked progress toward becoming a reality.
The development team of government and industry had to deal with a myriad of details in 1996: qualification and development testing on a major scale, completion of flight components, delivery of software releases, and bringing all this activity together to ensure the timely availability of components and modules to the Kennedy Space Center (KSC) for processing and \Jintegration\j with the Space Shuttle.
Phase 1. Phase 1 of the program, the joint U.S.-Russian effort to expand cooperation in human space flight, made significant progress with two successful Shuttle-Mir space missions, to add to earlier ones conducted in 1995. These link-ups (see Space shuttle) between Atlantis and Mir ferried supplies required on board the space station as well as U.S. astronauts Shannon Lucid and John Blaha as members of Russian station crews.
With Dr. Lucid s arrival aboard Mir in March 1996, the U.S. began a new period of permanent presence of U.S. astronauts in space that will continue through the operational duration of the ISS. Her 188 day mission set a new U.S. record for on-orbit duration and a new worldwide record for a woman in space.
Embodied within Phase 1 are a range of activities which provide the framework for U.S./Russian cooperation in space, and the mechanism for facilitating the \Jintegration\j of \JRussia\j into full partnership in the ISS.
Phase 1 also is already providing essential opportunities to test assumptions and validate models in the actual environment that the ISS will experience. In essence, with Phase 1 the space station era has already begun for the U.S., and the knowledge and experience gained has already demonstrated its value in the planning of training and operational procedures for ISS crews.
Phase 2/3. Design and fabrication of flight elements and software for the first six American assembly flights are essentially complete, with qualification testing under way. At the end of 1996, about half a year remained prior to shipment of the hardware to KSC in preparation for launch of the first U.S. assembly flight in December 1997.
Consisting of Node 1, two pressurized mating adapters and two stowage racks, the hardware has been structurally completed and is being outfitted and tested. Overall, NASA in 1996 passed the 45 percent completion mark of scheduled work, having built over 132,000 lb. (60 t) of hardware in the process.
The close of 1996 marked the third consecutive year in which the program performed within budgetary guidelines and schedule constraints, even if some concerns arose about future cost control, Russian- supplied elements, and a failure of Node 1 during pressure testing, requiring structural strengthening of the element. However, program-wide cost and schedule performance remained close to targets; where variances occurred, aggressive cost and schedule recovery plans were implemented.
Development programs continued in other ISS partner countries as well. Significant changes to the historic U.S./Russian cooperative plans were agreed to at the Joint Commission on Economic and Technological Cooperation (GCC) meeting between Russian Minister Chernomyrdin and Vice President Gore in July 1996.
Russia reaffirmed its willingness to meet its obligations with regard to critical early ISS elements, but will defer some subsequent elements until the year 2000 to relieve its launch burden. Three high-capacity Logistics Transfer Vehicles (LTV s) will take the place of the previously planned Progress resupply ships in 1998 and 1999.
Also, the Russian Solar Power Platform, critical for power and control functions at a later date, is now scheduled to be launched aboard the space shuttle. Thus, working together, a plan has emerged that maintains both agencies commitments without impacting the ISS program s $2.1 billion annual cost ceiling ( cap ) and its completion date.
In its ongoing research toward an advanced space launch vehicle to eventually replace the space shuttle and slash the high cost of space transportation (today: around $10,000 per pound mass to orbit) and reduce turnaround time, NASA in 1996 went from Phase I, a concept exploration effort to define both an operational Singe-Stage-to-Orbit (SSTO) reusable launch vehicle (RLV) and a salable, traceable SSTO demonstrator designated X-33, to Phase II (1996-99), a technology demonstration program involving design, fabrication and flight-test of the sub-scale (53% scale) X-33 and ground tests of full-scale RLV components.
Phase II is budgeted at $979 million, and target date for first flight of the remotely piloted, sub-orbital X-33, capable of altitudes up to 50 miles (80 km) and speeds of Mach 15, is March 1999. On July 2, NASA selected Lockheed Martin and their Skunk-Works- proposed VentureStar concept, eliminating the competing proposals of McDonnell Douglas Aerospace and Rockwell International. The full-scale version, expected to be developed with private-commercial funding, will go into production early in the next century.
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"Space Flight (U.S.) 1996: Space Sciences and Astronomy",452,0,0,0
Several automated and remotely-directed research missions added a staggering range of important and, in part, revolutionary discoveries to our growing body of knowledge about the universe. Science news were dominated by the announcements of discoveries of possible ancient life forms in two ancient life forms in two meteorites originating from the planet Mars.
Also, on the ground, NASA-funded research by an international team of scientists found evidence of prehistoric life in rock formations on Akilia Island in southern West \JGreenland\j that is at least 3.85 billion years ago, pushing back the conventional date of the emergence of life on Earth by 400 million years.
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"Space Flight (U.S.) 1996: Hubble Space Telescope",453,0,0,0
The performance of the Hubble Space \JTelescope\j (HST) in 1996 continued to exceed the original expectations of its builders. A wide range of new discoveries at distances reaching from billions of light-years in deep space to near-by planets of our Solar System surprised cosmologists and astronomers, again challenging some of their most accepted and treasured theories about the evolution, form, structure, processes, age, and future of the universe.
Stunning images received from the HST included the first direct image of a star other than the Sun when it took a picture in ultraviolet light of the red \Jsupergiant\j star \JBetelgeuse\j (Alpha Orionis) in the shoulder of the winter \Jconstellation\j Orion, showing an enormous hot spot at least 2,000 degrees K hotter than the surface of the star itself. Even though the apparent size of \JBetelgeuse\j is 20,000 times smaller than the width of the full Moon, the HST's ability to resolve it equates approximately to resolving a car's headlights at a distance of 6,000 miles (9,600 km).
Other highlights of the continuing exploration with HST included: Revelation of hundreds of previously undiscovered galaxies in various stages of evolution that probably date back to near the beginning of time. Astronomers were able to assemble a mosaic from 276 Hubble images covering an area just 1/30th the diameter of the full Moon,- the deepest and most detailed optical view ever taken of the universe.
HST took a picture of the planetary nebula NGC 7027, about 3000 light-years from Earth in the direction of the \Jconstellation\j \JCygnus\j, showing a star in its death throes - which last just a few thousand years even if the star lives up to 10 billion years. Later, it observed for the first time the surface of the small distant planet Pluto, discovered only 66 years ago, with a series of snapshots of nearly its entire surface as it rotated through a 6.4-day period, and showed it to be a complex object with more large-scale contrast than any planet except Earth.
The images revealed almost a dozen distinct \Jalbedo\j features, or provinces, never seen before, including a "ragged" northern polar cap bisected by a dark strip, a bright spot rotating with the planet, a cluster of dark spots, and a bright linear marking that is intriguing scientists.
Later, Hubble discovered thousands of gaseous tadpole-shaped fragments in the Helix Nebula, a ring of glowing gases blown off the surface of a sunlike star late in its life. Dubbed "cometary knots" because their glowing heads and gossamer tails superficially resemble comets, they are probably the result of a dying star's final outbursts.
When the \Jcomet\j Hyakutake passed by Earth just 9.3 million mi. (15 million km) away late in March, the super-telescope probed its inner region and the small icy nucleus at its center, estimated to measure only 1-3 km across. Discovered by a Japanese amateur astronomer on January 30, the \Jcomet\j could be seen clearly from Earth with a naked eye.
In 1996, Hubble pictures also revealed the mysterious heart of the Crab Nebula, the tattered remains of a stellar cataclysm witnessed more than 900 years ago, and produced numerous images over several months of observation which were assembled in a cosmic "movie" (567 KB) showing for the first time dramatic changes in the dynamic interior of the Nebula.
Hubble measured the diameters of a special class of pulsating stars called Mira variables which change size rhythmically, specifically of R Leonis and W Hydrae. The results suggest that these gigantic old stars, strangely enough, are not round at all but egg-shaped. Furthermore, it revealed galaxies under construction in the early universe, part of a long sought ancient population of "galactic building blocks".
The grouping of 18 gigantic star clusters appear to be the same distance from Earth and close enough to each other that they will eventually merge into a few galaxy-sized objects. At 11 billion light-years, they are so far away that they existed during the epoch when, as is commonly believed, galaxies started to form. This adds weight to a leading theory that galaxies grew by starting out as clumps of stars which through a complex series of encounters consolidated into larger assemblages seen by us today as fully- formed galaxies.
HST surveys of the "homes" of quasars, the universe's most energetic objects, revealed that quasars live in a remarkable variety of galaxies, many of which undergo violent collisions. Most of these mysterious objects lie at the cores of luminous galaxies, both spiral and elliptical in shape. According to current theory, the only credible possibility explaining how quasars can be so compact, variable and powerful are supermassive black holes in the process of swallowing up stars, gas and dust.
Closer to home, HST showed Texas-sized dust storms churning on Mars in September and October 1996 near the edge of the northern polar cap, probably a consequence of large temperature differences between the polar ice and the dark regions to the south, which are heated by the springtime Sun.
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"Space Flight (U.S.) 1996: Galileo",454,0,0,0
On December 7 of last year, after more than six years of travel, NASA s deep-space probe \JGalileo\j had inserted itself in an orbit around Jupiter. In early 1996, it completed transmitting the relayed data from its Probe which had entered the atmosphere of the belted gas-giant on 12/7/95, surviving entry speeds of over 106,000 mph (47 km/s), temperatures twice those on the surface of the Sun and deceleration forces up to 230 times the gravity on Earth.
During its 57-minute life, it sent measurements obtained on its descent back to \JGalileo\j more than 130,000 mi. (208,000 km) overhead for storage and transmission. The first condensed backup data were radioed back 2-4 times in December and January from the orbiter's solid-state memory. On 1/25, \JGalileo\j relayed the full data set from the tape recorder.
They revealed that the Probe had found the entry region of Jupiter to be much drier than anticipated (less water than even on the Sun), and had not detected the three-tiered cloud structure that most researchers had postulated. The amount of \Jhelium\j turned out to be about one-half of what was expected.
The Probe detected extremely strong winds and very intense turbulence, which indicates that the energy source driving much of Jupiter's distinctive circulation phenomena is probably heat escaping from the planet s deepest interior. The Probe also discovered an intense new radiation belt approximately 31,000 mi. (49,600 km) above Jupiter's cloud tops, and a virtual absence of \Jlightning\j (only about one-tenth as frequent as on Earth).
On March 14, \JGalileo\j used its main engine for the last time to rise the perijove (orbit low point) above the highly radioactive areas near Jupiter and sent the \Jspacecraft\j to its flybys with Jupiter's four largest moons, Ganymede, Io, Callisto and Europa. The German-built engine fired for 24 minutes, burning 408 lb. (185 kg) of propellants. On June 27 (2:29am EDT), \JGalileo\j passed just 519 mi. (835 km) above Ganymede's surface at a relative speed of 17,448 mph (7.8 km/s), providing the most detailed images and other information ever obtained about the icy satellite.
It revealed that the face of the huge satellite has been extensively bombarded by comets and \Jasteroids\j and dramatically wrinkled and torn by the same forces that make mountains and move continents on Earth, and that Ganymede, unlike our Moon, possesses its own magnetic field.
Also in June, the automated explorer started taking pictures of Io at a range of 1.4 million mi. (2.24 million km) showing new layers of sulfur and sulfur dioxide frost deposited from the \Jvolcano\j Masubi in Io's southern hemisphere since the Voyager flyby in 1979. \JGalileo\j also photographed a huge erupting geyser-like \Jvolcano\j on Io.
Truly sensational images from the moon Europa, in August, showed clear evidence of "warm ice" or even liquid water that may have existed or perhaps still exists today beneath Europa's cracked icy crust. In a statement released on August 13, NASA administrator Daniel Goldin said, "These fantastic new images of an icy moon of Jupiter are reminiscent of the ice- covered \JArctic\j Ocean on our planet.
The lack of craters, the cracks and signs of movement, all indicate that this might be young ice on a dynamic surface. It raises the possibility of a liquid ocean on Europa, the only other place in our solar system where we suspect such an ocean might exist." There are also theories that Europa has active subsurface volcanoes.
In a second pass by Ganymede on 9/6, the \Jspacecraft\j snapped three-dimensional pictures of giant, icy fissures from just 163 mi. (260 km) away, and on 11/4 (8:34am EST), for the first time, the \Jspacecraft\j flew close to the moon Callisto, passing within 686 mi. (1000 km) of the stark, crater- studded natural satellite.
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"Space Flight (U.S.) 1996: NEAR (Near Earth Asteroid Rendezvous)",455,0,0,0
With the launch of the NEAR (Near Earth Asteroid Rendezvous) \Jspacecraft\j on February 17, NASA began the world s first attempt to study an asteroid up close. The 1800 lb. (816 kg, half of it propellant) probe was also the first in an anticipated series of low-cost, quickly-built interplanetary excursion craft dubbed Discovery. NEAR is targeted for the 25- mi. (40-km) long asteroid 433 Eros, to spend a year in an orbit around the asteroid.
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"Space Flight (U.S.) 1996: POLAR",456,0,0,0
Launched on February 24 from Vandenberg AFB into a near-polar orbit around the Earth, POLAR is the last U.S. element of the International Solar Terrestrial Physics (ISTP) program. Among the scientific objectives of the mission are the measurement of the complete plasma, energetic particles and fields in the high-latitude polar region, the solar energy input, the characteristics of the auroral plasma acceleration outflow, etc., to better understand the Sun s myriad effects on Earth and its space environment. POLAR s sister probe, WIND, was launched November 1, 1994, toward the Lagrangian or libration point L1 between the Sun and Earth.
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"Space Flight (U.S.) 1996: Compton Gamma Ray Observatory",457,0,0,0
The Gamma Ray Observatory (GRO), named for American physicist (and Nobel Laureate) Arthur Holly Compton, marked its five year anniversary on April 5. The singular satellite, dedicated to observe gamma ray emissions across a broad spectrum of energies, has made a large number of significant discoveries in deep space after its launch in 1991 from the space shuttle Atlantis, among them, in December 1995, the sudden appearance of a never- before-seen type of object which bursts and pulses at the same time and is currently the brightest source of gamma rays and X-rays in the sky. All four instruments of the \Jspacecraft\j are still performing near their design specifications.
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"Space Flight (U.S.) 1996: Mars Exploration",458,0,0,0
In 1996, Mars suddenly leapt into the focus of increased public interest due to the media- sensationalized announcement of the recent discovery of possible life forms in its remote past. For the first time, we seem to have indications pointing, with some possibility, to the existence of primitive biota on Mars more than 3.6 billion years ago.
They were discovered deep inside a piece of Martian rock, designated ALH84001, which fell on Earth as a \Jmeteorite\j 13,000 years ago: organic molecules, next to traces of several mineral compounds associated with biological activity, as well as possible microscopic fossils of primitive organisms,- some ovoid, others tubular or worm-like and jointed, similar to but tiny in comparison to terrestrial microbes.
A similar discovery was reported in late October by British scientists at Open University, London, in a 17.4 lb. (7.9 kg) Mars \Jmeteorite\j, EETA79001, of considerable younger age: only 175 million years old, it had been blasted off the red planet about 600,000 years ago; thus, the organic traces found in it could point to much more recent and possibly even present-day life forms on Mars. Should these discoveries and their bold interpretation be confirmed, they it would doubtlessly be one of the most important scientific events of the outgoing century.
The discovery in ALH84001 and EETA79001 came too late to enable appropriate modifications to the research machines launched this year to Mars and those of the next synodic period two years hence. The modifications of later missions, however, for 2001 and thereafter, has already begun, and much is expected from them.
Mars Global Surveyor. On November 7, NASA launched Mars Global Observer (MGS) as the first U.S. \Jspacecraft\j bound for Mars since the ill-fated Mars Observer in 1992. It is also the first Surveyor-type Mars explorer of a series of such probes planned for the next decade, and with over a ton of mass, it is also the heaviest of the lot. MGS is scheduled to arrive at Mars on September 12, 1997, where it is to establish itself eventually in a polar, sun-synchronous orbit of 236 mi. (378 km) altitude and 92.9 degrees inclination to the equator. Its objectives include the complete mapping of the surface of the red planet with a sophisticated camera system.
Mars Pathfinder. Planned to be the first Mars lander since the highly successful Viking probes in 1976, NASA s Mars Pathfinder was launched and injected into a trans-Mars trajectory on December 4. Consisting of a lander and a remote-controlled six-wheeled minirover named Sojourner, the 1918 lb. (870 kg) \Jspacecraft\j is due to touch down in the \JAres\j Valley region of Mars on July 4, 1997, using braking rockets and inflated airbags to soften the impact. After its deployment, the little 24-lb. (11-kg) rover will explore the vicinity of the landing site, equipped with two black-and-white \Jtelevision\j cameras and an alpha-proton-x-ray-spectrometer for soil investigations.
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"Space Flight (U.S.) 1996: Ulysses",459,0,0,0
The NASA-ESA solar-polar explorer Ulysses continued its return-journey from its overflight of the Sun s north pole in June 1995, heading out to the orbit of Jupiter. Its measurements over the solar south pole in November 1994, reported in 1996, found a surprisingly small increase in the amount of the helium-3, a lighter isotope of the gas \Jhelium\j, since the formation of the solar system; this allows a more precise estimate of the amount of dark matter in the universe. The exact nature of this invisible matter remains one of the most intriguing mysteries of \Jastronomy\j, and at present as much as 90 percent of the universe may be composed of it, according to best estimates. The findings are not at all consistent with current models of the formation and abundance of matter in the universe.
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"Space Flight 1996: Summary",460,0,0,0
For space flight, 1996 proved to be a year of strong contrasts. While the United States space budget remained relatively stable, international space flight activities were again characterized by trends, already recognizable in 1995 and 1994, of incisive reductions in public spending for space programs and, consequently, increasing pressure on governments to shift responsibilities in the space flight area to the private sector, particularly for services and operations, as well as for new industrial commercialization initiatives, such as - in the United States - for the development of a new, lower-cost reusable launch vehicle to eventually replace the space shuttle.
Despite these difficult shifts in economic priorities, 1996 again was one of the most active years for sustaining operations and advanced developments in space, mainly in the United States with the progressing development of the International Space Station (ISS) as well as in \JRussia\j with the continuing utilization of the space station Mir and, after more than ten years of operation, its increasingly burdensome demands for repair and maintenance as well as its growing dependence on the shuttle for resupply flights to continue operations.
1996 also brought surprising new discoveries by the Hubble Space \JTelescope\j (HST), the European Infrared Space Obervatory (ISO) and the deep-space probe \JGalileo\j at Jupiter, as well as the beginning of NASA's intensive robotic exploration program of the planet Mars after a twenty-year hiatus.
Commercial companies engaged in providing space services with a variety of expendable carrier rockets again this year showed themselves firmly in the saddle, as the commercial satellite launch market boomed, boosting insurance profits and attracting new capital to the business. However some countries, particularly Europe, \JRussia\j and China, encountered painful setbacks in their commercial space activities, with considerable losses in resources and even lives.
The U.S. space shuttle and \JRussia\j s earth- orbiting Mir added another two rendezvous-and-docking flights to the two joint missions of 1995, bringing the total flights of Phase 1 of the three-phased ISS development program to four, with five more to come in 1997 and 1998. Purpose of these missions is to conduct joint on-board research preparatory to ISS operations, to gather practical experience in joint on-orbit activities in the interest of risk mitigation, and to test assembly procedures for the imminent ISS construction phase.
As in the preceding year, a total of nine crewed missions from the two major space-faring nations carried 48 humans into space, including four women, bringing the total number of people (counting repeaters) launched into space since 1958 to 715, incl. 63 women, respectively 362 individuals (30 female).
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"Space News Update 1997",461,0,0,0
\JHubble Witnesses the Final Blaze of Glory of Sun-like Stars (17 December, 1997)\j
\JClosest Europa Flyby Marks Start of Galileo Mission 'Part II' (16 December, 1997)\j
\JHubble Captures Brief Moment in Life of Lively Duo (4 December, 1997)\j
\JFirst Observation of Space-Time Distortion by Black Holes (6 November, 1997)\j
\JGalileo Finds Arizona-Sized Volcanic Deposit on Io (5 November, 1997)\j
\JMars Pathfinder Winds Down After Phenomenal Mission (4 November, 1997)\j
\JHubble Catches Up With A Blue Straggler Star (29 October, 1997)\j
"Hubble Witnesses the Final Blaze of Glory of Sun-like Stars (17 December, 1997)",462,0,0,0
The end of a sun-like star's life was once thought to be simple: the star gracefully casting off a shell of glowing gas and then settling into a long retirement as a burned-out white dwarf.
Now, a dazzling collection of detailed views released today by several teams of astronomers using NASA's Hubble Space Telescope reveals surprisingly intricate glowing patterns spun into space by aging stars: pinwheels, lawn sprinkler style jets, elegant goblet shapes, and even some that look like a rocket engine's exhaust.
"These eerie fireworks offer a preview of the final stage of our own Sun's life," says Bruce Balick of the University of Washington, Seattle. More than simply a stellar "light-show," these outbursts provide a way for heavier elements --predominantly carbon -- cooked in the star's core, to be ejected into interstellar space as raw material for successive generations of stars, planets and, potentially, life.
The astronomers say the incandescent sculptures are forcing a re-thinking of stellar evolution. In particular, the patterns may be woven by an aging star's interaction with unseen companions: planets, brown dwarfs, or smaller stars.
"The first time we looked at the Hubble's breathtaking pictures, we knew that our older and simpler ideas of how these objects are formed had to be overhauled," said Howard Bond of the Space Telescope Science Institute (STScI), Baltimore MD. "The basic question is: How do these nebulae shape themselves?"
"Hubble's colorful views are a feast for the eyes," said Mario Livio, also of the STScI. "Their beauty is matched only by the mystery."
Surprising new details revealed by the Hubble pictures include:
\B╖\b Unexplained disks and "donuts" of dust girdling a star, which pinch outflowing gas. These may be linked to the presence of invisible companions.
\B╖\b Remarkably sharp, inner bubbles of glowing gas -- like a balloon inside a balloon -- blown out by the violently outflowing gases called a "fast wind" (1000 miles/sec) ejected during the final stages of a star's death.
\B╖\b Strange, glowing "red blobs" placed along the edge of some nebulae may be chunks of older gas caught in the stellar gale of hot flowing material from the dying star.
\B╖\b Jets of high-speed particles that shoot out in opposite directions from a star, and plow through surrounding gas, like a garden hose stream hitting a sand pile.
\B╖\b Pinwheel patterns formed by symmetrical ejection of material so that intricate structures are mirrored on the opposite side of a star.
"We're still reveling in the quality of the data and the wealth of new details. In the longer term, we're going to have to confront these strikingly symmetric structures with some fundamentally revised ideas about the final stages of a star's life," said Balick. "The lovely patterns of gas argue that some highly ordered and powerful process orchestrates the ways stars lose their mass, completely unlike an explosion."
A long-standing puzzle is how these nebulae acquire their complex shapes and symmetries. The red giant stars that preceded their formation should have ejected simple, spherical shells of gas. "Hubble's ability to see very fine structural details -- usually blurred beyond recognition in ground-based images -- enables us to look for clues to this puzzle," said Balick.
Several teams of astronomers will be observing planetary nebulae using new infrared instruments installed on the Hubble Telescope last February. This way, astronomers can glimpse the ejection of material at a very early stage long before the expelled nebula starts to become visible optically.
Given Hubble's high resolution, astronomers also hope to revisit the same nebula in a few years to actually see how the shell has further expanded into space. Their observations will be compared to predictions and either refine or dismiss current ideas on the mass ejection mechanisms of dying stars.
"These nebulae observed by Hubble give us a preview of our own Sun's fate. Some five billion years from now, after the Sun has become a red giant and burned the Earth to a cinder, it will eject its own beautiful nebula and then fade away as a white dwarf star," warned Bond.
The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc. (AURA) for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency (ESA).
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"Closest Europa Flyby Marks Start of Galileo Mission 'Part II' (16 December, 1997)",463,0,0,0
NASA's Galileo spacecraft today successfully made its closest-ever flyby of Jupiter's icy moon Europa, marking the start of an extended mission that will focus on new and tantalizing scientific questions raised by its just-completed, highly successful two-year primary mission.
"Galileo has earned a place in history as the first mission to orbit an outer planet," said Dr. Wesley T. Huntress, Jr., NASA's associate administrator for space science, Washington, DC. "Galileo already has returned a wealth of new information in its two-year scientific exploration of Jupiter's atmosphere and system of moons. But the best yet may still be ahead of us as Galileo continues its mission at Jupiter with a focus on the moons Europa and Io in the next two years."
Galileo dipped over Europa at an altitude of only 124 miles (200 kilometers), with the signal received on Earth at 7:49 a.m. EST. This was the first encounter of the Galileo Europa mission, which began formally on Dec. 8, following the end of Galileo's primary mission. The Galileo Europa mission will study Jupiter's icy satellite in detail in hopes of shedding more light on the intriguing prospect that liquid oceans may lie under Europa's ice crust.
New images released today from Galileo's Europa encounter of Nov. 6 show more evidence that the moon has been subjected to intense geological deformation. The pictures show a mottled region of dark and splotchy terrain that scientists say represents some of the most recent geologic activity on Europa.
It is believed the mottled appearance was created when chaotic areas of the bright, icy crust broke apart and exposed darker material underneath. The new images also show a smooth, gray band where the Europan crust has been fractured, separated, and filled in with material from the interior. Numerous isolated mountains or "massifs" are visible.
The new images represent a small portion of the 1,800 images obtained during Galileo's primary mission, including hundreds of high-resolution images of Jupiter's moons. The images and other information gathered by Galileo's science instruments have dramatically revised our knowledge of Jupiter and its moons, according to mission scientists.
The Galileo Europa mission is designed to follow up on these discoveries and will include eight consecutive Europa flybys through February 1999, followed by four Callisto flybys and one or two Io encounters in late 1999, provided the spacecraft remains healthy.
"The Galileo Europa mission really builds upon the success of the prime mission which has forced us to re-think many of our perceptions of the Jovian system," said Galileo project scientist Dr. Torrence Johnson. "We've acquired a tremendous pool of knowledge about Jupiter, its magnetosphere and its four largest moons."
The key findings of Galileo's primary mission include:
\B╖\b The existence of a magnetic field on Jupiter's largest moon, Ganymede;
\B╖\b The discovery of volcanic ice flows and melting or "rafting" of ice on the surface that supports the premise of liquid oceans underneath at some point in Europa's history;
\B╖\b The observation of water vapor, lightning and aurora on Jupiter;
\B╖\b The discovery of an atmosphere of hydrogen and carbon dioxide on the moon Callisto;
\B╖\b The presence of metallic cores in Europa, Io and Ganymede and the lack of evidence of such a core in Callisto;
Evidence of very hot volcanic activity on Io and observations of dramatic changes compared to previous observations and even during the period of Galileo's observations.
"We look forward to providing even more fascinating science results over the next two years," said newly appointed Galileo Europa mission project manager Bob Mitchell of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
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"Hubble Captures Brief Moment in Life of Lively Duo (4 December, 1997)",464,0,0,0
Some stars in double-star systems have found a quick way to lose weight by dumping their extra pounds onto their companions. Astronomers using NASA's Hubble Space Telescope have discovered such a case in the double-star system Phi Persei.
A "rapid diet" program has trimmed an aging, once massive star to a lean one-solar mass, while the once mild-mannered, moderate-sized companion has bulked up to a hefty nine-solar masses and is spinning so violently that it's flinging gas from its surface. This observation has allowed astronomers to catch a glimpse of an unusual, fleeting moment in the life of a massive star in a double-star system.
"We have seen massive star binaries at the end of their lives, when one star has collapsed to become a neutron star," says Douglas Gies of the Center for High Angular Resolution Astronomy at Georgia State University, Atlanta.
"There are about six dozen of those known in our galaxy. But what we hadn't seen before is the phase just prior to the collapse of the aging star. The Hubble observations dramatically show how severely a star's material can be removed through the gravitational influence of a nearby companion."
What Gies and his team have seen is an opportunistic star that has taken advantage of an aging, ailing partner. After consuming most of its hydrogen -- the fuel that keeps its thermonuclear furnace running -- the aging star swelled up and began jettisoning its mass until only its bare core was left.
The companion star cannibalized the discarded material, thereby increasing in size. Gies calls the stripped-down star a subdwarf, a type of aging star that has passed the expansion phase -- by swelling and puffing away its outer layers -- and is on its way to becoming a fading white dwarf. Yet this aging, stripped-down star, which has the same mass as the Sun, is nine times hotter than the Sun at 95,000 degrees Fahrenheit and is very bright.
The subdwarf would be the brightest object of its class in the sky (a sixth-magnitude star) if it could be seen alone. If placed at the Sun's distance, it would appear 200 times brighter than the Sun. However, the beefed-up companion is ten times brighter in visible light than the subdwarf, which is lost in its glare and eluded detection for many years.
The subdwarf was detected by the Hubble telescope's Goddard High Resolution Spectrograph (which was removed from Hubble last February during the second servicing mission), allowing scientists to identify its spectral signature.
Gies' analysis of these observations will appear in the Jan. 20, 1998, issue of the Astrophysical Journal. With the Hubble data, he pieced together a better picture of life in this double-star system, especially how the beefed-up companion gained its extra mass.
By puffing away most of its mass, the aging, stripped-down star has given new life and a new identity to its companion. The roughly 10-million-year-old companion has potentially doubled its lifetime because it has gained a vast amount of hydrogen fuel, which is needed to maintain its thermonuclear furnace.
By beefing up, the companion also has changed its identity from a normal, moderately massive star to a "Be" star, a type of hot star with a broad, flattened disk of hydrogen gas swirling around it, much like the rings of Saturn. Based on measurements taken by astronomers at the U.S. Naval Observatory, the disk is eight times wider than the star.
The disk formed from gas spun off the rapidly rotating "Be" star. What causes the fast spin of "Be" stars has been a mystery to astronomers. Now the Hubble telescope observations of Phi Persei offer at least a partial explanation: The gas discarded from a nearby swelling star strikes the companion off-center, causing it to spin faster.
"The companion is now rotating so fast (one million mph or 450 kilometers per second at its equator) that the star is distorted into a flattened oblate figure in which gravity can barely maintain its hold on the star's outer layers," Gies explains.
This new information about the stripped-down star and its companion leads scientists to speculate about their past. Before the exchange of material, the stripped-down subdwarf was the more massive of the two, about six times more massive than the Sun. Its companion was slightly less bulky, about five solar masses. Such massive stars usually race through life at a faster pace than most stellar objects, ending their lives in one big supernova explosion. However, stars in binary systems live differently.
When the once massive subdwarf entered its twilight years about one million years ago, it swelled in size as it began using up its hydrogen fuel. A single massive star would have eventually exploded, but the presence of the companion prevented the once-massive star from suffering such a violent fate. Instead, the once-massive star dumped most of its outer layers onto its companion, and now may be heading to a quiet demise.
A curious destiny may await this pair. As long as the "Be" star doesn't break apart, it will live for another 10 million years because of the hydrogen fuel it acquired from its companion. Then it will swell during the expansion phase and possibly dump some of its mass back onto the subdwarf, which will have evolved into a white dwarf. The subdwarf then might grow in mass and eventually explode as a supernova. Or, the companion might swell up so much that it would engulf the white dwarf, eventually tossing out its material in mix-master action.
"The beefed-up 'Be' star has won a new lease on life, but its ultimate fate will be determined by the corpse of its former companion, which remains in orbit uncomfortably nearby," Gies concludes. Phi Persei is 720 light-years away in the constellation Perseus, visible in the autumn evening sky in the northern hemisphere, just north of the Andromeda Galaxy, M31. The double-star system is visible as a fourth-magnitude star.
The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc. (AURA) for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency (ESA).
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"First Observation of Space-Time Distortion by Black Holes (6 November, 1997)",465,0,0,0
Astronomers using NASA's Rossi X-ray Timing Explorer (RXTE) spacecraft reported today that they have observed a black hole that is literally dragging space and time around itself as it rotates. This bizarre effect, called "frame dragging," is the first evidence to support a prediction made in 1918 using Einstein's theory of relativity.
The phenomenon is distorting the orbit of hot, X-ray emitting gas near the black hole, causing the X-rays to peak at periods that match the frame-dragging predictions of general relativity.
The research team, led by Dr. Wei Cui of the Massachusetts Institute of Technology, is announcing its results in a press conference today during the American Astronomical Society's High Energy Astrophysics Division (HEAD) meeting in Estes Park, CO. Collaborators in the research include Dr. Wan Chen of NASA's Goddard Space Flight Center, Greenbelt, MD, and Dr. Shuang N. Zhang of NASA's Marshall Space Flight Center, Huntsville, AL.
"If our interpretation is correct, it could demonstrate the presence of frame dragging near spinning black holes," said Cui. "This observation is unique because Einstein's theory has never been tested in this way before."
Black holes are very massive objects with gravitational fields so intense that near them, nothing, not even light, can escape their pull. This effect shrouds the hole in darkness, and its presence can only be inferred from its effects on nearby matter. Many of the known or suspected black holes are orbiting a close "companion" star.
The black hole's gravity pulls matter from the companion star, which forms a disk around the black hole as it is drawn inward by the black hole's gravity, much like soap suds swirling around a bathtub drain. Gas in this disk gets compressed and heated and emits radiation of various kinds, especially X-rays.
The research team used these X-ray emissions to determine if frame dragging was present. The team found that the X-ray emissions were varying in intensity. By analyzing this variation, they found a pattern, or repetition, that was best explained by a perturbation in the matter's orbit.
This perturbation, called a precession, occurs when the orbit itself shifts around the black hole. This is evidence for frame dragging because as the matter orbits the black hole, the space-time that is being dragged around the black hole drags the matter along with it. This shifts the matter's orbit with each revolution.
Einstein's Theory of General Relativity has been highly successful at explaining how matter and light behaves in strong gravitational fields, and has been successfully tested using a wide variety of astrophysical observations.
The frame-dragging effect was first predicted using general relativity by Austrian physicists Joseph Lense and Hans Thirring in 1918. Known as the Lense-Thirring effect, it has not been definitively observed thus far, so scientists will scrutinize the new reports very carefully.
The possible detection of frame dragging around another type of very dense, quickly spinning objects, called neutron stars, was accomplished very recently by Italian astronomers, whose work led Dr. Cui's team to seek the effect near black holes.
The Italians, Drs. Luigi Stella of the Astronomical Observatory of Rome, and Mario Vietri of the Third University of Rome, will report their findings at the November 6 conference in Estes Park. These observations also were made using the RXTE, which is available for use by astronomers throughout the world.
"This is exciting work that needs further confirmation, as for any seemingly major advance in science," said Dr. Alan Bunner, Director of the Structure and Evolution of the Universe Program at NASA Headquarters, Washington, DC.
The RXTE spacecraft is a 6,700 pound observatory placed into orbit by NASA in December 1995. Its mission is to make astronomical observations from high-energy light in the X-ray range, which is emitted by powerful events in the universe. These events are often associated with massive, compact objects such as black holes and neutron stars.
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"Galileo Finds Arizona-Sized Volcanic Deposit on Io (5 November, 1997)",466,0,0,0
Observations taken by NASA's Galileo spacecraft five months apart reveal a new dark spot the size of Arizona on Jupiter's moon Io, indicating that dramatic volcanic activity occurred during that time.
"This is the largest surface change on Io observed by Galileo during its entire two-year tour of the Jovian system," said Galileo imaging team member Dr. Alfred McEwen, a research scientist at the University of Arizona in Tucson.
The visible change took place during the five months between Galileo's seventh and tenth orbits of Jupiter. The change is manifested as a dark spot about 249 miles in diameter, surrounding a volcanic center named Pillan Patera, which is named after the South American god of thunder, fire and volcanoes. Dark features at the center of the deposits may be new lava flows.
These changes appear in images taken by the Solid State Imaging system aboard Galileo, with marked differences between the pictures taken on April 4, 1997 and September 19, 1997. In June of 1997 an active plume was observed over Pillan by Galileo and the Hubble Space Telescope with a height of 75 miles, and both Galileo and ground-based astronomers observed an intense hot spot.
"Most of the volcanic plume deposits on Io show up as white, yellow or red due to sulfur compounds. However, this new deposit is gray, which tells us it has a different composition, possibly richer in silicates than the other regions," McEwen explained. "While scientists knew that silicate volcanism existed on Io from high temperatures, this may provide clues as to the composition of the silicates, which in turn tells us about Io's evolution."
"Io is probably primarily composed of silicates, which is the type of volcanic rock found on Earth, " McEwen added, "but the extreme volcanism of Io may have led to the creation of silicate compositions that are unusual on Earth."
The Io images showing the changes in Pillan Patera also reveal alterations in the plume deposit of Pele, the large red oval southwest of Pillan, which may indicate that both plumes were active at the same time and interacted with one another. A dark region southwest of Pele, which appears similar to the Pillan deposits, has been present since the Voyager flybys in 1979.
Io is the most volcanically active body in the Solar System. Scientists hope to learn more about the fiery satellite when Galileo continues its studies over the next two years, during a mission extension known as the Galileo Europa Mission.
The extended mission will include eight additional encounters of Europa, four of Callisto, and two close Io flybys in late 1999, depending on spacecraft health. Galileo will pass very close to Pillan Patera in the first of the two Io flybys, so high-resolution images can be acquired over a small portion of this area.
Galileo was launched in 1989 and entered orbit around Jupiter on Dec. 7, 1995. The final satellite encounter of its two-year primary mission will occur on Thursday, Nov. 6, 1997 at 3:32 p.m. EST, when the spacecraft swoops over Europa at an altitude of 1,269 miles.
"The Galileo Orbiter is performing flawlessly and all 11 of its sophisticated science instruments and the radio science investigations are still providing excellent data," said Galileo Project Manager Bill O'Neil of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
"A great bounty of Jupiter system science has been obtained and the continuing study of these data will surely add many important discoveries. While not all of the original objectives could be met due to the antenna failure, I believe that the overall science return from Galileo will easily exceed what was envisioned at project inception 20 years ago, because our team of scientists and engineers has done such a superb job of capturing the most important observations."
The Galileo mission is managed by JPL for NASA's Office of Space Science, Washington, DC. JPL is an operating division of California Institute of Technology, Pasadena, CA.
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"Mars Pathfinder Winds Down After Phenomenal Mission (4 November, 1997)",467,0,0,0
After operating on the surface of Mars three times longer than expected and returning a tremendous amount of new information about the red planet, NASA's Mars Pathfinder mission is winding down.
Flight operators at NASA's Jet Propulsion Laboratory, Pasadena, CA, made the announcement today after attempting to re-establish communications with the spacecraft over the last month. With depletion of the spacecraft's main battery and no success in contacting Mars Pathfinder via its main or secondary transmitters, the flight team cannot command the spacecraft or the small rover named Sojourner that had been roving about the landing site and studying rocks.
"We concede that the likelihood of hearing from the spacecraft again diminishes with each day," said Pathfinder Project Manager Brian Muirhead. "We will scale back our efforts to re-establish contact but not give up entirely.
"Given that, and the fact that Pathfinder is the first of several missions to Mars, we'll say 'see you later' instead of saying goodbye," he said.
At the time the last telemetry from the spacecraft was received, Pathfinder's lander had operated nearly three times its design lifetime of 30 days, and the Sojourner rover operated 12 times its design lifetime of seven days.
"I want to thank the many talented men and women at NASA for making the mission such a phenomenal success. It embodies the spirit of NASA, and serves as a model for future missions that are faster, better, and cheaper.
Today, NASA's Pathfinder team should take a bow, because America is giving them a standing ovation for a stellar performance," said NASA Administrator Daniel S. Goldin.
Since its landing on July 4, 1997, Mars Pathfinder has returned 2.6 billion bits of information, including more than 16,000 images from the lander and 550 images from the rover, as well as more than 15 chemical analyses of rocks and extensive data on winds and other weather factors.
The only remaining objective was to complete the high-resolution 360-degree image of the landing site called the "Super Pan," of which 83 percent has already been received and is being processed. The last successful data transmission cycle from Pathfinder was completed at 3:23 a.m. Pacific Daylight Time on Sept. 27, which was Sol 83 of the mission.
"This mission has advanced our knowledge of Mars tremendously and will surely be a beacon of success for upcoming missions to the red planet," added Dr. David Baltimore, president of the California Institute of Technology, which manages JPL for NASA. "Done quickly and within a very limited budget, Pathfinder sets a standard for 21st century space exploration."
The Mars Pathfinder team first began having communications problems with the spacecraft on Saturday, Sept. 27. After three days of attempting to re-establish contact, they were able to lock on to a carrier signal from the spacecraft's auxiliary transmitter on Oct. 1, which meant that the spacecraft was still operational.
They locked on to the same carrier signal again on Oct. 6, but were not able to acquire data on the condition of the lander. At that time, the team surmised that the intermittent communications were most likely related to depletion of the spacecraft's battery and a drop in the spacecraft's operating temperatures due to the loss of the battery, which kept the lander functioning at warmer temperatures.
Over the last month the operations team has been working through all credible problem scenarios and taking a variety of actions in attempting to recover the link with Pathfinder. With all of the most plausible possibilities exhausted, the team plans to continue sending commands and listening for a spacecraft signal on a less frequent basis.
"Basically we are shifting to a contingency strategy of sending commands to the lander only periodically, perhaps once a week or once per month," said Mission Manager Richard Cook. "Normal mission operations are over, but there is still a small chance of restabilising a link, so we'll keep trying at a very low level."
Although the true cause of the loss of lander communications may never be known, recent events are consistent with predictions made at the beginning of the extended mission in early August, Muirhead said.
When asked about the life expectancy of the lander, project team members predicted that the first thing that would fail on the lander would be the battery; this apparently happened after the last successful transmission September 27.
After that, the lander was expected to begin getting colder at night and go through much deeper day-night thermal cycles. Eventually, the cold or the cycling would probably render the lander inoperable.
According to Muirhead, it appears that this sequence of events has probably taken place. The health and status of the rover is also unknown, but since initiating its onboard backup operations plan a month ago, the rover is probably circling the vicinity of the lander, attempting to communicate with it.
The rover, which went into a contingency mode on Oct. 6, or Sol 92 of the mission, had completed an alpha proton X-ray spectrometer study of a rock nicknamed Chimp, to the left of the Rock Garden, when it was last heard from.
The rover team had planned to send the rover on its longest journey yet -- a 165-foot (50-meter) clockwise stroll around the lander -- to perform a series of technology experiments and hazard avoidance exercises when the communications outage occurred. That excursion was never initiated once the rover's contingency software began operating.
Now known as the Sagan Memorial Station, the Mars Pathfinder lander was designed primarily to demonstrate a low-cost way of delivering a set of science instruments and a free-ranging rover to the surface of the red planet. Landers and rovers of the future will share the heritage of spacecraft designs and technologies first tested in this "pathfinding" mission.
Part of NASA's Discovery program of low-cost planetary missions, the spacecraft used an innovative method of directly entering the Martian atmosphere. Assisted by a 36-foot-diameter (11-meter) parachute, the spacecraft descended to the surface of Mars on July 4 and landed, using airbags to cushion the impact. The spacecraft's novel entry was successful.
Scientific highlights of the Mars Pathfinder mission are: Martian dust includes magnetic, composite particles, with a mean size of one micron.
Rock chemistry at the landing site may be different from Martian meteorites found on Earth, and could be of basaltic andesite composition.
The soil chemistry of Ares Vallis appears to be similar to that of the Viking 1 and 2 landing sites.
The observed atmospheric clarity is higher than was expected from Earth-based microwave measurements and Hubble Space Telescope observations.
Dust is confirmed as the dominant absorber of solar radiation in Mars' atmosphere, which has important consequences for the transport of energy in the atmosphere and its circulation. Frequent "dust devils" were found with an unmistakable temperature, wind and pressure signature, and morning turbulence; at least one may have contained dust (on Sol 62), suggesting that these gusts are a mechanism for mixing dust into the atmosphere.
Evidence of wind abrasion of rocks and dune-shaped deposits was found, indicating the presence of sand.
Morning atmospheric obscurations are due to clouds, not ground fog; Viking could not distinguish between these two possibilities.
The weather was similar to the weather encountered by Viking 1; there were rapid pressure and temperature variations, downslope winds at night and light winds in general. Temperatures were about 10 degrees warmer than those measured by Viking 1.
Diversity of albedos, or variations in the brightness of the Martian surface, was similar to other observations, but there was no evidence for the types of crystalline hematite or pyroxene absorption features detected in other locations on Mars.
The atmospheric experiment package recorded a temperature profile different than expected from microwave measurements and Hubble observations.
Rock size distribution was consistent with a flood-related deposit.
The moment of inertia of Mars was refined to a corresponding core radius of between 807 miles and 1,242 miles (1,300 and 2,000 kilometers).
The possible identification of rounded pebbles and cobbles on the ground, and sockets and pebbles in some rocks, suggests conglomerates that formed in running water, during a warmer past in which liquid water was stable.
Engineering milestones of the mission included demonstrating a new way of delivering a spacecraft to the surface of Mars by way of direct entry into the Martian atmosphere. In addition, Mars Pathfinder demonstrated for the first time the ability of engineers to deliver a semi-autonomous roving vehicle capable of conducting science experiments to the surface of another planet.
The Mars Pathfinder mission is managed by the Jet Propulsion Laboratory for NASA's Office of Space Science, Washington, DC. The mission is the second in the Discovery program of fast track, low-cost spacecraft with highly focused science goals. JPL is managed by the California Institute of Technology, Pasadena, CA.
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"Hubble Catches Up With A Blue Straggler Star (29 October, 1997)",468,0,0,0
Astronomers have long been mystified by observations of a few hot, bright, apparently young stars residing in well-established neighborhoods where most of their neighbors are much older.
With the help of NASA's Hubble Space Telescope, astronomers now have evidence that may eventually help solve the 45-year-old mystery of how these enigmatic stars, called blue stragglers, were formed. For the first time, astronomers have confirmed that a blue straggler in the core of a globular cluster (a very dense community of stars) is a massive, rapidly rotating star that is spinning 75 times faster than the Sun. This finding provides proof that blue stragglers are created by collisions or other close encounters in an overcrowded cluster core.
Astronomers studied a blue straggler in the tumultuous heart of the nearby globular cluster 47 Tucanae (47 Tuc), located 15,000 light-years away in the southern constellation Tucana. The observation was made by astronomers Michael M. Shara of the Space Telescope Science Institute, Baltimore, MD; Rex A. Saffer of Villanova University, Villanova, PA; and Mario Livio, also of the Institute. Their analysis will appear in the Nov. 1 issue of the Astrophysical Journal Letters.
"This is an extremely exciting result," Saffer said, "because it may help distinguish between competing theories of blue straggler star formation and evolution."
"Since their discovery 45 years ago, blue stragglers have been assumed to be stars much like the Sun, although their bluer color and greater brightness imply that they are more massive and much younger than normal globular cluster stars. Our analysis confirms that, but without having to make any assumptions about the state of blue straggler star evolution," Saffer said.
Using Hubble's Faint Object Spectrograph (removed during the Second Servicing Mission in February), the astronomers analyzed the spectrum of one blue straggler, measuring its temperature, diameter, and rotation rate.
The team then combined these measurements with the blue straggler's apparent brightness, taken from a Hubble Telescope Wide Field and Planetary Camera archival image, to obtain the star's mass. The astronomers report the derived temperature and mass are consistent with the characteristics of a normal star with a mass about 1.7 times that of the Sun. However, the star is spinning at least two to three times faster than stars of its kind.
Astronomers now believe that blue stragglers are created by the merger of two low-mass stars. But they have two different views of how these stars interact to create blue stragglers. One merger theory proposes that a violent collision of two unrelated stars creates a blue straggler. Another hypothesizes that a slow coalescence of a gravitationally bound pair creates the straggler star.
Based on their analysis of the blue straggler in 47 Tuc, the team favors the slower, gentler merger scenario between binary stars. In double-star systems where the stars are close enough to touch each other, the more massive star can cannibalize its partner, producing a single, even more massive star.
This process, astronomers believe, more likely results in a rapidly spinning merger product where the fast orbital motions of the binary star produces the rapid spin of the consolidated pair.
The second merger scenario involves a collision between two unrelated stars, which run into each other by chance in the dense star cluster core.
"It's a bit like a head-on wreck between two tractor trailers," Saffer explained, "where the enormous energy carried by the fast-moving stars is deposited in the debris from the collision."
Saffer credits the Hubble telescope's superior spatial resolution with being able to peer into a swirling mix of stars to capture a blue straggler in the cluster core."While some blue stragglers are found in globular cluster outskirts, in 47 Tuc the blue stragglers are only found in the cluster core," Saffer said. "The crowding of the stars there is too severe for the current generation of ground-based telescopes to resolve them."
Globular clusters contain up to 1 million stars packed into a swarm only about 20 light-years in diameter. They also are among the oldest stellar systems in the Milky Way Galaxy. Stars speeding through the extremely crowded cluster core are far more likely to collide or pass near each other than stars in the sparse neighborhood of the Sun.
These processes can produce a zoo of stellar animals, such as X-ray binaries, pulsars, blue stragglers, and other exotic species, all of which have actually been observed in globular cluster cores.
The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc., for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency.
NASA's Hubble Space Telescope has uncovered over 1,000 bright, young star clusters bursting to life in a brief, intense, brilliant "fireworks show" at the heart of a pair of colliding galaxies.
"The sheer number of these young star clusters is amazing," said Dr. Brad Whitmore of the Space Telescope Science Institute, Baltimore, MD. "The discovery will help us put together a chronological sequence of how colliding galaxies evolve. This will help us address one of the fundamental questions in astronomy: why some galaxies are spirals while others are elliptical in shape."
"These spectacular images are helping us understand how globular star clusters formed from giant hydrogen clouds in space," added Dr. Francois Schweizer of the Carnegie Institution of Washington, Washington, DC. "This galaxy is an excellent laboratory for studying the formation of stars and star clusters since it is the nearest and youngest example of a pair of colliding galaxies."
By probing The Antennae galaxies (called The Antennae because a pair of long tails of stars formed by the encounter resembles an insect's antennae) and some of the other nearby galactic-scale collisions, Hubble is coming up with a variety of surprises:
\B╖\b Globular star clusters are not necessarily relics of the earliest generations of stars formed in a galaxy, as once commonly thought, but also may provide fossil records of more recent collisions.
\B╖\b The "seeds" for star clusters appear to be huge clouds (tens of light years across) of cold hydrogen gas, called giant molecular clouds, which are squeezed by surrounding hot gas heated during the collision and then collapse under their own gravity. Like a string of firecrackers being ignited by the collision, these reservoirs of gas light up in a great burst of star formation.
\B╖\b The ages of the resulting clusters provide a clock for estimating the age of a collision. This offers an unprecedented opportunity for understanding, step-by-step, the complex sequence of events which takes place during a collision, and possibly even the evolution of spiral galaxies into elliptical galaxies.
Earlier Hubble pictures show that nearly a third of very distant galaxies, which existed early in the history of the universe, appear to be interacting galaxies, like The Antennae.
In particular, the Hubble Deep Field (a "long-exposure" image from Hubble looking at galaxies far back into time) uncovered a plethora of odd-shaped, disrupted-looking galaxies. They offer direct visual evidence that galaxy collisions were more the rule than the exception in the early formation period of the universe.
These distant galaxy collisions are too faint and too small to study in much detail. Astronomers say we are fortunate to have such a nearby example as The Antennae to study, since collisions between galaxies are relatively rare today. "The degree of detail in these images is astounding, and represents both a dream come true and a nightmare when it comes to the analysis of such a large amount of data," Whitmore said.
In addition to providing a window into how stars and galaxies formed in the dim past, the Hubble views also might offer a glimpse of the future fate of Earth's home galaxy, the Milky Way, when it either sideswipes or plows head-on into the neighboring Andromeda galaxy.
The Hubble observations of The Antennae galaxies, as well as several other nearby colliding galaxies, were conducted by Whitmore and co-investigators Schweizer, Bryan Miller, Michael Fall, and Claus Leitherer over the past several years.
Hubble's resolution and sensitivity allowed the team to uncover over 1,000 exceptionally bright young star clusters, sometimes called super star clusters, within The Antennae -- the prototypical galaxy smashup.
Ground-based images only were able to see the brightest of these clusters, and even in these cases were not able to show that the clusters were very compact with the sizes of normal globular clusters.
Observing other galaxy collisions, the Hubble team discovered the presence of young star clusters which were very bright and blue in the case of ongoing collisions, but had faded to become fainter and redder for the older merger remnants. This allowed them to place the snapshots of galaxy collisions into a chronological sequence.
The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc. for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency.
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"Hubble Identifies What May Be the Most Luminous Star Known (7 October, 1997)",470,0,0,0
Astronomers using NASA's Hubble Space Telescope have identified what may be the most luminous star known -- a celestial mammoth which releases up to 10 million times the power of the Sun and is big enough to fill the diameter of Earth's orbit. The star unleashes as much energy in six seconds as our Sun does in one year.
The image, taken by a University of California, Los Angeles (UCLA)-led team with the recently installed Near-Infrared Camera and Multi-Object Spectrometer (NICMOS) aboard Hubble, also reveals a bright nebula, created by extremely massive stellar eruptions. The nebula is so big (four light-years) that it would nearly span the distance from the Sun to Alpha Centauri, the nearest star to Earth's solar system.
The astronomers estimate that when the titanic star was formed one to three million years ago, it may have weighed up to 200 times the mass of the Sun before shedding much of its mass in violent eruptions.
"This star may have been more massive than any other star, and now it is without question still among the most massive -- even at the low end of our estimates," says Don F. Figer of UCLA. "Its formation and life stages will provide important tests for new theories about star birth and evolution."
The UCLA astronomers estimate that the star, called the "Pistol Star" (for the pistol shaped nebula surrounding it), is approximately 25,000 light-years from Earth near the center of the Milky Way galaxy. The Pistol Star is not visible to the eye, but is located in the direction of the constellation Sagittarius, hidden behind the great dust clouds along the Milky Way.
The Pistol Star was first noted in the early 1990s, but its relationship to the nebula was not realized until 1995, when Figer proposed in his Ph.D. thesis that the "past eruptive stages of the star" might have created the nebula. The Hubble spectrometer results confirm this conclusion.
The astronomers believe that the Pistol nebula was created by eruptions in the outer layers of the star which ejected up to 10 solar masses of material in giant outbursts about 4,000 and 6,000 years ago. The star will continue to lose more material, eventually revealing its bare hot core, sizzling at 100,000 degrees.
Burning at such a dramatic rate, the Pistol Star is destined for certain death in a brilliant supernova in 1-3 million years. "Massive stars are burning their candles at both ends; they are so luminous that they consume their fuel at an outrageous rate, burning out quickly and often creating dramatic events, such as exploding as supernovae," said Mark Morris, a UCLA professor of astronomy and co-investigator. "As these stars evolve, they can eject substantial portions of their atmospheres -- in the case of the Pistol Star, producing the nebula and an extreme stellar wind (outflow of charged particles) that is 10 billion times stronger than our Sun's."
The Pistol Star would be visible to the naked eye as a fourth magnitude star in the sky (which is quite impressive given its distance of 25,000 light-years) if it were not for interstellar dust clouds of tiny particles between the Earth and the center of the Milky Way that absorb the star's light.
The most powerful telescopes cannot see the Pistol Star in visible wavelengths. However, ten percent of the infrared light leaving the Pistol Star reaches Earth, putting it within reach of infrared telescopes, which have seen rapid technological advances in recent years -- spurred by projects such as NICMOS.
The Pistol Star was so massive when it was born that it brings into question current thinking about how stars are formed, say the UCLA astronomers. In the current view, stars form within large dust clouds which contract under their own gravity, eventually forming hot clumps that ignite the hydrogen fusion process.
The star may radiate enough energy to halt the inward fall of material, thus limiting its maximum mass. The initial mass of the Pistol Star may have exceeded this theoretical upper limit. "It is perhaps no accident that this extreme-mass star is found near the center of the galaxy," says Morris. "Current evidence leads us to believe that the star formation process there may favor stars much more massive than our modest Sun."
Over the coming year, the team will be using the new near-infrared spectrometer that Ian S. McLean's team is building at UCLA for the giant 10-meter Keck II telescope in Hawaii. The new instrument will be used to measure the velocities of the expanding gas shells.
In addition to Figer, Morris, and McLean, the team also includes Caltech physicist Gene Serabyn and Columbia University astronomer R. Michael Rich.
The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc. (AURA), for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency.
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"NASA Receives Approval To Launch Cassini Mission (3 October, 1997)",471,0,0,0
NASA today received formal approval from the White House Office of Science and Technology Policy (OSTP) to proceed toward the launch of the robotic Cassini mission to explore Saturn and its moon Titan. "NASA and its interagency partners have done an extremely thorough job of evaluating and documenting the safety of the Cassini mission. I have carefully reviewed these assessments and have concluded that the important benefits of this scientific mission outweigh the potential risks," said OSTP Director Dr. John H. Gibbons, who signed the launch approval.
NASA Administrator Daniel S. Goldin said, "I am confident in the safety of the Cassini mission, and I fully expect that it will return spectacular images and scientific data about Saturn, in the same safe and successful manner as the Voyager, Galileo and Ulysses missions." White House launch approval is required by presidential directive due to the type of power source used to provide electrical power for the Cassini spacecraft and its scientific instruments, and the heater units that it carries to keep the spacecraft's instruments and electronics warm in deep space.
The Radioisotope Thermoelectric Generators (RTGs) and Radioisotope Heater Units used to power Cassini and keep its internal systems warm have been used in previous NASA missions ranging from Apollo to Galileo, and have been approved by five previous administrations ranging from Nixon to Bush. RTGs produce power by the heat generated through the natural radioactive decay of non-weapons grade plutonium dioxide, which is transformed into electricity by solid-state thermoelectric converters.
Before Administrator Goldin sent the request for launch approval to OSTP, two separate processes were completed to address the environmental and safety aspects of the mission. NASA completed an Environmental Impact Statement in June 1995 and a supplement in June 1997, as required by the National Environmental Policy Act and NASA policy. Consistent with long-standing Presidential policy, the Department of Energy (DOE) prepared over the past seven years a comprehensive Safety Analysis Report. In addition, an Interagency Nuclear Safety Review Panel, including safety experts from DOE, NASA, the Department of Defense (DOD), the Environmental Protection Agency (EPA), and a technical advisor from the Nuclear Regulatory Commission conducted a comprehensive evaluation of the safety analysis. This panel was supported by over 50 scientific experts from academia and industry.
DOD, EPA and DOE have written to the NASA Administrator confirming that, in their view, the safety analysis conducted for the mission is comprehensive and thorough. Cassini is a cooperative endeavor of NASA, the European Space Agency (ESA) and the Italian Space Agency, or Agenzia Spaziale Italiana. The mission will send a sophisticated robotic spacecraft, equipped with 12 scientific experiments, to orbit Saturn for a four-year period and study the Saturnian system in detail. The ESA- built Huygens probe that will parachute into Titan's thick atmosphere carries another six scientific instrument packages. Saturn is the second-largest planet in the solar system and is made up mostly of hydrogen and helium.
Its placid-looking, butterscotch-colored face masks a windswept atmosphere where jet streams blow at 1,100 miles per hour and swirling storms roil just beneath the cloud tops. Previous spacecraft passing by Saturn found a huge and complex magnetic environment, called a magnetosphere, where trapped protons and electrons interact with each other, the planet, rings and surfaces of many of the moons. Although it is believed to be too cold to support life, haze- covered Titan is thought to hold clues to how a primitive Earth evolved into a life-bearing planet. It has an Earth-like, nitrogen- based atmosphere and a surface that many scientists believe probably features chilled lakes of ethane and methane.
Scientists believe that Titan's surface is probably coated with the residue of a sticky brown organic rain. The launch of Cassini aboard a Titan IV-B/Centaur launch vehicle is scheduled for 4:55 a.m. EDT on October 13 from Cape Canaveral Air Station, FL. An on-time launch will deliver the Cassini mission to Saturn almost seven years later on July 1, 2004. Cassini's primary mission concludes in July 2008.
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"Science Team Chosen For Mission To Explore Asteroid (29 September, 1997)",472,0,0,0
Ten planetary scientists have been chosen to lead the analysis of measurements to be made by miniaturized instruments carried aboard a mission called Deep Space-1, the first flight in NASA's New Millennium Program. Scheduled for launch in July 1998, Deep Space-1 (DS-1) is intended to validate advanced instrument and spacecraft systems technologies required for low-cost space science missions.
The spacecraft will conduct flybys of an asteroid, a comet and Mars. DS-1 science investigation proposals were evaluated on the basis of their scientific ideas, and the unique theoretical and analytical capabilities that they would bring to bear in meeting the overall mission objectives and its cost constraints.
The selected scientists are: Frances Bagenal, University of Colorado, Boulder Daniel Boice, Southwest Research Institute, San Antonio, TX Daniel Britt, University of Arizona, Tucson Bonnie Buratti, Jet Propulsion Laboratory (JPL), Pasadena, CA Jurgen Oberst, the German Aerospace Research Establishment (DLR), Berlin Tobias Owen, University of Hawaii, Honolulu Laurence Soderblom, U.S. Geological Survey, Flagstaff, AZ Alan Stern, Southwest Research Institute, Boulder, CO Nicolas Thomas, Max-Planck-Institut fur Aeronomie, Lindau, Germany David Young, Southwest Research Institute, San Antonio.
DS-1's primary science goals include detailed studies of the characteristics of the solar wind, the stream of charged particles emitted by the Sun, and learning more about the physical properties of the asteroid McAuliffe (January 1999 flyby) and Comet P/West-Kohoutek-Ikemura (June 2000), including the comet's nucleus and its plasma properties. The DS-1 spacecraft science instrument package has two main components. The Miniature Integrated Camera Spectrometer (MICAS) encompasses a camera, an ultraviolet imaging spectrometer and an infrared imaging spectrometer, all within one 26-pound (12-kilogram) package.
The Plasma Experiment for Planetary Exploration (PEPE) combines multiple instruments into one compact 13-pound (six-kilogram) package designed to determine the three-dimensional distribution of plasma, or electrically charged particles, over its field of view. PEPE includes a very low-power, low-mass micro-calorimeter to help understand plasma-surface interactions and a plasma analyzer to identify the individual molecules and atoms in the immediate vicinity of the spacecraft that have been eroded off the surface of the asteroid and the comet "NASA could not have selected a better team of investigators. The results of the DS-1 investigations will make a significant contribution to our understanding of the conditions in the early Solar System," said Dr. Robert Nelson, the DS-1 project scientist at NASA's Jet Propulsion Laboratory, Pasadena, CA.
"We will learn more about the material from which planets condensed and life evolved. Ultimately, we will learn more about ourselves." The 12 advanced systems technologies to be validated by DS-1 include solar electric propulsion, high-power solar concentrator arrays, autonomous on-board optical navigation, and several telecommunications and microelectronics devices. "We are conducting science on Deep Space-1 in order to demonstrate that the technologies being tested are compatible with future science-focused missions, and to take full advantage of this rare opportunity to send a capable spacecraft to such interesting solar system targets," explained Dr. Marc Rayman, DS-1 Chief Mission Engineer at JPL.
A significant step toward revealing the mysteries of gamma- ray bursts was taken this week by The Johns Hopkins University Applied Physics Laboratory (APL), Laurel, MD, when NASA's Near Earth Asteroid Rendezvous (NEAR) spacecraft sent back unexpected data showing a major gamma-ray burst. APL manages the NEAR mission for NASA. The observation came after researchers reconfigured the gamma-ray spectrometer to make more frequent data returns as NEAR travels to a rendezvous with the asteroid Eros in February 1999.
If as distant as new evidence suggests, gamma-ray bursts are the most violent explosions known, emitting in one second as much energy as the Sun will emit in its lifetime. The gamma-ray spectrometer was not originally planned to begin its work until the spacecraft reached Eros. But while en route a simple software change was added that gave a new astrophysics capability to this planetary spectrometer, which resulted in detection of a gamma-ray burst on Sept. 15, that lasted for about 10 seconds. Since then six more bursts have been detected. Several of the bursts have been confirmed by the European Space Agency/NASA Ulysses spacecraft, now in a polar orbit around the sun and by two detectors on NASA's Wind spacecraft near the Earth.
These three spacecraft, along with other Earth-orbiting spacecraft, form a 3-dimensional interplanetary network for observing gamma-ray bursts that has not been possible since the loss of the Mars Observer in 1993. "Seeing this burst validates that the NEAR detector can be a true working partner in the interplanetary network for gamma-ray burst detection," says APL's lead gamma-ray instrument engineer John Goldsten, who was the first to see the gamma-ray burst data. Jacob Trombka, NASA's Science Team Leader for the gamma- ray instrument, says, "NEAR's instrument is more sensitive than we believed it would be. It's seeing bursts that other spacecraft aren't seeing." The success of the instrument is the result of a good design, he says.
"Originally we didn't have time to include a burst mode on the instrument, but the system was so well designed that we were able to upload such a system a few weeks ago." Gamma-ray bursts remain one of the great mysteries of astrophysics since their discovery more than 30 years ago. They tend to be randomly distributed over the sky and occur with a frequency of about one per day for the most sensitive detectors.
If they are of cosmological origin, they represent the most powerful events that are known in the universe. The debate as to their local or cosmological origin will most likely be resolved by locating sources of gamma-ray bursts and then identifying them with optical and radio telescopes. NASA's Hubble Space Telescope made the first observation of a fuzzy object associated with a burst that was detected last Feb. 28 by the Italian BeppoSAX satellite.
The sources of gamma-ray bursts can be located in the sky by timing the arrival of the gamma-rays at three well-separated spacecraft. Since 1993, the spacecraft instrumented to observe such bursts have been the Ulysses spacecraft plus several spacecraft near the Earth: the BeppoSAX and Wind as well as NASA's Compton Gamma-Ray Observatory (CGRO) and Rossi X-ray Timing Explorer. The near-Earth spacecraft are too close to each other to allow a unique determination of the location of the bursts.
The addition of the Near Earth Asteroid Rendezvous spacecraft to the interplanetary network will provide 3- dimensional triangulation of source locations and should greatly increase the probability of associating a gamma-ray source with a particular source from optical and radio telescopes. The new capability on NEAR will expand the network and enable it to obtain the locations of moderate and stronger bursts, which occur at least several times per month, to a position in the sky accurate to about a minute of arc (about a thirtieth of the size of the moon). NEAR, the first mission of NASA's Discovery Program for "faster, better, cheaper" planetary exploration, will be the first spacecraft to orbit an asteroid.
On June 27 NEAR sent back spectacular images of 253 Mathilde as it flew past the asteroid. In February 1999, NEAR will reach Eros and begin the first long- term, close-up look at an asteroid's surface composition and physical properties. NEAR was designed and built by The Johns Hopkins University Applied Physics Laboratory, in Laurel, MD, which also manages the program for NASA.
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"Mars Pathfinder Rover Exits Rock Garden (26 September, 1997)",474,0,0,0
After 83 days of atmospheric, soil and rock studies, NASA's Mars Pathfinder is moving into extended mission activities that will take the rover on its longest trek yet, while the lander camera completes its biggest and best landscape panorama. "The lander and rover performance continues to be nothing short of extraordinary," said Brian Muirhead, Mars Pathfinder project manager at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "We have proven that we know how to design robust robots to operate in the hostile environment of Mars." The rover has just completed its last alpha proton X-ray spectrometer study for a while, taking compositional measurements of a rock nicknamed Chimp, located just behind and to the left of an area scientists call the Rock Garden.
Once data from the spectrometer have been retrieved, Sojourner will begin a 164-foot (50-meter) clockwise stroll around the lander to perform a series of technology experiments and hazard avoidance exercises. Meanwhile, the Pathfinder lander camera is continuing to image the Martian landscape in full-resolution color as part of its goal to provide a "super panorama" image of the Ares Vallis landing site. Each frame of this panorama is imaged using 12 color filters plus stereo. "The super pan will be our biggest and best imaging data product," Muirhead said. "It is made up of 1 gigabit (1 billion bits) of data, of which we've received more than 80 percent. Given our limited downlink opportunities, we should have the full image by the end of October."
The 22-pound (10.5-kilogram) rover has survived 10 times longer than its primary mission design of seven days, while the lander has now been operating 2.5 times longer than it was originally expected to operate, according to Richard Cook, Mars Pathfinder mission manager. Both vehicles are solar-powered, but carried batteries to conduct night-time science experiments and keep the lander warm during the sub-freezing nights on Mars. Normal usage has fully depleted the rover's non-rechargeable batteries, limiting it to daylight activities only. The lander battery, which packed more than 40 amp-hours of energy on landing day, performed perfectly during the 30-day primary mission, but is now down to less than 30 percent of its original capacity.
"We expected to begin seeing this type of degradation on both vehicles and, of course, designed both the lander and rover to operate without batteries altogether," Cook said. "If everything else continues to operate properly, we could continue conducting surface experiments for months." About once every two weeks, the lander battery is used to perform some night-time science experiments, he added. The primary activity is acquiring meteorological data and images of morning clouds, as well as images of Mars' two small moons, Phobos and Deimos. Despite the lack of battery power, the rover has continued taking successful spectrometer readings during the day.
In the next two weeks, engineers will drive the vehicle back to a magnetic target on the ramp from which Sojourner first touched Martian soil. "This analysis of the dust on the ramp magnet is a very important science measurement," noted Dr. Matthew Golombek, Mars Pathfinder project scientist. "The results should give us a clue about how all this magnetic dust was formed."
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"Hubble Sees A Neutron Star Alone In Space (24 September, 1997)",475,0,0,0
Astronomers using NASA's Hubble Space Telescope have taken their first direct look, in visible light, at a lone neutron star. This offers a unique opportunity to pinpoint its size and to narrow theories about the composition and structure of this bizarre class of gravitationally collapsed, burned-out stars. By successfully characterizing the properties of an isolated neutron star, astrophysicists have an opportunity to better understand the transitions matter undergoes when subjected to the extraordinary pressures and temperature found in the intense gravitational field of a neutron star.
The Hubble results show the star is very hot, and can be no larger than 16.8 miles (28 kilometers) across. These results prove that the object must be a neutron star, for no other known type of object can be this hot and small. "This puts the neutron star uncomfortably close to the theoretical limit of how small a neutron star should be," says Fred Walter of the State University of New York (SUNY) at Stony Brook. "With this observation we can begin to rule out some of the many models of the internal structure of neutron stars."
The observation results, made by Walter and Lynn Matthews (also of SUNY), are reported in the Sept. 25 issue of Nature magazine. Neutron stars, which are created in some supernovae, are so dense because the electrons and protons that form normal matter have been squeezed into neutrons and other exotic subatomic particles. Neutron star matter is the densest form of matter known to exist. (Theoretically, a piece of neutron star surface weighing as much as a fleet of battleships would be small enough to be held in the palm of your hand.)
The Hubble observations, combined with earlier data, promise to help astronomers refine the mathematical description -- called the equation of state -- of the complex transformations matter undergoes at extraordinary densities not found on Earth. Equations of state are well understood for "everyday" matter such as water, which can transition between gaseous, liquid and solid states. But the behavior of matter under extreme temperature and pressure found on a neutron star is not well understood. Several hundred million neutron stars should exist in our galaxy. However, all neutron stars now known have either been found orbiting other stars in X-ray binary systems or emitting machine-gun blasts of radio energy as pulsars (a class of neutron star).
The neutron star seen by Hubble is not a member of a binary system, and is not known to pulse at X-ray or radio wavelengths (it has not been detected as a radio source). Pulsars are young neutron stars born with strong magnetic fields; non- pulsing neutron stars may be old, dead pulsars, with ages of more than a million years, or they may never have been pulsars. Only a few lone neutron star candidates have been pinpointed through X- ray observations, and this is the first optical counterpart to be identified.
The first clue that there was a neutron star at this location came in 1992, when the ROSAT (the Roentgen Satellite) found a bright X-ray source without any optical counterpart in optical sky surveys. It drew the attention of astronomers because objects this hot and bright, without counterparts at other wavelengths, are extremely rare. Hubble's Wide Field Planetary Camera 2 was used in October 1996 to undertake a sensitive search for the optical object, and found a stellar pinpoint of light within only 2 arc seconds (1/900th the diameter of the Moon) of the X-ray position. Astronomers haven't directly measured the neutron star's distance but fortunately the neutron star lies in front of a molecular cloud known to be about 400 light-years away in the southern constellation Coronae Australis.
Using the distance to the cloud as an upper limit, the astronomers calculated a diameter by next comparing the neutron star's brightness and color as measured by Hubble, along with X- ray brightness from the ROSAT and EUVE (Extreme Ultraviolet Explorer) satellites. The object is brightest at X-ray wavelengths. In the two Hubble images, the object is brighter at ultraviolet wavelengths than at visible wavelengths. They concluded they are directly seeing an ultracompact surface sizzling at about 1.2 million degrees Fahrenheit.
To be so hot, yet so dim (below 25th magnitude in visual light) and relatively close to Earth, the object must be extremely small -- below the size of a white dwarf, a more common stellar cinder. A hot white dwarf at this magnitude would lie 150,000 light-years away (outside our galaxy), and have 1/70,000 as much X-ray emission. The 16.8-mile diameter estimate comes from assuming the neutron star is at the farthest it can be, just in front of the obscuring "wall" of the molecular cloud. If instead the neutron star is significantly closer to us, say midway to the molecular cloud, it would be smaller still, and present an even bigger challenge to the theories of the equation of state of nuclear matter.
Although neutron stars in binary systems allow astronomers to measure their mass, which turn out to be consistent with theory, it's much harder for astronomers to estimate the diameter of the neutron stars. Since the neutron stars "feed" on their companion stars in these systems, the light does not come exclusively from the surface but from jets, disks and other phenomenon that occur around the star. This can lead to inaccurate size estimates. Over the next year, planned observation with the Hubble will be used in an attempt to determine exactly how far away and how large the star is.
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"Cassini Launch Rescheduled For Oct. 13 (19 September, 1997)",476,0,0,0
The launch of NASA's Cassini spacecraft aboard a U.S. Air Force Titan IVB rocket has officially been rescheduled on the Eastern Range for Monday, Oct. 13. The payload is now back at Complex 40 atop the Titan IV Centaur. The launch window for Cassini extends from 4:55 to 7:15 a.m. EDT. Cassini is a joint NASA-European Space Agency (ESA) mission to Saturn, which is scheduled to arrive at the ringed planet in 2004 after more than six years of interplanetary travel. After arrival, the spacecraft will orbit Saturn for four years studying the gas giant planet, its rings and moons, and the ESA-built Huygens probe will descend to the surface of the giant moon Titan.
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"Twin Telescopes With Near-Infrared 'Eyes' Begin All-Sky Survey (17 September, 1997)",477,0,0,0
The first of a pair of new telescopes, funded primarily by NASA, has begun an ambitious three-and-a-half year near- infrared survey of the entire celestial sky, peering through the curtain of interstellar dust in the Milky Way galaxy. The Two-Micron All-Sky Survey (2MASS), based at the University of Massachusetts, Amherst, MA, features two 1.3-meter telescopes, one at a Smithsonian Astrophysical Observatory site atop Mount Hopkins, near Tucson, AZ, and the other at a National Optical Astronomy Observatories site in Cerro Tololo, Chile. "The sky survey catalogues produced 100 years ago are still useful to astronomers," said Project Manager Rae Stiening.
"We expect this new, greatly updated survey will be an invaluable resource for the next 100 years." "Preliminary observations by 2MASS are already suggesting new infrared sources will be discovered," said Program Manager Dr. Michael Klein at NASA's Jet Propulsion Laboratory, Pasadena, CA. "Some of these will be targets for detailed studies for future space observatories, like the Advanced X-Ray Facility (AXAF), the Space Infrared Telescope Facility and the Next Generation Space Telescope."
The survey is designed to catalogue one million galaxies and 300 million stars in the local universe, along with quasars, which are strong, extremely bright radio sources, and galaxies with black holes, the intriguing entities with gravity so powerful not even light can escape. 2MASS will observe many known asteroids and possibly some comets, and it is uniquely sensitive to exotic objects like brown dwarfs, which lack the mass needed to ignite and become full-fledged stars.
The telescopes are equipped with near- infrared detector arrays that will provide the most complete census to date of cool stars in the Milky Way galaxy and provide new data for detailed studies of the galactic structure. Near- infrared emission is at wavelengths roughly two-to-four times longer than visible light and permits astronomers to "see through" the obscuring effects of interstellar dust in the Milky Way galaxy. As Stiening explained, "Sunsets on Earth look reddish because only red light makes it through the dust in our atmosphere. Infrared observations enable us to penetrate the dust in our galaxy and other galaxies and, therefore, they provide a much clearer view of interior regions." The 2MASS survey will measure accurately the positions and infrared brightness of stars and galaxies. Combined with complementary ground-based red shift surveys, the 2MASS extra- galactic data will provide a three-dimensional view of large-scale structures in the local universe. The enabling technology for this survey is the breakthrough in large-format infrared detector arrays.
These technologies, funded through the U.S. Department of Defense and NASA, are being adapted for astronomical purposes to increase sensitivity dramatically. It's expected the new survey will be some 25,000 times more sensitive than a precursor survey at the California Institute of Technology, Pasadena, CA, nearly 30 years ago. 2MASS uses the type of detectors developed for the Near Infrared Camera and Multi Object Spectrometer on NASA's Hubble Space Telescope. "Observing time at most telescopes is divided among a variety of scientific programs using a suite of different instruments.
2MASS telescopes will be completely dedicated to mapping the sky using one instrument, a three-color infrared camera," said Principal Investigator Dr. Michael Skrutskie, a University of Massachusetts physics and astronomy professor, who leads the science working group that will evaluate the data products. He also managed the design and fabrication effort for the infrared cameras, which are attached to an identical pair of telescopes. Data will be processed at JPL's Infrared Processing and Analysis Center at Caltech.
Every two nights, the center will process 60 gigabytes of data, which is more data than processed during the entire Infrared Astronomy Satellite (IRAS) mission of 1983. The 2MASS survey is funded by NASA's Office of Space Science, the National Science Foundation, the U.S. Naval Observatory and the University of Massachusetts.
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"Mars Global Surveyor Detects Martian Magnetic Field (17 September, 1997)",478,0,0,0
Scientists have confirmed the existence of a planet-wide magnetic field at Mars using an instrument on-board NASA's Mars Global Surveyor orbiter, as the spacecraft began to circle and study the planet from a highly elliptical orbit. "Mars Global Surveyor has been in orbit for only a few days, yet it already has returned an important discovery about the Red Planet," said Vice President Al Gore.
"This is another example of how NASA's commitment to faster, better, cheaper Mars exploration that began with Mars Pathfinder is going to help answer many fundamental questions about the history and environment of our neighboring planet, and the lessons it may hold for a better understanding of life on Earth." The spacecraft's magnetometer, which began making measurements of Mars' magnetic field after its capture into orbit on Sept. 11, detected the magnetic field on Sept. 15. The existence of a planetary magnetic field has important implications for the geological history of Mars and for the possible development and continued existence of life on Mars.
"Preliminary evidence of a stronger than expected magnetic field of planetary origin was collected and is now under detailed study," said Dr. Mario H. Acuna, principal investigator for the magnetometer/electron reflectrometer instrument at NASA's Goddard Space Flight Center, Greenbelt, MD. "This was the first opportunity in the mission to collect - more - - 2 - close-in magnetic field data. Much more additional data will be collected in upcoming orbits during the aerobraking phase of the mission to further characterize the strength and geometry of the field. The current observations suggest a field with a polarity similar to that of Earth's and opposite that of Jupiter, with a maximum strength not exceeding 1/800ths of the magnetic field at the Earth's surface."
This result is the first conclusive evidence of a magnetic field at Mars. "More distant observations obtained previously by the Russian missions Mars 2,3 and 5 and Phobos 1 and 2 were inconclusive regarding the presence or absence of a magnetic field of internal origin," said Acuna. The magnetic field has important implications for the evolution of Mars. Planets like Earth, Jupiter and Saturn generate their magnetic fields by means of a dynamo made up of moving molten metal at the core.
This metal is a very good conductor of electricity, and the rotation of the planet creates electrical currents deep within the planet that give rise to the magnetic field. A molten interior suggests the existence of internal heat sources, which could give rise to volcanoes and a flowing crust responsible for moving continents over geologic time periods. "A magnetic field shields a planet from fast-moving, electrically charged particles from the Sun which may affect its atmosphere, as well as from cosmic rays, which are an impediment to life," Acuna said.
"If Mars had a more active dynamo in its past, as we suspect from the existence of ancient volcanoes there, then it may have had a thicker atmosphere and liquid water on its surface." It is not known whether the current weaker field now results from a less active dynamo, or if the dynamo is now extinct and what the scientists are observing is really a remnant of an ancient magnetic field still detectable in the Martian crust. "Whether this weak magnetic field implies that we are observing a fossil crustal magnetic field associated with a now extinct dynamo or merely a weak but active dynamo similar to that of Earth, Jupiter, Saturn, Uranus and Neptune remains to be seen," Acuna said.
Mars Global Surveyor's magnetometer discovered the outermost boundary of the Martian magnetic field -- known as the bow shock -- during the inbound leg of its second orbit around the planet, and again on the outbound leg. The discovery came just before Mars Global Surveyor began its first aerobraking maneuver to lower and circularize its orbit around Mars, said Glenn Cunningham, Mars Global Surveyor project manager at NASA╣s Jet Propulsion Laboratory (JPL), Pasadena, CA.
"This first 'step down' into the upper atmosphere was performed in two stages," Cunningham said. "On Sept. 16, during the farthest point in the spacecraft's orbit, called the apoapsis, the spacecraft fired its main engine for 6.5 seconds, slowing Global Surveyor's velocity by 9.8 miles per hour (4.41 meters per second). This maneuver lowered the spacecraft's orbit from 163 miles (263 kilometers) to 93 miles (150 kilometers) above the surface of the planet. At its closest approach to Mars this morning, known as the periapsis, the spacecraft dipped into the upper fringes of the Martian atmosphere for 27 seconds, allowing the drag on its solar panels to begin the long aerobraking process of circularizing its orbit."
Mars Global Surveyor will continue aerobraking through the Martian atmosphere for the next four months, until its orbit has been circularized and it is flying about 234 miles (378 kilometers) above the Martian surface. All systems and science instruments onboard the spacecraft continue to perform normally after six days in orbit around the red planet.
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"Hubble Stays On Trail of Fading Gamma-Ray Burst Fireball (16 September, 1997)",479,0,0,0
New Hubble Space Telescope observations of the ever- fading fireball from one of the universe's most mysterious phenomena -- a gamma-ray burst -- is reinforcing the emerging view that these titanic explosions happen far away in other galaxies, and so are among the most spectacularly energetic events in the universe. The most recent finding from observations with Hubble's Imaging Spectrograph made on Sept. 5 -- nearly six months after the blast -- is being reported today at the Fourth Huntsville Symposium on Gamma Ray Bursts, at NASA's Marshall Space Flight Center, Huntsville, AL.
"Hubble is the only telescope capable of continuing to watch the aftermath of this explosion, because it has faded to 1/500th its brightness when first discovered by ground-based telescopes last March," says Andrew Fruchter of the Space Telescope Science Institute in Baltimore, MD. "These observations provide an unprecedented opportunity to better understand the catastrophe behind such incredible outbursts." Hubble's key findings: 1. The continued visibility of the burst, and the rate of its decline over time, support theories that produce the light from a gamma-ray burst in a "relativistic" fireball (expanding at nearly the speed of light) located at extragalactic distances.
A burst in our galaxy, at the observed brightness, would have been slowed by the interstellar medium within the first few weeks and faded from sight by now. 2. The observations contradict earlier claims, by some astronomers, that the gamma-ray burst is moving against the sky background (this offset is called proper motion). Had proper motion been detected, the gamma-ray burst would have had to be no farther than about 30,000 light years, or about the distance to the center of the galaxy. 3. The fuzzy companion object in which the fireball is embedded -- as first confirmed by Hubble in March 26 observations -- has not noticeably faded.
This means it is not a relatively nearby nebula produced by the explosion, but in all likelihood a host galaxy. 4. Since the burst did not occur at the center of the host galaxy, but near its edge, the gamma-ray burst phenomenon is not related to activity in the nucleus of a galaxy. The Hubble observations support the "fireball" model for a gamma-ray burst. "These observations are consistent with colliding neutron stars creating the fireball, but do not require it. The cause of that fireball is still not determined. Though colliding neutron stars is one theoretical means of producing such a fireball, it is not the only one," says Fruchter. Hubble observations over the past six months show the fireball is fading at a constant rate, as predicted by theory.
Eventually, gas plowed in front of the stellar tidal wave should build up enough resistance to bring the fireball to a halt -- like snow piling up in front of a plow -- and it should blink out. The fact that that hasn't happened yet, however, offers more clues to solving the gamma-ray burst mystery. If the burst had happened nearby, the resulting fireball should have had only enough energy to propel it into space for a month or so before "hitting the wall" of accumulated gas and dying out. The fact that this fireball has expanded to gargantuan size, sweeping out a bubble of space one light-year across, means the explosion was truly titanic and, to match the observed brightness, must have happened at the vast distances of galaxies.
When Hubble first observed the fireball on March 27 (several weeks after the initial discovery), it was at 26th magnitude. The magnitude scale is used to measure the brightness of objects in space. The lower the magnitude, the brighter the object. The unaided eye can detect objects of the 6th magnitude. By the Sept. 5 observation it had faded to one-fifth that brightness to 27.7 magnitude (approximately 1/500,000 the brightness of the faintest star). The suspected host galaxy has remained at approximately 25th magnitude. Only Hubble has the resolution and contrast capability to still distinguish the fading fireball from the now- brighter host galaxy.
The researchers hope for follow-up observations to continue keeping track of the burst's optical counterpart until it fades away. The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc. (AURA) for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency (ESA).
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"Repair Work on Cassini Huygens Probe Completed Successfully (12 September, 1997)",480,0,0,0
Engineers at Kennedy Space Center, FL, have completed repairs to damaged thermal insulation on the European Space Agency's Huygens probe, a part of the Cassini mission to Saturn. Based on the amount of time needed to return the spacecraft to the pad from a spacecraft checkout facility and complete the work necessary to be ready for launch, NASA managers have set a tentative launch date of no earlier than Oct. 13.
The launch date will be confirmed by the Air Force after the spacecraft has been mated to the Titan IV/Centaur next week. On Sept. 7, the spacecraft was removed from atop the Air Force Titan IV rocket at Space Launch Complex 40 to repair the insulation inside the Huygens probe. Damage was caused by a higher-than-acceptable flow rate from the air conditioning to the inside of the probe. To make the probe ready for launch again, engineers detached and disassembled the Huygens probe from the Cassini orbiter. Upon inspection, the damage was found to be limited to a few square inches of foam and Kapton insulation. The damaged material was removed and the insulation repaired. The Huygens probe has now been retested and reassembled. Closeouts are under way today in preparation for mating the probe once again to the Cassini orbiter on Saturday.
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"Hubble Reveals Huge Crater on the Surface of the Asteroid Vesta (4 September, 1997)",481,0,0,0
Astronomers have used NASA's Hubble Space Telescope to discover a giant impact crater on the asteroid 4 Vesta. The crater is a link in a chain of events thought responsible for forming a distinctive class of tiny asteroids as well as some meteorites that have reached the Earth. The giant crater is 285 miles across, which is nearly equal to Vesta's 330 mile diameter. If Earth had a crater of proportional size, it would fill the Pacific Ocean basin. Astronomers had predicted the existence of one or more large craters, reasoning that if Vesta is the true "parent body" of some smaller asteroids, then it should have the wound of a major impact that was catastrophic enough to knock off big chunks.
The observations are described in the Sept. 5 issue of Science Magazine. "In hindsight we should have expected finding such a large crater on Vesta," says Peter Thomas of Cornell University, Ithaca, NY. "But it's still a surprise when it's staring you in the face." Another surprising finding is that such a large crater, relative to Vesta's size, might have been expected to cause more damage to the rest of the minor planet. "This is a unique opportunity to study the effects of a large impact on a small object," says Michael Gaffey of Rensselaer Polytechnic Institute, Troy, NY.
"This suggests that more asteroids from the early days of the solar system may still be intact." The collision gouged out one percent of the asteroid's volume, blasting over one-half million cubic miles of rock into space. This tore out an eight-mile deep hole that may go almost all the way through the crust to expose the asteroid's mantle (Vesta is large enough to be differentiated like Earth -- with a volcanic crust, core and mantle, making it a sort of "mini- planet".) Because of the asteroid's small diameter and low gravity, the crater resembles smaller craters on the Moon that have a distinctive central peak.
Towering eight miles, this cone-shaped feature formed when molten rock "sloshed" back to the bullseye center after the impact. One clue for a giant crater came in 1994 when Hubble pictures showed that one side of Vesta's football shape appeared flattened. "We knew then there was something on Vesta that was unusual," says Thomas. The astronomers had to wait for a better view from Hubble when Vesta made its closest approach to Earth in a decade, in May 1996, when the asteroid was 110 million miles away.
A total of 78 Wide Field Planetary Camera 2 pictures were taken. The team then created a topographic model of the asteroid's surface by noting surface irregularities along the limb and at the terminator (day/night boundary) where shadows are enhanced by the low Sun angle. The immense crater lies near the asteroid's south pole. This is probably more than coincidental, say researchers. The excavation of so much material from one side of the asteroid would have shifted its rotation axis so that it settled with the crater near one pole. Unlike some other large asteroids that have jumbled surfaces due to the asteroids' breakup and recollapse, the rest of Vesta's surface is largely intact, despite the cataclysm.
This is based on previous measurements showing it has a surface of basaltic rock -- frozen lava -- which oozed out of the asteroid's presumably hot interior shortly after its formation 4.5 billion years ago, and has remained largely intact ever since. Approximately six percent of the meteorites that fall to Earth are similar to Vesta's mineralogical signature, as indicated by their spectral characteristics. Vesta's spectrum is unique among all the larger asteroids. The crater may be the ultimate source of many of these meteorites. Most meteorites are believed to come from other asteroids, but their specific objects of origin cannot be determined in most cases.
Thus the distinctive mineralogical makeup of these meteorites means that Vesta is the only world other than the Earth, the Moon and Mars, for which scientists have samples of specifically known origin. A mystery has been that the meteorites could not have traveled directly from Vesta because at Vesta's location in the asteroid belt, there are no perturbing gravitational forces that would cause pieces to fall into orbits intersecting the inner planets like apples shaken out of a tree. However, Vesta's "daughter" asteroids -- literally "chips off the block" which have color characteristics similar to Vesta -- are near a "chaotic zone" in the asteroid belt where Jupiter's gravitational tug can redirect fragments into orbits that intersect Earth's orbit.
A good determination of the shape of Vesta was necessary for the next step in interpretation, which will use multi-color images of Vesta obtained with Hubble to study the detailed mineralogy of surface regions, including the region of the giant crater. Also, a team led by Don McCarthy of the University of Arizona plans to obtain additional images of Vesta at longer wavelengths this fall using the new Near Infrared and Multi-Object Spectrometer science instrument onboard Hubble. Members of the Vesta research team are Principal Investigator Ben Zellner of Georgia Southern University; the Co- Investigators are Richard Binzel, MIT; Michael Gaffey, Rensselaer Polytechnic Institute; Alex Storrs, Space Telescope Science Institute; Peter Thomas, Cornell University, and Dr. Ed Wells, Computer Sciences Corporation.
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"Cassini To Survey Worlds of Saturn and Titan (3 September, 1997)",482,0,0,0
The planet Saturn, its famous icy rings, and its enigmatic moon, Titan, are the prime scientific targets of the international Cassini mission, the most ambitious and far-reaching planetary exploration ever mounted. Final preparation of Cassini is now underway for a launch from Cape Canaveral, FL, in October 1997. The mission marks the first time a space probe has attempted to land on the moon of another planet, providing the first direct sampling of the Earth-like atmosphere of Titan and the first detailed pictures of its previously hidden surface.
Titan is Saturn's largest moon, nearly the size of Mars and bigger than either Mercury or Pluto. Cassini, in development since October 1989, is a cooperative endeavor of NASA, the European Space Agency (ESA) and the Italian Space Agency, or Agenzia Spaziale Italiana. The mission will send a sophisticated robotic spacecraft equipped with 12 scientific experiments to orbit Saturn for a four-year period and study the Saturnian system in detail. The ESA-built Huygens probe that will parachute into Titan's thick atmosphere carries another six scientific instrument packages.
"With its bright, complex rings, 18 known moons and magnetic environment, Saturn is a lot like a solar system in miniature form," said Dr. Wesley T. Huntress, NASA's associate administrator for space science. "Saturn's family of rings and moons is a one- stop treasure trove, offering countless clues to the history of planetary and solar system evolution. Cassini and the Huygens probe represent our best efforts yet in our ongoing exploration of the solar system." The launch period for Cassini's nearly seven-year journey to Saturn opens on Oct. 6 at 5:39 a.m. EDT and closes Nov. 15, 1997. A U.S. Air Force Titan IVB/Centaur launch system, the most powerful launch vehicle in the U.S. fleet, will loft Cassini onto the interplanetary trajectory that will deliver the spacecraft to Saturn almost seven years later on July 1, 2004.
Cassini's primary mission concludes in July 2008. Saturn is the second-largest planet in the solar system and is made up mostly of hydrogen and helium. Its placid-looking, butterscotch-colored face masks a windswept atmosphere where jet streams blow at 1,100 miles per hour and swirling storms roil just beneath the cloud tops. Spacecraft passing by Saturn found a huge and complex magnetic environment, called a magnetosphere, where trapped protons and electrons interact with each other, the planet, rings, and surfaces of many of the moons.
Saturn's best known feature -- its bright rings -- consists not just of a few rings but of hundreds of rings and ringlets broad and thin, composed of ice and rock particles ranging in size from grains of sand to boxcars. "Shepherd moons" found orbiting near the edges of some of the rings gravitationally herd in ring particles that would otherwise spread out into deep space. Although it is believed to be too cold to support life, haze- covered Titan is thought to hold clues to how the primitive Earth evolved into a life-bearing planet. It has an Earth-like, nitrogen-based atmosphere and a surface that many scientists believe probably features chilled lakes of ethane and methane.
Scientists believe that Titan's surface is probably coated with the residue of a sticky brown organic rain. On Nov. 6, 2005, Huygens will descend by parachute into Titan's sky, providing our first direct sampling of Titan's atmosphere and the first detailed photos of its hidden surface. The Cassini spacecraft is the most complex interplanetary spacecraft ever built. Because of Cassini's challenging mission, the long distance Cassini must travel, and the value of its scientific return, each component and the system as a whole has undergone an unprecedented program of rigorous testing for quality and performance.
"Every phase of the mission has been reviewed and validated internally and externally by NASA and independent experts," said Huntress. Because of the very dim sunlight at Saturn's orbit, Cassini could not conduct its mission to Saturn on solar power. Electrical power is supplied to Cassini by a set of radioisotope thermoelectric generators (RTGs) which convert the heat from the natural decay of plutonium. RTGs have been used on 23 previous U.S. missions. Plutonium dioxide also is used in 117 radioisotope heater units placed on Cassini and Huygens to keep electronics systems at their operating temperatures.
These units were most recently used on the Mars Pathfinder mission's Sojourner rover to keep the system from failing during cold Martian nights. The mission is named for two 17th century astronomers: Italian-French astronomer Jean-Dominique Cassini made several key discoveries about Saturn, and Dutch scientist Christian Huygens discovered Titan. Development of the Huygens probe was managed by an ESA team located at the European Space Technology and Research Center (ESTEC) in Noordwijk, The Netherlands. The Cassini orbiter was designed, developed and assembled at NASA's Jet Propulsion Laboratory (JPL), located in Pasadena, CA.
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"Two Voyager Spacecraft Still Going Strong After 20 Years (2 September, 1997)",483,0,0,0
Twenty years after their launch and long after their planetary reconnaissance flybys have been completed, both Voyager spacecraft are now gaining on another milestone -- crossing that invisible boundary that separates our solar system from interstellar space, the heliopause. Since 1989 when Voyager 2 encountered Neptune, both spacecraft have been studying the environment of space in the outer solar system. Science instruments on both spacecraft are sensing signals that scientists believe are coming from the heliopause -- the outermost edge of the Sun's magnetic field that the spacecraft must pass through before they reach interstellar space.
"During their first two decades, the Voyager spacecraft have had an unequaled journey of discovery. Today, even though Voyager 1 is now more than twice as far from the Sun as Neptune, their journey is only half over, and more unique opportunities for discovery await the spacecraft as they head toward interstellar space," said Dr. Edward Stone, the Voyager project scientist and director of NASA's Jet Propulsion Laboratory, Pasadena, CA. "The Voyagers owe their ability to operate at such great distances from the Sun to their nuclear electric power sources which provide the electrical power they need to function."
The Sun emits a steady flow of electrically charged particles called the solar wind. As the solar wind expands supersonically into space, it creates a magnetized bubble around the Sun, called the heliosphere. Eventually, the solar wind encounters the electrically charged particles and magnetic field in the interstellar gas. The boundary created between the solar wind and interstellar gas is the heliopause. Before the spacecraft reach the heliopause, they will pass through the termination shock -- the place where the solar wind abruptly slows down from supersonic to subsonic speed.
Reaching the termination shock and heliopause will be major milestones for the spacecraft because no one has been there before and the Voyagers will gather the first direct evidence of their structure. Encountering the termination shock and heliopause has been a long sought-after goal for many space physicists, and exactly where these two boundaries are located and what they are like still remains a mystery. "Based on current data from the Voyager cosmic ray subsystem, we are predicting the termination shock to be in the range of 62 to 90 astronomical units (AU) from the Sun. Most 'consensus' estimates are currently converging on about 85 AU. Voyager 1 is currently at about 67 AU and moving outwards at 3.5 AU per year, so I would expect crossing the termination shock sometime before the end of 2003," said Dr. Alan Cummings, a co-investigator on the cosmic ray subsystem at the California Institute of Technology.
"Based on a radio emission event detected by the Voyager 1 and 2 plasma wave instruments in 1992, we estimate that the heliopause is located from 110 to 160 AU from the Sun," said Dr. Donald A. Gurnett, principal investigator on the plasma wave subsystem at the University of Iowa. (One AU is equal to 93 million miles (150 million kilometers), or the distance from the Earth to the Sun.) "The low-energy charged particle instruments on the two spacecraft continue to detect ions and electrons accelerated at the Sun and at huge shock waves, tens of AU in radius, that are driven outward through the solar wind.
During the past five years, we have observed marked variations in this ion population, but have yet to see clear evidence of the termination shock. We should always keep in mind that our theories may be incomplete and the shock may be a lot farther out than we think," said Dr. Stamatios M. Krimigis, principal investigator for the low energy charged particle subsystem at The Johns Hopkins University Applied Physics Laboratory. Voyager 2 was launched first on Aug. 20, 1977, and Voyager 1 was launched a few weeks later on a faster trajectory on Sept. 5. Initially, both spacecraft were only supposed to explore two planets -- Jupiter and Saturn. But the incredible success of those two first encounters and the good health of the spacecraft prompted NASA to extend Voyager 2's mission to Uranus and Neptune.
As the spacecraft flew across the solar system, remote-control reprogramming has given the Voyagers greater capabilities than they possessed when they left the Earth. There are four other science instruments that are still functioning and collecting data as part of the Voyager Interstellar Mission. The plasma subsystem measures the protons in the solar wind. "Our instrument has recently observed a slow, year-long increase in the speed of the solar wind which peaked in late 1996, and we are now observing a slow decrease in solar wind velocity," said Dr. John Richardson, of the Massachusetts Institute of Technology, principal investigator on the plasma subsystem. "We think the velocity peak coincided with the recent solar minimum.
As we approach the solar maximum in 2000, the solar wind pressure should decrease, which will result in the termination shock and heliopause moving inward towards the Voyager spacecraft." The magnetometer instrument onboard the Voyagers measures the magnetic fields that are carried out into interplanetary space by the solar wind. The Voyagers are currently measuring the weakest interplanetary magnetic fields ever detected and those magnetic fields being measured are responsive to charged particles that cannot be detected directly by any other instruments on the spacecraft, according to Dr. Norman Ness, principal investigator on the magnetometer subsystem at the Bartol Research Institute, University of Delaware.
Other science instruments still collecting data include the planetary radio astronomy subsystem and the ultraviolet spectrometer subsystem. Voyager 1 encountered Jupiter on March 5, 1979, and Saturn on Nov. 12, 1980, and then, because its trajectory was designed to fly close to Saturn's large moon Titan, Voyager 1's path was bent northward by Saturn's gravity sending the spacecraft out of the ecliptic plane, the plane in which all the planets but Pluto orbit the Sun. Voyager 2 arrived at Jupiter on July 9, 1979, and Saturn on Aug. 25, 1981, and was then sent on to Uranus on Jan. 25, 1986, and Neptune on Aug. 25, 1989.
Neptune's gravity bent Voyager 2's path southward sending it also out of the ecliptic plane and on toward interstellar space. Both spacecraft have enough electrical power and attitude control propellant to continue operating until about 2020 when the available electrical power will no longer support science instrument operation. Spacecraft electrical power is supplied by Radioisotope Thermoelectric Generators (RTGs) that provided approximately 470 watts of power at launch. Due to the natural radioactive decay of the plutonium fuel source, the electrical energy provided by the RTGs is continually declining.
At the beginning of 1997, the power generated by Voyager 1 had dropped to 334 watts and to 336 watts for Voyager 2. Both of these power levels represent better performance than had been predicted before launch. The Voyagers are now so far from home that it takes nine hours for a radio signal traveling at the speed of light to reach the spacecraft. Science data are returned to Earth in real-time to the 34-meter Deep Space Network antennas located in California, Australia and Spain. Voyager 1 will pass the Pioneer 10 spacecraft in January 1998 to become the most distant human-made object in our solar system. Voyager 1 is currently 6.3 billion miles (10.1 billion kilometers) from Earth, having traveled 7.4 billion miles (11.9 billion kilometers) since its launch.
The Voyager 1 spacecraft is departing the solar system at a speed of 39,000 miles per hour (17.4 kilometers per second). Voyager 2 is currently 4.9 billion miles (7.9 billion kilometers) from Earth, having traveled 6.9 billion miles (11.3 billion kilometers) since its launch. The Voyager 2 spacecraft is departing the solar system at a speed of 35,000 miles per hour (15.9 kilometers per second). JPL, a division of the California Institute of Technology, manages the Voyager Interstellar Mission for NASA's Office of Space Science, Washington, DC.
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"Scientists Discover Massive Jet Streams Flowing Inside the Sun (28 August, 1997)",484,0,0,0
Scientists using the joint European Space Agency (ESA)/NASA Solar and Heliospheric Observatory (SOHO) spacecraft have discovered "jet streams" or "rivers" of hot, electrically charged gas called plasma flowing beneath the surface of the Sun. They also found features similar to trade winds that transport gas beneath the Sun's fiery surface. These new findings will help them understand the famous sunspot cycle and associated increases in solar activity that can affect the Earth with power and communications disruptions.
The observations are the latest made by the Solar Oscillations Investigation (SOI) group at Stanford University, Palo Alto, CA, and they build on discoveries by the SOHO science team over the past year. "We have detected motion similar to the weather patterns in the Earth's atmosphere," said Dr. Jesper Schou of Stanford. "Moreover, in what is a completely new discovery, we have found a jet-like flow near the poles. This flow is totally inside the Sun. It is completely unexpected, and cannot be seen at the surface."
"These polar streams are on a small scale, compared to the whole Sun, but they are still immense compared to atmospheric jet streams on the Earth," added Dr. Philip Scherrer, the SOI principal investigator at Stanford. "Ringing the Sun at about 75 degrees latitude, they consist of flattened oval regions about 17,000 miles across where material moves about 10 percent (about 80 mph) faster than its surroundings. Although these are the smallest structures yet observed inside the Sun, each is still large enough to engulf two Earths."
Additionally, there are features similar to the Earth's trade winds on the surface of the Sun. The Sun rotates much faster at the equator than at the poles. However, Stanford researchers Schou and Dr. Alexander G. Kosovichev have found that there are belts in the northern and southern hemispheres where currents flow at different speeds relative to each other. Six of these gaseous bands move slightly faster than the material surrounding them. The solar belts are more than 40 thousand miles across and they contain "winds" that move about ten miles per hour relative to their surroundings.
The first evidence of these belts was found more than a decade ago by Dr. Robert Howard of the Mount Wilson Observatory. The Stanford researchers have now shown that, rather than being superficial surface motion, the belts extend down to a depth of at least 12,000 miles below the Sun's surface. "In one way, the Sun's zonal belts behave more like the colorful banding found on Jupiter than the region of tradewinds on the Earth," said Stanford's Dr. Craig DeForest.
"Somewhat like stripes on a barber pole, they start in the mid-latitudes and gradually move toward the equator during the eleven-year solar cycle. They also appear to have a relationship to sunspot formation as sunspots tend to form at the edges of these zones. "We speculate that the differences in speed of the plasma at the edge of these bands may be connected with the generation of the solar magnetic cycle which, in turn, generates periodic increases in solar activity, but we'll need more observations to see if this is correct," said DeForest.
Finally, the solar physicists have determined that the entire outer layer of the Sun, to a depth of at least 15,000 miles, is slowly but steadily flowing from the equator to the poles. The polar flow rate is relatively slow, about 50 miles per hour, compared to its rotation speed, about 4,000 miles per hour; however, this is fast enough to transport an object from the equator to the pole in a bit more than a year. "Oddly enough, the polar flow moves in the opposite direction from that of the sunspots and the zonal belts, which are moving from higher to lower latitudes," said DeForest.
Evidence for polar flow previously had been observed at the Sun's surface, but scientists did not know how deep the motion extended. With a volume equal to about 4 percent of the total Sun, this feature probably has an important impact on the Sun's activity, argue Stanford researchers Scherrer, with Dr. Thomas L. Duvall Jr., Dr. Richard S. Bogart, and graduate student Peter M. Giles. For the last year, the SOHO spacecraft has been aiming its battery of 12 scientific instruments at the Sun from a position 930,000 miles sunward from the Earth. The Stanford research team has been viewing the Sun's surface with one of these instruments called a Michelson Doppler Imager that can measure the vertical motion of the Sun's surface at one million different points once per minute.
The measurements show the effects of sound waves that permeate the interior. The researchers then apply techniques similar to Earth-based seismology and computer-aided tomography to infer and map the flow patterns and temperature beneath the Sun's roiling surface. "These techniques allow us to peer inside the Sun using sound waves, much like a doctor can look inside a pregnant woman with a sonogram," said Dr. Schou. Currently, the Stanford scientists have both identified new structures in the interior of the Sun and clarified the form of previously discovered ones.
Understanding their relationship to solar activity will require more observations and time for analysis. "At this point, we do not know whether the plasma streams snake around like the jet stream on Earth, or whether it is a less dynamic feature," said Dr. Douglas Gough, of Cambridge University, UK. "It is intriguing to speculate that these streams may affect solar weather like the terrestrial jetstream impacts weather patterns on Earth, but this is completely unclear right now. The same speculation may apply to the other flows we've observed, or they may act in concert. It will be especially helpful to make observations as the Sun enters its next active cycle, expected to peak around the year 2001."
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"Cassini Launch Remains On Schedule (20 August, 1997)",485,0,0,0
The Terminal Countdown Demonstration of the Air Force Titan IV rocket for NASA's Cassini mission has been successfully completed. Today's Terminal Countdown Demonstration was a retest after leaks were repaired on the Centaur upper stage identified during the initial demonstration on Aug. 5. "The success of the Titan test today keeps the launch of Cassini on target for Oct. 6," said Richard Spehalski, Cassini Program Manager. "The processing of the spacecraft here at KSC has gone well and we are also on schedule." The Cassini spacecraft is scheduled for liftoff from Cape Canaveral Air Station, Space Launch Complex 40, on Oct. 6 at 5:38 a.m. EDT. This will begin Cassini's 6.7 year journey to explore the planet Saturn.
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"Mars Pathfinder Results Generating New Picture of Mars (8 August, 1997)",486,0,0,0
NASA's Mars Pathfinder spacecraft -- a novel mission to send an inexpensive lander and roving prospector to the surface of Mars -- has concluded its primary mission, fulfilling all of its objectives and returning a wealth of new information about the red planet. The robotic lander, which continues to explore an ancient outflow channel in Mars' northern hemisphere, completed its milestone 30-day mission on Aug. 3, capturing far more data on the atmosphere, weather and geology of Mars than scientists had expected.
In all, Pathfinder has returned 1.2 gigabits (1.2 billion bits) of data and 9,669 tantalizing pictures of the Martian landscape to date. "The data returned by the Sagan Memorial Station and Sojourner has been nothing short of spectacular, and it will help provide a scientific basis for future Mars missions, including a sample return, for years to come," said Dr. Wesley Huntress, NASA associate administrator for space science.
"The Pathfinder team's "can do" attitude not only was critical to overcoming several complex technical challenges during development and cruise, but has carried through the uncharted territory of operating a solar- powered lander and mobile rover on the surface of a planet millions of miles from Earth." "This mission demonstrated a reliable and low-cost system for placing science payloads on the surface of Mars," said Brian Muirhead, Mars Pathfinder project manager at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "We've validated NASA's commitment to low-cost planetary exploration, shown the usefulness of sending microrovers to explore Mars, and obtained significant science data."
A new portrait of the Martian environment has begun to emerge in the 30 days since Pathfinder and its small, 23-pound rover began to record weather patterns, atmospheric opacity and the chemical composition of rocks washed down into the Ares Vallis flood plain. The rover's alpha proton X-ray spectrometer team, led by principal investigator Dr. Rudolph Rieder, has been able to analyze the first-ever in- situ measurements of Mars rocks. "We are seeing much more differentiation of volcanic materials than we expected to see," said Dr. Matthew Golombek, Mars Pathfinder project scientist at JPL.
"The high silica content of one of the rocks we've measured, nicknamed Barnacle Bill, suggests that there was more crustal activity -- heating and recycling of materials -- early in Mars' history than we thought." Similarly, atmospheric-surface interactions, measured by a meteorology package onboard the lander, are confirming some conditions observed by the Viking landers 21 years ago, while raising questions about other aspects of the planet's global system of transporting volatiles such as water vapor, clouds and dust, said science team leader Dr. Timothy Schofield.
The meteorology mast on the lander has observed a rapid drop-off in temperatures just a few feet above the surface, and one detailed 24-hour measurement set revealed temperature fluctuations of 30-40 degrees Fahrenheit in a matter of minutes. In addition, sweeping, color panoramas of the Martian landscape, created by the Imager for Mars Pathfinder team and principal investigator Peter Smith, are revealing clear evidence that the surface of Mars has been altered by winds and flowing water. Sojourner, a robust rover capable of semi-autonomous "behaviors," captured the imagination of the public, which followed the mission with great interest via the World Wide Web.
Twenty Pathfinder mirror sites, constructed by JPL web engineer Kirk Goodall and managed by Pathfinder webmaster David Dubov, recorded 565 million hits worldwide during the period of July 1 -- August 4. The highest volume of hits in one day occurred on July 8, when a record 47 million hits were logged, which is more than twice the volume of hits received on any one day during the 1996 Olympic Games in Atlanta.
The rover's performance has easily surpassed its designers' minimum expectations. Engineers designed the roving vehicle's electronics, battery power and hazard avoidance features to see it through at least a week of safe roving, not knowing beforehand what conditions it might encounter on Mars. After 30 days, the rover is still healthy and has traveled 171 feet in distance, circumnavigating the lander and taking 384 spectacular views of rocks and the lander.
"Sojourner's capabilities to detect hazards and then act on its own to overcome those hazards have been remarkable," said Dr. Jacob Matijevic, Sojourner project manager. "The technology experiments we have been able to perform with the rover's wheels have given us more information about the composition of the Martian soil, as well as rocks around the landing site.
Sojourner's durability in this frigid, hostile environment also is showing us that we are on the right track to building smarter, even more durable rovers for future missions." Pathfinder's primary objective was to demonstrate a low-cost way of delivering an instrumented lander and free-ranging rover to the surface of the red planet. Landers and rovers of the future will share the heritage of spacecraft designs and technologies tested in this "pathfinding" mission.
Part of NASA's Discovery program of low-cost planetary missions with highly focused science goals, the spacecraft used an innovative method of directly entering the Martian atmosphere. Assisted by a 36-foot-diameter parachute, the spacecraft descended to the surface of Mars and landed, using airbags to cushion the impact. This novel method of diving into the Martian atmosphere worked like a charm. "Every event during the entry, descent and landing went almost perfectly," said Richard Cook, Pathfinder mission manager.
"The sequences were executed right on time and well within our margins." Pathfinder landed right on the money, within 13 miles of the targeted landing site. The landing site coordinates in Ares Vallis were later identified as 19.33 degrees North latitude, 33.55 degrees West longitude. The spacecraft's terminal velocity as it parachuted to the ground was higher than expected, said Rob Manning, Pathfinder flight system chief engineer. "Interestingly, we estimated our descent on the parachute at about 134 miles per hour.
Software controlling the retro rockets recorded Pathfinder's speed at about 140 miles per hour at the time the rocket-assisted deceleration rockets fired." Pathfinder's performance in the Martian atmosphere will be of great value to Mars Global Surveyor, which will aerobrake through the Martian atmosphere to circularize its orbit when it reaches Mars on September 11. The Pathfinder navigation team, led by Pieter Kallemyn of JPL, estimated that horizontal wind velocities in the upper atmosphere helped accelerate the spacecraft's descent velocity by about 20 to 25 miles per hour.
After being suspended from a 65-foot bridle and firing its retro rockets, a 19-foot diameter cluster of airbags softened Pathfinder's landing, marking the first time this airbag technique has been used on another planet. The spacecraft hit the ground at a speed of about 40 miles per hour and bounced about 16 times across the landscape for about six-tenths of a mile before coming to a halt. The airbag seems to have performed perfectly and sustained little or no damage. To top it off, the spacecraft even landed on its base petal, consequently allowing its thumb-sized antenna to communicate the successful landing to a jubilant team on Earth only three minutes after touchdown.
Science data from the surface of Mars will continue to be collected and transmitted to Earth, then analyzed by scientists, as Pathfinder enters its extended mission. The lander was placed in a two-day hibernation period earlier this week to recharge its battery after the conclusion of the primary mission, and the flight team now will begin to power the lander battery off each Martian night to conserve energy. The rover's batteries remain in good condition, but are not rechargeable. The Mars Pathfinder mission is managed by the Jet Propulsion Laboratory for NASA's Office of Space Science, Washington, DC. JPL is a division of the California Institute of Technology, Pasadena, CA.
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"Countdown Test Reveals Fuel Leaks On Cassini Mission Centaur Upper Stage (7 August, 1997)",487,0,0,0
During the Tuesday, August 5, terminal countdown demonstration test, Air Force and Lockheed Martin engineers observed leakage in the Centaur stage of the Titan IV-B rocket for the Cassini mission to Saturn. This test, in which the Centaur is fully fueled, is normally conducted to identify problems which could affect the performance of the Titan IV.
Leakage of this nature can occur on occasion when the Centaur is first tanked with cryogenic propellants. During this test, engineers observed some liquid hydrogen and liquid oxygen leakage in the thrust section. Engineering assessments are currently being performed to determine the cause of the leakage and what corrective action is necessary to ready the vehicle for the Cassini launch. Until this has been done, what impact this might have on the planned October 6 launch date, if any, cannot be definitely determined. A repeat test will be performed to assure that there are no additional leaks or other issues.
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"Hubble Separates Stars in the Mira Binary System (6 August, 1997)",488,0,0,0
Although the giant star Mira has been known for 400 years, astronomers have had to wait for NASA's Hubble Space Telescope to provide the first ultraviolet images of the extended atmosphere of the cool red giant star and its nearby hot companion. By giving astronomers a clear view of the individual members of this system, Hubble has provided valuable insights into other types of double star systems where the stars are so close they interact with one another.
The separation between Mira and its companion is about 70 times more than that between Earth and the Sun, (equal to an angular size of only 0.6 arcseconds -- the apparent diameter of a dime at four miles away) even smaller than the typically fuzzy ground-based telescopic image of a single star as smeared out by Earth's turbulent atmosphere. Using the European Space Agency's Faint Object Camera aboard Hubble, Margarita Karovska and John Raymond of the Harvard- Smithsonian Center for Astrophysics, Cambridge, MA; Warren Hack of the Space Telescope Science Institute, Baltimore, MD; and Edward Guinan of Villanova University, Villanova, PA, obtained both ultraviolet and visible light images and spectra of the two separate stars in the Mira system.
The results appear in the June 20 Astrophysical Journal Letters. In ultraviolet light, Hubble has resolved a small hook-like appendage extending from Mira in the direction of the companion, which might be material from Mira being gravitationally drawn toward the smaller star. Alternately, it could be material in Mira's upper atmosphere being heated due to the companion's presence. Hubble's visible-light images show that Mira has an odd, asymmetrical shape resembling a football. This may be tied to dramatic changes occurring during its expansion-contraction cycles, or to the presence of unresolved spots on its surface.
Hubble allows astronomers to measure the star's size at about 60 milliarcseconds, corresponding to a diameter some 700 times larger than our Sun. If Mira were at the center of our solar system, it would extend out more than 300 million miles, well beyond Mars' orbit and nearly two-thirds of the way to Jupiter. Mira (officially called Omicron Ceti in the constellation Cetus) is the prototype for an entire class of stars known as "Mira-type variables." Although once like our Sun, Mira is now at the end of its life, and has evolved into a cool red giant star that is highly variable in brightness. Contracting and expanding every 332 days, Mira sheds vast amounts of material through its powerful "wind" of gas and dust.
Mira's companion is a burned-out star called a white dwarf that is surrounded by material captured from Mira's wind. At a distance of about 400 light-years, Mira is the closest wind-accreting binary system to Earth. Separating the spectra of Mira and its companion -- something astronomers previously have tried to do through indirect means -- is a crucial step for studies of physical processes associated with wind accretion in binaries. Mira was discovered on August 13, 1596, by Dutch astronomer David Fabricus, who mistook it for a nova because it later faded from view. He called it Mira, meaning "The Wonderful." Astronomers later realized it was really the first case of a variable star.
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"Remote Control Robot Breaks Rough Terrain Travel Record (5 August, 1997)",489,0,0,0
A hardy traveler named "Nomad" recently set a record by traveling farther than any remotely controlled robot has before over rough territory. The robot's four wheels logged more than 133 miles (215 kilometers) across Chile's rugged Atacama Desert from June 15 to July 31, during a field experiment designed to prepare for future missions to Antarctica, the Moon and Mars. Scientists from NASA's Ames Research Center, Moffett Field, CA, and Carnegie Mellon University's Robotics Institute in Pittsburgh performed experiments with Nomad for 45 days, conducting both technology demonstrations and scientific activities.
Nomad often worked on its own to avoid obstacles and, in a clear foreshadowing of the future duties of similar robots, it recognized meteorites planted in the desert as a test and may even have found a fossil. "The Atacama trek is a quantum leap for the planetary robotics culture, where the historical standard of travel has been yards, not miles," said principal investigator Dr. William L. "Red" Whittaker of Carnegie Mellon. "Although the 'straight-line' distance on a map was only about 13 miles, Nomad had to weave through very difficult terrain, and it made numerous sidetrips for science and to test the meteorite sensors.
It is a pioneer laying a trail toward future planetary robots, who will be challenged for thousands of miles and years of operations, in bold missions like searching for signs of life." The 1,600-pound robot, developed at Carnegie Mellon and funded by NASA, validated the use of color stereo video cameras with human-eye resolution for geology. A separate panospheric camera returned more than a million video panoramas from the Atacama, a cold, arid region located above 7,000 feet.
"During different phases of testing, we configured the robot to simulate wide-area exploration of the Moon, the search for past life on Mars and for the gathering of meteorite samples in the Antarctic," said Dave Lavery, telerobotics program manager at NASA Headquarters, Washington, DC. "Nomad met or exceeded all of our objectives for this project." "We want to give planetary scientists experience using mobile robots, so that they can develop the skills necessary for performing remotely guided investigations," added Dr. David Wettergreen, Nomad project leader at Ames.
Nomad is about the size of a small car. To maneuver through rough terrain, the robot has four-wheel drive and four-wheel steering with a chassis that expands to improve stability and travel over various terrain conditions. Four aluminium wheels with cleats provide traction in soft sand. For this terrestrial experiment, power was supplied by a gasoline generator that enabled the robot to travel at speeds up to about one mile per hour. "Nomad drove itself through about 12 miles (20 kilometers) of the 133 miles it traveled," said Dr. Mark Maimone, Nomad software and navigation lead at Carnegie Mellon. "Autonomous driving is critical for planetary exploration because the communications delay between Earth and planets can be many minutes.
With autonomous driving, a robot can explore a much greater distance because it doesn't have to wait for a person to decide a safe route. The rover is able to see obstacles and recognize them on its own," he said. Nomad's unique onboard panospheric camera provided live 360- degree, video-based still images of the robot's surroundings. "Experimentation with the panospheric camera validated the use of immersive imagery for remote driving," Maimone said. The camera takes a 360-degree picture -- one frame per second -- and did so throughout the mission. The high-resolution video camera focuses up into a hemispheric mirror similar to a store security mirror.
The video view includes all of the ground up to the horizon in the circle surrounding Nomad. "The camera is a new technology, and it gave members of the public as well as scientists a new way to drive with peripheral, or side vision," he explained. "We sent the Nomad pictures to a theater at the Carnegie Science Center in Pittsburgh that has a 200-degree, semi-circular screen. Fifty people at a time pushed a button to vote on whether the robot should look to the left, center or right." On June 25, NASA scientists were driving the robot remotely from their laboratory at Ames, more than 5,455 miles (8,780 kilometers) away, when the scientists in California found a rock that appeared to contain algae fossils.
Using the rover's cameras, scientists noticed a light- colored, three-inch diameter rock with a darker, intricately shaped marking in a rock outcrop in the Chilean desert. The rock was retrieved by Chilean scientists and was brought to Ames for scientific analysis. "The rock is sedimentary and was formed in an ancient sea bed. However, the consensus is that this rock does not contain fossilized algae," said Dr. Nathalie Cabrol, the expedition's NASA science team leader. The science team was excited to learn that the outcrop was an undiscovered geologic deposit from the Jurassic Period. "This experience is one of the most important of the science tests," Cabrol said.
"I am not sure that we can get much closer to what may happen with the research of interesting rocks on Mars and the related search for life in the coming Mars exploration program. We are most likely to face this exact situation of selecting a rock because it looks interesting to us. Once in the lab, we were unable to tell conclusively if there had been life in the rock at one time or not." "The first-level interpretation from the rover camera was close enough, fossil or not," she added. "The team was able to reconstruct the geology of the site, often matching or at least getting very close to the conclusions of the back-up field team."
The total cost of developing Nomad and conducting the desert trek is $1.6 million. The project is funded by NASA with in-kind support from corporate sponsors and educational foundations. NASA and Carnegie Mellon are formulating plans to use Nomad to look for meteorites in Antarctica in 1998 and 1999.
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"World's Most Powerful Telescopes To Discover Farthest Galaxy in the Universe (30 July, 1997)",490,0,0,0
An international team of astronomers has discovered the most distant galaxy found in the universe to date, by combining the unique sharpness of the images from NASA's Hubble Space Telescope with the light-collecting power of the W. M. Keck Telescopes -- with an added boost from a gravitational lens in space. The results show the young galaxy is as far as 13 billion light years from us, based on an estimated age for the universe of approximately 14 billion years.
This would place the galaxy far back in time during the "formative years" of galaxy birth and evolution, less than a billion years after the birth of the universe in the Big Bang. The detailed image shows that bright dense knots of massive stars power this object. Due to the firestorm of starbirth within it, the galaxy is intrinsically one of the brightest young galaxies in the universe, blazing with the brilliance of more than ten times our own Milky Way.
"We are fascinated to be witnessing the very early stages of the construction of what could well become a massive galaxy like our own Milky Way," says Garth Illingworth of the University of California, Santa Cruz. "This object is a pathfinder for deciphering what is happening in young galaxies, and offers a rare glimpse of the powerful events that transpired during the formation of galaxies." "We were excited by the possibility that we may have found a unique example of a galaxy in formation at the time of the earliest quasars," said Marijn Franx of the University of Groningen in the Netherlands.
Predicted by Einstein's theory of general relativity, gravitational lenses are collections of matter (such as clusters of galaxies) that are so massive they warp space in their vicinity, allowing the light of even more-distant objects to curve around the central lens-mass and be seen from Earth as greatly magnified. The object is so far away, observing it in such detail would tax the capabilities of both Hubble and Keck without the magnification of the gravitational lens, provided by a foreground cluster of galaxies that is much closer to us at five billion light-years.
Due to a rare and fortunate alignment of the young galaxy behind the foreground cluster, astronomers gain a magnified view that is five to ten times better than Hubble alone can yield for an object at such a great distance. A telltale sign of the lensing is the smearing of the remote galaxy's image into an arc-shape by the gravitational influence of the intervening galaxy cluster . The smeared image of the galaxy stood out because of its unusual reddish color.
"Such magnified galaxies had been observed before, but never with such a color. The special color of the galaxy in the arc is due to absorption by the matter in the universe between us and the galaxy, and suggested to us that it was at a great distance," says Franx. The suspected remoteness of the lensed object was confirmed when the team of astronomers made spectroscopic observations with one of the twin 10-meter Keck telescopes on Mauna Kea, HI, to measure its redshift, and therefore its distance, based on the shifting of its light towards the red end of the visible light spectrum.
The resulting high redshift corresponds to a very early era when the universe was just beginning to form galaxies. Though candidates for still more distant objects have been proposed, they have not been confirmed spectroscopically. The previous most-distant known object was the quasar PC1247+34. "Based on this image we can begin to make some conclusions about the early growth of galaxies," says Illingworth. "The knots show that starbirth happens in very tiny regions compared with the size of the final galaxy." This helps clarify the astronomer's view of the formation of galaxies as occurring within a cauldron of hot gas, with knots of intense star formation, strong winds, and "mergers" -- collisions of the dense star-forming knots.
Using Keck's spectroscopic capabilities, the astronomers have also, for the first time, been able to measure the motions of the gas within such a distant galaxy. The observations reveal gas flowing at nearly 500,000 miles per hour (200 km/sec), presumably accelerated by energy from supernova explosions going off like a string of firecrackers. "The strong winds that we observe suggest that galaxies may lose a lot of material when they are young and thereby enrich the empty space around them," says Franx. "Many astronomers had speculated about the existence of such winds in such distant galaxies, and we now have an object where we can see them directly. It is striking that the most distant galaxy found to date is also the one that provides us the most detailed picture of events in such distant galaxies."
The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc. (AURA) for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency (ESA). The W.M. Keck observatory is operated by the University of California, the California Institute of Technology and NASA.
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"Advanced Composition Explorer Set To Study Matter From Sun, Milky Way and Beyond (23 July, 1997)",491,0,0,0
The Earth is constantly being bombarded by a stream of accelerated particles arriving not only from the Sun, but also from interstellar and galactic sources. The study of these energetic particles by NASA's Advanced Composition Explorer (ACE) observatory will contribute to the understanding of the formation and evolution of the solar system as well as the astrophysical processes involved. The space science observatory is scheduled for launch on a Delta II rocket at 10:39 a.m. EDT August 25, from Launch Complex 17 at the Cape Canaveral Air Station, FL.
"The Advanced Composition Explorer observatory is designed to sample the matter that comes near the Earth from the Sun, from the apparently, but not actually, empty space between the planets, and from the Milky Way galaxy beyond the solar system," said Don Margolies, ACE Mission Manager at NASA's Goddard Space Flight Center, Greenbelt, MD. "While previous missions have studied these particles, the instruments on ACE have a collecting power 10 to 1,000 times greater and will be at least 100 times more sensitive than anything we've ever flown," Margolies said.
"We will be able to study known phenomena in much greater detail than previously possible, and discover new ones to give us a better understanding of the interaction between the Sun, the Earth, and the galaxy." The Advanced Composition Explorer has six high-resolution particle detection sensors and three monitoring instruments. It will sample low-energy particles of solar origin and high-energy galactic particles.
The observatory will be placed into an orbit at the L1 libration point, which is almost a million miles (or 1.5 million kilometers) away from the Earth, about 1/100 the distance from the Earth to the Sun. The ACE payload includes four brand-new, state-of-the-art spectrometers. They are the Cosmic Ray Isotope Spectrometer; Solar Isotope Spectrometer; Solar Energetic Particle Ionic Charge Analyzer; Ultra-Low-Energy Isotope Spectrometer. In addition, there are four spare instruments from other NASA missions that, with appropriate modifications, are being flown on ACE. They are the Solar Wind Electron, Proton, and Alpha Monitor; Solar Wind Ionic Charge Spectrometer; Electron, Proton, and Alpha Monitor; and a Magnetometer.
An additional instrument, the Solar Wind Ion Mass Spectrometer, is a newly built copy of a previously flown instrument. Also onboard are two secondary instruments, the Spacecraft Loads and Acoustic Measurements, designed to measure spacecraft environments during the first five minutes of launch, and the Real Time Solar Wind experiment, which will provide real-time data to the National Oceanic and Atmospheric Administration (NOAA). NASA and nine universities in the U.S. and Europe built the instruments. The ACE spacecraft's instruments and experiments will work together to add to our understanding of solar events ranging from "solar storms" to the origin and evolution of solar and galactic matter.
The scientific goal of the ACE mission is to measure accurately the composition of several different types of matter, including particles coming from the Sun, the very thin gas between the planets, the even thinner gas just outside the solar system, and matter from distant parts of the galaxy. The particles that ACE measures are moving very fast, up to 3.5 million miles per hour, and are atomic and subatomic. ACE also has an Earth applications goal. It will provide NOAA's Space Environment Center, Boulder, CO, with continuous real-time solar wind "space weather" information, which will give an advance warning (about one hour) of geomagnetic storms that can affect electric power grids, Earth-orbiting spacecraft, and radio communications on Earth.
The 1,730-pound observatory was built by the Johns Hopkins Applied Physics Laboratory, Laurel, MD, where its instruments were integrated. ACE was tested at the Applied Physics Laboratory and at Goddard. The science payload was provided under the direction of the Payload Management Office at the California Institute of Technology, Pasadena, CA. Flight operations will be conducted from Goddard; the Science Center is located at Caltech. The ACE mission is managed by the Explorer Program at Goddard for the Sun-Earth Connections Program in the Office of Space Science, NASA Headquarters, Washington, DC.
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"NASA Statement on the Passing of Gene Shoemaker (19 July, 1997)",492,0,0,0
Planetary scientist Dr. Eugene ("Gene") Shoemaker, 69, was killed in a two-car accident near Alice Springs, Australia, on the afternoon of July 18. His wife Carolyn Shoemaker suffered broken bones, and reportedly is hospitalized in stable condition. A geologist by training, Shoemaker is best known for discovering, with his wife Carolyn and colleague David Levy, a comet near Jupiter.
Comet Shoemaker-Levy 9 was broken up by tidal forces from Jupiter, and its fragments collided with the planet in July 1994. Together, the Shoemakers were the leading discoverers of comets this century. "Gene was one of the most renowned planetary scientists in the world, and a valued member of the NASA family since the earliest days of lunar exploration," said NASA Administrator Daniel S. Goldin. "His work on the history of meteor impacts and the role that they play in the evolution of the Solar System is a fundamental milestone in the history of space science.
"Gene was an extremely articulate man who could explain the wonders of the planets in simple language that anyone could understand and get excited about," Goldin added. "Although he never realized his dream of doing field geology on the surface of the Moon, all future exploration of that rocky world owes a debt to his pioneering spirit. Our warmest thoughts are with his dear wife Carolyn as she recovers from her injuries."
Shoemaker's signature work was his research on the nature and origin of the Barringer Meteor Crater near Winslow, AZ, which helped provide a foundation for cratering research on the Moon and planets. This work led to the establishment of a lunar chronology, allowing the dating of geological features of its surface. Shoemaker took part in the Ranger lunar robotic missions, was principal investigator for the television experiment on the Surveyor lunar landers (1963-1968), and led the geology field investigations team for the first Apollo lunar landings (1965-1970).
In 1961, he organized the Branch of Astrogeology of the U.S. Geological Survey in Flagstaff, AZ, and acted as its director from 1961 to 1966. On his retirement from the U.S.G.S. in 1993, Shoemaker became a staff member at Lowell Observatory in Flagstaff. An early supporter of the idea that an asteroid or comet impact had doomed much of Earth's life (including the dinosaurs) 65 million years ago, Shoemaker chaired key NASA working groups on how best to survey such near-Earth objects in 1981 and 1994. Most recently, he was active in the Clementine mission that imaged the Moon, and was science team leader on the planned Clementine 2 mission. Shoemaker won numerous awards during his career, and in 1980 became a member of the National Academy of Sciences.
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"STS-94 Going Smoothly (2 July 1997)",493,0,0,0
Voss and Crouch in Spacelab
The astronauts on board the Space Shuttle Columbia have begun the operation of the nearly three dozen experiments that make up the science payload of this \JMicrogravity\j Science Laboratory mission, while setting up housekeeping on the shuttle for a planned 16 days on orbit.
This morning Columbia's orbit brought the space shuttle to within about 60 miles of the Space Station Mir, and the astronauts reported getting a very good view of that orbiting Russian outpost.
The astronauts on board Columbia are less than one day into their planned 16-day mission, but they're already returning science data to the Spacelab Mission Operations Control Center at the Marshall Spaceflight Center. Mission managers say the orbiter and its systems are operating in good shape with no issues being worked.
Commander Jim Halsell and pilot Susan Still concluded their first full work day on orbit, seeing to the proper operation of the shuttle systems. Mission specialist Don Thomas and payload specialist Greg Linteris spent their workday in the Spacelab module.
Susan Still has set up the shuttle's ham radio equipment on the flight deck. It will be used for shuttle amateur radio experiment contacts with radio operators around the world throughout the mission.
The blue team astronauts-payload commander Janice Voss, mission specialist Mike Gernhardt, and payload specialist Roger Crouch --took over from the red team shortly after noon. About an hour later payload commander Janice Voss and payload specialist Roger Crouch began their day's experiment work in the Spacelab module.
At 3:30 p.m. CDT the four red team astronauts will start their sleep period, which will last until about 11:30 p.m. CDT when they will get their wake-up call from Mission Control. At 1:02 a.m. tomorrow morning they'll begin taking a handover from the blue team; Voss, Gernhardt, and Crouch will be sent to bed an hour later, and shortly thereafter payload specialist Greg Linteris will begin his day's work in the Spacelab module. Reveille will sound for the astronauts of the blue team shortly after 10:00 a.m. CDT tomorrow.
\BMars Pathfinder\b
On July 4, during the STS-94 mission, the Mars Pathfinder will land on Mars. Renewed interest in the exploration of Mars was sparked by the suggestion last year that evidence of Life on Mars may have been found in a Martian \Jmeteorite\j.
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"NASA Adjusts To Loss Of Data From Japanese Adeos Satellite (2 July, 1997)",494,0,0,0
"The failure of \JJapan\j's Advanced Earth Observing Satellite (ADEOS or Midori) \Jspacecraft\j with the two NASA instruments aboard it is a real blow to NASA's science program," said Mike Mann, Deputy Associate Administrator, NASA's Mission to \JPlanet\j Earth Strategic Enterprise, Washington, DC.
"Fortunately, much of the ozone data provided by the Total Ozone Mapping Spectrometer (TOMS) science instruments aboard ADEOS can be provided by instruments on another \Jspacecraft\j. However, the sea-surface winds data provided by the NASA Scatterometer (NSCAT) will be harder to replace and were opening essentially new opportunities for research and operational users worldwide," Mann said.
The two NASA instruments were aboard the ADEOS \Jspacecraft\j, which on June 30 was declared lost by the National Space Development Agency of \JJapan\j (NASDA).
"The collaboration between NASDA and NASA on this mission has been outstanding and is reflective of the great partnership that exists between \JJapan\j and the U.S in the area of global change research," Mann said.
"NASDA has performed in an exemplary and open manner in the development of the \Jspacecraft\j and in dealing with us. However, space operations is a risky business; those of us involved in the business strive to limit the risk but sometimes mishaps do occur," Mann said.
"The data we have obtained to date are extremely valuable," said Jim Graf, NSCAT project manager at NASA's Jet Propulsion Laboratory, Pasadena, CA. "If we knew we were limited to just nine months of data, we would have chosen the period we actually got. We obtained coverage over the summer and winter monsoon seasons and what may be the onset of an El Nino. Perhaps the largest loss is the discontinuity of the long-term data set, which is being used to understand interannual and decadal variations in our climate."
The scatterometer measured wind speed and direction over the world's oceans. The data set is extremely valuable and versatile and is being used by climate change researchers, operational weather forecasters, and commercial ship routing firms. During its flight, the instrument gathered 42 weeks' worth of data.
Within a very few short months after launch, the value of ADEOS data was seen in U.S. weather forecasting. "NOAA had begun using ocean surface wind products, derived from NSCAT, in weather forecasting," said Helen Wood, Director, Office of Satellite Data Processing and Distribution, National Oceanic and Atmospheric Administration. "Ocean surface wind measurements are used in numerical weather prediction models and help forecasters more accurately determine the path and intensity of tropical storms and hurricanes."
Because this instrument provided measurements that will be needed over the long term, NASA was already developing a second scatterometer instrument to continue this vital data set. That instrument, called "SeaWinds," will be delivered to NASDA for \Jintegration\j on the \Jspacecraft\j next April and is scheduled for launch in 1999 on ADEOS II.
The launch of a Total Ozone Mapping Spectrometer sensor aboard ADEOS was helping to extend the unique data set of global total column ozone measurements begun by a similar instrument carried aboard NASA's Nimbus-7 satellite in 1978 and extended until December 1994 with the Meteor-3 TOMS.
"The ADEOS spectrometer, along with the TOMS Earth Probe (EP) instruments also observed the unusual loss of \JArctic\j polar ozone reported earlier this year," said Dr. Arlin J. Krueger, Principal Investigator and Instrument Scientist for TOMS/ADEOS at NASA's Goddard Space Flight Center, Greenbelt, MD.
Although it also provided ozone coverage, NASA's Total Ozone Mapping Spectrometer/Earth Probe instrument had also been providing high ground resolution research data to complement the global data of the spectrometer on ADEOS. As a result, its orbit is different than TOMS/ADEOS. The EP satellite has adequate fuel to raise its present 500 km orbit to an orbit near the 800 km ADEOS orbit, where contiguous Earth coverage is possible for monitoring of ozone and volcanic eruption clouds. NASA is considering raising TOMS/EP to a higher orbit.
With this adjustment, much more complete global coverage of total ozone measurements previously provided by TOMS/ADEOS could be received. However, some of the unique smaller-scale aerosols and ozone research being done by TOMS/EP would be lost. The next Total Ozone Mapping Spectrometer mission is planned for launch on a Russian Meteor-3M \Jspacecraft\j in 2000.
The loss of the ADEOS platform has a particularly serious impact on oceanographic research since two instruments, the Ocean Color and Temperature Sensor and the Polarization and Directionality of the Earth's \JReflectance\j, both capable of providing routine global estimates of \Jphytoplankton\j pigment concentrations, were lost. These instruments were providing the first routine global observations of ocean color and were initiating the much-needed, long-term time series of such measurements for global change studies.
Future routine global ocean-color information will be provided by SeaWIFS, a commercial mission from which NASA will purchase data, currently scheduled for launch July 18.
The NASA Scatterometer and Total Ozone Mapping Spectrometer/ADEOS were developed under NASA's strategic enterprise called Mission to \JPlanet\j Earth, a comprehensive research effort to study Earth's land, oceans, atmosphere, ice and life as an interrelated system.
NASA is cooperating with NASDA to identify the cause of the ADEOS failure and recommend a solution for future missions.
#
"Asteroid Mathilde Reveals Her Dark Past (30 June, 1997)",495,0,0,0
More than 100 years after her discovery, asteroid 253 Mathilde has been sharing her secrets with scientists in the Science Data Center at the Johns Hopkins University Applied Physics Laboratory in Laurel, MD. A 25-minute flyby of the asteroid by NASA's Near Earth Asteroid Rendezvous (NEAR) \Jspacecraft\j on June 27 has resulted in spectacular images of a dark, crater-battered little world assumed to date from the beginning of the solar system.
The Mathilde flyby is the closest encounter with an asteroid to date and the first with a C-type asteroid. The asteroid's mean diameter was found to be 33 miles (52 kilometers), which is somewhat smaller than researchers originally estimated.A study of the asteroid's \Jalbedo\j (brightness or reflective power) shows that it reflects three percent of the Sun's light, making it twice as dark as a chunk of charcoal. Such a dark surface is believed to consist of carbon-rich material that has not been altered by planet-building processes, which melt and mix up the solar system's original building block materials.
The Mathilde flyby met all its initial goals: getting a clear image of the sunlit side of the asteroid, getting color images that will give clues to the types of rock that make up the asteroid, and getting images that will help researchers determine if Mathilde has any moons. In the next month, scientists expect to complete initial analysis of their data and have improved measurements of Mathilde's volume, mass, and density.
"The Mathilde encounter was one of the most successful flybys of all time," said Dr. Robert W. Farquhar, of the Applied Physics Laboratory, NEAR Mission Director. "We got images that were far better than we thought possible, especially since the \Jspacecraft\j was not designed for a fast flyby."
Only the multispectral imager, one of six instruments on the \Jspacecraft\j, was used during the flyby in order to conserve power provided by solar-powered panels. The \Jspacecraft\j was approximately 186 million miles from the Sun, too far to provide power for NEAR's other instruments.
"Even though this was a very difficult undertaking," said Dr. Stamatios M. Krimigis, head of the APL Space Department that managed the program for NASA, "the NEAR Operations Team was so well prepared there was little doubt that it would succeed; not only that, but this was the smallest operations team of any planetary encounter, proving that the Discovery Program paradigm of 'smaller, faster, cheaper' is alive and well."
Although Mathilde proved to be rounder than \Jasteroids\j such as Gaspra and Ida, Dr. Joseph Veverka of Cornell University, Ithaca, NY, who leads the mission's imaging science team, said, "Mathilde turned out to be more irregularly shaped than most of us expected. The degree to which the asteroid has been battered by collisions is astounding. At first glance there are more huge craters than there is asteroid."
The imager found at least five craters larger than 12 miles (20 kilometers) in diameter just on the lighted side of the asteroid. Scientists wonder how the asteroid can remain intact after having been hit by this many projectiles, each probably at least a mile wide.
The craters reveal evidence of the asteroid's makeup. "We knew that C-asteroids are black, but we did not expect their surfaces to be as uniformly black and colorless as Mathilde's surface turned out to be," Veverka said. "This global blandness is an important clue telling us that \Jasteroids\j such as Mathilde are made of the same dark, black rock throughout because none of the craters, which are punched deep into the asteroid, show evidence of any other kind of rock." Such uniformity seems to confirm that C-type \Jasteroids\j are in fact pristine samples of the primitive building blocks of the larger planets.
Dr. Donald K. Yeomans of the Jet Propulsion Laboratory, Pasadena, CA, who heads the radio science team formed to determine Mathilde's mass said, "Mathilde is an asteroid with a very tortured past." By determining the bulk density of the asteroid, researchers will have a clue to how it was formed. A composite of objects would have a lower density than a solid chunk from a larger asteroid. Data analysis to determine density will not be complete until later this year, but Dr. Yeomans said, "Preliminary results suggest that Mathilde is much less dense than we had thought."
One mystery that remains is Mathilde's extraordinarily slow (17.4 days) rotation rate. Its collision history could be a factor, but more research needs to be done to determine what role such collisions have played. The search for Mathilde moons continues; none has yet been discovered.
The next major event of the NEAR mission will occur on July 3, when the \Jspacecraft\j's bi-propellant engine is fired to head NEAR back toward Earth. This deep-space maneuver will be the first time the engine has been fired and will keep both engineers and scientists in suspense for 11 minutes before they know if the maneuver was successful. An Earth gravity-assist maneuver on Jan. 23, 1998, will send the \Jspacecraft\j toward its primary target, asteroid 433 Eros. NEAR will reach Eros nearly a year later and will remain locked in orbit around the asteroid until Feb. 6, 2000, when the mission ends.
Commenting on the success of the Mathilde flyby soon after the first images were received, Dr. Wesley T. Huntress Jr., NASA Associate Administrator, Office of Space Science, said, "It's today that the Discovery Program really begins. NEAR was the first of our Discovery missions to be launched and it's the first to return scientific results." He said the APL-led team that managed the NEAR program proved the concept behind the Discovery Program:that exciting planetary missions can be done at low cost, in a short time.
The NEAR \Jspacecraft\j was launched Feb. 17, 1996, from Cape Canaveral Air Station in \JFlorida\j. NEAR Science Team Group Leaders are: Joseph Veverka, Cornell University; Jacob I. Trombka, NASA/Goddard Space Flight Center, Greenbelt, MD; Mario H. Acuna, NASA/Goddard; Maria T. Zuber, MIT and NASA/Goddard; and Donald K. Yeomans, NASA/Jet Propulsion Laboratory, Pasadena, CA. Andrew Cheng, JHU/APL, is the Project Scientist. The Johns Hopkins University Applied Physics Laboratory operates the mission for NASA's Office of Space Science, Headquarters, Washington, DC.
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"U.S. Space Station (First Component) Begins Launch Preparations (26 June, 1997)",496,0,0,0
The Internatonal Space Station Program passed a major milestone this week as the first U.S.-manufactured component began a year of launch preparations at the Kennedy Space Center, FL.
A connecting module, called Node 1, was shipped by cargo \Jaircraft\j to \JFlorida\j on Sunday from the Marshall Space Flight Center╣s Space Station Manufacturing Facility in \JHuntsville\j, AL. The node will be the first U.S-built segment for the station to reach orbit when it is launched in July 1998 aboard the Space Shuttle Endeavour on the STS-88 mission.
"The International Space Station has begun moving from the factory floor to the launch pad," program manager Randy Brinkley said. "By the time Node 1 is launched next year, pieces of the station will be leaving factories in locations worldwide to be readied for launch, and the first piece already will be in orbit. From now through the turn of the century, the processing of station components will be a major focus at the Kennedy Space Center."
The crew of Endeavour will use the Shuttle's robotic arm to dock Node 1 with the Functional Cargo Block as the node sits atop the orbiter docking system in the Shuttle's cargo bay. The Functional Cargo Block is a component that supplies early power and propulsion systems for the station. It will be the first element to be placed in orbit and will be launched two weeks before the STS-88 mission on a Russian Proton rocket from the Baikonur Cosmodrome in \JKazakhstan\j. After the two components are linked together, three spacewalks will be performed from the Shuttle to connect power, data and utility lines and install exterior equipment.
Node 1 is now in Kennedy's Space Station Processing Facility, a new facility completed in 1994 and designed specifically for preparing International Space Station elements for launch. The node will be joined by two pressurized mating adapters, the first arriving at Kennedy in July from the McDonnell Douglas manufacturing facility in Huntington Beach, CA. Prior to launch, the two conical mating adapters will be attached to either end of the node at Kennedy. In orbit, the two adapters will serve as the connecting point for the U.S. and Russian station segments and as a docking location for the Space Shuttle.
"The Kennedy team at the Space Station Processing Facility has been preparing for several years for this occasion," said Glenn Snyder, Kennedy payload manager for STS-88. "We are looking forward to getting started with the processing of the first element as well as the others that will follow."
Work on Node 1 at Kennedy will include the completion of assembly and checkout tasks; acceptance testing of the node and mating adapters; communications testing with Mission Control; leak testing; and \Jtoxicology\j testing. Also, optical targets will be installed on the node that will assist the Shuttle's robotic arm operator during the docking in orbit.
The Functional Cargo Block, a U.S.-funded and Russian-built component, is currently undergoing modifications and enhancements at the Krunichev State Research and Production Space Center in Moscow. It is scheduled to be shipped to \JKazakhstan\j via a special rail car in January 1998 to begin final launch preparations.
On Friday, June 27, NASA will issue its Draft Environmental Impact Statement on the development and flight testing of the X-33 Advanced Technology Demonstrator. NASA also has announced plans to hold public meetings in 11 communities in five states to present the study's findings and seek public comment.
The environmental study examines the potential impacts of X-33 development and flight testing. Major issues addressed in the environmental study -- which began Oct. 7, 1996 -- include noise and sonic booms, flight safety, and effects on airspace and air traffic patterns. Many of the issues addressed originated from comments NASA gathered during public scoping meetings held in the same communities during the fall and early winter.
The 273,000-pound, wedge-shaped X-33 is being developed under a cooperative agreement with NASA by Lockheed Martin Skunk Works, Palmdale, CA. As many as 15 flight tests of the X-33 are planned to originate from Edwards Air Force Base, CA, and land at sites in Southern \JCalifornia\j, \JUtah\j, and Montana or Washington, beginning in 1999.
The objective of this technology development and demonstration effort is to support government and private sector decisions by the end of the decade on development of an operational next- generation space launch vehicle. A full-scale, single-stage-to- orbit reusable launch vehicle will dramatically increase reliability and lower the cost of putting a pound of payload into space from $10,000 to $1,000. By reducing the cost associated with transporting payloads into low Earth orbit, a commercial reusable launch vehicle would create new opportunities for space access and significantly improve U.S. economic competitiveness in the worldwide launch marketplace. NASA will be a customer, not the operator, of an industry-developed reusable launch vehicle.
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"Tropical Rainfall Measuring Mission Set For October 31 Launch (25 June, 1997)",498,0,0,0
NASA and the National Space Development Agency of \JJapan\j (NASDA) have set October 31 at 3:40 p.m. EST (Nov. 1, 1997, 5:40 a.m., JST) as the official launch date for the Tropical Rainfall Measuring Mission.
The first Earth science satellite dedicated to studying the properties of tropical and subtropical rainfall, the Tropical Rainfall Measuring Mission (TRMM) carries microwave and visible/infrared sensors, and the first spaceborne rain radar. Tropical rainfall comprises more than two-thirds of global rainfall and is the primary distributor of heat through the circulation of the atmosphere. More precise information about this rainfall and its variability is crucial to understanding and predicting global climate change.
"We're very excited about this major opportunity for cooperation with \JJapan\j, which is NASA's largest international partner in Earth science," said William Townsend, Acting Associate Administrator for NASA's Mission to \JPlanet\j Earth enterprise, Washington, DC. "The Tropical Rainfall Measuring Mission has great potential to improve scientific understanding of climate processes related to the heat released by tropical rainfall. In turn, this knowledge improves the global atmospheric circulation computer models that are used to make weather and climate forecasts."
NASDA will provide the Precipitation Radar for TRMM and an H-II rocket to launch the observatory on a three-year mission from the Tanegashima Space Center in \JJapan\j.
"We are very happy to provide the Precipitation Radar for TRMM and launch this first space mission to measure a driving force of the global atmosphere, tropical rainfall. We hope this U.S.-Japan joint mission provides important data for predicting global climate change and weather anomalies," said Dr. Kazuyoshi Yoshimura, Executive Director of NASDA in \JTokyo\j. "We will launch TRMM in November, and hereafter we can launch a rocket in each fall season. This is a good opportunity to expand the cooperation between the U.S. and \JJapan\j, and we expect a further cooperation in various fields, such as Earth observation satellites, Earth science, and global change research."
NASA's Goddard Space Flight Center in Greenbelt, MD, fabricated the observatory's structure and support systems, integrated and tested the \Jspacecraft\j and is providing two science instruments. Two other instruments are being provided by NASA's Langley Research Center, Hampton, VA, and its Marshall Space Flight Center, \JHuntsville\j, AL.
Goddard also will operate TRMM via NASA's Tracking and Data Relay Satellite System. NASA and NASDA will share responsibility for science data processing and distribution to the global change research community.
Current knowledge of rainfall is limited, especially over the oceans. By flying in a low-altitude orbit of 217 miles (350 kilometers), TRMM's complement of state-of-the-art instruments will provide extremely accurate measurements of the distribution and variability of tropical rain and \Jlightning\j, and the balance of solar radiation absorbed and reflected by Earth's atmosphere.
Extensive prelaunch testing of TRMM was completed recently and the observatory currently is undergoing final preparations for its shipment to the Japanese launch site in late August.
The TRMM launch window opens at 5:40 a.m. JST on Nov. 1, and with an approximate two-hour launch window daily for a 40- day period. TRMM's companion payload on the H-II rocket will be \JEngineering\j Test Satellite-7, a Japanese robotics experiment.
The TRMM project is part of NASA's Mission to \JPlanet\j Earth enterprise, a long-term, coordinated research effort to study the total Earth system and the effects of natural and human- induced changes on the global environment. TRMM is managed by Goddard for NASA's Office of Mission to \JPlanet\j Earth, Washington, DC.
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"Mir Collision (25 June, 1997)",499,0,0,0
Mir Complex Facing the Docking Mechanism
At 4:27 a.m. CDT on Wednesday, June 25, the Progress resupply \Jspacecraft\j collided with the Spektr module of the Russian Mir Space Station during practice docking operations.
The crew, including U.S. \JAstronaut\j Mike Foale, is fine and there are no plans to evacuate the Russian outpost.
The collision with Spektr caused a loss of pressure aboard the station. The crew shut Spektr's hatch and repressurized the remaining modules on Mir and cabin pressure on the station is holding steady.
NASA management is working closely with the Russian flight crew to assess the damage to Spektr and take appropriate action.
A Progress resupply vessel will be launched on July 5 and will dock with Mir on July 11. An "internal EVA" will then be performed by Tsibliev and Lazutkin to try to repair the damage to Spektr.
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"Mir Conditions Stable",500,0,0,0
Space Station Mir continues to orbit in stable condition as Russian flight controllers are working with \JAstronaut\j Michael Foale, Commander Vasiliy Tsibliev and Flight Engineer Alexander Lazutkin to develop methods for possible repairs to the damaged Spektr Module. Yuri Koptev, head of the Russian Space Agency, said that a special "plate" is under construction that will house 22 cables that will transmit power from Spektr's solar panel arrays back to the Mir Core Module for battery recharging.
Mir's gyrodynes are now functional and Mir is maintaining its orientation in the normal fashion. Batteries in the Core Module and in the Kvant-1 module are fully charged; the batteries in Kvant-2 are near a full charge.
Environmental conditions aboard Mir are good, the air conditioners are back on line, air pressure is stable, oxygen levels are nominal and CO2 scrubbing is being handled as normally by the Vosdukh CO2 scrubber supplemented by U.S.-supplied LiOH canisters. The Elektron oxygen generator aboard Kvant-1 was shut down due to overheating; the cause is believed to be a faulty coolant flow control valve. The crew will supplement the oxygen supply by burning three \Jlithium\j perchlorate oxygen-generating candles daily. It is hoped that the Elektron can be operated for up to 10 hours daily; longer operations are not recommended because the orientation of Mir that is most efficient for the generation of solar electricity causes overheating in Kvant-1's docking node. When Progress M-35 docks with Mir this will no longer be a problem.
Progress M-35 is expected to be ready for launch July 5, and will carry up materials necessary to make the necessary repairs.
The \JAntares\j satellite receiver delivered by Atlantis aboard STS-84 was successfully installed Monday and, following installation of a second unit shipped to Mir aboard Progress M-35, this should restore full communications via \JRussia\j's \JAltair\j data relay satellites.
On Wednesday, June 25, at 5:18 a.m. EDT, Progress M-34 crashed into the Mir Space Station Spektr Module during a test of the TORU, a newly installed Progress guidance system. The Spektr module, which houses many U.S. experiments and all of American \Jastronaut\j Michael Foale's personal effects, sprung a leak and had to be sealed off completely with its power shut down.
When the Spektr was punctured, air pressure in the Mir complex dropped from its normal 750 mm Hg down to 675 mm Hg. It has now risen back to 692 mm Hg (nominal range: 660-850 mm Hg) and has stabilized.
Spektr is the primary electricity generating module for the entire Mir complex. It is estimated by American sources that the loss of power to the overall Mir complex currently is in the neighborhood of 50%. The internal spacewalk to attempt repairs to Spektr is being planned for around 9:00 p.m. EDT July 11; if the crew is not yet ready by then, it will be put off until the evening of July 14.
Progress M-34 has now been de-orbited and has burned up in the atmosphere. The Progress M-34 resupply vessel, which was launched on April 6, 1997 and had delivered (along with other cargo) the parts used to repair the malfunctioning Elektron oxygen generator, was at the time of the collision loaded with garbage and was only being used to test the new docking equipment before being jettisoned on Saturday to burn up in the atmosphere.
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"Mir Conditions Improving",501,0,0,0
Space Station Mir continues to orbit in stable condition as Russian flight controllers are working with \JAstronaut\j Michael Foale, Commander Vasiliy Tsibliev and Flight Engineer Alexander Lazutkin to develop methods for possible repairs to the damaged Spektr Module. Yuri Koptev, head of the Russian Space Agency, said that a special "plate" is under construction that will house 22 cables that will transmit power from Spektr's solar panel arrays back to the Mir Core Module for battery recharging.
Mir's gyrodynes are now functional and Mir is maintaining its orientation in the normal fashion. Batteries in the Core Module and in both Kvant modules are fully charged.
Environmental conditions aboard Mir are good, the air conditioners are back on line, air pressure is stable, oxygen levels are nominal and CO2 scrubbing is being handled as normally by the Vosdukh CO2 scrubber supplemented by U.S.-supplied LiOH canisters. Because of electricity usage constraints, only the Elektron oxygen generator aboard Kvant-1 is operating and the one aboard Kvant-2 is dormant; the crew is supplementing the oxygen supply by burning one \Jlithium\j perchlorate oxygen-generating \Jcandle\j daily.
Telemetry indicates that air pressure has not yet fallen to zero aboard Spektr; the exact location and size of the leak is not presently known.
Progress M-35 is expected to be ready for launch July 5, and will carry up materials necessary to make the necessary repairs.
On Wednesday, June 25, at 5:18 a.m. EDT, Progress M-34 crashed into the Mir Space Station Spektr Module during a test of the TORU, a newly installed Progress guidance system. The Spektr module, which houses many U.S. experiments and all of American \Jastronaut\j Michael Foale's personal effects, sprung a leak and had to be sealed off completely with its power shut down.
When the Spektr was punctured, air pressure in the Mir complex dropped from its normal 750 mm Hg down to 675 mm Hg. It has now risen back to 692 mm Hg (nominal range: 660-850 mm Hg) and has stabilized.
Because the Russian communications satellite \JAltair\j is also out of service, information on this event is only available during the station's passes over Russian communication sites.
The Russian news agency Novosti reported on June 25 that the crew is safe and in no danger, but the return of the crew aboard the Soyuz escape craft is a possible option.
Spektr is the primary electricity generating module for the entire Mir complex. It is estimated by American sources that the loss of power to the overall Mir complex currently is in the neighborhood of 50%. The potential for effecting repairs to Spektr is being evaluated by the Russian Mir team; unless its power generating capacity can be restored, almost all scientific experimentation aboard Mir will have to be curtailed.
Progress M-34 has moved away from the Mir complex and is now under Russian ground control. The Progress M-34 resupply vessel, which was launched on April 6, 1997 and had delivered (along with other cargo) the parts used to repair the malfunctioning Elektron oxygen generator, was loaded with garbage and was only being used to test the new docking equipment before being jettisoned on Saturday to burn up in the atmosphere. Plans now are to de-orbit Progress M-34 on Monday, June 30.
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"Shuttle Mission Reflight In Quest Of Scientific Mysteries (23 June, 1997)",502,0,0,0
Searching to solve Earth-bound scientific mysteries in space, teams of researchers are taking NASA's \JMicrogravity\j Science Laboratory back into orbit aboard the Space Shuttle Columbia, targeted for launch on July 1.
This Shuttle mission will be a reflight of NASA's \JMicrogravity\j Science Laboratory-1, dedicated to 33 experiments concentrated in the areas of combustion science, protein crystals and study of the properties of metals and alloys important to many industrial processes.
In April, the previous flight of the \JMicrogravity\j Science Laboratory was cut short after four days because of a faulty fuel cell. The \Jastronaut\j team and investigators at Marshall Space Flight Center, \JHuntsville\j, AL, were only able to begin their schedule of experiments, which had been planned for 16 days.
"Those four days allowed our science team to barely open the door to tantalizing scientific research," said Joel Kearns, manager of NASA's \JMicrogravity\j Research Program Office at Marshall. "We were able to verify that we are headed in the right direction. But we were not able to reach our destination because of the shortened mission."
Kearns, nevertheless, cited important accomplishments during the abbreviated April mission. He said researchers tested their experiment hardware under flight conditions and it performed "extraordinarily." NASA and commercial research teams also acquired their first glimpses of some "never-before-seen phenomena," said Kearns.
The first observations of free-floating flame balls during the April flight were described by Dr. Paul Ronney of the University of Southern \JCalifornia\j in Los Angeles as "successful beyond my wildest dreams." His experiment, called the Structure of Flame Balls at Low Lewis-number, is designed to increase understanding of the characteristics of fuels and fires.
Dr. Gerard Faeth of the University of \JMichigan\j at Ann Arbor said that during the four-day flight, scientists got their first glimpse of the concentration and structure of soot from a fire burning in near-weightless conditions. The initial findings were an important step forward in understanding combustion and soot formation, he said. Like the flame ball experiment, this investigation could lead to improvements in fuel efficiency for all types of burning processes -- from car and jet engines to heating and cooking appliances.
"The success we've glimpsed from the shortened Shuttle mission in April makes it clear that we're heading in the right direction," said Kearns. "All activated research apparatus functioned in an outstanding manner. This upcoming mission has the potential to add considerably to our basic scientific knowledge and our quality of life here on Earth," Kearns pointed out.
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"NASA Research Provides New Insights on Black Holes (20 June, 1997)",503,0,0,0
A NASA scientist has made the first-ever observation of spinning black holes -- confirming Einstein's theory that black holes spin. The new observations from several orbiting \Jspacecraft\j adds to the growing body of knowledge on how these mysterious objects are formed and behave.
Black holes -- predicted by Einstein's General Theory of Relativity -- are believed to result from the collapse of a star or a group of stars. A black hole is an extremely compact and massive object with such a powerful gravitational field that nothing -- not even light -- can escape.
In a paper to be published today by The Astrophysical Journal, Letters, Dr. Shuang Nan Zhang of the Universities Space Research Association at NASA's Marshall Space Flight Center, \JHuntsville\j, AL, and his research associates report that two of the black holes they have studied are rapidly spinning - - rotating 100,000 times per second -- while others are spinning very slowly or not at all.
By comparison, before this discovery, the Crab Pulsar was considered to be among the most rapidly spinning objects in the universe -- rotating 33 times per second.
"Black holes have always been difficult objects to define. We can only characterize them with three properties -- mass, charge and spin," said Zhang.
"In the past, we've only been able to measure a black hole's mass. But now that we've learned how to measure a second property -- spin rate -- one might say that we are two-thirds of the way to understanding black holes. This is a major leap in unraveling the black hole mystery," said Zhang.
"Determining the spin of black holes is of enormous importance, not only that the spin gives us an idea of how much angular momentum the black hole has 'swallowed' during its lifetime, but also we can examine whether the spin is related to the formation of powerful jets," said Dr. Mario Livio, senior scientific staff member at the Space \JTelescope\j Science Institute. "The two rapidly spinning black holes also occasionally eject streams of high-speed material called relativistic jets from the black hole region -- at roughly the same speed at which the hole is spinning," said Zhang.
Since a black hole emits no light, the best way to observe it and learn about its properties is to study its interaction with the environment around it.
"The Theory of Relativity explains that there should be a last stable orbit around the black hole," said Zhang. "Material inside this orbit cannot survive and is consumed by the black hole."
The researchers determined the spin rate by accurately measuring the closest stable orbit of material around the black hole.
"The size of this orbit is related to the spin of the black hole. By looking at the material occupying this orbit and measuring the orbit's size, we can learn how fast the black hole is spinning," said Zhang.
To measure the orbit, Zhang and his colleagues, Dr. Wei Cui of the \JMassachusetts\j Institute of Technology, Cambridge, MA, and Dr. Wan Chen associated with both the University of Maryland at College Park and NASA's Goddard Space Flight Center, Greenbelt, MD, used data from several satellites -- the Compton Gamma Ray Observatory, the Rossi X-Ray Timing Explorer, the Roentgen Satellite, and the Advanced \JSpacecraft\j for \JCosmology\j Astrophysics -- which collect X-rays emitted from the material orbiting near the black hole.
Using this technique, Zhang and his colleagues have measured the spins of several stellar mass black holes -- small black holes with masses comparable to stars.
Those measured reside in binary systems, comprised of a black hole and an ordinary star. The black hole, with its massive gravitational force, consumes material from the atmosphere of the companion star.
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"NASA To Test Adaptable Mobile Robot In South American Desert (17 June, 1997)",504,0,0,0
From laboratories and a science center in North America, a group of NASA and Carnegie Mellon University scientists will control a robotic rover this summer as it explores a desert in South America to learn more about driving automated vehicles on Mars and the Moon.
During the 45-day field experiment from June 15 to July 31, scientists from NASA's Ames Research Center, Moffett Field, CA, and Carnegie Mellon's Robotics Institute, \JPittsburgh\j, PA, will conduct an unprecedented 120-mile robotic trek in the Atacama Desert in northern \JChile\j. The scientists will test the ability of the robot, nicknamed Nomad, to navigate, explore and perform science tasks remotely.
"The primary objective of the Atacama Desert trek is to develop, evaluate and demonstrate a robot capable of long distance and long duration planetary exploration," said David Wettergreen, NASA Ames project manager.
"During different phases of this test, we will configure the robot to simulate wide-area exploration of the Moon, the search for signs of past life on Mars and the gathering of \Jmeteorite\j samples in the Antarctic, which makes for a really unique and challenging experiment," said Dave Lavery, Telerobotics Program Manager at NASA Headquarters, Washington, DC.
Chile's Atacama Desert, a cold, arid region located above 7,000 feet, was chosen for the field experiment because its harsh terrain is analogous to that found on Mars and the Moon. The desert's barren landscape features craters, rocks and loose sand without any vegetation due to the lack of rain.
"This site is pretty much what we expect to find on Mars," said Nathalie Cabrol, the expedition's NASA science team leader. "Our goal is to simulate several NASA planetary exploration missions, and this will provide some good training for future missions," Cabrol added. The desert trek also will test and validate tools and techniques that NASA has been developing for future planetary missions.
Nomad was designed and built by researchers at Carnegie Mellon's Robotics Institute. About the size of a small car, the robot weighs 1,600 pounds and features four-wheel drive/four-wheel steering with a chassis that expands to improve stability and travel over various terrain conditions. Four aluminum wheels with cleats provide traction in soft sand. Power is supplied by a \Jgasoline\j generator and enables the robot to travel at speeds up to 20 inches per second. Nomad also contains onboard navigation sensors and computers to enable it to avoid obstacles without relying on a human operator.
Nomad's unique onboard panospheric camera provides live 360-degree, video-based still images of the robot's surroundings. The images will be displayed on large screens at Ames and Carnegie Science Center in \JPittsburgh\j, where the public will have an opportunity to control the rover every day throughout the trek. The rover carries additional color video cameras to provide stereo vision for detecting obstacles and high-resolution color video cameras for experiments in remote \Jgeology\j to be conducted by NASA.
The total cost of developing Nomad and conducting the desert trek is $1.6 million. The project is funded by NASA with in-kind support from corporate sponsors and educational foundations.
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"Pathfinder Sets New World Record (12 June, 1997)",505,0,0,0
The Pathfinder, a remotely piloted \Jaircraft\j, set a new world record this week for high-altitude flight for solar-powered \Jaircraft\j at the Pacific Missile Range Facility, \JKauai\j, Hawaii.
The study of solar-powered, remotely piloted \Jaircraft\j could lead to revolutionary technologies to conduct low-cost, high- altitude atmospheric and remote sensing studies.
During its flight on Monday, June 9, the \Jaircraft\j reached an altitude of 67,350 feet, breaking the previous record of 50,500 feet it set on Sept. 11, 1995, at NASA's Dryden Flight Research Center, Edwards, CA.
Pathfinder is part of NASA's Environmental Research \JAircraft\j and Sensor Technology (ERAST) Program, and was designed, manufactured, and operated by Aerovironment Inc., Simi Valley, CA. Pathfinder is a lightweight flying wing whose six electric motors each drive a propeller. Solar cells covering the 98-foot-wing span wing provide power to the motors and other systems.
Pathfinder's record-breaking flight began with a takeoff into light ocean breezes. After completion of low-level system checks, Pathfinder began climbing, and six hours into its flight, passed its previously set record of 50,500 feet. After the \Jaircraft\j climbed to 67,350 feet, the decision was made to land. Pathfinder spent over 90 minutes flying above 60,000 feet. The solar-powered vehicle completed its mission with a perfect landing.
"This achievement is a direct result of the teamwork of NASA, the Pacific Missile Range Facility, Aerovironment, and the \JKauai\j Community College," said Jennifer Baer-Riedhart, the Pathfinder Program Manager at Dryden.
A unique aspect of the Pathfinder flight test was the extensive use of special-purpose computational techniques developed by NASA's Ames Research Center, Mountain View, CA. These techniques were used to validate the \Jaircraft\j's stability characteristics and to adjust the on-board control system characteristics "on the fly" at regular altitude intervals.
The Environmental Research \JAircraft\j and Sensor Technology Alliance is made up of NASA, industry, and the environmental science community. Participants have access to NASA technology and expertise in return for giving assistance in NASA projects. \JKauai\j Community College has personnel and equipment that assist in the maintenance, repair, and operation of the program's \Jaircraft\j for which NASA provides technical work experience and additional equipment, including spare solar panels.
The Environmental Research \JAircraft\j and Sensor Technology Program is NASA's response to growing scientific requirements for measurements at higher altitudes and longer durations than the current fleet of scientific platforms permits. The predominant focus of the program consists of advancing the technology of remotely piloted \Jaircraft\j. Concurrent efforts in development, miniaturization, and \Jintegration\j of science sensors can be later applied to important science missions. Pathfinder is one of many projects and \Jaircraft\j in the program.
Pathfinder's entire mission was monitored, and the world record will be certified, by the National Aeronautic Association's Capt. E.E. "Buck" Hilbert, Chairman of the Contest and Records Board.
Additional missions for Pathfinder this year include other high-altitude flights and several science missions with payloads supplied by Ames for monitoring \Jcoral\j reef degradation and deforestation around the island of \JKauai\j.
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"Mars Rover (Prototype) Completes Simulated Mars Trek (10 June, 1997)",506,0,0,0
NASA's newest, six-wheeled prototype Martian rover -- nicknamed Rocky 7 -- has successfully passed its most rigorous field test yet, traveling six-tenths of a mile over rugged, Mars- like terrain, while conducting science experiments and snapping 580 photographs along the way.
The week-long series of field tests, carried out May 23-30 at Lavic Lake, an ancient lake bed about 175 miles east of Los Angeles, CA, was designed to simulate several weeks of a real Mars rover mission and to test the rover's ability to drive much greater distances than current rovers. In addition, Rocky 7 conducted five simulated science experiments in real-time and collected samples of soil and rocks that would be retrieved and returned to Earth by a later Mars mission.
"One of the chief objectives of these tests was to test Rocky 7's ability to traverse farther over a wide variety of terrain with more Mars-like characteristics than we did in the last set of tests in December 1996," said Dr. Samad Hayati, Rocky 7 task manager at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "The rover actually traveled about 80 percent farther than it traveled in the last set of tests, over three distinct terrains, using a minimum of instructions from us to guide its way."
The Rocky 7 rover represents the newest model of rovers that may be sent to Mars in the years 2001 and 2003. It looks, however, very similar to its predecessor, Sojourner, which will land on Mars on July 4. Rocky 7 weighs slightly more than Sojourner at 33 pounds and has about the same dimensions -- measuring 19 inches wide by 25 inches long by 12.5 inches tall. Rocky 7 also sports the same six-wheeled chassis and spring-less "rocker-bogie" mobility system, which allows the vehicle to conform to the contours of the surface and scale objects almost as tall as itself without tipping over.
Continued robotic exploration of Mars in the next century will focus on the search for water and evidence to confirm hints that life may have existed once in Mars' early history. Successive Mars missions will be designed not only to examine the planet's composition, atmosphere and weather, but also to identify natural resources that could be mined and used for eventual human expeditions to the red planet.
The southern side of Lavic Lake, located in the Twenty-Nine Palms Marine Corps Base near Palm Springs, CA, was chosen for the field tests because it is a playa, analogous to some regions of Mars, with areas of \Jlava\j flow, cracked mud, terrain strewn with \Jbasalt\j rocks and an alluvial fan.
Rocky's travels began on a \Jbasalt\j flow covered with cobblestones resting in a layer of wind-blown silt, which offered a variety of obstacles for the robot to hurdle. Engineers tested some of the rover's new features, such as a 12.5-inch manipulator arm with four degrees of freedom. Mounted on the front of the vehicle, the arm carried a "point reflectance" spectrometer that could be extended four inches in any direction to study the color of various surfaces. In future rover missions on Mars, science instruments on the rover arm will help researchers determine the composition of surface soils and rocks.
Engineers also tested a 4.5-foot, antenna-like mast, which would be deployed once the future rover was out and about on Mars. The mast has three degrees of freedom and can be used in much the same way as an arm to deploy science instruments against rocks or align them in the nadir, or down-pointing, position. Two science instruments -- a Moessbauer spectrometer and a nuclear magnetic \Jresonance\j spectrometer -- were mounted on the mast to study surface rocks with different types of coatings, such as red iron oxide and desert varnish, which might be found on Mars. To carry out the variety of science experiments performed during the week, Rocky 7 had to raise its mast 85 times.
Rocky 7 carried a pair of stereo imagers on the front and back of the vehicle, which acted as its "eyes." The rover was furnished with simulated descent imaging to recreate landing, then asked to deploy its mast and begin each traverse and sequence of imaging and science experiments.
"Images and science measurements were obtained in several regions of the \Jbasalt\j flow," said Dr. Richard Volpe, chief engineer on the rover development team at JPL. "This pavement of \Jbasalt\j boulders and outcrops offered many terrain obstacles for rover navigation and numerous targets for the rover to measure cobbles and the underlying dust."
In the second journey, the rover set out over the playa, strewn with craters and ejecta fields, and traveled into a crater. Using its mast and arm, the vehicle was able to measure properties of the mud-cracked floor. Rocky 7 also took images of its own tire tracks to help scientists update its location.
"The rover conducted several long traverses across the playa floor, taking images of the tracks left by its wheels so that we could trace its path," said Dr. Raymond Arvidson, science team lead and chairman of the Earth and Planetary Sciences Department at Washington University, St. Louis, MO. "The tracks are used to update positional information, after the observations are completed and help us map out the vehicle's next route."
The last excursion was the most challenging -- an obstacle course taking the rover over an alluvial fan extending from the nearby mountains. There, Rocky 7 was asked to use its science instruments to look for evidence that water had been transported to the sediment and to explore the region for cobbles and boulders that had come from volcanic rocks, just as it will do on Mars some day.
"Imaging and \Jspectroscopy\j data were acquired for the fan rocks and fine-grained sediment, and samples of the sediment were collected," Arvidson said. "The data are currently being analyzed and will be used to fine-tune rover designs and operations and to evaluate what can be learned about ancient lake environments and desert pavement formation."
By the end of the week, the rover had returned 580 images to remote operators in the field and those stationed at JPL. The field test simulated 32 days of a real Mars rover mission.
Classrooms across the country and as far away as \JFinland\j participated in a remote driving test on the last day of the field work. The demonstration was designed to determine how well the vehicle could be controlled remotely using a World Wide Web operator interface called the Web Interface for Telescience. Six schools in \JCalifornia\j, Oregon, Georgia, \JIdaho\j, \JTexas\j and \JFinland\j participated in the exercise to command the rover from their classrooms, as scientists will do one day from their home institutions.
The Rocky 7 rover development and field testing was supported by JPL's Robotics and Mars Exploration Technology Program Office for NASA's Office of Space Science, Washington, DC.
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"Hubble Is First To Spot Colliding Supernovas (10 June, 1997)",507,0,0,0
Astronomers using the Hubble Space \JTelescope\j have taken the first images clearly showing interactions between two or more exploding stars, called supernovas, which are producing a tremendous display in a galaxy 17 million light years from Earth.
Debris speeding out from the supernovas is slamming together in a cosmic collision, the likes of which have never before been seen. The images are especially startling because the collision is taking place over a time period of perhaps a few hundred years, a fleeting blink of an eye in the ancient cosmos, according to William P. Blair, a Johns Hopkins University astrophysicist who led a team of scientists making the discovery.
Blair worked with two other astrophysicists: Robert A. Fesen, from Dartmouth College, and Eric M. Schlegel, from the Harvard-Smithsonian Center for Astrophysics. Their findings were detailed in a poster paper presented today during a meeting of the American Astronomical Society. The paper is on display from 10 a.m. to 6:30 p.m. in the South Main Hall of the Benton Convention Center, \JWinston-Salem\j, NC.
The astronomers were puzzled when they first spotted the object with a \Jtelescope\j at the Kitt Peak National Observatory in \JArizona\j and with the ROSAT X-ray satellite. It was extremely bright in optical and X-ray light -- just like a young \Jsupernova\j - - a star much more massive than the Sun destroying itself in a \Jtitanic\j explosion. But further analysis showed that it did not have the proper mixture of elements, and it was expanding too slowly to be a young \Jsupernova\j. Instead, it had all the characteristics of a much older remnant of a \Jsupernova\j, in which the expanding bubble of debris has spread far into space, diffusing into the interstellar gas.
How, then, could it be so bright?
They found the answer after using the Hubble \JTelescope\j's Wide Field and Planetary Camera 2. The space \Jtelescope\j's superior resolution -- its ability to distinguish separate objects that are close to each other -- brought the matter into clearer focus. Whereas the bright point of light had looked like a single \Jsupernova\j from the ground, the Hubble image clearly showed the remnants of two or more objects colliding.
When a massive star explodes, gas and debris are thrown in all directions at speeds of up to 22 million miles an hour (36 million kilometers per hour), producing a shock wave and compressing the gas into a dense "shell" of material.
The Hubble image captured the shells from one or more supernovas crashing into each other, like "a train wreck," and producing a tremendous light display for an object so far away, Blair said.
Astronomers had predicted the process, but because the phenomenon is so short-lived, it had never been seen directly.
"It's the first time that we've identified one of these interactions right when the shells are in the process of slamming into each other," Blair said. "The reason this object is so bright is that we caught it at a very specific time in its evolution. And Hubble's resolution is what allowed us to see it."
The \Jsupernova\j interaction is taking place in a galaxy known as NGC 6946, located 17 million light years away in the northern \Jconstellation\j \JCepheus\j. Like the Milky Way, it is a spiral galaxy, but it's about half the Milky Way's mass and size. The exploding stars probably were about 40 light years apart.
Supernova explosions have been seen in NGC 6946 at an extraordinary rate: Astronomers have observed six supernovas in that galaxy since 1917. Only one other galaxy has displayed so many supernovas.
"It indicates that not only is there a lot of star formation going on, but a lot of those stars are massive," Blair said. "They are evolving quickly, and they are exploding as supernovas."
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"Hubble Instruments Provide Steady Stream Of Discoveries (9 June, 1997)",508,0,0,0
Seeing a group of baby Sun-like stars surrounding their "mother star"; detecting a \Jtitanic\j shock wave smashing into unseen gas around a \Jsupernova\j; finding a disk at the heart of a galactic collision: these are among what are now becoming routine discoveries for astronomers as they finish checking out the Hubble Space \JTelescope\j's new instruments.
Like someone trying out a new camera, Hubble instrument teams have looked at a variety of objects and made some unexpected new findings along the way.
These results are being presented at the 190th Meeting of the American Astronomical Society in \JWinston-Salem\j, NC, by the principal investigator for the Near Infrared Camera and Multi-Object Spectrometer, Dr. Rodger Thompson of the University of \JArizona\j, and Dr. Bruce Woodgate, principal investigator for the Space \JTelescope\j Imaging Spectrograph at NASA's Goddard Space Flight Center, Greenbelt, MD.
"We were very pleased to see this new generation of ultraviolet light sensors come up to full operation so successfully. That's a major step forward for space \Jastronomy\j. We still have a lot of work to do to get them fully commissioned, but these first ultraviolet science images taken with the spectrograph are very promising," said Dr. David Leckrone, Hubble Project Scientist at Goddard.
NEW FINDINGS FROM THE NEW INSTRUMENTS
PARENT AND OFFSPRING STARS Peering through the dust in a nearby star-forming region called NGC 2264 that contains the Cone Nebula, the infrared camera has provided direct confirmation of a type of starbirth called "triggered" star formation. Though such a starbirth scenario has been theorized for years, this is the first direct observation on such a small-distance scale. This form of starbirth happens when a gale of high speed particles from a young massive star (called NGC 2264 IRS or Allen's Source) compresses nearby dust and gas until it is dense enough to trigger the formation of six much smaller and fainter Sun-like stars only a fraction of a light-year away from the massive "parent."
BLAST WAVE FROM \JSUPERNOVA\j The spectrograph has discovered the first direct evidence for material from \JSupernova\j 1987A colliding with an outer ring of gas which was ejected before the star exploded. The fastest debris, moving at 9,000 miles per second or 1/20th the speed of light, is slamming into the slower moving gas which was ejected when the doomed star was a red giant. Astronomers report this is a new component of material around the star which was predicted to exist, but was never directly seen before.
DUST DISK DEEP IN THE HEART OF GALACTIC COLLISION Probing the nucleus of the peculiar galaxy called Arp 220, the infrared camera has discovered a 300-light-year diameter dust disk and other remarkably complex structures around the galaxy's unique "double nucleus" of two bright compact star clusters. The disk, which may fuel a supermassive black hole or rapid star- formation activity, appears silhouetted against one of the twin star clusters at the nucleus. BEAM FROM A BLACK HOLE The imaging spectrograph, which just last month demonstrated its efficiency as a black hole hunter, now shows the results when the power of a black hole is unleashed into its surrounding environment. In a single observation, the spectrograph measured the velocities of hundreds of gas blobs caught up in a twin-cone beam of radiation emanating from a supermassive black hole at the core of galaxy NGC 4151. Follow-up observations reveal hot gas emanating from deep within the throat of the beam, near the vicinity of the black hole. These observations also allow scientists to map the mass outflows near the black hole. The surprisingly complex motion may offer clues to the galaxy's stellar population, the orientation of the beam in the past, or evidence of some kind of backflow of gas into the central cone regions.
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"Spacewalkers Named For Space Station Assembly Flights (9 June, 1997)",509,0,0,0
A cadre of 14 Space Shuttle astronauts has begun intensive training in preparation for the spacewalks required for on-orbit construction of the International Space Station.
"It is important for us to begin work now to train the crews who will support Space Station assembly flights," said David C. Leestma, director of Flight Crew Operations. "These crew members will be exceptionally busy preparing for some challenging and demanding tasks, from initial assembly through installation of the robotic arm and an airlock for station-based spacewalks. We expect that these assignments may be refined in the future, with additional tasks as mission requirements dictate."
The first astronauts assigned to train for space station assembly tasks were Jerry L. Ross (Col., USAF) and James H. Newman, Ph.D., who were named to the STS-88 crew in August 1996. Training will now expand to include Leroy Chiao, Ph.D.; Robert L. Curbeam, Jr., (Lt. Cmdr., USN); Michael L. Gernhardt, Ph.D.; Canadian \JAstronaut\j Chris A. Hadfield (Lt. Col., CAF); Thomas D. Jones, Ph.D; Mark C. Lee (Col., USAF); Michael Lopez- Alegria (Cmdr., USN); William S. "Bill" McArthur, Jr., (Col., USA); Carlos I. Noriega (Major, USMC); James F. Reilly, II, Ph.D.; Joseph R. Tanner; and Peter J.K. "Jeff" Wisoff, Ph.D.
The specific assignments for these U.S. assembly flights, through August 1999, are:
STS-88/Jerry Ross and Jim Newman, July 9, 1998, Node 1; ISSA-2A, Connect power cables
STS-92/Leroy Chiao; Jeff Wisoff, January 1999, Integrated Truss, ISSA-3A, Michael Lopez-Alegria, Portable Mating, and Bill McArthur Adapter
STS-97/Joe Tanner and Carlos Noriega, March 1999, Photovoltaic module, ISSA-4A
STS-98/Mark Lee and Tom Jones, May 1999, U.S. Lab Module, ISSA-5A
STS-99/Chris Hadfield and, June 1999, Lab outfitting; ISSA-6A, Robert Curbeam Remote Manipulating System
STS-100/Mike Gernhardt and James Reilly, August 1999, Joint Airlock, ISSA-7A, High Pressure, Gas Assembly
"The assignment of these crew members is a critical element in our ability to bring together the elements of the space station on the ground and then successfully assemble them in orbit," said Randy Brinkley, Space Station program manager. "I am highly pleased with the level of expertise and dedication these crew members bring to the program."
Each of the astronauts named for these assembly flights has previous spaceflight experience, with the exception of Curbeam, who will make his first flight in July 1997, and Reilly, who will make his first flight in January 1998. Ross, Newman, Chiao, Wisoff, Tanner, Lee and Gernhardt all have performed spacewalks during their previous flights. The remaining astronauts, Lopez- Alegria, McArthur, Noriega, Jones, Curbeam, Reilly and Hadfield are training for their first spacewalks. When Hadfield conducts his scheduled spacewalk in June 1999, he will become the first Canadian to walk in space.
NASA's Space Shuttle program passed a major milestone today on its way to reflying the orbiter Columbia and the first reflight of the same payload and crew in Space Shuttle history. Columbia, which saw an abbreviated mission in April due to indications of a faulty fuel cell, was transported from \JFlorida\j's Kennedy Space Center (KSC) Orbiter Processing Facility (OPF) to the Vehicle Assembly Building today where it will be mated with an external tank and solid rocket boosters in preparation for roll-out to Launch Pad 39A next week.
With the Spacelab payload secure in the orbiter's cargo bay, NASA remains on track for a targeted July 1 launch date for reflight of the \JMicrogravity\j Science Laboratory mission.
Suspicious readings from one of Columbia's fuel cells compelled NASA managers to cut the STS-83 mission short after only four days in space, marking only the third time in Shuttle history that a mission was curtailed for mechanical reasons. The original mission was expected to last 16 days.
Since the return of Columbia following the shortened STS-83 mission, the suspect fuel cell has undergone extensive anaylsis. The conclusion is that an undetermined and isolated incident caused a slight change in the voltage of about one-fourth of the 96 cells that make up each fuel cell.
To ensure the health of the fuel cells pre-launch, the power plants will be started earlier than usual to allow for additional monitoring before liftoff. Also, the program is reviewing the possibility of installing new fuel cell performance monitors that will indicate individual cell "health" rather than a single monitor for each of three 32-cell substacks.
This will provide additional insight into pinpointing large voltage shifts in a single cell, which could indicate a potential problem, or a small voltage shift in a number of cells, which is a benign situation. Presently, the performance monitor provides a gross indication of fuel cell health, which caused the team to assume the worst in the case of STS-83.
As with all hardware issues on the Shuttle, fuel cell anomalies are taken seriously and reviewed extensively prior to clearing future missions for launch. Additionally, the flight rules are being reviewed to ensure that proper insight is provided to flight controllers in making decisions on the health of the fuel cells.
Columbia launched on April 4 and landed in \JFlorida\j on April 8 without completing the mission's science objectives. About two weeks later, Shuttle program managers decided to refly the \JMicrogravity\j Science Laboratory mission on STS-94 as soon as possible within safety guidelines.
"This decision demonstrated the Shuttle program's confidence in the KSC processing team," said Bob Sieck, Director of Shuttle Processing. "Special credit goes to the workers in Orbiter Processing Facility Bay 1. They produced a quality product in record time."
When marching orders were given, NASA's Shuttle and payload communities teamed up to give Columbia and the Spacelab payload a speedy turnaround. Once in the OPF, replacement of the problem fuel cell was the first order of business and that was completed the week after landing. Managers then put into motion a strategy that minimized the amount of rework performed on the Shuttle and reduced the time required to service the payload.
The ambitious schedule required that all experiment reservicing be done while the Spacelab remained in the Shuttle's payload bay. Between flights, Spacelab is normally removed and then transported to KSC's Operations and Check-out Building for rework in a spacious environment. Payload technicians overcame the Shuttle's cramped conditions and successfully completed many critical tasks such as replenishing the flammable fluids of a combustion experiment.
"This is the first time that a payload has remained in an orbiter between flights," said KSC Payload Manager Scott Higginbotham. "We are excited about having accomplished something that has never been tried before."
Working side-by-side with the payload team, Shuttle technicians and managers faced some challenges of their own. Normally an orbiter visits the OPF for about 85 days in preparation for its next launch, but this reflight called for about 56 days in the facility. Managers saved some time by deferring certain routine structural inspections until Columbia's next mission, but other work could not wait and had to be accomplished before launch.
For example, the Shuttle's forward reaction control system, located in the nose of the vehicle, had to be removed with three out of sixteen steering thrusters requiring replacement. Also, two of the three 85-pound auxiliary power units that provide hydraulic power to Columbia's flight control systems were replaced having reached their run-time limit between overhauls.
An important part of this time-saving strategy was to minimize the burden on the Shuttle processing team. "Most of the time savings in the OPF was the result of a concerted planning effort between NASA and our contractor partners," said Grant Cates, NASA flow director for Columbia. "Once the plan was in place, the team approached this challenge in much the same way that they approach every flow."
To further speed up Columbia's processing for reflight, managers took one main engine scheduled to fly on Atlantis in September and two engines from Columbia's November flight. The external fuel tank and solid rocket boosters being used on STS-94 were originally slated for mission STS-85.
With the original STS-83 astronauts slated to fly again on STS-94, additional time savings were achieved by leaving the crew compartment set-up virtually unchanged. The crew equipment interface test and the terminal countdown demonstration test, both familiarization exercises previously completed by the crew, were deemed unnecessary for this reflight mission. Columbia is scheduled to roll out of the Vehicle Assembly Building on June 11, bound for launch pad 39A. The STS-94 launch is currently targeted for July 1 at 2:37 p.m. EDT.
Regardless of the efforts necessary to perform a quick turn- around of Columbia and its \Jmicrogravity\j science payload, NASA managers and engineers are confident that no safety margins were compromised.
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"X-38 Atmospheric Vehicle To Begin Flight Tests In July (4 June, 1997)",511,0,0,0
The first X-38 atmospheric test vehicle, which carries applications for future space vehicles, was shipped today from the Johnson Space Center, Houston, TX, to the Dryden Flight Research Center, Edwards, CA, to begin unpiloted flight tests in July.
The X-38 represents an innovative new \Jspacecraft\j design as a technology testbed, with possible use as an International Space Station emergency crew return "lifeboat."
Once operational, the successors to the X-38 may become the first new piloted \Jspacecraft\j to travel to and from orbit in more than 20 years, and the X-38 is being developed at a fraction of the cost of past human space vehicles.
The primary application of the new \Jspacecraft\j would be as an International Space Station "lifeboat," which would be delivered to the station by the Space Shuttle. The project also aims to develop a design that could be modified easily for other uses, such as a possible joint U.S. and European human \Jspacecraft\j that could be launched on the French Ariane 5 booster.
"Beginning full-scale flight tests is a big milestone for us that our team has been looking forward to with a lot of excitement," said X-38 project manager John Muratore. "No one has ever done anything like this before ╨ deploying a parafoil from a lifting body and flying a lifting body with an all-electric flight control system ╨ and there are unknowns. We expect surprises. But we have done a lot of work to minimize the unknowns, and we are confident this vehicle can perform well."
The atmospheric test vehicle, designated vehicle 131, is the first of three sub-scale vehicles largely built at Johnson planned for such testing. The unpiloted flight testing will begin at Dryden with "captive carry" flights, during which the vehicle remains attached to the NASA B-52 \Jaircraft\j, in July and early August. The first free-flight drop test of the vehicle, in which it will be released at an altitude of 25,000 feet, is planned for late August. Similar free-flight drop testing will continue at Dryden periodically through late 1999. An unpiloted space flight test is scheduled for launch aboard a Space Shuttle in the spring of 2000. The X-38 space flight test vehicle also will be built largely at Johnson.
The X-38 is being developed with an unprecedented eye toward efficiency, taking advantage of available equipment and already- developed technology for as much as 80 percent of the \Jspacecraft\j's design. The design uses a lifting body concept originally developed by the Air Force's X-24A project in the mid-1970s. Following the jettison of a deorbit engine module, the X-38 would glide from orbit unpowered like the Space Shuttle and then use a steerable parafoil parachute for its final descent to landing.
In the early years of the International Space Station, a Russian Soyuz \Jspacecraft\j will be attached to the station as a crew return vehicle. As the size of the station crew increases, however, a return vehicle like the X-38 that can accommodate up to six passengers will be needed.
NASA has awarded contracts for $1 million each to Lockheed Martin and the Boeing Company to study a possible, future upgrade to the Space Shuttle in which the rocket boosters that power the Shuttle would fly back to the launch site, rather than drop into the ocean for later recovery.
The study effort being led by NASA's Marshall Space Flight Center, \JHuntsville\j, AL, is part of a continuing program to upgrade the current Shuttle fleet -- providing additional safety, improved operability, enhanced performance and reduced costs.
To complement its in-house examination, Marshall is asking each of the contractors for an in-depth concept definition on the liquid fly-back booster.
Proposals from industry will provide data and configuration studies for both the booster and its engine, focusing on the liquid fly-back booster concept -- including analysis and evaluation, model fabrication and wind-tunnel testing.
If the concept is implemented, the unpiloted, liquid fly-back boosters would become the first-stage boosters of the Space Shuttle system.
Under the systems \Jintegration\j concept being studied, a Shuttle launch using the upgraded booster would appear similar to the current system to an observer on the ground.
After separation from the Shuttle, however, the two booster rockets would begin coasting for nine minutes, rather than parachuting into the ocean. Then jet engines would be started, and the unpiloted boosters would fly back and land at the Kennedy Space Center, FL.
Other principal elements of the present Space Shuttle system - - including the orbiter, Space Shuttle main engines and external tank -- would remain essentially unchanged if the new boosters are incorporated.
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"NASA Studies High Altitude Radiation With Upgraded ER-2",513,0,0,0
Using an upgraded NASA ERR2 \Jaircraft\j, researchers at NASA's Langley Research Center, Hampton, VA, have begun a month-long campaign to measure radiation at high altitudes.
This campaign, funded by NASA's High-Speed Research program, is the first of several that will measure naturally occurring radiation from cosmic and solar rays at altitudes between 52,000 and 70,000 feet.
The data will be used to characterize the radiation environment for the aircrew and frequent-flying public on a future High-Speed Civil Transport. The High-Speed Civil Transport, a conceptual \Jsupersonic\j airliner, would carry 300 passengers at 2.4 times the speed of sound, at altitudes of up to 68,000 feet.
"The broad aim of the Atmospheric Ionizing Radiation ER-2 flight-measurements campaign is to understand the composition, distribution and intensities of cosmic and solar radiation at commercial \Jsupersonic\j transport-cruise altitudes," said Allen Whitehead, the High-Speed Research environmental impact manager.
"Our primary concern is the level of uncertainties in the knowledge of the upper atmosphere's radiation environment and the human body's response to that type of environment," said Dr. John Wilson, the Atmospheric Ionizing Radiation chief scientist.
"Radiation measurements will be obtained by an array of instruments from the United States, Canada, \JGermany\j, United Kingdom and \JItaly\j in a collaborative effort devised by Dr. Wilson," said Donald Maiden, the Atmospheric Ionizing Radiation project manager. "The instrument types which make up the array were recommended by the National Council on Radiation Protection in a study sponsored by the High-Speed research program."
"The primary thrust is to characterize the atmospheric radiation and to define dose levels at high altitude flight. A secondary thrust is to develop and validate dosimetric techniques and monitoring devices for protection of the aircrew who work many hours at cruise altitudes," Maiden added.
According to Maiden, "Even though the exposure levels are higher at the higher cruise altitude, the typical flying public will actually receive less radiation exposure than on today's subsonic transports because of the higher speed of the High-Speed Civil Transport. This is another advantage for speed."
The flight program is a collaborative effort with the Department of Energy's Environmental Measurements Laboratory; NASA's Johnson Space Center; the Canadian Defense Research Establishment and Royal Military College; the German Aerospace Research Establishment; the United Kingdom's National Radiation Protection Board; the Boeing Company; and several domestic and foreign university guest investigators.
Recent modifications to the NASA ER-2, sponsored by NASA's Mission to \JPlanet\j Earth program, increased its altitude capabilities, allowing it to reach easily those altitudes where the High-Speed Civil Transport will fly. The NASA ER-2 is based at NASA's Ames Research Center, Mountain View, CA.
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"Astronaut Named To Station Assembly Flight (2 June, 1997)",514,0,0,0
Japanese \JAstronaut\j Koichi Wakata will fly on STS-92, the third Space Shuttle mission to assemble the International Space Station, set for a January 1999 launch on Atlantis.
Wakata's assignment to the mission was announced by NASA Administrator Daniel S. Goldin and Japanese Science and Technology Agency Minister Riichiro Chikaoka, in \JTokyo\j, \JJapan\j, today.
"NASA is honored to have Mr. Wakata participate in such an early and significant space station assembly mission," Goldin said. "His participation on this flight is symbolic of the close bond that has developed between the American and Japanese space programs, and the extent to which we rely upon one another to meet our mutual objectives in space."
Wakata was selected as an \Jastronaut\j in 1992 and has one previous space flight to his credit. He flew as a Mission Specialist on STS-72 in January 1996 aboard Endeavour. During that flight, the crew retrieved the orbiting Space Flyer Unit satellite which was launched from \JJapan\j 10 months earlier, and deployed and retrieved the OAST-Flyer satellite.
On STS-92, he will be the primary operator of the Shuttle's Remote Manipulator System robot arm supporting space station assembly tasks to be performed during four scheduled spacewalks. STS-92 will be the fifth assembly flight and will build on previous American and Russian flights beginning with the launch of the Functional Cargo Block (FGB) in June 1998. Prior to the arrival of Atlantis and the STS-92 crew, space station elements already delivered to orbit will include the FGB; Node 1 and two Pressurized Mating Adapters; the Service Module; and various logistical cargos that will be carried aboard the second Shuttle assembly mission in December 1998.
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"Satellite Measurements Suggest El Nino Is Brewing Again (29 May, 1997)",515,0,0,0
Simultaneous ocean measurements taken by two orbiting NASA science instruments suggest that another weather-disrupting El Nino condition may be developing in the Pacific Ocean, with the potential of altering global weather patterns next winter.
Sea-surface height measurements taken by the radar \Jaltimeter\j onboard the joint U.S.-French TOPEX/POSEIDON satellite and wind data collected by the NASA Scatterometer on \JJapan\j's Advanced Earth Observing Satellite (ADEOS) are being used together for the first time to diagnose changing oceanographic and atmospheric conditions in the tropical Pacific Ocean.
The El Nino phenomenon is thought to be triggered when steady westward blowing trade winds weaken and even reverse direction. This change in the winds allows the large mass of warm water that normally is located near \JAustralia\j to move eastward along the equator until it reaches the coast of South America. This displaced pool of unusually warm water affects where rain clouds form and, consequently, alters the typical atmospheric jet stream patterns around the world. The change in the wind strength and direction also impacts global weather patterns.
"NSCAT has observed two episodes of the reversal of the trade winds in the western Pacific, one at the end of December and one at the end of February. Both generated warm water masses, called Kelvin waves, that traveled across the Pacific and were measured by TOPEX/POSEIDON," said Dr. Lee-Lueng Fu, TOPEX/POSEIDON project scientist at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "Kelvin waves are often a precursor to a warm state of the tropical Pacific, sometimes leading to an El Nino.
"Whether an El Nino event will occur cannot be determined by just examining the satellite data," Fu continued. "A computer model that couples ocean-atmosphere data, like the one used by the National Oceanic and Atmospheric Administration (NOAA), is a necessary tool to issue scientifically based predictions. Now, for the first time, both TOPEX/POSEIDON and NSCAT are observing and providing the best, near real-time view of global ocean winds and sea level ever obtained. These observations will help NOAA's model to predict the occurrence of El Nino."
NOAA has issued an advisory regarding the presence of the early indications of El Nino conditions. A number of El Nino forecast activities supported by NOAA indicate the likelihood of a moderate or strong El Nino in late 1997. The forecast model operated at NOAA's National Centers for Environmental Prediction (NCEP) used data collected by the TOPEX/POSEIDON satellite.
"The use of TOPEX/POSEIDON data clearly improved our forecast for the winter of 1996-1997," said Dr. Ants Leetmaa, chief scientist at NCEP. "We currently use the data continuously for our operational ocean analyses and El Nino forecasts. The use of this data set enabled a clearer picture to be developed of the multi-year evolution of ocean conditions in the tropical Pacific that have resulted in the onset of the current warm episode. We have not yet had a chance to utilize the NSCAT data in the models but we anticipate that its use also will improve our forecast system."
"Since the beginning of the instrument's operation in September 1996, NSCAT has observed stronger than normal easterly winds in the central and western tropical Pacific, which might have piled up warm water in the west as indicated by the higher than normal sea level and sea surface temperature," said Dr. W. Timothy Liu, NSCAT project scientist at JPL. "This is usually a precursor of subsequent anomalous warming in the east. Kelvin waves moving across the Pacific do not necessarily mean El Nino, but we are studying how these seasonal phenomena like Kelvin waves are related to events like El Nino that occur over several years. TOPEX/POSEIDON and NSCAT will provide continuous, near real-time observations of the critical developments in the Pacific in the months to come."
The climatic event has been given the name El Nino, a Spanish term for a "boy child," because the warm current first appeared off the coast of South America around Christmas. Past El Nino events have caused unusually heavy rain and flooding in \JCalifornia\j, unseasonably mild winters in the Eastern United States and severe droughts in \JAustralia\j, Africa and \JIndonesia\j. Better predictions of extreme climate episodes like floods and droughts could save the United States billions of dollars in damage costs. El Nino episodes usually occur approximately every two to seven years.
The TOPEX/POSEIDON satellite uses an \Jaltimeter\j to bounce radar signals off the ocean's surface to get precise measurements of the distance between the satellite and the sea surface. These data are combined with measurements from other instruments that pinpoint the satellite's exact location in space. Every ten days, scientists produce a complete map of global ocean \Jtopography\j, the barely perceptible hills and valleys found on the sea surface. With detailed knowledge of ocean \Jtopography\j, scientists can then calculate the speed and direction of worldwide ocean currents.
The NASA scatterometer uses an array of stick-like antennas that radiate radar pulses in the Ku-band across broad regions of the Earth's surface. The way the radar signal bounces off the ocean's surface allows scientists to calculate both wind speed and direction. At any given time, NSCAT's antennas scan two swaths of ocean -- one on either side of the satellite's near-polar, sun-synchronous 500-mile (800- kilometer) orbit. The scatterometer takes 190,000 wind measurements per day, mapping more than 90 percent of the world's ice-free oceans every two days.
Both the TOPEX/POSEIDON \Jaltimeter\j and the NASA scatterometer are radar instruments, which allows them to operate 24 hours a day, collecting data day or night, regardless of sunlight or weather conditions.
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"Thermal Material Performs Well During First Flight Test (29 May, 1997)",516,0,0,0
A new thermal protection material designed to prevent \Jspacecraft\j from burning up during reentry into Earth's atmosphere performed extremely well during its first flight test. The ultra-high temperature ceramic material may someday revolutionize the approach that engineers take to the design and protection of aerospace vehicles.
A large amount of data on the thermal performance of the new material was collected on May 21 when a Mk 12A reentry vehicle blasted off from Vandenberg Air Force Base, near Lompoc, CA, at 4:27 a.m. EDT aboard a U.S. Air Force Minuteman III missile. The reentry vehicle, equipped with a 0.141-inch radius nose tip made from the new ceramic material, took an approximately 30-minute, 4,200-mile ride. Once out in space, the reentry vehicle was released, returning through Earth's atmosphere at blistering speed to a watery impact in the Kwajalein missile range of the Pacific Ocean. The sharp nose tip was instrumented with five heat sensing devices designed to provide information on how well the new ceramic material stood up to the burning temperatures of reentry. NASA engineers report that data was collected right up to the moment of splashdown.
"This was just the first data-gathering flight test for this new material," said Joan Salute, project manager for the mission located at NASA's Ames Research Center, Mountain View, CA. "However, initial results suggest that it was a complete success. After extensive data analysis, we should have good information on the potential of this ultra-high temperature ceramic material to support a radical new concept in aerospace vehicle design -- the use of hypersonic sharp leading edges."
This first test flight -- authorized only last December -- was accomplished in less than six months at a cost to NASA of $1.1 million. Plans for additional flights currently are under discussion. The potential benefits to be derived from the use of sharp leading-edge designs for \Jspacecraft\j and transatmospheric vehicles are tremendous. Sharp-body designs offer reduced drag, thereby providing substantial savings in the cost per pound expended to put objects into orbit. In addition, they provide a greatly enhanced lift-to-drag ratio, enabling what is called cross-range capability. This means that \Jspacecraft\j and transatmospheric vehicles can reenter Earth's atmosphere from any orbit and land at any location, unlike present blunt body vehicles. Finally, sharp leading edges minimize the number of free electrons that interfere with radio frequency transmissions and cause the communications blackout associated with the reentry of blunt body vehicles.
The history of leading edge design is instructive. In the 1940s, \Jaircraft\j routinely featured wings with sharp leading edges since that configuration was found to reduce drag at \Jsupersonic\j velocities (above the speed of sound, Mach 1, approximately 740 mph at sea level). However, with the coming of hypersonic flight (Mach 5 and faster) in the 1950s, the buildup of heat on the sharp leading edges of vehicle wings was so severe that available materials tended to burn up and their leading edges became blunted. Engineers took their cue from this natural blunting process in addressing the problem.
"All along, Mother Nature was trying to tell us how to deal with this situation," said Paul Kolodziej, lead engineer for the project at Ames. "Engineers realized that if they blunted the leading edge themselves in the design process, this would move the shock wave created during hypersonic flight forward and out, away from the vehicle, creating an air pocket in front of it." This air absorbs much of the heat associated with reentry, preventing the leading edge from becoming so hot and melting. "Because these new ceramic materials operate at ultra-high temperatures, we can now build sharp leading edges that don't melt during reentry along trajectories such as those flown by the Space Shuttle," Kolodziej said.
Blunt body design remains the norm today. In fact, it was the blunt body concept, originally developed at Ames by the late H. Julian Allen, that enabled development of space vehicles capable of withstanding the rigors of reentry. The problem is that blunt body vehicles have high drag and are not efficient. They therefore require large and expensive propulsion systems that impose a severe penalty in terms of cost, capability and performance. Developing an approach that permits the safe use of sharp leading-edge configurations for vehicles flying at hypersonic speeds holds the promise of yielding huge benefits. The key to achieving that payoff is the development of highly efficient thermal protection materials like those currently being evaluated by NASA engineers.
The ultra-high temperature ceramic material that just completed its first flight test has already proven out in ground-based testing in the arcjet facilities at Ames. In fact, the materials were shown to be very stable at temperatures in the range of 1,700 - 2,800 degrees \JCelsius\j (3,092-5,072 degrees Fahrenheit) in the presence of high- velocity dissociated air, such as is encountered during conditions of reentry. They also have been shown to be resistant to thermal shock and fatigue failure and, hence, reliable for repetitive operation and use over multi-mission lifecycles.
The development and testing of the new material is part of a joint program among NASA, Sandia National Laboratories and the US Air Force called Slender Hypervelocity Aerothermodynamic Research Probes, or SHARP.
NASA's participation in the SHARP program is funded by the Headquarters Office of Aeronautics and Space Transportation Technology. The objective of the effort is to demonstrate the viability of sharp leading edges for space vehicles and possible future transatmospheric passenger \Jaircraft\j that may have the capability to fly out of Earth's atmosphere and into the fringes of space and back. The SHARP program is a key element supporting NASA's overall mission to expand the nation's frontiers and capabilities in the aeronautics and space domain.
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"Polar Spacecraft Images Support Theory Of Interplanetary Snowballs Spraying Earth's Upper Atmosphere (28 May, 1997)",517,0,0,0
Images from NASA's Polar \Jspacecraft\j provide new evidence that Earth's upper atmosphere is being sprayed by a steady stream of water-bearing objects comparable to small comets.
Using Polar's Visible Imaging System (VIS), a research team led by Dr. Louis A. Frank of the University of \JIowa\j in \JIowa\j City has detected objects that streak toward Earth, disintegrate at high altitudes and deposit large clouds of water vapor in the upper atmosphere. Frank's research is being reported in a news briefing at 10 a.m. today at the spring meeting of the American Geophysical Union at the Convention Center in Baltimore, MD.
The incoming objects, which Frank estimates to be the size of a small house, pose no threat to people on Earth, nor to astronauts in orbit. "They break up and are destroyed at 600 to 15,000 miles above the Earth," Frank noted. "In fact, this relatively gentle 'cosmic rain' -- which possibly contains simple organic compounds -- may well have nurtured the development of life on our planet."
"This is an intriguing result that requires further scientific investigation," said Dr. George Withbroe, science director for the Sun-Earth Connection program in NASA's Office of Space Science. "We need to look closely at measurements from other sensors to find out if they see related signatures in the atmosphere, now that we have learned more about what to look for."
The Polar cameras have imaged trails of light in both ultraviolet and visible wavelengths as the objects disintegrate above the atmosphere. Using a filter that detects visible light emitted only by fragments of water molecules, Frank has shown that the objects consist primarily of water.
"The Polar results definitely demonstrate that there are objects entering the Earth's upper atmosphere that contain a lot of water," commented Dr. Thomas M. Donahue, a noted atmospheric physicist and professor at the University of \JMichigan\j in Ann Arbor.
"The images show that we have a large population of objects in the Earth's vicinity that have not been detected before," said Frank, who designed the VIS instrument. "We detect these objects at a rate that suggest Earth is being bombarded by five to 30 small comets per minute, or thousands per day." Comets are known to contain frozen water and are sometimes called "dirty snowballs".
Frank's new observations are consistent with a controversial theory he proposed in 1986 to explain the existence of dark spots, which he termed "atmospheric holes", in images of the sunlit atmosphere of the Earth. He first detected these holes while analyzing data from an ultraviolet imager flown on NASA's Dynamics Explorer 1 \Jspacecraft\j. He theorized that the holes were caused by the disintegration of small icy comets in the upper atmosphere. The water vapor they produce momentarily absorbs the ultraviolet solar radiation scattered from oxygen atoms in the upper atmosphere, preventing it from reaching his camera and resulting in a dark spot on the image. These holes have diameters of 15 to 25 miles.
His theory of a new class of objects in the Solar System ignited a wide-ranging controversy. Many colleagues discounted the appearance of the holes as an instrumental problem. But the new images from Polar also include observations of atmospheric holes in much greater detail than before, suggesting that they are real. "These results certainly vindicate Lou Frank's earlier observations", said Donahue.
"These remarkable images cap a great first year for Polar," added Dr. Robert Hoffman, Project Scientist for Polar, which is operated and managed by NASA's Goddard Space Flight Center, Greenbelt, MD. "I am pleased that Polar's instruments were able to actually detect these objects streaking towards the Earth and disintegrating into clouds of water vapor. They give scientists a fascinating new and important phenomenon to take into account in theories of Solar System evolution."
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"X-34 Systems Design Freeze Completed (22 May, 1997)",518,0,0,0
Government and industry managers yesterday successfully completed a three-day systems design freeze of X-34, the next technology demonstrator vehicle in NASA's Reusable Launch Vehicle stable.
X-34, a reusable, suborbital, air-launched vehicle, will fly at speeds approaching Mach 8 (eight times the speed of sound) at altitudes up to 50 miles to continue demonstrations of new, reusable launch vehicle technologies aimed, ultimately, at dramatically reducing the cost of transporting payloads into space.
The review, conducted at Orbital Sciences Corporation in Dulles, VA, essentially anchors the design of X-34 systems -- for example, structures, guidance, navigation and control, avionics, thermal protection systems and main propulsion systems -- that will allow the program to proceed with fabrication and manufacturing of these systems. Initial flights, scheduled for launch and landing within air space at the White Sands Missile Range, NM, will begin in late 1998. Up to 25 flights within a year's time are scheduled; some also will take place in \JFlorida\j to demonstrate subsonic landing and thermal protection system performance through inclement weather conditions.
X-34 bridges the gap between earlier subsonic Clipper Graham (DC-XA) flights in 1996, and the larger and higher performance X- 33 Mach 15 demonstrator, which will begin flights in early 1999. The winged, reusable, single-stage X-34 vehicle -- propelled by a kerosene/liquid oxygen engine developed, tested and provided by NASA -- will demonstrate key technologies, including composite primary and secondary airframe structures; cryogenic \Jinsulation\j and propulsion system elements; advanced thermal protection systems and materials; low cost avionics, including differential Global Positioning and Inertial Navigation Systems; as well as key operations technologies such as integrated vehicle health- monitoring and automated checkout systems.
Six NASA Centers, Department of Defense installations (White Sands Missile Range and Holloman Air Force Base) and an industry team led by the prime contractor, Orbital Sciences Corporation, are playing key roles in the development and eventual flight testing of the X-34 technology testbed demonstrator. The program is managed by the Marshall Space Flight Center, \JHuntsville\j, AL.
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"Tailless Aircraft (Remotely Piloted) Completes First Flight",519,0,0,0
A NASA/McDonnell Douglas remotely piloted, tailless \Jaircraft\j successfully completed its first flight on May 17 at NASA's Dryden Flight Research Center, Edwards, CA. The lack of vertical tails greatly enhances the stealthy characteristics of the airplane, and holds promise for greater agility than is currently available in existing military fighter \Jaircraft\j.
Called the X-36, the subscale research \Jaircraft\j lifted off from Rogers Dry Lake at 7:08 a.m., PDT. The \Jaircraft\j flew for five minutes and reached an altitude of approximately 4,900 feet. An additional 24 test flights of the X-36 are scheduled at Dryden during the next six months.
"We thought the flight was outstanding; we are beginnning to show what the fighter \Jaircraft\j of the future will look like," said Rod Bailey, X-36 program manager. When we saw this airplane lift off, we saw the shape of airplanes to come."
NASA's Ames Research Center, Moffett Field, CA, leads the X- 36 program, and has technical responsibility for continued development of some of the critical technologies needed for future tailless, stealthy fighter \Jaircraft\j.
There are two 28-percent-scale X-36s, which are remotely piloted jets built by the McDonnell Douglas Corporation's Phantom Works division in St. Louis, MO, and are designed to fly without the traditional vertical and horizontal tails found on most \Jaircraft\j. Each \Jaircraft\j measures 18 feet long, 3 feet high, has a 10-foot wing span, and weighs 1,250 pounds. Each \Jaircraft\j is powered by a Williams Research F112 turbofan engine that provides 700 pounds of thrust.
The X-36 \Jaircraft\j are remotely controlled by a pilot in a ground station cockpit, complete with a heads-up display . The pilot-in-the-loop approach eliminates the need for expensive and complex autonomous flight control systems. The design reduces weight and drag of the \Jaircraft\j and explores new flight control technologies. The \Jaircraft\j use split ailerons to provide yaw control, as well as raising and lowering in a normal fashion to provide roll control. The X-36 also incorporates a thrust vectoring system.
"The flight control system functioned flawlessly and we look forward to subsequent flights to demonstrate the full range of maneuverability of the aircraft," said Mark Sumich, X-36 project manager.
"We knew within five to ten seconds into the flight that we had a good flying airplane," said Gary Jennings, McDonnell Douglas X-36 program manager. "Flying in a simulator is one thing, but until you actually fly the airplane, you don't really know how it will handle. Today we found out that it handled extremely well."
NASA Ames and McDonnell Douglas developed the technologies required for a tailless fighter beginning in 1989. In 1993, McDonnell Douglas proposed the remotely piloted \Jaircraft\j technology demonstration to validate the technologies in a real flight environment. In 1994, McDonnell Douglas began fabrication of the two \Jaircraft\j in their rapid prototyping facility in St. Louis. The project was jointly funded under a roughly 50/50 cost- sharing arrangement between NASA and McDonnell Douglas. The combined program cost for the development, fabrication, and flight testing of the two prototype \Jaircraft\j is approximately $20 million.
"The first flight went very well; it was just textbook perfect," said Larry Walker, X-36 Project pilot. "It was a nice takeoff and the handling was great. I knew instantly that it was a nice flying airplane. I see no obstacles in the future for this type of technology."
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"NASA Technology Spawns New Civilian Tiltrotor Aircraft (20 May, 1997)",520,0,0,0
After 20 years of research and development, an \Jaircraft\j known as the tiltrotor, which combines the speed and range of a turboprop airplane with the vertical takeoff and landing capability of a \Jhelicopter\j, is poised to enter commercial service.
NASA's Ames Research Center, Mountain View, CA, which manages NASA's Short Haul Civil Tiltrotor program, will look back on 20 years of development during a special program there, celebrating the 20th anniversary of the first flight of the XV-15 Tiltrotor Research \JAircraft\j, on Thursday, May 22 starting at 11:30 a.m. EDT in the Main Auditorium (N-201). Former NASA Deputy Administrator Dr. Hans Mark will discuss "The Tiltrotor Story: A Classic American Technology Development" from noon until 1 p.m. EDT, followed by tours of key tiltrotor research facilities from 1:30 p.m. to 3 p.m. EDT. Media are invited to attend.
After pioneering work in the past, Ames is continuing its work on the tiltrotor concept with studies on the development of a 40-passenger civil tiltrotor. A 1995 Department of Transportation study concluded that a new air transportation system based on civil tiltrotor technology could be created and that a 40- passenger civil tiltrotor is feasible technically, economically and environmentally.
In July 1999, Bell \JHelicopter\j Textron Inc., Forth Worth,TX, and Boeing Defense and Space Group's Helicopters Division, Philadelphia, PA, plan to conduct the first flight of the nation's first civilian tiltrotor, the Bell-Boeing 609, a nine-passenger executive transport \Jaircraft\j. The company is currently building the world's first production tiltrotor, the V- 22 Osprey, a military \Jaircraft\j that will carry 24 combat-ready troops. Bell-Boeing estimates a domestic market of about 1,000 \Jaircraft\j for the 609 over the next 20 years.
In addition to the V-22 and 609, Sikorsky \JAircraft\j Corp., a subsidiary of United Technologies Corp., Stratford, CT, is actively pursuing the technology for the next generation of tiltrotors, a variable-diameter tiltrotor, which could further develop the tiltrotor's capability to match those of both a \Jhelicopter\j and a turboprop airplane. Both the 609 and V-22 are direct descendants of the XV-15 Tiltrotor Research \JAircraft\j, developed under a joint NASA/Army/Bell program at Ames.
"The XV-15 was the proof-of-concept vehicle for tiltrotor aircraft," said William J. Snyder, manager of the Advanced Tiltrotor Transport Technology Office at Ames. "The XV-15, which is still flying at Bell \JHelicopter\j Textron's flight test center in Arlington, TX, established the technology \Jdatabase\j for development of the world's first production tiltrotor, the Bell-Boeing V-22 Osprey and the Bell-Boeing 609," Snyder said. "It's the single- most important aeronautical development fully attributable to Ames."
Tiltrotor technology, as exemplified by the XV-15, has been one of the most successful aeronautics research programs in NASA's history. "The XV-15 research \Jaircraft\j's research contributions did not end with the initiation of the V-22; it continues today to be the primary civil testbed for flight evaluation of new tiltrotor technology, almost 20 years after its first flight," Snyder said.
The XV-15 combines standard \Jaircraft\j cruise flight with vertical takeoff and landing (VTOL) and short takeoff and landing (STOL) capabilities. A unique feature of the XV-15 are the two large, three-bladed proprotors mounted at the tips of the wings. For takeoff, the proprotors and their engines are rotated straight up, enabling the \Jaircraft\j to lift off vertically, like a \Jhelicopter\j. Once off the ground, the proprotors and engines rotate forward like a conventional \Jaircraft\j enabling it to cruise for more than two hours.
These features enable the XV-15 to take off and land at small landing terminals or vertiports rather than at conventional airports and thus give them a significant advantage over turboprop \Jaircraft\j. "By removing short haul turboprops from the nation's major airports and replacing them with 40-passenger tiltrotor transports, airport congestion and delays could be substantially reduced and substantial savings in time and dollars could be realized by passengers and by the air transport system," Snyder said.
"The XV-15 is much quieter in cruise flight than turboprops. Further, flight research has shown that by converting the proprotors from airplane to \Jhelicopter\j mode in a certain way, the noise footprint of the XV-15 has been found to be reduced by 30 percent over that made by conventional helicopters -- a very significant accomplishment that is important for community acceptance," said John Zuk, national lead for civil tiltrotor studies in the Advanced Tiltrotor Transport Technology Office at Ames.
Twenty years ago this month, the XV-15 successfully completed its first flight at the Bell \JHelicopter\j plant in \JTexas\j under Ames' management. The \Jaircraft\j had a 15-year career in flight research at Ames before being sent to Bell in 1994 for further research.
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"IMAX To Document Space Station Assembly in 3-D (20 May, 1997)",521,0,0,0
The historic, on-orbit construction of the International Space Station will be documented in 3-D by the Imax Corporation in a large-format (70-mm) feature film to be seen around the world.
This will be the first 70-mm space film to be captured in 3- D, a breakthrough made possible by Imax's current development of a 3-D movie camera that will meet the exacting requirements and strict limitations of flying on \Jspacecraft\j. The film also will be distributed in Imax's 2-D format.
"Our astronauts have said that previous Imax films are the closest thing to actually being in space," NASA Administrator Daniel S. Goldin said. "Capturing the assembly of the International Space Station in this realistic and compelling format will help NASA share this experience with the public. After all, the station belongs to the public, and they have a right to watch it become reality."
The feature film will be made under a Space Act Agreement between NASA and Imax. NASA will own the copyright on all film footage shot in space under the agreement. The agency, in turn, has granted Imax a limited license to create a large-format feature film using the footage. All of the footage eventually will be made available publicly. Meanwhile, NASA retains the right to disseminate, at any time, still photos and videotape segments made from the 70-mm footage.
Under previous Space Act Agreements, Imax has produced three documentaries about the space program: "The Dream is Alive," "Blue Planet" and "Destiny in Space." Collectively, these films have been viewed by more than 60 million people and still are being shown across the United States and internationally. The announcement of the International Space Station film was made in conjunction with today's premiere in Washington, DC, of a fourth Imax space film, "Mission to Mir."
"Mission to Mir" tells the story of the Space Age relationship between the United States and \JRussia\j, beginning with fierce Cold War competition and leading to today's cooperative space program with \JRussia\j, including the docking missions between the Space Shuttle and Space Station Mir. The footage from those docking missions--part of the ongoing "Phase One" program that will pave the way for U.S.-Russian cooperation on the International Space Station -- was obtained by Imax under contract to NASA and is the centerpiece of the newly released feature film.
The International Space Station is a partnership of 15 countries, including the United States, \JRussia\j, Canada, \JJapan\j and eleven member nations of the European Space Agency. On-orbit assembly is scheduled to begin with the launch of the first element by October 1998, and will be completed by 2002.
NASA will acquire images of the assembly process in other, non-Imax formats--including \Jtelevision\j, videotape and still photos--and disseminate them publicly as soon as they are available.
There are now 153 Imax theaters in 22 countries, including 74 in the United States. Twenty-nine of the theaters are capable of showing 3-D films.
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"Astronaut Shannon Lucid Receives Russian Order Of Friendship Medal (19 May, 1997)",522,0,0,0
Astronaut Shannon Lucid received the Order of Friendship medal from Russian President Boris Yeltsin during a ceremony at the \JKremlin\j honoring Russian Cosmonauts and international astronauts who have flown on the Russian space station Mir in 1996.
The Order of Friendship medal is one of the highest Russian civilian awards and the highest award that can be presented to a non-citizen.
Dr. Lucid spent a U.S. record of 188 days in space last year. She was launched to Mir on the Space Shuttle Atlantis (STS-76) on March 22, 1996, and returned to Earth on Atlantis (STS-79) on September 26, 1996.
Lucid was selected as an \Jastronaut\j in 1978, the first \Jastronaut\j class to include women to train for Space Shuttle missions. Lucid has flown on five space missions including her Mir mission. Dr. Lucid holds pilot ratings in commercial, instrument and multi-engine \Jaircraft\j.
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"Hubble Finds Cloudy, Cold Weather Conditions For Mars Bound Spacecraft (20 May, 1997)",523,0,0,0
As two NASA \Jspacecraft\j speed toward a mid-year rendezvous with Mars, astronomers using the Hubble Space \JTelescope\j are providing updated planetary weather reports to help plan the missions.
Hubble's new images show that the "Martian invasion" of \Jspacecraft\j will experience considerably different weather conditions than seen by the last U.S. \Jspacecraft\j to land on Mars 21 years ago.
Martian atmospheric conditions will affect the operation of both the Mars Pathfinder landing on July 4, and the September 11 arrival of the Mars Global Surveyor which will map the planet from orbit. Hubble images taken barely three weeks apart, on March 10 and March 30, reveal dramatic changes in some local conditions, and show overall cloudier and colder conditions than Viking encountered two decades ago.
"Because Pathfinder uses the atmosphere to decrease its velocity for landing, and because the lander and rover are solar- powered, understanding the state of the atmosphere prior to landing is important," said Dr. Matthew Golombek, Pathfinder project scientist at NASA's Jet Propulsion Laboratory, Pasadena, CA.
"On July 4, Mars Pathfinder will enter the atmosphere directly from approach and slow itself behind an aeroshell with a parachute, small solid rockets and giant airbags. The lander carries a small rover to explore the surface and investigate the kinds of materials present. Hubble images of Mars are helping us to adjust our flight path for landing and effectively plan surface operations," said Golombek.
"It's not the dusty Mars of the Project Viking days (mid 1970s to early 1980s) or the habitable oasis of science fiction stories," says Todd Clancy of the Space Science Institute in Boulder, CO. "We're finding a Mars that's colder, clearer, cloudier. Hubble is rapidly changing our view of Mars' environment. The planet's weather apparently has a flip-side to it."
Hubble's findings also offer new insights into the differences and similarities of weather on the other terrestrial planets. "The planets are similar in many important ways, so the very major differences between them are interesting from a viewpoint of understanding \Jmeteorology\j better," said team leader Phil James of the University of Toledo in Ohio. "Hubble is allowing us to look at Mars in ways never before seen."
In September, NASA's Mars Global Surveyor will skim across the upper Martian atmosphere to slow down by friction and enter orbit around the red planet. Atmospheric density is a key factor in precisely executing this complex and delicate aerobraking maneuver. Hubble is ideal for tracking regional dust storms which could pose a threat to Surveyor by drastically changing the planet's air density. Such storms can cause a tenfold increase in the Martian atmosphere's drag at 60 miles above the surface.
Comparing the appearance of Mars to that in earlier \Jspacecraft\j observations, Hubble has found some areas of the Martian surface that have been changed dramatically by wind-blown dust. The most prominent example is the "classic dark feature" called Cerberus, which is roughly the size of \JCalifornia\j (800 by 250 miles). This feature has been seen as a low \Jalbedo\j (dark) region by ground-based telescopic observers since early this century, and was studied in detail by the Mariner 9 and Viking orbiters in the 1970s.
In Hubble's view, only three dark splotches remain, probably related to dark sand being carried out of craters by the wind. The astronomers think that dust storms in the region have covered the formerly dark surface with bright dust, effectively erasing Cerberus from the map.
Hubble is ideally suited for long-term study of Mars. When Mars has been closest to Earth, Hubble has resolved surface details as small as 25 miles across. This allows astronomers to track subtleties in the shifting cloud patterns and periodic dust storms. This planet-wide "weather satellite" view is complementary to the close-up views which will be provided by Mars Pathfinder and Mars Global Surveyor.
The Space \JTelescope\j Science Institute is operated by the Association of Universities for Research in \JAstronomy\j, Inc. (AURA), for NASA, under contract with the Goddard Space Flight Center. The Hubble Space \JTelescope\j is a project of international cooperation between NASA and the European Space Agency (ESA).
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"Shuttle Mission, Ukranian Payload Specialists Selected (16 May, 1997)",524,0,0,0
NASA today announced the selection of Col. Leonid Kadenyuk, the first Ukrainian to fly on the U.S. Space Shuttle, as the primary payload specialist for the fourth U.S. \JMicrogravity\j Payload flight scheduled for a November 1997 launch on the Space Shuttle Columbia as mission STS-87.
The announcement comes as the first session of the U.S./Ukraine Binational Commission convened today at the White House. The Commission is co-chaired by Vice President Al Gore and Ukrainian President Leonid Kuchma.
NASA also named another Ukrainian, Dr. Yaroslav Pustovyi, to serve as an alternate. As an alternate payload specialist, Pustovyi will undergo the same training as Kadenyuk and will be ready to serve on the mission crew if necessary.
Kadenyuk will conduct the Collaborative Ukrainian Experiment (CUE), a series of 11 Shuttle middeck experiments focusing on the effects of \Jmicrogravity\j on plant growth and \Jpollination\j.
The CUE project resulted from the November 1994 summit meeting between President Bill Clinton and Ukrainian President Leonid Kuchma. At that time, the two presidents signed the Agreement Between the United States of America and \JUkraine\j on Cooperation in the Exploration and Use of Outer Space for Peaceful Purposes and announced that a Ukrainian representative would fly on a future Space Shuttle mission. NASA and the National Space Agency of \JUkraine\j (NSAU) were designated as implementing agencies for carrying out U.S./Ukrainian cooperative space activities.
U. S. and Ukrainian scientists will collaborate on the plant experiments to be carried out by the Ukrainian payload specialist during the flight. The 11 CUE investigations, which use the Shuttle's Plant Growth Facility and Biological Research in Canisters hardware, consist of five primary and six secondary experiments.
Six U.S. and 16 Ukrainian principal investigators are collaborating on the experiments, which will study the effect of \Jmicrogravity\j on \Jpollination\j, \Jfertilization\j and flowering of plants and seedlings. The research also furthers the study of previously observed abnormal growth and developmental phenomena involving plants in space.
A significant element of CUE is its large educational component. High school students in both countries will perform a plant \Jpollination\j experiment on the ground at the same time the experiment is conducted on the Shuttle. The payload specialist will discuss the experiment with students during scheduled educational downlinks from the Shuttle.
Kadenyuk was selected as a Russian \Jcosmonaut\j in 1976. He has trained for and studied systems for the Soyuz, Soyuz-TM, MTKK, Buran, and Space Stations Salyut and Mir. He was born in the village of Klishkovtsiy in the Khotinsky Region of the Chernovitsky District. He is currently employed by the \JKiev\j Botanical Institute of the National Academy of Sciences of \JUkraine\j.
Pustovyi is a graduate of the Military Space-Engineer Academy, and received a doctorate in Radio Physics from Kharkiv State University. He is employed by the Institute of Magnetism at the National Academy of Sciences of \JUkraine\j.
The other members of the STS-87 crew are Commander Kevin Kregel, Pilot Steven Lindsey (Major, USAF) and Mission Specialists Winston Scott (Captain, USN), Kalpana Chawla, Ph.D. and Takao Doi, Ph.D. (NASDA).
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"Space Station Control Board Approves New Assembly Schedule (15 May, 1997)",525,0,0,0
The International Space Station Control Board has approved a new baseline schedule that keeps the assembly sequence intact and targets the first station launch for June 1998 -- an eight-month delay from the previous schedule.
As announced by NASA in April, the revision in the station's assembly schedule is the direct result of funding delays in the construction of the Service Module, the primary Russian contribution to the early assembly of the station and a component that will supply the early living quarters, life support systems and propulsion. Russian-funded work on the Service Module now has fully resumed as a result of Russian government funding, and it is rapidly progressing.
"The recent completion of a major Russian general designers review for the Service Module, in which I participated, and full Russian funding of the work, gives us high confidence that the Service Module can meet a revised launch date of December 1998," Program Manager Randy Brinkley said. "The Russian Space Agency has been extremely forthcoming in its dealings with NASA on this subject, and they and their contractors have gone out of their way to demonstrate their resolve to meet their commitment. Based on what I saw and heard during my most recent visit to \JRussia\j, I have every confidence that RSA and the Russian space industry are fully committed to meeting their obligations for the Service Module and ISS."
Although the first station launch, the launch of the Functional Energy Block (FGB) on a Russian Proton booster, is delayed by eight months in the new schedule, the beginning of full-fledged research flights to the station in August 1999 -- the end of Phase 2 of the program -- is a delay of only four months from what previously had been planned. To enhance the station's capabilities, modifications will be made to the FGB to allow it to be refueled and to accommodate dockings by Russian Soyuz capsules.
Despite delays in the Russian hardware, work has continued on all U.S. station components, and the first U.S.-built component, Node 1, will be delivered to the Kennedy Space Center this summer for pre-launch testing and processing. Node 1 will be launched on Space Shuttle mission STS-88 in July 1998 to be mated to the already-orbiting FGB. Because U.S. components such as the laboratory module, the first truss segment and the first solar array remain on schedule, NASA will take advantage of the extra time in assembly to pursue integrated testing of components after they are shipped to KSC.
"A little more than a year from now, we'll launch the first component. About a year and a half from now, we will launch the first crew. Only two years from today, that first crew will be finishing up the first tour onboard. Four Shuttle assembly flights will already have been completed. And we'll be only a few months from completing Phase 2 of the program," Brinkley said. "This \Jspacecraft\j is on deck, and we are number one on the runway."
Other highlights of the new schedule, called the ISS Assembly Sequence, Rev. C, include:
* In January 1999, the second Space Shuttle assembly mission, designated STS-92 and assembly Flight 3A, will be launched and later followed by a Russian Soyuz \Jspacecraft\j carrying the first crew -- ISS Commander Bill Shepherd, Soyuz Commander Yuri Gidzenko and Flight Engineer Sergei Krikalev -- to begin a permanent human presence on the ISS.
* Two Space Shuttle flights have been added to the assembly sequence to increase margin and add flexibility. The flights, called flight 2A.1 in late 1998 and flight 7A.1 in late 1999, may be used to offload cargo from adjacent assembly flights and assist with U.S. outfitting of the station.
* At present, NASA plans to continue the conversion of a Naval Research Laboratory stage into an Interim Control Module (ICM), that could be used to augment the station's future propulsion capabilities if needed by being attached to either the Functional Cargo Block (FGB) or the Service Module.
* Assembly flight 13A, a shuttle mission that carries two additional solar arrays, has been realigned earlier in the assembly sequence and will provide additional power for scientific activities and station assembly.
* Launch date options for the European Space Agency's Columbus Orbital Facility remain under evaluation. While these options are analyzed, the launch dates for all flights after Utilization Flight 5 in June 2002 will remain under review; however, the U.S. Habitat Module will be fully outfitted by December 2002 regardless of the launch options chosen. These launch dates are expected to be set at a Space Station Control Board meeting in Fall 1997.
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"NASA And Japan To Cooperate On Asteroid Sample Return Mission (14 May, 1997)",526,0,0,0
NASA and \JJapan\j's Institute of Space and Astronautical Science (ISAS) have agreed to cooperate on the first mission to collect samples from the surface of an asteroid and return them to Earth for in-depth study.
Known as MUSES-C, the mission will be launched on a Japanese M-5 launch vehicle in January 2002 from Kagoshima Space Center, \JJapan\j, toward a touchdown on the asteroid \JNereus\j in September 2003. A NASA-provided miniature robotic rover will conduct in- situ measurements on the rocky surface.
The asteroid samples will be returned to Earth by MUSES-C via a parachute-borne recovery capsule in January 2006, just weeks after a NASA mission named Stardust is expected to return collected \Jcomet\j dust samples to Earth.
NASA and ISAS will cooperate on several aspects of the mission, including mission support and scientific analysis. Dr. Atsuhiro Nishida, Director General of ISAS, and Dr. Wesley T. Huntress Jr., NASA Associate Administrator for Space Science, signed a summary of discussions outlining the cooperation on MUSES-C during a May 2 meeting in Washington, DC.
"This ambitious mission is an opportunity for two spacefaring nations to combine their expertise and achieve something truly fantastic," said Dr. Jurgen Rahe, director of Solar System Exploration at NASA Headquarters. "The rover will be the smallest ever flown in space. With a successful mission, we will have the first direct insight into the composition of the materials that helped form the rocky inner planets more than four billion years ago."
With a mass of less than 2.2 pounds, the asteroid rover technology experiment would be a direct descendant of the technology used to build the Sojourner rover due to land on Mars with the Mars Pathfinder lander on July 4 of this year. The rover will carry two science instruments: a visible imaging camera and a near-infrared point spectrometer. It will be designed and built by NASA's Jet Propulsion Laboratory, Pasadena, CA.
Other U.S. contributions to MUSES-C include testing of its reentry capsule heat shield at NASA's Ames Research Center, Mountain View, CA, and navigation and tracking support from the ground-based Deep Space Network. NASA also will provide co- investigators to join in the mission's science, and the Agency will share in access to the asteroid samples.
Nereus is a small, near-Earth asteroid roughly one mile in diameter. It was discovered in 1982. At its closest point to the Sun, its orbit takes it just inside the orbit of the Earth.
MUSES-C will continue a recent string of missions focused on \Jasteroids\j. NASA's \JGalileo\j mission, now in looping orbit around Jupiter, flew by two \Jasteroids\j -- Gaspra and Ida -- on its way to the giant gas planet, discovering a small moonlet around one of them. The Near Earth Asteroid Rendezvous \Jspacecraft\j, a NASA Discovery Program mission built and operated by the Johns Hopkins University's Applied Physics Laboratory, will fly by the asteroid Mathilde on June 27 on its way to orbit the large asteroid Eros in 1999.
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"NASA Earth Science Research Aircraft Soars To New Heights (13 May, 1997)",527,0,0,0
A NASA ER-2 \Jaircraft\j, complete with a full array of science instrument packages, recently conducted its first operational mission at an altitude of 70,000 feet, a key region for atmospheric research.
The vehicle currently is on deployment to \JAlaska\j for missions over the North Pole in support of a project known as POLARIS, which stands for Photochemistry of Ozone Loss in the \JArctic\j Region In Summer. NASA has two such vehicles in its ER-2 fleet based at the Agency's Ames Research Center, Mountain View, CA. The ER-2 is a civilian version of the U-2 aerial reconnaissance plane.
A program to modernize the vehicles by making them lighter, more fuel efficient and more productive was completed recently. Over the next year, these improvements will increase significantly the size of science payloads and enhance the altitude performance of the ER-2s in support of NASA's Mission to \JPlanet\j Earth enterprise.
Following a recent flight, Jim Barrilleaux, an ER-2 pilot and acting chief of Ames' High Altitude Missions branch, expressed his surprise at the magnitude of difference in feel and performance of the vehicle. "It flies like a completely new aircraft," he said. "It feels really tight."
Earth scientists also are excited about the enhanced capability. "It is really critical that we have access to consistent measurements at this key altitude, which is an intermediate region between \Jaerosol\j particle-driven processes measured by standard aircraft-based sensors and gas-phase processes monitored by orbiting satellites," said Dr. Michael Kurylo, manager of the Upper Atmosphere Research Program at NASA Headquarters, Washington, DC.
The first deployment of an upgraded ER-2 currently is being conducted over the North Pole through May 15. The present POLARIS payload utilizes two large superpods attached to the wings. This more than doubles the available volume for science instruments, while still permitting operation at enhanced altitudes of 70,000 feet and above, according to flight engineers.
The POLARIS mission is seeking to understand the fundamental chemistry that dominates the naturally occuring seasonal reduction of ozone over the pole in the course of the \JArctic\j summer. Many of the chemical reactions in which project scientists are interested in occur at altitudes in the 75,000-foot range. Now, even a fully loaded ER-2 can operate approximately 2,500 feet higher than previously possible due to lower fuel requirements and lighter \Jaircraft\j weight. This increased altitude capability permits extension of in-place measurements for validating and upgrading existing models of the upper atmosphere.
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"Hyakutake X-Rays Show Ability To Monitor Comets And Solar Wind (9 May, 1997)",528,0,0,0
A supercomputer simulation of \JComet\j Hyakutake's interaction with the solar wind demonstrates that resulting X-ray emissions can be used to monitor comets and solar wind phenomena, NASA- funded researchers write in today's issue of "Science."
The simulation was conducted using an Earth sciences supercomputer at the NASA Goddard Space Flight Center, Greenbelt, MD. The results match and explain March 27, 1996, observations of \JComet\j Hyakutake by \JGermany\j's ROSAT satellite, the first detection of X-ray emissions from any \Jcomet\j. The model also supports a leading theory for how the X-rays are generated.
"Cometary X-rays present a potentially powerful new tool to monitor \Jcomet\j activity far from Earth, as well as the composition and flux of the solar wind," said co-author Dr. Tamas Gombosi of the University of \JMichigan\j, Ann Arbor. "By capturing these X-rays' detailed energy spectrum, it might be possible to monitor the propagation and evolution of spectacular solar wind phenomena, such as the coronal mass ejections seen this January and April."
About one percent of the solar wind, which flows from the Sun out past Pluto, is composed of minor ions: atoms (such as oxygen, carbon and neon) that have been nearly stripped of their electrons and thus have a high positive charge. Dr. Thomas Cravens of the University of Kansas theorizes that these minor ions steal electrons from neutral atoms and molecules of cometary origin. The electrons are first seized in excited states, traveling in the ions' outer orbitals. As the electrons fall to lower orbitals, Cravens' theory asserts that X-rays are emitted, in addition to other forms of radiation.
"Considering the magnitude and shape of the emission, we believe the most satisfactory theory to be this mechanism of charge exchange excitation," Gombosi said. "Other explanations produce neither the crescent pattern nor the intensity observed by ROSAT and duplicated by our simulation."
Within this pattern, some \Jelectron\j orbital transitions emit distinct wavelengths of X-rays that can be measured. The computer simulation shows that the overall X-ray spectrum for \JComet\j Hyakutake depends mainly on the solar wind composition, and not on the \Jcomet\j. Because of this independence, researchers can determine the relative size of the \Jcomet\j's atmosphere from the proximity of the brightest X-rays to the icy nucleus.
"In Hyakutake, the brightest X-ray region was 18,700 miles (30,000 kilometers) ahead of the \Jcomet\j, on the Sun side," said University of \JMichigan\j co-author Dr. Michael Combi. "If the \Jcomet\j has enough of an atmosphere, the solar wind minor ions recombine with electrons far from the nucleus. If the \Jcomet\j were producing less atmospheric gas, the place of maximum emission would be closer to the nucleus," Combi said.
This theory will be tested on \JComet\j Hale-Bopp, which is scheduled to be observed by \JJapan\j's ASCA X-ray satellite this September. "Comet Hale-Bopp should have the emission shifted further sunward; it is bigger than Hyakutake," Combi said.
Active comets are typically first observed in visible light at large distances from the Sun. After discovery, the orbits of comets can be established with very high accuracy as they pass through the inner solar system. "If X-rays are observed from the known location of a \Jcomet\j, one can conclude with great confidence that the X-rays originated from the comet," Gombosi said.
The University of \JMichigan\j team used March 27, 1996, solar wind density measurements from NASA's WIND \Jspacecraft\j. Their model first considers the global interaction of the solar wind with the \Jcomet\j. It projects the \Jcomet\j into a three-dimensional grid that automatically applies finer resolution where more activity occurs. This physics component predicts the deflective paths and speed of the solar wind traveling through the \Jcomet\j.
Other co-authors of the "Science" paper are Roman Haberli, Darren De Zeeuw and Kenneth Powell. The University of \JMichigan\j team is one of nine Grand Challenge Investigations funded by the NASA High Performance Computing and Communication Program's Earth and Space Sciences Project. Additional funding comes from NASA's Office of Space Science, the National Science Foundation and the Swiss National Science Foundation.
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"Spacecraft Watch For Comet Hale-Bopp Tail Disruption (5 May, 1997)",529,0,0,0
A fleet of \Jspacecraft\j for the International Solar Terrestrial Physics (ISTP) program is watching for a break in \JComet\j Hale- Bopp's plasma ion tail.
"Preliminary estimates indicate that it may happen in the next few days," said Dr. Mario Acuna, lead scientist for ISTP at NASA's Goddard Space Flight Center (GSFC), Greenbelt, MD. Goddard is the focal point for many of the ISTP investigations.
Amateur astronomers around the world were put on watch last week when Dr. Bill Farrell, co-investigator for NASA's Wind \Jspacecraft\j at GSFC, placed a notice on an Internet E-mail list, after scientists studying data from ISTP \Jspacecraft\j estimated that \JComet\j Hale- Bopp's ion tail likely would be disrupted when it enters a region around the Sun known as the "current sheet." Observations from amateur astronomers monitoring changes in the \Jcomet\j's tails will provide near-real-time data to scientists to complement observations from \Jspacecraft\j.
Scientists explain the disruption as a complicated interaction between the \Jcomet\j and the Sun's influence and magnetic fields. As a \Jcomet\j comes closer to the Sun, ices from the nucleus (a porous structure of dust and ice composed of frozen gases) are continually vaporized, dislodging the dust, which is formed by the \Jcomet\j's weak gravity into a cloud, called a coma, surrounding the \Jcomet\j. While pressure from the visible sunlight "pushes" the coma dust into a diffuse dust tail, the ultraviolet portion of the sunlight gives the coma an electrical charge, or ionizes it, turning it into a plasma of electrically charged particles of ions and electrons.
The solar wind (also a plasma), flowing from the Sun at speeds from 240-450 miles per second and carrying an embedded magnetic field, smashes into the coma gas, causing additional \Jionization\j. The magnetic field in the solar wind picks up \Jcomet\j ions and accelerates them into a long, blue plasma tail. Since this tail is stretched very long, it is much fainter than the dust tail and consists mostly of long-lived (stable) ionized carbon monoxide. The magnetic field is draped around the \Jcomet\j coma and controls the formation of the plasma tail. If the magnetic field is disrupted, the plasma tail may be disconnected.
Hale-Bopp's orbit is tilted relative to the Sun's equator with the \Jcomet\j moving from the Sun's northern hemisphere to its southern hemisphere, crossing the Sun's equatorial plane. This plane is the location of the "current sheet," a place where the Sun's magnetic field lines change direction. As Hale-Bopp passes through this plane, its ion tail may disconnect because of the change in direction of the magnetic field.
"Other events on the Sun may disrupt Hale-Bopp's tail," adds Dr. Farrell. "For example, at any time, the Sun may eject large amounts of hot, electrically charged material in the form of plasma, called Coronal Mass Ejections, or CME's. The magnetic fields associated with a CME may disrupt the ion tail, particularly if the CME is from the Sun's eastern limb in the direction of Hale-Bopp. Also, the solar wind is more gusty around the equatorial regions, and this could cause a disruption as well," he said.
"Monitoring this \Jcomet\j tail disruption is more than anticipating an intriguing astronomical phenomenon," said Dr. Farrell. "The stronger solar events can have a tremendous impact on Earth. The plasma ejected by these events smashes into the Earth's magnetic field and compresses it. This generates a magnetic storm which can disrupt power grids and radio communications. Additionally, the effects can damage microcircuits in satellites. With ISTP, if we can monitor disruption events for comets, we can do the same for Earth, providing a warning when they occur," he said.
When Hale-Bopp crosses the current sheet, it will provide additional data about its structure where no ISTP \Jspacecraft\j exist. "It could cost about a billion dollars to build and place a \Jspacecraft\j where Hale-Bopp is," said Dr. Adam Szabo, senior scientist with Hughes STX on the Wind project. "Comet Hale-Bopp will give us interesting information about this region of space for virtually no cost, except our time to watch and study it. It's a bonus which can really help us understand the most powerful forces which are affecting the Earth."
The ISTP \Jspacecraft\j involved in this study are NASA's Polar and Wind missions and the European Space Agency/NASA Solar and Heliospheric Observatory mission.
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"Mars Global Surveyor To Aerobrake In Modified Configuration (30 April, 1997)",530,0,0,0
NASA's Mars Global Surveyor \Jspacecraft\j can safely and successfully aerobrake into its final orbit around Mars this fall with its one partially deployed solar panel in a modified configuration, mission managers have decided.
No special maneuvers will be conducted to attempt to force the array to latch, and the focus of the Surveyor \Jengineering\j team now will turn to minor modifications to the critical aerobraking phase that will circularize the \Jspacecraft\j's orbit for the beginning of two years of science operations.
"After careful analysis of the situation, we've determined that the solar panel on Mars Global Surveyor that is not fully deployed presents very little risk to the mission," said Glenn E. Cunningham, Mars Global Surveyor project manager at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
The decision by NASA's flight team at JPL and its partners at Lockheed Martin Astronautics, \JDenver\j, CO, was reached after several months of extensive analysis of \Jspacecraft\j data, ground-based computer simulations and a series of very slight \Jspacecraft\j maneuvers that were carried out in January and February to characterize the situation.
"Thanks to an early launch that gave us an advantageous trajectory, we will not have to aerobrake into the Martian atmosphere as fast as we had originally planned to reach the mapping orbit, and that will reduce the amount of heating that the solar panels undergo during this gradual descent," Cunningham explained.
"We will rotate the solar-cell side of the panel that is not fully deployed by 180 degrees, so that it faces into the direction of the air flow that exerts drag force on the \Jspacecraft\j as it dips repeatedly into the atmosphere," he said. "This way, the unlatched panel will not be in danger of folding up onto the \Jspacecraft\j's main structure, nor will the panel be at any greater risk of heating up too much."
The solar panel in question is one of two 11-foot wings that were unfolded shortly after Surveyor's Nov. 7, 1996, launch from Cape Canaveral Air Station, FL. Data suggest that a piece of metal called the "damper arm," which is part of the solar array deployment mechanism located at the "elbow" joint where the entire panel is attached to the \Jspacecraft\j body, probably was sheared off during deployment in the first day of flight. The lever that turns the shaft became wedged in a two-inch space between the shoulder joint and the edge of the solar panel, leaving the panel tilted at 20.5 degrees from its fully deployed and latched position.
Although the situation was never considered a serious threat to accomplishing the science objectives of the mission, the tilted array caused the JPL/Lockheed Martin flight team to re-evaluate the aerobraking phase, in which the \Jspacecraft\j must rely almost solely on its solar panels for the drag needed to lower it into a nearly circular mapping orbit over the poles of the planet. This phase of the mission will begin a week after Mars Global Surveyor is captured in orbit around Mars on Sept. 11, and will last approximately four months.
Aerobraking was first tested in the final days of the Magellan mission to Venus in October 1994. The technique is an innovative method of braking which allows a \Jspacecraft\j to carry less fuel to a planet and take advantage of the planet's atmospheric drag to descend into a low-altitude orbit.
Mars Global Surveyor will use an aerobraking phase much like that used to circularize Magellan's orbit. The solar wings -- which feature a Kapton flap at the tip of each wing for added drag -- supply most of the surface area that will slow the \Jspacecraft\j by a total of more than 2,684 miles per hour during the four-month phase. Surveyor's orbit around Mars will shrink during this phase from an initial, highly elliptical orbit of 45 hours to a nearly circular orbit taking less than two hours to complete.
Engineers determined that the deployment springs currently holding the tilted solar panel in its nearly deployed position will not be strong enough to withstand the forces of aerobraking. To solve that problem, they designed a new configuration in which the tilted solar panel, along with the deployment springs, will be rotated 180 degrees, using a motor- driven inner gimbal actuator, and held in position with force applied by an outer gimbal actuator. Sequencing software will be modified to turn the gimbal actuators on before each closest approach to the planet and off at the conclusion of each drag pass.
As a consequence of the new aerobraking configuration, the more sensitive cell-side of the unlatched wing will be exposed directly to the wind flow of atmospheric entry, requiring that aerobraking be done in a more gradual, gentle manner. Ground tests have demonstrated that the unlatched solar panel will have more than adequate thermal margin to withstand additional heating as the \Jspacecraft\j circularizes its orbit for the beginning of science mapping in March 1998.
Meanwhile, Mars Global Surveyor continues to perform very well on its arcing flight path toward the red planet and its arrival in orbit. A third, very minor trajectory correction maneuver, planned for April 21, was deemed unnecessary and canceled. In addition, science instrument calibrations continue to go well, and plans are being prepared to take an approach image of Mars a few days before the July 4 landing of Mars Pathfinder, which passed Mars Global Surveyor enroute to Mars on March 14, 1997.
Mars Global Surveyor is the first mission in a sustained program of robotic exploration of Mars, managed by JPL for NASA's Office of Space Science, Washington, DC.
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"Antimatter Clouds And Fountain Discovered In The Milky Way (28 April, 1997)",531,0,0,0
Scientists using data from an instrument on NASA's Compton Gamma Ray Observatory (CGRO) have discovered two unexpected clouds of antimatter in the Milky Way Galaxy which scientists call "antimatter annihilation radiation."
Scientists from Northwestern University, Evanston, IL, the Naval Research Laboratory (NRL), Washington, DC, and other institutions used CGRO's Oriented Scintillation Spectrometer Experiment (OSSE) to make the discovery, which points to the existence of a hot fountain of gas filled with antimatter electrons rising from a region that surrounds the center of the Milky Way galaxy. The nature of the furious activity producing the hot antimatter-filled fountain is unclear, but could be related to massive star formation taking place near the large black hole at the center of the galaxy. Other possibilities include winds from giant stars or black hole antimatter factories.
The researchers used maps of gamma ray sources from CGRO which they expected to show a large cloud of antimatter near the galactic center and along the plane of the galaxy. The maps, surprisingly, also show a second cloud of antimatter well off the galactic plane. The second cloud may be caused by the explosions of young massive stars.
"The origin of this new and unexpected source of antimatter is a mystery," said William R. Purcell, research scientist and assistant professor of physics and \Jastronomy\j at Northwestern University.
"The antimatter cloud could have been formed by multiple star bursts occurring in the central region of the galaxy, jets of material from a black hole near the galactic center, the merger of two neutron stars, or it could have been produced by an entirely different source," said James D. Kurfess, head of the Gamma and Cosmic Ray Astrophysics Branch at the Naval Research Laboratory.
The researchers presented their findings today at the fourth Compton Symposium in Williamsburg, VA. The results have been submitted for publication in the Astrophysical Journal. A second paper presented at the conference, titled "The Annihilation Fountain in the Galactic Center Region," examines theoretical models for one possible source of the antimatter -- star bursts in the central region.
The second paper is authored by Dr. Charles Dermer and Dr. Jeffrey Skibo of NRL. They note that the gamma-ray observations permit us to see clearly, for the first time, a new part of our galaxy made of a hot column of gas filled with antimatter electrons (also called positrons by scientists), and they argue that the antimatter electrons come from newly created elements produced by exploding stars formed near the center of our galaxy.
"It is like finding a new room in the house we have lived in since childhood," comments Dr. Dermer. "And the room is not empty -- it has some engine or boiler making hot gas filled with annihilating antimatter. No one is certain whether the antimatter comes from exploding stars, black holes or something entirely different, and that is what makes this discovery so exciting."
Evidence points to the existence of a black hole with the mass of a million Suns at the very center of our galaxy. Unlike in other galaxies which harbor huge black holes, very little light comes from this source. Huge dense clouds of gas also surround the galactic center. Prolific star formation, powerful stellar winds from massive stars, and supernovae are all found here. Another theory, based on observation of radio emissions showing some black holes produce X-rays and jets, is that such outflowing jets could be made of antimatter.
The Compton Gamma Ray Observatory, launched from the Space Shuttle in 1991, views the universe in a search for gamma rays and their source. Gamma rays are extremely energetic light photons produced by high-energy particles, by the decay of excited nuclei, and when matter and antimatter annihilate each other. Antimatter cannot be found in large quantities on Earth because it would instantly vaporize anything it came into contact with. All evidence points to the universe being composed almost entirely of normal matter, though opinions differ on this.
Using the OSSE experiment, the OSSE team found antimatter positrons to be annihilating with normal matter electrons at an astonishing rate. Scientists are speculating on the origin of this antimatter, with a "black-hole lobby" favoring antimatter production in the jets of black holes.
Other scientists favor freshly synthesized radioactive material in stellar explosions being ejected up above our galaxy in an annihilating fountain of gas. Drs. Dermer and Skibo favor the latter scenario, because exploding stars will eject large quantities of hot gas made up of normal matter. This hot gas provides a target with which the antimatter electrons can annihilate.
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"Microgravity Science Laboratory Mission Set For July (25 April, 1997)",532,0,0,0
Columbia's \JMicrogravity\j Science Laboratory (MSL) mission will fly again in early July to complete the mission cut short earlier this month because of a fuel cell problem. The remaining Space Shuttle flights in 1997 have been adjusted to accommodate Columbia's mission, which will fly as STS-94. Air Force Lt. Col. Jim Halsell and the rest of the STS-83 crew will fly this mission and will conduct proficiency training until the flight.
Space Shuttle Program managers today formally baselined the STS-94 mission to follow Atlantis' sixth docking with the Russian space station Mir next month. \JAstronaut\j Jerry Linenger will return home on STS-84 following a four-month stay on Mir, and Mike Foale will replace him as a station crew member. Managers will formally select the launch date following the Flight Readiness Review on April 30.
"While shortening STS-83 was disappointing, we now are in a position to do everything possible to complete the MSL mission with minimal impact to downstream flights," said Space Shuttle Program Manager Tommy Holloway. "Also, it provides us with a unique opportunity to demonstrate our ability to respond to challenges such as this one."
Reflying Columbia in July dictated that downstream flights for the remainder of the year change slightly. Following STS-94, Discovery will fly in early- to mid-August on the STS-85 mission to deploy and retrieve a science satellite to study Earth's atmosphere. The flight also will demonstrate the use and operational capability of a robot arm that will be deployed outside the Japanese Experiment Module of the International Space Station.
The seventh Shuttle-Mir docking mission on Atlantis is targeted for mid- to late-September. STS-86 will include the return of Foale from Mir and delivery of his replacement, \Jastronaut\j Wendy Lawrence.
The eighth and final mission scheduled in 1997 will be the STS-87 flight of Columbia slated for mid- to late-November. The 16-day mission includes the conduct of science experiments associated with the fourth flight of the U.S. \JMicrogravity\j Payload and the deployment and retrieval of a science satellite.
#
"U.S. Astronomy Ready For Milestone Spacewalk (25 April, 1997)",533,0,0,0
U.S. \Jastronaut\j Jerry Linenger, more than 100 days into his four-month research mission on the Space Station Mir, is scheduled to conduct his first spacewalk next Tuesday, April 29, to place experiments on two of the station's modules and to retrieve two others deployed outside the Mir last year. He will be joined by Mir 23 Commander Vasily Tsibliev.
The spacewalk, which is supposed to last about six hours, marks the first time a U.S. \Jastronaut\j will conduct a spacewalk wearing a Russian spacesuit. Russian \Jcosmonaut\j Vladimir Titov is scheduled to conduct a spacewalk wearing a U.S. spacesuit during Atlantis' docking mission to the Mir in September.
Astronauts Linda Godwin and Rich Clifford conducted a spacewalk outside the Shuttle Atlantis last year to place suitcase-sized sensors and materials designed to catch debris on the Mir's docking module to determine how often debris hits the surface of the Mir and to assess outside environmental factors.
During his spacewalk, Linenger will deploy an experiment to collect data on how the space environment affects the Mir's outersurface (Optical Properties Monitor Experiment). He will also retrieve the Partial Impact Experiment and the Mir Sample Experiment from the Kvant-2 science module; both experiments were deployed by Russian cosmonauts during a spacewalk last year and are designed to monitor the outside environment of the Mir.
Linenger and Tsibliev also will place a radiation detection device called the Benton radiation dosimeter on the Kvant-2 near the end of their spacewalk and evaluate a safety tether which may be used by both American and Russian spacewalkers in the assembly of the International Space Station.
Tsibliev conducted five spacewalks during his previous flight on the Mir in 1993. This will be Linenger's first spacewalk. Tsibliev plans to conduct two more spacewalks with Flight Engineer Alexander Lazutkin, in June, about a month after U.S. \Jastronaut\j Mike Foale has replaced Linenger on board the Russian outpost.
The spacewalk is expected to be seen on NASA \JTelevision\j from 6 a.m. to 6:30 a.m. EDT and from 7:30 a.m. to 8:30 a.m. EDT. Highlights will be replayed on NASA TV's Video File at noon EDT.
NASA \JTelevision\j is available through the GE2 satellite, located on Transponder 9C, at 85 degrees West longitude, frequency 3880.0 MHz, audio 6.8 MHz.
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"NASA Concurs With Independent Review Of Bion 11 Mission (22 April, 1997)",534,0,0,0
NASA is suspending its participation in primate research on the Bion 12 mission, part of an international project to study the physiological effects of low gravity and space radiation. NASA's decision is based on the recommendations of an independent review board requested by the Agency to look into the post-flight death of a rhesus monkey following the successful flight and landing of the Bion 11 satellite.
The panel found that there was an unexpected mortality risk associated with anesthesia for surgical procedures (biopsy of bone and muscle) on the day following return from space. NASA has determined that this risk is unacceptable and is therefore discontinuing its participation in the primate experiments on Bion 12.
The independent review was led by Dr. Ronald Merrell, chairman, Department of Surgery, Yale University, New Haven, CT. Dr. Merrell closely consulted with the Russian Bioethics Commission of the Russian Academy of Sciences, which conducted the Russian inquiry.
Based on the difficulty encountered with post-flight anesthesia on the Bion 11 mission, the research protocols originally developed for the Bion 12 mission cannot be conducted without an unacceptable risk to the primates. NASA therefore plans to:
╖ incorporate lessons learned from this mission into ongoing scientific research, reviews and medical considerations for space flight;
╖ in concert with the biomedical community, conduct research with the appropriate models to investigate medical care in relation to space physiology;╖ work with the biomedical research community to develop new technologies for collecting critical data needed to continue this important research.
The Bion program is a cooperative space venture among the U.S., Russian and French space agencies for conducting biomedical research using Russian-owned rhesus monkeys. The 14-day Bion 11 mission, carrying two rhesus monkeys as well as other life science and \Jmicrogravity\j experiments, began on Dec. 24, 1996, with its launch from \JRussia\j's Plesetsk launch site. The flight was successfully completed when the \Jspacecraft\j landed in \JKazakhstan\j on Jan. 7, 1997.
Experiments flown on the Bion missions encompass a broad range of important investigations that expand our understanding of a variety of fundamental and applied life sciences questions.
In space, as on the ground, biomedical research on animals plays a vital role in expanding NASA's capacity to understand and treat medical problems. NASA is deeply concerned with the welfare of its animals and is fully committed to conducting its animal research programs in conformance with the highest ethical standards. The Bion experiments were thoroughly reviewed four times by NASA and outside panels to ensure that they met ethical standards, and that they pursued worthwhile and important scientific objectives that could not be achieved without the use of animals.
#
"Solar Storm Spotted By Spacecraft (9 April, 1997)",535,0,0,0
A large eruption on the Sun was detected at 10 a.m. EDT, on Monday, April 7 by the Solar and Heliospheric Observatory (SOHO) \Jspacecraft\j, and scientists say the ejected matter, traveling through space as a so-called interplanetary magnetic cloud, is likely to reach the Earth at 8 p.m. EDT today. The storm is large in size; however, it is moderate in strength compared to many others that have reached the Earth in the past.
Although the SOHO findings are of interest to scientists, the National Oceanic and Atmospheric Administration Space Environment Center (SEC), provider of official government forecasts, has predicted, based on "classical indicators," that the solar storm will be an ordinary event and pose no danger to the population at large.
SEC notes the absence of energetic solar particles which would be expected to accompany a strong solar storm. Such particles, if present, can increase radiation levels for astronauts and cosmonauts in orbit. "We are forecasting an event with conditions below the threshold of concern for most of our users,"said Dr. Ernie Hildner, head of the Space Environment Center.
The center of the storm will miss the Earth, but it could be broad enough to affect the Earth's space environment and could cause increased auroral activity (Northern and Southern lights) at high latitudes. In the past, some solar storms have affected \Jspacecraft\j in orbit, shortwave communications and electric power grids.
SOHO, a joint project of NASA and the European Space Agency, is positioned about 900,000 miles sunward of the Earth. The \Jspacecraft\j has a continuous telescopic view of the Sun and also is equipped with sensors to sample solar particles as they sweep past. A vast network of satellites, space probes, and ground sensors, part of the International Solar Terrestrial Physics (ISTP) program, is monitoring the approaching interplanetary storm and preparing to observe its possible effects on the Earth's space environment.
The solar eruption, called a coronal mass ejection, was first spotted by Shane Stezelberger, a ground controller with SOHO's Large Angle \JCoronagraph\j and Spectrometer (LASCO) team at the SOHO Experiment Operations Facility at NASA's Goddard Space Flight Center in Greenbelt, MD. "I knew it was a big one when I saw it," said Stezelberger, a recent graduate of Virginia \JPolytechnic\j Institute and State University in Blacksburg. He promptly notified SOHO scientists, including Dr. Barbara Thompson, whose immediate reaction was "Wow." Dr. Donald Michels of the U.S. Naval Research Laboratory said, "We've never before seen one headed directly at Earth that was as big and bright, and loaded with complex details, as this one." SOHO was launched on Dec. 2, 1995.
"The eruption seemed to blow open a hole in the Sun's corona that had opened and then healed previously," said Dr. Thompson, who is a physicist with SOHO's Extreme-Ultraviolet Imaging \JTelescope\j (EIT) team. Accompanying the coronal mass ejection was a powerful solar flare explosion, according to Dr. Arthur \JPoland\j, NASA's SOHO Project Scientist. "The flare triggered a \Jsupersonic\j wave that swept through the corona like a \Jtsunami\j on the surface of the sea," Dr. \JPoland\j said.
A sensor called the WAVES experiment on NASA's WIND satellite, positioned near the SOHO, detected bursts of radio emissions from high speed electrons associated with the explosion, also beginning at about 10 a.m. EDT Monday.
The interplanetary storm, traveling toward Earth at a speed of over 1,500,000 miles per hour, is expected to pass the WIND and SOHO satellites at about 7 p.m. EDT Wednesday, and to strike the Earth's magnetosphere an hour later, according to Dr. Nicola Fox, Global Geospace Science program coordinator at Goddard. At that time, according to her projections, the Geotail satellite, a joint Japanese-NASA \Jspacecraft\j that is part of the ISTP program, will be passing through the center of the Earth's magnetic tail, on the side opposite the Sun. "Geotail will be ideally positioned to observe storm activity like dipolarizations and bursty bulk flows of plasma," she added. Plasma is the scientific term for electrified gas.
NASA's Earth-orbiting POLAR satellite will train its battery of visible, ultraviolet and X-ray cameras on the north and south polar regions of the Earth to observe "a likely considerable enhancement of the aurora," said Dr. Robert Hoffman, POLAR Project Scientist. And, with other POLAR onboard sensors, "we'll look for possible large increases in the intensity of the radiation belts." The likely impact of the interplanetary cloud on the magnetosphere could be to compress it, driving the radiation belts closer to the Earth, as occurred in January after a smaller coronal mass ejection lifted off the Sun on Jan. 6.
Analysis of the \Jspacecraft\j and ground sensor data on this week's solar storm and its effects on the Earth should lead to a better understanding of the basic physical processes involved and how such disturbed conditions of "space weather" can be predicted.
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"NASA Revises International Space Station Schedule (9 April, 1997)",536,0,0,0
NASA will begin its on-orbit assembly of the International Space Station (ISS) no later than October 1998, and is looking at options that will allow the Agency to work around the delay caused by the late arrival of a key station module.
"We knew from the outset that building an International Space Station was going to be tremendously challenging. Space exploration is not easy or predictable," said NASA Administrator Daniel S. Goldin. "We will work through this schedule issue, and we undoubtedly will face additional problems in the future. But we are well on our way to the realization of this world-class facility," he said.
The on-orbit assembly of the International Space Station originally was scheduled to begin in November 1997 with the launch of the NASA-financed/Russian-built and launched Functional Cargo Block (FGB). Inadequate funding by the Russian government to the Russian Space Agency (RSA) and its contractors for building another key station element -- the Service Module (SM) -- has put construction up to eight months behind schedule.
NASA managers and engineers have been reviewing various options to mitigate the impact to the ISS program of the current schedule slip of the Service Module, and to begin the steps necessary to mitigate the impact of potential additional Russian delays. RSA has been a joint participant in the effort to identify these steps. Options under consideration are:
ò Modify the FGB to allow for on-orbit refueling and upgrade of its avionics capability. These changes will give the FGB the capability to augment the early control and reboost capabilities to protect for a Service Module delay.
ò Develop an Interim Control Module (ICM) in conjunction with the Naval Research Laboratory to provide reboost capability and attitude control in the event that the SM experiences further delays, or propellant storage/reboost capability if the SM is launched on time.
ò Consider the installation of life support systems in the U.S. lab to allow early human presence on the ISS.
ò Define options involving the ICM to provide the functions of a permanent propulsion module in order to complement Russian logistics capability and to provide roll control to replace or complement the Russian Science Power Platform functions.
NASA will determine the timing for decisions which need to be made in the event that \JRussia\j is unable to provide its agreed contributions to the ISS program. These decision points will be selected to allow for the timely provision of an alternative capability.
NASA has begun initial steps at the working level to accommodate changes to the Space Shuttle manifest. NASA has reassigned the Space Shuttle Endeavour to fly the September 1997 STS-86 mission to the Mir space station instead of the Space Shuttle Atlantis. This change will allow Atlantis to begin its scheduled orbiter maintenance down period several months earlier, while permitting Endeavour a mission prior to flying the first ISS assembly flight in July 1998.
Additional adjustments to the remainder of the assembly sequence will be worked in consultation with the other International Partners and research community over the next several weeks.
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"Low Ozone Measured Over North Pole (8 April, 1997)",537,0,0,0
Unusually low levels of ozone over the \JArctic\j were measured during March by satellite-based monitoring instruments operated by NASA and the National Oceanic and Atmospheric Administration (NOAA).
"These are the lowest ozone values ever measured by the TOMS instruments during late-March and early-April in the Arctic," said TOMS Project Scientist, Dr. Pawan K. Bhartia, of NASA's Goddard Space Flight Center (GSFC), Greenbelt, MD. "However, these low ozone amounts are still nearly a factor of two greater than the lowest values seen by TOMS in the Antarctic during Southern hemisphere Spring."
Centered in a stable, nearly circular region over the North Pole, the average March 1997 ozone amounts were 40 percent lower than the average March amounts observed between 1979 and 1982. This follows ozone amounts in March 1996 that were 24 percent lower than the 1979-82 average, although this low was off center of Earth's pole toward the North Atlantic.
The minimum in total column ozone fell to 219 Dobson units on March 24, 1997, from values near 280 units earlier in March. Two NASA Total Ozone Mapping Spectrometer (TOMS) instruments, one aboard NASA's Earth Probe (TOMS-EP) satellite and the other aboard \JJapan\j's Advanced Earth Observing Satellite (ADEOS) made the measurements of the rapid decrease, supported by similar data from the Solar Backscatter Ultraviolet instruments aboard the NOAA-9 and NOAA-14 satellites.
The Halogen \JOccultation\j Experiment (HALOE) aboard NASA's Upper Atmosphere Research Satellite (UARS) measured vertical distributions of ozone that confirm these low \JArctic\j values. On March 26, HALOE measured a very low ozone concentration of less than one part per million of ozone (normal concentrations are near 3-4 parts per million) at an altitude of 12.4 miles slightly northeast of Hudson's Bay.
Over the Alaskan \JArctic\j, where NOAA monitors ozone from the surface at Barrow, \JAlaska\j, March average ozone was about 375 Dobson units, slightly below the March average for the past ten years of 413 Dobson units, according to David Hofmann, Director of NOAA's Climate Monitoring and Diagnostics Laboratory in Boulder, CO. "On March 17 and 18 ozone dipped to values below the 300 Dobson unit range when the edge of the low ozone region extended to the latitude of Barrow (71 degrees north)," he said. "This is not a typical occurrence and indicates the unusual conditions this year."
"The unusual meteorological conditions played a significant role in the March ozone lows," according to Paul A. Newman of the GSFC. "The reason or reasons behind these unusual stratospheric weather patterns are unclear, and figuring out why this pattern occurred will be a significant component of our further research efforts."
Furthermore, measurements from balloon-based ozone instruments operated by Environment Canada and launched from Eureka (80 degrees North latitude) and Resolute Bay (75 degrees North) reveal 60 percent ozone losses between the altitudes of 6.2-15.5 miles during March, in comparison to historical March observations.
The TOMS data show that the region of low ozone amounts below 280 DU exceeded 5.3 million square kilometers, covering a substantial fraction of the \JArctic\j region. These low ozone amounts are found inside the \JArctic\j polar vortex, a part of the stratospheric circumpolar jet stream.
The 1996/1997 winter polar vortex has been unusually strong and persistent into March. Data from NOAA's Climate Prediction Center show cold temperatures low enough to form polar stratospheric clouds during late March. These clouds, common in January and February but rare in late March, helped convert certain forms of stratospheric \Jchlorine\j into forms which are highly reactive to ozone destruction. The combination of reactive \Jchlorine\j compounds and sunlight from the March rising sun at polar latitudes leads to destruction of ozone.
The most recent observations indicate that temperatures had become too warm by March 30 for polar stratospheric clouds to form. In addition, the minimum ozone amounts in the \JArctic\j have begun to slowly increase from the unusually low March amounts, and the area covered by the low has begun to decrease.
"The appearance of this well-defined region of low ozone is consistent with our expectations following detailed chemical analyses of the \JArctic\j winter \Jstratosphere\j in early 1989 and 1992," said Dr. Michael Kurylo, manager of NASA's Upper Atmosphere Research Program, which organized those airborne experiments. "The persistence of such cold temperatures within the \JArctic\j vortex well into the sunlit period is an essential ingredient for driving many of the chemical cycles for ozone destruction. We will now be examining these low ozone air masses into their recovery period using our best satellite and airborne instruments."
An international treaty on ozone-depleting substances is leading to reductions in their concentrations in the atmosphere and hence to reduced \Jchlorine\j levels in the \Jstratosphere\j. As we move into the next century, chlorine-catalyzed ozone losses resulting from CFCs and other chlorine-containing species will be reduced.
Ozone, a molecule made up of three atoms of oxygen, absorbs harmful ultraviolet radiation from the Sun. Most atmospheric ozone is found in a thin layer between 6-18 miles. A Dobson unit is related to the physical thickness of the ozone layer if it could be brought down to the Earth's surface. The global average ozone layer thickness is 300 Dobson units, which equals 1/8th of an inch, approximately the thickness of two stacked pennies. In contrast, the ozone layer thickness in the Antarctic ozone hole is about 100 Dobson units (1/25th of an inch), approximately the thickness of a single dime.
TOMS-EP, ADEOS, and UARS, and the \Jaircraft\j, balloons and ground-based ozone-measurement programs are key parts of a global environmental effort which includes NASA's Mission to \JPlanet\j Earth, a long-term, coordinated research effort to study the Earth as a global environmental system.
The TOMS instruments are managed by Goddard. The HALOE instrument is managed by NASA's Langley Research Center, Hampton, VA. NOAA-9 and NOAA-14 are managed by the NOAA National Environmental Satellite, Data, and Information Service, Suitland, MD.
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"Shuttle Experiment To Study Medicinal Properties Of Plants (1April, 1997)",538,0,0,0
Studies of plants on the next Space Shuttle mission may someday lead to the production of life- saving medicines and other important compounds.
The experiments conducted by Dr. Gerard Heyenga at NASA's Ames Research Center, Mountain View, CA, will be part of the 16-day STS-83 mission, scheduled for launch this week.
"A fundamental objective of this research is to evaluate whether \Jmicrogravity\j may be used to alter specific metabolic pathways in plants, and ultimately apply this technology for Earth-based benefits," Heyenga said.
Heyenga hypothesizes that extended exposure of plants to \Jmicrogravity\j may reduce their expenditure of energy on structural components, thereby increasing flow through other metabolic pathways, many of which produce materials of important medicinal value. Of even greater significance is the possibility that corresponding changes may occur at a genetic level, he said.
A comparison between space- and Earth-grown plants would give a unique opportunity to obtain a greater understanding of how these pathways are controlled at the gene level, Heyenga said. In turn, "such knowledge would allow us to manipulate or genetically engineer plants with desired metabolic traits," he added. "For example, this information could be applied to the lumber industry in the production of trees with a low lignin content, greatly reducing the cost of paper production both economically and environmentally." Conversely, it might be applied to improving timber quality in fast-growing softwoods, reducing the need to harvest slow-growing hardwoods, he said. "If this hypothesis is correct and achievable, it obviously represents the basis for a multi-billion dollar industry and certainly highlights the value of space-related research and such facilities as the Space Station."
A critical requirement in the investigation is the ability to maintain well-characterized and high- quality plant-growth conditions during space flight and corresponding Earth experiments. "To achieve a meaningful understanding of the effects of \Jmicrogravity\j on plants, it is essential that we minimize or avoid additional factors that may cause any stress and that complicate the evaluation," Heyenga said.
To this end, the flight experiment will involve the use of a new, advanced plant growth facility known as the Plant Generic Bioprocessing Apparatus (PGBA), built by BioServe, a NASA Commercial Space Center located at the University of \JColorado\j in Boulder. The chamber was first flown on the Space Shuttle in 1996. "While it was essentially a hardware verification test, the PGBA produced a particularly high quality of plant material over the ten-day mission, which provided a good basis for further research," Heyenga said.
The PGBA chamber maintains a highly controlled environment, supplying appropriate light, temperature and gas exchange conditions. The chamber will utilize the new modular "nutrient pack system" designed by Heyenga to supply plants with water and \Jnutrients\j throughout the mission. Thirty packs will be used to support the growth of nine plant species.
The packs offer several advantages over existing systems. Depending on the type of supporting substrate used, packs may reabsorb water from the chamber's condensate recovery system, closing the water loop and presenting an important opportunity for long-term plant cultivation. A number of packs will utilize a gel matrix that will allow the examination of the roots' spatial orientation. Since the matrix is safely encapsulated in a protective membrane, the nutrient pack system has been certified for the first radiolabelling tracing experiment of higher plants in space. "This technology will open an entirely new area of space plant \Jphysiology\j, allowing the study of issues not previously possible," Heyenga said. "It is likely to lead to some very exciting results."
The plant species chosen for the flight experiment include a member of the black pepper family. This choice was based on a collaborative effort with a research group in \JBrazil\j. "I believe it is highly important that we utilize every possible means to expand our understanding in space research," Heyenga said. "The use of such tropical species, with their unique and specific metabolic pathways, hopefully will provide us with an early indicator of whether our hypothesis is correct while the plants are exposed to the relatively short period of \Jmicrogravity\j experienced during a typical Space Shuttle mission."
Despite the complexity of the research program, Heyenga is pleased with progress so far. "The work has involved a particularly broad multidisciplinary effort by a number of organizations. As a visiting scientist on a National Research Council fellowship, it is unique to find a place like Ames that can support this type of activity," he said. "Earlier work by Dr. Robert MacElroy and Dr. Mark Kliss at Ames in the area of enclosed plant growth systems has provided important support for the present flight experiment."
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"NASA Hosts Space Station Radiator Tests (31 March, 1997)",539,0,0,0
An innovative radiator, designed to provide cooling for the International Space Station, is undergoing testing at the NASA Lewis Research Center's Plum Brook Station in Sandusky, OH.
The Photovoltaic Radiator (PVR) system is being tested in Plum Brook's Space Power Facility, the world's largest vacuum chamber. Lockheed Martin Vought Systems of \JDallas\j, TX, manufactured the PVR hardware and is conducting the tests.
The tests, which are evaluating the radiator system and its deployment mechanism, qualification thermal \Jcycling\j and thermal heat rejection performance, are scheduled for completion in early April. This is one of the final tests of the radiator system prior to its installation on the International Space Station.
The first round of tests confirmed that the radiator's deployment mechanism would operate properly in the cold void of space.
"Validating the deployment mechanism of a photovoltaic radiator system in space conditions is critical to ensure successful deployment in space," said Lewis program manager James Mullins. "The Plum Brook vacuum chamber enabled us to test the radiator in temperatures ranging from -250 to 125 degrees Fahrenheit."
The next phase of testing will use a NASA- designed and-assembled ammonia flow system to evaluate the radiator's performance. Ammonia pumped through the system will collect heat from the Space Station's electronic equipment or module cooling components and transport it to the radiator panels where it will be dissipated.
The PVR is a critical component of the Space Station's thermal control system. It will cool the photovoltaic-power-system electronic equipment and the batteries used for power storage. The PVR also will provide environmental cooling for the service module during early phases of the Space Station.
Each PVR Orbital Replacement Unit (ORU) consists of seven radiator panels, each about 6 feet by 12 feet in size. The radiator panels are designed to deploy on orbit from a stowed position about two feet high to an extended position about 50 feet in length. Each ORU weighs about 1,600 pounds, with a total of four units in the final Space Station configuration. Fabrication of the hardware is being performed by Lockheed Martin Vought Systems under contract to the Rocketdyne Division of Boeing North America, the prime contractor for the International Space Station power system.
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"Shuttle's New Lighter, Stronger External Tank Completes Major Pressure Tests (28 March, 1997)",540,0,0,0
The first new, super lightweight, external fuel tank for the Space Shuttle is set for final assembly after successfully completing proof pressure tests that verify its design. The achievement is a significant step toward the first launch of the International Space Station. The new external tank is the same size as the one currently used on the Space Shuttle -- but about 7,500 pounds lighter.
"Each pound we remove from the external tank is a pound that can be added to the payload," said Parker Counts, manager of the External Tank Project at the Marshall Space Flight Center, \JHuntsville\j, AL. "The lighter tank is essential for launching the Space Station because the Station components will be assembled in a more demanding orbit than previously planned."
The 154-foot-long external tank, higher than a 15-story building and as wide as a silo with a diameter of about 27 feet, is the largest single component of the Space Shuttle. The external tank holds the liquid \Jhydrogen\j and liquid oxygen propellants in two separate tanks for the Shuttle's three main engines.
The two major changes to the external tank involve materials and design. Both the liquid \Jhydrogen\j and the liquid oxygen tank are constructed of aluminum \Jlithium\j -- a lighter, stronger material than the metal alloy used for the Space Shuttle's current external tank. The tank's structural design also has been improved. The walls of the redesigned \Jhydrogen\j tank are manufactured in an orthogonal waffle-like pattern, providing more strength and stability than the previous design.
"The new external tank has passed one of the most innovative structural verification test programs ever designed, culminating with these proof tests," Counts said.
The proof test for the liquid oxygen tank was a hydrostatic, or water pressure test. The tank was placed vertically on the test stand at NASA's Michoud Assembly Facility in New Orleans, LA, and filled with water, which has similar density to liquid oxygen. The tests simulated conditions encountered during flights and validated the design changes.
The liquid \Jhydrogen\j tank was pressurized with gaseous \Jnitrogen\j and subjected to conditions simulating the thrust of the orbiter's main engines and solid rocket boosters. Tests checked the new design by exposing the tank to harsher conditions than it will encounter in flight.
After the tests, comprehensive X-ray and dye penetrant inspections will be performed to further verify the tank's flight worthiness.
The proof tests completed March 25 were the final in a series of rigorous certification and structural verification tests.
"Our team of dedicated employees -- both at Marshall and Lockheed Martin Corporation -- completed the ambitious task of designing and building the improved external tank and verifying its design," said Counts.
"All of this was achieved successfully on a tight schedule of about three-and-a-half years. They pushed the state of the art of test technology, looking beyond current program requirements and squeezing all the information possible out of each test to understand the capability of the new tank," he added.
After thermal protection \Jfoam\j is sprayed on its exterior, the first super lightweight tank will be shipped by barge from Louisiana to the Kennedy Space Center, FL, for its launch with the first elements of the International Space Station. The changes to the external tank will not affect the assembly process when the orbiter is mated to its tank and solid rocket boosters.
The Shuttle's current external tank and the new super lightweight tank are manufactured by Lockheed Martin.
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"Solar-Powered Aircraft To Be Developed (27 March, 1997)",541,0,0,0
Aeronautical engineers in Southern \JCalifornia\j are developing an \Jaircraft\j -- called Centurion -- which they believe will push solar-powered \Jaircraft\j concepts literally to new heights.
Engineers for AeroVironment, Inc., Simi Valley, CA, are designing the \Jaircraft\j to fly at 100,000 feet altitude. The company is developing this concept as a member of NASA's Environmental Research \JAircraft\j and Sensor Technology (ERAST) program, which is sponsored by NASA's Dryden Flight Research Center, Edwards, CA.
Like its predecessor, the AeroVironment-developed Pathfinder, the Centurion will be an ultralight flying wing with multiple electric motors along its wingspan, powered by solar cells spread across the wing's upper surface. Centurion's wingspan, however, will be more than twice that of Pathfinder.
According to John Del Frate, Dryden's ERAST deputy project manager, recent flight tests of a quarter-scale battery-powered model of the craft at El Mirage Dry Lake in Southern \JCalifornia\j's high desert have answered questions about the Centurion's aerodynamics and stability.
"We saw it fly, and it flew quite well," said Del Frate. "It has given us confidence that we can go ahead with the design of the full-scale vehicle."
"We'll take the data from these flights and incorporate them into the design of the full-scale proof-of-concept vehicle," added Bill Parks, Centurion's chief designer and operations manager for the subscale flight tests.
"We're essentially scaling the \Jaircraft\j up, designing new airfoils that are more efficient for high altitudes and optimizing the systems," said Rik Meininger, AeroVironment's Centurion project manager.
Both cost and efficiency considerations have driven building and flying a subscale model, then a full-scale prototype before developing the final solar-powered Centurion.
"We find that we can make configuration changes very quickly and very cost effectively, then immediately test it and come back and change if necessary," he said. "It allows us, in a very short period of time, to get a lot of test data, and also do the risky things that normally you wouldn't want to do with a full-scale \Jaircraft\j. By the time we get to the final \Jaircraft\j stage, we should only be doing minor changes and fine-tuning for optimization," Meininger said.
The final solar-powered Centurion will be designed to reach an ultra-high altitude of 100,000 feet for a relatively short duration -- about two hours -- while carrying a small 200-pound payload of scientific sensors. The full-scale Centurion will span between 210 and 240 feet.
The subscale Centurion spans 62.5 feet but has only a two- foot chord. The straight wing is in five "spanwise" sections that are supported on the ground by four underwing pods. The model weighs in at a feathery 25 pounds, giving the kite-like craft a wing loading of only two-tenths of a pound per square foot and a design airspeed of about 11 knots (12.5 miles per hour).
Centurion officials had a chance to assess the lightweight craft's stability during a planned flight in adverse conditions. Although turbulence tossed the model around like a cork, causing the wing to flex significantly, remote test pilot Wyatt Sadler was able to maintain safe altitude by adding differential power. After a short duration flight, he brought the craft to a safe landing. The model flew more than an hour and 40 minutes on 13 flights.
The Centurion is one of several unpiloted \Jaircraft\j being developed by an alliance between NASA and several small aeronautical development companies and universities under NASA's ERAST program. The goal of the ERAST program is to develop aeronautical technologies that will lead to development of a new family of high-flying remotely piloted \Jaircraft\j for scientific missions.
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"Hubble And Iue Hale-Bopp Observations Surprise Astronomers (27 March, 1997)",542,0,0,0
Completing an unprecedented year-long study of \JComet\j Hale- Bopp using two NASA observatories, the Hubble Space \JTelescope\j and the International Ultraviolet Explorer, astronomers report that they are surprised to find that the different ices in the nucleus seem to be isolated from each other. They also report seeing unexpectedly brief and intense bursts of activity from the nucleus during the monitoring period. The Hubble observations suggest that the nucleus is huge, 19 to 25 miles across.
The findings, by a team of scientists led by Johns Hopkins astrophysicist Dr. Harold Weaver, are being published in the March 28 issue of the journal Science.
"Hale-Bopp will probably provide the most revealing portrait of the workings of a cometary nucleus since the \Jspacecraft\j missions to \Jcomet\j Halley in 1986," said Weaver. "This is a unique opportunity; we have never had the chance to examine a \Jcomet\j in this much detail, over this large a range of distance from the Sun."
The key results:
Violent Eruptions on the \JComet\j's Surface
During the course of long-term observations, which began in August 1995, astronomers unexpectedly caught the \Jcomet\j going through a sudden brief outburst, where, in little more than an hour, the amount of dust being spewed from the nucleus increased at least eight-fold. "The surface of Hale-Bopp's nucleus must be an incredibly dynamic place, with 'vents' being turned on and off as new patches of icy material are rotated into sunlight for the first time," Weaver said.
A Complex, Mottled Nucleus
To their surprise, astronomers found that water ice sublimates (turns directly from a frozen solid into a gas) at a different rate than the trace ices, implying that those components are not contained within the water on the \Jcomet\j. This conclusion is further supported by Hubble data showing that the rate at which dust left the nucleus was much different than the sublimation rate of water. This result is contrary to previous models for a \Jcomet\j's nucleus, which suggest that the trace components, such as carbon disulfide ice, are contained inside of the most abundant ice on the \Jcomet\j, frozen water. As water sublimates, the trace components and dust should be released at similar rates, but this is not what Hubble observed.
A Monstrous Nucleus
By studying Hubble Space \JTelescope\j images, the astronomers have estimated that its nucleus may be about 19 to 25 miles in diameter. The average \Jcomet\j is thought to have a nucleus of about three miles in diameter, or even smaller. The \Jcomet\j or asteroid that struck the Earth 65 million years ago, possibly causing the \Jextinction\j of the dinosaurs, was probably about six to nine miles across.
Because Hale-Bopp was unusually bright when it was still a great distance away, well outside the orbit of Jupiter, it has given scientists their best view ever of the changes in a \Jcomet\j's nucleus as it gets closer to, and is progressively heated by, the Sun. Those changes, in turn, provide information about the composition and structure of comets, which are believed to be remnants from the formation of the solar system, about 4.6 billion years ago. Learning more about comets can provide important information about the materials and processes that formed the solar system.
The Space \JTelescope\j Science Institute is operated by the Association of Universities for Research in \JAstronomy\j, Inc. (AURA), for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space \JTelescope\j is a project of international cooperation between NASA and the European Space Agency (ESA).
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"Space Science And Human Space Flight Enterprises Agree To Joint Robotic Mars Lander Mission (25 March, 1997)",543,0,0,0
In a cooperative activity intended to advance scientific knowledge and help lay the groundwork for a future decision on whether to send humans to Mars, NASA's Space Science and Human Exploration and Development of Space (HEDS) enterprises have agreed to jointly fund and manage two robotic missions to Mars due for launch in 2001.
"For the first time since the 1960s, NASA's space science and human space flight programs are cooperating directly on the exploration of another planetary body," said Dr. Wesley T. Huntress Jr., NASA associate administrator for space science. "Mars is a challenging destination for any type of \Jspacecraft\j to reach, and it makes a great deal of sense for us to pursue the maximum possible return of knowledge from any chance to go there."
"This joint effort is a sign that NASA is acquiring the information that will be needed for a national decision, perhaps in a decade or so, on whether or not to send humans to Mars," said Wilbur Trafton, associate administrator for space flight. "Early in the next century, once the International Space Station is deployed and operating, the question of our next major goal in human space flight will come up. This partnership is a major step toward ensuring that we have the information needed to answer that question."
NASA intends to launch two separate \Jspacecraft\j to Mars, a small orbiter and a small lander, in March and April 2001, respectively.
The Mars Surveyor 2001 Lander will deliver a small, advanced technology rover capable of traveling several tens of miles across the Martian highlands. The rover will be able to collect rock and soil samples for later return to Earth by a future robotic mission.
Under the new internal NASA agreement, the 2001 Lander will now also be a platform for instruments and technology experiments designed to provide key insights to decisions regarding successful and cost-effective human missions to Mars. Hardware on the lander will be used for an in-situ demonstration test of rocket propellant production using gases in the Martian atmosphere. Other equipment will characterize the planet's soil properties and surface radiation environment.
"Before we can send humans into deep space, we need to understand the nature of the space environment and its effect on living systems," said Arnauld Nicogossian, M.D., acting associate administrator for life and \Jmicrogravity\j sciences. "The Mars 2001 mission will give us invaluable information about the radiation environment of space and the surface on Mars."
Analyses of Martian dust and soil are necessary to understand any interactions with the systems currently planned that will supply the habitation and working environment for future human explorers.
A companion mission to the lander known as the Mars Surveyor 2001 Orbiter will be launched in March 2001. The 2001 Orbiter will be the first to use the atmosphere of Mars to slow down and directly capture the \Jspacecraft\j into orbit, in a technique called aerocapture. The scientific objectives of the mission are to conduct mineralogical mapping of the entire planet and characterize its orbital radiation environment. The 2001 Orbiter also will carry a radio relay to support the lander and a possible Russian robotic rover mission.
The preliminary cost estimate for both integrated missions is approximately $311 million, not including launch costs. An integrated team of the Jet Propulsion Laboratory (JPL), Pasadena, CA; the Johnson Space Center, Houston, TX; and Lockheed Martin Astronautics, \JDenver\j, will develop the missions, led by JPL.
Both of the 2001 missions are part of an ongoing NASA series of robotic Mars exploration \Jspacecraft\j that began with the launches of the Mars Global Surveyor orbiter and the Mars Pathfinder lander in November and December 1996, respectively. Mars Pathfinder and its 25-pound rover, named Sojourner, will land on Mars in a region called \JAres\j Vallis on July 4, 1997.
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"NASA Managers Set April 3 As Launch Date For The Microgravity Science Laboratory Mission (21 March, 1997)",544,0,0,0
Following completion of the Flight Readiness Review yesterday, NASA managers set April 3 as the official launch date for NASA's \JMicrogravity\j Science Laboratory (MSL-1) mission.
The focus of the upcoming mission, designated STS-83, will concentrate on NASA's efforts to further understand the subtle and complex phenomena associated with the influence of gravity in many aspects of daily life.
The STS-83 flight agenda resembles future work set to take place aboard the International Space Station. STS-83 will be the 22nd flight of Space Shuttle Columbia and the 83rd mission flown since the start of the Space Shuttle program in April 1981.
The crew of mission STS-83 includes: Commander Jim Halsell; Pilot Susan Still; Mission Specialists Janice Voss, Michael Gernhardt and Donald Thomas; and Payload Specialists Roger Crouch and Greg Linteris. Thomas, who suffered a broken ankle following a routine training exercise on Jan. 29, has officially been cleared to fly as planned.
"We are very pleased that Don has been cleared for flight and are confident in his ability to carry out his mission responsibilities," said David C. Leestma, director of Flight Crew Operations.
Cady Coleman, who was training with the STS-83 crew as a backup mission specialist, will return to her previous duties supporting crew habitability activities for the \JAstronaut\j Office.
The launch window for STS-83 on April 3 opens at 2:01 p.m. EST and extends for 2 hours, 30 minutes. Columbia's mission duration is planned for 15 days, 16 hours. The STS-83 mission will conclude with Columbia's landing at Kennedy Space Center, FL, on April 19 at about 7:30 a.m. EDT.
Performing beyond expectations, the high- resolution mirrors for NASA's most powerful orbiting X-ray \Jtelescope\j have successfully completed initial testing at Marshall Space Flight Center's X-ray \JCalibration\j Facility, \JHuntsville\j, AL.
"We have the first ground test images ever generated by the \Jtelescope\j's mirror assembly, and they are as good as -- or better than -- expected," said Dr. Martin Weisskopf, Marshall's chief scientist for NASA's Advanced X-ray Astrophysics Facility (AXAF).
The mirror assembly, four pairs of precisely shaped and aligned cylindrical mirrors, will form the heart of NASA's third great observatory.
The X-ray \Jtelescope\j produces an image by directing incoming X-rays to detectors at a focal point some 30 feet beyond the \Jtelescope\j's mirrors. The greater the percentage of X-rays brought to focus and the smaller the size of the focal spot, the sharper the image.
Tests show that on orbit, the mirror assembly of the Advanced X-ray Astrophysics Facility will be able to focus approximately 70 percent of X-rays from a source to a spot less than one-half arc second in radius. The \Jtelescope\j's resolution is equivalent to being able to read the text of a newspaper from half a mile away.
"The \Jtelescope\j's focus is very clear, very sharp," said Weisskopf. "It will be able to show us details of very distant sources that we know are out there, but haven't been able to see clearly."
In comparison, previous X-ray telescopes -- Einstein and Rosat -- were only capable of focusing X- rays to five arc seconds. The Advanced X-ray \JTelescope\j's resolving power is ten times greater.
"Images from the new \Jtelescope\j will allow us to make major advances toward understanding how exploding stars create and disperse many of the elements necessary for new solar systems and for life itself," said Dr. Harvey Tananbaum, director of the Advanced X- ray Astrophysics Facility Science Center at the Smithsonian Astrophysical Observatory, in Cambridge, MA -- responsible for the \Jtelescope\j's science mission.
"We will observe X-rays generated when stars are torn apart by the incredibly strong gravity around massive black holes in the centers of galaxies," added Tananbaum.
On a larger scale, the \Jtelescope\j will play a vital role in answering fundamental questions about the universe. "The superior quality of the mirrors will allow us to see and measure the details of hot gas clouds in clusters of galaxies, giving us a much better idea of the age and size of the universe," said Dr. Leon Van Speybroeck, \JTelescope\j Scientist at the Smithsonian Observatory.
"These same observations also will measure the amount of dark matter present, providing unique insight into one of nature's great puzzles," said Van Speybroeck.
A second phase of testing is now underway at Marshall. \JCalibration\j of the observatory's science instruments began in mid-February. "This phase of testing," said Weisskopf, "includes two focal plane instruments and two sets of gratings used to analyze images and energy distributions from cosmic sources seen by the telescope."
Working around the clock, test teams are taking measurements and studying results. "It is very exciting," said Weisskopf. "With more than 1,200 measurements taken, there is already a tremendous amount of information for study."
The \Jcalibration\j process will end around late April. The mirror assembly then will be shipped to TRW Space and \JElectronics\j Group, Redondo Beach, CA -- NASA's prime contractor for the program -- for \Jintegration\j into the \Jspacecraft\j. The science instruments will remain at Marshall for several more weeks of testing before being shipped to Ball Aerospace and Technologies Corporation in Boulder, CO, where they will be integrated into the science instrument module before being shipped to TRW.
The Advanced X-ray Astrophysics Facility is scheduled for launch in August 1998 and will join NASA's Hubble Space \JTelescope\j and Compton Gamma-ray Observatory in exploring the universe.
Marshall manages development of the observatory for the Office of Space Science, NASA Headquarters, Washington, DC. Using glass purchased from Schott Glaswerke,
Mainz, \JGermany\j, the \Jtelescope\j's mirrors were built by Hughes Danbury Optical Systems, Danbury, CT. The mirrors were coated by Optical Coating Laboratory, Inc., Santa Rosa, CA; and assembled by Eastman-Kodak Company, Rochester, NY.
The AXAF CCD Imaging Spectrometer instrument was developed by \JPennsylvania\j State University, University Park, and the \JMassachusetts\j Institute of Technology (MIT), Cambridge, MA. One of the two gratings was developed by MIT. The other was developed by the Space Research Organization Netherlands, \JUtrecht\j, Netherlands, in collaboration with the Max Planck Institute, Garching, \JGermany\j. The High Resolution Camera instrument was built by the Smithsonian Astrophysical Observatory.
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"U.S. Presence In Space Anniversary (17 March, 1997)",546,0,0,0
Saturday, March 22, marks the one-year anniversary of a continuous U.S. presence in space, which began with the launch of \Jastronaut\j Shannon Lucid aboard Space Shuttle Atlantis on the STS- 76 mission to the Mir space station.
Since Lucid arrived on Mir, astronauts John Blaha and Jerry Linenger have followed in her footsteps, conducting continuous scientific experiments aboard the Russian complex as a precursor to the development and occupancy of the International Space Station.
Linenger will remain aboard Mir until mid-May, when he will be replaced by \Jastronaut\j Mike Foale, who, in turn, will be replaced in September by \Jastronaut\j Wendy Lawrence. The final U.S. \Jastronaut\j scheduled for a tour of duty on the Mir is David Wolf in early 1998.
Former \Jastronaut\j Norm Thagard was the first U.S. \Jastronaut\j to live and work on the Mir. Thagard spent four months on the Russian outpost in 1995. Lucid spent a U.S. - record 188 days in space from the time of her launch on March 22, 1996, to her return to Earth on the STS-79 mission on Sept. 26, 1996. Blaha, who arrived on the STS-79 mission on Sept. 16 last year, spent 128 days in space, returning to Earth aboard Atlantis at the completion of the STS-81 mission on Jan. 22, 1997. Linenger was launched on the STS-81 mission on Jan. 12.
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"Lunar Prospector Spacecraft Construction Complete (12 March, 1997)",547,0,0,0
Construction and assembly of NASA's Lunar Prospector \Jspacecraft\j, designed to obtain the first complete compositional and gravity maps of the Moon, has been completed in preparation for its scheduled September 1997 launch.
Functional and environmental \Jspacecraft\j tests will be conducted over the next several months, according to project manager Tom Dougherty of Lockheed Martin Missiles & Space, Sunnyvale, CA. Once this activity is successfully completed, current plans call for the \Jspacecraft\j to be shipped to Spaceport \JFlorida\j in late August for launch on September 24, 1997.
"We're delighted with progress to date," said Scott Hubbard, NASA Lunar Prospector mission manager at Ames Research Center, Mountain View, CA. "Lockheed Martin and its construction team put a detailed program into place and executed it well within the established schedule and with tight cost control."
The total cost of the mission to NASA, including launch, mission operations and data analysis, is $63 million.
Why is NASA going back to the Moon? Despite a high level of scientific and public interest, particularly during the Apollo era, major gaps remain in scientific knowledge about Earth's nearest planetary neighbor, according to project scientists. Over 75 percent of the lunar surface is not mapped in detail, and important questions about the Moon's history, composition and internal processes remain unanswered.
During its planned one-year polar orbiting mission, Lunar Prospector will map the Moon's surface composition, gravity and magnetic fields, and try to detect volatile release activity. This information should provide insights into the origin and evolution of the Moon. Lunar Prospector also should directly determine the existence or absence of water ice in the Moon's polar regions, which has been suggested by analysis of indirect, radar-based data from the Clementine mission.
As the first peer-reviewed, competitively selected mission in NASA's "faster, better, cheaper" Discovery Program series, Lunar Prospector is an embodiment of the Agency's new way of doing business. With an emphasis on minimized risk, lowered costs, and rapid turnaround time, and its prime focus on delivery of science data, Lunar Prospector will help usher in a new era of Solar System exploration missions.
"Lunar Prospector is serving as a pathfinder in many different ways," said Hubbard. The mission has "already made history in terms of management style, technical approach, cost management and focused science. Technical insight rather than detailed programmatic oversight was used to ensure innovation and maximum return on investment. The Ames program office paid close attention to the progress of the project and its schedule, cost and science return, but provided no detailed specifications. The Principal Investigator was given the flexibility to implement the best available approach," he said.
The Lunar Prospector \Jspacecraft\j is a small, spin-stabilized vehicle with a fully fueled mass of 660 pounds. It is 4.5 feet high and 4 feet in diameter, with three 8-foot booms or masts. Solar cells mounted on its outer surface will provide more than 200 watts of power.
Five scientific instruments are mounted on the booms to isolate them from the main structure and \Jelectronics\j. A neutron spectrometer will have the capability to locate as little as one cup of water in about a cubic yard of lunar soil (regolith). The discovery of water ice in the lunar polar regions would mean that water, necessary for life support and a potential source of both oxygen and \Jhydrogen\j to produce rocket propellant, could be available for use by future lunar explorers.
A gamma-ray spectrometer will provide global maps of the elemental composition of the surface layer of the Moon. Improved knowledge of the concentrations of such elements as uranium, thorium, \Jpotassium\j, iron, \Jtitanium\j, oxygen, silicon, aluminum, \Jmagnesium\j and \Jcalcium\j will aid in understanding the composition and evolution of the lunar crust.
An alpha particle experiment will provide information on the level of tectonic and volcanic lunar out-gassing activity. It will map the locations and frequency of \Jradon\j gas release events on the Moon, a body thought to be tectonically and volcanically dead until Apollo provided evidence that it may still have some limited activity.
A magnetometer and \Jelectron\j reflectometer will map local lunar magnetic fields, known to be weak compared to the global magnetic field of the Earth. This will help determine the origin of such fields and may provide information on the size and composition of the lunar core.
The Doppler gravity experiment will provide the first global gravity map of the Moon, essential for planning follow-on robotic and human exploration missions. It also will provide data on density differences in the crust, internal densities and the nature of the core.
When Lunar Prospector is launched, it will take five days to reach the Moon, making two midcourse maneuvers, deploying booms, and collecting \Jcalibration\j data via its science instruments en route. Once the \Jspacecraft\j reaches the Moon, it will be put into a circular, 118-minute, 62-mile altitude, polar-mapping orbit to begin its mission.
If fuel is available at the end of the one-year nominal mission, lunar mapping may be extended at altitudes as low as 6.2 miles over areas of special interest. After the fuel needed for orbital maintenance is depleted, the \Jspacecraft\j will eventually impact on the lunar surface.
The Lunar Prospector mission is being implemented for NASA by Lockheed Martin, Sunnyvale, CA, with important contributions from Los Alamos National Laboratory, the University of California- Berkeley Space Science Laboratory, the Goddard Space Flight Center, Greenbelt, MD, and the Jet Propulsion Laboratory, Pasadena, CA.
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"NASA-ESA Agreement Enhances Station With Additional Node (10 March, 1997)",548,0,0,0
The planned final configuration of the International Space Station has been enhanced under a recent agreement signed by NASA with the European Space Agency (ESA) that will have \JESA\j construct two station docking nodes, one of which is a new addition, in exchange for the planned NASA launch of the station's ESA-supplied Columbus laboratory module.
The launch-offset barter agreement, a type of agreement common within the International Space Station Program, exchanges \JESA\j services to construct the nodes as payment to NASA for the launch of the Columbus module. Under the agreement, \JESA\j will supply Node 2, a docking node planned to be launched in mid-2000 and connected to the United States Laboratory Module, and Node 3, a new addition to the station planned to be launched after station assembly is completed in the current assembly sequence.
Node 3 will attach to the station's habitation module and provide valuable additional docking ports to the orbital outpost.
Construction of the two nodes will be delegated to the Italian Space Agency. Plans to convert the Node 1 structural test article, located at Boeing's facilities at the Marshall Space Flight Center, \JHuntsville\j, AL, into the Node 2 flight article will be revised accordingly.
Although the construction of the two nodes by \JESA\j is the primary service supplied to NASA under the barter agreement, the agreement also includes requirements for \JESA\j to supply a crew refrigerator/freezer unit for the station's habitation module; a cryogenic freezer unit for the U.S. laboratory module; and a variety of other minor hardware.
The next step is for NASA and \JESA\j to negotiate an implementing arrangement for this activity.
Astronaut Terrence W. Wilcutt (Lt. Colonel, USMC) will command the eighth of nine planned missions to dock the Space Shuttle with \JRussia\j's Mir space station. STS-89 is targeted for a January 1998 launch on Discovery.
Three members of the 1995 \JAstronaut\j Class and a veteran space flyer round out the crew. Joining Wilcutt on the flight deck will be Pilot Joe Frank Edwards, Jr. (Cmdr., USN). Mission Specialists for the flight are Bonnie J. Dunbar, Ph.D., Michael P. Anderson (Major, USAF) and James F. Reilly, II, Ph.D.
Discovery also will carry Mission Specialist David Wolf, M.D, to Mir to begin a planned four-month stay. Wolf will replace \JAstronaut\j Wendy Lawrence, scheduled to arrive at Mir in September 1997 during the STS-86 mission aboard Atlantis. Lawrence will return to Earth on board Discovery as a member of the STS-89 crew.
STS-89 will mark Wilcutt's third space flight and his second visit to Mir. He was the pilot on STS-79, the fourth Shuttle/Mir flight which saw the first exchange of U.S. crew members on the Mir, leaving John Blaha and returning Shannon Lucid to Earth. Wilcutt's first flight was STS-68 in 1994.
Dunbar is a veteran \Jastronaut\j with four space flights to her credit. She flew on STS-61A in 1985, STS-32 in 1990, STS-50 in 1992, and the first Shuttle/Mir docking mission, STS-71 in 1995. Edwards, Anderson and Reilly will be making their first trip to space during STS-89.
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"Pioneer 10 Spacecraft Nears 25TH Anniversary, End Of Mission",550,0,0,0
A major milestone for humanity's most distant and longest- lived interplanetary explorer will occur on March 2, 1997, when NASA's hardy Pioneer 10 \Jspacecraft\j reaches its 25th anniversary in space.
"Pioneer 10 exemplifies the American pioneering spirit of exploration far beyond the frontier," said Dr. Wesley T. Huntress, Jr., Associate Administrator for Space Science at NASA Headquarters, Washington, DC. "Not only has it made many major scientific discoveries in the far reaches of space, we're proud that it has managed to stay alive almost ten times longer than the original mission called for, a tribute to the designers and builders at TRW, and the operators at NASA's Ames Research Center.
"NASA operated the Pioneer 10 mission as long as it had enough power to return science data about the conditions in space as far from Earth as possible. We will end the science mission at the end of March because the power has finally become too weak to do significant science," Huntress said.
Launched from Cape Kennedy on March 2, 1972, aboard an Atlas \JCentaur\j rocket for what had been planned as a two-year mission to Jupiter, Pioneer 10 is now so far away that its radio signal, traveling at the speed of light (186,000 miles per second), takes nine hours and ten minutes to reach the Earth. Currently twice as far from the Sun as Pluto, Pioneer is returning data about the farthest reaches of the Sun's atmosphere.
Now 6.2 billion miles from Earth, Pioneer 10, built by TRW Space and \JElectronics\j Group, Redondo Beach, CA, was the first \Jspacecraft\j to travel through the asteroid belt and explore the outer solar system, the first \Jspacecraft\j to visit Jupiter, the first to use a planet's gravity to change its course and to reach solar-system-escape velocity, and the first \Jspacecraft\j to pass beyond the known planets.
Now traveling at 28,000 miles per hour, Pioneer 10 is recording the intensity of galactic cosmic rays in the outer heliosphere (the region of solar wind influence) and searching for the heliopause, the true outer boundary of the solar system where the flow of hot gas from the Sun (the solar wind) bumps into the interstellar medium. The \Jspacecraft\j is heading in the direction of the long "tail" of the teardrop-shaped heliosphere.
Many scientists rank the first crossing of the asteroid belt between Mars and Jupiter as one of Pioneer 10's major achievements. Before the crossing, no one knew how many rocks, as well as grains of sand, speeding through space at thousands of miles per hour would impact and possibly disable the \Jspacecraft\j. Pioneer 10 made the crossing nearly unscathed, thus opening the way for other \Jspacecraft\j to explore beyond Mars, including its sister Pioneer 11, the twin Voyager \Jspacecraft\j, the \JGalileo\j mission to Jupiter and, later this year, the Cassini mission to Saturn.
The \Jspacecraft\j also survived Jupiter's intense radiation belts when it flew safely by the giant planet in December 1973. Providing the best information of the planet obtained to that date, Pioneer 10's instruments studied ultraviolet and infrared radiation and charged particles, and provided the first "close- ups" of Jupiter and its moons. It also was the first to measure the planet's giant radiation belts, magnetic field and magnetosphere, as well as its atmosphere and interior. The \Jspacecraft\j measured the exact masses of Jupiter and its four planet-sized moons, Io, Europa, Ganymede and Callisto.
After its Jupiter flyby, the \Jspacecraft\j continued its mission of exploration for over two decades, and in 1983 became the first \Jspacecraft\j to travel beyond the outermost planets, Neptune and Pluto. Pioneer 10 will have its first "near-star- encounter" in about 30,000 years when it will pass within approximately three light years of the red dwarf star Ross 248 in the \Jconstellation\j \JTaurus\j. In the next million years, Pioneer 10 will pass ten stars at distances ranging from three to nine light years, and will probably still be traveling through the Milky Way galaxy when the Sun becomes a red giant and destroys the Earth five billion years from now.
Pioneer carries a message for any intelligent life forms that it might encounter on its trek across the galaxy. A gold- anodized aluminum plaque designed by Dr. Frank Drake and the late Dr. Carl Sagan is bolted to the \Jspacecraft\j. The plaque's \Jengraving\j depicts a man and a woman, a map of Earth's solar system, and other symbols which may help intelligent beings interpret the message and understand something about the \Jspacecraft\j's creators, and where they lived.
The 570-lb. \Jspacecraft\j carries 11 instruments that have been used to measure magnetic fields, solar wind, high energy cosmic rays, cosmic and asteroidal dust, and Jupiter's ultraviolet and infrared radiation. Pioneer 10 obtains power from four \Jradioisotope\j thermoelectric generators that originally supplied 160 watts, but are now two-thirds degraded. The \Jspacecraft\j is spin-stabilized and has a nine-foot dish antenna. Pioneer's 8- watt signal, equal to the power of a night light, now reaches the 70-meter antennae and sensitive receivers at NASA's Deep Space Network with the strength of .3 billionths of a trillionth of a watt.
The Pioneer 10 mission is managed by the Ames Research Center, Mountain View, CA, for the Office of Space Science, NASA Headquarters, Washington, DC.
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"Mir, Fire Extinguished (24 February, 1997)",551,0,0,0
A problem with an oxygen-generating device on the Mir space station last night set off fire alarms and caused minor damage to some hardware on the station. No injuries to any of the six crewmembers on board were reported. The fire was located in the Kvant 1 module.
The fire, which began at 10:35 p.m. Sunday, Moscow time, burned for about 90 seconds. The crew was exposed to heavy smoke for five to seven minutes and donned masks in response. After completing physical exams of everyone on board, U.S. \Jastronaut\j Jerry Linenger, a physician, reported that all crewmembers are in good health. Medical personnel have directed them to wear goggles and masks until an analysis of the Mir atmosphere has been completed.
Lithium perchlorate candles are burned to generate supplemental oxygen when more than three people are on board the space station. The oxygen-generating candles usually burn for five to 20 minutes. Russian officials believe the problem began when a crack in the oxygen generator's shell allowed the contents of the cartridge to leak into the hardware in which it was located. Crewmembers extinguished the fire with \Jfoam\j from three fire extinguishers, each containing two liters of a water-based liquid.
The damage to some of Mir's hardware resulted from excessive heat rather than from open flame. The heat destroyed the hardware in which the device, known as a "candle," was burning, as well as the panel covering the device. The crew also reported that the outer \Jinsulation\j layers on various cables were melted by the heat. It is reported by Russian flight controllers that all Mir systems continue to operate normally, however.
"It is unfortunate that this incident occurred, but we are thankful that there were no injuries," said Frank Culbertson, Director of the Phase One Shuttle-Mir program. "Russian management and operations specialists have been very informative as to what happened, and we are working closely with them on evaluating the health of the crew and how best to respond to the damage," added Culbertson.
"The crew did a great job handling the fire, and the ground support has been excellent on both sides."
In addition to Linenger, the Mir crewmembers include Mir 22 cosmonauts Valery Korzun and Alexander Kaleri, Mir 23 cosmonauts Vasily Tsibliev and Alexander Lazutkin, and German researcher Reinhold Ewald, representing the German space agency, DARA. Korzun, Kaleri and Ewald are scheduled to return to Earth on Sunday as previously planned to wrap up a six-month mission for Korzun and Kaleri and three weeks of scientific experiments for Ewald. Linenger will remain aboard Mir until mid-May with Tsibliev and Lazutkin.
Officials are evaluating possible impacts to the mission and its science activities, as technical experts at the Russian Mission Control Center investigate the incident. The burned panel and other materials may be returned to Earth with Korzun, Kaleri and Ewald on Sunday for further analysis.
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"Sun's Place In The Galaxy (14 February, 1997)",552,0,0,0
An action-packed movie assembled from images taken by an instrument aboard the NASA-European Space Agency Solar Heliospheric Observatory (SOHO) has provided a remarkable galactic perspective on the Sun and its place in the Milky Way.
Taken during Dec. 22-27, 1996, the series of images show the Sun drifting in front of the stars of the \Jconstellation\j \JSagittarius\j, as the constant solar wind blows outward in all directions. Soon, a \Jcomet\j passes into view from the south and disappears behind the Sun. Finally, in an unrelated event, a plainly visible giant puff of solar gas is emitted, representing a large mass ejection in a direction away from the Earth.
The remarkable images come from SOHO's visible-light \Jcoronagraph\j, LASCO, which is able to mask the intense rays from the Sun's surface in order to reveal the much fainter glow of the solar atmosphere, or corona. Operated with its widest field of view, LASCO's unprecedented sensitivity enabled it to see the thin ionized gas of the solar wind out to the edges of the picture, 13 million miles from the Sun's surface. Many stars are brighter than the gas, and they create the background scene.
The results may alter human perceptions of the Sun in the same way that the Apollo lunar mission photographs revealed the Earth to be a beautiful but isolated planet in space, according to leader of the LASCO team, Dr. Guenter Brueckner of the Naval Research Laboratory, Washington, DC.
"I spend my life examining the Sun, but this movie is a special thrill," Brueckner said. "For a moment, I forget the years of effort that went into creating LASCO and SOHO, and leave aside the many points of scientific importance in the images. I am happy to marvel at a new impression of our busy star that gives us life, and which affects our environment in many ways that we are only now beginning to understand."
For many centuries even astrologers knew that the Sun was in \JSagittarius\j in December and drifting towards the next zodiacal \Jconstellation\j, Capricorn. This was a matter of calculation only, because the Sun's own brightness prevented a direct view of the starfield. The SOHO-LASCO movie makes this elementary point of \Jastronomy\j a matter of direct observation for the first time.
"This is an especially dramatic sample of the data that scientists are now starting to gather routinely from SOHO," said George Withbroe, director of the Sun-Earth Connection science program at NASA Headquarters, Washington, DC. "It really helps drive home the idea that the Sun is a typical star, although we certainly have a special relationship with it."
In the movie, north is at the top of the scene, which corresponds with the orientation of the Sun as seen at midday in the northern hemisphere of Earth. SOHO's progress in orbit around the Sun remains in step with the Earth's motion. SOHO travels towards the right (west) in relation to the stars during the period of observation. As a result, the Sun's position appears to shift to the left (eastwards) in front of the stars. In this mode, LASCO observes an area of the sky 32 times wider than the visible Sun itself.
At the time of the observations, SOHO is looking towards the heart of the Milky Way Galaxy, which lies in the \Jconstellation\j of \JSagittarius\j. The Milky Way, made by the light of billions of distant stars, forms a luminous band slanting down and to the right. Dark lanes seen in the Milky Way are real features familiar to astronomers. They are created by dust clouds in the disk of the galaxy that obscure the distant stars.
A doomed \Jcomet\j, previously unknown, enters on the left of the image on Dec. 22. Its path curves towards the Sun and on Dec. 23 it disappears behind the occulting mask of the \Jcoronagraph\j. It fails to reappear on the far side of the Sun. Whether or not its trajectory took it directly towards the visible surface, the \Jcomet\j must have evaporated in Sun's atmosphere. It was one of a family of comets known as sungrazers, believed to be remnants of a large \Jcomet\j that broke up perhaps 900 years ago. Other fragments were responsible for spectacular \Jcomet\j apparitions in 1843, 1882 and 1965.
Called \JComet\j SOHO 6, it is one of seven sungrazers discovered so far by LASCO. Analyses of these cometary orbits, now in progress, are a prerequisite for their inclusion in the official record of \Jcomet\j discoveries. LASCO also provided unique pictures of \JComet\j Hyakutake passing behind the Sun in early May 1996.
Debris strewn from the tails of many comets makes a disk of dust around the Sun, in the \Jecliptic\j plane where the planets orbit. It scatters sunlight and is sometimes visible at twilight on the Earth, known as the Zodiacal Light. In the raw images obtained by LASCO, the Zodiacal Light is brighter than the solar corona. Image processing subtracted these effects precisely, to bring the solar wind and the Milky Way into plain view.
Random flickers of light in the images are due to cosmic rays striking the detector. Cosmic rays are energetic particles coming from exploded stars in the Milky Way, and variations in the solar wind influence their intensity in the vicinity of SOHO and the Earth. Operating beyond the Earth's protective magnetic field, which repels many particles, SOHO is more exposed to the cosmic rays.
In the largest outburst from the Sun seen in the December movie, a mass ejection causes billions of tons of gas to race out into space on the right-hand (western) side of the Sun. The origin of this event much lower in the Sun's atmosphere was evident in an expanding bubble seen in processed images from the SOHO extreme ultraviolet imager.
Coronal mass ejections are the hurricanes of space weather. SOHO is ideally placed to report and even anticipate their origins in the Sun's atmosphere. Although the Sun is supposedly very quiet at present, being close to the minimum count of sunspots, LASCO has observed so many outbursts large and small -- roughly one a day -- that scientists are having to think again about how to define a coronal mass ejection. Later LASCO images, from Jan. 6, 1997, revealed a large mass ejection directed towards the Earth.
A mission of international scientific cooperation between NASA and the European Space Agency (ESA), SOHO was launched on Dec. 2, 1995. It is situated near the L-1 Lagrangian point, where the Earth's and Sun's gravitational forces balance, some one million miles sunward from the Earth. This vantage point enables solar astronomers to use SOHO's 12 science instruments to observe the Sun continuously, with no intervening "night." Major areas of research include studies of the solar interior via helioseismology, the solar atmosphere in ultraviolet and visible light, and the solar wind and energetic particles.
SOHO and its instruments are operated from the Goddard Space Flight Center, Greenbelt, MD, for the NASA's Office of Space Science, Washington, DC.
Brueckner will discuss the video and its contents in detail in a briefing at 6 p.m. EST on Saturday, Feb. 15, during the annual meeting of the American Association for the Advancement of Science at the Seattle Convention Center. For more information, contact the AAAS press room at 206/624- 9087.
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"Radio Telescope Larger Than Earth (7 February, 1997)",553,0,0,0
NASA and the National Radio \JAstronomy\j Observatory are joining with an international consortium of space agencies to support the launch of a Japanese satellite next week that will create the largest astronomical "instrument" ever built -- a radio \Jtelescope\j more than two-and-a-half times the diameter of the Earth that will give astronomers their sharpest view yet of the universe.
The launch of the Very Long Baseline Interferometry (VLBI) Space Observatory Program (VSOP) satellite by \JJapan\j's Institute of Space and Astronautical Science (ISAS) is scheduled for Feb. 10 at 11:50 p.m. EST (1:50 p.m. Feb. 11, \JJapan\j time.)
The satellite is part of an international collaboration led by ISAS and backed by \JJapan\j's National Astronomical Observatory; NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA; the National Science Foundation's National Radio \JAstronomy\j Observatory (NRAO), Socorro, NM; the Canadian Space Agency; the \JAustralia\j \JTelescope\j National Facility; the European VLBI Network and the Joint Institute for Very Long Baseline Interferometry in Europe.
Very long baseline interferometry is a technique used by radio astronomers to electronically link widely separated radio telescopes together so they work as if they were a single instrument with extraordinarily sharp "vision," or resolving power. The wider the distance between telescopes, the greater the resolving power. By taking this technique into space for the first time, astronomers will approximately triple the resolving power previously available with only ground-based telescopes. The satellite system will have resolving power almost 1,000 times greater than the Hubble Space \JTelescope\j at optical wavelengths. The satellite's resolving power is equivalent to being able to see a grain of rice in \JTokyo\j from Los Angeles.
"Using space VLBI, we can probe the cores of quasars and active galaxies, believed to be powered by super massive black holes," said Dr. Robert Preston, project scientist for the U.S. Space Very Long Baseline Interferometry project at JPL. "Observations of cosmic masers -- naturally-occurring microwave radio amplifiers -- will tell us new things about the process of star formation and activity in the heart of other galaxies."
"By the 1980s, radio astronomers were observing the universe with assemblages of radio telescopes whose resolving power was limited only by the size of the Earth. Now, through a magnificent international effort, we will be able to break this barrier and see fine details of celestial objects that are beyond the reach of a purely ground-based \Jtelescope\j array. We anticipate a rich harvest of new scientific knowledge from VSOP," said Dr. Paul Vanden Bout, Director of NRAO.
In the first weeks after launch, scientists and engineers will "test the deployment of the reflecting mesh \Jtelescope\j in orbit, the wide-band data link from the satellite to the ground, the performance of the low noise amplifiers in orbit, and the high-precision orbit determination and attitude control necessary for VLBI observations with an orbiting telescope," according to Dr. Joel Smith, manager of the U.S. Space VLBI project at JPL. Scientific observations are expected to begin in May.
The 26-foot diameter orbiting radio \Jtelescope\j will observe celestial radio sources in concert with a number of the world's ground-based radio telescopes. The 1,830-pound satellite will be launched from ISAS' Kagoshima Space Center, at the southern tip of \JKyushu\j, one of \JJapan\j's main islands, and will be the first launch with ISAS' new M-5 series rocket.
The satellite will go into an elliptical orbit, varying between 620 to 12,400 miles above the Earth's surface. This orbit provides a wide range of distances between the satellite and ground-based telescopes, which is important for producing a high- quality image of the radio source being observed. One orbit of the Earth will take about six hours.
The satellite's observations will concentrate on some of the most distant and intriguing objects in the universe, where the extremely sharp radio "vision" of the new system can provide much-needed information about a number of astronomical mysteries.
For years, astronomers have known that powerful "engines" in the hearts of quasars and many galaxies are pouring out tremendous amounts of energy. They suspect that supermassive black holes, with gravitational fields so strong that not even light can escape them, lie in the centers of these "engines." The mechanism at work in the centers of quasars and active galaxies, however, remains a mystery. Ground- based radio telescopes, notably NRAO's Very Long Baseline Array (VLBA), have revealed fascinating new details in recent years, and VSOP is expected to add a wealth of new information on these objects, millions or billions of light-years distant from Earth.
Many of these same objects act as super-powerful particle accelerators to eject "jets" of subatomic particles at nearly the speed of light. Scientists plan to use VSOP to monitor the changes and motions in these jets to learn more about how they originate and interact with their surroundings.
The satellite also will aim at regions in the sky where giant collections of water and other molecules act as natural amplifiers of radio emission much as lasers amplify light. These regions, called cosmic masers, are found in areas where new stars are forming and near the centers of galaxies. Observations can provide the detail needed to measure motions of individual \Jmaser\j "spots" within these regions, and provide exciting new information about the star-forming regions and the galaxies where the masers reside. In addition, high-resolution studies of cosmic masers can allow astronomers to calculate distances to them with unprecedented accuracy, and thus help resolve continuing questions about the size and age of the universe.
The project is a major international undertaking, with about 40 radio telescopes from more than 15 countries having committed time to co-observe with the satellite. This includes the National Science Foundation's Very Long Baseline Array (VLBA), an array of 10 telescopes spanning the United States from Hawaii to Saint Croix; NASA's Deep Space Network (DSN) sites in \JCalifornia\j, \JSpain\j, and \JAustralia\j; the European VLBI Network, more than a dozen telescopes ranging from the United Kingdom to China; a Southern Hemisphere array of telescopes stretching from eastern \JAustralia\j to South Africa; and \JJapan\j's network of domestic radio telescopes.
In the United States, NASA is funding critical roles in the VSOP mission at both JPL and NRAO. JPL has built an array of three new tracking stations at its DSN sites in Goldstone, CA; Madrid, \JSpain\j; and near \JCanberra\j, \JAustralia\j. A large existing tracking station at each of these sites has also been converted to an extremely sensitive radio \Jtelescope\j for simultaneous observations with the satellite. JPL also is providing precision orbit determination, scientific and operational planning support to the Japanese, and advice to U.S. astronomers who wish to observe with the satellite. NRAO is building a new tracking station at Green Bank, WV; contributing observing time on the VLBA array of telescopes; modifying existing data analysis hardware and software, and aiding astronomers with the analysis of the VSOP data. Much of the observational data will be processed at NRAO's facility in Socorro, NM, using the VLBA Correlator, a special purpose high-performance computer designed to process VLBI data.
VSOP is the culmination of many years of planning and work by scientists and engineers around the world. Tests using NASA's Tracking and Data Relay Satellite System (TDRSS) proved the feasibility of space VLBI in 1986. Just last year, those old data were used again to test successfully the data- reduction facilities for VSOP.
JPL manages the U.S. Space Very Long Baseline Interferometry project for NASA's Office of Space Science, Washington, DC. The VLBA, headquartered in Socorro, NM, is part of the National Radio \JAstronomy\j Observatory, a facility of the National Science Foundation, operated under cooperative agreement by Associated Universities, Inc.
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"NASA Managers Set Feb. 11 As Launch Date For Second Shuttle Servicing Mission To Hubble Space Telescope (31 January, 1997)",554,0,0,0
Following completion of a flight readiness review meeting, NASA managers set Feb. 11 as the official launch date for NASA's second Shuttle mission of the year.
The mission, designated STS-82, is the second in a series of planned servicing missions to the orbiting Hubble Space \JTelescope\j (HST). Following rendezvous with and retrieval of the HST on the third day of the mission, four space walks on four successive days will take place as the astronauts remove and replace various HST components. Work performed on the \Jtelescope\j will significantly upgrade the scientific capabilities of the HST and keep the \Jtelescope\j functioning smoothly until the next scheduled servicing mission in 1999.
The launch window on Feb. 11 opens at 3:56 a.m. EST and extends for 65 minutes. Discovery's mission duration is planned for nine days, 22 hours, 47 minutes. The STS-82 mission will conclude with Discovery returning home to the Kennedy Space Center, FL, on Feb. 21 at about 2:43 a.m. EST.
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"Planetary Astronomers Start Year With Two Discoveries (29 January, 1997)",555,0,0,0
Two newly detected members of the Solar System -- a rare asteroid orbiting close to Earth and a distant \Jcomet\j making its only appearance -- mark the first discoveries of the year for a team of astronomers at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
The discoveries, reported Jan. 10 by JPL planetary scientists Eleanor Helin, Steve Pravdo, David Rabinowitz and Ken Lawrence, were made possible with a few nights of clear observing weather and use of a sensitive, charge-coupled device (CCD) camera called the Near-Earth Asteroid Tracking (NEAT) system at Mt. Haleakala, Maui, HI. Since their initial sightings, both objects have become the focus of worldwide observations by astronomers in \JJapan\j, China, \JAustralia\j, Canada, \JItaly\j and the Czech Republic.
"This asteroid is a member of a rare class of \Jasteroids\j, called Atens, which stay within Earth's orbit most of their lifetimes," said Helin, principal investigator of the NEAT project. "The object has a higher inclination to the plane of Earth's orbit than most Atens; in fact, at 31 degrees, it has the second highest inclination of all the Atens we've discovered."
The highly inclined orbit, which is unusual, may result from long-range interactions with the planets, or may be the outcome of previous orbits passing near the Earth. With the discovery of more Atens, the relative importance of these competing influences may be better understood.
Dubbed 1997 AC11, the asteroid is a faint object with an absolute magnitude of 21, and probably measures about 600 feet in diameter. It is only the 24th Aten to be discovered in 21 years, since Helin found and named the first Aten in January 1976. With orbits that are smaller than Earth's, and short periods, Atens are in the vicinity of Earth frequently. This closeness to Earth makes them more likely to impact the planet than other types of \Jasteroids\j.
"Atens never wander far from the orbit of Earth and can cross Earth's orbit as many as four times a year," Helin said. "1997 AC11, for instance, has a period of 8/10ths of a year, or roughly 9.5 months. As we continue to observe it in coming months, we will be able to characterize its orbital path with more precision. With more precise data, we will be able to examine its potential for collision with Earth at some time in the future."
Along with the newest Aten, astronomers also discovered a new \Jcomet\j, still distant but moving toward the Earth and Sun, as it passed through the \Jconstellation\j of Leo. Designated \JComet\j 1997 A1, the celestial snowball is expected to make its closest approach to Earth on Feb. 6, passing at a distance of about 230 million miles, but remaining visible in the night sky for several months thereafter.
"This \Jcomet\j has traveled a long distance, originating in the Oort Cloud, a region far beyond Pluto's orbit which is believed to house trillions of incipient comets," Helin said. "It has a parabolic orbit, which means it will travel through our Solar System once and probably never be seen again. Parabolic comets do not present their calling cards before arriving in the inner Solar System. They appear without warning."
At discovery, 1997 A1 was fairly dim at magnitude 19, and showed a weakly condensed nucleus with a diffuse halo and short tail, Helin said. The Minor \JPlanet\j Center at the Smithsonian Astrophysical Observatory in Cambridge, MA, announced the discovery, reporting it as a parabolic \Jcomet\j, with an orbital inclination of 145 degrees from the \Jecliptic\j plane, and indicated that it would not pass any closer than 3.17 astronomical units (295 million miles) from the Sun.
JPL's NEAT team, in conjunction with another observing effort under way at the Laboratory's Table Mountain Observatory in San Bernardino, CA, will continue to track and characterize the \Jcomet\j over the next several months until it is no longer visible.
During its closest approach on Feb. 6, the newly discovered \Jcomet\j will be visible in the \Jconstellation\j of Cancer and brighten to a magnitude of about 18. Moderate-sized telescopes with CCD chips will be able to observe the \Jcomet\j, Helin said. Astronomers report that the \Jcomet\j is continuing to outgas, or warm up and boil off some of its ices, as it moves toward the Sun.
Discoveries of very faint or distant objects, and those surprisingly close by, are increasing due to the introduction of technologically advanced, fully autonomous CCD telescopes. The NEAT camera, for example, employs a very large, very sensitive 4,096-by-4,096-pixel CCD. The camera is installed on a 39-inch \Jtelescope\j operated at the summit of Mt. Haleakala by the U.S. Air Force.
Using this powerful, fully automated system, astronomers are discovering many more objects than was possible in the past. The January observing run, for instance, produced more than 700 asteroid sightings, including high-inclination inner-belt \Jasteroids\j and a number of potential Mars-crossers, which will be confirmed after more observations become available. Total detections since NEAT began operations in late 1995 have climbed to more than 9,000 objects, of which more than 50 percent are new objects and more than 800 of those have received new designations.
NEAT was built and is being managed by the Jet Propulsion Laboratory for NASA's Office of Space Science, Washington, DC.
#
"STS-82 Set For Second Hubble Telescope Servicing Mission (31 January, 1997)",556,0,0,0
Astronauts on the Space Shuttle Discovery STS-82 mission will significantly upgrade the scientific capabilities of NASA's Hubble Space \JTelescope\j (HST) during the ten-day servicing mission by installing two state-of-the-art instruments. They also will perform maintenance to keep HST functioning smoothly until the next scheduled servicing mission in 1999. STS-82, scheduled to launch Feb. 11, 1997, is the second servicing mission to HST since its deployment in April 1990.
The seven-member crew will conduct at least four spacewalks (also called Extravehicular Activities or EVAs) to remove two older instruments and install two new \Jastronomy\j instruments, as well as other servicing tasks. The two older instruments being replaced are the Goddard High Resolution Spectrometer and the Faint Object Spectrograph. Replacing these instruments are the Space \JTelescope\j Imaging Spectrograph (STIS) and the Near Infrared Camera and Multi-Object Spectrometer (NICMOS). HST's current complement of science instruments includes two cameras, two spectrographs, and fine guidance sensors.
In addition to installing the new instruments, astronauts will replace other existing hardware with upgrades and spares. Hubble will get a refurbished Fine Guidance Sensor, an optical device that is used on HST to provide pointing information for the \Jspacecraft\j and is used as a scientific instrument for astrometric science. The Solid State Recorder (SSR) will replace one of HST's current reel-to-reel tape recorders. The SSR provides much more flexibility than a reel-to-reel recorder and can store ten times more data.
One of Hubble's four Reaction Wheel Assemblies (RWA) will be replaced with a refurbished spare. The RWA is part of Hubble's Pointing Control Subsystem. The RWAs use spin momentum to move the \Jtelescope\j into position. The wheels also maintain the \Jspacecraft\j in a stable position. The wheel axes are oriented so that the \Jtelescope\j can provide science with only three wheels operating, if required.
The STS-82 crew will be commanded by Ken Bowersox, who will be making his fourth Shuttle flight. The pilot is Scott Horowitz, who will be making his second flight. There are five mission specialists assigned to this flight. Joe Tanner, Mission Specialist-1, is making his second flight. Mission Specialist-2, Steve Hawley, is making his fourth flight. Greg Harbaugh, Mission Specialist-3, is making his fourth flight. Mark Lee, Mission Specialist-4, is making his fourth flight. Mission Specialist-5, Steve Smith, is making his second flight. The crew members who will conduct the planned EVAs are Mark Lee, Greg Harbaugh, Steve Smith and Joe Tanner.
Training for this mission began nearly two years ago. Extensive training for the EVAs was conducted at the Johnson Space Center in Houston, TX, in the 25-foot deep Weightless Environment Training Facility; at the Marshall Space Flight Center in \JHuntsville\j, AL, in the 40-foot deep Neutral \JBuoyancy\j Simulator; and at the Goddard Space Flight Center in Greenbelt, MD, in a 12,500 square-foot cleanroom.
During the training, the astronauts practiced every detail of every task they will have to perform during the four spacewalks. They also rehearsed using more than 150 specialized tools and crew aids developed specifically for this mission, ranging from a simple bag for carrying some of the smaller tools to sophisticated, battery-operated power tools. The hand-operated devices allow astronauts to more efficiently perform intricate, labor-intensive tasks. Tools will allow access to equipment bays on both HST and the Shuttle and will help remove and install scientific instruments and other components.
Discovery is targeted for an early morning launch on Feb. 11 from NASA's Kennedy Space Center (KSC) Launch Complex 39-A at 3:56 a.m. EST. The launch window is 61 minutes. With an on-time launch on Feb. 11 and a nominal 10-day mission, Discovery will land back at KSC on Feb. 21 at about 2:43 a.m. EST.
STS-82 will be the 22nd Flight of Discovery and the 82nd mission flown since the start of the Space Shuttle program in April 1981.
New \JAstronomy\j Instruments
The HST was designed to allow new instruments to be easily installed as old ones become obsolete. This was demonstrated during the first servicing mission in December 1993, when, during an 11-day mission that included a record five EVAs, astronauts successfully installed a new camera which had its corrective \Joptics\j built right in, and a special instrument, called the COSTAR (Corrective \JOptics\j Space \JTelescope\j Axial Replacement) that would properly refocus light from the flawed main mirror to the other instruments.
The new instruments installed during this mission will again dramatically expand Hubble's scientific capabilities. The Space \JTelescope\j Imaging Spectrograph (STIS) provides unique and powerful spectroscopic capabilities for the HST. A spectrograph separates the light gathered by the \Jtelescope\j into its spectral components so that the composition, temperature, motion, and other chemical and physical properties of astronomical objects can be analyzed.
STIS's two-dimensional detectors allow the instrument to gather 30 times more spectral data and 500 times more spatial data than existing spectrographs on Hubble which look at one place at a time. One of the greatest advantages to using STIS is in the study of supermassive black holes.
STIS will search for massive black holes by studying the star and gas dynamics around galactic centers. It also will measure the distribution of matter in the universe by studying \Jquasar\j absorption lines, use its high sensitivity and spatial resolution to study star formation in distant galaxies, and perform spectroscopic mapping of solar system objects.
The Near Infrared Camera and Multi-Object Spectrometer (NICMOS) promises to gain valuable new information on the dusty centers of galaxies and the formation of stars and planets. NICMOS consists of three cameras. It will provide the capability for infrared imaging and spectroscopic observations of astronomical targets.
NICMOS will give astronomers their first clear view of the universe at near-infrared wavelengths between 0.8 and 2.5 micrometers -- longer wavelengths than the human eye can see. The expansion of the universe shifts the light from very distant objects toward longer red and infrared wavelengths.
NICMOS's near infrared capabilities will provide views of objects too distant for research by current Hubble optical and ultraviolet instruments. NICMOS's detectors perform more efficiently than previous infrared detectors.
Technological Advances
Besides rapidly advancing scientific understanding of the universe, HST is making direct contributions to the health, safety, and quality of people's lives through a variety of technological spinoffs.
For instance, a new, non-surgical breast \Jbiopsy\j technique using a device originally developed for HST's Imaging Spectrograph (STIS) is now saving women pain, scarring, radiation exposure, time and money. This technique, called stereotactic automated large-core needle \Jbiopsy\j, enables a doctor to precisely locate a suspicious lump and use a needle instead of surgery to remove tissue for study. This precise process is possible because of a key improvement in digital imaging technology known as a Charge Coupled Device or CCD.
Looking Toward the Future
The Hubble Space \JTelescope\j was designed to operate in space for 15 years. Since its deployment in 1990, HST has revolutionized astronomers' vision of the universe more than any prior telescopes. Many new details about planets, stars and galaxies have been revealed in the short span of six years. Hubble provided dramatic and detailed views of \Jcomet\j fragments smashing into Jupiter; clues about the existence of black holes in the core of galaxies; and has made significant progress in determining the age and size of the universe. With astronauts geared to embark on the second mission to service the \Jtelescope\j, scientists are looking forward to even greater capabilities to look deeper into the universe.
Two more servicing missions are planned for 1999 and 2002 to keep the \Jtelescope\j functioning efficiently and to improve its scientific capability. Among the tasks for the 1999 mission are installation of the Advanced Camera, new solar arrays, and the \Jtelescope\j will be reboosted into a higher orbit. In 2002, plans call for installation of an as-yet undefined advanced scientific instrument as well as maintenance to keep HST functioning until at least 2005.
The HST is a joint project between NASA and the European Space Agency.
Rollout of the Space Shuttle Discovery has resumed following evaluation of a 24-foot long crack on the Mobile Launch Platform (MLP). Structural engineers have determined the integrity of the MLP has not been compromised. Discovery's trip to the launch pad resumed shortly after noon, EST.
Rollout of Discovery began shortly after 7 a.m. EST from the Vehicle Assembly Building at the Kennedy Space Center, FL. Rollout was stopped at approximately 8:25 a.m. EST after engineers heard a "loud bang" and noticed that a crack had developed on the MLP. The "Y"-shaped crack is on the surface of the MLP and runs from near the left hand SRB flame hole toward the near corner of the MLP.
Discovery is scheduled for launch on mission STS-82 on Feb. 11.
Ice-spewing volcanoes and the grinding and tearing of tectonic plates have reshaped the chaotic surface of Jupiter's frozen moon Europa, images from NASA's \JGalileo\j \Jspacecraft\j reveal.
The images, captured when \JGalileo\j flew within just 430 miles (692 kilometers) of Europa on Dec. 19, were released at a news briefing today at NASA Headquarters, Washington, DC.
Although the images do not show currently active ice volcanoes or geysers, they do reveal flows of material on the surface that probably originated from them, said \JGalileo\j imaging team member Dr. Ronald Greeley of \JArizona\j State University, Tempe.
"This is the first time we've seen actual ice flows on any of the moons of Jupiter," said Greeley. "These flows, as well as dark scarring on some of Europa's cracks and ridges, appear to be remnants of ice volcanoes or geysers."
The new images appear to enhance Europa's prospects as one of the places in the Solar System that could have hosted the development of life, said Greeley.
"There are three main criteria to consider when you are looking for the possibility of life outside the Earth -- the presence of water, organic compounds and adequate heat," said Greeley. "Europa obviously has substantial water ice, and organic compounds are known to be prevalent in the Solar System. The big question mark has been how much heat is generated in the interior.
"These new images demonstrate that there was enough heat to drive the flows on the surface. Europa thus has a high potential to meet the criteria for exobiology," Greeley added.
"This doesn't prove that there is an ocean down there under the surface of Europa, but it does demonstrate that it is a scientifically exciting place," said \JGalileo\j imaging team member Dr. Robert Sullivan, also of \JArizona\j State University.
The images also reveal a remarkable diversity in the geological age of various regions of Europa's surface. Some areas appear relatively young, with smooth, crater-free terrain, while others contain large craters and numerous pits, suggesting that they are much older.
The icy crust bears the signs of having been disrupted by the motion of tectonic plates. "There appear to be signs of different styles of tectonism," said Greeley. "In many areas we see that the crust was pulled apart in a spreading similar to the processes on the sea floor on Earth. This is different from the tectonic processes at work on, say, Jupiter's moon Ganymede. This suggests that Europa's interior may be different from Ganymede's."
Galileo scientists will have a better chance to understand Europa's interior when the \Jspacecraft\j gathers gravity data on another flyby next November. The gravity field is measured by tracking how the frequency of \JGalileo\j's radio signal changes as it flies past the moon. This was not possible during the recent flyby because radio conditions were degraded as Jupiter passed behind the Sun from Earth's point of view.
Europa is crisscrossed by an amazingly complex network of ridges, according to Sullivan. "Ridges are visible at all resolutions," he explained. "Closely paired ridges are most common. With higher resolution, ridges seen previously as singular features are revealed to be double."
Some of the ridges may have formed by tension in the icy crust: as two plates pull apart slightly, warmer material from below might push up and freeze to form a ridge. Other ridges may have been formed by compression: as two plates push together, the material where they meet might crumple to form the ridge.
In addition to ice flows and tectonics, Greeley and Sullivan noted that some areas on Europa seem to have been modified by unknown processes that scientists are still debating. Greeley said that some areas, for example, seem to have been modified by "sublimation erosion" -- the \Jevaporation\j of water and other volatiles such as ammonia and \Jmethane\j into the vacuum of space. "Something is destroying the topography," said Greeley, "and this sublimation erosion is a good candidate for what is at work."
During last month's encounter, \JGalileo\j flew more than 200 times closer to Europa than the Voyager 2 \Jspacecraft\j did in 1979. After a swing past Jupiter next week in what mission engineers call a "phasing orbit," \JGalileo\j's next targeted flyby will take it again past Europa as it passes within 364 miles (587 kilometers) on Feb. 20.
The Jet Propulsion Laboratory, Pasadena, CA, manages the \JGalileo\j mission for NASA's Office of Space Science, Washington, DC.
#
"Monkey Dies After Completing 14-Day Bion Mission (9 January, 1997)",559,0,0,0
A rhesus monkey, one of two just returned to Earth after the Russian Bion 11 flight, died Jan. 8, 1997, post-operatively after all post-mission tests were completed at the Institute for Biomedical Problems in Moscow. The cause of the animal's death is unknown at this time. The death will be investigated separately by both the Russian Space Agency and NASA.
Ronald Merrell, M.D., Chairman, Department of Surgery, Yale University, New Haven, CT, will determine a process for an independent investigation of the incident. Merrell is the Chair of the NASA Bion Task Force.
The Bion program is a cooperative space venture among the U.S., Russian and French space agencies for the conduct of biomedical research using Russian-owned rhesus monkeys. The 14-day Bion 11 mission, carrying two rhesus monkeys, began on Dec. 24, 1996, with its launch from \JRussia\j's Plesetsk launch site. The mission landed in \JKazakhstan\j on Jan. 7. The monkeys "were alert, active and knew the people who were there to greet them," according to Dr. Joseph Bielitzki, NASA's Chief Veterinary Officer, who observed the landing.
The monkeys were then transported to the Institute for Biomedical Problems where postflight testing was conducted. The data collected are still being analyzed and will help U.S., French and Russian investigators understand how space flight affects the musculoskeletal system, as well as animal behavior and \Jphysiology\j. Following the recovery period, the remaining monkey will be retired to the Institute of Medical Primatology (Russian Primate Center) at Sochi/Adler.
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"Clipper Graham Incident Report Released (7 January, 1997)",560,0,0,0
An unconnected hose led to the destruction of the Clipper Graham technology demonstrator last summer. The Clipper Graham (DC-XA) Incident Investigation Board has released its final report concerning the July 31 post-landing tip-over and fire which destroyed the 43-foot vertical takeoff and landing vehicle at White Sands Missile Range, NM.
The Board, Chaired by former \JAstronaut\j Vance Brand, concluded "The primary cause of the vehicle mishap was that the brake line on the \Jhelium\j pneumatic system for landing gear 2 was not connected. This unconnected brake line prevented the brake mechanism from being pressurized to release the brake and resulted in landing gear 2 not extending. The vehicle became unstable upon landing, toppled onto its side, exploded and burned."
The July 31 flight demonstration was the fourth in last summer's series of tests with Clipper Graham by NASA and its industry partner, McDonnell Douglas. The vehicle flew the planned flight profile successfully and uneventfully until landing. With landing gear 2 failing to deploy -- there were four landing gears on Clipper Graham -- the vehicle tipped over and, according to the report, "The Clipper Graham DC-XA vehicle was totally destroyed by ground impact and ensuing explosions and fires."
Contributing causes of the mishap were identified as follows:
-Design of the system for gear stowage required McDonnell Douglas technicians to break the integrity of the \Jhelium\j brake line after integrity had been already verified. No other check was conducted to re-verify the integrity of the system after disconnection and reconnection of the line was completed;
-Landing gear stowage was never identified as a critical process. No special steps were taken to ensure the readiness of this system for flight;
-During the gear stowage process, there was no record of checking off steps or evidence of cross-checking;
-Distraction or interruption of the mechanical technician during gear stowage operations may have contributed to the non- connection of the brake line.
Gary E. Payton, NASA's Director of Space Transportation, commended the Board's thorough review of the incident. "Even though the Clipper Graham Program itself is now behind us, the technology demonstrations from last summer's four flights were outstanding. 'Lessons learned' from the Board's report, and the observations and recommendations made will play an important role in the Agency's continuing Reusable Launch Vehicle activities. In the X-33 and X-34 programs, for example, cost reduction and efficient reusability will continue to be our major objectives, along with safety and reliability that the proper mix of automation and human control can deliver."
The five member Board consisted of Chairman Vance Brand, Dryden Flight Research Center, CA; George Hopson, Marshall Space Flight Center, AL; Charles E. Harris, Langley Research Center, VA; Lt. Col. David Sharp, USAF Safety Center, NM; and Warren Wiley, Kennedy Space Center, FL.
The continuing cooperative effort in space exploration between the United States and \JRussia\j will be the focus of NASA's first Shuttle mission of 1997 with the launch of Space Shuttle Atlantis on Mission STS-81.
This is the fifth of nine planned missions to Mir and the second one involving an exchange of U.S. astronauts. \JAstronaut\j John Blaha, who has been on Mir since September 19, 1996, will be replaced by \Jastronaut\j Jerry Linenger. Linenger will spend more than four months on Mir. He will return to Earth on Space Shuttle Mission STS-84, scheduled for launch in May 1997.
Atlantis will again be carrying the SPACEHAB module in the payload bay of the orbiter. The double module configuration will house experiments to be performed by Atlantis' crew along with logistics equipment to be transferred to Mir.
The STS-81 crew will be commanded by Michael A. Baker who will be making his fourth Shuttle flight. The pilot, Brent W. Jett, Jr., will be making his second flight. There are four mission specialists assigned to this flight. Peter J.K. "Jeff" Wisoff, serving as Mission Specialist-1, is making his third flight. Mission Specialist-2 John M. Grunsfeld is making his second space flight. Marsha S. Ivins serving as Mission Specialist-3 is making her fourth space flight. Jerry M. Linenger will be Mission Specialist-4 for launch through docking with Mir. Shortly after docking, Linenger and Blaha will conduct their handover with Linenger becoming a member of the Mir crew and Blaha becoming Mission Specialist-4 through the end of the flight.
Atlantis is targeted for an early morning launch on or about January 12, 1997 from NASA's Kennedy Space Center Launch Complex 39-B. The current launch time of 4:27 a.m. EST may vary by a few minutes based on calculations of Mir's precise location in space at the time of liftoff due to Shuttle rendezvous phasing requirements. The STS-81 mission is scheduled to last 10 days, 3 hours, 30 minutes. An on-time launch on January 12 and nominal mission duration would have Atlantis landing back at Kennedy Space Center on January 22 at about 8 a.m. EST.
Atlantis' rendezvous and docking with the Mir actually begin with the precisely timed launch setting the orbiter on a course for rendezvous with the orbiting Russian facility. Over the next two to three days, periodic firings of Atlantis' small thruster engines will gradually bring the Shuttle within closer proximity to Mir.
The STS-81 mission is part of the NASA/Mir program which consists of nine Shuttle-Mir dockings and seven long duration flights of U.S. astronauts aboard the Russian space station. The U.S. astronauts will launch and land on a Shuttle and serve as Mir crew members while the Mir cosmonauts use their traditional Soyuz vehicle for launch and landing. This series of missions will expand U.S. research on Mir by providing resupply materials for experiments to be performed aboard the station as well as returning experiment samples and data to Earth.
The current Mir 22 mission began when cosmonauts Valeri Korzun and Aleksandr Kaleri were launched on August 17, 1996, in a Soyuz vehicle and docked with the Mir two days later. John Blaha joined the Mir 22 crew with the September 19, 1996, docking of STS-79. Blaha will complete his stay on Mir and return with the STS-81 crew. Linenger will work with the Mir 22 crew until the arrival of Mir 23 cosmonauts Vasili Tsibliev, Aleksandr Lazutkin and German researcher Reinhold Ewald in early February 1997. After the Mir 22 crew and Ewald return to Earth in a Soyuz, Linenger will complete his tour with the Mir 23 crew. Linenger will be replaced by NASA \JAstronaut\j Mike Foale when Atlantis again docks with Mir in May.
The STS-81 mission also will include several experiments in the fields of advanced technology, Earth sciences, fundamental \Jbiology\j, human life sciences, \Jmicrogravity\j, and space sciences. Data also will supply insight for the planning and development of the International Space Station, Earth-based sciences of human and biological processes, and the advancement of commercial technology.
STS-81 will involve the transfer of 5,975 pounds of logistics to and from the Mir, the largest transfer of items to date. During the docked phase, 1,400 pounds of water, 1,137.7 pounds of U.S. science equipment, 2,206.1 pounds of Russian logistics along with 268.2 pounds of miscellaneous material will be transferred to Mir. Returning to Earth aboard Atlantis will be 1,256.6 pounds of U.S. science material, 891.8 pounds of Russian logistics and 214.6 pounds of miscellaneous material.
STS-81 will be the 18th flight of Atlantis and the 81st mission flown since the start of the Space Shuttle program in April 1981.
#
"Month in Space - July",562,0,0,0
July 1
In 1770, \JComet\j Lexell passed a mere 2.3 million km (0.0151 AU) from the Earth. This was the closest ever Earth approach of a \Jcomet\j in recorded history.
In 1847, the 6th asteroid, Hebe, was discovered.
In 1965, Tiros X, the first US Weather Bureau-funded satellite, was launched into orbit to obtain maximum coverage of the 1965 hurricane and typhoon season.
In 1976, the Smithsonian's National Air & Space Museum in Washington DC is opened.
In 1991, the \JGalileo\j \Jspacecraft\j entered the asteroid belt. It would make the first flyby of an asteroid, Gaspra, in October 1991.
July 2
In 1985, the European Space Agency launched Giotto to encounter \JComet\j Halley in 1986.
In 1990, the Giotto \Jspacecraft\j made an Earth flyby for a gravity assist for an encounter with \JComet\j Grigg-Skjellerup in 1992.
July 3
In 1992, SAMPEX was launched. Carrying four cosmic ray monitoring instruments, SAMPEX studied the solar energetic particles, anomoulas cosmic rays, galactic cosmic rays and magnetospheric electrons.
July 4
In the year 1054, the brightest known \Jsupernova\j, the Crab Nebula, starts shining for 23 days and is observed by Chinese and Japanese astronomers.
In 1819, William Herschel made his last telescopic observation, of a \Jcomet\j. Herschel was the discoverer of Uranus in 1781.
July 6
In 1687, Edmond Halley published Isaac Newton's "Principia Mathematica".
Lysithea, the 10th Moon of Jupiter, was discovered in 1938 by S. Nicholson.
July 7
The USSR launched \JPhobos\j 1 in 1988 towards Mars. A wrong command was sent to the \Jspacecraft\j on August 29 causing the \Jspacecraft\j to fail. \JPhobos\j 1 is now orbiting the Sun.
In 1914, Robert Goddard received a patent for a two-stage rocket.
July 8
Radio \Jastronomy\j was born in 1933 when Karl Jansky determined that the Milky Way is a source of radio noise.
July 9
In 1979, Voyager 2 flew past Jupiter and collected data on the Jovian system and its moon. Voyager 2 used the Jupiter flyby as a slingshot to Saturn, Uranus and Neptune.
July 10
In 1990, \JJapan\j's Hiten \Jspacecraft\j made its 2nd flyby of the Moon.
In 1992, the Giotto \Jspacecraft\j made a successful flyby of \JComet\j Grigg-Skjellerup.
In 1962, the US launched Telstar 1, the first active real-time communications satellite.
July 11
Skylab burned up in Earth's atmosphere in 1979 and pieces of the space station landed in western \JAustralia\j.
July 12
In 1977, the High Engery Astronomical Observatory (HEAO-1) was launched.
In 1988, the USSR launched \JPhobos\j 2 to Mars. The \Jspacecraft\j went into Mars orbit on January 29, 1989 and transmitted back some data and photos of Mars and its moon \JPhobos\j. However, contact was lost with the \Jspacecraft\j on March 27, 1989 before it could fullfill its primary mission.
In 1995, the \JGalileo\j probe separated from the \JGalileo\j orbiter in preparation for arrival at Jupiter. Both \Jspacecraft\j would encounter Jupiter in December 1995.
July 13
In 1917, Langley Research Center was founded in Hampton, Virginia.
In 1969, the USSR launched Luna 15 just two days before Apollo 11. Luna 15 was to land on the Moon and return soil samples to Earth before Apollo 11 returned, but instead crashed on the Moon in the Sea of Crises on July 21 after a 4 minute descent.
In 1972, the largest lunar quake was recorded.
July 14
In 1965, Mariner 4 made the first ever Mars flyby. It returned the first close-up images of that planet and confirmed the existence of surface craters.
In 1967, Surveyor 4 was launched to the Moon. Contact was lost 2.5 minutes before landing in Sinus Medii, and the \Jspacecraft\j presumably crashed.
July 16-22
In 1994, 21 fragments from \JComet\j Shoemaker-Levy 9 impacted on Jupiter in a spectacular fashion.
July 16
Giuseppe Piazzi was born in 1746. Piazzi would discover the first asteroid, Ceres, in 1801.
The USSR launched the first Proton rocket in 1965. This heavy lift rocket launches 50,000 pounds to orbit and is still in use today.
In 1969, Apollo 11 is launched to the Moon on a Saturn V rocket. Carrying, Neil Armstrong, Buzz Aldrin and Michael Collins, this would be the first manned landing on the Moon.
In 1990, \JPakistan\j's first satellite, Badr-A, was launched on the maiden flight of China's Long March 2E rocket.
July 17
In 1975, the USSR's Soyuz 19 linked with US's Apollo 18 for the first ever joint manned flight effort between the two countries. Known as Apollo-Soyuz, the two crews shared meals, conducted experiments together and held a joint news conference from space.
The first photograph of a star was taken by Harvard Observatory of Vega in 1850.
July 18
In 1965, Zond 3 was launched by the USSR for a Moon flyby. It returned 25 images of the far side of the Moon.
In 1966, Gemini 10 was launched on a Titan rocket into Earth orbit. John Young and Michael Collins were the astronauts on this mission, and they performed the first docked vehicle maneuvers.
In 1980, India becomes the 7th nation to launch a satellite into Earth orbit with the launch of Rohini 1.
During an EVA outside Salyut 7 in 1984, Svetlana Savitskaya becomes the first woman to walk in space.
July 19
In 1967, Explorer 35 was launched to the Moon, going into lunar orbit 3 days later. Explorer 35's mission was to study the solar wind and interplanetary fields at lunar distances. Results indicated no shock front precedes the Moon, and no evidence of a lunar magnetic field, radiation belts or \Jionosphere\j.
July 20
In 1964, SERT 1 was launched in a suborbital flight to test an ion engine performance in space. The flight confirmed that high prevalence ion beams could be neutralized in space.
In 1969, Apollo 11 lands on the Moon. Neil Armstrong becomes the first man on the Moon. Buzz Aldrin becomes the second man. They spend 21.5 hours on the lunar surface, including 2.5 hours outside their lunar excursion module (LEM).
In 1976, Viking 1 becomes the first \Jspacecraft\j to successfully land on Mars. Landing in the Plain of Chryse, Viking 1 returned the first images from the surface of the red planet showing an orange-red plain strewn with rocks and sand dunes.
July 21
Sinope, Jupiter's 9th moon, was discovered in 1914 by S. Nicholson.
In 1961, Mercury 4 was launched carrying Gus Grissom on the second manned suborbital flight by the US. Grissom had to be rescued from the ocean when the capsule sank after splashdown.
The USSR launched Mars 4 in 1973. Originally intended to be a Mars orbiter, the \Jspacecraft\j flew by Mars on February 10, 1974 when its main engine failed to fire. Though failing to go into orbit, the \Jspacecraft\j did return some images of Mars and some radio \Joccultation\j data.
July 22
Astronomer Friedrich Bessel was born in 1784.
In 1962, Mariner 1 was launched in the first Venus mission attempt by the United States. A launch vehicle problem caused the \Jspacecraft\j to veer off course, and forced the Range Safety Officer to destroy the \Jspacecraft\j 293 seconds into the launch.
In 1972, Venera 8 arrived at Venus. The descent module parachuted to the surface and returned data on temperature, pressure, light levels and descent rates. Venera 8 lasted 50 minutes on the surface of Venus.
In 1972, the first Earth resources technology satellite, Landsat 1, was launched.
July 23
In 1980, the USSR launched Soyuz 37 carrying cosmonauts Viktor Gorbatko and Pham Tuan. Tuan was from Vietnam and become the first third-world \Jcosmonaut\j.
In 1984, a possible ring (later confirmed) was discovered around Neptune.
July 24
In 1950, a captured German \JV-2\j rocket became the first rocket to be launched from Cape Canaveral, \JFlorida\j.
In 1992, Geotail was launched to study the structure and dynamics of the tail region of the Earth's magnetosphere.
July 25
In 1973, the USSR launched Mars 5 in 1973. The \Jspacecraft\j went into Mars orbit on February 12, 1974, and was one of the few successful Mars missions by the USSR. Mars 5 returned images of a small portion of the southern hemisphere of Mars.
In 1990, the Combined Release and Radiation Effects Satellite (CRRES) was launched. A joint NASA/DOD program, CRRES released chemicals in Earth orbit to study the Earth's magnetic fields and the plasmas that travel through them.
July 26
In 1659, Christaan Huygens published his book, Systema Saturnium, in which he explained why Saturn's rings appear to disappear every 14 to 15 years.
In 1958, Explorer 4 was launched into Earth orbit.
In 1971, Apollo 15 was launched, the fourth manned Moon landing. Astronauts David Scott, Alfred Worden and James Irvin arrived in lunar orbit on July 30. 170 pounds of rocks were collected and a lunar rover was driven across the Moon's surface for the first time.
In 1963, Syncom 2 was launched, the first operational geosynchronous satellite.
In 1995, new moons around Saturn discovered by the Hubble Space \JTelescope\j was announced.
July 28
The first photograph of a total eclipse was taken in 1851 in Konigsbert, East \JPrussia\j.
In 1964, Ranger 7 was launched to the Moon. Impacting on the Moon three days later, Ranger 7 returned 4,316 images of the Moon including the first close-up images of the Moon's surface.
Astronauts Alan Bean, Jack Lousma and Owen Garriot make the second visit to Skylab in 1973.
In 1992, ALEXIS was launched.
July 29
NASA was created when President Eisenhower signed into law the NASA Act of 1958.
In 1960, a Mercury capsule (MA-1) was launched in a suborbital test. The Atlas launch vehicle blew up 65 seconds after launch.
In 1993, the first test flight of the Delta-Clipper X was successfully performed.
July 30
In 1610, \JGalileo\j observes Saturn's rings for the first time through a \Jtelescope\j. The true nature of Saturn's rings was not known then, and \JGalileo\j described his observations as a possible triple planet.
Carme, Jupiter's 11th moon, was discovered in 1938 by S. Nicholson.
In 1965, Pegasus 3 was launched. Pegasus 3 determined that the flux of large particles in space was less than expected.
July 31
In 1969, Mariner 6 flew by Mars and provide high-resolution images of the Martian surface, concentrating on the equatorial region of the red planet.
During the Apollo 15 mission in 1971, David Scott becomes the first person to drive a rover on the Moon.
In 1995, the Ulysses reached its maximum northern latitude (80.22 degrees) above the Sun.
#
"Month in Space - June",563,0,0,0
June 1
The ROSAT satellite is launched in 1990.
The USSR launched Soyuz 9 in 1970 with cosmonauts Andrian Nikolayev and Vitali Sevastyanov. Their 18-day stay in space broke a five-year-old U.S. Gemini 7 record of 14 days.
June 2
In 1966, Surveyor 1 becomes NASA's first \Jspacecraft\j to soft land on the Moon.
In 1983, the USSR launched Venera 15, a radar mapping orbiter to Venus.
June 3
In 1965, Edward White makes the first American spacewalk during the Gemini 4 mission.
The 200-inch Hale \JTelescope\j becomes operative in 1948.
In 1966, Gemini 9 is launched carrying astronauts Thomas Stafford and Eugene Cernan.
June 6
Georgi Dobrovolksy, Vladislav Volkov, Viktor Patsayev are launched in Soyuz 11 in 1971 and docked with the Salyut space station. They stayed 23 days in space, but died on the flight back to Earth. See June 30 for more details.
In 1983, the USSR launched Venera 16, a radar mapping orbiter to Venus.
June 8
Astronomer Giovanni Cassini was born in 1625.
NASA's X-15 was flown for the first time in 1959. The X-15 was carried to an altitude of 38,000 feet by a B-52 bomber and dropped. Pilot Scott Crossfield made an unpowered five minute gliding descent.
In 1965, the USSR launched Luna 6 for a Moon flyby. It passed within 100,000 miles of the Moon.
In 1975, the USSR launched Venera 9, a Venus orbiter and lander.
June 9
In 1931, Robert Goddard patents the first rocket-powered \Jaircraft\j design.
June 10
In 1973, Explorer 49 was launched into orbit around the Moon. The \Jspacecraft\j used the Moon to block radio interference from Earth so that it could perform precise radio astronomical measurements of the Sun, Jupiter and deep space. Transmission were terminated in 1977, and Explorer 49 still holds the record for the longest a \Jspacecraft\j has remained in lunar orbit - 4 years.
June 11
In 1985, USSR's Vega 1 drops off a lander and \Jballoon\j at Venus. Vega 1 then went on to encounter \JComet\j Halley on March 6, 1986.
June 12
In 1967, Venera 4 is launched by the USSR to Venus. The \Jspacecraft\j parachuted to the surface of Venus in October the same year.
June 13
Pioneer 10 leaves the solar system by crossing the orbit of Neptune in 1983.
The Jet Propulsion Lab's official birthday is today. This is the date of the oldest document found dated in 1944 that makes reference to the "Jet Propulsion Laboratory".
June 14
In 1967, Mariner 5 is launched for a Venus flyby in October. It flies by Venus at 4,100 km in October.
In 1975, Venera 10 lands on Venus and transmits images from the surface.
Like its twin a couple of days before, Vega 2 drops off its lander and ballon at Venus. It also flew on to \JComet\j Halley for an encounter on March 9, 1986.
June 16
In 1963, USSR's Valentina Tereshkova is launched in Vostok 6 and becomes the first woman in space.
June 17
Space shuttle Columbia is launched in 1985 on flight STS-51G and deploys three communications satellites. Prince Sultan Salman Abdel Aziz Al-Saud of Saudi Aribia becomes the first Arab in space as well as the first royalty in space.
June 18
Sally Ride becomes the first US woman in space in 1983 on the space shuttle Challenger on STS-7. Two satellites were deployed.
June 22
The Royal Greenwich Observatory was founded in 1675.
In 1976, the USSR launched the Salyut 5 space station.
In 1978, Pluto's moon Charon is discovered by Jim Christy.
Astronomer Edwin Hubble was born in 1903.
June 24
In 1974, the USSR launched the Salyut 3 space station.
June 26
Astronomer Charles Messier was born in 1730.
In 1978, SEASAT 1 is launched to monitor ocean currents, wave heights and surface temperatures.
June 27
The first rockects were launched in 1945 from a station on Wallops Island off Virginia's eastern shore. This station would eventually become Wallops Flight Facility.
In 1982, space shuttle Columbia was launched on STS-4. This was the shuttle's last test flight.
In 1994, the Ulysses \Jspacecraft\j begins its first solar passage as it passes underneath the Sun's southern pole.
June 29
Astronomer George Hale was born in 1868.
June 30
A small \Jcomet\j or stony asteroid exploded over a remote area near Tunguska in central \JSiberia\j in the USSR in 1908. Known as the Tunguska Event, the explosion flattened trees up to 40 km away, and shattered windows and forced doors off their hinges in a trading station 65 km from the explosion.
Three cosmonauts died on their way back to Earth in Soyuz 11. A USSR government commission found that, during re-entry, a faulty valve had allowed pressure to escape from the \Jspacecraft\j, killing them 30 minutes before landing. Not wearing space suits, the cosmonauts were found dead in the Soyuz 11 capsule when it was opened after landing.
March 13, 1998: Two scientists exploring a microworld locked in ancient ice have found a wide range of lifeforms from fungi, algae, and bacteria to a few diatoms - and a few items with strange shapes.
"We've found some really bizarre things - things that we've never seen before," said Richard Hoover of NASA's Marshall Space Flight Center. Hoover and Dr. S.S. Abyzov of the Russian Academy of Sciences have been examining deep ice core samples from the Vostok Station about 1,000 km (620 miles) from the South Pole.
The objects have fanciful names - like Mickey Mouse and Klingon (see picture) - based on passing resemblances. Hoover expects that most will fall into known categories of microorganisms as he and Abyzov study the images.
We're exploring a new world," Hoover said. "Until we get a lot more experience, we're going to see brand new things all the time."
The ice harboring these finds is as old as 400,000 years, depending on the depth. Russian scientists at the St. Petersburg Mining Institute in St. Petersburg, Russia, developed the technology for drilling ice cores without contaminating the samples. Since 1974, they have worked at Vostok Station, extracting cores from ever greater depths.
In 1996, the Russian Academy of Sciences announced that a large lake of liquid water lies beneath the 3 km-deep glacier at Vostok.
Meanwhile, the ice samples from above the lake's surface (which has not been breached) are stirring interest in the scientific community. In the 1970s, Abyzov discovered - and in some cases revived - microorganisms in ice that conventional wisdom had said was sterile.
Now the discovery of ice and slush on Europa and mounting evidence of water on Mars and on our own Moon are leading scientists to rethink the possibilities of life elsewhere in the universe.
"These are very important questions for future cosmic research on places like Europa, comets, the Martian ice caps," Abyzov said of the mysteries in the Antarctic ice.
Hoover and Abyzov are using Marshall's Environmental Scanning Electron Microscope (ESEM) to examine material found in the ice. Their work is a collaborative effort between NASA/Marshall and the Institute of Microbiology of the Russian Academy of Sciences working in collaboration with the St. Petersburg Mining Institute and the Institute of Arctic and Antarctic Research in St. Petersburg, Russia.
"This is very useful," Abyzov said, "because we do not have this equipment in our laboratory. We have scanning electron microscopes, but without the additional equipment you have."
As might be expected, they have found a lot of atmospheric dust and debris, and possibly some cosmic dust.
"There are some dust particles with unusual spectra," Hoover said. "Which may be cosmic dust particles." The ESEM allows the operator to designate a point on a specimen and then scan with X-rays to determine what elements are present. The ratios found in some of the dust particles do not match ratios expected in terrestrial dust grains.
"Mickey Mouse" and other colonies of small microbes appear to be out of the ordinary. These are fluffy white objects, about 1 micron wide and resembling cotton balls.
"Here's the shocker," said Hoover, pointing at the ESEM monitor, "these small coccoid bodies are covered with all this incredible fibrous structure." The filaments appear to be about 30 to 40 nanometers wide (that's about 1/10th a wavelength of visible light).
"It's difficult for me to say what it is," Abyzov said, "but I tend to agree that this is biological."
"There are all sorts of microorganisms in the ice. Some are readily recognizable as cyanobacteria, bacteria, fungi, spores, pollen grains, and diatoms, but some are not recognizable as anything we've ever seen before," Hoover said. Many will turn out to be known. It's just that they look different under the ESEM, which provides details that are not available through other microscopes.
Familiar items include bits of sponge and feather, and diatom fragments (see picture), Hoover's other area of personal interest and expertise (he works at Marshall as an X-ray astronomer).
They have also found a number of large cyanobacteria with nanobacteria attached.
"What is clearly going on is that when microorganisms freeze, they shut down and go into this anabiotic state," Hoover explained. Anabiotic means alive but inactive, like suspended animation. Russian scientists have been able to revive and culture bacteria, yeast, fungi, and other microbes found in ice cores.
"One of the things that was really exciting was that many of the cyanobacteria from 1,243 meters down had lots of antimony," Hoover said. The X-ray spectrum showed carbon, oxygen, zinc, silicon, aluminum, and potassium - all chemicals common to life. But it also showed an abnormal amount of antimony, a toxic heavy metal.
"It was not just one of these that had it," Hoover said, "but microorganism after microorganism."
Gregory Jerman, the ESEM operator, noted that the metal content has varied with depth. At some levels the microorganisms show large quantities of antimony, while in others zinc rings the bell.
With more than 150 ESEM images and almost as many spectra recorded, Hoover and Abyzov next go to the Jet Propulsion Laboratory in Pasadena, California. There, Dr. Ken Nealson will try to extract genetic material from the microorganisms.
As Hoover talked, another new image appeared on screen. "It's pretty big," he said of an object that looked like a porpoise and probably is a protozoan. "The work of identifying and classifying everything in the ice will be long and challenging," Hoover said. He compared it to his own initiation in the world of diatoms where for years everything looked new until he became familiar with it. Then he could quickly recognize the rare unusual specimen.
"It's necessary to know what to look for and the kinds of things you can see," he said.
Like the Klingon's forehead, a wrinkled object resembling a character from Star Trek, or the porpoise. For now some of them just have nicknames, until Hoover, Abyzov, and their colleagues analyze their exciting images and obtain more definitive identifications of these microscopic beasts of the frozen underworld.
#
"Cassini Mission Status Report (8 January, 1998)",565,0,0,0
Having clocked in more than 200 million kilometers (124 million miles) since launch last Oct. 15, the Cassini spacecraft remains in excellent health as it progresses on its long flight path to Saturn.
The Cassini flight team has nearly completed all scheduled spacecraft configuration activities for the mission's early phase. Onboard software stored in the command data subsystem has been directing spacecraft activities as scheduled. The spacecraft's attitude in space is being maintained using the spacecraft's small hydrazine thrusters. For the first 14 months of flight, Cassini will fly so that its 4-meter-diameter (13-foot) high-gain antenna always faces the Sun, shading most of the spacecraft from the more intense sunlight that characterizes the inner solar system.
Throughout this period, spacecraft communications are conducted through one of Cassini's two low-gain antennas; the antenna selected depends on the relative geometry of the Sun, Earth and spacecraft.
In December and through January and beyond, periodic instrument maintenance and health checks, along with other routine spacecraft "housekeeping" activities, dominate Cassini's schedule. Cassini spacecraft operations and science team personnel continue to develop software command "loads" that will direct spacecraft activities and scientific instrument observations later in the mission.
Today Cassini is traveling at a speed of more than 109,200 kilometers per hour (67,800 miles per hour) and is more than 22,450,000 kilometers (13,950,000 miles) from Earth. It is 27,500,000 kilometers (17,088,000 miles) from Venus, where the spacecraft will perform its first gravity-assist maneuver on April 26. As the spacecraft swings around Venus, the planet's gravity will boost Cassini's speed for the next phase of the journey to Saturn.
#
"Cassini Mission Status Report (3 February, 1998)",566,0,0,0
The Cassini spacecraft remains in excellent health as it travels on its Saturn-bound trajectory at a speed of approximately 120,000 kilometers per hour (about 75,000 miles per hour). The spacecraft has traveled more than 271 million kilometers (about 168 million miles) since launch last October 15.
Cassini continues to fly with its 4-meter (13-foot) diameter high-gain antenna pointed toward the Sun so that the rest of the spacecraft is shaded. It will maintain in this attitude, except during planned trajectory adjustments, for the first 14 months of flight as it travels through the inner solar system.
Radio communications with the spacecraft are currently through low-gain antenna 2, one of the spacecraft's two low-gain antennas. Low-gain antenna 2 is located at the end of the spacecraft opposite the high-gain antenna. The low-gain antenna that is selected for a given period depends on the relative geometry of the Sun, Earth and the spacecraft. The telemetry data rate from Cassini is currently 40 bits per second.
For about the next month, there will be an increase in the amount of telecommunications time allotted to Cassini by NASA's Deep Space Network to meet the data needs of spacecraft navigators as they prepare for two long-scheduled trajectory refinements. In late January, Deep Space Network antennas were trained on Cassini about four times a week. In February, those episodes of tracking will be approximately doubled. The adjustments will be made in preparation for Cassini's flyby of the planet Venus on April 26. The additional data gathered through the extra telecommunications time are used to refine knowledge of the spacecraft's location, which will aid navigators in setting precise parameters for the trajectory adjustments, such as the duration of thruster firings.
#
"Cassini Mission Status Report (3 March, 1998)",567,0,0,0
The Cassini spacecraft successfully performed the second scheduled trajectory adjustment of its mission last week, fine- tuning its flight path in preparation for its flyby of Venus on April 26. The trajectory adjustment needed was so minor that the maneuver was performed using Cassini's small hydrazine thrusters instead of the spacecraft's large main engine. Engineering data recorded during the thruster firing confirmed that the maneuver went as planned, with all spacecraft and ground components performing perfectly. A final trajectory adjustment prior to the Venus flyby is scheduled in early April.
Cassini remains in excellent health, flying at a speed relative to the Sun of approximately 137,000 kilometers per hour (about 85,000 miles per hour). It is slowly gaining speed as it feels the tug of gravity from Venus. The spacecraft will gain a significant boost in speed when it swings around Venus next month. Cassini has traveled approximately 362 million kilometers (about 224 million miles) since launch on October 15, 1997.
#
"Cassini Mission Status (26 March, 1998)",568,0,0,0
Cassini's fault protection system worked as planned Tuesday after the spacecraft detected a very small orientation difference between its two stellar reference units, according to preliminary data received from the spacecraft yesterday. The event has had no impact on the mission and all planned activity for Cassini's Venus flyby next month remains on schedule.
The second stellar reference unit, part of Cassini's attitude and articulation control subsystem, was in the process of taking over while the first stellar reference unit was to undergo a routine maintenance "bake-out" heating to remove normal post-launch contaminants from its aperture.
This was the first use of the second stellar reference unit, and as it began operating, Cassini's attitude control software found a very small difference in the orientation of the units. Sensing this difference, the computer that controls the attitude control subsystem executed pre-programmed commands that brought the spacecraft into a low-activity state to await instruction from ground controllers. The spacecraft executed this response exactly as designed and the spacecraft remains healthy, project officials said. Cassini has remained in contact with ground controllers throughout, and engineering data on the event and the spacecraft's overall operations were received yesterday.
Preliminary analysis indicates that the discrepancy was within specifications, but that control limits were set too tightly, triggering the preprogrammed commands that set the spacecraft in its low-activity state. This problem is expected to be solved easily with an adaptation made to the spacecraft's attitude and articulation control software.
By early this afternoon, ground controllers will have sent commands to return the spacecraft to normal operations. Cassini remains on course for its April 26 flyby of Venus, with the last fine-tuning of the flight path, if it is needed, scheduled for early April. During its Venus flyby, Cassini's radio and plasma wave science instrument will take advantage of the opportunity to search for lightning in Venus's atmosphere.
Today, the spacecraft is approximately 17 million kilometers (about 10.5 million miles) from Venus and is traveling at a speed of about 143,000 kilometers (about 88,800 miles per hour). It has traveled about 440 million kilometers (about 273 million miles) on its Saturn-bound trajectory since launch on October 15, 1997.
#
"Cassini Mission Status (3 April, 1998)",569,0,0,0
The Cassini spacecraft remains on track for its flyby of Venus on April 26, which due to the effects of Venus' gravity, will give the spacecraft a 26,280 kilometer-per-hour boost (16,330 mile-per-hour) in speed. Cassini's navigators have determined that the spacecraft is already so accurately targeted for its 284-kilometer altitude swingby of Venus that a scheduled fine-tuning maneuver is unnecessary and has been cancelled.
Cassini is feeling the Sun's gravitational tug since the spacecraft last week reached its perihelion, or closest point to the Sun, and is now flying in outbound direction. The spacecraft is traveling at a speed of approximately 143,000 kilometers per hour (about 89,000 miles per hour) relative to the Sun, and has traveled approximately 464 million kilometers (about 288 million miles) since launch on October 15, 1997.
Over the past week, Cassini began transmitting previously- recorded data from last week's engineering checkout of the European Space Agency's Huygens probe. This health-check of Huygens occurs every six months. The data are forwarded to the Huygens team in Europe for analysis. The remainder of the engineering checkout data from Huygens is scheduled to be transmitted from the spacecraft next week.
#
"Cassini Mission Status (29 April, 1998)",570,0,0,0
Cassini spacecraft navigators reported that Sunday's successful flyby of Venus was on time and on target.
"The accuracy achieved by our navigators is roughly equivalent to shooting a basketball from Los Angeles to London and making a swish shot," said Richard Spehalski, Cassini program manager at NASA's Jet Propulsion Laboratory.
As planned, Cassini flew 284 kilometers (176 miles) above the Venusian surface at 6:52:14 Pacific Daylight Time, (Earth- received time) on April 26. Because of Venus' distance (136 million kilometers or 85 million miles from Earth), it took Cassini's radio signals 7.5 minutes to travel to Earth.
The tug of Venus' gravity gave the Cassini spacecraft a boost in speed of about 7 kilometers per second (about 4 miles per second). This will help the spacecraft reach Saturn in July 2004. Leaving Venus, Cassini was moving at more than 141,000 kilometers per hour (87,000 miles per hour) relative to the Sun.
The Venus flyby is the latest of dozens of similar "gravity- assist" flybys of planets and moons performed by JPL-controlled spacecraft over the past three decades. Cassini will perform three more similar gravity-assist flybys: Venus again in June of 1999, Earth in August of 1999, and Jupiter in December 2000. All the flybys use the gravitational pull of the target planets to impart more speed to the spacecraft to help it reach Saturn. The Venusian flyby was the lowest-altitude gravity-assist planetary pass Cassini will make in its mission.
Science instruments on the spacecraft searched for lightning in Venus' atmosphere during the flyby, and the radar instrument onboard was activated to test a bounced signal off Venus' surface.
After it enters orbit around Saturn in 2004, Cassini will study the ringed planet, its moons and ring system for at least four years. It will also deliver a scientific probe called Huygens to parachute to the surface of Saturn's largest moon, Titan.
#
"Cassini Mission Status (5 January, 1998)",571,0,0,0
The Cassini spacecraft is presently traveling at a speed of approximately 147,000 kilometers/hour (~91,000 mph) relative to the sun and has traveled approximately 555 million kilometers (~344 million miles) since launch on October 15, 1997. Cassini's first planetary gravity assist, a technique used to increase spacecraft velocity, occurred early Sunday morning, April 26th. Cassini is now traveling approximately 11,000 kph (7,000 mph) faster than it was a week ago. The Venus-1 flyby was a tremendous success. Post-flyby Navigation tracking has indicated that the spacecraft is precisely on the desired trajectory. This information has allowed the Program to cancel the next trajectory correction maneuver, which had been planned for May 14, as it is no longer needed.
The most recent Spacecraft status is from the DSN tracking pass on Tuesday, 04/28, over Canberra. There have been two additional no-telemetry passes on Wednesday 04/29 and Thursday 04/30. The Cassini spacecraft is in an excellent state of health and is executing the C7 sequence nominally.
Inertial attitude control is being maintained using the spacecraft's hydrazine thrusters (RCS system). The spacecraft continues to fly in a High Gain Antenna-to-Sun attitude. It will maintain the HGA-to-Sun attitude, except for planned trajectory correction maneuvers, for the first 14 months of flight.
Communication with Earth during early cruise is via one of the spacecraft's two low-gain antennas; the antenna selected depends on the relative geometry of the Sun, Earth and the spacecraft. The downlink telemetry rate is presently 40 bps.
\BSpacecraft Activity Summary:\b
On Friday, 04/24, there were no changes to spacecraft configuration.
On Saturday, 04/25, the Solid State Recorder (SSR) record and playback pointers were reset, according to plan. This housekeeping activity, done approximately weekly, maximizes the amount of time that recorded engineering data is available for playback to the ground should an anomaly occur on the spacecraft.
On Sunday, 04/26, the Venus-1 flyby and associated activities took place, with Venus closest approach occurring at 6:44am PDT. The Radio & Plasma Wave Science (RPWS) instrument conducted a several-hour search for signals which could indicate the occurrence of lightning in the Venusian atmosphere. The Radar instrument conducted an engineering test near closest approach to attempt to acquire its first "bounce" from a target body - a closed-loop test very difficult to accomplish on the ground.
Finally, during the portion of flyby when the spacecraft flew behind Venus (as seen from the Earth), some of NASA's Deep Space Network equipment was employed to conduct an atmospheric occultation experiment to obtain data which can be used by the Radio Science team. Results of the RPWS and Radar activities are going to be played back from the Cassini Solid State Recorders (SSRs) this coming Saturday (5/2) and Sunday (5/3), respectively.
On Monday, 04/27, as part of the onboard sequence, Cassini executed a Memory Readout of Mass Properties in preparation for TCM 4. As it turns out, TCM 4 will not be needed and has since been cancelled.
On Tuesday, 04/28, a maintenance activity was performed on the SSR Flight Software Partitions. This activity repairs any SSR double bit errors (DBEs) which have occurred in the code-containing portions of the Flight Software partitions during the preceding period. The real-time command based portion of the activity, which clears telemetry flags and reads out the results of the maintenance activity, is scheduled for this Friday, 5/1.
On Wednesday, 04/29, there were no changes to spacecraft configuration.
On Thursday, 04/30, the Solid State Recorder (SSR) record and playback pointers were reset, according to plan.
#
"Cassini Mission Status (4 June, 1998)",572,0,0,0
Now more than 20 million kilometers (about 12.5 million miles) beyond Venus, Cassini is zooming along at more than 129,300 kilometers per hour (about 80,000 miles per hour), having received a big boost in speed from the effects of Venus' gravity when the spacecraft flew past that planet in late April.
The international Cassini and Huygens team this week bids a fond farewell to Cassini Program Manager Richard J. ("Spe") Spehalski, who retires from Cassini's helm on June 5. Since 1992, Spehalski has managed the successful development, launch and operation of Cassini. Robert T. Mitchell, currently Project Manager of the Galileo Europa Mission, assumes the post of Cassini Program Manager on June 8.
Earlier this week, NASA bestowed its highest honor on Spehalski, awarding him the agency's Distinguished Service Medal for his leadership of NASA's two historic flagship planetary missions - Cassini, and the Galileo Mission to Jupiter, which he managed through final development and launch and early operations from 1988 to 1990.
"The Galileo and Cassini missions are the last great flagship planetary flights of discovery of the 20th century, and their success has been critical to support for a vigorous program of solar system exploration," said JPL Director Dr. Edward C. Stone, commenting on Spehalski's retirement. "JPL, NASA, and the nation owe Spe and the teams he led a debt of gratitude for their accomplishments."
Cassini remains in excellent health. So far, the spacecraft has traversed more than 675 million kilometers of space since launch on Oct. 15, 1997. The mission's operations schedule over the past month was dominated by a variety of engineering tests and routine maintenance activities on Cassini and the European Space Agency's (ESA) Huygens Titan probe. Tests and engineering analysis for the Huygens probe are performed at ESA's Huygens Probe Operations Center (HPOC) in Darmstadt, Germany, via a telecommunications link to JPL. Last week, engineers conducted an in-flight calibration test that pointed the high-gain antenna 12 degrees from the Sun. Data from that test confirmed the excellent health of the Huygens probe radio receiver.
Cassini's international science team gathered for the 17th Cassini Project Science Group in Pasadena this week. The primary topics covered at the meeting were planned and proposed operation of the science instruments during the flight to Saturn, and plans for the approach to Saturn and the four-year mission in orbit at Saturn.
#
"Comet Hale-Bopp Holds Clues to Creation of Comet Ices",573,0,0,0
March 12, 1998
AMHERST, Mass. -- University of Massachusetts astronomer Matthew Senay is part of a research team reporting that Comet Hale-Bopp, which passed by the Earth last year, provides evidence that comet ices are created inside huge interstellar clouds of gas, dust, and ice. The findings, detailed in the current issue of Science, suggest that comet ices are not, as previously thought, created in the outer regions of the solar system, where comets themselves are formed. The study also offers new evidence which implies that the temperature of the interstellar cloud that formed our solar system could not have been colder than 30 degrees Kelvin; a temperature considerably higher than scientists previously believed. The research team includes astronomers from the University of Hawaii, the National Research Council of Canada, and the Observatoire de Paris-Meudon.
Astronomers studied the chemical composition of Hale-Bopp's nucleus during its journey past the Earth in April of 1997, and compared their findings with the chemistry of interstellar clouds, which are comprised of gas and ice-mantled dust. These clouds can become massive, fragmenting and collapsing into smaller clouds or "protostellar nebulae." Chemical reactions within the smaller clouds cause significant amounts of deuterium, a form of hydrogen sometimes called "heavy hydrogen," to latch onto the ice-encrusted dust particles. An individual protostellar nebula continues collapsing until the beginnings of a star appear at the center, with planets forming in the flattened disk of gas, dust, and ice -- called the "protoplanetary nebula" -- around it.
Using a radio telescope located in Hawaii, astronomers determined the ratio of deuterium to hydrogen in the hydrogen cyanide molecules ejected by Hale-Bopp. These molecules cannot be seen visually; however, they emit radio waves at specific frequencies, enabling scientists to measure molecular abundances with precision, Senay says. The team then compared this ratio with the ratio of the same substances in the hydrogen cyanide of interstellar clouds. The large deuterium to hydrogen ratio in the hydrogen cyanide ejected from the cometary ices of Hale-Bopp is evidence for interstellar ices being the forerunner of comets, Senay says.
This research provides evidence that cometary ice is created within interstellar clouds that have not yet collapsed; the ice is then incorporated into the comet during its formation.
#
"Comet Hale-Bopp's Unusual Tail",574,0,0,0
Information from the European Southern Observatory
While famous Comet Hale-Bopp continues its long voyage towards the outer reaches of the solar system, observations proceed with telescopes in the southern hemisphere. These research programmes aim at a better understanding of the further development of this very active comet as it moves away from the Sun and slowly cools. Among the key questions are for instance: "When will it cease to display a dust tail?" and "Will the nucleus undergo outbursts during which much fresh material will be dispensed into space, as this has occasionally happened by other comets (e.g. Halley)?"
Hale-Bopp is now too faint (magnitude 8) to be seen with the unaided eye. At the end of January 1998, it is about 635 million kilometres from the Sun.
Recent observations at the ESO La Silla observatory with the 1-metre Schmidt Telescope have shown that the comet is still very active and, in particular, that it still possesses an impressive tail structure. Two photos taken in early January 1998 demonstrated an unusual appearance and continuing, enormous extent of this tail.
Both images showed the normal dust-tail in the anti-solar direction (upwards - to the North), extending over a distance of more than 4░ to the edge of the field. Since the photos were made near the time that the Earth was crossing the comet's orbital plane, they also showed the so-called \Ineck-line\i structure, a narrow and very straight feature in the same direction which is caused by sunlight reflected in the thin dust sheet in this plane. The true length of this feature depends on the exact geometry of the dust distribution, but may well be of the order of 1 AU (150 million km) or even more.
Interestingly, there is also a narrow, sunward \Ispike\i, or \Ianti-tail\i, that may be followed at least 0.5░ in the other direction. The tail/anti-tail phenomenon is interpreted by ESO astronomer \IHermann Boehnhardt\i as due to the presence of large and old dust grains that were released by the nucleus at least 100 days before these observations, possibly much earlier.
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"Mars Orbiter Camera Views the 'Face on Mars'",575,0,0,0
Shortly after midnight Sunday morning (5 April 1998 12:39 AM PST), the Mars Orbiter Camera (MOC) on the Mars Global Surveyor (MGS) spacecraft successfully acquired a high resolution image of the "Face on Mars" feature in the Cydonia region. The image was transmitted to Earth on Sunday, and retrieved from the mission computer data base Monday morning (6 April 1998). The image was processed at the Malin Space Science Systems (MSSS) facility 9:15 AM and the raw image immediately transferred to the Jet Propulsion Laboratory (JPL) for release to the Internet.
The picture was acquired 375 seconds after the spacecraft's 220th close approach to Mars. At that time, the "Face", located at approximately 40.8░ N, 9.6░ W, was 275 miles (444 km) from the spacecraft. The "morning" sun was 25░ above the horizon. The picture has a resolution of 14.1 feet (4.3 meters) per pixel, making it ten times higher resolution than the best previous image of the feature, which was taken by the Viking Mission in the mid-1970's. The full image covers an area 2.7 miles (4.4 km) wide and 25.7 miles (41.5 km) long.
#
"Third Cydonia Observation",576,0,0,0
The Mars Surveyor Operations Project is proceeding with the implementation of its third and final cluster of targeted imaging at Mars. This cluster will again target the two Viking Lander sites, a refined Mars Pathfinder landing site, and a new area in Cydonia.
On Tuesday, April 21st, at 1:43 PM PDT and on Wednesday morning, April 22, at 1:22 AM PDT, MGS will again attempt to image the sites of the Viking Landers on two consecutive orbits. Recall that on the first attempt, Viking 1 was slightly outside the camera's field of view. However, on the second attempt the site was in the image, but it was not possible to see the lander. The Viking 2 site has been covered with clouds on both previous attempts.
Then on Wednesday afternoon, April 22 at 1:00 PM PDT, MGS will again attempt to image the site of the Mars Pathfinder landing. This site was missed on the two previous attempts.
On Thursday afternoon at 12:17 PM PDT, MGS will again image a portion of the Cydonia region. Global Surveyor will again target to capture an image of the features known as "The City". This area contains features identified as "mounds", "city square", "pyramid" and the "fortress". The image will be targeted to capture portions of the "pyramid" and the "fortress", as well as "mounds".
As with the two previous images of the Cydonia region, the camera will be set to produce an image 1024 pixels wide so that the length of the image can be maximized to include as many features as possible. With a range from Cydonia to the spacecraft of 392 kilometers (244 miles), this will enable a resolution of 3.46 m/pixel (11.4 feet/pixel) and an image 3.5 km (2.2 miles) in width by 33 km (20.5 miles) in length.
The same probabilities of success of 30% to 50% will apply to each of these attempts based on navigation uncertainties and spacecraft attitude control performance. Experience with the first and second clusters of targeted images has shown that winter weather in the northern hemisphere of Mars at this time causes haze, dust storms, surface frost and heavy cloud cover to be significant factors in the success of seeing the targets clearly. The weather effects are not included in the probability of success estimates.
Results of the Cydonia imaging will be posted on the Internet, in the same manner as following the first and second observation attempts, at approximately mid-morning Pacific Time on Friday, April 24th. (When the playback of data from the spacecraft occurs overnight, as it does in this case, the image will be released shortly after the opening of business the following day.) If the landers are within the resulting images and can be identified, the image(s) containing it (them) will be released.
#
"El Nino Visualizations",577,0,0,0
NASA-Goddard Space Flight Center
Released 3/26/98
The animations and images below show a combined perspective of the El Ni±o induced change in the sea surface topography as viewed from space by the NASA TOPEX radar altimetry satellite and sea surface temperature observed by NOAA's AVHRR sensor for the period from January 1997 to March 1998. The three dimensional relief map shows a sea level rise along the Equator in the eastern Pacific Ocean of up to 34 centimeters with the red colors indicating an associated change in sea surface temperature of up to 5.4 degrees Celsius.
Also shown is a combined perspective of the the weakening of the Trade Winds across the Pacific Ocean and the associated increase in sea surface temperature. The convergence of the surface wind field into the anomalous warm water regions indicates a continued strengthening of this El Ni±o event. The sea temperature below the surface as measured by the TOGA TAO moorings is shown illustrating how the thermocline (the boundary between warm and cold sea water at 20 degrees Celcius) is flattened out by El Ni±o.
\B1997-98 Temperatures Beneath Sea\b
Views of sea surface height (represented by the bumps) and sea temperature (represented by the color). Red is 30 degrees C, blue is 8 degrees C. The thermocline is the border between the dark blue at the bottom and the cyan. The thermocline exists at 20 degrees C. \IData from 1/1/97 to 3/10/98\i.
(see picture sequence)
\B1997-98 Temperature Anomalies Beneath the Sea\b
Views of sea surface height (represented by the bumps) and sea temperature (represented by the color). Red is 10 degrees C above normal, blue is 10 degrees C below normal. Notice the large area of colder than normal water shutting off El Ni±o towards the end of the animation. \IData from 1/1/97 to 3/10/98\i.
(see picture sequence)
#
"High Wire Act May Be Best Way To Explore Europa",578,0,0,0
March 13, 1998: As NASA works to make space missions cheaper, it is looking at the possibility of using a long wire to power spacecraft exploring space around Jupiter where Galileo is gathering more hints that icebound Europa may have the right conditions for life.
Today, Dr. Dennis Gallagher of NASA's Marshall Space Flight Center will discuss electrodynamic tethers at the Ninth Annual Advanced Propulsion Research Workshop and Conference being held at the Jet Propulsion Laboratory in Pasadena, California.
In theory, a spacecraft could use a 10 km (6.2-mile) wire to augment rockets for propulsion once it reaches Jupiter.
"These are exciting possibilities that are worth exploring. The physics is wonderful," Gallagher said. "The engineering will be a challenge, though."
Which is to say that some very sophisticated controls will be needed to operate an electrical tether in Jupiter's dynamic environment.
An electrical tether uses the same principles as electric motors and generators. Move a wire through a magnetic field and you get an electrical current for power. Send electricity through a wire and you get a magnetic field that drags or pushes on any outside magnetic field.
This runs motors inside toys, appliances, disk drives, and generators in power plants, automobiles, and so on.
It can also generate electrical power for a satellite orbiting a planet with a magnetic field, or raise or lower the satellite's orbit - if the satellite has an electrically conducting tether.
NASA tested a Tethered Satellite System on the Space Shuttle in 1995 and 1996. Although it broke on the second mission, the tether produced some surprises in how electrical currents are produced and conducted by extended objects in space. Marshall Space Flight Center is now developing a Propulsive Small Expendable Deployer System - ProSEDS - that will speed a rocket stage's return to Earth.
If successful, it may be followed by an Electrodynamic Tether Upper Stage that would use the same principles to boost satellites to higher orbits, or a similar system on the International Space Station to help maintain its orbit.
"Jupiter is another path the program could take," said Gallagher, a plasma physicist at NASA/Marshall. "What we're suggesting is getting together with the Jet Propulsion Laboratory and doing an advanced tether study for a Europa orbiter mission."
The concept is to use a tether to propel the spacecraft and power its electrical system, thus saving the most precious of space resources, money. By reducing the amount of propellant needed once the spacecraft arrives at Jupiter, or the size of the electrical power system, the cost of the spacecraft also can be reduced, and it can be launched with a smaller, cheaper rocket.
An electrical tether will work only where nature provides both a magnetic field and a plasma (electrified gas). The motion of the wire through the magnetic field provides the energy, and the electrons in the plasma provide the return path that completes the electrical circuit.
The Earth's magnetic field and its ionosphere, which extends well into "empty" space, would do well for satellites here.
Jupiter is a bit more of a challenge, Gallagher explained.
Near the planet, where the plasma is densest, a 10 km (6.2 mile) tether would produce a 50,000-volt potential and a 20 amp current. That would be 1 megawatt of power flowing through a line just 1 mm (1/25th of an inch) thick.
"This would become a tremendous fuse and vaporize the tether," Gallagher said. This is also where engineering steps in and has to deal with the numbers developed by physics.
"You could only use the tether to conduct for brief intervals," Gallagher said. Theoretically, it could bring the satellite down from a high, 100-day orbit to a tighter, 5-day orbit. And the megawatt of power would be far more than than the 100 watts that the spacecraft would need during normal operations.
While the planet has a large magnetic field, its strength drops out towards the four large Galilean that are of greatest interest to scientists; the plasma density also drops. Europa is 9 Rj - nine Jovian radii, or 630,000 km (391,000 mi) - out.
"If you get that far out, densities have fallen substantially, and the field is pretty weak," Gallagher said. That means a much longer tether would be needed. The extra weight might offset the gains, and the tether would have a greater risk of being hit by a micrometeorite.
Oddly enough, another difficulty is the gravity gradient. The slight difference in gravitational pull across the length of the tether is what keeps it taut. But while Jupiter is the most massive planet in our solar system, it is also the largest. That means its gravity gradient is shallow more than 4 Rj where the probe would need to work.
The solution might be to spin the spacecraft so centripetal force keeps the tether taut. That, of course, complicates the electrical controls.
As for exploring Europa itself, Gallagher said that more needs to be known.
"Europa has a thin atmosphere and may have an ionosphere," he said. "Perhaps it has its own built-in blanket of current carriers." On the other hand, its magnetic field is very weak, so a longer tether might be required to generate enough current to power the spacecraft.
It might even be possible to extend a tether skyward from a Europa science station and power the the craft that way, Gallagher said.
So, the bottom line for now is a definite "maybe."
"One of the objectives of this study was to figure out whether it was worth looking at seriously," Gallagher said. "This study could just as easily have said, 'Don't bother.'" But it didn't.
"Europa is a potentially exciting place to use electrodynamic tethers."
#
"Galileo Mission Update 98",579,0,0,0
\JGalileo Starts Two-Year Europa Extended Mission\j
\JGalileo's Anomalous Behavior Corrected\j
\JGalileo's Test Data Analyzed\j
\JGalileo Transmits Information\j
\JGalileo Flies Past Europa\j
\JGalileo Communication Re-established\j
\JGalileo's Information Processed\j
\JNew Class of Dust Ring Discovered Around Jupiter\j
\JGalileo Completes Flyby of Europa\j
\JGalileo Update (13-19 April, 1998)\j
\JGalileo Update (20-26 April, 1998)\j
\JGalileo Update (27 April - 03 May, 1998)\j
\JGalileo Update (4-10 May, 1998)\j
\JGalileo Update (11-17 May, 1998)\j
\JGalileo Update (14 May, 1998)\j
\JGalileo Update (18-24 May, 1998)\j
\JGalileo Update (25-29 May, 1998)\j
\JGalileo Update (29 May, 1998)\j
\JGalileo Update (31 May, 1998)\j
\JGalileo Mission Finds Strange Interior of Jovian Moon\j
\JGalileo Update (29 June - 5 July, 1998)\j
\JGalileo Spacecraft Sees Volcanic Fireworks On Jupiter's Moon Io\j
\JGalileo Update (6 - 12 July, 1998)\j
\JGalileo Update (13 - 19 July, 1998)\j
\JGalileo Update (27 July - 2 August 1998)\j
\JGalileo Update (29 July, 1998)\j
#
"Galileo Starts Two-Year Europa Extended Mission",580,0,0,0
January 9, 1998
After yielding a rich harvest of science results in 1997, NASA's Galileo spacecraft wrapped up its primary mission on Dec. 7 and began a two-year follow-on journey, known as the Galileo Europa mission.
The transition from primary to extended mission brought a change in management. Bob Mitchell, who served as mission director for the last year of Galileo's primary mission, was appointed project manager for the Galileo Europa Mission, taking over from Bill O'Neil, who served as Galileo project manager for the flight to Jupiter and the two-year primary mission at Jupiter. O'Neil will serve as a consultant on the senior staff of JPL's Telecommunications and Mission Operations Directorate pending his next assignment at the Laboratory.
"A great bounty of Jupiter system science has been obtained and the continuing study of these data will surely add many more important discoveries," O'Neil said of the mission. "I've been involved with the Galileo mission since its beginning in 1977, and have been at the helm since 1990 for the flight to Jupiter, the first-ever outer planet entry and orbit insertion, and throughout the two-year primary mission tour of the Jovian system. I feel extraordinarily fortunate to have had this priceless, truly unique experience. But it is time for new challenges. I am delighted to turn the reins over to Bob Mitchell. Having worked closely with Bob for more than 25 years, I know that he will do a superb job leading the team."
"Accomplishing what we have set out to do with such a small team is going to be a real challenge," Mitchell said. "But we have an excellent team in place, and I'm looking forward to it."
The first flyby of the Galileo Europa Mission took place on Dec. 16, when the spacecraft swooped past Europa at an altitude of 200 kilometers (124 miles), making it the closest Europa encounter of the entire Galileo mission. The extended mission will include seven more Europa flybys, four encounters with Callisto, and one or two close flybys of Io, depending on spacecraft health.
Pictures and other data returned by Galileo during its primary mission continued to fascinate the public in 1997. New images of Europa revealed evidence of ice flows, a complex network of crisscrossed ridges, chunky ice rafts and relatively smooth, crater-free patches. The areas of rafting added to the mounting evidence of liquid oceans under Europa's icy crust at some point in its history. The presence of oceans would increase the odds that life could have existed there.
"We're intrigued by these blocks of ice, similar to those seen on Earth's polar seas during springtime thaws," said Dr. Ronald Greeley, an Arizona State University geologist and Galileo imaging team member. "The size and geometry of these features lead us to believe there was a thin icy layer covering water or slushy ice, and that some motion caused these crustal plates to break up."
Galileo investigators discovered a hydrogen atmosphere around Ganymede and both hydrogen and carbon dioxide in an atmosphere on Callisto. The spacecraft also found that Europa has an ionosphere, produced by ionization of its tenuous oxygen atmosphere. This finding came after a series of six occultation experiments, when the radio signal was interrupted while Europa was positioned between Galileo and Earth. These experiments were performed during Galileo's encounters with Europa in December 1996 and February 1997.
"While this discovery does not relate to the question of possible life on Europa, it does show us there are complex surface and atmospheric processes occurring there, and Europa is not just some dead hunk of material," said lead investigator Dr. Arvydas Kliore of JPL.
Galileo also transmitted new evidence of numerous high- temperature volcanoes on Jupiter's volatile moon, Io. One recent discovery revealed a new dark spot the size of Arizona on Io. The visible change occurred between Galileo's seventh and tenth orbits of Jupiter, and produced a dark area about 400 kilometers (249 miles) in diameter, surrounding a volcanic center named Pillan Patera.
"This is the largest surface change on Io observed by Galileo during its entire two-year tour of the Jovian system," said Galileo imaging team member Dr. Alfred McEwen, a University of Arizona research scientist.
Other significant results from Galileo this past year included the confirmation of the spacecraft's 1996 discovery of a magnetic field and magnetosphere on Ganymede, and the discoveries that all the Galilean moons except Callisto have a core. Callisto did show signs of surface erosion and blanketing at fine scale.
"Before Galileo, we could only make educated guesses about the structure of the Jovian moons," said Dr. John Anderson, a JPL planetary scientist. "Now, with the help of the spacecraft, we can measure the gravitational fields of the satellites and determine their interior structure and density. We can determine how the matter is distributed inside."
Galileo's instruments also detected some interesting, Earth- like phenomena on Jupiter, including the presence of lightning and aurora. Recent findings confirm the suspicion that the thunderstorms provide energy for the low pressure centers on Jupiter, which in turn feed the Great Red Spot, white ovals and other large storms.
In 1997, Galileo also found clusters of volcanic vents and hot spots in greatest concentration on Io in the areas closest and farthest from Jupiter. Other discoveries include evidence of salt and carbon dioxide in Europa's icy crust and landslides on Callisto.
While the spacecraft was busy making scientific history, Galileo team members made history of their own in January. O'Neil, Johnson, and others met with Pope John Paul II during a trip to Italy.
"None of us ever anticipated that Project Galileo would result in a papal audience, "O'Neil said. "The Pope seemed very interested in learning about Galileo results. He encouraged continuing exploration of the universe."
O'Neil and Johnson received honorary doctorates from the University of Padova and attended the Three Galileos Conference, a meeting designed to honor Galileo the man, Galileo the mission, and Galileo the new national telescope of Italy.
Members of the Galileo flight team are busy studying data on two incidents of anomalous behavior that have occurred, one during the spacecraft's December 16, 1997 flyby of , and one since. In both cases, the anomalies involved the attitude control subsystem, which controls where the spacecraft and scan platform are pointing. Team members believe the culprit may be one of the spacecraft's two gyroscopes. The gyroscopes are used to point the spacecraft when there is a need for very precise pointing control and knowledge of the spacecraft, usually for remote sensing science observations or orbit trim maneuvers.
Although the anomalies are not considered serious, until the situation was corrected, Galileo's radio antenna was pointing in a direction about 10 degrees from Earth, about eight degrees greater than the normal attitude for ideal transmission of information to Earth. Nonetheless, Galileo was still able to transmit pictures and other information stored on its tape recorder during the Europa flyby, but the data was sent at a lower rate to compensate for the large angle.
To help correct the spacecraft's attitude and speed up the data transmission, the Galileo team reduced the angle between Earth and the radio antenna by performing a Sun acquisition turn, which turned the spacecraft toward the Sun by using the Sun's bright light as a guide. This turn, which was performed Sunday night, means that the antenna is now about 3.1 degrees from Earth. The spacecraft is now operating normally, but the cause of the anomalies is still not fully understood, and the team is continuing its investigation.
This week, Galileo is expected to transmit to Earth high time resolution information on the interaction between Europa and the magnetic and electric field environment of Jupiter. Also on the playback schedule are images of Europa's Conamara Chaos region, an area of bright, icy crust that has broken apart, exposing darker underlying material. There may be some pictures of Europa's regions of mottled or "blotchy" terrain.
Scientists are particularly interested in seeing the pictures taken by Europa during the December 16 encounter, because the flyby was the closest ever to be performed by Galileo, with the spacecraft dipping down to 200 kilometers (124 miles) above the icy moon's surface.
The spacecraft recently began a two-year extended mission, known as the , which will include a total of eight Europa flybys, four of , and one or two of , as long as the spacecraft remains healthy.
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"Galileo's Test Data Analyzed",582,0,0,0
January 20, 1998
Members of the Galileo flight team are analyzing data from a test performed Friday night, which they hope will shed light on the cause of two recent incidents of anomalous behavior by the spacecraft. While the investigation continues, the spacecraft has resumed normal transmission to Earth of pictures and other science information stored on its onboard tape recorder.
While one anomaly occurred during the spacecraft's December 16, 1997 flyby of , and the other after the flyby, both involved the attitude control subsystem, which controls where the spacecraft and scan platform are pointing. Team members suspect the cause may have been one of the spacecraft's two gyroscopes. The gyroscopes are used to point the spacecraft when very precise pointing control and knowledge of the spacecraft's position and orientation are needed, usually for camera and other remote sensing science observations or for maneuvers that adjust the spacecraft's flight path.
The anomalies were not considered serious, but they did cause a temporary slowdown in the rate at which information was transmitted to Earth. That's because the anomalies caused Galileo's radio antenna to point in a direction about 10 degrees from Earth, about eight degrees greater than the normal attitude for ideal data transmission. However, information is now being transmitted at a normal rate once again, thanks to a spacecraft turn performed last week which pointed Galileo's antenna within 3 degrees of Earth, a normal angle.
Galileo has begun sending back to Earth some high-resolution pictures taken during the Dec.16 Europa encounter. That flyby was the closest ever to be performed by Galileo, with the spacecraft dipping down to 200 kilometers (124 miles) above the icy moon's surface. This week, Galileo will also return fields and particles information on the interaction between Europa and Jupiter's magnetic and electric field environment.
A flight path maneuver is planned for Thursday evening, Jan. 22 to prepare for Galileo's upcoming Europa encounter on Feb. 10. Special precautions have been taken in the design of this maneuver to minimize its vulnerability to any gyro problems. Because of its proximity to solar conjunction, when the Sun will be between Galileo and Earth, no data collection is planned except for radio science information.
The spacecraft recently began a two-year extended mission, known as the , which will include a total of eight Europa flybys, four of Callisto, and one or two of Io, depending on spacecraft health.
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"Galileo Transmits Information",583,0,0,0
January 30, 1998
NASA's Galileo spacecraft has been transmitting to Earth this past week pictures and other science information gathered during the Dec. 16, 1997 flyby of Jupiter's icy moon, Europa. The information, which had been stored on the spacecraft's onboard tape recorder, includes fields and particles observations of the interaction between Europa and Jupiter's magnetic and electric field environment.
Also included are pictures and observations of Europa's wedged regions and hot regions. Important observations of surface changes on the volcanic moon Io will help later on in the Galileo Europa Mission, when one or two close Io flybys are planned.
Next week, Galileo will transmit science information and pictures of Europa's regions of fretted shapes and regions of dark lines, as well as images of the Gilgamesh region on another Jovian moon, Ganymede.
There have been no further occurrences of anomalous behavior by Galileo's attitude control subsystem, which controls where the spacecraft and scan platform are pointing. Team members have confirmed that a hardware error in one of the spacecraft's two gyroscopes caused the two anomalies, which temporarily slowed down the rate at which data could be transmitted to Earth.
The gyroscopes are used to point the spacecraft when very precise pointing control and knowledge of the spacecraft's position and orientation are needed, usually for camera and other remote sensing science observations or for maneuvers that adjust the spacecraft's flight path. The team will continue studying the anomalies to determine whether they may recur and to design ways to work around the situation for the remainder of the mission.
A flight path maneuver was performed successfully on Thursday, Jan. 22, to prepare for Galileo's upcoming Europa encounter on Feb. 10. Special precautions were taken in the design of this maneuver to minimize its vulnerability to any gyro problems.
Another flight path maneuver will be performed on Sat., Feb. 7, if it is deemed necessary for fine-tuning before Galileo's Europa flyby on Feb. 10. Because of the solar conjunction, when the Sun is between Galileo and Earth, that flyby will include no data collection except for radio science information.
On Tues. Feb. 3, the spacecraft will be turned slightly to adjust the antenna position. The turn will be executed in normal mode with some precautions built in, but no problem is anticipated. This will be the final attitude adjustment needed before the end of February, when Galileo leaves its solar conjunction period.
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"Galileo Flies Past Europa",584,0,0,0
February 17, 1998
NASA's Galileo spacecraft successfully flew over at an altitude of 3,552 kilometers (2,207 miles) on Tues., Feb. 10. Because of the approaching , the spacecraft gathered some , but no other science information. During solar conjunction, the Sun passes between Earth and the spacecraft, hindering communication for a period of about 2-1/2 weeks. Therefore, the spacecraft will not transmit science data to Earth until the conjunction period ends.
During the Feb. 10 flyby, the spacecraft did transmit to Earth an observation taken during its closest-ever Europa flyby, which occurred on Dec. 16. During that encounter, Galileo passed above Europa's icy surface at an altitude of only 200 kilometers (124 miles). The latest observation returned to Earth was designed to help scientists learn more about the composition of an area of Europa characterized by the Pwyll crater.
The Galileo flight team adjusted the spacecraft's flight path on Sat., Feb. 7, to prepare for the Feb. 10 Europa flyby. Another flight path adjustment was made on Fri., Feb. 13 to make sure Galileo is on track for its next Europa encounter, scheduled for March 29. Both maneuvers were successful, even though special measures were utilized to counteract the effects of a gyroscope anomaly which has occurred twice since the Dec. 16 Europa flyby. Engineers believe solid state switches in one of the gyro units may be reacting to the effects of radiation from Jupiter, but they continue their analysis of the situation.
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"Galileo Communication Re-established",585,0,0,0
March 11, 1998
NASA's Galileo spacecraft is once again sending pictures and science information to Earth following a 2-1/2 week-hiatus caused by a period of Radio communications were hampered during this period, when the Sun passed between Earth and Galileo.
Now that full communications have been re-established, Galileo is sending to Earth pictures and information stored on its onboard tape recorder during the close flyby this past December. Included are observations of the Pwyll impact crater region, the Conamara Chaos region, and fields and particles information on Europa's interaction with Jupiter's magnetic and electric fields.
This batch of information was actually transmitted once before, but this retransmission of the recorded data allows scientists to fill gaps caused by transmission problems, and to replay particularly interesting observations and additional data.
A turn for attitude maintenance was performed successfully on Sat., March 7. On Tues., March 10, the spacecraft's attitude control system was tested. This test was designed to determine how Galileo was affected by intense radiation exposure during the February 10 Europa flyby. Intense radiation in the Jovian system is considered a prime candidate as a cause of recent anomalous behavior by an attitude control system gyroscope. After analyzing results of the test, it was determined that the gyro's performance had degraded further. Although this may affect future pointing accuracy and stability, the Galileo team believes strongly that spacecraft still can collect additional science information during future flybys.
Later this week, the spacecraft will perform a flight path adjustment to ensure that it is aimed correctly for its next encounter with Europa, scheduled for .
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"Galileo's Information Processed",586,0,0,0
March 23-27, 1998
Galileo spends the first five days of this week preparing for its with the Jupiter system, a flyby of scheduled for Sunday, March 29, at an altitude of about 1,645 kilometers (1,022 miles). Preparations include processing and transmitting to Earth all remaining science information from the spacecraft's previous complete science encounter with the Jupiter system in December 1997. In addition, on Wednesday, the flight team completes constructing and transmits to the spacecraft the set of computer instructions required to perform the final flight path correction prior to the Europa flyby.
The flight path correction is scheduled for execution on Thursday, three days prior to the close flyby. Also on Thursday, the flight team will transmit to the spacecraft the set of computer commands that will control all spacecraft activity during the encounter period. Finally, on Friday, regular maintenance is performed on the spacecraft's onboard tape recorder which is used to store all of the science information gathered during Galileo's encounters.
The science information that is processed and transmitted to Earth this week includes data from three observations: two by Galileo's and one by its . The first of the spectrometer data contains information further describing the surface composition on a region of Europa characterized by the much publicized wedge-like features. These features are believed to be evidence of the existence of a liquid ocean, or, at least, soft ice, under Europa's surface. Not unrelated to the possibility of a subsurface ocean, the photopolarimeter radiometer observation contains information on the temperature of the surface of Europa.
The observation is part of a series designed for the to search for possible hot spots on Europa. These hot spots, if they are found, could be further evidence of a heat source on Europa that could lead to the creation of a liquid ocean or soft ice environment. The second set of spectrometer data deviates from the Europa focus and contains data describing volcanic activity on . This data is important for keeping tabs on Io and will help scientists plan their observations later in the Galileo Europa Mission.
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"New Class of Dust Ring Discovered Around Jupiter",587,0,0,0
April 3, 1998
Scientists have found evidence for a new ring of dust that occupies a backward orbit around Jupiter, based on computer simulations and data from NASA's Galileo spacecraft, it is reported in today's issue of Science magazine.
A team led by researchers at the University of Colorado at Boulder reported that a faint, doughnut-shaped ring of interplanetary and interstellar dust some 1,126,000 kilometers in diameter (about 700,000 miles) appears to be orbiting the giant planet. Evidence for the new ring's existence comes from computer simulations that correlate with data collected by a dust detector aboard the Galileo spacecraft has detected this ring by capturing some of its dust, said Dr. Joshua Colwell, a research associate at the university's Laboratory for Atmospheric and Space Physics.
Surprisingly, the researchers say, most of the interstellar and interplanetary dust particles appear to be in a "retrograde" orbit -- that is, moving in the opposite direction of the rotating planet and its moons, Colwell said. The reason for the backward orbit of the tiny particles is not yet clear, he said.
The paper in Science magazine was authored by Colwell, research associate Dr. Mihaly Horanyi, also of the Laboratory for Atmospheric and Space Physics, and planetary scientist Dr. Eberhard Grun of the Max Planck Institute for Astrophysics in Heidelberg, Germany, who is the principal investigator on Galileo's dust detector.
NASA's Voyager 2 detected an uneven dust ring around Jupiter in 1979 that scientists believe was created by the collisions of small moonlets with micrometeoroids in the Jovian system. But the newly identified ring of dust with smoke-size particles originating from beyond the Jovian system appears to be much larger, more sparse and, possibly unique in the solar system.
"I suspect we may wind up seeing something similar at Saturn," said Colwell. Launched in 1997, NASA's Cassini spacecraft will reach the ringed planet in 2004.
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"Galileo Completes Flyby of Europa",588,0,0,0
April 9, 1998
The Galileo spacecraft successfully completed its most recent flyby of Jupiter's moon Europa on March 29, and indications are there was no change to the gyroscope performance. Because one of the two gyros had been acting up, the closest approach to Europa was carried out in cruise mode, with the gyros turned off; the spacecraft used only stars to orient itself and point its instruments. However, an attitude-control system performance test showed that the gyros did not degrade further during this latest pass through Jupiter's intense radiation environment. Galileo project engineers have pinpointed a single computer chip as the cause of the anomalous behavior. This particular chip has received more radiation exposure than other similar chips in the gyro electronics.
This week, Galileo transmitted to Earth pictures and other science information gathered during the latest Europa flyby. This includes one of three observations by the photopolarimeter radiometer designed to refine temperature variation maps of Europa's surface. This will help scientists understand surface ages and composition and the process that may have formed the surface. In addition, there is information from instruments that study magnetic fields and charged particles on the interaction between Europa and Jupiter's magnetic and electric field environment. The camera and the near infrared mapping spectrometer have returned information on a region of dark lines and the Mannann'an crater on Europa. Data gathered by the spectrometer of the south pole of Jupiter's volcanic moon, Io, provides the spacecraft's best view of the area until late 1999.
On Friday, April 10, the spacecraft will perform regular propulsion system maintenance and perform a turn to keep its radio antenna pointed toward Earth.
Galileo's next Europa flyby will take place on May 31, 1998, at an altitude of 2,521 kilometers (1,566 miles). The spacecraft successfully completed its primary mission in December 1997 and is now in its two-year extension, the Galileo Europa Mission. Current plans include four more Europa flybys after the May encounter, four Callisto flybys, and one or two of Io, depending on spacecraft health.
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"Galileo Update (13-19 April, 1998)",589,0,0,0
Processing and transmission to Earth of pictures and other science information, a process also known as playback, is the sole order of business this week on Galileo. The data was gathered and stored on the spacecraft's onboard tape recorder during the last few days of March as the spacecraft flew through the Jupiter system and within 1,645 kilometers (1022 miles) of the surface of Jupiter's moon Europa.
Most of the data return this week was obtained by the spacecraft camera, or Solid State Imaging (SSI) subsystem, during the close flyby of Europa. Among the pictures returned this week we find a couple containing the Mannann'an crater. Taken from slightly different angles, this pair of images will be combined to form a stereo image of the region. A similar pair of pictures containing a region of dark spots on Europa is also returned this week. This pair will also be combined to form a stereo image of the region. Regional coverage of this same dark spot area was obtained in November 1997.
Another SSI observation provides photometric information describing Europa's surface. These photometric measurements will tell us how intensely light is reflected from the surface and provide additional information on its makeup. Finally, a picture of a region of triple-bands is returned. These triple-band features are believed to be formed when Europa's surface cracks, material upwells from below the surface and spills to both sides of the central crack. They are considered evidence for the possible existence of a sub-surface ocean or, at the least, soft ice.
Sprinkled throughout the week, similar to last week's playback schedule, is information from the fields and particles instruments describing the interaction of Jupiter's magneto sphere with Europa. Remember that Jupiter's magneto sphere is that region of space where Jupiter's magnetic and electric fields dominate those fields generated by the solar wind. The charged particles that make up the magneto sphere co-rotate with Jupiter, at a rate of about 1 revolution every 10 hours, and get disturbed as they sweep past each of the satellites in orbit around Jupiter. Each new bit of data describing this interaction will help scientists understand the phenomena at work.
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"Galileo Update (20-26 April, 1998)",590,0,0,0
Galileo continues on the outbound leg of its orbit around Jupiter as it processes and transmits to Earth science information safely stored on the spacecraft's on-board tape recorder. The information was gathered by the spacecraft's 11 instruments as they flew past Jupiter and its moons late last month. On Thursday, the spacecraft performs a small turn to keep its antenna pointed toward Earth. These turns are required to maintain the right telecommunications conditions and science data flowing to the ground.
The batch of information processed and transmitted this week is dedicated primarily to observations of Jupiter's icy moon Europa. The spacecraft camera, or solid state imaging subsystem, returns two pictures of Europa this week. The first is a high-resolution picture of a circular feature known as Tyre Macula. This feature is believed to have been created by the impact of a mountain-sized asteroid or comet. This region was imaged at lower resolutions during Galileo's primary mission in April 1997 and is scheduled for imaging again at the end of May. The other picture that is returned this week includes a region of bright plains that transitions to a series of wedge shaped features. The spacecraft's Near Infrared Mapping Spectrometer also returns two observations of the same region of Europa. The region is characterized by pull-apart wedges, but also contains dark spots.
Finally, continuing from previous weeks, is the return of science information from the fields and particles instruments describing the interaction of Jupiter's magnetosphere with Europa.
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"Galileo Update (27 April - 03 May, 1998)",591,0,0,0
Many different activities are on Galileo's to-do list this week as the spacecraft passes through apojove (the furthest distance from Jupiter for each orbit) and starts heading back toward Jupiter again. Processing and transmission to Earth of science information, also known as playback, continues throughout the week. Included on this week's schedule is data from Europa, Ganymede, Io and Callisto. Playback of data stored on the onboard tape recorder is interrupted several times this week to perform important navigation and engineering activities.
On this week's playback schedule, we find information collected by the spacecraft's camera, or solid-state imaging subsystem (SSI), the Near Infrared Mapping Spectrometer (NIMS) and the photopolarimeter radiometer (PPR). Europa information that is retrieved this week includes data on the materials, obtained by NIMS, and temperatures, obtained by PPR, that are found on its surface. Global views of Europa, by SSI and NIMS, are also processed and transmitted to Earth.
A global view of Ganymede obtained by SSI is returned to Earth this week. The image is expected to provide more information on the radius, shape, color and photometry of the satellite. Remember that photometry is the measurement of light intensity and it helps to identify different materials. The image is also expected to tell scientists whether frost on the surface of Ganymede is mobile enough to be noticed at a global scale. SSI also returns an image of Io obtained while in eclipse.
Due to the lack of sunlight, these types of images have proven to be the best way to discover and monitor lava temperatures and interactions of plumes from Io with Io's atmosphere and Jupiter's magnetic and electric field environment. Finally, NIMS returns a global map of Callisto. Together with an observation planned for May 1999, the information gathered from this observation is expected to shed some light on open questions regarding differences in materials found in different regions of Callisto.
Engineering and navigation activities are initiated this week when the spacecraft performs regular maintenance on its onboard tape recorder. On Thursday, the spacecraft will perform a flight path correction in preparation for its next encounters with Europa and Jupiter in late May. Regular maintenance of the spacecraft's propulsion system is performed on Friday. And on Sunday, the flight team will transmit commands to the spacecraft to change onboard attitude control software. The changes to the software will allow the attitude control computer to, on its own, interpret the anomalous behavior of the gyroscopes. Back here on Earth, the behavior of the gyroscopes will continue to be monitored and updates to the onboard software will be made as required.
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"Galileo Update (4-10 May, 1998)",592,0,0,0
Galileo continues to process and transmit to Earth science information stored on the onboard tape recorder during the spacecraft's pass through the heart of the Jovian system at the end of March 1998. Science data processing is interrupted once this week as the spacecraft runs a performance test on the attitude control subystem to check on computer software changes made over the weekend. The software changes should make it possible for the attitude control computer to correct and use the anomalous data being produced by one of the spacecraft's two gyroscopes. The gyroscope has been behaving anomalously since December 1997.
This week's processing and transmission schedule includes observations that have already been processed and transmitted to Earth once. A normal part of operations on Galileo, this reprocessing and retransmission opportunity provides the chance to fill all data gaps caused by transmission problems. It also allows any particularly interesting information to be reprocessed under different processing conditions and retransmitted to Earth. Finally, it allows new parts of observations, that would not otherwise have been returned, to be processed and transmitted to Earth.
The solid state imaging, or camera, team has three observations on this week's schedule. The first contains a region of Io that scientists hope to image at much higher resolution toward the end of the Galileo Europa Mission in October 1999. The image that is returned this week will provide a context for the October 1999 image. The two other images returned by the camera team contain different regions of Europa.
The first looks at the Mannann'an crater region and the second at a region of dark spots. The near infrared mapping spectrometer team returns the other two observations scheduled for this week. The first is the highest resolution map of Io that will be obtained during the Europa Campaign (December 1997 - April 1999) of the Galileo Europa Mission. It also contains the best view of the south pole of Io to date, although a better view is planned for observations in November 1999. The second observation captures Europa and a region of ice rifts.
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"Galileo Update (11-17 May, 1998)",593,0,0,0
Galileo spends most of the week processing and transmitting to Earth science information stored on the spacecraft's onboard tape recorder. All of the data on this week's schedule contains information describing Jupiter's icy moon Europa. The data set returned this week was gathered by the spacecraft's camera, near infrared spectrometer and suite of fields and particles instruments during the spacecraft's close flyby of the moon, just over 6 weeks ago.
Some data from last week's schedule slipped into this week when Galileo released antenna time at the Deep Space Network's 70-meter antenna in Canberra, Australia. The antenna time was released to support radio frequency observations of the newly identified gamma ray burst you may have heard about in the news. The burst was located in a relatively close galaxy and the Canberra antenna was considered a key part for very long baseline interferometry (VLBI) observations required to study the event.
Last week's efforts to change the attitude control computer's onboard software have encountered a glitch. It appears that the onboard software was successfully modified, but the effect on the gyroscopic data is not as expected. The flight team will continue to gather and analyze spacecraft engineering data to determine what went wrong. Another gyroscope performance test is scheduled this Friday.
This week's information processing and transmission activities continue to retrieve data from a section of the onboard tape recorder that has already been accessed once this orbit. This second processing and transmission opportunity allows data gaps to be filled, re-processing of data with different parameters, or selection of entirely new data.
On the data return schedule we find two observations by the spacecraft's camera of a region of Europa characterized by dark spots. Together they will provide a stereo topographic view of the area. The camera team also returns an observation of a region containing characteristic triple bands. Finally, the camera team returns an observation designed to provide photometric measurements of Europa's surface. Photometry is the measurement of light intensities which can then be used to help identify the different materials on the surface of Europa.
The near-infrared spectrometer team contributes to this week's schedule by scheduling the return of an observation of a region of Europa containing dark spots and pull-apart wedges. The observation is expected to provide more information on the materials that make up this region of Europa. Finally, the fields and particles instruments return measurements that will add to the repository of information describing the interaction of Jupiter's magnetic and electric fields with Europa.
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"Galileo Update (14 May, 1998)",594,0,0,0
The Galileo spacecraft has spent the week processing and transmitting to Earth pictures and science information gathered during its March 29 flyby of Jupiter's moon Europa. The material had been stored on the spacecraft's onboard tape recorder.
Some of the information would have been transmitted last week, but it was delayed in the name of science when the Galileo team gave up some antenna time at the Deep Space Network's 70- meter (230-foot) antenna in Canberra, Australia. The antenna was needed to support radio frequency observations of a newly identified gamma ray burst.
Included in this week's batch of information transmitted to Earth by Galileo are two images of a region of Europa notable for its dark spots. Together, these images provide a stereo topographic view of the area. Another observation measures the varying light intensities on Europa, information which helps scientists identify different surface materials. An observation from Galileo's near-infrared spectrometer should provide more information on the materials that make up the region of Europa which has dark spots and pull-apart wedge sections.
The spacecraft is sending back previously recorded information that will beef up knowledge of the interaction between Jupiter's magnetic and electric fields and Europa.
Last week, the Galileo team modified the spacecraft's onboard attitude control software. However, the adjustment did not change the gyroscope's behavior as the team had hoped. A second modification was made Wednesday, May 13, and early tests indicate that the procedure was a success. Another gyro performance test is scheduled this Friday, May 15, and the Galileo team expects it will confirm that the attitude control system is now performing as planned. The attitude control system has been behaving anomalously since the spacecraft's closest flyby to Europa last December 16. The Galileo team has been able to operate the spacecraft in such a way that the anomaly has had very little effect on the spacecraft's performance.
Nonetheless, engineers continue to analyze the situation, which they believe is related to the spacecraft's repeated exposure to Jupiter's strong radiation.
Galileo's next Europa flyby will take place on May 31, 1998, at an altitude of 2,521 kilometers (1,566 miles). The spacecraft successfully completed its primary mission in December 1997 and is now in its two-year extension, the Galileo Europa Mission. Current plans include four more Europa flybys after the May encounter, four Callisto flybys, and one or two of Io, depending on spacecraft health.
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"Galileo Update (18-24 May, 1998)",595,0,0,0
With just about two weeks remaining before its next close encounter with Jupiter's moon Europa, Galileo spends this week continuing to process and transmit to Earth science data gathered during its previous encounter in late March. Data processing is interrupted once this week, on Wednesday, to turn the spacecraft to keep the radio antenna pointed toward Earth, and to perform regular maintenance on the spacecraft's propulsion system.
Only two observations are on the processing and transmission schedule this week. Both were performed by the spacecraft's camera, or solid state imaging subsystem, and contain science information describing Europa. The first is a high-resolution picture of the Tyre Macula crater region. This region contains a circular feature about 140 kilometers (87 miles) in diameter (about the size of the island of Hawaii) and is thought to be the site where an asteroid or comet hit Europa's ice crust. The second observation contains a region that shows a transition from bright plains to pull-apart wedges. These features suggest that the surface crust has been separated and filled with material from below the surface.
Last week, flight team engineers successfully identified and corrected a minor error in an update to the attitude control computer's onboard software that had been performed on Sunday, May 3. The original software update had not been performing as expected. Testing performed after this latest update has shown that the software is now performing as designed and the attitude control computer should be able to correct and use the output from the gyroscope that has been behaving anomalously since December 1997.
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"Galileo Update (25-29 May, 1998)",596,0,0,0
Galileo spends this week preparing for its next passage through the heart of the Jupiter system. The encounter is scheduled to start this Saturday, May 30, and features a flyby of Jupiter's moon Europa at an altitude of 2516 kilometers (1564 miles).
Preparation for the encounter includes completion of processing and transmission to Earth of science information still stored on the spacecraft's onboard tape recorder. The data was acquired during Galileo's previous flyby of Europa in late March.
Data processing and transmission is interrupted once this week, on Thursday, when the spacecraft executes the final flight path correction prior to the close flyby of Europa.
Five observations remain on the data processing schedule. The first is a global color observation of Europa, performed by the spacecraft's camera. The image will provide information describing the global geology of Europa, specifically the origin, composition, and distribution of materials on the surface. The camera team also returns a global color observation of Ganymede.
Similar to the Europa observation, the information returned will describe the radius, shape, color, and composition of Ganymede's surface. An observation of Io is also returned by the camera team. Acquired while Io was eclipsed from the sun by Jupiter, this type of data has proved to be the best way to discover and monitor lava temperatures and barely visible interactions between volcanic plumes, Io's atmosphere and Jupiter's magnetosphere.
The remaining two observations are returned by the near infrared mapping spectrometer team. One is the second of a set of three distant observations of Europa. The other is a global map of Callisto designed to provide more data describing the materials found on the surface.
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"Galileo Update (29 May, 1998)",597,0,0,0
The Galileo spacecraft is preparing for its next encounter with Jupiter's moon Europa, scheduled for this Sunday, May 31 at 2:12 p.m. Pacific Daylight Time, at an altitude of 2,521 kilometers (1,566 miles).
On Thursday, May 28, Galileo completed most of its approach maneuver sequence of commands, but the process was stopped because of an error in the construction of the sequence. The spacecraft aborted the sequence, including the remaining playback of recorded data, and put itself in a "safe mode." That means the spacecraft places itself in a low-activity state until it receives new instructions from the ground. The spacecraft is set to resume its planned activities after a new Europa encounter sequence is radioed to Galileo via the Deep Space Network tonight.
Despite the halt in the spacecraft's planned sequence, Galileo's onboard tape recorder transmitted to Earth 97 percent of the pictures and information that had been scheduled for playback during this orbit. The data were gathered during Galileo's flyby of Europa in late March.
Galileo's navigation team has determined that the encounter activities, which start on Saturday, May 30, will not be affected significantly by the safing event. This will be the first encounter since December of last year to use the full gyroscope capability. The gyro anomaly has been corrected, and at this point, the attitude control system is believed to be fully functional. Continual calibrations will be needed to maintain this capability.
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"Galileo Update (31 May, 1998)",598,0,0,0
The level of activity increases dramatically today, the second day of the Europa-Orbit 15 encounter period. The bulk of today's activity is geared toward the acquisition of science information to study Europa. The spacecraft flies over Europa's surface at a distance of 2516 kilometers (1564 miles) at 2:13 PM (PDT).
The Europa flyby is preceded by passage through the point of closest approach to Io for this orbit, which occurs at 1:59 PM (PDT) at a distance of 313,000 kilometers (194,000 miles). In the evening, the spacecraft passes through the point of closest approach to Jupiter. Jupiter closest approach occurs at 7:35 PM and a distance of 633,000 kilometers (393,000 miles).
The radio science team spends the day measuring changes in Galileo's radio frequency as the spacecraft flies past Europa. Using the Doppler effect, the change in frequency can be related to change in the speed of the spacecraft and Europa's gravitational pull on the spacecraft. These measurements are taken almost every orbit and are used to improve the map of the gravity field produced by Europa.
Today's remote sensing activity on Europa involves all four remote sensing instruments. The spacecraft camera focuses on a region containing the largest known massif on Europa. A massif is a block of crust that is surrounded by faults and has been displaced or moved without having broken apart. The region under observation is named Cilix. Two images of this region are obtained to allow the construction of a stereo image. An observation of the same region is also taken by the near infrared spectrometer.
The near infrared spectrometer data will provide scientists with information on what types of materials are found in this region. The camera also looks at an unusually rugged region containing pits and mounds with a very prominent ridge. The region is located east of the Tyre Macula impact crater and will also be imaged twice to allow for the construction of a stereo image. Both the near infrared spectrometer and the ultraviolet spectrometer also look at this region in order to provide composition information.
Observations of Europa at global and regional scales are sprinkled throughout the day. The camera looks at a previously unexplored region of mottled, or blotchy, terrain and also takes photometric measurements (light intensity that provides clues for material identification) on a global scale. The near infrared mapping spectrometer looks for surface composition information, also on a global scale, and is accompanied on the search by the ultraviolet spectrometer.
Surface temperatures and thermal characteristics are examined by the photopolarimeter radiometer instrument. This data is expected to help determine how compact the surface on Europa is, how old it is, what makes it up and how it got the way it is. Both high and low resolution observations are planned.
Last, but not least, for about an hour surrounding the Europa flyby, the fields and particles instruments perform a high time resolution (several hundreds of bits per second) recording of measurements of the dust, magnetic and electric field environment. These data will help scientists improve their understanding of the interaction between Europa and its surroundings.
A few observations of Io are scattered throughout the day. Two of them are performed by the near infrared mapping spectrometer in conjunction with the ultraviolet spectrometer. A third is performed by the ultraviolet spectrometer alone. The joint observations are the second highest resolution observations of Io planned for the Europa orbits of the Galileo Europa Mission. The Pele, Marduk and Reiden volcanic regions are targeted for observation.
These same regions are planned for observation again in October 1999. The remaining ultraviolet spectrometer observation is planned to provide data that will complement information gathered during Galileo's primary mission. The data should provide information that will allow scientists to better understand the density of Io's atmosphere and how it varies with location on the moon and the passage of time. Scientists also hope to uncover relationships to volcanic activity and sulfur frost sublimation.
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"Galileo Mission Finds Strange Interior of Jovian Moon",599,0,0,0
4 June, 1998
New data from NASA's Galileo spacecraft have prompted scientists to modify their concept of the interior structure of Jupiter's moon, Callisto, and suggest that Callisto has evolved differently than the other largest Jovian moons -- Io, Ganymede and Europa. The new findings, to be published in the journal Science on Friday, June 5, will be presented Monday, June 8 at the American Astronomical Society meeting in San Diego, CA.
"Previous Galileo data had indicated that Callisto's interior was totally undifferentiated," said Dr. John Anderson, planetary scientist at NASA's Jet Propulsion Laboratory, Pasadena, CA. "But new information suggests Callisto has a strange interior--it's not completely uniform nor does it vary dramatically. There are signs that interior materials, most likely compressed ice and rock, have settled partially, with the percentage of rock increasing toward the center of Callisto."
The new information was collected during Galileo's third Callisto encounter in September 1997. Anderson reported on the findings, along with UCLA geophysics and planetary physics professor Gerald Schubert, a Galileo gravity investigator, and Dr. William B. Moore, also of UCLA; and Dr. Robert A. Jacobson, Eunice L. Lau, and William L. Sjogren of JPL.
Scientists now believe Callisto is different from Io, Ganymede and Europa, which have differentiated structures with separated layers. There is strong evidence that Ganymede is separated into a metallic core, rock mantle, and ice-rich outer shell, while Io has a metallic core and a rock mantle but no ice.
"The fact that Callisto is the only one of the four large Jovian moons that is not completely differentiated raises an intriguing possibility," said Schubert. "Because Io, Ganymede and Europa are closer to Jupiter, they have been more affected by gravitational squeezing and subsequent heating. Over time, the forces exerted on the three inner moons have caused different constituents such as water ice, rock, and metal to separate into different layers. However, because Callisto is farther from Jupiter, it is "half-baked" compared to the other moons, with its ingredients somewhat separated but still largely mixed together," he said.
"Learning about the structure of these celestial bodies enhances our knowledge of how all planets and moons form and evolve, including our own Earth and Moon," Schubert added.
Scientists had previously reported a differentiated interior for Europa, consisting of a metallic core surrounded by a rock mantle and a water ice-liquid outer shell. They are now refining the model by studying the newest Galileo data, including that gathered during the closest-ever Europa flyby in December 1997, at an altitude of 205 kilometers (127 miles).
Europa's metallic core could be up to half the size of the moon's radius, with the water ice-liquid shell estimated to be between 80 to 170 kilometers thick (50 to 106 miles), with 100 km (62 miles) considered the most likely thickness. As more data become available from additional flybys, scientists hope to learn more about Europa's structure. Europa is of particular interest because of the prospect that liquid oceans may lie beneath its icy crust.
Information about the interior structure of Jupiter's moons is obtained by studying radio Doppler data that is gathered when the Galileo spacecraft flies by the satellites. Each moon exerts a gravitational tug, and the strength of that tug is affected by the distribution of rock inside. The tug, in turn, changes the spacecraft's speed and the radio frequency of its signals. By studying those changes, scientists can characterize the rock content and structure of the body.
The Galileo spacecraft entered orbit around Jupiter on December 7, 1995, and spent two years studying Jupiter, its four largest moons and its magnetosphere during its primary mission. The spacecraft is now in the midst of a two-year extension, known as the Galileo Europa Mission.
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"Galileo Update (29 June - 5 July, 1998)",600,0,0,0
This week Galileo's processing and transmission of pictures and science information to Earth continues, but is interrupted to perform important engineering tasks.
On Tuesday, the spacecraft will perform a gyroscope performance test. The test will provide information that will allow engineers here on Earth to determine whether performance of the gyroscope continues to degrade. The gyroscope, one of two on the spacecraft, has been producing anomalous data since December 1997.
Since March 1998, a modification to the software used by the attitude control computer has allowed correct interpretation of the gyroscope data. If the data from this test shows that the gyroscope performance has changed, engineers may decide to make minor adjustments to parameters in the modified attitude control software.
The other engineering activities for this week are performed on Wednesday. Data processing is interrupted to perform regular maintenance on the spacecraft's propulsion system, and to turn the spacecraft to keep the radio antenna pointed toward Earth.
This week's data return activity begins with the retrieval of data from a section of the onboard tape recorder that has already been accessed once during this orbit. This second processing and transmission opportunity allows the Galileo project team to fill in data gaps, select entirely new data, or re-process data with different parameters.
On this week's data return schedule, the near infrared spectrometer returns two observations. Both of the observations cover the Pele, Marduk and Reiden regions of Jupiter's fiery moon Io. Observations of these regions are also planned for late 1999 when the spacecraft is scheduled, if healthy, for two close flybys of Io.
The spacecraft's camera returns the other three observations scheduled this week. One is another moderate-resolution image taken near Europa's terminator. The other is a photometry observation of a region observed in April 1997 during Galileo's primary mission. The photometric data will provide information about how intensely light is reflected from Europa's surface. This will provide scientists with clues for identifying the surface materials found in the region.
Finally, the camera returns an image of the Cilix region of Europa. In low resolution Voyager images, Cilix appeared to be an impact crater. Images obtained during Galileo's first orbit caused scientists to think that the feature was instead a large mound. The data recently returned during the first pass through the data stored on the tape recorder have settled the issue: Cilix is, in fact, an impact crater. The last of this observation will be returned in this second pass.
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"Galileo Spacecraft Sees Volcanic Fireworks On Jupiter's Moon Io",601,0,0,0
July 2, 1998
New observations by NASA's Galileo spacecraft reveal dozens of volcanic vents on Jupiter's fiery moon Io where lava sizzles hotter than any surface temperatures recorded on any planetary body in our solar system. Temperatures this high are not known to have occurred on Earth for billions of years. At one such volcanic vent, known as Pillan Patera, two of Galileo's instruments have indicated the lava temperature may be 2,000 Kelvin (3,140 degrees Fahrenheit).
"The most likely explanation for these very high temperatures is that the eruptions contain magnesium-rich silicates," said Dr. Alfred McEwen of the University of Arizona, Tucson, AZ, a member of Galileo's solid state imaging camera team. "We've tentatively identified magnesium-rich orthopyroxene in lava flows around these hot spots. This leads us to conclude that silicate volcanism is taking place with lava compositions expected to melt at a very high temperature. We must now think of Io's volcanoes in terms of the type of very high-temperature silicate volcanism which was found on Earth during its early days, and which we suspect occurred also on Venus and Mars."
The new findings by the Galileo camera and the spacecraft's near infrared mapping spectrometer have updated scientists' information on Io's volcanic processes. Previously, Io observations made by the Voyager spacecraft in 1979 put the highest temperature estimates at about 650 Kelvin (710 degrees Fahrenheit). This led many scientists to believe that Io's volcanic activity was caused by low-temperature sulfur volcanism. In 1986, ground-based telescope observations increased the temperature estimates to above 900 Kelvin (1,160 degrees Fahrenheit), which suggested that silicate volcanism was occurring at least occasionally, just as it does on Earth today. In 1996 and 1997, Galileo identified 30 locations with temperatures higher than 700 Kelvin (800 degrees Fahrenheit).
"This new data indicate that high-temperature eruptions on Io are a basic and common part of its active volcanic processes," said Dr. Torrence Johnson of JPL, Galileo project scientist. Johnson led the group that found the high temperature eruption in 1986. He is also a member of the near infrared mapping spectrometer team. "Io's current volcanic activity may have a lot in common with ancient volcanic processes on Earth and other planets. Since the geologic record from those times is very sparse, it's quite exciting to be able to study this type of volcanism going on today."
"This discovery of high-temperature silicate volcanism provides us with an extremely important clue to understanding the geophysical processes within Io," McEwen explained. Io is heated by periodic tides as it orbits Jupiter, along with the other Galilean satellites (Europa, Ganymede and Callisto).
Armed with this new information, scientists also hope to learn more about the composition of Io's crust. "Io's extreme volcanic activity is expected to result in a low-density crust rich in silica, sodium and potassium," said McEwen. "However, the high-temperature volcanism suggests that the crust may be composed of heavier lavas."
Galileo's solid state imaging camera observed Io during 11 eclipses in 5 orbits, when Io was in Jupiter's shadow, and sunlight was blocked so the camera could better see the glowing volcanic vents. Io's hot spots were also studied by the spacecraft's near infrared mapping spectrometer during 11 orbits, mostly when Io was not in eclipse. The camera provides high spatial resolution to image the hottest features and map color variations, while the spectrometer can observe at many wavelengths and is sensitive to a wider temperature range. Thus, the combination of both instruments provides a powerful means to study Io's volcanism. The camera and spectrometer together have discovered a total of 41 hot spots on Io.
Scientists hope to gather more detailed information about Io with two planned close flybys in late 1999, as long as the Galileo spacecraft remains healthy. Galileo has been orbiting Jupiter and its four largest moons, including Io, for 2-1/2 years. It is currently in the midst of an extended journey, known as the Galileo Europa Mission, with eight flybys of Europa and four of Callisto, in addition to the Io flybys.
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"Galileo Update (6 - 12 July, 1998)",602,0,0,0
Galileo spends all of this week processing and transmitting to Earth pictures and science information stored on its onboard tape recorder. The data was acquired during the spacecraft's close flyby of Jupiter's icy moon, Europa, in late May.
This week's playback schedule continues to access sections of the tape recorder from which data has already been retrieved once this orbit. A part of normal operations on Galileo, this second cut at the tape recorder provides the Project team the chance to fill all data gaps caused by transmission problems. It also allows the team to select entirely new data, or re-process data with different parameters.
This week's playback schedule includes three observations obtained by the camera, two observations by the near infrared mapping spectrometer and a portion of an observation recorded by the fields and particles instruments.
Two of the images returned by the camera capture an unusually rugged region of Europa. The region is east of the better known Tyre Macula impact crater and is characterized by pits, mounds and a very prominent ridge. The remaining image contains a previously unexplored region of Europa. The region's mottled or blotchy appearance is believed to indicate the presence of contaminants in the ice.
The near infrared mapping spectrometer observations focused on two different regions on Europa. The regions are located on Europa's leading side, which is shielded from impacts by particles carried by Jupiter's magnetosphere. You see, as the magnetosphere rotates with Jupiter, every 10 hours, it overcomes the satellite from behind, resulting in particle impacts with the trailing side. The absence of magnetospheric particles may reveal new information on the composition of the surface of Europa.
The fields and particles instruments return a portion of a high time resolution recording which measures dust, plasma, and electromagnetic fields near Europa. These data will expand the available knowledge of how Europa interacts with Jupiter's magnetosphere.
Finally, results from last week's gyroscope performance test are in. They indicate a minor improvement in the gyroscope performance. However, flight engineers believe that the improvement is too small to warrant any adjustments to the parameters in the attitude control computer's software. The software was modified in March to allow the computer to correctly interpret and use the anomalous gyroscope data. The gyroscope has been producing anomalous data since December 1997.
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"Galileo Update (13 - 19 July, 1998)",603,0,0,0
Galileo spends this week preparing for its next encounter with Jupiter and its moons. The encounter features the fifth close flyby of Jupiter's icy moon Europa since the start of the Galileo Europa Mission. Encounter activities are scheduled to start late Sunday night, July 19 (Pacific Time).
Before encounter begins, Galileo must complete the playback of data written to its tape recorder during the spacecraft's previous passage through the heart of the Jupiter system in early June. All of this data will be overwritten by new science observations during the upcoming encounter.
Preparations for the next encounter include, regular maintenance on the spacecraft's onboard tape recorder and propulsion system, and the spacecraft's anomalous gyroscope is tested to keep track of how it is performing. Finally, the spacecraft executes any final flight path corrections required prior to the close Europa flyby.
This week's playback activities are designed to fill gaps in previously returned data, select entirely new data, or re-process data with different parameters. Science data from Europa, Jupiter and Io are returned this week.
Regional and global scale information on Europa is returned by the near infrared spectrometer, and a mosaic of unexplored terrain near Europa's terminator is returned by the spacecraft camera. Three other observations returned by the near infrared mapping spectrometer were taken 16, 20, and 22 hours after the closest approach to Europa in early June. The spectrometer's 16, 20, and 22 hour observations will provide high resolution spectra that is low in noise due to the fact that radiation noise decreases roughly in proportion to the spacecraft's distance from Jupiter. These spectra will be used to help identify non-ice components of the surface.
The spacecraft camera returns two observations of Io taken while the moon was eclipsed from the sun by Jupiter. The third and fourth in a series, they are designed to allow scientists to study the changes in Io's surface temperature as the eclipse progresses. Toward the end of the week, the near infrared spectrometer returns portions of two observations of Io's surface. Included in these data is information that may lead to the discovery of a newly formed hot spot on Io.
Finally, three observations of Jupiter, performed by the near infrared spectrometer, are returned this week. All three observations will provide information on the differences in temperature and composition that can be found across Jupiter's cloud belts and cloud zones.
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"Galileo Update (27 July - 2 August 1998)",604,0,0,0
Having recovered from last week's anomaly, Galileo is expected to return to normal operations this week. Last week's encounter was cut short when Galileo's main computer placed the spacecraft in safe mode after experiencing an erroneous power reset signal. An anomaly of this type has occured nine times in Galileo's past, but not since 1993.
The main computer is expected to start executing a revised cruise command sequence on Wednesday. Shortly after, a post-encounter test will be performed to keep track of changes to the performance of the spacecraft's anomalous gyroscope. On Friday, the spacecraft will execute commands to fire its thrusters and correct any errors to its orbital flight path.
Playback of data stored on the spacecraft's on-board tape recorder is expected to start toward the end of the week. Several observations were recorded successfully prior to last week's anomaly and are on this week's playback schedule.
The first observation on the schedule belongs to the near-infrared mapping spectrometer and looks for changes in Io's surface, in particular within the Prometheus volcano region. The photopolarimeter radiometer also returns an observation of Io designed to provide information on how rapidly temperatures change on Io's surface.
Observations of Jupiter's atmosphere containing data describing hot spots, white ovals and atmospheric temperature variations are also on this week's schedule. The hot spot information is contained in a single observations performed by the near-infrared mapping spectrometer. The white oval and temperature variation data are returned in one and two observations, respectively, all performed by the photopolarimeter radiometer.
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"Galileo Update (29 July, 1998)",605,0,0,0
The Galileo spacecraft has resumed transmitting science data to Earth in real time. Last night, engineers uplinked command sequences that should enable the spacecraft to resume playing back recorded science information as of 7:15 tomorrow morning, Pacific time. This will restore complete functioning of the spacecraft's science operations, which were disrupted last week when the spacecraft put itself into "safing" mode because of an anomaly.
The anomaly was caused by multiple resets, triggered when debris shorted a signal line in one of the spacecraft's two onboard command and data subsystems. The two subsystems receive commands from Earth and transmit information to the ground. Because the anomaly occurred during a flyby of Jupiter's moon Europa, nearly all data from that encounter were lost. However, Galileo scientists expect to receive some remnants of the data that were recorded, including observations of the volcanic moon Io made by the spacecraft's near infrared mapping spectrometer.
The spacecraft is undergoing a gyroscope performance test today and engineers will analyze the results through the evening. This is in preparation for Friday morning's scheduled flight path correction maneuver. This will ensure that Galileo is aimed correctly as it heads toward another Europa flyby on September 26, with two more Europa encounters scheduled for the coming months.
During the past 2-1/2 years, the spacecraft has flown by Europa eight times, gathering intriguing pictures and science information. Galileo successfully completed its two-year primary mission in December 1997, and is in the midst of a two-year extension, known as the Galileo Europa Mission.
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"Sun's Magnetic Field",606,0,0,0
Magnetic field lines loop through the solar atmosphere and interior to form a complicated web of magnetic structures. Many of these structures are visible in the chromosphere and corona, the outermost layers of the Sun's atmosphere. However, we measure the magnetic field itself in the photosphere, the innermost layer of the Sun's atmosphere. A variety of techniques are then used to mathematically map these magnetic field measurements into the outer layers where they can be compared with the observed structures.
An outstanding problem is our lack of understanding about how the Sun chooses different field lines to populate with hot coronal plasma. MSFC Solar Astronomer Dr. Allen Gary has devised a program that wraps each selected field line with a tube of plasma and then compares an image containing that tube with the X-ray observations to determine the density of the plasma. The set of tubes that best match these observations are shown here.
The three-dimensionality of these loops is better illustrated by the movie in which the loop system is rotated in a manner similar to how the Sun would rotate it through a series of observations over a two week period.
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"Mars Surveyor 2001",607,0,0,0
The Mars Surveyor 2001 missions will follow two other robotic Mars missions to be launched in late 1998 and early 1999. All are part of NASA's long-term, systematic exploration of Mars in which two missions are launched to the planet approximately every 26 months. NASA's Office of Space Science has selected the following investigations for the Mars 2001 Orbiter, due for launch in March of that year, and the Mars 2001 Lander/Rover, due for launch in April:
The two robotic spacecraft scheduled for launch in mid-2001 to orbit and land on Mars will carry a descent camera, a multispectral imager, and a robotic rover capable of traversing tens of miles across the red planet's rocky highlands.
The Mars 2001 Thermal Emission Imaging System (THEMIS) will map the mineralogy and morphology of the Martian surface using a high resolution camera and a thermal infrared imaging spectrometer. Dr. Phil Christensen from Arizona State University in Tempe is the principal investigator for THEMIS.
The Mars 2001 Lander will carry a small, advanced technology rover capable of traveling several tens of miles across the Martian highlands. The rover will be slightly larger than the Pathfinder Sojourner rover and will be designed to go farther (100 km vs. 100 m for Sojourner) and to last longer (1 year vs. 7 days for Sojourner).
The rover will carry a payload called Athena, which is an integrated suite of instruments which will conduct in-situ scientific analyses of surface materials. It also will be able to collect and analyze core samples for later return to Earth by a future robotic mission. Dr. Steven Squyres from Cornell University, Ithaca, NY, is the principal investigator for Athena.
The 2001 Orbiter will be the first to use the atmosphere of Mars to slow down and directly capture a spacecraft into orbit in one step, using a technique called aerocapture. The 2001 Lander will carry an imager to take pictures of the surrounding terrain during the lander's rocket-assisted descent to the surface.
The descent imaging camera will provide images of the landing site for geologic analyses, and will aid planning for initial operations and traverses by the Athena rover. Dr. Michael Malin of Malin Space Science Systems Inc. in San Diego, CA, is the team leader for the Descent Imager science team and Dr. Ken Herkenhoff of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA is a team member.
The Orbiter also will carry the Gamma Ray Spectrometer (GRS), the last of the remaining Mars Observer science investigations. The GRS will achieve global mapping of the elemental composition of the surface and the abundance of hydrogen in the shallow subsurface.
An integrated team consisting of the Jet Propulsion Laboratory and Lockheed Martin Astronautics, Denver, CO, will develop the missions, led by JPL. Both of the 2001 missions are part of an ongoing NASA series of robotic Mars exploration spacecraft that began with the launches of the Mars Global Surveyor orbiter and the Mars Pathfinder lander in November and December 1996, respectively.
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"Mars Surveyor Report (July 1997)",608,0,0,0
Friday, July 25, 1997
Mars Global Surveyor is rapidly approaching Mars, now just 48 days away from entering orbit around this planetary neighbor. Today the orbiter is about 11 million kilometers (7 million miles) from Mars, traveling at a speed of about 21.7 kilometers per second (about 49,000 miles per hour) with respect to the Sun.
The spacecraft and flight teams recently conducted a live simulation of spacecraft telecommunications procedures with one of NASA's Deep Space Network tracking stations in Madrid, Spain, to prepare for the beginning of mapping operations.
During mapping, Surveyor will circle Mars once every two hours. On each orbit, the spacecraft will pass behind Mars and will not be able to maintain a communications link with Earth. Upon entry and exit into this occultation zone, Surveyor's radio signal will pass through the thin Martian atmosphere on its way to Earth. An analysis of the variation of the signal's strength and frequency caused by the atmosphere as it fades and reappears will enable scientists to determine atmospheric properties at specific locations on Mars.
The recent simulation was designed to allow the flight and tracking teams to become accustomed to the spacecraft's performance as its signal repeatedly passes in and out of the Martian atmosphere. During the simulation, the spacecraft turned its radio transmitter on and off over the course of a six-hour period in order to simulate three orbits worth of communications events. Participants found the simulation beneficial and have planned several more of these dry runs over the next two months.
Mars Global Surveyor will reach the orbit of Mars on Sept. 11, 1997, then spend six months aerobraking through the thin Martian atmosphere to achieve a low-altitude mapping orbit. Mapping and science operations will begin on March 15, 1998.
After 260 days in flight, all systems onboard Mars Global Surveyor remain in excellent condition. The spacecraft is on target to intercept Mars at 6:39 p.m. Pacific Daylight Time on Sept. 11 (1:39 a.m., Sept. 12 in Universal Time).
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"Mars Surveyor Report (September 1997)",609,0,0,0
September 11, 1997
8 p.m. Pacific Daylight Time\b
NASA's Mars Global Surveyor , the first in a series of orbiters and landers to explore Mars in the next decade, performed a critical engine burn this evening and successfully entered orbit around the red planet.
The spacecraft executed a 22-minute engine burn at 01:17 Universal Time (6:17 p.m. Pacific Daylight Time) and slowed its speed by more than 3,200 kilometers per hour (2,000 miles per hour) to be captured in Martian orbit. (Because radio signals traveling at the speed of light take 14 minutes, 6 seconds to reach Earth from Mars, the start of the engine burn was detected by ground controllers at 6:31 p.m. PDT.)
Doppler data after the burn indicated that the spacecraft is now in a highly elliptical orbit which takes it within about 250 kilometers (155 miles) of Mars at its closest point and about 56,000 kilometers (34,800 miles) at the most distant.
Global Surveyor is the first U.S. spacecraft to orbit Mars in more than 20 years and the first to use aerobraking rather than propulsive maneuvers to adjust its orbit upon arrival. The technique was demonstrated during the final months of the Magellan mission to Venus in the summer of 1993, and found to be a plausible design for circularizing a spacecraft's orbit while saving fuel.
Preparations for this evening's orbit insertion burn began two days ago, on September 9, when Global Surveyor pressurized its propellant tanks. Pressurization occurred at 9:15 a.m. PDT and was nearly instantaneous. The propellant tanks reached the required value of 18.6 bars (270 pounds per square inch), which was necessary to perform the capture burn tonight.
Today's activities began about 17 minutes before the orbit insertion burn, when Mars Global Surveyor powered up its small 2.3-kilogram (5-pound) thrusters and prepared to turn its main engine in the direction of the spacecraft's motion, or toward Mars. After this reorientation, and at an altitude of 1,490 kilometers (926 miles) above Mars, the spacecraft fired its 660-newton main engine for approximately 22 minutes, 39 seconds.
Twelves minutes after the start of the engine burn, the spacecraft passed behind Mars as seen from Earth and was temporarily blocked from communications with Earth. This occultation lasted 12 minutes. The spacecraft emerged from behind Mars nearly four minutes after the engine burn had been completed.
Two of NASA's Deep Space Network tracking facilities at Goldstone, CA and Canberra, Australia, were monitoring the spacecraft's orbit insertion burn and reacquired Mars Global Surveyor's signal within seconds of its reappearance from the back side of Mars at 6:57 p.m. PDT.
Global Surveyor has arrived at Mars during fall in the northern hemisphere and spring in the southern hemisphere, which usually coincides with the start of the dust storm season. Although dust storms are a concern for the navigation team, the spacecraft will be able to raise its orbit and fly over these storms if it becomes necessary.
Data from the surface of Mars, furnished by the highly successful Mars Pathfinder lander and rover mission, in addition to data from the orbiting Hubble Space Telescope and the National Radio Astronomy Observatory microwave antenna in Boulder, CO, will assist the Mars Global Surveyor team as they begin to dive into the upper atmosphere and understand the dynamics of the Martian environment.
Mars Global Surveyor will complete three revolutions around Mars in its initial, highly elliptical orbit, and gather some science data. On September 17, the spacecraft will perform its first aerobraking maneuver.
Each time the spacecraft reaches the farthest point in its orbit around Mars -- known as the apoapsis -- it will perform an engine burn to trim the orbit. After four engine burns at apoapsis -- on September 17, 20, 22 and 24 -- Global Surveyor's orbit will be reduced to about three hours, meaning the spacecraft will be completing one revolution around Mars about every three hours.
During the next three months, the navigation team will continue to fine-tune the spacecraft's apoapsis and periapsis, or farthest and closest points over Mars, respectively. In January 1998, the navigation team will begin three weeks of final orbital adjustments. Then the science instruments will be turned on around March 10 and the mapping mission will begin on March 15, 1998.
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"Mars Surveyor 98 Update",610,0,0,0
\JMars Surveyor 98 Introduction\j
\JMars '98: Next Generation\j
\JMars Surveyor Project Status Report 1998\j
\JMars '98: Landing Zone Images Reveal Strange New Terrain\j
\JMars '98: Scientists View First Close-Ups of Strange, Layered Polar Terrain\j
\JMars '98: Global Surveyor Ready To Image 'The City'\j
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"Mars Surveyor 98 Introduction",611,0,0,0
The Mars Surveyor '98 program is the next generation of spacecraft to be sent to Mars. Consisting of an orbiter and lander launched separately on Med-Lite launch vehicles (Delta 7425 configuration), the Mars '98 mission will add to the knowledge gained by the Mars Global Surveyor and Mars Pathfinder missions.
The general science theme for the 1998 Surveyor missions is "Volatiles and Climate History." The Mars 98 orbiter will launch in December 1998 and arrive at Mars 10 months later. Upon arrival at Mars, the spacecraft will use a series of aerobraking maneuvers to achieve a stable orbit, and then use atmospheric instruments and cameras to provide detailed information about the surface and climate of Mars.
The Mars 98 lander will launch a month after the orbiter and will land near the southern polar cap on Mars. The lander is equipped with cameras, a robotics arms and instruments to measure the Martian soil composition. Two small microprobes are also piggybacking on the lander, which will penetrate into the Martian subsurface to detect water ice.
The lander and orbiter missions are both managed by the 1998 Mars Surveyor Project at the Jet Propulsion Laboratory for NASA's Office of Space Science. Lockheed Martin Astronautics in Denver, Colorado, is NASA's industrial partner in the mission.
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"Mars '98: Next Generation",612,0,0,0
January 9, 1998
The Mars Surveyor '98 program is the next generation of spacecraft to be sent to Mars. Consisting of an orbiter--to be launched Dec. 10, 1998, and lander, set for launch on Jan. 3, 1999--the Mars '98 mission will add to the knowledge gained by the Global Surveyor and Pathfinder missions. The general science theme for the 1998 Surveyor missions is "Volatiles and Climate History."
The Mars '98 orbiter will arrive at Mars Sept. 23, 1999, while the lander will touch down Dec. 3, 1999.
Upon arrival at Mars, the spacecraft will use a series of aerobraking maneuvers to achieve a stable orbit, and then use atmospheric instruments and cameras to provide detailed information about the surface and climate of Mars.
The '98 orbiter mission will carry a rebuilt version of the Mars Observer Pressure Modulated Infrared Radiometer (PMIRR), as well as the Mars color imaging system. PMIRR will observe the global distribution and time variation of temperature, pressure, dust, water vapor and condensates in the Martian atmosphere.
The imaging system will observe synoptically Martian atmospheric processes at global scale and study details of the interaction of the atmosphere with the surface at a variety of scales in both space and time. In addition to the science payload, the orbiter spacecraft will provide an on-orbit data relay capability for future U.S. and/or international surface stations.
The lander will land near the southern polar cap and is equipped with cameras, a robotics arm and instruments to measure the composition of the Martian soil. Two small microprobes are also piggybacking on the lander, which will penetrate into the Martian subsurface to detect water ice.
The science package for the lander includes the Mars Volatile and Climate Surveyor (MVACS) integrated lander payload, the Mars Descent Imager (MARDI) and an atmospheric lidar experiment provided by the Russian Space Agency Institute for Space Science.
The integrated lander payload includes a surface stereo imager with Mars Pathfinder heritage; a meteorology package; an instrumented robotic arm for sample acquisition, soil manipulation and closeup imaging of the surface and subsurface; and the thermal and evolved gas analysis experiment for determining the nature and abundance of volatile material in the Martian soil.
The images obtained while the lander descends to the surface will establish the geological and physical context of the landing site. The atmospheric lidar experiment will determine the dust content of the Martian atmosphere above the landing site.
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"Mars Surveyor Project Status Report 1998",613,0,0,0
February 1, 1998
The 1998 Mars Surveyor orbiter is 312 days from launch and the lander is 335 days from launch (as of 2/1/98). The following provides a summary description of the current Project status as of February 1, 1998.
\BOrbiter Integration and Test\b
The orbiter spacecraft has been fully assembled and checked out functionally. Major tests successfully completed include the Deep Space Network (DSN) Compatibility Test, the Bus Functional Test (BFT), Mission System Test 1 (MST 1), and the modal survey. Descriptions of the Bus Functional Test and the Mission System Test are included below.
The Pressure Modulator InfraRed Radiometer (PMIRR) flight instrument is integrated and all interfaces and functionality have been verified. The Mars Color Imager (MARCI) electrical interface unit (i.e., electrical equivalent camera less optics) is integrated and interfaces are verified. The flight camera is on schedule for integration in February prior to thermal vacuum testing.
The refurbishment of the orbiter flight processor and power distribution and drive unit ATLO Test Units is complete and all orbiter electronics are in their final flight configuration. The orbiter is in the Reverberant Acoustics Lab (RAL) at Lockheed Martin in Denver and is being prepared for acoustic testing.
The schedule maintains 71 days of margin prior to shipment and 20 days of margin at KSC.
Significant mass margin for the orbiter exists with essentially all elements weighed. Recent orbit insertion analyses indicate that by utilizing just a portion of the expected margin at launch to further fill the hydrazine and ox tanks will allow the orbiter to insert directly into a 15 hour or shorter period orbit. This provides significant margin against aerobraking uncertainties and increases the probability that the orbiter will be in place to support the lander at lander arrival.
\BLander Integration and Test\b
The lander spacecraft is fully integrated electrically and all interfaces have been verified. Major tests successfully completed include the Deep Space Network (DSN) Compatibility Test and the Bus Functional Test (BFT). In addition, the lander to orbiter UHF relay was tested successfully using both spacecraft in the high bay at LMA. Installation of the lander science payload is in process.
The flight Lidar, Meteorology package, Surface Stereo Imager, Robotic Arm, and Robotic Arm Camera, have been installed on the spacecraft and functional testing is in progress. The flight Mars Descent Imager will be installed next week and the flight Thermal and Evolved Gas Analyzer experiment will be installed in June prior to landed configuration Thermal Balance testing. The lander is in the high bay facility at Lockheed Martin and will be moved to the Reverberant Acoustics Lab in March for the start of environmental testing.
The current schedule for the lander is:
3/21/98 - Start Acoustic Test
3/31/98 - Start EMI/EMC Tests
5/4/98 - Start Cruise Thermal Vacuum Test
6/23/98 - Start Landed Thermal Balance Test
10/14/98 - Ship lander to KSC
1/3/99 - Launch
The current lander schedule maintains 41 days of margin prior to shipment with 20 days of margin at KSC.
Significant mass margin for the lander exists with full tanks and most elements weighed. In all likelihood the lander will be well below the maximum design mass providing significant margin against launch vehicle performance shortfalls, entry heating limits, parachute deployment limits, and landing site elevation uncertainties.
\BBus Functional Test\b
The purpose of the Bus Functional Test (BFT) is to verify the end-to-end core command and control functionality in an integrated vehicle configuration. The major parts of the test include complete testing of uplink and downlink command and data capability (data rates, file sizes, etc.), non-volatile memory read and write functionality, and test of the Command and Data Handling Module Interface Card (C-MIC) functions which control heartbeat, fault recovery, and vehicle state.
Also, end-to-end phasing of the Attitude Control Subsystem from sensor inputs through actuator commands is tested in a quasi-closed loop manner. The BFT is conducted on both A&B sides of the hardware. Some miscellaneous functions also will be verified as part of the BFT, including voltage drop verification throughout the longest harness cable run, launch event detection logic, and critical command aliveness functions to be used in later tests. The BFT represents a major milestone in the integration and test phase of development.
\BMission System Test\b
The Mission System Tests (MST), which consist of the System Aliveness Test (SAT), the Mission Profile Test (MPT), and Sequence Verification Test (SVT), are run before and after each environmental test and after the spacecraft is shipped to KSC. The SAT is a standard, multiple use test that will be run as the first part of each MST.
Its purpose is to demonstrate continuity of the electrical signal and power paths, meaning that all components are operational and able to receive commands and send telemetry. The significance of this first SAT is in demonstrating that the spacecraft has achieved a fully integrated system of software and hardware and to establish the baseline data for subsequent runs of the SAT.
Unlike earlier subsystem level tests on the spacecraft, this end to end test was performed without the aid of ground support equipment and thereby establishes a higher level of verification of the hardware / software interfaces. Orbiter MST 1 also includes a limited mission profile test (MPT).
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"Mars '98: Landing Zone Images Reveal Strange New Terrain",614,0,0,0
February 10, 1998
New images of the Mars '98 landing zone, located in the frigid south polar region of Mars, reveal the first close-up views of this strange, layered terrain, which will become the site of NASA's 1998 Mars Polar Lander mission.
Ground fog obscures the surface in the raw images, while image-processing techniques, expose ground features in great detail. The images show an array of light and dark mottled patterns, and swirling bands of eroded, layered rock reminiscent of the edges of Alaskan ice sheets.
The Mars Polar Lander is one of two spacecraft to be launched to Mars in late 1998 and early 1999 to continue NASA's program of robotic exploration of the red planet. The 1998 lander will touch down in the southern polar region of Mars in December 1999 and begin a three-month surface mission to dig for traces of subsurface water. Meanwhile, its sister ship, the Mars Climate Orbiter, will begin a two-year mapping mission to profile the Martian atmosphere and map the surface.
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"Mars '98: Scientists View First Close-Ups of Strange, Layered Polar Terrain",615,0,0,0
February 11, 1998
Swirling bands of eroded, layered rock, reminiscent of the edges of Alaskan ice sheets, and an array of light and dark mottled patterns blanket the frigid floor of Mars' south pole, where NASA's newly named Mars Polar Lander will touch down in late 1999.
The new images of the landing zone for the Mars Polar Lander, taken by the camera aboard NASA's Mars Global Surveyor, confirm that this strange, layered terrain in the south polar region represents a dramatic departure from the now-familiar Martian landscapes observed by the Viking landers and Mars Pathfinder. In December 1999, the next lander in a steady series begun by Pathfinder will set down in this uncharted territory to dig for traces of frozen, subsurface water.
"Despite ground fog that obscures part of the surface in these images, we can see much more surface detail than we've ever seen before, which suggests that the 75-degree south latitude landing zone is quite a bit more rugged and geologically diverse than we had previously thought," said Dr. Michael Malin of Malin Space Science Systems, Inc., San Diego, CA. Malin is principal investigator of the Global Surveyor camera and the cameras on the 1998 missions, the Mars Polar Lander and its newly named partner, the Mars Climate Orbiter.
In the current images from Mars Global Surveyor, obtained during an aerobraking orbit from about 2,800 kilometers (1,700 miles) above the planet's surface, objects about 15 meters (48 feet) across can be resolved.
Once the spacecraft reaches its final mapping orbit early next year, at an average of 378 kilometers (234 miles) above the surface, the camera will be able to resolve ground features as small as 2 to 3 meters (7 to 9 feet) across. This greater clarity will enable views of objects as small as boulders or as subtle as sand dunes.
Over the next year, the Global Surveyor images will be used in concert with other spacecraft data such as that obtained by the thermal emission spectrometer to better characterize the geology of Martian south pole.
After Global Surveyor has reached its mapping orbit, data from the spacecraft's laser altimeter, which measures the height and roughness of Martian surface features, will be combined with imaging data to aid the final choice of landing sites.
"We have a wonderful opportunity in the next year to study this region with data from Mars Global Surveyor, which underscores the true advantage of conducting a continuing program of Mars exploration," said Dr. John McNamee, Mars Surveyor '98 project manager at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
"We will be able to characterize the geology of the whole region and find the best spot to land, one that presents a balance between lander safety and scientific interest. This process does not have to be finalized until June 1999, five months after the lander has been launched and six months before it lands."
The images are being studied while the 1998 Mars Climate Orbiter and Mars Polar Lander are undergoing key hardware integration and testing at Lockheed Martin Astronautics, Denver, CO. The spacecraft are currently being prepared for transfer to the Lockheed Martin environmental test chambers to ensure that they can survive and operate in the extreme conditions at Mars. At the completion of this testing, the spacecraft will be flown separately to NASA's Kennedy Space Center, FL, for integration with their launch vehicles.
The 1998 Mars lander and orbiter missions are designed to learn more about the history of Mars' climate and the behavior of related Martian volatiles, such as water vapor and ground ice. The orbiter, scheduled for launch on Dec. 10, will conduct a two- year primary mission to profile the Martian atmosphere and map its surface.
The lander, scheduled for liftoff on Jan. 3, 1999, will carry out a three-month mission to search for traces of subsurface water in this frozen, layered terrain and any evidence of a physical record of climate change.
To meet these scientific objectives, the orbiter will carry a rebuilt version of the Pressure Modulated Infrared Radiometer (PMIRR) that was lost with Mars Observer in 1993. This atmospheric sounder will observe the global distribution and time variation of temperature, dust, water vapor and condensates in the Martian atmosphere. PMIRR is a collaboration between JPL, Oxford University and Russia's Space Research Institute.
Like Mars Global Surveyor, the Mars Climate Orbiter carries a dual camera system, but this one is contained in an amazingly compact package about the size of a pair of binoculars. The Mars color imager's 0.5-kilogram (1 pound) wide-angle camera will return daily low-resolution global views of the planet's atmosphere and surface, while its medium-angle camera will provide higher resolution (40 meters or 30 feet per pixel) images.
The medium-angle camera will build global and regional maps of Mars in multiple colors over the course of the mission. These maps will be used to characterize surface properties and changes in the distribution of dust.
The 1998 lander carries three scientific packages: the Mars descent imager, again provided by Malin, which will view the landing site at increasingly higher resolution as the lander descends to the surface of Mars; the atmospheric lidar experiment, provided by the Russia space institute, which will monitor the presence and height of atmospheric hazes, coupled with a miniature microphone furnished by The Planetary Society, Pasadena, CA, to record the sounds of Mars; and the Mars Volatile and Climate Surveyor (MVACS) package, led by principal investigator Dr. David Paige of the University of California, Los Angeles.
MVACS includes a surface stereo imager based on the Mars Pathfinder camera, both built at the University of Arizona; a meteorology package, built at JPL; a robotic arm, also built at JPL, to acquire soil samples and close-up images of the surface and subsurface; and the thermal and evolved gas analysis experiment, built at the University of Arizona. JPL will oversee mission operations with the spacecraft team at Lockheed Martin Astronautics and the instrument teams located at their home institutions during the lander and orbiter missions.
"MVACS and the other science experiments are tailor-made for the exploration of Mars' south pole," said Dr. Richard Zurek, project scientist at JPL. "The robotic arm, which is reminiscent of the Viking arm and scoop that were used to carry out biology experiments in the mid-1970s, is, in fact, much more versatile. It can reach farther out, dig up to 1 meter (3 feet) below the surface and then place soil samples in a miniature oven, called the evolved gas analysis experiment, where the samples are 'cooked' and analyzed for chemical and gas content."
Piggybacking on the Mars Polar Lander are two small, 2- kilogram (4.5-pound) microprobes, provided by NASA's New Millennium validation program. Deployed before landing, they will penetrate and embed themselves beneath the Martian surface to study subsurface materials.
The Mars Polar Lander and Mars Climate Orbiter are the second set of launches in a long-term NASA program of Mars exploration known as the Mars Surveyor Program. The missions are managed by JPL, Pasadena, CA, for NASA's Office of Space Science, Washington, DC. Lockheed Martin Astronautics, Denver, CO, is NASA's industrial partner in the mission.
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"Mars '98: Global Surveyor Ready To Image 'The City'",616,0,0,0
April 10, 1998
The Mars Global Surveyor operations team is gearing up to begin imaging a second set of specifically targeted geologic features on Mars, after completing the first set of images last week and successfully capturing the so-called "Face on Mars."
At the direction of NASA Administrator Daniel Goldin, the flight team has developed a schedule of new targets. On Tuesday, April 14, Global Surveyor will image a second portion of the Cydonia region known as "The City." This area of Cydonia contains geological features that have been referred to as "mounds," a "city square," "pyramids" and "the fortress." The spacecraft's high-resolution camera will use the "city square" portion of this geologic formation as the target point.
Last week's attempts to image the landing sites of the Viking 1, Viking 2 and Mars Pathfinder landers were unsuccessful. Global Surveyor will make new attempts to image the Viking sites on two consecutive orbits on Sunday, April 12. On Monday, April 13, the spacecraft will image the Mars Pathfinder landing site, using refined coordinates obtained during the first attempt.
Winter weather in the northern hemisphere of Mars was a significant factor in preventing a view of the landing sites during the first series of attempts. The site of the Viking Lander 1 in Chryse Planitia, for instance, was covered with a thick cloud layer, which reduced but did not eliminate surface visibility. However, data showed that the spacecraft's pointing was off just enough to miss that target.
The spacecraft was able to target the Viking 2 lander site in Utopia Planitia, which is farther north and on the other side of Mars from Viking 1. However, this area was heavily overcast with clouds and haze which reduced surface visibility by 70 to 80 percent and rendered the image unusable. The spacecraft missed the Mars Pathfinder site due to the inaccuracy of landing site coordinates.
The project team estimates that Global Surveyor has about a 30 to 50 percent of imaging each target on a given attempt, due to navigation uncertainties and spacecraft performance.
A third and final set of high-resolution imaging of the Viking, Pathfinder and Cydonia regions will be attempted on April 21-23.
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"Mars 98 Lander Introduction",617,0,0,0
The lander spacecraft will launch from the Cape Canaveral Air Force Station (CCAFS) Space Launch Complex 17 (SLC-17) during a 25 day launch period beginning on January 3, 1999. The second pad at SLC-17 will be used for the lander launch to enable the quick two week turnaround from the end of the orbiter launch period.
The lander will enter the Martian atmosphere directly from the hyperbolic transfer orbit at 7 km/s in December 1999. The lander spacecraft will decelerate to a soft landing using a heat shield to aerobrake, a parachute, and actively guided propulsion to reduce vertical velocity to less than 2.4 m/s and horizontal velocity to less than 1 m/s at surface touchdown.
The lander will be targeted to the northernmost boundary of the polar layered deposits at a high southern latitude site, between 75 degrees and 80 degrees south latitude. The surface science mission will be conducted over the course of a 3 month primary mission. The landing will occur during late spring in the southern hemisphere and extend through the early summer season. The timing of the landing is optimal for a high southern latitude site because the sun is always above the horizon during the course of the primary mission providing maximum solar insolation and a relatively benign thermal environment.
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"Mars 98 Lander Mission Overview",618,0,0,0
The MSP 98 Lander will be launched on a Delta 7425 in January 1999, and will arrive at Mars in December 1999. Burnout of the 3rd stage will be followed by yo-yo despin of the entire stack, followed by spacecraft separation. At this point both the spacecraft and upper stage will have been injected onto a Type 2 trajectory whose aimpoint is biased away from the nominal entry aimpoint, to assure that the upper stage has less than a 1E-4 probabilty of impacting Mars, as required by Planetary Protection regulations.
After separation, the solar panels will be deployed and pointed to the sun, and initial acquisition achieved by the DSN. Throughout cruise, contact will be maintained via the Medium Gain Antenna, and the solar panels pointed at the sun [with a small offset in inner cruise].
Approximately 15 days after launch, the largest Trajectory Correction Maneuver [TCM-] will be executed. This maneuver will remove launch vehicle injection errors and the spacecraft's injection aimpoint bias. Depending on the size of the maneuver, it may be necessary to divide this into two smaller maneuvers.
Provisions have been made to execute up to 4 additional small TCM's during the remainder of cruise, including one 7 hours prior to entry for final control of the entry angle and landing footprint. Precision approach navigation will be effected via near simultaneous tracking of the approaching Lander and an orbiter at Mars (either the MSP98 Orbiter or Mars Global Surveyor).
After a direct atmospheric entry, the Lander will be slowed by a Mars Pathfinder-heritage aeroshell and parachute, and a controlled propulsive landing effected. For launch during the Lander's primary launch period, landing will occur between 75 and 78S, on the southern polar layered terrain. The first landed day's activities will include deployment of the solar panels, functional checkout, and establishment of communication with the Orbiter and time critical science activities.
Routine science activities will commence on the second day following landing. The Lander will be equipped with a UHF relay for downlink via the MSP98 Orbiter and/or MGS, and command uplink via the MSP98 Orbiter. A direct to Earth (DTE) link will also be available for Lander commanding and as a backup downlink.
The Lander will carry the Mars Volatiles and Climate Surveyor (MVACS) instrument suite, which will perform in situ investigations to address the science theme "Volatiles and Climate History" on Mars. The Lander will also provide descent imaging with the Mars Descent Imager (MARDI), and accommodate a LIDAR instrument supplied by the Russian Space Agency.
The Lander will search for near-surface ice and possible surficial records of cyclic climate change, and characterize physical processes key to the seasonal cycles of water, carbon dioxide and of dust on Mars. The duration of the landed science phase is expected to last no more than approximately 90 days.
#
"Mars 98 Orbiter Introduction",619,0,0,0
The orbiter spacecraft will launch from the Cape Canaveral Air Force Station (CCAFS) Space Launch Complex 17 (SLC-17) during a 14 day launch period beginning on December 10, 1998. The Mars Orbit Insertion (MOI) propulsive maneuver will occur in September 1999 and will place the orbiter into a highly elliptical, near polar orbit around Mars.
Peripapse will be lowered to approximately 110 km altitude to initiate the aerobraking maneuvers. Successive passes of the orbiter through the upper atmosphere of Mars will slow the vehicle and lower the apoapse of the orbit to 450 km over the course of the 2 month aerobraking phase.
The orbit then will be circularized using the orbiter's onboard propulsion resulting in the design 400 km altitude, near circular, polar science mapping orbit. Science operation of the PMIRR and MARCI instruments will be conducted over the course of the one Martian year (687 Earth day) mapping mission. The orbiter will continue operations in a relay only mode following the science mission in support of any future U.S. or international Mars surface missions.
#
"Mars 98 Orbiter Mission Overview",620,0,0,0
(Last updated 05 Jan 1998)
The MSP 98 Orbiter launches aboard a Delta 7425 in December 1998, and arrives at Mars in September 1999. Burnout of the 3rd stage is followed by yo- yo despin of the entire stack, followed by spacecraft separation.
At this point both the spacecraft and upper stage have been injected onto a Type 2 trajectory whose aimpoint is biased away from the nominal Mars Orbit Insertion (MOI) aimpoint, to assure that the upper stage has less than a 1E- 4 probability of impacting Mars, as required by Planetary Protection regulations.
After separation, the solar panels are deployed and pointed to the sun, and initial acquisition achieved by the Deep Space Network (DSN). During inner cruise, the solar panel is sun pointed, and contact maintained via the Medium Gain Antenna.
Approximately 15 days after launch, the largest Trajectory Correction Maneuver [TCM- 1] is executed. This maneuver removes launch vehicle injection errors and the spacecraft's injection aimpoint bias. Provisions have been made to execute up to 3 additional small TCM's during the remainder of cruise, as needed, to shape the orbit and direct the spacecraft to the proper aimpoint for MOI.
All TCM's are performed with the monopropellant hydrazine thrusters. As the heliocentric distance increases during cruise, communications moves to the High Gain Antenna.
At Mars arrival, the Orbiter bipropellant engines are used to propulsively insert the spacecraft into an elliptical capture orbit. The biprop engines burn for approximately 16 minutes, until all the loaded oxidizer is exhausted. One minute later, an additional maneuver is executed by the Hydrazine thrusters, if needed, to reduce the orbit period further.
Depending on launch date and propellant mass consumed during cruise, the resultant orbit period lies between 19 and 40 hrs, with a nominal periapse altitude of 160 km. A maneuver to lower periapse in preparation for aerobraking occurs at the first apoapse of the final capture orbit.
Over the next two months, the energy of the orbit is reduced via successive passes through the atmosphere of Mars, controlled by small Orbit Trim Maneuvers near apoapse. At aerobrake termination, two maneuvers transfer the Orbiter to its final, frozen, near sun- synchronous mapping orbit, at a descending node of approximately 4 PM. This occurs some time prior to the MSP98 Lander arrival in December, 1999.
During the Lander's surface lifetime, the Orbiter provides command and data relay support, and also engages in a limited amount of orbital science. In its mapping phase, the Orbiter performs systematic daily global sounding and imaging of the Mars atmosphere for approximately one Mars year (687 days).
The nadir- mounted science payload consists of a rebuilt Pressure Modulator Infrared Radiometer (PMIRR), and the Mars Color Imager (MARCI). Once its mapping mission is complete, the Orbiter will be available as a communication relay for future Mars landers for up to 3 additional years. Upon completion of its relay mission, the Orbiter may perform a maneuver or be placed in a low- drag attitude to satisfy Planetary Protection regulations.
#
"Shuttle Columbia 1998 Launch",621,0,0,0
The space shuttle Columbia is scheduled to launch April 16, 1998, at 1:19 p.m. CST from launch pad 39-B at Kennedy Space Center.
The primary payload is Neurolab, which consists of investigations focusing on the effects of microgravity on the nervous system. Specifically, experiments will study the adaptation of the vestibular system and space adaptation syndrome, the adaptation of the central nervous system and the pathways that control the ability to sense location in the absence of gravity, and the effect of microgravity on a developing nervous system.
The Neurolab payload consists of 26 human and nonhuman scientific experiments and associated hardware in a Spacelab long module and the orbiter middeck. The experiment disciplines are primarily involved with life science investigations utilizing human subjects and laboratory animals.
\BSTS-90 (90)\b
Columbia (25)
Pad 39-B (41) (estimated)
90th Shuttle Mission (estimated)
25th Flight OV-102 (estimated)
\BCrew:\b
Richard A. Searfoss (3), Commander
Scott D. Altman (1), Pilot
Richard M. Linnehan DVM (2), Mission Specialist
Dafydd Rhys Williams MD (1) (CSA), Mission Specialist
Kathryn P. Hire (1), Mission Specialist
Dr. Jay C. Buckey (1), Payload Specialist
Dr. James A. Pawelczyk (1), Payload Specialist
Alternate Payload Specialists -
Dr. Alexander W. Dunlap (0), Alternate Payload Specialist
Dr. Chiaki Mukai (1) (NASDA), Alternate Payload Specialist
\BMission Objectives:\b
Neurolab is a Spacelab module mission focusing on the effects of microgravity on the nervous system. The goals of Neurolab are to study basic research questions and to increase the understanding of the mechanisms responsible for neurological and behavioral changes in space. Specifically, experiments will study the adaptation of the vestibular system and space adaptation syndrome, the adaptation of the central nervous system and the pathways which control the ability to sense location in the absence of gravity, and the effect of microgravity on a developing nervous system.
The mission is a joint venture of six space agencies and seven U.S. research agencies. Investigator teams from nine countries will conduct 31 studies in the microgravity environment of space. Other agencies participating in this mission include six institutes of the National Institutes of Health, the National Science Foundation, and the Office of Naval Research, as well as the space agencies of Canada, France, Germany, and Japan, and the European Space Agency.
\BLaunch:\b
Launch April 16, 1998 2:19 p.m. EDT (ESTIMATED). Launch window is 2 hours, 30 minutes.
On Monday, March 16, 1998, Columbia rolled over to the VAB.
On Thursday, March 12, 1998, work on Columbia's right inboard elevon was completed and both forward and aft compartment close-outs were in work. Tomorrow, technicians will complete work to install the aft doors. Aft compartment structural leak tests will follow.
On Monday, 3/2/98, The Neurolab transfer tunnel has been mechanically and electrically mated and the tunnel interface verification test was completed on Friday. Space Shuttle main engine close-out operations and leak checks are complete. Main engine heat shields are being installed and a landing gear functional test is scheduled for today. In the Vehicle Assembly Building, the external tank was successfully mated to the solid rocket boosters on Thursday.
On 2/26/98, work to install the Space Shuttle Main Engines (SSME's) were complete. The STS-90 crew participated in the crew equipment and interface test (CEIT) and a sharp edge inspection of the orbiter's crew module and Neurolab.
On Thursday, 2/5/98, installation of Columbia's airlock hatch "D" was completed. The hatch provides access from the tunnel adapter to the Spacelab transfer tunnel. On Friday, 2/6/98, functional tests on Columbia's airlock hatch "D" and inspections of micrometeorite hits on the orbiter radiator and checklife support system were completed. Aft compartment closeouts and payload premate testing continued. Technicians proceeded with main engine heat shield rework efforts. The Neurolab payload is scheduled to arrive at the OPF on Wednesday (2/11/98) for installation into the orbit's cargo bay and the Shuttle main engines will be installed on Thursday.
On Monday, February 2, 1998, replacement of a relief valve on Columbia's auxiliary power unit No. 2 and water spray boiler checkout were completed. Leak checks on the Spacelab water line are also completed and work on payload bay flood light No. 3 continued. Tunnel adapter flow rate leak testing and airlock ducting reconfiguration were also in work.
On Monday, January 12, 1998 Technicians completed removal of Columbia's window No. 6 yesterday and installed the new window on 1/13/98. Replacement of floodlights No. 1 and No. 5, in the orbiter's payload bay was in work as well as aft flight deck reconfiguration. Fuel cell voltage tests and checks of the flash evaporator system were also scheduled.
On Monday, January 5, 1998, Shuttle Columbia's payload bay doors were opened yesterday to accommodate Ku band antenna testing yesterday afternoon. The orbiter's nose and main landing gear tires were also installed yesterday. Orbiter maneuvering system testing and replacement of fuel cells No. 1 and 3 are in work. Window polishing efforts begin tomorrow and inspections of the forward reaction control system are scheduled for Friday 1/9/98.
Orbiter Thermal Protection System (TPS) Tile damage was found on Columbia following its inspection after STS-87 landed on 12/5/97. 99 tile removals are planned with 40 tile removals under evaluation.
"Shuttle Neurolab Payload: Benefits of Neurolab Science",622,0,0,0
\BAquatic Team\b
Data from the aquatic experiments on Neurolab may disclose the mechanisms at work in various forms of motion sickness experienced by many people on Earth. In particular, studies of the mechanisms controlling otoconial production may provide further insight into what may have gone wrong in certain conditions of human pathology, such as benign pavoxysmal positional vertigo (BPPV), an abnormality involving the otoliths. The studies may also help explain why aging otoliths become smaller.
Further benefits include the use and perfection of the sieve or wafer electrode that is used to record nerve impulses. This electrode offers potential use as a connection to the nervous system in people with deafness caused by hair cell damage. It also could be used as an interface to signal motor prostheses how and when to move.
A quote from Dr. Wiederhold:
"Looking at changes that occur in organisms in microgravity is very practical. Long term space missions and space exploration point toward a time when it will be necessary to begin generating life-including human beings-away from Earth. Studies such as those on Neurolab are, therefore, vital."
\BAutonomic Nervous System Team\b
It is estimated that over one-half million Americans suffer from disorders of autonomic blood pressure control that often result in orthostatic intolerance. Data from the Neurolab studies may provide insight into orthostatic intolerance syndromes, such as postural orthostatic tachycardia syndrome, similar to the phenomenon observed in astronauts after space flight. In addition, therapeutic countermeasures developed from the studies could improve the treatment of affected patients.
A Quote from Dr. Robertson:
"The 500,000 Americans with orthostatic intolerance (OI) have a genuine stake in the outcome of Neurolab. Results from microneurography and norepinephrine spillover will help define the nature of their disorder. It should also bring us closer to the day when OI will yield to therapeutic intervention."
\BMammalian Development Team\b
Just as we have the ability at birth to learn whatever language we are exposed to, we may also have the ability to adapt to whatever gravitation field we experience in early life. These abilities often exist only during a "critical period". Afterward, our potential is limited. Understanding the nature of these critical periods in important in pediatrics.
Knowledge of the critical period for developing normal vision has already changed how strabismus ("lazy eye") is treated in children. The Neurolab mammalian development experiments will expand this knowledge to the control of movement, regulation of blood pressure and maintaining balance. The experiments explore the great potential of our nervous system to adapt to whatever environmental conditions are found at birth. Information from these studies can also be applied to the development of treatments for individuals suffering from childhood neuromuscular diseases, such as muscular dystrophy, or from sustained trauma to their nerves, muscles or spinal cord.
A Quote from Dr. Riley:
"To develop or gain strength in a muscle, a person must exert that muscle against a force. On Earth, gravity presents a natural environment for muscle strengthening or 'loading'. One steps up to a curb, lifts a pail of water, walks, and stands, for example, against the tug of gravity. Muscle loading begins, however, even before birth as the unborn infant begins kicking against its mother's womb. If gravity is required for proper muscle loading, and muscle loading begins in the womb, will mammals which develop in microgravity have normally developed muscles? The mammalian studies on Neurolab will help answer this intriguing question, and at the same time, for example, increase our understanding of the impact of premature birth on the muscle development in infants."
A Quote from Dr. Walton:
"We are really excited about the Neurolab mission because we will find out if 'walking' develops the same way in space as on Earth. The young rats will take their first steps in space."
\BNeurobiology Team\b
How much of normal development is preprogrammed in our genetic code, and how much can be modified by our environment and experiences?
These are two old and puzzling questions.
The Neurobiology team will provide information on the relative importance of the environment and other external stimuli (like gravity) on nervous system development. The gravity system of crickets serves as a model to investigate the general effects of altered gravitational conditions on the development of the structure, function and efficiency of a gravity sensory system. Insects offer useful model systems because their central neurons can be easily identified.
A Quote from Dr. Horn:
"Although each principal investigator has a unique objective, we share a single view of the importance of Neurolab. It is a spectacular, but highly scientific, endeavour that will be of great benefit to humanity. The neurobiology experiments will reveal much about how the nervous system develops in microgravity."
\BNeuronal Plasticity Team\b
Experiments on vestibular adaptation will yield a better understanding of balance disorders, which affect more than 90 million Americans.
Experiments on circadian rhythms could yield valuable data to researchers seeking the causes of jet lag, insomnia and mental disorders such as winter depression. This data is also applicable to aging populations and shift workers, both of whom experience changes in circadian rhythms.
Reconstructing the images from the maculas in Dr. Ross' experiment has led to the development of a Biocomputation Center at the Ames Research Center. The computers there now are also being used to develop vital surgery techniques, a key application and NASA spinoff for the three-dimensional reconstruction software developed there.
A Quote from Dr. Holstein
"The Neurolab mission provides the international scientific community with an extraordinary and unique environment in which to study the cellular bases for learning in the adult brain."
A Quote from Dr. Fuller
"Discovering-through space flight-that the circadian system is sensitive to changes in gravity has been a surprise. Our Neurolab studies were designed to use gravity as a tool to help us unravel the mysteries of the circadian clock. We will use our findings to help assure normal circadian function for our astronauts as well as the rest of us who remain Earth-bound."
\BSensory Motor and Performance Team\b
On Earth, anticipatory mechanisms in individuals with neurological diseases such as Parkinson' disease, basal ganglia disorders, or cerebellar deficiency can be studied using Dr. Berthoz's ball catching experiment. The experiment could also be used for studying motor function development in children. New methods for programming the movements of robots were generated during the development of Dr. Bock's VCF experiment. With further refinement, these methods could improve the capacity of robots to perform complex tasks.
In clinical medicine, the development of new methods for evaluating a patient's ability to use visual and pressure cues for maintaining balance and orientation could be useful. Portable head-mounted virtual reality displays such as the one developed for Dr. Oman's Neurolab experiment may prove useful for patients, perhaps someday providing visual prostheses for individuals with vestibular impairment. The results from the studies could also be applicable to the design of flight simulator and virtual reality vision systems.
A Quote from Dr. Oman:
"On Neurolab, we will be using immersive virtual reality (VR) techniques in space for the first time. The VR helmet gives us complete control of the visual stimulus. The sustained weightless environment of Neurolab provides a truly unique opportunity to understand the interaction between visual and gravireceptive cues in human spatial orientation."
\BSleep Team\b
The Neurolab sleep studies are expected to benefit not only astronauts but Earth-based individuals as well. Whether through an expanded understanding of the physiological effects of melatonin, or the causes of sleep disruption, the studies will be applicable to many groups of individuals with a high incidence of insomnia,
#
"Pathfinder Mission Status Update 97",623,0,0,0
\JPathfinder Status (July97)\j
\JPathfinder Status (August97)\j
\JPathfinder Status (October97)\j
\JPathfinder Status (November97)\j
#
"Pathfinder Status (July97)",624,0,0,0
20 July 1997
1 p.m. Pacific Daylight Time
The Pathfinder mission operations team commanded the lander early this morning and did obtain a carrier signal over the high- gain antenna starting at 3:14 a.m. PDT for the normal period of about 66 minutes, but the signal strength was below expected levels and no scientific data was received.
"This told us that the spacecraft was basically healthy but that there was a problem with the telecommunications link," said Project Manager Brian Muirhead. A later attempt to communicate with the lander through its high-gain antenna from 7:03 to 7:27 a.m. PDT was not successful.
"The flight team is assessing the possible causes of the communication problem, said Muirhead. "This morning's problem may be related to some extent to configuration problems between the spacecraft and the Deep Space Network, but more data is needed to fully assess the problem. We are trying to troubleshoot a problem with very little information," he said.
The flight team is preparing sequences for a low-gain antenna communications session for about midnight tonight (July 20, Pacific Time). A communications session with the high-gain antenna is planned for about 3:30 a.m. PDT tomorrow, July 21.
"Since we have only limited windows to communicate with the spacecraft we must wait patiently for our next opportunity, Muirhead said. "We will go through the usual steps that have worked for us before, and then we will get to the bottom of the problem as we have before." The telecommunications problem is not thought to be related to the reset problem previously experienced by the lander's computer.
The rover remains safely at the rock called Scooby Doo. Earth will rise over the Sagan Memorial Station at 8:07 p.m. PDT today July 20, and sunrise will be at 11:15 p.m. Earth set is at 9:45 a.m. July 21.
21 July 1997
10 a.m. Pacific Daylight Time
The Mars Pathfinder flight team successfully reestablished contact with the Pathfinder lander and rover early this morning, completing several communications sessions using both the low- gain and high-gain antennas.
"What a difference a day makes," said Brian Muirhead, Pathfinder project manager. "The project team has successfully regained full communication capability on both the low-gain and high-gain antennas. The team is extremely pleased with our current status."
Most of the communications problem experienced over the weekend was associated with ground operations, not with the spacecraft on Mars, Muirhead said. "We'll be working to eliminate the cause of these problems in the coming days, as we return to a more normal mode of operations."
The flight team successfully initiated its first low-gain communications session of the Martian day at 10:38 p.m. Pacific Daylight Time on July 20, then began a second low-gain session at 1:36 a.m. July 21. Both sessions were returning data at the low data rate of 40 bits per second. At 3:22 a.m. PDT, the team conducted a third, brief low-gain session at a slightly higher data rate of 150 bits per second.
"All sessions worked perfectly, and we gained all of the basic engineering and telemetry data that had been stored onboard," Muirhead reported. "We verified that all spacecraft subsystems were healthy."
At 4:50 a.m. PDT, the team conducted a brief high-gain antenna session to make sure the high-gain antenna was pointed at Earth. A full high-gain antenna session at 8,200 bits per second was later performed beginning at 6:43 a.m. PDT.
The team acquired all data on lander and rover health and completed acquisition of all of the spacecraft engineering data. They also sent a software update to correct sequences onboard the flight computer which have caused it to automatically reset itself.
Tonight's science activities will include downlinking measurements of a white-colored rock named Scooby Doo and continuing to acquire data from a full resolution color panoramic photograph of the landing site.
On this Martian day, Sol 17, Earth rose over the newly named Sagan Memorial Station at 8:07 p.m. PDT yesterday July 20. Sunrise was at 11:15 p.m. July 20 and Earth set occurred this morning at 9:45 a.m. July 21.
22 July 1997
12 Noon Pacific Daylight Time
Two-and-a-half weeks after landing in an ancient Martian flood basin known as Ares Vallis, Mars Pathfinder has fulfilled all of its primary science goals and continues to operate nearly flawlessly, the flight team reported at today's press briefing.
More than 300 megabits of data have been returned just in the last week, said Dr. Matthew Golombek, Pathfinder project scientist. The rover continues to follow an aggressive series of maneuvers to study rocks and soils identified by the science teams for their interesting features.
In addition, the rover's wheel tracks and soil abrasion experiments are beginning to yield new information about the Martian soil, which appears to be finer than talcum powder.
A communications problem experienced last weekend has been resolved, reported Richard Cook, Mars Pathfinder mission manager. The problem was associated with ground operations, which has been required to reconfigure equipment and software on a daily basis, and the necessity of establishing communications links only during the short periods of time each day when the lander's transmitter is on.
Scientists are beginning to learn more about the Martian soil by studying the rover's wheel tracks, asking it to perform soil abrasion experiments and measuring the material properties of dust and soil through these wheel-soil interactions.
Dr. Henry Moore, a rover scientist with the U.S. Geological Survey in Menlo Park, CA, likened the Martian soil to a very fine-grained silt that could be found in Nebraska. The Martian particles are less than 50 microns in diameter, which is finer than talcum powder.
Dust coverage on some of the spacecraft instruments is accumulating at a very low rate of about a quarter of a percent per day, added Dr. Geoffrey Landis, NASA Lewis Research Center in Cleveland, which is very close to the team's original predictions.
These measurements also indicated that the dust was not moving toward the Martian poles right now. Additional study of dust patterns in the Martian environment may shed more light on the ways in which dust leaves the Martian atmosphere.
Dr. Peter Smith, University of Arizona, who is principal investigator of the lander camera, described more about the Martian landscape, pointing out a shallow riverbed crossing through the landing site and rocks in the distance that were washed into this outflow channel from the Martian highlands.
About four distinct impressions left by the airbags were evident in the images presented today, noted Dr. Tim Parker, a science team member at JPL. The disturbed soil suggested that the spacecraft was nearly rolling, rather than bouncing, by the time it came to a stop. Parker estimated that the spacecraft bounced 15 to 20 times over a kilometer (6/10ths of a mile) of the landing site before stopping.
Science activities tonight will take the rover through the "cabbage patch," an area of soil in between Scooby Doo and a light-colored rock named Lamb. The rover will conduct a soil experiment , then turn and move toward Lamb. Scientists will take measurements of the dark soil near that rock before moving Sojourner close enough to place its spectrometer against the rock.
On this Martian day, Sol 18, Earth rose over the Sagan Memorial Station at 8:47 p.m. PDT yesterday, July 21. Sunrise was at 11:54 p.m. July 21 and Earth set occurred this morning (July 22) at 10:25 a.m. PDT.
23 July 1997
1:30 p.m. Pacific Daylight Time
The Mars Pathfinder lander and rover continue to operate flawlessly on the surface of Mars, 19 days after landing in an ancient outflow channel called Ares Vallis.
Pathfinder's 1-foot-tall roving geologist -- named Sojourner -- continues to collect data on crustal materials and rocks in the immediate vicinity to provide scientists with new information on the geology of this region. The Pathfinder lander, on the other hand, has become a virtual weather station, using its wind socks, wind sensors and image magnets around-the-clock now to profile the pressure, temperature, density and opacity of the Martian atmosphere.
Two downlink sessions were successfully completed by 11 a.m. today, using the 70-meter (230-foot) antenna of NASA's Deep Space Network facility in Madrid, Spain, reported David Gruel, Mars Pathfinder flight director for Sol 19. The flight team retrieved a total of 45 megabits of data over night, most of which was imaging data from the ongoing science experiments.
"The lander and rover are in excellent health and continue to operate flawlessly," Gruel said. "Meteorological data are being gathered around the clock."
First on Sojourner's list of activities tonight is a wheel abrasion experiment, in which the 10.5-kilogram (23-pound) vehicle will turn and dig some of its wheels into the fine Martian sand to measure material properties of the surface. Next the rover will position its alpha proton X-ray spectrometer face- down in the soil next to a rock called "Lamb" and make measurements of the rock's chemical composition.
On this Martian day, Sol 19, Earth rose over the Sagan Memorial Station at 9:30 p.m. PDT yesterday, July 22. Sunrise was at 12:30 a.m. July 23, and Earth set occurred at 11:04 a.m. PDT today.
24 July 1997
2:30 p.m. Pacific Daylight Time
All communications sessions between the Pathfinder lander and rover were successfully completed today, one day short of the mission's three-week anniversary on the surface of Mars.
Sol 20 began when the Earth rose over Mars' horizon at 10:30 p.m. Pacific Daylight Time last night (July 23), enabling the flight team to initiate communications with the spacecraft. The Sun later rose at 1:15 a.m. PDT this morning, supplying the lander and rover with the energy needed to carry out specific tasks.
Communications were carried out using the 70-meter (230- foot) antenna of NASA's Deep Space Network facility in Madrid, Spain. Forty-seven megabits of data during two downlink sessions were returned on Sol 20.
The data indicated that both the lander and rover remain in excellent health and are continuing to operate masterfully. Flight Director Dave Gruel reported that no further flight software resets have occurred since the team sent modified flight software three sols, or days, ago.
Today's data included numerous images taken for ongoing science experiments. The Imager for Mars Pathfinder (IMP) also completed another section of the 12-color super panorama image of the landing site, then imaged the rover to add to an ongoing "rover movie" that is being assembled. IMP took a final, end-of- the-day photo of Sojourner following completion of its activities.
Sojourner traveled a total of 7/10ths of a meter (2.3 feet) today and performed another soil mechanics experiment that involved staging a "wheely." The last of its activities was to lower the alpha proton X-ray spectrometer onto the soil near the rock named Lamb. Presently, because it is night on Mars, the rover is powered down and using only its battery to operate the spectrometer and gather data on the Martian soil near Lamb. That data will be transmitted to Earth via the lander during the next Martian day, Sol 21, which begins when Earth rises over Mars tonight at 8:48 p.m. PDT.
Activities for Sol 21 will include another rover soil mechanics test, some more autonomous driving and repositioning of Sojourner's spectrometer against the side of Lamb in preparation for data-gathering the following night.
The lander's meteorological experiment reported highs today of minus 2 degrees Celsius (28 degrees Fahrenheit) and morning low temperatures of minus 73 degrees Celsius (minus 99 degrees Fahrenheit). The weather detectors also recorded large fluctuations of 3/10ths millibars in total pressure on the surface of Mars.
On this Martian day, Sol 20, the Earth set at 11:45 a.m. PDT, ending spacecraft communications with Earth for the day. The Sun set at 1 p.m. PDT.
25 July 1997
1:30 p.m. Pacific Daylight Time
Mars Pathfinder celebrated its three-week anniversary on the surface of Mars today, with all spacecraft systems, science instruments and rover activities continuing to go exceptionally well.
On this Martian day, Sol 21, Earth rose at 10:48 p.m. PDT July 24 and Sunrise occurred at 1:53 a.m. PDT today.
The science team finished analyzing alpha proton X-ray spectrometer data from the rock nicknamed "Scooby Doo," the third rock measured by the rover since rolling off its ramp on July 5. ("Barnacle Bill" and "Yogi" were the first two rocks to be measured.) "Scooby Doo," of interest to scientists because of its light color, has a chemical signature very similar to other soils measured at the Pathfinder landing site. However, initial analysis shows that it contains slightly higher amounts of calcium and silicon.
Data returned during successful communications sessions last night indicated that the lander and rover remain in excellent health, reported Guy Beutelschies, Pathfinder flight director for Sol 21.
Sojourner performed a "self-guided" traverse today, receiving a minimum of instructions from Earth before driving off to find its own way to the next rock. Up until now, Sojourner has relied on detailed instructions and "way points," or X-axis and Y-axis coordinates, to find its way to the next rock target.
Today's 3-meter (10-foot) excursion, however, involved only two sets of way point instructions and an additional command to "find the rock." Sojourner used its own hazard avoidance system to locate the two way points, as it usually does, but then relied only on its laser light beams to find the next rock and line up with it. By 11 a.m. PDT, Sojourner had stopped just 25 centimeters (10 inches) in front of "Souffle," the next rock to be studied.
The rover will begin making measurements of "Souffle" on Sol 22, using its alpha proton X-ray spectrometer.
Meanwhile, atmospheric and meteorological data on the temperatures and density of the Martian atmosphere continue to be received during daily telecommunications sessions. Data stored onboard Pathfinder last week, while the flight computer was automatically resetting itself, were returned on Sol 21.
The lander camera snapped images of the disturbed soil near the rock called "Lamb," and photographed three more rocks: "Half Dome," "Shark" and "Pumpkin."
The Earth set today -- Sol 21 -- at 12:24 p.m. PDT. The Sun set at 2:46 p.m. PDT.
26 July 1997
3 p.m. Pacific Daylight Time
The Mars Pathfinder lander and rover remain healthy and are continuing to carry out science experiments on this Martian day, Sol 22. The Earth rose over Mars at 11:28 p.m. PDT July 25. The sun rose today at 2:33 a.m. PDT.
Sojourner's self-guided journey to the rock "Souffle" was interrupted briefly by a software sequencing error, which was identified and corrected immediately. A sequencing error is easily corrected by modifying the numerical coding in the program responsible for executing the command, just as a computer user would modify coding in a program that runs the main menu or desktop functions of a personal computer.
"The problem was corrected immediately and a new sequence was radiated to the rover during the second downlink session," said Becky Manning, flight director for Sol 22. "By the end of that session, ground controllers had received confirmation that the rover had successfully received and was executing the instructions to continue its traverse to Souffle."
Sojourner will leave Souffle on Sol 23 and circumnavigate the lander. When that journey has been completed, the rover will be in the vicinity of three new rocks named "Baker Bench," "Desert Princess" and "Marvin."
The Mars Pathfinder lander imager (IMP) returned more data from the "insurance" panorama and "super" panorama today. It is preparing to take the standard end-of-the-day photograph of the rover before surface operations conclude in 30 minutes.
On Sol 22, the Earth set over Mars at 1:04 p.m. today. The sun will set in about 25 minutes, at 3:25 p.m. PDT.
27 July 1997
4:30 p.m. Pacific Daylight Time
All lander and rover systems and science instruments continue to operate well on Sol 23 of the Mars Pathfinder mission. Earth rise at the landing site occurred at midnight PDT July 26; sunrise followed at 3:13 a.m. PDT today.
In keeping with the tradition of playing wake-up songs for the space shuttle astronauts, the rover and flight teams were awakened on Sol 23 to the music of the Blues Brothers' version of "Raw Hide," a 1960s television western. The song was chosen to match "a long day of driving" for the rover.
Flight Director Jennifer Harris reported that start-of-the- day images showed the rover had begun to climb up the side of the rock named "Souffle," but was not able to position its science spectrometer against the rock. Consequently, no alpha proton X- ray spectrometer data were acquired today.
"However, this did not deter the rover from executing a long traverse which took it past the lander, through the 'Rock Garden' and past a rock named 'Casper,' before coming to a stop near the rocks 'Desert Princess' and 'Baker's Bench,'" Harris said. In all, Sojourner traveled six meters (nearly 20 feet) to complete the traverse, the longest excursion it has taken yet.
Images of the traverse, as well as routine beginning and end-of-day images, were taken by the Imager for Mars Pathfinder (IMP) camera. These images will go into a "rover movie," which is being compiled by the imaging team.
The IMP imaged sunrise on Mars, Phobos, one of Mars' two small moons, and the next portion of the super panorama, Harris said. The flight team also completed its downlink of the IMP stowed-position "insurance pan," which will enable them to begin downloading another portion of the super panorama.
Tomorrow's activities will include sending the rover to a way point beyond the rocks "Calvin" and "Hobbes." There it will be instructed to turn toward a rock named "Mini Matterhorn," take a picture of it and then image the lander.
On this Martian day, Sol 23, the Earth set at 1:43 p.m. PDT and the sun set at 4:04 p.m. PDT.
28 July 1997
3 p.m. Pacific Daylight Time
The Mars Pathfinder lander and rover remain healthy and are continuing science experiments on the surface of Mars. The Earth rose over Mars on this Martian day -- Sol 24 -- at 12:48 a.m. PDT. The sun rose at 3:53 a.m. PDT.
The Imager for Mars Pathfinder (IMP) camera focused its lens on the sky today to photograph dust in the upper atmosphere and to search for clouds. IMP also imaged the wind socks onboard the Pathfinder lander to give scientists more information on wind direction and strength. Also included in today's photography session were images of Phobos, one of Mars' two small moons, and plans to image the Martian sunset later today.
The flight team continued to downlink data for the super panorama of the landing site, which is being assembled by the IMP team, said Flight Director Guy Beutelschies.
Sojourner was awakened this morning with the pop song "Radar Love," and executed a 7-meter (23-foot) traverse, the longest trip yet to be completed. The rover began its journey near the rock "Souffle" and ended it near the rock called "Mini Matterhorn." Next the rover imaged the rock and then the lander. Plans for tomorrow (Sol 25) call for more imaging of "Mini Matterhorn," after which the rover will begin a new traverse toward a rock called "Mermaid."
The Earth set at 2:23 p.m. PDT today and the sun will set at 4:43 p.m. PDT.
29 July 1997
4:30 p.m. Pacific Daylight Time
Imaging the atmosphere of Mars -- how clear or dusty it is and whether there are traces of water vapor -- was the focus of science activities on the surface of Mars today.
The Mars Pathfinder imaging team also photographed the lander's wind socks, three small socks attached at different heights to a 1-meter mast. Visual images of these small socks provide scientists with information on wind strength and direction.
Temperatures on Sol 25 were typical, ranging from highs near minus 12 degrees Celsius (10 degrees Fahrenheit) and lows of minus 79 degrees Celsius (minus 110 degrees Fahrenheit). Today the Earth rose over Mars at 1:28 a.m. PDT and the Sun rose at 4:32 a.m. PDT.
The Atmospheric Science Instrument/Meteorology Package (ASI/MET) instrument team reported a very successful day of data return, said Flight Director Jennifer Harris, receiving more information than ever before on the pressure of the Martian atmosphere. Also included in the downlink sessions was more imaging data for the high-resolution "super panorama" of the landing site. In all, a total of 48 megabits of data was successfully returned.
A sequencing transmission error prevented the rover from executing its daily traverse, Harris said. The situation was quickly corrected and the rover was able to complete an accelerometer diagnosis sequence, which involved making a 120- degree turn in place. Sojourner will complete its traverse to the rock nicknamed Mini Matterhorn tomorrow and then turn to image the lander.
The Earth set today at 3:03 p.m. PDT and the Sun will set at 5:22 p.m. PDT.
30 July 1997
4:30 p.m. Pacific Daylight Time
Pathfinder's 10.5-kilogram (23-pound) rover called Sojourner stalled today during the last stretch of its journey to a rock nicknamed Mermaid, but was quick to recover and prepare for completion of the traverse tomorrow.
Data returned this morning from the Sagan Memorial Station indicated that the rover's left front wheel stalled during the third of four waypoint maneuvers. Waypoints are navigational instructions -- consisting of x- and y-axis coordinates -- used by the rover to travel from one rock to another.
To complete today's traverse to the rock nicknamed Mermaid, for instance, the rover had to make four short trips based on four sets of waypoint coordinates. Its wheel jammed during the third segment of the journey.
"This stall was probably caused by a small rock becoming jammed in one of the rover wheel's cleats," said Flight Director David Gruel. "Once the rover detected the stall on its own, she was able to autonomously clear the problem by backing up a short distance. Since the stall did not exist after the backup was performed, there's a high probability that Sojourner is ready to continue the drive around the lander tomorrow."
Approximately 55 megabits of engineering and science data were returned today. All data indicated the lander and rover are healthy and the lander's battery continues to power the craft through the subfreezing nights on Mars. The rover returned a "spectacular" image today of the rear of the lander and the rock nicknamed Mini Matterhorn.
Temperatures on Mars today ranged from a balmy minus 13 degrees Celsius (8 degrees Fahrenheit) at 5:35 p.m. local solar time to minus 79 degrees Celsius (minus 105 degrees Fahrenheit) at 5:30 a.m. local time. Winds were light and from the west.
On this Martian day, Sol 26, Earthrise occurred at 2:09 a.m. PDT and sunrise followed at 5:12 a.m. PDT. The Earth later set at 3:43 p.m. PDT and Pathfinder will observe its 26th sunset at 6 p.m. PDT.
31 July 1997
4:30 p.m. Pacific Daylight Time
The Mars Pathfinder flight team has completed all of its science and engineering goals, four days before the primary mission draws to a close, said Dr. Matthew Golombek, Pathfinder project scientist, at today's press briefing.
Atmospheric-surface interactions were the focus of today's presentation. To set the stage, Dr. Mark Lemmon, a member of the Imager for Mars Pathfinder (IMP) camera team from the University of Arizona, presented new images of the Martian sunrise and sunset.
True to color, the dawn images revealed pale pink sunrises and clouds floating overhead. The reddish tint is the result of Martian dust, composed of oxidized iron, which is present in the atmosphere. The sunset images -- color-enhanced to bring out structural detail in the atmosphere -- showed a sky darkening to salmon-colored hues.
These spectacular images of the Martian summer are possible by return of an unprecedented amount of science and engineering data -- on the order of 400 megabits just in the last nine days -- Golombek pointed out.
Temperature highs and lows at the landing site have not varied much, said Dr. Robert Haberle, a participating scientist from the NASA Ames Research Center. They range from highs of about minus 12 degrees Celsius (8 degrees Fahrenheit) to lows near minus 76 Celsius (minus 105 Fahrenheit). Frozen water-ice clouds are evident in the Martian sky during the early morning hours, but evaporate once temperatures rise.
"We expect late night and early morning clouds, but we expect those clouds will burn off fairly rapidly with sunrise, giving way to a dusty Martian day," Haberle said. Although there has not been much variation in these weather conditions since Pathfinder arrived, they are expected to begin changing in about a month, as fall arrives and ushers in the dust storm season.
Atmospheric pressures, on the other hand, are fluctuating dramatically, sometimes peaking two, three or four times a day, Haberle noted. Pressure oscillations are indicative of a global scale thermal tidal system that is moving dust, water-ice or vapor clouds and other volatiles through the atmosphere. On Mars, these atmospheric variations are sizable, whereas on Earth they almost never occur.
Since data-gathering began, the maximum change in pressure over the course of a day has been 0.3 millibars, which is about 4.5 percent of the average pressure on Mars. On Earth, pressures that low might occur during a severe hurricane. A better understanding of these pronounced pressure oscillations will help scientists understand the processes by which volatiles enter and escape the Martian atmosphere, and may shed more light on the rise of regional and global dust storms.
Wind speeds have been increasing with altitude, reported Dr. Robert Sullivan of Arizona State University. And temperatures will vary dramatically with elevation. When ground temperatures are 16 to 21 degrees Celsius (60 to 70 degrees Fahrenheit), they can drop to minus 23 to 27 degrees Celsius (minus 10 to minus 15 degrees Fahrenheit) just five and a-half feet above the ground.
Images of the Martian landscape also revealed a shiny object about 1,200 meters (7/10ths of a mile) away from the lander. Dr. Michael Malin, a participating scientist, said the object is about the same dimensions and is probably the spacecraft's discarded backshell, which separated just before the spacecraft landed.
Although the Pathfinder lander and rover remain healthy, engineers plan to recharge the lander's battery during a two-day hiatus beginning Sunday, Aug. 3. The lander will perform some science experiments during the day, but will use most of its solar energy to charge the battery. At night, the craft essentially goes to sleep.
The rover will continue its daily traverses and spectrometer studies, rolling off to a smooth, dark region of soil called Mermaid Dune tomorrow. After taking measurements of the soil, scientists will identify one of three large, dust-free rocks -- Shark, Half Dome and Wedge -- as the next target for study.
On this Martian day, Sol 27, Earthrise occurred at 2:49 a.m. PDT and sunrise followed at 5:52 a.m. PDT. The Earth later set at 4:23 p.m. PDT and the sun set at 6:41 p.m. PDT.
#
"Pathfinder Status (August97)",625,0,0,0
02 August 1997
5 p.m. Pacific Daylight Time
Mars Pathfinder lander and rover operations were curtailed today when the lander downlink session, scheduled to begin at 1:20 p.m. PDT, was not initiated. The cause of the problem is currently unknown, said Mars Pathfinder Mission Manager Richard Cook.
The flight team, led by Flight Director Carl Steiner, subsequently reestablished communications with the lander and received a brief carrier blip at 3:30 p.m. PDT. Additional downlink sessions were not attempted, however, because of the lack of time before the Earth set over the landing site at 5:42 p.m. PDT.
Data on the health of the lander and rover, in addition to other engineering telemetry, will be acquired tomorrow. The flight team expects Sol 30 operations to proceed normally.
Additional information will be posted on the JPL home page at http://www.jpl.nasa.gov as it becomes available. An audio update is also available by calling 1-800-391-6654.
On this Martian day, Sol 29, Earthrise occurred at 4:09 a.m. PDT and sunrise followed at 7:11 a.m. PDT. The sun set over the landing site at 7:59 p.m. PDT.
03 August 1997
7 p.m. Pacific Daylight Time
Although the reason for yesterday's loss of downlink opportunities has not yet been identified, science activities proceeded normally today on the surface of Mars. Today, Sol 30, marks the end of the Mars Pathfinder primary mission, 30 days after the spacecraft landed in an ancient outflow channel called Ares Vallis.
The Imager for Mars Pathfinder (IMP) continued to image the thin Martian atmosphere, the lander's wind socks, the Sun and the rover as it roamed to another destination, said Carl Steiner, Mars Pathfinder flight director. Acting as a weather station, the Pathfinder lander -- now called the Sagan Memorial Station -- gathered weather data for the 12th consecutive day.
Data from yesterday's surface operations had been stored onboard the lander and were downlinked today. Highs on Mars rose to minus 10 degrees Celsius (14 degrees Fahrenheit) today and dipped to minus 70 degrees Celsius (minus 94 degrees Fahrenheit).
The rover finished its soil analysis of Mermaid Dune before heading toward the Rock Garden. An onboard tilt protection circuit caused the rover to shut down after reaching 10 centimeters (0.3 feet) of motion.
"An especially noisy accelerometer had caused this problem in past, but had successfully guarded the rover against excessive tilt," Steiner said. "The rover team thought it prudent to activate this device, even with the possibility of inadvertent shut-down, because of the uneven path to the Rock Garden and the long traverse." Sojourner will resume this traverse on Sol 32, as Pathfinder's extended mission gets under way.
Three downlink sessions were successfully carried out today using the low-gain antenna once and the high-gain antenna for the next two sessions, Steiner said. The operations team, however, was unable to complete its planned downlink of an eighth (octant) of the so-called "super pan" before the end of the day and the beginning of a two-day sleep period for the lander.
"This will be the first time in nearly 240 days that the lander's electronics will be powered off," Steiner said. "At the conclusion of today's activities, all lander electronics, with the exception of a few computer chips that comprise the hybernate circuit, will be powered off to conserve energy through the Martian evening and prolong our waning battery."
The hybernate circuit has been programmed to wake up the lander at 7:30 a.m. local solar time tomorrow. A backup circuit will wake the lander at 8 a.m. if the lander is still asleep. Tomorrow's activities will focus on recharging the battery to the fullest capacity possible. No science experiments are planned.
On this Martian day, Sol 30, Earthrise occurred at 4:49 a.m. PDT and sunrise occurred at 7:51 a.m. PDT. The Earth set over the landing site at 6:22 p.m. PDT and the sun set at 8:39 p.m. PDT.
18 August 1997
4 p.m. Pacific Daylight Time
Daily communications with the Mars Pathfinder lander and rover have resumed after an interruption on Saturday, Aug. 16, that was caused by an automatic reset of the lander's flight computer. The cause of the reset is not known, said Mars Pathfinder Flight Director Rob Manning.
The flight team was able to reactivate the lander by sending it instructions to reinitialize high-gain pointing and then begin its scheduled downlink session on Sunday evening. The downlink session began right on time at 10:09 p.m. PDT Aug. 17.
New images indicated that the rover had stopped its traverse to a rock nicknamed Shark after partially climbing up a rock called Wedge. The rover's hazard avoidance software is designed to stop the vehicle when it begins to tilt too much.
Wedge is one of many small rocks forming a gateway to the Rock Garden, which is Sojourner's next destination. This portion of the landing site is more challenging than other regions because it is much rockier than any terrain explored to date by the rover.
The flight team will instruct Sojourner to continue its trek to Shark tonight. Shark is of interest to scientists because it is a large, smooth rock, which is relatively dust-free and, therefore, an ideal candidate for the next spectrometer analysis.
Shark is part of a cluster of large rocks standing straight up from the ground, but leaning slight to the left. Scientists have theorized that these large, angular rocks may have been tilted by a huge flood which swept through Ares Vallis early in Mars' history.
Twenty-six megabits of data were received last night, on Sol 44 of the Pathfinder mission. The data included rover health status data, meteorological data and three additional sections of the "super" panorama.
27 August 1997
Images of the Martian sunrise and sunset, with water ice clouds floating through the atmosphere, were unveiled today at a Mars Pathfinder press briefing, held on Sol 53 of the mission, at NASA's Jet Propulsion Laboratory.
Today's collection of photographs included one portion of the super panorama view looking to the north-northeast from the Sagan Memorial lander. The super panorama of the landing site, which is being constructed from high resolution color images taken by the Imager for Mars Pathfinder (IMP) instrument, will be comprised of about 3,000 images when it is completed in about eight weeks, said Dr. Matthew Golombek, Pathfinder project scientist at JPL.
This mammoth color and stereo data set, which is now about 65 percent finished, will be used to derive high quality topographic maps of the Martian surface and detailed shapes of rocks and other surface features. Scientists will also be able to examine subtle chemical, mineralogical and textural variations in rocks and soils from this panorama.
Temperatures on Mars today remained in roughly the same temperature range. Today's low was minus 75 degrees Celsius (-103 degrees Fahrenheit) and the high was minus 10 degrees Celsius (14 degrees Fahrenheit).
The highest pressure measurements seen yet on Mars were recorded yesterday (Sol 52) at 6.8 millibars, said Dr. Tim Schofield, atmospheric structure/meteorology package team leader at JPL. In addition to temperature and pressure measurements, Pathfinder has observed a total of 12 dust devils, small swirls of dust kicked up by winds blowing down through the canyons in the Ares Vallis landing zone.
Scientists are finding that Martian temperatures are cooler at higher altitudes (about 80 kilometers or 50 miles) than on the ground. They think that ice clouds forming about 10 kilometers to 15 kilometers (6 to 9 miles) above the surface are responsible for this cooling trend higher in the atmosphere. The clouds are thought to be made of very small ice particles, about one-tenth the size of Martian dust or one-thousandth the thickness of a human hair.
Dr. Mark Lemmon, a member of the lander camera imaging team at the University of Arizona, noted color variations in some of the sunset pictures. The blue color is not caused by clouds of water ice but by Martian dust in the atmosphere, Lemmon said.
The dust absorbs blue light, giving the sky its red color, but it also scatters some of the blue light into areas that looked very blue around the Sun. The blues only show up near sunrise and sunset, when the light has to pass through the largest amount of dust.
Sojourner, which remains in excellent health, began exploring the Rock Garden yesterday, after spending about a week en route to the region. The Rock Garden is an assemblage of several large boulders and many smaller rocks near the lander.
After conducting a chemical analysis yesterday of the rock nicknamed Shark, the rover moved toward another rock called Half Dome today, but climbed too high up on the rock and automatically shut itself off.
Tomorrow the rover team will instruct the rover to back down the rock and reposition the alpha proton X-ray spectrometer against the side of Half Dome. Chemical analyses of all of the rocks studied so far indicate that at least two types of rocks are present in the Pathfinder landing zone: those with high levels of silicon and those with high levels of sulfur, reported Dr. Tom Economou, co-investigator of the alpha proton X-ray spectrometer team at the University of Chicago.
Soil mechanics experiments using the rover's wheels and cleats to dig below the surface have revealed different layers of material, Howard Eisen, principal investigator on the soil mechanics technology experiment at JPL, pointed out.
Soil surfaces differ near the lander, where the soil contains a mixture of pebbles, fine-grained sand and clods, from regions a bit farther out. There, the surface is covered with a bright drift material, Eisen said. Using the rover's cleats to dig below the surface, scientists have discovered that cloddy material was present underneath the drift.
After traveling a total of about 80 meters (263 feet) around the landing site, Sojourner will continue to explore the Rock Garden for the next several days, taking as many chemical analyses as possible of the large boulders in the vicinity.
After these rocks have been studied, the rover will head back to the ramp on which it exited the lander and study a dust sample that has been accumulating on a magnet, Golombek said. This study may provide new information about magnetic properties that might be present in the Martian soil.
Longer range plans for the rover may take it much farther away from the lander, so that it may peer over the rim of what appears to be a shallow riverbed, and photograph a region that cannot be seen by the lander.
The Earth rose over Mars on Sol 53 at 8:15 p.m. Pacific Daylight Time and the Sun rose at 11:05 p.m. PDT yesterday (Aug. 26). The Earth set this morning at 9:55 a.m. PDT and the Sun set at 11:35 a.m. PDT.
#
"Pathfinder Status (October97)",626,0,0,0
01 October 1997
After experiencing difficulties in communicating with the Mars Pathfinder spacecraft for the past three days, the operations team was able to reestablish a brief two-way communications session Tuesday using the lander's auxiliary transmitter. Receipt of this beacon signal indicated that the spacecraft is still operational.
The team began having communications problems with the spacecraft on Saturday, Sept. 27. These problems could be related to degradation of the spacecraft's battery. The last successful data transmission cycle from Pathfinder was completed at 3:23 a.m. Pacific Daylight Time on Sept. 27, which was Sol 83 of the mission.
No signal was received from the spacecraft on the next Martian day, Sol 84, which began in the evening of Sept. 27. The team's transmission session began at 11:15 p.m. PDT. The lack of a signal, at that time, was thought to be caused by a possible computer reset incident, ground system problem or low voltage condition.
A reset or a low voltage condition, caused by the aging of the battery, would cause the spacecraft sequence to automatically stop and not execute its planned communication with Earth.
The team attempted to communicate with the spacecraft again on Sept. 29 (Sol 85) and Sept. 30 (Sol 86) with no success.
Tonight, on Sol 88 of the mission, the team will use the auxiliary transmitter again to attempt to acquire engineering data that will help them assess the cause of the communications problem. Meanwhile, the rover, which receives its instructions from Earth via the lander, is currently running a contingency program which has instructed it to stand still rather than begin its trek around the lander.
The team will repeat these activities on subsequent days and attempt to receive telemetry that will give them more information about the health of the lander and rover.
If Pathfinder operations do not return to normal tonight, a Mars Pathfinder team representative will provide an update on the situation at the beginning of the planned Mars Global Surveyor science news briefing at 9 a.m. PDT on Thursday, Oct. 2.
07 October 1997
The Mars Pathfinder operations team reestablished communications with the lander today, on Sol 92 of the mission, after four days of silence from the spacecraft. The team received a transmission from the spacecraft's main transmitter. The signal was detected using the Madrid, Spain 34-meter antenna.
No data was received, but receipt of a spacecraft signal indicates that the lander is operational and the battery is off- line. Meanwhile, the rover, which is programmed to begin a contingency sequence when it has not heard from the lander for five days, started that activity on Sol 90. In this mode of operation, the rover is instructed to return to the lander and begin circling it.
The Mars Pathfinder operations team will repeat commands tomorrow night, on Sol 93, to verify two-way communications with the lander's main transmitter and attempt to return engineering data on the health of the lander and rover. If successful, that information would be returned the following day, on Sol 94 of the mission.
09 October 1997
\BPathfinder Team Paints an Earth-Like Picture of Early Mars\b
Mars is appearing more and more like a planet that was very Earth-like in its infancy, with weathering processes and flowing water that created a variety of rock types and a warmer atmosphere that generated clouds, winds and seasonal cycles.
Those observations, along with new images taken by the Mars Pathfinder rover and lander, and an update on the condition of the spacecraft, were presented at an Oct. 8 press briefing originating from NASA's Jet Propulsion Laboratory.
"What the data are telling us is that the planet appears to have water-worn rock conglomerates, sand and surface features that were created by liquid water," said Dr. Matthew Golombek, Mars Pathfinder project scientist at JPL.
"If, with more study, these rocks turn out to be made of composite materials, that would have required liquid water flowing on the surface to round the edges in pebbles we see on the surface or explain how they were embedded in larger rocks. That would be a very important finding."
Golombek also stressed the amount of differentiation -- or heating, cooling and recycling of crustal materials -- that appears to have taken place on Mars. "We're seeing a much greater degree of differentiation -- the process by which heavier elements sink to the center of the planet while lighter elements rise to the surface -- than we previously thought, and very clear evidence that liquid water was stable at one time in Mars' past.
"Water, of course, is the very ingredient that is necessary to support life," he added, "and that leads to the $64,000 question: Are we alone in the universe? Did life ever develop on Mars? If so, what happened to it and, if not, why not?"
Despite recent communications problems with Earth, the Mars Pathfinder lander and rover are continuing to operate during the Martian days, when they can receive enough energy to power up spacecraft systems via their solar panels. The mission is now into Sol 94, or the 94th Martian day of operations, since landing on July 4.
"Everything that we have seen over the last 10 days (with respect to communications) is like a twisty little maze with passages all alike," said Jennifer Harris, acting mission manager. "I am happy to report that we have made contact with the spacecraft using its main transmitter. We were able to confirm that we could send a command to the spacecraft to turn its transmitter on and then turn it off.
"We don't know yet whether we are receiving that signal over the low-gain or high-gain antenna," she added, "but we should be able to determine this over the next few days."
The Mars Pathfinder team began having communications problems with the spacecraft on Saturday, Sept. 27. After three days of attempting to reestablish contact, they were able to lock on to a beacon signal from the spacecraft's auxiliary transmitter on Oct. 1, which meant that the spacecraft was still operational.
At that time they surmised that the communications problems were most likely related to depletion of the spacecraft's battery and uncertainties in the onboard clock. The last successful data transmission cycle from Pathfinder was completed at 3:23 a.m. Pacific Daylight Time on Sept. 27, which was Sol 83 of the mission.
Since then, efforts have been made during each Martian day to reestablish contact with both the primary and auxiliary transmitter and obtain engineering telemetry that would tell the team more about the health of the lander and rover. On Oct. 7, the team was able to lock on to Pathfinder's signal, via NASA's Deep Space Network 34-meter-diameter (112-foot) dish antenna in Madrid, Spain, for about 15 minutes, using the main transmitter. However, in repeating the process on Oct. 8, they did not receive a signal.
The rover, which receives its instructions from Earth via the lander, is currently running a contingency software program that was preprogrammed to start up if the vehicle did not hear from the lander after five Martian days. That program was powered on Oct. 6, on Sol 92 of the mission.
In this contingency mode, the rover is instructed to return to the lander and begin circling it. This precaution is designed to keep Sojourner close to the lander in the event that the spacecraft was able to begin communicating with it again.
If normal communications are reestablished, the rover team will send new commands to Sojourner to halt the contingency circling and begin a traverse to a specific location.
Dr. William Folkner, an interdisciplinary scientist at JPL, presented data on the rotation and orbital dynamics of Mars, which are being obtained from two-way ranging and Doppler tracking of the lander as Mars rotates.
Measurements of the rate of change in Mars' spin axis have important implications for learning more about the density and mass of the planet's interior. Eventually, scientists may be able to determine whether Mars' core is presently molten or fluid. The size of the core also can be used to characterize the thickness, or radius, of Mars' mantle.
"By measuring the spin axis of Mars, we can learn something about the interior of the planet, because the speed of the change in its orientation is related to how the mass is distributed inside," Folkner said. "If the core is fluid, its spin and the way in which the planet wobbles slightly will be different from the spin and wobble of a planet with a solid core.
"If Mars' core is solid, then it can't be less than about 1,300 kilometers (807 miles) in radius, out of the planet's total radius of 3,400 kilometers (2,112 miles)," Folkner added. "If the core is made up of something less dense than iron, if it's a mixture of, say, iron and sulfur, then the core would be bigger, but it couldn't be bigger than about 2,000 kilometers (1,242 miles) in radius."
New close-up images of dunes around the landing site are showing some scientists clear evidence that there is sand on the surface of Mars. Identification of sand, as opposed to dust or pebbles, is a significant factor in establishing that weathering processes such as erosion, winds and flowing water all contributed to Mars' present landscape.
"We've made significant progress in establishing that water was a dominant agent in forming the surface, and now we can say that there is another agent at work, and that is the wind, that has created and modified some of the landforms on a smaller and medium scale," said Dr. Wes Ward of the U.S. Geological Survey, Flagstaff, AZ, a member of the Imager for Mars Pathfinder team. "And because the water is no longer there, wind probably is the dominant agent shaping the Martian surface at this moment."
Ward showed images of Ares Vallis, taken by the rover and Viking 1 orbiter images to point out the structural difference in these surface features. While Viking 1 surface features around a rock nicknamed "Big Joe" showed drifts, the dune-like surfaces in the Ares Vallis flood basin resemble sand that has been blown southwest over the landing site.
The presence of sand also points to the likely presence of liquid water, needed to create these small, 1-millimeter-diameter granules, and weathering agents such as wind to blow them into small ridges and moats present around the Ares Vallis rocks.
"The wind is quite an active agent," Ward said. "Sand is the smoking gun, and as far as I'm concerned, the gun is smoking and has Colonel Mustard's prints all over it. We are seeing sand at the landing site."
Dr. Greg Wilson, of Arizona State University, who is on the Pathfinder atmospheric experiment team, reported increases in the pressure of the Martian atmosphere and a drop in surface temperatures.
"We expect to see a continued increase in pressure and decrease in temperatures as the dust season approaches and winds begin to lift more dust into the Martian atmosphere," he said. "The dust season on Mars usually begins in the next few weeks."
16 October 1997
The Mars Pathfinder operations team is continuing its efforts to attempt to send commands to the Pathfinder lander while, at the same time, investigating possible scenarios to explain what might be occurring onboard the spacecraft. The last signal received from the spacecraft was on Sol 93, which was Tuesday, October 7, at 7:21 a.m. Pacific Daylight Time.
There is no indication at this time that the spacecraft is no longer operating. The difficulty in communications is thought to be related to the degradation of the spacecraft battery. In the "no battery" mode of operations, the spacecraft cannot keep track of time accurately and will also be powered on for a smaller portion of each day.
As time progresses, spacecraft hardware will become colder. In regular operations, by turning on the transmitter, the spacecraft hardware warms up sufficiently to operate normally. It is possible that because the team has not been able to turn on the transmitter for long periods of time over several days that the spacecraft temperatures are colder than normal. Those lower temperatures could cause the spacecraft hardware to operate differently than expected.
In their attempts to communicate with the spacecraft recently, the operations team is focusing on both the non- operational battery scenario and the colder temperatures at which the spacecraft is probably operating. They are also experimenting with the timing of their commands in the belief that the spacecraft may be waking up later than normal due to the faulty onboard clock.
22 October 1997
The Mars Pathfinder operations team is continuing its efforts to reestablish communications with the Pathfinder lander. Although they are experiencing communications difficulties, the team is confident that the spacecraft is still operating on the surface of Mars, according to Mission Manager Richard Cook.
The last time they were able to send a command to the Pathfinder lander instructing it to transmit a signal back to Earth was on Sol 93, which was Tuesday, October 7, at 7:21 a.m. Pacific Daylight Time.
Team members suspect that the spacecraft may not be receiving commands from Earth properly because the lander's hardware has become much colder than normal. In regular operations, when the lander's transmitter is turned on, spacecraft hardware warms up sufficiently to operate normally.
Since the transmitter has not been on for several days, engineers suspect that temperatures within the lander are considerably colder than normal. Predicted internal temperatures drop to as low as -50 ║C (-58 ║F) in the early morning and only rise to about -30 ║C (-22 ║F) in the late afternoon. These temperatures are about 20 ║C (38 ║F) colder than the coldest previous operational temperatures.
The lower temperatures cause the spacecraft radio hardware to operate outside the range of radio frequencies that ground controllers have used in the past. During the past three weeks the operations team has been transmitting to the spacecraft at a lower frequency and sweeping through a wider frequency range, a technique that has been used on other missions to attempt to cause the spacecraft receiver to lock on to the transmitted signal. Once ground controllers finish this, they send commands instructing the lander to turn on its transmitter and send a signal back to Earth.
To be certain that they investigate all possibilities, team members are also consulting with experts knowledgeable about the radio and other key elements of the spacecraft. They have identified some new scenarios that are being pursued to regain communications.
These recommendations include doing more testing of the engineering model hardware in the laboratory to better understand how the spacecraft might be behaving. Another recommendation has suggested shifting and increasing the range of frequencies being swept through much more than previously attempted.
According to Project Manager Brian Muirhead, the possibility exists that an unrecoverable problem may have occurred. Team members expected that, once the lander's onboard battery died, cold and thermal cycling could result in a failure of some other element of Pathfinder and thereby end the mission.
"However, the team will continue to do everything possible to reestablish communications until all options have been exhausted," Muirhead said. The mission has already exceeded all of its goals in terms of spacecraft lifetime and data return.
The science team, meanwhile, continues to process and analyze the large volume of data sent back by Pathfinder's lander and rover. Further science products are planned and new results will continue to be presented as they develop.
#
"Pathfinder Status (November97)",627,0,0,0
November 4, 1997
After operating on the surface of Mars three times longer than expected and returning a tremendous amount of new information about the red planet, NASA's Mars Pathfinder mission is winding down.
Flight operators at NASA's Jet Propulsion Laboratory, Pasadena, CA, made the announcement today after attempting to reestablish communications with the spacecraft over the last month. With depletion of the spacecraft's main battery and no success in contacting Mars Pathfinder via its main or secondary transmitters, the flight team cannot command the spacecraft or the small rover named Sojourner that had been roving about the landing site and studying rocks.
"We concede that the likelihood of hearing from the spacecraft again diminishes with each day," said Pathfinder Project Manager Brian Muirhead. "We will scale back our efforts to reestablish contact but not give up entirely.
"Given that, and the fact that Pathfinder is the first of several missions to Mars, we'll say 'see you later' instead of saying goodbye," he said.
At the time the last telemetry from the spacecraft was received, Pathfinder's lander had operated nearly three times its design lifetime of 30 days, and the Sojourner rover operated 12 times its design lifetime of seven days.
"I want to thank the many talented men and women at NASA for making the mission such a phenomenal success. It embodies the spirit of NASA, and serves as a model for future missions that are faster, better, and cheaper. Today, NASA's Pathfinder team should take a bow, because America is giving them a standing ovation for a stellar performance," said NASA Administrator Daniel S. Goldin.
Since its landing on July 4, 1997, Mars Pathfinder has returned 2.6 billion bits of information, including more than 16,000 images from the lander and 550 images from the rover, as well as more than 15 chemical analyses of rocks and extensive data on winds and other weather factors. The only remaining objective was to complete the high-resolution 360-degree image of the landing site called the "Super Pan," of which 83 percent has already been received and is being processed. The last successful data transmission cycle from Pathfinder was completed at 3:23 a.m. Pacific Daylight Time on Sept. 27, which was Sol 83 of the mission.
"This mission has advanced our knowledge of Mars tremendously and will surely be a beacon of success for upcoming missions to the red planet," added Dr. David Baltimore, president of the California Institute of Technology, which manages JPL for NASA. "Done quickly and within a very limited budget, Pathfinder sets a standard for 21st century space exploration."
The Mars Pathfinder team first began having communications problems with the spacecraft on Saturday, Sept. 27. After three days of attempting to reestablish contact, they were able to lock on to a carrier signal from the spacecraft's auxiliary transmitter on Oct. 1, which meant that the spacecraft was still operational. They locked on to the same carrier signal again on Oct. 6, but were not able to acquire data on the condition of the lander. At that time, the team surmised that the intermittent communications were most likely related to depletion of the spacecraft's battery and a drop in the spacecraft's operating temperatures due to the loss of the battery, which kept the lander functioning at warmer temperatures.
Over the last month the operations team has been working through all credible problem scenarios and taking a variety of actions in attempting to recover the link with Pathfinder. With all of the most plausible possibilities exhausted, the team plans to continue sending commands and listening for a spacecraft signal on a less frequent basis.
"Basically we are shifting to a contingency strategy of sending commands to the lander only periodically, perhaps once a week or once per month," said Mission Manager Richard Cook. "Normal mission operations are over, but there is still a small chance of reestablishing a link, so we'll keep trying at a very low level."
Although the true cause of the loss of lander communications may never be known, recent events are consistent with predictions made at the beginning of the extended mission in early August, Muirhead said. When asked about the life expectancy of the lander, project team members predicted that the first thing that would fail on the lander would be the battery; this apparently happened after the last successful transmission September 27.
After that, the lander was expected to begin getting colder at night and go through much deeper day-night thermal cycles. Eventually, the cold or the cycling would probably render the lander inoperable. According to Muirhead, it appears that this sequence of events has probably taken place. The health and status of the rover is also unknown, but since initiating its onboard backup operations plan a month ago, the rover is probably circling the vicinity of the lander, attempting to communicate with it.
The rover, which went into a contingency mode on Oct. 6, or Sol 92 of the mission, had completed an alpha proton X-ray spectrometer study of a rock nicknamed Chimp, to the left of the Rock Garden, when it was last heard from. The rover team had planned to send the rover on its longest journey yet -- a 165-foot (50-meter) clockwise stroll around the lander -- to perform a series of technology experiments and hazard avoidance exercises when the communications outage occurred. That excursion was never initiated once the rover's contingency software began operating.
Now known as the Sagan Memorial Station, the Mars Pathfinder lander was designed primarily to demonstrate a low-cost way of delivering a set of science instruments and a free-ranging rover to the surface of the red planet. Landers and rovers of the future will share the heritage of spacecraft designs and technologies first tested in this "pathfinding" mission.
Part of NASA's Discovery program of low-cost planetary missions, the spacecraft used an innovative method of directly entering the Martian atmosphere. Assisted by a 36-foot-diameter (11-meter) parachute, the spacecraft descended to the surface of Mars on July 4 and landed, using airbags to cushion the impact. The spacecraft's novel entry was successful.
Scientific highlights of the Mars Pathfinder mission are:
\B╖\b Martian dust includes magnetic, composite particles, with a mean size of one micron.
\B╖\b Rock chemistry at the landing site may be different from Martian meteorites found on Earth, and could be of basaltic andesite composition.
\B╖\b The soil chemistry of Ares Vallis appears to be similar to that of the Viking 1 and 2 landing sites.
\B╖\b The observed atmospheric clarity is higher than was expected from Earth-based microwave measurements and Hubble Space Telescope observations.
\B╖\b Dust is confirmed as the dominant absorber of solar radiation in Mars' atmosphere, which has important consequences for the transport of energy in the atmosphere and its circulation. Frequent "dust devils" were found with an unmistakable temperature, wind and pressure signature, and morning turbulence; at least one may have contained dust (on Sol 62), suggesting that these gusts are a mechanism for mixing dust into the atmosphere.
\B╖\b Evidence of wind abrasion of rocks and dune-shaped deposits was found, indicating the presence of sand.
\B╖\b Morning atmospheric obscurations are due to clouds, not ground fog; Viking could not distinguish between these two possibilities.
\B╖\b The weather was similar to the weather encountered by Viking 1; there were rapid pressure and temperature variations, downslope winds at night and light winds in general. Temperatures were about 10 degrees warmer than those measured by Viking 1.
\B╖\b Diversity of albedos, or variations in the brightness of the Martian surface, was similar to other observations, but there was no evidence for the types of crystalline hematite or pyroxene absorption features detected in other locations on Mars.
\B╖\b The atmospheric experiment package recorded a temperature profile different than expected from microwave measurements and Hubble observations.
\B╖\b Rock size distribution was consistent with a flood-related deposit.
\B╖\b The moment of inertia of Mars was refined to a corresponding core radius of between 807 miles and 1,242 miles (1,300 and 2,000 kilometers).
\B╖\b The possible identification of rounded pebbles and cobbles on the ground, and sockets and pebbles in some rocks, suggests conglomerates that formed in running water, during a warmer past in which liquid water was stable.
Engineering milestones of the mission included demonstrating a new way of delivering a spacecraft to the surface of Mars by way of direct entry into the Martian atmosphere. In addition, Mars Pathfinder demonstrated for the first time the ability of engineers to deliver a semi-autonomous roving vehicle capable of conducting science experiments to the surface of another planet.
#
"Solar Flare",628,0,0,0
\BApril 9, 1997
updated April 10, 1997\b
Astronomers have observed a giant eruption of material on the Sun that is headed for the Earth. Detected by NASA's Solar and Heliospheric Observatory (SOHO), the million-mile per hour stream of particles is predicted to slam into the Earth sometime Wednesday afternoon. When these events reach the Earth, they can produce spectacular images of the Earth's aurora, or "Northern Lights."
\BApril 15, 1997\b
Astronomers observed a giant eruption of material on the Sun early last week heading for the Earth. Detected by NASA's Solar and Heliospheric Observatory (SOHO), the million-mile per hour stream of particles that slammed into the Earth produced a spectacular aurora, or "Northern Lights" display. Solar eruptions can cause a brilliant Earthly aurora that lasts for days, and this eruption did just that.
Scientists at NASA's Marshall Space Flight Center watched the Northern Lights in real-time with the Ultraviolet Imager Experiment aboard the POLAR spacecraft. The current auroral pictures are posted as they are received from the instrument.
#
"Ulysses Mission",629,0,0,0
The Ulysses Mission is the first spacecraft to explore interplanetary space at high solar latitudes. Ulysses is a joint endeavor of the European Space Agency (ESA) and the National Aeronautics and Space Administration (NASA) of the USA. The spacecraft and spacecraft operations team are provided by ESA, the launch of the spacecraft, radio tracking, and data management operations are provided by NASA. Scientific experiments are provided by investigation teams both in Europe and the USA. The spacecraft was launched on Oct. 6, 1990 by the shuttle Discovery with two upper stages.
To reach high solar latitudes, the spacecraft was aimed close to Jupiter so that Jupiter's large gravitational field would accelerate Ulysses out of the ecliptic plane to high latitudes; no man made launch vehicle could by itself provide the needed velocity for Ulysses to achieve high latitudes. Encounter with Jupiter occurred on February 8, 1992, and since then Ulysses travelled to higher latitudes with maximum Southern latitude of 80.2 degrees being achieved on Sept. 13, 1994.
Ulysses will travel through high Northern latitudes during June through September of 1995. These high latitude observations are being obtained during the quiet (minimum) portion of the 11 year solar cycle. The spacecraft's orbital period is six years and, if kept in operation, the high latitude observations will be obtained during the active (maximum) portion of the solar cycle.
\BA Deep Space Voyage to High Latitudes over the Solar Poles\b
Expanding gasses from the solar corona dominate the properties of interstellar space over a large region around the Sun known as the heliosphere. Launched by the Space Shuttle Discovery in October, 1990, Ulysses flew by Jupiter in February, 1992, where a gravity assist manoeuvre placed the spacecraft in a unique solar polar orbit, allowing it to fly over the south pole of the Sun in 1994 and over the north pole in 1995. The Ulysses mission explored for the first time the high latitude heliosphere away from the plane of the ecliptic.
The primary results of the mission have been to discover at these high latitudes the properties of the solar corona, the solar wind, the heliospheric magnetic field, solar energetic particles, galactic cosmic rays, solar radio bursts and plasma waves. Other investigations include study of cosmic dust, gamma ray bursts, and studies of the Jovian magnetosphere obtained during the Jupiter fly-by. With the first phase of its mission successfully completed, Ulysses has now embarked on a second orbit of the Sun; it will for the first time investigate the high latitude properties of the solar wind during the maximum of the solar activity cycle. Ulysses is a joint project of the European Space Agency (ESA) and the National Aeronautics and Space Administration (NASA)..
#
"Ulysses -- Scientific Objectives",630,0,0,0
The primary mission of the Ulysses spacecraft is to characterize the heliosphere as a function of solar latitude. The heliosphere is the vast region of interplanetary space occupied by the Sun's atmosphere and dominated by the outflow of the solar wind. The periods of primary scientific interest are when Ulysses is at or higher than 70 degrees latitude at both the Sun's south and north poles. On 26 June 1994, Ulysses reached 70 degrees south. There it began a four-month observation from high latitudes of the complex forces at work in the Sun's outer atmosphere -- the corona.
Scientists have long studied the Sun from Earth using Earth-based sensors. More recently, solar studies have been conducted from spaceborne platforms; however, these investigations have been mostly from the ecliptic plane (the plane in which most of the planets travel around the Sun) and no previous spacecraft have reached solar latitudes higher than 32 degrees. Now that Ulysses high latitude data is available, scientists from the joint National Aeronautics and Space Administration (NASA)-European Space Agency (ESA) mission are obtaining new and better understanding of the processes going on at high solar latitudes.
Scientists have long been aware of differences between the polar regions of the Sun and lower latitudes. Sun spots are only seen at lower latitudes, and photographs of the solar corona take during solar eclipses often showed dark regions over the poles. The solar corona consists of hot gasses (over 1,000,000 degrees); at this temperature the graviational field of the Sun can not prevent escape of coronal gas as the solar wind.
However, the Sun has a global magnetic field. Many of the solar magnetic field lines that leave the solar surface return to the surface, but some of the field lines, particularly those over the poles, extend deep into interplanetary space. The solar wind expands into interplanetary space along these field lines, and the regions (known as coronal holes) of the corona from which the hot gas escapes are dark because of the low gas density.
The properties of the Sun's polar magnetic field are poorly understood, and it has an important influence on the escape of the solar wind. The complex processes that heat and accelerate the solar wind are not well understood, and Ulysses observations over the poles should provide important new information on how the solar wind expands from the Sun that will aid scientists in understanding these processes.
The magnetic field also exerts a crucial influence on matter arriving near the Sun from the Milky Way galaxy and from the nearby interstellar medium. Incoming cosmic rays are subjected to forces exerted by the magnetic field. The structure of the Sun's magnetic field is thought to favor entry of cosmic rays by way of the Sun's polar regions.
Scientists hope that Ulysses can shed some light on the extent to which the galactic cosmic rays observed at Earth use this route and on the ways in which their properties are modified as a result. Scientists also hope to gain more knowledge of the intensity and properties of the cosmic rays far from the Sun.
#
"Ulysses' Second Solar Orbit",631,0,0,0
With the spacecraft and its scientific payload in excellent condition, Ulysses has now embarked on its second orbit of the Sun. The ultimate goal of this next phase of the mission is the study of the Sun's polar regions under conditions of high solar activity, culminating in polar passes in 2000 and 2001 (south polar pass: Sep 2000 to Jan 2001; north polar pass: Sep to Dec 2001).
Much before this, however, with Ulysses descending slowly in latitude after the northern polar pass, there is a unique opportunity to make coordinated observations with ESA's SOHO spacecraft, which carries an extensive complement of experiments dedicated to studying the Sun's corona and the solar wind. The period around aphelion (1997-98) will also be of great interest.
During this interval Ulysses will spend many months close to the ecliptic at almost constant radial distance (~ 5 AU) from the Sun, enabling the temporal evolution of many interplanetary phenomena to be studied free of concern about spatial variations. Measurements of interstellar pick-up ions (atoms of interstellar gas that have become singly ionised) will also benefit from the "dwell" at moderately large heliocentric distance, since many species are unable to reach the inner solar system.
Ulysses' out-of-ecliptic orbit has a period of 6.2 years, corresponding to approximately half a solar cycle. As noted above, a fortuitous consequence of this is that the high-latitude passes of the second solar orbit will occur close to the maximum of solar cycle 23. The conditions in the polar regions are expected to be dramatically different from those encountered during the prime mission.
In particular, the rather simple configuration of the corona found at solar minimum, with large coronal holes over the polar caps, will have been replaced by a much more complex arrangement, probably including high-latitude streamers. Transient events (solar flares, coronal mass ejections, etc.) related to the increase in solar activity will dominate, greatly disturbing the underlying structure of the solar wind and influencing the transport of cosmic rays and energetic solar particles.
Ulysses is clearly in a unique position, literally and figuratively, to study the evolution of the three-dimensional heliosphere from the current solar minimum to maximum activity conditions. No other space mission in the foreseeable future will address these goals.
#
"Voyager 1 Now Most Distant Human-Made Object in Space",632,0,0,0
February 13, 1998
In a dark, cold, vacant neighborhood near the very edge of our Solar System, the Voyager 1 spacecraft is set to break another record and become the explorer that has traveled farthest from home.
At approximately 5:10 p.m. EST on Feb. 17, 1998, Voyager 1, launched more than two decades ago, will cruise beyond the Pioneer 10 spacecraft and become the most distant human-created object in space, at 6.5 billion miles (10.4 billion kilometers) from Earth. The two are headed in almost opposite directions away from the Sun.
"For 25 years, the Pioneer 10 spacecraft led the way, pressing the frontiers of exploration, and now the baton is being passed from Pioneer 10 to Voyager 1 to continue exploring where no one has gone before," said Dr. Edward C. Stone, Voyager project scientist and Director of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
"At almost 70 times farther from the Sun than the Earth, Voyager 1 is at the very edge of the Solar System. The Sun there is only 1/5,000th as bright as here on Earth, so it is extremely cold, and there is very little solar energy to keep the spacecraft warm or to provide electrical power.
The reason we can continue to operate at such great distances from the Sun is because we have radioisotope thermal electric generators (RTGs) on the spacecraft that create electricity and keep the spacecraft operating," Stone said. "The fact that the spacecraft is still returning data is a remarkable technical achievement."
Voyager 1 was launched from Cape Canaveral on Sept. 5, 1977. The spacecraft encountered Jupiter on March 5, 1979, and Saturn on Nov. 12, 1980. Then, because its trajectory was designed to fly close to Saturn's large moon Titan, Voyager 1's path was bent northward by Saturn's gravity, sending the spacecraft out of the ecliptic plane -- the plane in which all the planets except Pluto orbit the Sun.
Launched on March 2, 1972, the Pioneer 10 mission officially ended on March 31, 1997. However, NASA's Ames Research Center, Moffett Field, CA, intermittently receives science data from Pioneer as part of a training program for flight controllers of the Lunar Prospector spacecraft now orbiting the Moon.
"The Voyager mission today presents an unequalled technical challenge. The spacecraft are now so far from home that it takes nine hours and 36 minutes for a radio signal traveling at the speed of light to reach Earth," said Ed B. Massey, project manager for the Voyager Interstellar Mission at JPL. "That signal, produced by a 20 watt radio transmitter, is so faint that the amount of power reaching our antennas is 20 billion times smaller than the power of a digital watch battery."
Having completed their planetary explorations, Voyager 1 and its twin, Voyager 2, are studying the environment of space in the outer Solar System. Although beyond the orbits of all the planets, the spacecraft still are well within the boundary of the Sun's magnetic field, called the heliosphere. Science instruments on both spacecraft sense signals that scientists believe are coming from the outermost edge of the heliosphere, known as the heliopause.
The heliosphere results from the Sun's emitting a steady flow of electrically charged particles called the solar wind. As the solar wind expands supersonically into space in all directions, it creates a magnetized bubble -- the heliosphere -- around the Sun.
Eventually, the solar wind encounters the electrically charged particles and magnetic field in the interstellar gas. In this zone the solar wind abruptly slows down from supersonic to subsonic speed, creating a termination shock. Before the spacecraft travel beyond the heliopause into interstellar space, they will pass through this termination shock.
"The data coming back from Voyager now suggest that we may pass through the termination shock in the next three to five years," Stone said. "If that's the case, then one would expect that within 10 years or so we would actually be very close to penetrating the heliopause itself and entering into interstellar space for the first time."
Reaching the termination shock and heliopause will be major milestones for the mission because no spacecraft have been there before and the Voyagers will gather the first direct evidence of their structure. Encountering the termination shock and heliopause has been a long-sought goal for many space physicists, and exactly where these two boundaries are located and what they are like still remains a mystery.
Science data are returned to Earth in real-time to the 34-meter Deep Space Network antennas located in California, Australia and Spain. Both spacecraft have enough electricity and attitude control propellant to continue operating until about 2020, when electrical power produced by the RTGs will no longer support science instrument operation. At that time, Voyager 1 will be almost 150 times farther from the Sun than the Earth -- almost 14 billion miles (more than 20 billion kilometers) away.
On Feb. 17, Voyager 1 will be departing the Solar System at a speed of 39,000 miles per hour (17.4 kilometers per second ). At the same time, Voyager 2 will be 5.1 billion miles (8.1 billion kilometers) from Earth and is departing the Solar System at a speed of 35,000 miles per hour (15.9 kilometers per second).
The Voyager spacecraft will be the third and fourth human artifacts to escape entirely from the solar system. Pioneers 10 and 11, which preceded Voyager in outstripping the gravitational attraction of the Sun, both carried small metal plaques identifying their time and place of origin for the benefit of any other spacefarers that might find them in the distant future. With this example before them, NASA placed a more ambitious message aboard Voyager 1 and 2-a kind of time capsule, intended to communicate a story of our world to extraterrestrials.
The Voyager message is carried by a phonograph record-a 12-inch gold-plated copper disk containing sounds and images selected to portray the diversity of life and culture on Earth. The contents of the record were selected for NASA by a committee chaired by Carl Sagan of Cornell University. Dr. Sagan and his associates assembled 115 images and a variety of natural sounds, such as those made by surf, wind and thunder, birds, whales, and other animals. To this they added musical selections form different cultures and eras, and spoken greetings from Earth-people in fifty-five languages, and printed messaged from President Carter and U.N. Secretary General Waldheim. Each record is encased in a protective aluminum jacket, together with a cartridge and a needle.
Instructions, in symbolic language, explain the origin of the spacecraft and indicate how the record is to be played. The 115 images are encoded in analog form. The remainder of the record is in audio, designed to be played at 16-2/3 revolutions per second. It contains the spoken greetings, beginning with Akkadian, which was spoken in Sumer about six thousand years ago, and ending with Wu, a modern Chinese dialect. Following the section on the sounds of Earth, there is an eclectic 90-minute selection of music, including both Eastern and Western classics and a variety of ethnic music.
Once the Voyager spacecraft leave the solar system (by 1990, both will be beyond the orbit of Pluto), they will find themselves in empty space. It will be forty thousand years before they make a close approach to any other planetary system. As Carl Sagan has noted, "The spacecraft will be encountered and the record played only if there are advanced spacefaring civilizations in interstellar space. But the launching of this bottle into the cosmic ocean says something very hopeful about life on this planet."
\BInformation About the Earth and Its Inhabitants Included on the Record\b:
\B╖\b Images
\B╖\b Sounds
\B╖\b Languages
\B╖\b Music
#
"Voyager Mission",634,0,0,0
\JVoyager Planetary Mission Fact Sheet\j
\JVoyager Mission - History\j
\JVoyager Operations\j
\JVoyager - Jupiter Encounter\j
\JVoyager - Saturn Encounter\j
\JVoyager - Uranus Encounter\j
\JVoyager - Neptune Encounter\j
\JVoyager's Interstellar Mission\j
#
"Voyager Planetary Mission Fact Sheet",635,0,0,0
The twin spacecraft Voyager 1 and Voyager 2 were launched by NASA in separate months in the summer of 1977 from Cape Canaveral, Florida. As originally designed, the Voyagers were to conduct closeup studies of Jupiter and Saturn, Saturn's rings, and the larger moons of the two planets.
To accomplish their two-planet mission, the spacecraft were built to last five years. But as the mission went on, and with the successful achievement of all its objectives, the additional flybys of the two outermost giant planets, Uranus and Neptune, proved possible -- and irresistible to mission scientists and engineers at the Voyagers' home at the \IJet Propulsion Laboratory\i in Pasadena, California.
As the spacecraft flew across the solar system, remote-control reprogramming was used to endow the Voyagers with greater capabilities than they possessed when they left the Earth. Their two-planet mission became four. Their five-year lifetimes stretched to 12 and more.
Eventually, between them, Voyager 1 and 2 would explore all the giant outer planets of our solar system, 48 of their moons, and the unique systems of rings and magnetic fields those planets possess.
Had the Voyager mission ended after the Jupiter and Saturn flybys alone, it still would have provided the material to rewrite astronomy textbooks. But having doubled their already ambitious itineraries, the Voyagers returned to Earth information over the years that has revolutionized the science of planetary astronomy, helping to resolve key questions while raising intriguing new ones about the origin and evolution of the planets in our solar system.
#
"Voyager Mission - History",636,0,0,0
The Voyager mission was designed to take advantage of a rare geometric arrangement of the outer planets in the late 1970s and the 1980s which allowed for a four-planet tour for a minimum of propellant and trip time. This layout of Jupiter, Saturn, Uranus and Neptune, which occurs about every 175 years, allows a spacecraft on a particular flight path to swing from one planet to the next without the need for large onboard propulsion systems. The flyby of each planet bends the spacecraft's flight path and increases its velocity enough to deliver it to the next destination. Using this "gravity assist" technique, first demonstrated with NASA's Mariner 10 Venus/Mercury mission in 1973-74, the flight time to Neptune was reduced from 30 years to 12.
While the four-planet mission was known to be possible, it was deemed to be too expensive to build a spacecraft that could go the distance, carry the instruments needed and last long enough to accomplish such a long mission. Thus, the Voyagers were funded to conduct intensive flyby studies of Jupiter and Saturn only. More than 10,000 trajectories were studied before choosing the two that would allow close flybys of Jupiter and its large moon Io, and Saturn and its large moon Titan; the chosen flight path for Voyager 2 also preserved the option to continue on to Uranus and Neptune.
From the NASA \IKennedy Space Center\i at Cape Canaveral, Florida, Voyager 2 was launched first, on August 20, 1977; Voyager 1 was launched on a faster, shorter trajectory on September 5, 1977. Both spacecraft were delivered to space aboard Titan-Centaur expendable rockets.
The prime Voyager mission to Jupiter and Saturn brought Voyager 1 to Jupiter on March 5, 1979, and Saturn on November 12, 1980, followed by Voyager 2 to Jupiter on July 9, 1979, and Saturn on August 25, 1981.
Voyager 1's trajectory, designed to send the spacecraft closely past the large moon Titan and behind Saturn's rings, bent the spacecraft's path inexorably northward out of the ecliptic plane -- the plane in which most of the planets orbit the Sun. Voyager 2 was aimed to fly by Saturn at a point that would automatically send the spacecraft in the direction of Uranus.
After Voyager 2's successful Saturn encounter, it was shown that Voyager 2 would likely be able to fly on to Uranus with all instruments operating. NASA provided additional funding to continue operating the two spacecraft and authorized JPL to conduct a Uranus flyby. Subsequently, NASA also authorized the Neptune leg of the mission, which was renamed the Voyager Neptune Interstellar Mission.
Voyager 2 encountered Uranus on January 24, 1986, returning detailed photos and other data on the planet, its moons, magnetic field and dark rings. Voyager 1, meanwhile, continues to press outward, conducting studies of interplanetary space. Eventually, its instruments may be the first of any spacecraft to sense the heliopause -- the boundary between the end of the Sun's magnetic influence and the beginning of interstellar space.
Following Voyager 2's closest approach to Neptune on August 25, 1989, the spacecraft flew southward, below the ecliptic plane and onto a course that will take it, too, to interstellar space. Reflecting the Voyagers' new transplanetary destinations, the project is now known as the Voyager Interstellar Mission.
Voyager 1 is now leaving the solar system, rising above the ecliptic plane at an angle of about 35 degrees at a rate of about 520 million kilometers (about 320 million miles) a year. Voyager 2 is also headed out of the solar system, diving below the ecliptic plane at an angle of about 48 degrees and a rate of about 470 million kilometers (about 290 million miles) a year.
Both spacecraft will continue to study ultraviolet sources among the stars, and the fields and particles instruments aboard the Voyagers will continue to search for the boundary between the Sun's influence and interstellar space. The Voyagers are expected to return valuable data for two or three more decades. Communications will be maintained until the Voyagers' nuclear power sources can no longer supply enough electrical energy to power critical subsystems.
The cost of the Voyager 1 and 2 missions -- including launch, mission operations from launch through the Neptune encounter and the spacecraft's nuclear batteries (provided by the Department of Energy) -- is $865 million. NASA budgeted an additional $30 million to fund the Voyager Interstellar Mission for two years following the Neptune encounter.
#
"Voyager Operations",637,0,0,0
Voyagers 1 and 2 are identical spacecraft. Each is equipped with instruments to conduct 10 different experiments. The instruments include television cameras, infrared and ultraviolet sensors, magnetometers, plasma detectors, and cosmic-ray and charged-particle sensors. In addition, the spacecraft radio is used to conduct experiments.
The Voyagers travel too far from the Sun to use solar panels; instead, they were equipped with power sources called radioisotope thermoelectric generators (RTGs). These devices, used on other deep space missions, convert the heat produced from the natural radioactive decay of plutonium into electricity to power the spacecraft instruments, computers, radio and other systems.
The spacecraft are controlled and their data returned through the \IDeep Space Network (DSN)\i, a global spacecraft tracking system operated by JPL for NASA. DSN antenna complexes are located in California's Mojave Desert; near Madrid, Spain; and in Tidbinbilla, near Canberra, Australia.
The Voyager project manager for the Interstellar Mission is George P. Textor of JPL. The Voyager project scientist is Dr. Edward C. Stone of the California Institute of Technology. The assistant project scientist for the Jupiter flyby was Dr. Arthur L. Lane, followed by Dr. Ellis D. Miner for the Saturn, Uranus and Neptune encounters. Both are with JPL.
#
"Voyager - Jupiter Encounter",638,0,0,0
Voyager 1 made its closest approach to Jupiter on March 5, 1979, and Voyager 2 followed with its closest approach occurring on July 9, 1979. The first spacecraft flew within 206,700 kilometers (128,400 miles) of the planet's cloud tops, and Voyager 2 came within 570,000 kilometers (350,000 miles).
Jupiter is the largest planet in the solar system, composed mainly of hydrogen and helium, with small amounts of methane, ammonia, water vapor, traces of other compounds and a core of melted rock and ice. Colorful latitudinal bands and atmospheric clouds and storms illustrate Jupiter's dynamic weather system. The giant planet is now known to possess 16 moons. The planet completes one orbit of the Sun each 11.8 years and its day is 9 hours, 55 minutes.
Although astronomers had studied Jupiter through telescopes on Earth for centuries, scientists were surprised by many of the Voyager findings.
The Great Red Spot was revealed as a complex storm moving in a counterclockwise direction. An array of other smaller storms and eddies were found throughout the banded clouds.
Discovery of active volcanism on the satellite Io was easily the greatest unexpected discovery at Jupiter. It was the first time active volcanoes had been seen on another body in the solar system. Together, the Voyagers observed the eruption of nine volcanoes on Io, and there is evidence that other eruptions occurred between the Voyager encounters.
Plumes from the volcanoes extend to more than 300 kilometers (190 miles) above the surface. The Voyagers observed material ejected at velocities up to a kilometer per second.
Io's volcanoes are apparently due to heating of the satellite by tidal pumping. Io is perturbed in its orbit by Europa and Ganymede, two other large satellites nearby, then pulled back again into its regular orbit by Jupiter. This tug-of-war results in tidal bulging as great as 100 meters (330 feet) on Io's surface, compared with typical tidal bulges on Earth of one meter (three feet).
It appears that volcanism on Io affects the entire jovian system, in that it is the primary source of matter that pervades Jupiter's magnetosphere -- the region of space surrounding the planet influenced by the jovian magnetic field. Sulfur, oxygen and sodium, apparently erupted by Io's many volcanoes and sputtered off the surface by impact of high-energy particles, were detected as far away as the outer edge of the magnetosphere millions of miles from the planet itself.
Europa displayed a large number of intersecting linear features in the low-resolution photos from Voyager 1. At first, scientists believed the features might be deep cracks, caused by crustal rifting or tectonic processes. The closer high-resolution photos from Voyager 2, however, left scientists puzzled: The features were so lacking in topographic relief that as one scientist described them, they "might have been painted on with a felt marker." There is a possibility that Europa may be internally active due to tidal heating at a level one-tenth or less than that of Io. Europa is thought to have a thin crust (less than 30 kilometers or 18 miles thick) of water ice, possibly floating on a 50-kilometer-deep (30-mile) ocean.
Ganymede turned out to be the largest moon in the solar system, with a diameter measuring 5,276 kilometers (3,280 miles). It showed two distinct types of terrain -- cratered and grooved -- suggesting to scientists that Ganymede's entire icy crust has been under tension from global tectonic processes.
Callisto has a very old, heavily cratered crust showing remnant rings of enormous impact craters. The largest craters have apparently been erased by the flow of the icy crust over geologic time. Almost no topographic relief is apparent in the ghost remnants of the immense impact basins, identifiable only by their light color and the surrounding subdued rings of concentric ridges.
A faint, dusty ring of material was found around Jupiter. Its outer edge is 129,000 kilometers (80,000 miles) from the center of the planet, and it extends inward about 30,000 kilometers (18,000 miles).
Two new, small satellites, Adrastea and Metis, were found orbiting just outside the ring. A third new satellite, Thebe, was discovered between the orbits of Amalthea and Io.
Jupiter's rings and moons exist within an intense radiation belt of electrons and ions trapped in the planet's magnetic field. These particles and fields comprise the jovian magnetosphere, or magnetic environment, which extends three to seven million kilometers toward the Sun, and stretches in a windsock shape at least as far as Saturn's orbit -- a distance of 750 million kilometers (460 million miles).
As the magnetosphere rotates with Jupiter, it sweeps past Io and strips away about 1,000 kilograms (one ton) of material per second. The material forms a torus, a doughnut-shaped cloud of ions that glow in the ultraviolet. The torus's heavy ions migrate outward, and their pressure inflates the jovian more energetic sulfur and oxygen ions fall along the magnetic field into the planet's atmosphere, resulting in auroras.
Io acts as an electrical generator as it moves through Jupiter's magnetic field, developing 400,000 volts across its diameter and generating an electric current of 3 million amperes that flows along the magnetic field to the planet's ionosphere.
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"Voyager - Saturn Encounter",639,0,0,0
The Voyager 1 and 2 Saturn flybys occurred nine months apart, with the closest approaches falling on November 12 and August 25, 1981. Voyager 1 flew within 64,200 kilometers (40,000 miles) of the cloud tops, while Voyager 2 came within 41,000 kilometers (26,000 miles).
Saturn is the second largest planet in the solar system. It takes 29.5 Earth years to complete one orbit of the Sun, and its day was clocked at 10 hours, 39 minutes. Saturn is known to have at least 17 moons and a complex ring system. Like Jupiter, Saturn is mostly hydrogen and helium. Its hazy yellow hue was found to be marked by broad atmospheric banding similar to but much fainter than that found on Jupiter. Close scrutiny by Voyager's imaging systems revealed long-lived ovals and other atmospheric features generally smaller than those on Jupiter.
Perhaps the greatest surprises and the most puzzles were found by the Voyagers in Saturn's rings. It is thought that the rings formed from larger moons that were shattered by impacts of comets and meteoroids. The resulting dust and boulder- to house-size particles have accumulated in a broad plane around the planet varying in density.
The irregular shapes of Saturn's eight smallest moons indicates that they too are fragments of larger bodies. Unexpected structure such as kinks and spokes were found in addition to thin rings and broad, diffuse rings not observed from Earth. Much of the elaborate structure of some of the rings is due to the gravitational effects of nearby satellites. This phenomenon is most obviously demonstrated by the relationship between the F-ring and two small moons that "shepherd" the ring material. The variation in the separation of the moons from the ring may the ring's kinked appearance. Shepherding moons were also found by Voyager 2 at Uranus.
Radial, spoke-like features in the broad B-ring were found by the Voyagers. The features are believed to be composed of fine, dust-size particles. The spokes were observed to form and dissipate in time-lapse images taken by the Voyagers. While electrostatic charging may create spokes by levitating dust particles above the ring, the exact cause of the formation of the spokes is not well understood.
Winds blow at extremely high speeds on Saturn -- up to 1,800 kilometers per hour (1,100 miles per hour). Their primarily easterly direction indicates that the winds are not confined to the top cloud layer but must extend at least 2,000 kilometers (1,200 miles) downward into the atmosphere. The characteristic temperature of the atmosphere is 95 kelvins.
Saturn holds a wide assortment of satellites in its orbit, ranging from Phoebe, a small moon that travels in a retrograde orbit and is probably a captured asteroid, to Titan, the planet-sized moon with a thick nitrogen-methane atmosphere. Titan's surface temperature and pressure are 94 kelvins (-292 Fahrenheit) and 1.5 atmospheres. Photochemistry converts some atmospheric methane to other organic molecules, such as ethane, that is thought to accumulate in lakes or oceans. Other more complex hydrocarbons form the haze particles that eventually fall to the surface, coating it with a thick layer of organic matter. The chemistry in Titan's atmosphere may strongly resemble that which occurred on Earth before life evolved.
The most active surface of any moon seen in the Saturn system was that of Enceladus. The bright surface of this moon, marked by faults and valleys, showed evidence of tectonically induced change. Voyager 1 found the moon Mimas scarred with a crater so huge that the impact that caused it nearly broke the satellite apart.
Saturn's magnetic field is smaller than Jupiter's, extending only one or two million kilometers. The axis of the field is almost perfectly aligned with the rotation axis of the planet.
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"Voyager - Uranus Encounter",640,0,0,0
In its first solo planetary flyby, Voyager 2 made its closest approach to Uranus on January 24, 1986, coming within 81,500 kilometers (50,600 miles) of the planet's cloud tops.
Uranus is the third largest planet in the solar system. It orbits the Sun at a distance of about 2.8 billion kilometers (1.7 billion miles) and completes one orbit every 84 years. The length of a day on Uranus as measured by Voyager 2 is 17 hours, 14 minutes.
Uranus is distinguished by the fact that it is tipped on its side. Its unusual position is thought to be the result of a collision with a planet-sized body early in the solar system's history. Given its odd orientation, with its polar regions exposed to sunlight or darkness for long periods, scientists were not sure what to expect at Uranus.
Voyager 2 found that one of the most striking influences of this sideways position is its effect on the tail of the magnetic field, which is itself tilted 60 degrees from the planet's axis of rotation. The magnetotail was shown to be twisted by the planet's rotation into a long corkscrew shape behind the planet.
The presence of a magnetic field at Uranus was not known until Voyager's arrival. The intensity of the field is roughly comparable to that of Earth's, though it varies much more from point to point because of its large offset from the center of Uranus. The peculiar orientation of the magnetic field suggests that the field is generated at an intermediate depth in the interior where the pressure is high enough for water to become electrically conducting.
Radiation belts at Uranus were found to be of an intensity similar to those at Saturn. The intensity of radiation within the belts is such that irradiation would quickly darken (within 100,000 years) any methane trapped in the icy surfaces of the inner moons and ring particles. This may have contributed to the darkened surfaces of the moons and ring particles, which are almost uniformly gray in color.
A high layer of haze was detected around the sunlit pole, which also was found to radiate large amounts of ultraviolet light, a phenomenon dubbed "dayglow." The average temperature is about 60 kelvins (-350 degrees Fahrenheit). Surprisingly, the illuminated and dark poles, and most of the planet, show nearly the same temperature at the cloud tops.
Voyager found 10 new moons, bringing the total number to 15. Most of the new moons are small, with the largest measuring about 150 kilometers (about 90 miles) in diameter.
The moon Miranda, innermost of the five large moons, was revealed to be one of the strangest bodies yet seen in the solar system. Detailed images from Voyager's flyby of the moon showed huge fault canyons as deep as 20 kilometers (12 miles), terraced layers, and a mixture of old and young surfaces. One theory holds that Miranda may be a reaggregration of material from an earlier time when the moon was fractured by an violent impact.
The five large moons appear to be ice-rock conglomerates like the satellites of Saturn. Titania is marked by huge fault systems and canyons indicating some degree of geologic, probably tectonic, activity in its history. Ariel has the brightest and possibly youngest surface of all the Uranian moons and also appears to have undergone geologic activity that led to many fault valleys and what seem to be extensive flows of icy material. Little geologic activity has occurred on Umbriel or Oberon, judging by their old and dark surfaces.
All nine previously known rings were studied by the spacecraft and showed the Uranian rings to be distinctly different from those at Jupiter and Saturn. The ring system may be relatively young and did not form at the same time as Uranus. Particles that make up the rings may be remnants of a moon that was broken by a high-velocity impact or torn up by gravitational effects.
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"Voyager - Neptune Encounter",641,0,0,0
When Voyager flew within 5,000 kilometers (3,000 miles) of Neptune on August 25, 1989, the planet was the most distant member of the solar system from the Sun. (Pluto once again will become most distant in 1999.)
Neptune orbits the Sun every 165 years. It is the smallest of our solar system's gas giants. Neptune is now known to have eight moons, six of which were found by Voyager. The length of a Neptunian day has been determined to be 16 hours, 6.7 minutes.
Even though Neptune receives only three percent as much sunlight as Jupiter does, it is a dynamic planet and surprisingly showed several large, dark spots reminiscent of Jupiter's hurricane-like storms. The largest spot, dubbed the Great Dark Spot, is about the size of Earth and is similar to the Great Red Spot on Jupiter. A small, irregularly shaped, eastward-moving cloud was observed "scooting" around Neptune every 16 hours or so; this "scooter," as Voyager scientists called it, could be a cloud plume rising above a deeper cloud deck.
Long, bright clouds, similar to cirrus clouds on Earth, were seen high in Neptune's atmosphere. At low northern latitudes, Voyager captured images of cloud streaks casting their shadows on cloud decks below.
The strongest winds on any planet were measured on Neptune. Most of the winds there blow westward, or opposite to the rotation of the planet. Near the Great Dark Spot, winds blow up to 2,000 kilometers (1,200 miles) an hour.
The magnetic field of Neptune, like that of Uranus, turned out to be highly tilted -- 47 degrees from the rotation axis and offset at least 0.55 radii (about 13,500 kilometers or 8,500 miles) from the physical center. Comparing the magnetic fields of the two planets, scientists think the extreme orientation may be characteristic of flows in the interiors of both Uranus and Neptune -- and not the result in Uranus's case of that planet's sideways orientation, or of any possible field reversals at either planet. Voyager's studies of radio waves caused by the magnetic field revealed the length of a Neptunian day. The spacecraft also detected auroras, but much weaker than those on Earth and other planets.
Triton, the largest of the moons of Neptune, was shown to be not only the most intriguing satellite of the Neptunian system, but one of the most interesting in all the solar system. It shows evidence of a remarkable geologic history, and Voyager 2 images showed active geyser-like eruptions spewing invisible nitrogen gas and dark dust particles several kilometers into the tenuous atmosphere. Triton's relatively high density and retrograde orbit offer strong evidence that Triton is not an original member of Neptune's family but is a captured object. If that is the case, tidal heating could have melted Triton in its originally eccentric orbit, and the moon might even have been liquid for as long as one billion years after its capture by Neptune.
An extremely thin atmosphere extends about 800 kilometer (500 miles) above Triton's surface. Nitrogen ice particles may form thin clouds a few kilometers above the surface. The atmospheric pressure at the surface is about 14 microbars, 1/70,000th the surface pressure on Earth. The surface temperature is about 38 kelvins (-391 degrees Fahrenheit) the coldest temperature of any body known in the solar system.
The new moons found at Neptune by Voyager are all small and remain close to Neptune's equatorial plane. Names for the new moons were selected from mythology's water deities by the International Astronomical Union, they are: Naiad, Thalassa, Despina, Galatea, Larissa, Proteus.
Voyager 2 solved many of the questions scientists had about Neptune's rings. Searches for "ring arcs," or partial rings, showed that Neptune's rings actually are complete, but are so diffuse and the material in them so fine that they could not be fully resolved from Earth. From the outermost in, the rings have been designated Adams, Plateau, Le Verrier and Galle.
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"Voyager Information",642,0,0,0
The Voyager mission was officially approved in May 1972, has received the dedicated efforts of many skilled personnel for over two decades, and has returned more new knowledge about the outer planets than had existed in all of the preceding history of astronomy and planetary science. And the two Voyager machines are still performing like champs.
It must come as no surprise that there are many remarkable, "gee-whiz" facts associated with the various aspects of the Voyager mission. These tidbits have been summarized below in appropriate categories. Several may seem difficult to believe, but they are all true and accurate.
\BOverall Mission\b
The total cost of the Voyager mission from May 1972 through the Neptune encounter (including launch vehicles, nuclear-power-source RTGs, and DSN tracking support) is 865 million dollars. At first, this may sound very expensive, but the fantastic returns are a bargain when we place the costs in the proper perspective. It is important to realize that:
1. on a per-capita basis, this is only 20 cents per U.S. resident per year, or roughly half the cost of one candy bar each year since project inception.
2. the daily interest on the U.S. national debt is a major fraction of the entire cost of Voyager.
A total of 11,000 workyears will have been devoted to the Voyager project through the Neptune encounter. This is equivalent to one-third the amount of effort estimated to complete the great pyramid at Giza to King Cheops.
A total of five trillion bits of scientific data will have been returned to Earth by both Voyager spacecraft at the completion of the Neptune encounter. This represents enough bits to encode over 6000 complete sets of the Encyclopedia Brittanica, and is equivalent to about 1000 bits of information provided to each person on Earth.
The sensitivity of our deep-space tracking antennas located around the world is truly amazing. The antennas must capture Voyager information from a signal so weak that the power striking the antenna is only 10 exponent -16 watts (1 part in 10 quadrillion). A modern-day electronic digital watch operates at a power level 20 billion times greater than this feeble level.
\BVoyager Spacecraft\b
Each Voyager spacecraft comprises 65,000 individual parts. Many of these parts have a large number of "equivalent" smaller parts such as transistors. One computer memory alone contains over one million equivalent electronic parts, with each spacecraft containing some five million equivalent parts. Since a color TV set contains about 2500 equivalent parts, each Voyager has the equivalent electronic circuit complexity of some 2000 color TV sets.
Like the HAL computer aboard the ship Discovery from the famous science fiction story \I2001: A Space Odyssey\i, each Voyager is equipped with computer programming for autonomous fault protection. The Voyager system is one of the most sophisticated ever designed for a deep-space probe. There are seven top-level fault protection routines, each capable of covering a multitude of possible failures. The spacecraft can place itself in a safe state in a matter of only seconds or minutes, an ability that is critical for its survival when round-trip communication times for Earth stretch to several hours as the spacecraft journeys to the remote outer solar system.
Both Voyagers were specifically designed and protected to withstand the large radiation dosage during the Jupiter swing-by. This was accomplished by selecting radiation-hardened parts and by shielding very sensitive parts. An unprotected human passenger riding aboard Voyager 1 during its Jupiter encounter would have received a radiation dose equal to one thousand times the lethal level.
The Voyager spacecraft can point its scientific instruments on the scan platform to an accuracy of better than one-tenth of a degree. This is comparable to bowling strike-after-strike \Iad infinitum\i, assuming that you must hit within one inch of the strike pocket every time. Such precision is necessary to properly center the narrow-angle picture whose square field-of-view would be equivalent to the width of a bowling pin.
To avoid smearing in Voyager's television pictures, spacecraft angular rates must be extremely small to hold the cameras as steady as possible during the exposure time. Each spacecraft is so steady that angular rates are typically 15 times slower than the motion of a clock's hour hand. But even this will not be quite steady enough at Neptune, where light levels are 900 times fainter than those on Earth. Spacecraft engineers have already devised ways to make Voyager 30 times steadier than the hour hand on a clock.
The electronics and heaters aboard each nearly one-ton Voyager spacecraft can operate on only 400 watts of power, or roughly one-fourth that used by an average residential home in the western United States.
A set of small thrusters provides Voyager with the capability for attitude control and trajectory correction. Each of these tiny assemblies has a thrust of only three ounces. In the absence of friction, on a level road, it would take nearly six hours to accelerate a large car up to a speed of 48 km/h (30 mph) using one of the thrusters.
The Voyager scan platform can be moved about two axes of rotation. A thumb-sized motor in the gear train drive assembly (which turns 9000 revolutions for each single revolution of the scan platform) will have rotated five million revolutions from launch through the Neptune encounter. This is equivalent to the number of automobile crankshaft revolutions during a trip of 2725 km (1700 mi).
The Voyager gyroscopes can detect spacecraft angular motion as little as one ten-thousandth of a degree. The Sun's apparent motion in our sky moves over 40 times that amount in just one second.
The tape recorder aboard each Voyager has been designed to record and playback a great deal of scientific data. The tape head should not begin to wear out until the tape has been moved back and forth through a distance comparable to that across the United States. Imagine playing a two-hour video cassette on your home VCR once a day for the next 22 years, without a failure.
The Voyager magnetometers are mounted on a frail, spindly, fiberglass boom that was unfurled from a two-foot-long can shortly after the spacecraft left Earth. After the boom telescoped and rotated out of the can to an extension of nearly 13 meters (43 feet), the orientations of the magnetometer sensors were controlled to an accuracy better than two degrees.
\BNavigation\b
Each Voyager used the enormous gravity field of Jupiter to be hurled on to Saturn, experiencing a Sun-relative speed increase of roughly 35,700 mph. As total energy within the solar system must be conserved, Jupiter was initially slowed in its solar orbit---but by only one foot per trillion years. Additional gravity-assist swing-bys of Saturn and Uranus were necessary for Voyager 2 to complete its Grand Tour flight to Neptune, reducing the trip time by nearly twenty years when compared to the unassisted Earth-to-Neptune route.
The Voyager delivery accuracy at Neptune of 100 km (62 mi), divided by the trip distance or arc length traveled of 7,128,603,456 km (4,429,508,700 mi), is equivalent to the feat of sinking a 3630 km (2260 mi) golf putt, assuming that the golfer can make a few illegal fine adjustments while the ball is rolling across this incredibly long green.
Voyager's fuel efficiency (in terms of mpg) is quite impressive. Even though most of the launch vehicle's 700 ton weight is due to rocket fuel, Voyager 2's great travel distance of 7.1 billion km (4.4 billion mi) from launch to Neptune results in a fuel economy of about 13,000 km per liter (30,000 mi per gallon). As Voyager 2 streaks by Neptune and coasts out of the solar system, this economy will get better and better!
\BScience\b
The resolution of the Voyager narrow-angle television cameras is sharp enough to read a newspaper headline at a distance of 1 km (0.62 mi).
Pele, the largest of the volcanoes seen on Jupiter's moon Io, is throwing sulfur and sulfur-dioxide products to heights 30 times that of Mount Everest, and the fallout zone covers an area the size of France. The eruption of Mount St. Helens was but a tiny hiccup in comparison (admittedly, Io's surface-level gravity is some six times weaker than that of Earth).
The smooth water-ice surface of Jupiter's moon Europa may hide an ocean beneath, but some scientists believe any past oceans have turned to slush or ice. In \I2010: Odyssey Two\i, Arthur C. Clarke wraps his story around the possibility of life developing within the oceans of Europa.
The rings of Saturn appeared to the Voyagers as a dazzling necklace of 10,000 strands. Trillions of ice particles and car-sized bergs race along each of the million-kilometer-long tracks, with the traffic flow orchestrated by the combined gravitational tugs of Saturn, a retinue of moons and moonlets, and even nearby ring particles. The rings of Saturn are so thin in proportion to their 171,000 km (106,000 mi) width that, if a full-scale model were to be built with the thickness of a phonograph record the model would have to measure four miles from its inner edge to its outer rim. An intricate tapestry of right-article patterns is created by many complex dynamic interactions that have spawned new theories of wave and particle motion.
Saturn's largest moon Titan was seen as a strange world with its dense atmosphere and variety of hydrocarbons that slowly fall upon seas of ethane and methane. To some scientists, Titan, with its principally nitrogen atmosphere, seemed like a small Earth whose evolution had long ago been halted by the arrival of its ice age, perhaps deep-freezing a few organic relics beneath its present surface.
The rings of Uranus are so dark that Voyager's challenge of taking their picture was comparable to the task of photographing a pile of charcoal briquettes at the foot of a Christmas tree, illuminated only by a 1 watt bulb at the top of the tree, using ASA-64 film. And Neptune light levels will be less than half those at Uranus.
\BThe Future\b
The solar system does not end at the orbit of Pluto, the ninth planet. Nor does it end at the heliopause boundary, where the solar wind can no longer continue to expand outward against the interstellar wind. It extends over a thousand times farther out where a swarm of small cometary nuclei, termed Oort's Cloud, is barely held in orbit by the Sun's gravity, feeble at such a great distance. Voyager 1 passed above the orbit of Pluto in May 1988, and Voyager 2 will pass beneath Pluto's orbit in august 1990. But even at speeds of over 35,000 mph, it will take nearly 20,000 years for the Voyagers to reach the middle of the comet swarm, and possibly twice this long for them to pass the outer boundaries of cometary space. By this time, they will have traveled a distance of two light-years, equivalent to half of the distance to Proxima Centauri, the nearest star.
Barring any serious spacecraft subsystem failures, the Voyagers may survive until the early twenty-first century, when diminishing power and hydrazine levels will prevent further operation. Were it not for these dwindling consumables and the possibility of losing lock on the faint Sun, our tracking antennas could continue to "talk" with the Voyagers for another century or two!
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"Voyager's Interstellar Mission",643,0,0,0
\BMission Objective\b
The mission objective of the Voyager Interstellar Mission (VIM) is to extend the NASA exploration of the solar system beyond the neighborhood of the outer planets to the outer limits of the Sun's sphere of influence, and possibly beyond. This extended mission is continuing to characterize the outer solar system environment and search for the heliopause boundary, the outer limits of the Sun's magnetic field and outward flow of the solar wind. Penetration of the heliopause boundary between the solar wind and the interstellar medium will allow measurements to be made of the interstellar fields, particles and waves unaffected by the solar wind.
\BMission Characteristics\b
The VIM is an extension of the Voyager primary mission that was completed in 1989 with the close flyby of Neptune by the Voyager 2 spacecraft. Neptune was the final outer planet visited by a Voyager spacecraft. Voyager 1 completed its planned close flybys of the Jupiter and Saturn planetary systems while Voyager 2, in addition to its own close flybys of Jupiter and Saturn, completed close flybys of the remaining two gas giants, Uranus and Neptune.
At the start of the VIM, the two Voyager spacecraft had been in flight for over 12 years having been launched in August (Voyager 2) and September (Voyager 1), 1977. Voyager 1 was at a distance of approximately 40 AU (Astronomical Unit - mean distance of Earth from the Sun, 150 million kilometers) from the Sun, and Voyager 2 was at a distance of approximately 31 AU.
Voyager 1 is escaping the solar system at a speed of about 3.5 AU per year, 35 degrees out of the ecliptic plan to the north, in the general direction of the Solar Apex (the direction of the Sun's motion relative to nearby stars). Voyager 2 is also escaping the solar system at a speed of about 3.1 AU per year, 48 degrees out of the ecliptic plane to the south. (see picture 1)
It is appropriate to consider the VIM as three distinct phases: the termination shock, heliosheath exploration, and interstellar exploration phases. The two Voyager spacecraft began the VIM operating, and are still operating, in an environment controlled by the Suns magnetic field with the plasma particles being dominated by those contained in the expanding supersonic solar wind. This is the characteristic environment of the termination shock phase. At some distance from the Sun, the supersonic solar wind will be held back from further expansion by the interstellar wind. The first feature to be encountered by a spacecraft as a result of this interstellar wind/solar wind interaction will be the termination shock where the solar wind slows from supersonic to subsonic speed and large changes in plasma flow direction and magnetic field orientation occur.
Passage through the termination shock ends the termination shock phase and begins the heliosheath exploration phase. While the exact location of the termination shock is not known, it is very possible that Voyager 1 will complete the termination shock phase of the mission about the year 2000 or 2001 when it will be about 80 AU from the Sun. After passage through the termination shock, the spacecraft will be operating in the heliosheath environment which is still dominated by the Suns magnetic field and particles contained in the solar wind. The heliosheath exploration phase ends with passage through the heliopause which is the outer extent of the Suns magnetic field and solar wind. The thickness of the heliosheath is uncertain and could be tens of AU thick taking several years to traverse. Passage through the heliopause begins the interstellar exploration phase with the spacecraft operating in an interstellar wind dominated environment. This interstellar exploration is the ultimate goal of the Voyager Interstellar Mission. (see picture 2)
\BScience Investigations\b
There are currently five science investigation teams participating in the VIM.
\B╖\b Magnetic field investigation
\B╖\b Low energy charged particle investigation
\B╖\b Plasma investigation
\B╖\b Cosmic ray investigation
\B╖\b Plasma wave investigation
The science teams for these investigations are currently collecting and evaluating data on the strength and orientation of the Sun's magnetic field; the composition, direction and energy spectra of the solar wind particles and interstellar cosmic rays; the strength of radio emissions that are thought to be originating at the heliopause, beyond which is interstellar space; and the distribution of hydrogen within the outer heliosphere.
There are seven operating instruments on-board each Voyager spacecraft although the Plasma instrument on Voyager 1 is not returning useful data. Five of these instruments directly support the five science investigation teams.
These five instruments are:
\B╖\b MAG Magnetic field investigation
\B╖\b LECP Low energy charged particle investigation
\B╖\b PLS Plasma investigation
\B╖\b CRS Cosmic ray investigation
\B╖\b PWS Plasma wave investigation
In addition, there are data being collected from two science instruments that do not have official science investigation teams associated with them. These instruments are:
\B╖\b PRA Planetary Radio Astronomy Subsystem
\B╖\b UVS Ultraviolet Spectrometer Subsystem
While there are not science investigation teams associated with these instruments, the captured data is made available to interested scientists.
\BScience Data Acquisition Strategy\b
Science data are returned to earth in real time at 160 bps. Real time data capture uses 34 meter Deep Space Network (DSN) resources with the project goal to acquire at least 16 hours per day of real time data per spacecraft. This goal is not always achieved due to the competition for DSN resources with prime mission projects and other extended mission projects.
Once a week per spacecraft, 48 seconds of high rate (115.2 kbps) PWS data are recorded onto the Digital Tape Recorder (DTR) for later playback. These data are played back to Earth once every 6 months per spacecraft and require 70 meter DSN support for data capture. After transmission of the data (either real time or recorded) to JPL, it is processed and made available in electronic files to the science teams located around the country for their processing and analysis.
\BSpacecraft Lifetime\b
The two Voyager spacecraft continue to operate, with some loss in subsystem redundancy, but still capable of returning science data from a full complement of VIM science instruments. Both spacecraft also have adequate electrical power and attitude control propellant to continue operating until around 2020 when the available electrical power will no longer support science instrument operation. At this time science data return and spacecraft operations will end.
Spacecraft electrical power is supplied by Radioisotope Thermoelectric Generators (RTGs) that provided approximately 470 w of 30 volt DC power at launch. Due to the natural radioactive decay of the Plutonium fuel source, the electrical energy provided by the RTGs is continually declining. At the beginning of 1997, the power generated by Voyager 1 had dropped to 334 w and to 336 w for Voyager 2.
Both of these power levels represent better performance than the pre-launch predictions, which included a conservative degradation model for the bi-metallic thermocouples used to convert thermal energy into electrical energy. As the electrical power becomes less and less, power loads on the spacecraft must be turned off in order to avoid having demand exceed supply. As loads are turned off spacecraft capabilities are eliminated. The table on the left identifies the year when specific capabilities will end as a result of the available electrical power limitations.
In order to maximize the duration of the fields and particles data acquisition capability, the first spacecraft loads to be turned off are instrument heaters on the scan platform. As these heaters are turned off the UVS, which is mounted on the scan platform, cools down until the point is reached when it can no longer function.
Termination of gyro operations ends the capability to calibrate the magnetometer instrument with magnetometer roll maneuvers (MAGROLs). These maneuvers are performed 6 times a year, on each spacecraft, and consist of a spacecraft attitude maneuver of 10 successive 360 degree turns about the roll axis. Data from a MAGROL allow the spacecraft magnetic field to be determined and subtracted from the magnetometer science data. This is important since the spacecraft magnetic field is larger than as the solar magnetic field being measured. The termination of gyro operations also means an end to the attitude maneuvers used to check the combined calibration of the Sun Sensor and the High Gain Antenna pointing direction for maintaining communications with the ground.
Instrument power sharing limits the number of science instruments that can be on at any given time. This instrument power sharing will continue until the available power will no longer support any instrument operation. At that time the Voyager Interstellar Mission will end.
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"Voyager To Maintain Data Acquisition for 30 Years (1)",644,0,0,0
\B1. INTRODUCTION\b
The Voyager Interstellar Mission (VIM) has the potential for obtaining useful interplanetary, and possibly interstellar, fields, particles, and waves (FPW) science data until around the year 2020 when the spacecraft's ability to generate adequate electrical power for continued science instrument operation will come to an end. In order to capitalize on this lengthy data acquisition potential, it is imperative that the spacecraft have a continuing sequence of instructions for acquiring the desired science data, and that the spacecraft High Gain Antenna (HGA) remain boresighted on the Earth for continuous data transmission.
Because of the long mission duration and the likelihood of periodic spacecraft anomalies, it is also advantageous to continue the use of the on-board fault protection capability for automated responses to specific subsystem anomalies, and to provide an onboard sequence to continue spacecraft operation in the specific event of the future loss of command reception capability. All of these factors are considered in the VIM sequencing strategy. The following sections describe the overall spacecraft sequencing strategy being used for VIM operations and the four sequencing elements (Baseline Sequence, Overlay Sequence, Mini-Sequence, and Backup Mission Load) used to implement the strategy.
The VIM sequencing strategy for spacecraft operations is significantly different from the strategy used during the primary mission which ended with the Voyager 2 flyby of Neptune in 1989. Only the current VIM strategy is discussed.
\B2. SEQUENCING STRATEGY\b
A spacecraft sequence is the set of instructions that are stored in an on-board computer memory and control the operation of the engineering subsystems and science instruments for the purpose of gathering and transmitting science data to the ground stations. Typically, a spacecraft sequence is designed to function for a specific period of time. Some of the key factors that determine the duration of a spacecraft sequence are:
\B╖\b size of the on-board memory;
\B╖\b characteristics of the science observations; a series of one of a kind type of observations, or a series of observations that are repetitive for an extended period of time;
\B╖\b ability to store HGA pointing information on-board the spacecraft.
With the Voyager spacecraft being mid-1970's technology, the available on-board memory for storing spacecraft sequence information is very limited by today's standards. A total of less than 1700 words are available between the two Computer Command Subsystem (CCS) memories for storing sequence instructions and HGA pointing information. This limited memory space can be a significant factor if the sequence is composed of a series of one of a kind activities.
Fortunately, the science observations being made during VIM are very repetitive in nature. This allows a science observation or instrument calibration to be repeated (cyclicized), essentially for the duration of the mission, at practically the same memory word cost as a single observation or calibration. This implementation of events as a cyclic sequence that repeats an observation or calibration on a regular basis is key to the VIM sequencing strategy.
Another key factor is the ability to store HGA pointing information out to approximately the year 2020 in the CCS memory. This not only simplifies the routine sequence generation support by not requiring HGA pointing information to be included in individual sequences transmitted to the spacecraft, but also enables the ability to provide a 30 year science data return mission even if command capability to the spacecraft is lost.
As long as command capability exists between the ground and the spacecraft, new spacecraft sequences can be stored on-board the spacecraft, or existing on-board sequences can be modified. If command capability is lost, the spacecraft must continue to operate with the sequence that is already stored on-board. Both of these command capability considerations are part of the VIM sequencing strategy.
The VIM sequencing strategy relies on having stored on-board each spacecraft a continuously executing sequence of repetitive science observations that satisfies the basic VIM heliospheric data acquisition requirements. HGA pointing information (HPOINTS), to keep the HGA pointed at the Earth until approximately the year 2020, are also stored in sequence memory to provide for continual communications capability. These two sequence capabilities are referred to as the "baseline sequence." Augmentation of the baseline sequence with non-repetitive science or engineering events use either an "overlay sequence," or a "mini-sequence."
The difference between these two types of augmentation sequences is that the overlay sequence operates for a fixed interval of time, currently six months, and contains all of the baseline sequence augmentations for that time interval. The mini-sequence is focused on accomplishing a single augmentation need and is not a regularly scheduled activity but is done on an as needed basis. Both types of sequences are developed and transmitted from the ground.
The baseline sequence, overlay sequence, and mini-sequence provide the basic operational sequence elements for normal operations while command capability with the spacecraft is available. In the event command capability is lost, another sequence element, the "Backup Mission Load, " (BML) provides the mechanism for continued spacecraft data acquisition without further ground interaction. A BML is stored on-board each spacecraft and contains the necessary instructions to modify the continuously executing baseline sequence to maintain the continued return of basic FPW data.
The BML also terminates both the use of the gyros when the available electrical power will no longer support their operation, and Digital Tape Recorder (DTR) utilization when telecom performance is no longer sufficient to support the minimum DTR playback data rate of 1.4 kbps. The termination of gyro and DTR operations are unconditional events and will occur at the programmed times even if command capabilty is available. If command capability is available, the times of these events can be updated based on observed actual performance.
All of these sequence elements use pre-defined blocks of commands to accomplish specific spacecraft functions. This is in contrast to the prime mission method of essentially programming each sequence in assembly language to maximize the science data return for each sequence. While the ability to use pre-defined blocks of commands greatly reduces the effort required to generate and validate a sequence of commands, there is an increase in the number of memory words needed to accomplish a given function. Fortunately, the VIM science data acquisition requirements, sequencing strategy, and available CCS memory space support the use of pre-defined blocks of commands.
The remainder of this write-up describes each of the four sequence elements. Also discussed is:
1. the interaction of the sequence elements with the on-board Fault Protection Algorithms (FPAs) that continually monitor selected subsystem performance parameters and contain programmed responses which are initiated in the event of out of-tolerance performance;
2. the planned modifications to the spacecraft configuration in response to the continual reduction in electrical power availability throughout the remainder of the mission.
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"Voyager To Maintain Data Acquisition for 30 Years (2)",645,0,0,0
\B3. BASELINE SEQUENCE DESCRIPTION\b
The baseline sequence is stored in CCS memory and is the continuously executing set of spacecraft operating instructions that controls the collection and return of the bulk of VIM science data acquired during normal operations (when command capability exists). It also provides the foundation for continuing the return of science data in the event of loss of command capability and the resulting BML entry (described in the following BACKUP MISSION LOAD DESCRIPTION).
Baseline sequence implementation relies on the use of eleven "spacecraft block routines" that are also stored in CCS memory. These routines are the equivalent of software program subroutines and are called by the baseline sequence for execution.
During normal operations, each spacecraft performs the following repetitive science and engineering activities under the control of the baseline sequence:
\B╖\b collection and transmission of continuous real time cruise science telemetry data, primarily at 160 bps data rate;
\B╖\b recording of one frame of high rate Plasma Wave Subsystem (PWS) data on the DTR each week
\B╖\b playback of six months of recorded high rate PWS data every 6 months;
\B╖\b execution of a magnetometer calibration roll maneuver (MAGROL) every 3 months;
\B╖\b execution of a HGA/sun sensor calibration maneuver (ASCAL) every 6 months;
\B╖\b perform a PMPCAL once a month. The PMPCAL consists of a combined Plasma Subsystem (PLS), Magnetometer Subsystem (MAG), and FPW Periodic Engineering and Science Calibration (PESCAL), calibration ;
\B╖\b perform DTR maintenance twice a year;
\B╖\b perform gyro conditioning (on/off) and a CCS timing test every 3 months
In addition, the baseline sequence contains the HGA pointing information to keep the antenna pointed at Earth for communication purposes. Commands are stored on-board to maintain HGA pointing on Voyager 1 until Day-Of-Year (DOY) 342, 2020, and until DOY 204, 2017 on Voyager 2. HGA pointing accuracy is assessed every 6 months with the ASCAL maneuver. If necessary, the stored pointing information can be updated, via ground command, with a new pointing file. No updates have been required since the HGA pointing information was initially stored in both spacecraft in 1990 and none are expected in the future based on performance trends to date.
All of the Voyager 1 baseline sequence spacecraft events which require 70m station coverage for capturing the telemetry data on the ground (MAGROL, playback of recorded high rate PWS data, & ASCAL) are scheduled to occur over the Goldstone station complex and the Voyager 2 events over the Canberra complex. These combinations provide the best telecom performance for data return. In order to keep baseline sequence spacecraft events in sync with Deep Space Station (DSS) view periods (requires a near constant sidereal time), spacecraft events are shifted in time at the rate of approximately 4 minutes earlier per day. This adjusts for the difference between a solar day (24 hours) and a sidereal day (23 hours, 56 min, 04.09 sec).
As a result of the accumulated effect of this 4 minute per day shift, spacecraft events occur one day earlier for each year of operation. After six consecutive years, the spacecraft events have shifted 6 days earlier in time. At the start of the seventh year, the timing of the events are delayed 7 days and the 4 minute per day shifting begins again.
This periodic 7 day timing shift keeps the spacecraft events occurring within a one week calendar time window for the duration of VIM, thus avoiding critical spacecraft (DTR playbacks or attitude maneuvers) from occurring during designated quiet weeks (i. e. Thanksgiving, Christmas/New Years). It should be noted that the described timing shifts are designed to maintain the proper Earth Received Time (ERT) of the spacecraft telemetry data within the desired station view period. This is important since the one-way light time increases approximately 1/2 hour per year of flight.
Once a year, in November for Voyager 1 and September for Voyager 2, the baseline sequence cyclics end and restart with the previously described timing adjustments being made. At this time, any ground commanded modifications to the baseline sequence also take effect.
Because of the continual reduction in the electrical power available to operate the spacecraft, gyro operations have to be terminated at some future time. DTR operations also have to be terminated when the downlink telecom performance will no longer support the minimum DTR playback data rate of 1.4 kbps. Based on expected power and telecom subsystem performance, the dates for the termination of gyro and DTR operations are stored in the long term events table of the BML for each spacecraft. These timed unconditional execute commands are used to terminate the baseline sequence gyro activities (MAGROLs, ASCALs, and gyro conditioning), and to terminate the recording and playback of high rate PWS data. The current dates stored in the long term events table are:
"Voyager To Maintain Data Acquisition for 30 Years (3)",646,0,0,0
\B4. OVERLAY SEQUENCE DESCRIPTION\b
Overlay sequences are used to augment the continuously executing baseline sequence and are prepared on a regularly scheduled basis. At the start of VIM (1990), overlay sequences were of 3 month duration, with a 6 week development period. Considering both spacecraft, with overlay sequence execution staggered 1 1/2 month between spacecraft, there was essentially an overlay sequence being developed at all times. Initial VIM flight team staffing levels supported two Sequence Integration Engineers (SIEs).
This allowed one SIE to be developing an overlay sequence while the second SIE could be performing other tasks (real time command requests, documentation updates, software maintenance, etc.). With the 1993 reduction in flight team staffing (reduction to one full time SIE), the duration of the overlay loads was extended to 6 months. The development time remained at 6 weeks providing time for the single SIE to perform the other necessary SIE tasks between overlay sequence development periods.
The intent of the overlay sequence is to provide a mechanism for incorporating non-repetitive science or engineering events into the spacecraft sequence of activities in combination with the baseline sequence. Typical events in overlay sequences have included:
\B╖\b Ultraviolet Spectrometer Subsystem (UVS) stellar and heliospheric observations;
\B╖\b two additional MAGROLs per year per spacecraft;
\B╖\b additional high rate PWS records and DTR playbacks;
\B╖\b DTR playbacks to recover data when the baseline sequence playback was not captured on the ground;
\B╖\b CCS/Flight Data Subsystem (FDS)/Attitude and Articulation Control Subsystem (AACS) memory readouts;
\B╖\b updates to BML and FPAs;
\B╖\b modifications to the baseline sequence.
The need for the 3 month and then 6 month overlay sequences has been driven primarily by the requirement for acquiring UVS stellar and heliospheric observations. The termination of UVS observations will eliminate the necessity for these regularly scheduled sequence loads. However, there remains a need to continue to provide a capability for augmentation of the baseline sequence for either science data acquisition or spacecraft engineering needs. This augmentation will be accomplished by small individual mini-sequences that perform a specific function and are implemented on an as needed basis.
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"Voyager To Maintain Data Acquisition for 30 Years (4)",647,0,0,0
\B5. MINI-SEQUENCE DESCRIPTION\b
Mini-sequences also augment the baseline sequence but are prepared on an as needed basis rather than on a regularly scheduled basis. These sequences are usually intended to perform a single function rather than the multiple functions performed by an overlay sequence. The development time for a mini-sequence varies from a day to a week or so depending on the complexity of the sequence.
A common use of a mini-sequence is to sequence a second DTR playback of high rate PWS data if the baseline sequence playback is not captured on the ground. Another common use is in response to a spacecraft anomaly (i. e. the performance of a memory readout, a CCS timing test, or other diagnostic or anomaly resolution effort). When the termination shock is expected to be encountered, a mini-sequence will be used to enable an on-board stored routine that will increase the number of high rate PWS records from once a week to approximately every nine hours. A mini-sequence will also be used to play back the recorded data.
\B6. BACKUP MISSION LOAD DESCRIPTION\b
The BML provides automated on-board protection against the loss of command capability to the spacecraft. Without command capability, the spacecraft must continue to operate with the instructions previously stored in the CCS memory. Key to the automated protection against the loss of command capability is the implementation of a command loss timer on-board each spacecraft. The command loss timer is simply a programmable timer that resets to a programmed duration each time a command is received by the spacecraft.
The current duration of the command loss timer is six weeks. When a command is received by the spacecraft, the timer is reset to a six weeks duration and immediately begins counting down towards zero. If the timer reaches zero, as a result of a command not being received by the spacecraft within the programmed six week duration, the command loss timer will have expired and the Command Loss (CMDLOS) routine will be activated which leads to the initiation of the BML.
The implementation of BML-7 (the seventh BML to be loaded on-board Voyager 2), in conjunction with the baseline sequence, provides this automated protection against loss of command capability. BML-7, with some differences in implementation for the two spacecraft, is loaded on-board both Voyager 1 and 2.
BML-7 is designed to meet the following science and mission objectives:
1. Continue to provide baseline sequence FPW data after a loss of command cability, for as long as the spacecraft continues to operate. In order to achieve this, BML-7 modifies the baseline sequence events to limit spacecraft activities (i. e. terminate MAGROLs and ASCALs) and configures the spacecraft FPAs for compatibility with the loss of command reception capability.
2. Continue to provide baseline sequence periodic calibrations of the science instruments (PMPCALs) in order to allow science teams to properly interpret the data.
3. Continue to perform baseline sequence DTR maintenance in support of the PWS high rate data acquisition, until DTR operations are terminated in 2010 for Voyager 1 and 2012 for Voyager 2.
4. Provide periodic gyro conditioning to maintain gyro integrity for FPA responses. Gyro conditioning is continued until the sequenced termination of gyro operations in 2007 for Voyager 1 and 2004 for Voyager 2.
5. Include reasonable provisions for maintaining an acceptable telecom posture for receipt of telemetry data via the X-band link.
6. Include additional reasonable fault protection enhancements.
7. Be transparent to the normal VIM sequence activities.
In response to these objectives, the design of BML-7 is organized into three independent tables, the BML Long-Term Events Table, the BML Initialization Table 1 (BMLIT1) and the BML Initialization Table 2 (BMLIT2). BMLIT1 and BMLIT2 are initiated after entry into the CMDLOS routine and contain commands that are only needed in the event command reception capability is lost. The BML Long-Term Events Table is a continuously active time event region in the CCS that executes independently of BML (the events contained in this table will occur whether BML-7 is entered or not).
BML is initiated with the execution of BMLIT1. The BMLIT1 events will be executed two weeks after CMDLOS entry if the CMDLOS timer has not been reset, but not prior to the end of a current overlay sequence plus 24 hours. The former condition indicates that BMLIT1 will not execute if command ability is restored. The latter condition is necessary to maintain the transparency of BML-7 from the normal sequencing activities. It is critical for BML to be independent of an active overlay sequence to avoid the necessity of constraining the contents of the overlay sequence to be compatible with BML-7.
A flag (CCS location 6714B) used in normal sequence design to allow the multiple uplink of a sequence load is also used in BML-7 to check for the end of an overlay sequence. This flag is set to 1 by the successful reception of all command blocks of an overlay sequence transmission. It remains set to 1 during execution of the overlay sequence and is automatically reset to 0 when the next planned uplink window opens at the end of an overlay sequence. This flag constrains the BMLIT1 events to not occur until 24 hours after the flag has been reset to 0 by the completion of the overlay sequence, even if the time from CMDLOS entry has exceeded 2 weeks.
Once initiated, BMLIT1 will:
1. Disable future attitude control maneuvers (ASCALs and MAGROLs). ASCALs will not be needed if the HPOINTS (sun sensor bias commands to update HGA pointing) cannot be reloaded, and MAGROLs may require update of the gyro scale factor, neither of which can be done if spacecraft command capability does not exist.
2. Initiate gyro conditioning. One pair of gyros will be powered on for a one week duration with the gyro fault test enabled once a year.
3. Select the 0.05 degree yaw deadband and the 0.25 degree roll deadband. All the future attitude control deadband will be 0.05/0.05/0.25 degrees, pitch/yaw/roll to maximize telecom performance.
4. Select the 160 bps downlink data rate for cruise data return.
5. Select the Attitude Propulsion (AP) thruster pulse width to the nominal 10 ms duration.
Prior to the start of the A022 sequence (October 16,1995) for Voyager 1 and the start of the B025 sequence (January 15,1996) for Voyager 2, the command loss timer for each spacecraft was set to four weeks. Beginning with sequences A022 and B025, the command loss timer for each spacecraft is set to six weeks. If command reception is not achieved during the duration of the command loss timer following the last successful command reception, entry into the CMDLOS routine automatically occurs. CMDLOS entry initiates a programmed response to try and establish a command reception capability using different combinations of redundant command reception hardware.
The reception of a ground command by the spacecraft during CMDLOS execution terminates CMDLOS and resets the CMDLOS timer while maintaining the successful command reception hardware configuration. After CMDLOS entry, the CMDLOS routine is executed four consecutive times. After the fourth execution of the CMDLOS routine, the routine is permanently disabled and BML will take effect. The BMLIT2 events will occur approximately three weeks after the initiation of the fourth, and final, CMDLOS execution (17 weeks, 4 days and 3 hours from the first CMDLOS entry if the command loss timer is set for four weeks and 23 weeks, 4 days and 3 hours if the command loss timer is set for six weeks).
Once initiated, BMLIT2 issues commands to power down the scan platform and configure the spacecraft for the long-term BML mission. Specifically, BMLIT2 will:
1. Power down the scan platform. The following electrical loads will be powered off if not already off:
\B╖\b ISS Narrow Angle (NA) Vidicon Replacement Heater
\B╖\b ISS NA Electronics Replacement Heater
\B╖\b Scan Platform Azimuth Actuator Heater
\B╖\b Azimuth Actuator Coil Heater
\B╖\b Infrared Interferometer Spectrometer and Radiometer Subsystem (IRIS) Replacement Heater
\B╖\b Ultraviolet Spectrometer Subsystem (UVS)
2. Initialize the Low Energy Charged Particle Subsystem (LECP) for data gathering.
3. Initialize the CRS for data gathering.
4. Select data rate of 1.4 kbps for DTR playback of high rate PWS data.
5. Inhibit switch of X-band power level from high to low power after DTR playback of high rate PWS data.
6. Configure the RF subsystem for the BML mission by establishing X-high power for all BML downlink operations if the prime X-band traveling wave tube (TWT) is operating on Voyager 1 and either X-band TWT is operating on Voyager 2. Because of the previous TWT switch on Voyager 1, X-Hi is maintained only if the current prime TWT is operating. If the current prime TWT fails and a switch to the backup TWT is made, X-Lo will be used for telemetry transmission.
7. Modify PWRCHK fault protection algorithm to select X-band TWT to the BML configuration after a PWRCHK entry.
8. Link the AACS Trajectory Correction Maneuver (TCM) thrusters as a second backup to provide pitch/yaw attitude control if the primary and backup thruster branches have failed.
The BML Long-Term Events Table is a separate continuously executing time event region that contains all of the unconditional long-term events for baseline sequence and BML operation. These events have absolute time values associated with them and will occur independent of whether the BML has been initiated. Key spacecraft reconfiguration events contained in this table include the termination of gyro and DTR operations.
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"Voyager To Maintain Data Acquisition for 30 Years (5)",648,0,0,0
\B7. SEQUENCE INTERACTION WITH ON BOARD FAULT PROTECTION\b
The FPAs on the Voyager spacecraft are designed to recover the spacecraft from specific potential mission-catastrophic failures. They are implemented primarily in the CCS, while some are in the AACS. The FPAs in the CCS are invoked by interrupts received from external sources, and followed by preprogrammed responses.
The five FPAs that currently reside in the CCS are:
\B╖\b AACS Power Code Processing (AACSIN)
\B╖\b Command Loss (CMDLOS)
\B╖\b Radio Frequency Power Loss (RFLOSS)
\B╖\b CCS Error (ERROR)
\B╖\b Power Check (PWRCHK)
The main function of the AACSIN routine is responding to AACS anomalies by the processing of Power Codes (PCs) received from the AACS. A PC is a regularly transmitted message from the AACS to the CCS providing status information about the AACS. Some of those PCs relating to a faulty AACS condition include heartbeat failure, celestial reference loss/acquisition, power supply failure, gyro failure, scan slew abort, command parity error, and bad/no echo response. The PCs are either functional or informational: the functional PC is used to generate a command to the power subsystem to initiate the proper response, and the informational PC is used by the CCS to invoke the preprogrammed responses. In most instances, the CCS echoes the PC back to the AACS to confirm receipt.
The purpose of the CMDLOS routine is to provide a means for the spacecraft to automatically respond to an on-board failure resulting in the inability to receive ground commands. Whenever a specified number of hours have elapsed without the CCS receiving a valid command, the CCS assumes a spacecraft failure and attempts to correct that failure by systematically switching to redundant hardware elements until a valid command is received. CMDLOS will be executed four consecutive times if command reception is not successful. After four unsuccessful executions, CMDLOS will be permanently disabled and BML will be activated.
The RFLOSS routine provides the spacecraft a means of automatically recovering from a failure of an X-band exciter or transmitter. The CCS monitors four input signals from the RFS that are associated with a failure of the exciters and transmitters. Whenever one or more of the above signals is input to the CCS, RFLOSS will systematically interrogate the four interrupts and attempt to correct the failure by swapping to a redundant unit.
If an AACSIN, CMDLOS, or RFLOSS entry occurs, the response will be integrated into any on-going sequencing activities which typically includes the baseline load, overlay load, HPOINTS and any unconditional long term events that are contained in the BML Long-Term Events Table. The commands from an FPA response and the regular sequencing activities will be interleaved.
The ERROR routine provides the capability to respond to certain anomalous CCS hardware and software conditions. In the event of a detected anomaly, the routine puts the CCS in a known, quiescent state and waits for ground action. It also stops any on-going sequencing activities including baseline load, overlay load, HPOINTS, and BML Long-Term Events Table. However, if an ERROR entry occurs in the BML mission, the sequencing activities will be restarted by the programmed rollback feature as described below.
The purpose of the PWRCHK is to provide a means for the spacecraft to automatically configure itself to a safe, low-power operating mode following a power subsystem undervoltage condition or a CCS tolerance detector trip. If a PWRCHK entry is due to an undervoltage condition, the routine issues commands reducing the spacecraft power load in an attempt to recover from the failure condition and configure the spacecraft to a safe, low-power mode.
If the CCS tolerance detector trip occurs indicating that the CCS input power has dropped below a level where the processor can reliably function, the spacecraft assumes that all other loads have experienced power-on-resets and issues the commands to re-enable the essential functions. In addition, due to the long round trip light time (RTLT), reduced Deep Space Network (DSN) coverage and BML vulnerability, a set of commands was designed and implemented on the spacecraft for VIM to configure the science instruments and DTR back to the nominal mission.
Since the PWRCHK is entered via ERROR routine, all the sequencing activities including heartbeat, baseline load, overlay load, HPOINTS and BML Long-Term Event Table will be terminated as in the ERROR entry. However, the baseline load and HPOINTS will be restarted by rollback, and the heartbeat and the BML Long-Term Events Table by PWRCHK.
The rollback is a special feature designed to allow one time event region per processor to be automatically restarted in the event of a PWRCHK entry or an ERROR entry in the BML mission. The rollback keeps track of timing of that particular event so it can be restarted as if there were no interruption. The CCS processor A rollback is used to restart the HPOINT table and the CCS processor B is for the baseline load.
In addition, one location in PWRCHK is reserved for restarting the BML. Unlike the baseline load or HPOINTs, timing for BML Long-Term Events is not kept track of, and that location has to be periodically updated in the BML Long-Term Events Table to point to the next long-term event. As a result, some of BML long-term events may be repeated. However, this will not pose any harm to the spacecraft.
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"Voyager To Maintain Data Acquisition for 30 Years (6)",649,0,0,0
\B8. MODIFICATIONS IN RESPONSE TO ELECTRICAL POWER LIMITATIONS\b
Radioisotope Thermoelectric Generators (RTGs) provide electrical power to the Voyager spacecraft. Due to the radioactive decay of the Plutonium fuel source, the electrical power provided by the RTGs is continually declining. The current rate of decay is approximately 5 watts per year. Because of the continual decline in the amount of power that is available, it is necessary to periodically reduce power consumption in order to maintain an adequate power margin. This is accomplished by turning off spacecraft power loads. Modifications planned are listed in the following tables by specific events and times for each spacecraft.
\BVOYAGER 1 \b
1998 - Reduction in Scan Platform power - preserve UVS and Elevation Actuator temperature (+11.0 W)
2010 - Termination of DTR operations (+5.8 W for DTR turnaround)
This power load reduction step is currently sequenced to occur on DOY 007, 2004 and could be changed if the current performance trends continue.
2011 - Discontinue gyro operations (+14.4 W steady state, +3.6 W turn on transient). This power load reduction step is currently sequenced to occur on DOY 305, 2010.
2015 - Pyro Instrumentation Power OFF (+2.4 W)
2016 - Turn off PLS instrument
Further responses to decreasing electrical power, beginning in 2018, will consist of either turning instrument off sequentially or turning instruments off and on in a power sharing mode to maintain an adequate power margin.
\BVOYAGER 2 \b
1996 - Reduction in Scan Platform power - preserve UVS and Elevation Actuator temperature (+11.0 W)
This power reduction step is currently sequenced to occur on DOY 267, 2006.
2010 - Discontinue gyro operations (+14.4 W steady state, +3.6 W turn on transient)
This power load reduction step is currently sequenced to occur on DOY 007, 2004 and could be changed if the current performance trends continue.
2012 - Termination of DTR operations (+5.8 W for DTR turnaround)
This power load reduction step is currently sequenced to occur on DOY 251, 2012.
2016 - Pyro Instrumentation Power OFF (+2.4 W)
Further responses to decreasing electrical power, beginning in 2016, will consist of either turning instrument off sequentially or turning instruments off and on in a power sharing mode to maintain an adequate power margin.
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"Space News Update 1998",650,0,0,0
\JTraining Begins For Crew Of Next Hubble Space Telescope Servicing Mission (30 July, 1998)\j
\JSOHO Spacecraft Located With Ground-Based Radar (27 July, 1998)\j
\JNearby Star Cluster Yields Insights Into Early Universe (23 July, 1998)\j
\JStatement of NASA Administrator Daniel S. Goldin on the Death of Alan Shepard (22 July, 1998)\j
\JNewly Discovered Stellar Cannibal Provides Missing Link (22 July, 1998)\j
\JAlan Shepard, First American Astronaut, Dies At 74 (22 July, 1998)\j
\JNASA Satellite Sheds New Light on the La Nina Phenomenon (16 July, 1998)\j
\JEfforts To Recover SOHO Spacecraft Continue (16 July, 1998)\j
\JNew NASA Facility Will Complete Worldwide Communications Coverage (13 July, 1998)\j
\JNASA Instruments On Japanese Planet-B Spacecraft Will Aid Studies Of Martian Upper Atmosphere (1 July, 1998)\j
\JEfforts To Recover SOHO Spacecraft Continue (30 June, 1998)\j
\JWater History, Rock Composition Among Latest Findings a Year after Mars Pathfinder (29 June, 1998)\j
\JNew Satellite Data Show Retreat of El Nino, Pacific Ocean in Transition (26 June, 1998)\j
\JNASA Starts Work on New Space Infrared Telescope Facility (25 March, 1998)\j
\JNASA Set to Study the Sun's Turbulent Upper Atmosphere (19 March, 1998)\j
\JNew Global Surveyor Data Reveals More (13 March, 1998)\j
\JMore Evidence Points To Impact As Dinosaur Killer (12 March, 1998)\j
\JAstronomers Track Down Asteroids in Hubble Archive (9 March, 1998)\j
\JLunar Prospector Finds Evidence of Ice at Moon's Poles (5 March, 1998)\j
\JCollins Named First Female Shuttle Commander (5 March, 1998)\j
\JBackground on Europa Data From The Galileo Mission To Jupiter (2 March, 1998)\j
\JDetailed Images From Europa Point To Slush Below Surface (2 March, 1998)\j
\JNASA Radar Reveals Hidden Remains At Ancient Angkor (12 February, 1998)\j
\JScientists View First Close-Ups of Strange, Layered Polar Terrain (11 February, 1998)\j
\JShock Wave Sheds New Light on Fading Supernova (10 February, 1998)\j
\JSpace Station Agreements To Be Signed In Washington (29 January, 1998)\j
\JSpace Pioneers Recall First U.S. Satellite Launch Upon 40th (28 January, 1998)\j
\JFast-Spinning Pulsar Discovery Provides Evolutionary Link (26 January, 1998)\j
\JAdditional Experiments Selected For Mars 2001 Missions (22 January, 1998)\j
\JEarth Swingby Puts NEAR Spacecraft on Final Approach to Eros (20 January, 1998)\j
\JFirst Station Element to be Shipped to Russian Launch Site (16 January, 1998)\j
\JAstronomers Discover an Infrared Background Glow in the Universe (9 January, 1998)\j
\JOld Faithful Black Holes Ejects Mass Equal to an Asteroid (7 January, 1998)\j
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"Training Begins For Crew Of Next Hubble Space Telescope Servicing Mission (30 July, 1998)",651,0,0,0
A team of veteran astronauts will begin training to install new instruments and upgrade systems to enhance the scientific capabilities of the orbiting Hubble Space Telescope. Crew members Steven L. Smith; C. Michael Foale, Ph.D.; European Space Agency astronaut Claude Nicollier; and John M. Grunsfeld, Ph.D., will conduct a record six space walks during the STS-104 mission, scheduled for launch in May 2000. Smith will be the payload commander, coordinating the astronauts' space-walking activities.
"The ambitious nature of this mission, with its six space walks, made it important for the payload crew to begin its training as early as possible," said David C. Leestma, director of Flight Crew Operations at NASA's Johnson Space Center, Houston, TX. The crew will rendezvous with and capture the orbiting Hubble Space Telescope, and secure it in Columbia's payload bay using the Shuttle's robot arm. Then, working in teams of two, the veteran astronauts will venture into the payload bay performing a variety of tasks that will improve the productivity and reliability of the telescope.
To enhance Hubble's scientific capability, the astronauts will remove the Faint Object Camera and replace it with the Advanced Camera for Surveys. With its three electronic cameras and complement of filters, this camera is expected to improve the telescope's sensitivity in the ultraviolet range by a factor of ten. Other primary tasks to be accomplished during the flight include the replacement of Fine Guidance Sensor 2, one of three such devices that help to accurately point the telescope; replacement of the existing solar arrays with rigid, high efficiency arrays; and replacement of a tape recorder with a solid state recorder.
Secondary tasks include the installation of an aft-shroud cooling system to upgrade the thermal protection of some of the telescope's systems; the installation of a new technology cryogenic cooler for the Near Infrared Camera and Multi-Object Spectrometer instrument, known as the NICMOS Cooling System; and the installation of six voltage/temperature improvement kits, which will improve Hubble's battery charge capability.
In addition, the astronauts will repair and replace much of the multi-layer exterior thermal insulation on the sun-facing side of the telescope. In February 1997, the STS-82 crew noticed peeling on several areas of the insulation and applied four patches to the most affected areas.Both Smith and Nicollier have previous experience with Hubble. Smith performed three space walks during the second Hubble servicing mission in February 1997. Nicollier operated the Shuttle's robot arm during the first visit to the telescope during the STS-61 mission in 1993. Foale has conducted space walks from both the Space Shuttle and Russia's Mir space station, accumulating more than 10 hours of space-walking experience. Grunsfeld has two previous space flights to his credit.
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"SOHO Spacecraft Located With Ground-Based Radar (27 July, 1998)",652,0,0,0
Ground-based radio telescopes have been able to detect the Solar and Heliospheric Observatory (SOHO) spacecraft and have found it rotating slowly near its original position in space, a potentially important step toward possible recovery of direct communications with the spacecraft. Radio contact with SOHO, a joint mission of the European Space Agency (ESA) and NASA, was interrupted on June 24, an event under review by a joint ESA/NASA investigation board.
With the encouragement of Dr. Alan Kiplinger of the National Oceanic and Atmospheric Administration's Space Environment Center in Boulder, CO, researchers at the U.S. National Astronomy and Ionosphere Center (NAIC) in Arecibo, Puerto Rico, used the facility's 305-meter (990-foot) diameter radio telescope to transmit a signal toward SOHO on July 23. The 70-meter dish of NASA's Deep Space Network in Goldstone, CA, acted as a receiver, locating the spacecraft's echo and tracking it using radar techniques for more than an hour.
Preliminary analysis of the radar data, which is ongoing, indicates that SOHO is still in its nominal halo orbit near the so-called "L-1" Lagrangian point in space, (a gravitationally stable vantage point 1.5 million kilometers ahead of the Earth) and is turning slowly at a rate of roughly one revolution per minute. Staff members of NAIC and the Deep Space Network, in close cooperation with ESA and NASA, are continuing to analyze the radar data to extract more precise information on SOHO's location and motion, which in turn could help in future recovery efforts, as SOHO's solar panels turn toward the Sun. ESA and NASA engineers also are continuing their efforts to re-establish radio data communication with the spacecraft, encouraged by the radar measurement of a slow spin rate, which suggests minimal structural damage has occurred.
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"Nearby Star Cluster Yields Insights Into Early Universe (23 July, 1998)",653,0,0,0
NASA's Hubble Space Telescope has taken a "family portrait" of young, ultra-bright stars nested in their embryonic cloud of glowing gases. The celestial maternity ward, called N81, is located 200,000 light-years away in the Small Magellanic Cloud, a small irregular satellite galaxy of our Milky Way. These are probably the youngest massive stars ever seen in the nearby galaxy. The nebula offers a unique opportunity for a close-up glimpse of the "firestorm" accompanying the birth of extremely massive stars, each blazing with the brilliance of 300,000 of our suns.
Such galactic fireworks were much more common billions of years ago in the early universe, when most star formation took place. "This is giving us new insights into the physical mechanisms governing star formation in far away galaxies that existed long ago," says Mohammad Heydari-Malayeri, who headed the international team of astronomers who made the discovery using Hubble's Wide Field and Planetary Camera 2. Because these stars are deficient in heavier elements, they also evolve much like the universe's earliest stars, which were made almost exclusively of the primordial elements hydrogen and helium that were created in the big bang.
The Small Magellanic Cloud is a unique laboratory for studying star formation in the early universe since it is the closest and best seen galaxy containing so-called "metal-poor" first- and second -generation type stars. These observations show that massive stars may form in groups. "As a result, it is more likely some of these stars are members of double and multiple star systems," says Heydari-Malayeri. "The multiple systems will affect stellar evolution considerably by ejecting a great deal of matter into space." This furious rate of mass loss from these stars is evident in the Hubble picture, which reveals dramatic shapes sculpted in the nebula's wall of glowing gases by violent stellar winds and shock waves. "This implies a very turbulent environment typical of young star formation regions," Heydari-Malayeri adds.
He believes one of the members of the cluster may be an extremely rare and short-lived class of super-hot star (50,000 degrees Kelvin) called a Wolf-Rayet. This star represents a violent, transitional phase in the final years of a massive star's existence - before it ultimately explodes as a supernova. "If confirmed by future Hubble observations, this finding will have a far-reaching impact on stellar evolutionary models," says Heydari-Malayeri. "That's because the Wolf-Rayet candidate is fainter than other such stars in that galaxy, in contrast with the predictions of these models." The team's work will be shortly submitted for publication in the European journal Astronomy and Astrophysics.
Hubble's resolution allows astronomers to pinpoint 50 separate stars tightly packed in the nebula's core within a 10 light-year diameter -- slightly more than twice the distance between Earth and the nearest star to our sun. The closest pair of stars is only one-third of a light-year apart. Before the Hubble observations, N81 was simply dubbed, "The Blob" because its features were indistinguishable by other telescopes.
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"Statement of NASA Administrator Daniel S. Goldin on the Death of Alan Shepard (22 July, 1998)",654,0,0,0
The entire NASA family is deeply saddened by the passing of Alan Shepard. NASA has lost one of its greatest pioneers; America has lost a shining star. Alan Shepard will be remembered, always, for his accomplishments of the past: being one of the original seven Mercury astronauts, for being the first American to fly in space, and for being one of only 12 Americans ever to step on the Moon.
He should also be remembered as someone who, even in his final days, never lost sight of the future. On behalf of the space program Alan Shepard helped launch, and all those that space program has and will inspire, we send our deepest condolences to his wife Louise, their children, and the rest of the Shepard family. Alan Shepard lived to explore the heavens. On this his final journey, we wish him Godspeed.
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"Newly Discovered Stellar Cannibal Provides Missing Link (22 July, 1998)",655,0,0,0
A newly-discovered star that is emitting rapid pulses of X-rays may be the long-sought missing link between old neutron stars that emit powerful flashes of X-rays, and older, rapidly spinning neutron stars that emit mainly radio waves. The new star, called an X-ray pulsar, is designated SAX J1808.4-3658. It has greatly accelerated its own rotation at the expense of a nearby "companion" star by pulling gas from the companion onto its surface in a process called accretion.
The fastest-spinning pulsar of its type ever seen, the newly discovered star is now rotating at more than 400 times per second (corresponding to a spin period of 2.5 milliseconds), making it the first known accretion-powered millisecond pulsar. Millisecond pulsars are neutron stars (extremely dense, city-sized stars) that rotate very rapidly; most complete one rotation in less than eight milliseconds (8/1000 of a second). Accretion occurs when gas from a nearby star gets pulled into the pulsar's strong gravitational field. Two competing teams used NASA's Rossi X-ray Timing Explorer (RXTE) spacecraft to make the discovery.
The first team, led by Dr. Michiel van der Klis and Rudy Wijnands of the University of Amsterdam, the Netherlands, discovered the pulsar and measured the time between rapid pulses of X-rays from the star to derive its rotation rate. The second team, led by Dr. Deepto Chakrabarty and Dr. Edward Morgan of the Massachusetts Institute of Technology, Cambridge, MA, discovered the two-hour orbital period of the pulsar and measured the size of the orbit, inferring the presence of a companion star. The results are being presented in the July 23 edition of the journal Nature.
"Astrophysicists have theorized for a long time that the only reason millisecond pulsars exist at all is that they get spun up by taking material from a companion star, but this is the first time one has been caught in the act. This has sometimes been called the Holy Grail of X-ray astronomy, and Rudy has at last found it!" said van der Klis. "This 'stellar cannibal' is a leisurely diner," added Chakrabarty. "We estimate that it has been pulling material from its companion star for the last 100 million to one billion years. Over that time, the companion star may have lost up to half its mass. Currently, the companion is about 15 percent of the mass of the Sun."
However, not all the companion's mass loss is due to accretion. "Millisecond pulsars may throw away material they can't capture by 'vaporizing' their companion stars with X-rays and particle beams. As accreting gas falls on to the surface of the pulsar, it heats up and emits X-rays. The X-rays blow material from the companion star. After the accretion phase ends, the pulsar may emit a high velocity beam of subatomic particles that continues to blow material off the companion. Over a billion years, this bombardment may cause the companion to vanish altogether," said Dr. Tod Strohmayer, a member of the RXTE team located at NASA's Goddard Space Flight Center, Greenbelt, MD.
"This X-ray and particle beam ablation may explain why millisecond pulsars are often found alone, despite the fact that they required a companion star to speed up. By 'vaporizing' the companion, they hide the evidence - it's a stellar version of the perfect crime," added Strohmayer. "In the case of the newly discovered pulsar, we found that its X-ray intensity is slightly fainter when it is on the far side of its orbit (with its companion between us and the pulsar). This is probably caused by an intervening 'fog' of material blown off the companion's surface -- direct evidence for 'vaporization' by the pulsar," said Chakrabarty.
The new pulsar helps scientists resolve a mystery. Prior to the discovery, two populations of neutron stars with relatively weak magnetic fields but with otherwise different characteristics were known. There were old, accreting neutron stars, which are powerful sources of X-rays generated from the material they are gobbling up from their companions, and the group of radiowave emitting pulsars that are rotating very rapidly and slowing down gradually. Scientists suspected there was a connection between the two, and the discovery of this pulsar that is both emitting X-rays and spinning rapidly provides the link. Although the magnetic fields of these two neutron star types are much stronger than the Earth's field, they are relatively weak by pulsar standards. Scientists think the weak magnetic field allows the accretion process to spin the star up to a high rotation rate.
After the accretion phase, X-ray emission from the pulsar ceases because there is no longer any infalling material to generate X-rays. The rotation speed begins to slow down at this point, because the accreting material was responsible for keeping the spin up as well. The pulsar's magnetic field rotates along with the star. The newly spun-up millisecond pulsar starts to emit radio waves as subatomic particles from its surface are accelerated into space by the pulsar's rotating magnetic field.
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"Alan Shepard, First American Astronaut, Dies At 74 (22 July, 1998)",656,0,0,0
Alan B. Shepard, Jr., the first American to fly in space and one of only 12 humans who walked on the Moon, died Tuesday night after a lengthy illness in Monterey, CA. He was 74. Shepard died at Community Hospital on the Monterey Peninsula, according to his family. The cause of death was not disclosed. Funeral services are pending.
"The entire NASA family is deeply saddened by the passing of Alan Shepard. NASA has lost one of its greatest pioneers; America has lost a shining star," said NASA Administrator Daniel S. Goldin. "Alan Shepard will be remembered, always, for his accomplishments of the past; being one of the original Mercury astronauts, for being the first American to fly in space, and for being one of only 12 Americans ever to step on the Moon. He should also be remembered as someone who, even in his final days, never lost sight of the future," Goldin added. "On behalf of the space program Alan Shepard helped launch, and all those that the space program has and will inspire, we send our deepest condolences to his wife, Louise, their children, and the rest of the Shepard family. Alan Shepard lived to explore the heavens. On this final journey, we wish him Godspeed."
"Alan Shepard is a true American hero, a pioneer, an original. He was part of a courageous corps of astronauts that allowed us to reach out into space and venture into the unknown," said George W.S. Abbey, Director of the Johnson Space Center, Houston, TX. "Alan Shepard gave all of us the privilege to participate in the beginnings of America's great adventure of human space exploration. He will be greatly missed. The program has lost one of its greatest supporters and a true friend. Our thoughts and prayers are with his wife, Louise, and their family."
Named as one of the nation's original seven Mercury astronauts in 1959, Shepard became the first to carry America's banner into space on May 5, 1961, riding a Redstone rocket on a 15-minute suborbital flight that took him and his Freedom 7 Mercury capsule 115 miles in altitude and 302 miles downrange from Cape Canaveral, FL. His flight followed by three weeks the launch of Soviet cosmonaut Yuri Gagarin, who on April 12, 1961, became the first human space traveler on a one-orbit flight lasting 108 minutes. Although the flight of Freedom 7 was brief, it nevertheless was a major step forward for the U.S. in a rapidly-accelerating race with the Soviet Union for dominance in the new arena of space.
Buoyed by the overwhelming response to Shepard's flight, which made the astronaut an instant hero and a household name, President John F. Kennedy set the nation on a course to the Moon, declaring before a joint session of Congress just three weeks later, "I believe this nation should commit itself to achieving the goal, before the decade is out, of landing a man on the Moon and returning him safely to the Earth." Over a three and a half year period from July 1969 to December 1972, a dozen Americans explored the lunar surface.
Shepard was the fifth man to walk on the Moon, and the oldest, at the age of 47. Shepard, however, was almost bypassed for a trip to the moon. He had to overcome an inner ear problem called Meuniere's syndrome that grounded him for several years following his initial pioneering flight. An operation eventually cured the problem and Shepard was named to command the Apollo 14 mission.
On January 31, 1971, Shepard, Command Module pilot Stuart Roosa and Lunar Module pilot Edgar Mitchell embarked for the Moon atop a Saturn 5 rocket. Shepard and Mitchell landed the lunar module Antares on February 5 in the Fra Mauro highlands while Roosa orbited overhead in the command ship Kitty Hawk. Shepard planted his feet on the lunar surface a few hours later, declaring, "Al is on the surface, and it's been a long way, but we're here."
During two excursions on the surface totaling nine hours, Shepard and Mitchell set up a science station, collected 92 pounds of rocks and gathered soil samples from the mountainous region. Near the end of the second moonwalk, and just before entering the lunar module for the last time, Shepard - an avid golfer - hit two golf balls with a makeshift club. The first landed in a nearby crater. The second was hit squarely, and in the one-sixth gravity of the moon, Shepard said it traveled "miles and miles and miles." Shepard's death leaves only four survivors among the original Mercury 7 astronauts: Sen. John Glenn, Scott Carpenter, L. Gordon Cooper and Walter Schirra.
Born Alan Bartlett Shepard, Jr. on Nov. 18, 1923, in East Derry, NH, he received a Bachelor of Science degree from the United States Naval Academy in 1944. Upon graduation, he married Louise Brewer, whom he met while at Annapolis. Shepard received his wings as a Naval aviator in 1947 and served several tours aboard aircraft carriers. In 1950, he attended Naval Test Pilot School at Patuxent River, MDS, and became a test pilot and instructor there. He later attended the Naval War College at Newport, RI, and after graduating, was assigned to the staff of the commander-in-chief, Atlantic Fleet, as an aircraft readiness officer.
In August 1974, Shepard, then a rear admiral, retired from both NASA and the Navy and became chairman of Marathon Construction Corp. in Houston. He later founded his own business company, Seven Fourteen Enterprises, named for his two missions on Freedom 7 and Apollo 14. In 1984, he and the other surviving Mercury astronauts, along with Betty Grissom, the widow of astronaut Virgil I. (Gus) Grissom, founded the Mercury Seven Foundation to raise money for scholarships for science and engineering students in college. In 1995, the organization was renamed the Astronaut Scholarship Foundation. Shepard was elected president and chairman of the foundation, posts he held until October 1997, when he turned over both positions to former astronaut James A. Lovell. Survivors include his widow, Louise, daughters Julie, Laura and Alice and six grandchildren.
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"NASA Satellite Sheds New Light on the La Nina Phenomenon (16 July, 1998)",657,0,0,0
Research scientists using data from the recently launched Tropical Rainfall Measuring Mission (TRMM) satellite, a joint U.S/Japanese mission, are shedding new light on the phenomenon known as La Nina. TRMM research team members have successfully retrieved sea-surface temperature data from the TRMM Microwave Imager (TMI) instrument onboard the spacecraft. This temperature data is giving scientists new insight into the complex evolution of the La Nina event -- the TMI is the only spaceborne microwave instrument observing sea-surface temperature in the tropics. The images show changes in sea-surface temperature, and ocean current movement and the dissipation of El Ni±o.
While it is too early to draw definite conclusions, the results to date appear to confirm the onset of La Nina type conditions. "TMI is an all-weather measuring instrument that can see through clouds," said Dr. David Adamec, oceanographer at the Goddard Space Flight Center, Greenbelt, MD. "The standard instrument (infrared radiometer), used to measure sea-surface temperature, must contend with clouds and atmospheric aerosols. Clouds block the flow of data, yet an uninterrupted consistent data stream is crucial for long-term climate study."
La Nina is essentially the opposite of the El Ni±o phenomenon and is characterized by unusually cold ocean temperatures in the equatorial Pacific, as compared to El Ni±o, where ocean temperatures are warmer than normal. La Nina and El Ni±o often are spoken of together and termed the El Ni±o/Southern Oscillations, or "ENSO." La Nina sometimes is referred to as the cold phase of the ENSO. At the Earth's surface, La Nina effects on the world's climate tend to be opposite those of El Ni±o.
At higher latitudes, El Ni±o and La Nina are just two of several factors that influence climate. However, the impacts of El Ni±o and La Nina at higher latitudes are most clearly seen in winter. During a typical La Nina year, winter temperatures are warmer than normal in the Southeast and cooler in the Northwest. Knowledge of La Nina is not as mature as that for El Ni±o. For example, every strong El Ni±o is not necessarily followed by a La Nina. Scientists at Goddard are performing advanced studies of El Ni±o and La Nina through information obtained from satellites in space and instruments in the oceans.
Acquiring quality sea-surface temperature data via a microwave scanner has been a long-term aspiration among oceanographers for more than a decade, when the last microwave imager ceased operations. In addition, none of the previously existing microwave scanners had the capability of the TRMM Microwave Imager. Ideally, this information will be used for the improvement of weather forecasting, anomalous weather study, and a better understanding of ocean current alteration.
Several NASA missions study the effects of El Ni±o and La Nina with orbiting satellites. The joint U.S.-French TOPEX/Poseidon satellite measures sea surface height; the Sea-Viewing Wide Field-of-View Sensor (SeaWiFS) measures ocean color; and TRMM measures precipitation and sea-surface temperature. The Tropical Atmosphere-Ocean Array consists of nearly 70 moored buoys in the tropical Pacific designed by the National Oceanic and Atmospheric Administration (NOAA).
The devices take real-time measurements of air temperature, relative humidity, surface winds, sea surface temperatures and subsurface temperatures down to a depth of 500 meters. Data from these moored buoys is processed by NOAA and then made available to scientists. The TRMM Microwave Imager instrument was provided by NASA. TRMM was developed jointly by NASA and NASDA and launched last November from NASDA's Tanegashima Space Center, Japan. This La Nina research is part of NASA's Earth Science Enterprise, a long-term research program designed to study the Earth's land, oceans, air, ice and life as a total system.
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"Efforts To Recover SOHO Spacecraft Continue (16 July, 1998)",658,0,0,0
NASA and European Space Agency (ESA) engineers, reasoning that over the next two-to-three months the spacecraft's solar panels will increasingly face the Sun and generate power, are continuing their efforts to contact the Solar and Heliospheric Observatory (SOHO) spacecraft. Meanwhile, the NASA/ESA investigation board concentrates its inquiry on three errors that appear to have led to the interruption of communications with SOHO on June 24. Officials remain hopeful that, based on ESA's successful recovery of the Olympus spacecraft after four weeks under similar conditions in 1991, recovery of SOHO may be possible.
The SOHO Mission Interruption Joint ESA/NASA Investigation Board has determined that the first two errors were contained in preprogrammed command sequences executed on ground system computers, while the last error was a decision to send a command to the spacecraft in response to unexpected telemetry readings. The spacecraft is controlled by a joint ESA/NASA Flight Operations Team, based at NASA's Goddard Space Flight Center, Greenbelt, MD. The first error was in a preprogrammed command sequence that lacked a command to enable an onboard software function designed to activate a gyro needed for control in Emergency Sun Reacquisition (ESR) mode. ESR mode is entered by the spacecraft in the event of anomalies.
The second error, which was in a different preprogrammed command sequence, resulted in incorrect readings from one of the spacecraft's three gyroscopes, which in turn triggered an Emergency Sun Reacquisition. At the current stage of the investigation, the board believes that the two anomalous command sequences, in combination with a decision to send a command to SOHO to turn off a gyro in response to unexpected telemetry values, caused the spacecraft to enter a series of Emergency Sun Reacquisitions, and ultimately led to the loss of control. The efforts of the investigation board are now directed at identifying the circumstances that led to the errors, and at developing a recovery plan should efforts to regain contact with the spacecraft succeed.
ESA and NASA engineers believe the spacecraft is currently spinning with its solar panels nearly edge-on towards the Sun, and thus not generating any power. Since the spacecraft is spinning around a fixed axis, as the spacecraft progresses in its orbit around the Sun, the orientation of the panels with respect to the Sun should gradually change. The orbit of the spacecraft and the seasonal change in the spacecraft-Sun alignment should result in the increased solar illumination of the spacecraft solar arrays over the next few months. The engineers predict that in late September 1998 illumination of the solar arrays and, consequently, power supplied to the spacecraft, should approach a maximum.
The probability of successfully establishing contact reaches a maximum at this point. After this time, illumination of the solar arrays gradually diminishes as the spacecraft-Sun alignment continues to change. In an attempt to recover SOHO as soon as possible, the Flight Operations Team is uplinking commands to the spacecraft via NASA's Deep Space Network, managed by NASA's Jet Propulsion Laboratory, Pasadena, CA, approximately 12 hours per day with no success to date. A recovery plan is under development by ESA and NASA to provide for orderly restart of the spacecraft and to mitigate risks involved.
The recovery of the Olympus spacecraft by ESA in 1991 under similar conditions leads to optimism that the SOHO spacecraft may be recoverable once contact is re-established. In May 1991, ESA's Olympus telecommunications satellite experienced a similar major anomaly which resulted in the loss of attitude, leading to intermittent power availability. As a consequence, there was inadequate communication, and the batteries and fuel froze. From analysis of the data available prior to the loss, there was confidence that the power situation would improve over the coming months. A recovery plan was prepared, supported by laboratory tests, to assess the characteristics of thawing batteries and propellants.
Telecommand access of Olympus was regained four weeks later, and batteries and propellant tanks were thawed out progressively over the next four weeks. The attitude was then fully recovered and the payload switched back on three months after the incident. Equipment damage was sustained as a result of the low temperatures, but nothing significant enough to prevent the successful resumption of the mission. The experience of Olympus is being applied, where possible, to SOHO and increases the hope of also recovering this mission. Estimating the probability of recovery is made difficult by a number of unknown spacecraft conditions.
Like Olympus, the hydrazine fuel and batteries may be frozen. Thermal stress may have damaged some of the scientific instruments as well. If the rate of spin is excessive, there may have been structural damage. SOHO engineers can reliably predict the spacecraft's orbit through November 1998. After that time, the long-term orbital behavior becomes dependent on the initial velocity conditions of the spacecraft at the time of the telemetry loss. These are not known precisely, due to spacecraft thruster activity that continued after loss of telemetry, so orbital prediction becomes very difficult.
Summing up the scientific returns from SOHO, which completed its two-year primary mission in April, Dr. George Withbroe, NASA's Director of the Sun-Earth Connections science program at NASA Headquarters said, "In the last two years, SOHO revolutionized our understanding of the Sun in many ways. It was a unique set of instruments devoted to the study of the most important star to us on Earth -- our Sun -- and we are very hopeful that the project engineers will be able to return this world-class observatory to science operations again."
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"New NASA Facility Will Complete Worldwide Communications Coverage (13 July, 1998)",659,0,0,0
Guam Island will be the site for a ribbon-cutting ceremony on July 15, 1998, to officially open a new terminal that will effectively complete NASA's vital communications and data-gathering support for NASA Earth-orbiting missions. Providing global, full-time and real-time communications support for NASA's Space Network customers, including the Space Shuttle, International Space Station and Hubble Space Telescope, the new ground terminal will be capable of communicating with geosynchronous tracking and data relay satellites stationed out of view of the existing Cacique and Danzante ground stations in White Sands, NM. NASA's Goddard Space Flight Center, Greenbelt, MD, manages the overall system.
"NASA built the Guam ground station to significantly expand the quantity and quality of services we provide to all our customers," said Goddard's ground terminal project manager, Tom Gitlin. Cost of funding the Guam station will be provided by NASA's Space Network operations budget and mitigated in part by the deactivation of the Canberra station. The Guam Remote Ground Terminal was conceived after NASA's Compton Gamma Ray Observatory suffered an onboard tape recorder failure in March 1992, and required full-time, real-time communications support. NASA established a limited capability ground terminal in Canberra, Australia, in late 1993 to provide continued support for the observatory's science mission. Goddard project officials quickly realized that an enhanced ground station was needed in the Pacific to better serve NASA's Space Network customers who traverse the Indian Ocean area.
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"NASA Instruments On Japanese Planet-B Spacecraft Will Aid Studies Of Martian Upper Atmosphere (1 July, 1998)",660,0,0,0
A NASA instrument to measure the gas composition of the upper atmosphere of Mars and hardware to support a radio science experiment will fly on a Japanese spacecraft known as Planet-B. The Neutral Mass Spectrometer (NMS) instrument and Ultra Stable Oscillator are scheduled for launch aboard Planet-B on July 3, 1998, from the Kagoshima Space Center on Kyushu Island, Japan.
"The Neutral Mass Spectrometer will enable us to measure the chemical composition of the upper atmosphere of Mars on a global scale, which has never been done before," said Dr. Hasso B. Niemann, the NMS principal investigator at NASA's Goddard Space Flight Center's Laboratory for Atmospheres in Greenbelt, MD. Previous upper atmospheric composition measurements were done in only two locations as NASA's Viking landers entered the Martian atmosphere on July 20 and Sept. 3, 1976, respectively.
The radio science hardware was built by the Johns Hopkins University Applied Physics Laboratory in Laurel, MD, under contract to NASA. The ultra-precise signals generated by the oscillator serve as a very accurate clock to enable analysis of the Martian atmosphere and to help guide the spacecraft as it orbits the red planet.
Planet-B is designed to perform long-term studies of the upper Martian atmosphere and ionosphere, and its interaction with the solar wind. Launch of Planet-B is scheduled for 2:12 p.m. EDT on July 3. After launch, the Planet-B spacecraft will be placed into Earth orbit and will use two swingbys past the Moon to establish conditions for a final trajectory to Mars.
Once the spacecraft reaches Mars, which is now scheduled for Oct. 11, 1999, it will be placed into a highly elliptical or "egg-shaped" orbit stretching from 93-186 miles (150-300 kilometers) to about 17,000 miles (27,300 kilometers) above the surface. The low-altitude portion of the orbit will be used for remote sensing of the lower atmosphere and surface, and for direct measurements of upper atmosphere and ionosphere. The more distant parts of the orbit will allow instruments to probe the ions and neutral gas escaping from Mars, which interact with the charged-particle "wind" blowing outward from the Sun. Ionization of the upper atmospheric gas by solar radiation produces the charged-particle atmosphere (ionosphere) that acts as an obstacle to the solar wind.
This radiation produces species of gas not seen in Mars' lower atmosphere, such as nitric oxide, or dissociates the atmosphere into single atomic species, such as atomic oxygen. If these neutral or ionized species possess enough energy, they can escape the gravitational pull of Mars, resulting in a net atmospheric loss. Measurements of lighter species such as atomic hydrogen and deuterium also can provide clues about the evolution of the Martian atmosphere.
Mars has little or no intrinsic magnetic field to interact with this process, making it more like Venus in this respect than Earth. The upper atmosphere of Venus and its solar wind environment were studied for almost 14 years by the U. S. Pioneer Venus Orbiter spacecraft from a similar, highly elliptical orbit. The Planet-B NMS instrument is a state-of-the-art enhancement of the Pioneer Venus mass spectrometer, weighing only six pounds (2.8 kilograms). To conserve space and weight, electronic items such as transistors and integrated circuits were removed from their outer casings and placed in larger packages called hybrid circuits.
Data from previous Mars exploration spacecraft such as Mariner 9 indicate that dust storms near the surface can heat the lower atmosphere and increase the gas density in the upper atmosphere where Planet-B will make its measurements. The U.S. Mars Surveyor 1998 mission known as the Mars Climate Orbiter, due for launch this December, carries an instrument called the Pressure Modulated Infrared Radiometer, which will provide complementary information on the lower atmosphere and its response to dust storms.
The Planet-B project is managed by the Institute of Space and Astronautical Science (ISAS) within the Japanese Ministry of Education. Planet-B carries 14 instruments from Japan, Canada, Sweden, Germany and the United States. ISAS personnel will operate the spacecraft and its instruments. The spacecraft was built by the Nippon Electric Corporation and will be launched by the new M-5 rocket. This rocket is designed to expand Japan's launch capability for the inner planets and beyond.
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"Efforts To Recover SOHO Spacecraft Continue (30 June, 1998)",661,0,0,0
Engineers are continuing efforts to reestablish contact with the NASA/European Space Agency (ESA) Solar and Heliospheric Observatory (SOHO) spacecraft using NASA's Deep Space Network (DSN). Contact with SOHO was lost on June 24 during maintenance operations. A team of experts from ESA and Matra Marconi Space, prime contractor for the SOHO spacecraft, gathered at NASA's Goddard Space Flight Center, Greenbelt, MD, to assist the NASA Flight Operations Team in assessing the situation and analyzing the spacecraft status should contact be reestablished.
Engineers are concentrating on gaining a full understanding of the events which led to the loss of signal, information which might help them devise procedures which may recover contact with SOHO. Commands are being sent to SOHO about once per minute through the DSN's 34-meter antennas instructing the spacecraft to activate its transmitters. Based on the last telemetry data received from SOHO, engineers said it appears most likely that the spacecraft is slowly spinning in such a way that its solar arrays, which generate power, either do not face the Sun at all or do not receive adequate sunlight to generate power.
However, based on the last data received, it appears that SOHO's solar panels may be exposed to an increasing amount of sunlight each day as it orbits the Sun. If this assumption is correct, within a few weeks enough sunlight might be hitting the solar panels to generate power to charge its batteries. The incident will be the subject of a joint ESA/NASA inquiry board co-chaired by Prof. Massimo Trella, ESA Inspector General, and Dr. Michael Greenfield, Deputy Associate Administrator for the Office of Safety and Mission Assurance, NASA Headquarters, Washington, DC. The other members of the board will be selected from ESA and NASA as well as from the scientific community. The board is expected to convene later this week at NASA's Goddard Space Flight Center, Greenbelt, MD.
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"Water History, Rock Composition Among Latest Findings a Year after Mars Pathfinder (29 June, 1998)",662,0,0,0
A year after the landing of Mars Pathfinder, mission scientists say that data from the spacecraft paint two strikingly different pictures of the role of water on the red planet, and yield surprising conclusions about the composition of rocks at the landing site. "Many of the things that we said last summer during the excitement after the landing have held up well," said Dr. Matthew Golombek, Pathfinder project scientist at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "But we have now had more time to study the data and are coming up with some new conclusions." Similar to ongoing science results from NASA's Mars Global Surveyor spacecraft currently in orbit around Mars, Pathfinder data suggest that the planet may have been awash in water three billion to 4.5 billion years ago.
The immediate vicinity of the Pathfinder landing site, however, appears to have been dry and unchanged for the past two billion years. Several clues from Pathfinder data point to a wet and warm early history on Mars, according to Golombek. Magnetized dust particles and the possible presence of rocks that are conglomerates of smaller rocks, pebbles and soil suggest copious water in the distant past.
In addition, the bulk of the landing site appears to have been deposited by large volumes of water, and the hills on the horizon known as Twin Peaks appear to be streamlined islands shaped by water. But Pathfinder images also suggest that the landing site is essentially unchanged since catastrophic flooding sent rocks tumbling across the plain two billion years ago. "Since then this locale has been dry and static," he said.
While the area appears to have been untouched by water for eons, wind appears to have been steadily eroding rocks at the landing site. Analysis of Pathfinder images shows that about one to two inches (three to five centimeters) of material has been stripped away from the surface by wind, Golombek noted. "Overall, this site has experienced a net erosion in recent times," said Golombek. "There are other places on Mars that are net 'sinks,' or places where dust ends up being deposited. Amazonis Planitia, for example, probably has about three to six feet (one to two meters) of fine, powdery dust that you would sink into if you stepped on it."
Chemical analysis of a number of rocks by the alpha proton X- ray spectrometer (APXS) instrument on Pathfinder's mobile Sojourner rover, meanwhile, reveals an unexpected composition that scientists are still trying to explain. The current assessment of data from this instrument suggests that all of the rocks studied by the rover resemble a type of volcanic rock with a high silicon content known on Earth as andesite, covered with a fine layer of dust. All of the rocks appear to be chemically far different from meteorites discovered on Earth that are believed to have come from Mars.
"The APXS tells us that all of these rocks are the same thing with different amounts of dust on them," said Golombek. "But images suggest that there are different types of rocks. We don't yet know how to reconcile this." When molten magma oozes up from a planet's mantle onto the surface of the outer crust, it usually freezes into igneous rock of a type that geologists call a basalt. This is typical on the floors of Earth's oceans, as well as on the maria or "seas" of the Moon and in many regions of Mercury and Venus.
By contrast, andesites typically form on Earth in tectonically active regions when magma rises into pockets within the crust, where some of its iron and magnesium-rich components are removed, leaving rock with a higher silicon content. 'We don't believe that Mars has had plate tectonics, so these andesites must have formed by a different mechanism," Golombek said. The rocks studied by Pathfinder most closely resemble andesites found in Iceland and the Galapagos Islands, tectonic spreading centers where plates are being pushed apart, said Dr. Joy Crisp of JPL. Andesites from these areas have a different chemical signature from andesites formed at subduction zones (areas where one edge of a crustal plate descends below another), mostly because wet ocean sediments carry more water down into the mantle at the subduction zones.
"On Mars, where the water content is probably lower and there is no evidence of subduction, we would expect a closer chemical similarity to Iceland andesites," said Crisp. The Martian rocks may have other origins, however. They could be sedimentary and influenced by water processes; they could be formed by melting processes resulting from a meteor impact; or, a third alternative is that the rocks might be basaltic, but covered by a silicon-rich weathering coating. "In any event, the presence of andesites on Mars is a surprise, if it is borne out as we study the data further," said Crisp. "Most rocks on Mars are expected to be basalts lower in silicon. If these are in fact andesites, they are probably not very abundant."
Scientists are looking forward to more data from the Thermal Emission Spectrometer instrument on the Mars Global Surveyor to reveal more about the chemical composition of the planet's surface, especially once the orbiting spacecraft begins its prime circular mapping mission in spring 1999. In other recent Pathfinder science findings, Dr. Steven Metzger of the University of Nevada found direct evidence of gusting winds called "dust devils" in images from Pathfinder's lander. Such dust devils had been seen in some Viking orbiter images and inferred from measurements of atmospheric pressure and winds by other instruments on the Pathfinder lander, but were not spotted in actual surface images until Metzger's discovery.
JPL planetary scientist Dr. Diana Blaney has been using data from Pathfinder, other spacecraft missions and ground-based observations to study weathering on Mars. Her work suggests that Mars is uniformly covered by a fine coating of dust formed by an unusual process involving meteor impacts and volcanic gases that add sulfur. NASA's next Mars missions, the 1998 Mars Climate Orbiter and Mars Polar Lander, are in testing now for launch in December and January, respectively. Whereas Pathfinder's science focus was on exploring rocks with its mobile robotic geologist, the Mars Polar Lander will focus on a search for water under the planet's surface, equipped with a robot arm that will dig into the soil at the landing site near the planet's south pole.
Launched on December 4, 1996, Pathfinder reached Mars on July 4, 1997, directly entering the planet's atmosphere and bouncing on inflated airbags as a technology demonstration of a new way to deliver a lander and rover to Mars. The lander operated nearly three times its design lifetime of 30 days, while the rover operated 12 times its design lifetime of seven days. During the mission, the spacecraft relayed 2.3 gigabits of data to Earth. This unexpectedly large volume of information included 16,500 images from the lander's camera, 550 images from the rover camera, 16 chemical analyses of rocks and soil, and 8.5 million measurements of atmospheric pressure, temperature and wind.
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"New Satellite Data Show Retreat of El Nino, Pacific Ocean in Transition (26 June, 1998)",663,0,0,0
New sea surface height measurements taken by the ocean-observing TOPEX/Poseidon satellite show the equatorial Pacific in a state of flux with the warm, high sea level El Nino-spawned waters in retreat and areas of colder, low sea level waters on the increase. "Sea level is a measure of the heat stored in the ocean. In the last month or so, the tropical Pacific has been switching from warm to cold. Lower sea level indicates less heat, hence a colder ocean," said Dr. Lee-Lueng Fu, the project scientist for the U.S.-French TOPEX/Poseidon mission at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
"It appears now the central equatorial Pacific Ocean will stay colder than normal for some time to come because sea level is about seven inches below normal, creating a deficit in the heat supply to the surface waters. It is not clear yet, however, if this current cooling trend will eventually evolve into a long-lasting La Nina situation." An El Nino condition begins when steady westward blowing trade winds weaken and even reverse direction. This change in the winds allows a large mass of warm water that is normally located near Australia to move eastward along the equator until it reaches the coast of South America. This displaced pool of unusually warm water affects evaporation, where rain clouds form -- and, in turn, alters the typical atmospheric jet stream patterns around the world.
The change in the wind strength and direction also impacts global weather patterns. The climatic event has been given the name El Nino, a Spanish term for "the Christ child," because the warm current first appeared off the coast of South America around Christmas. The 1997-98 El Nino has been the strongest ever recorded. This phenomenon was responsible for record rainfall in California, heavy flooding in Peru, drought and wildfires in Indonesia, tornadoes in the southeast United States and loss of life and property damage worldwide.
TOPEX/Poseidon's sea surface height measurements have provided scientists with their first detailed view of how El Nino's warm pool behaves because the satellite measures the changing sea surface height with unprecedented precision. A "La Nina" (Spanish for "little girl") is essentially the opposite of an El Nino, where the trade winds are stronger than normal and the cold water that normally exists along the coast of South America extends to the central equatorial Pacific.
A La Nina situation also changes global weather patterns and is associated with less moisture in the air, resulting in less rain along the coasts of North and South America. TOPEX/Poseidon will be able to track a potentially developing La Nina with the same accuracy. "It may be too soon to say 'good-bye' El Nino and 'hello' La Nina, because the effects of El Nino will remain in the climate system for a long time," said Dr. Bill Patzert, a research oceanographer at JPL.
"However, if the Pacific is transitioning to a La Nina, we'd expect to see clear, strong indication of it by late summer or early fall -- in approximately August or September -- just like we did last year with El Nino. The strongest impacts of a potential La Nina wouldn't be felt in the U.S. until next winter." A La Nina does not automatically follow an El Nino, Patzert added. Developed by NASA and the Centre National d'Etudes Spatiales (CNES), the TOPEX/Poseidon satellite uses an altimeter to bounce radar signals off the ocean's surface to get precise measurements of the distance between the satellite and the sea surface.
These data are combined with measurements from other instruments that pinpoint the satellite's exact location in space. Every ten days, scientists produce a complete map of global ocean topography, the barely perceptible hills and valleys found on the sea surface. Ocean temperatures affect ocean topography, which is why the TOPEX/Poseidon radar altimeter is able to monitor the changing El Nino and La Nina conditions. With detailed knowledge of ocean topography, scientists can then calculate the speed and direction of worldwide ocean currents.
Ground controllers lost contact with the NASA/European Space Agency (ESA) Solar and Heliospheric Observatory (SOHO) spacecraft June 24 during maintenance operations. SOHO went into emergency sun reacquisition mode, and ground controllers lost contact with the spacecraft at 7:16 p.m. EDT on June 24. This mode is activated when an anomaly occurs and the spacecraft loses its orientation toward the Sun. When this happens, the spacecraft automatically tries to point itself toward the Sun again by firing its attitude control thrusters under the guidance of an onboard Sun sensor.
Efforts to re-establish contact with SOHO did not succeed and telemetry was lost. Subsequent attempts using the full NASA Deep Space Network capabilities have so far also not been successful. Engineers from NASA and ESA are attempting to reestablish contact with the spacecraft. SOHO is a joint mission of the European Space Agency and NASA. It was launched aboard an Atlas IIAS rocket from Cape Canaveral Air Station, FL, on Dec. 2, 1995, and mission operations are directed from NASA's Goddard Space Flight Center, Greenbelt, MD.
In April 1998, SOHO successfully completed its nominal two-year mission to study the Sun's atmosphere, surface and interior. Major science highlights include the detection of rivers of plasma beneath the surface of the Sun; the discovery of a magnetic "carpet" on the solar surface that seems to account for a substantial part of the energy that is needed to cause the very high temperature of the corona, the Sun's outermost layer; the first detection of flare-induced solar quakes; the discovery of more than 50 sungrazing comets; the most detailed view to date of the solar atmosphere; and spectacular images and movies of coronal mass ejections, which are being used to improve the ability to forecast space weather.
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"Hubble Space Telescope Finds Evidence That Triton Is Warming Up (June 24, 1998)",665,0,0,0
Observations obtained by NASA's Hubble Space Telescope and ground-based instruments reveal that Neptune's largest moon, Triton, seems to have heated up significantly since the Voyager spacecraft visited it in 1989. "Since 1989, at least, Triton has been undergoing a period of global warming -- percentage-wise, it's a very large increase," said James L. Elliot, an astronomer at the Massachusetts Institute of Technology (MIT), Cambridge, MA. The warming trend is causing part of Triton's frozen nitrogen surface to turn into gas, thus making its thin atmosphere denser. Dr. Elliot and his colleagues from MIT, Lowell Observatory, and Williams College published their findings in the June 25 issue of the journal Nature. Even with the warming, no one is likely to plan a summer vacation on Triton, which is a bit smaller than Earth's moon.
The five percent increase means that Triton's temperature has risen from about 37 degrees on the absolute (Kelvin) temperature scale (-392 degrees Fahrenheit) to about 39 degrees Kelvin (-389 degrees Fahrenheit). If Earth experienced a similar change in global temperature over a comparable period, it could lead to significant climatic changes. Triton, however, is a very different and simpler world than Earth, with a much thinner atmosphere, no oceans, and a surface of frozen nitrogen. But the two share some contributing factors to global warming, such as changes to the Sun's heat output, how much sunlight is absorbed and reflected by their surfaces, and the amount of methane and carbon monoxide (greenhouse gases) in the atmosphere.
"With Triton, we can more easily study environmental changes because of its simple, thin atmosphere," Elliot explained. By studying these changes on Triton, the scientists hope to gain new insight into Earth's more complicated environment. Elliot and his colleagues explain that Triton's warming trend may be driven by seasonal changes in its polar ice caps.
Triton is approaching an extreme southern summer, a season that occurs every few hundred years. During this special time, the moon's southern hemisphere receives more direct sunlight, which heats the polar ice caps. "For a northern summer on Earth, it would be like the Sun being directly overhead at noon north of Lake Superior," Elliot said.
The scientists are basing a rise in Triton's surface temperature on the Hubble telescope's detection of an increase in the moon's atmospheric pressure, which has at least doubled in bulk since the time of the Voyager encounter. Any nitrogen ice on Triton that warms up a little results in a considerable leap in atmospheric pressure as the vaporized nitrogen gas joins the atmosphere. Because of the unusually strong link between Triton's surface ice temperature and its atmospheric pressure, Elliot says scientists can infer a temperature rise of three degrees Fahrenheit over nine years.
The scientists used one of Hubble's three Fine Guidance Sensors (used to keep the telescope pointed at a celestial target by monitoring the brightness of guide stars) in November 1997 to measure Triton's atmospheric pressure. When Triton passed in front of a star known as "Tr180" in the constellation Sagittarius, the guidance sensor measured the star's gradual decrease in brightness as Triton passed in front of it. The starlight became fainter as it traveled through Triton's thicker atmosphere.
Elliot and his colleagues list two other possible explanations for Triton's warmer weather. Because the frost pattern on Triton's surface may have changed over the years, it may be absorbing a little more of the Sun's warmth. Alternatively, changes in reflectivity of Triton's ice may have caused it to absorb more heat.
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"Radar System Aboard NASA DC-8 Aims To Unveil Mysteries of Cloud Structures (June 17, 1998)",666,0,0,0
In an effort to better understand clouds and how they affect our environment, a NASA DC-8 aircraft is flying this week carrying an airborne radar system designed to study the structure of clouds, including cloud liquid water content. During the current flights over the Southern United States, NASA's Airborne Cloud Radar is looking at clouds in an attempt to better understand how clouds warm or cool the Earth's atmosphere and how the presence of clouds influences the world's climate.
"Clouds represent a scientific puzzle that researchers have been trying to piece together for centuries," said Dr. Fuk Li, the principal investigator for the cloud radar at NASA's Jet Propulsion Laboratory, Pasadena, CA. "Scientists still don't know very much about the internal, vertical structures of clouds, and that leads to uncertainties in weather and climate predictions. Using the cloud radar, we will be able to study clouds in a new way that will help us understand their structure like never before. Once we have the cloud vertical structure information, atmospheric scientists will have a much better handle on long-term predictions of weather and climate change."
The cloud radar experiment was installed last month in the tail area (looking downward) of the DC-8, based at NASA's Dryden Flight Research Center, Edwards, CA. The DC-8 then flew to Tinker Air Force Base near Oklahoma City, OK, the origination point of this series of missions in which the radar collects cloud data while the plane flies above the clouds. Scientists will compare these data with measurements taken by satellite and ground-based sensors, including the Department of Energy's Southern Great Plains Cloud and Radiation Testbed, a series of instruments spread across north central Oklahoma and south central Kansas.
The radar, taking vertical measurements of the clouds from above, operates at 94 gigahertz, making it sensitive to cloud particles. The instrument transmits radar energy, which bounces off the cloud particles and is reflected back towards the aircraft. The radar measurements will be combined with information provided by other sensors to help analyze the properties of the clouds observed. This experiment has flown twice before aboard the DC-8 while still under development.
The DC-8's unique features make it an ideal platform for examining cloud structures, according to DC-8 mission manager Chris Jennison of Dryden. "The DC-8 was selected because it is the only aircraft that is capable of this mission in terms of altitude, speed, range and capacity for carrying scientists onboard. Since scientists can fly on the aircraft, they can operate their experiments themselves," Jennison said.
This Airborne Cloud Radar flight series is expected to total 20 flight hours. The experiment is designed to test several hypotheses and techniques related to satellite remote sensing of extensive, long-lasting, non-precipitating layers of cloud in the middle and upper troposphere -- atmosphere up to seven miles from the Earth's surface. It is expected that this instrument will be used in upcoming field experiments to better understand cloud-climate processes. One such planned experiment is the Tropical Cirrus Experiment called CRYSTAL planned for 2001. Eventually this instrument will be flown on satellite platforms designed to observe Earth's climate processes from space.
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"Arctic Crater Expedition To Seek Mars Science Insights (June 16, 1998)",667,0,0,0
NASA scientists soon will explore a barren Arctic meteorite impact crater to attempt to learn more about Mars and its early history, while testing technologies useful for future robotic and human exploration of the planet. From June 22 to July 26, a 20-member science team from NASA and several other research organizations will explore the Haughton Impact Crater and its surroundings on Devon Island in the Arctic Circle.
Scientists consider the site a potential Mars analog because many of its geologic features, such as the crater's ice-rich terrains, its ancient lake sediments and nearby networks of small valleys, resemble those reported at the surface of Mars. The site may shed light in particular on the early history of Mars, when the planet's climate may have been wetter and warmer. "The cold, relatively dry, windy and unvegetated environment at the Haughton site is milder and wetter than present-day Mars, but it may give us an idea of what early Mars was like and how some of its surface features were formed," said Principal Investigator Dr. Pascal Lee of NASA's Ames Research Center, Moffett Field, CA.
During the expedition, Dr. Omead Amidi and other engineers from Carnegie Mellon University's Robotics Institute, Pittsburgh, PA, will conduct field tests of an experimental, robotic helicopter. "The mission provides a great opportunity to demonstrate the feasibility and the value of robotic aircraft for mapping and surveying applications," Amidi said. Carnegie Mellon's small, 160-pound autonomous helicopter has vision-based stability and position control, as well as an onboard navigation computer, laser rangefinder and video system for site mapping.
In addition to the tests with the autonomous helicopter, scientists also will conduct experiments with a ground-penetrating radar system, a field spectrometer, drilling equipment and a stereo camera. The radar system will be deployed in an attempt to map ground-ice and other subsurface conditions within and outside the crater's 12- mile (20-kilometer) diameter. "The ability to find underground ice, both for human consumption and geologic studies, will be critical in the exploration of Mars," said Dr. Aaron Zent of Ames, Dr. Lee's post- doctoral research advisor.
Scientists will use a field spectrometer to determine the site's reflective qualities and better understand the crater's compositional evolution. In another experiment, scientists will use a portable drill to obtain core samples from ten feet deep in the frozen ground. Core samples of sediments from a lake that once occupied the crater will provide information about local climate evolution. Since the use of liquid drilling lubricants might be precluded on Mars, none will be used in this test.
A portable stereo camera system previously used by Carnegie Mellon's Nomad rover during its unprecedented 133-mile wheeled trek through Chile's Atacama Desert last summer will provide high- resolution images of the site, and produce images for a 360 degree photo-realistic virtual reality project being developed by Ames' Intelligent Mechanisms Group. Using laptop computer systems and "mobile workstations" developed by Ames' Intelligent Mobile Technologies Team, scientists will communicate with other field team members and send live images via a wireless link. Team members will operate from a base camp on a terrace of the Haughton River within the crater's perimeter and explore the site with All-Terrain Vehicles. Supplies will be brought in by Twin Otter airplane, while a helicopter will aid exploration of remote sites.
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"SOHO Spacecraft Sees Two Comets Plunge Into Sun (June 3, 1998)",668,0,0,0
In a rare celestial spectacle, two comets have been observed plunging into the Sun's atmosphere in close succession, on June 1 and 2. This unusual event on Earth's own star was followed on June 2 by a likely unrelated but also dramatic ejection of solar gas and magnetic fields on the southwest (or lower right) limb of the Sun.
The observations of the comets and the large erupting prominence were made by the LASCO coronagraph on the Solar and Heliospheric Observatory (SOHO) spacecraft. Science instruments on SOHO have discovered more than 50 comets, including many so- called sun grazers, but none in such close succession. The eruption of solar gas was directed away from Earth and does not pose a hazard to our planet or orbiting astronauts.
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"Space Grown Insulin Crystals Provide New Data On Diabetes (June 3, 1998)",669,0,0,0
Diabetic patients may someday reduce their insulin injections and lead more normal lives because of new insights gained through innovative space research in which the largest insulin crystals ever studied were grown on the Space Shuttle. Results from a 1994 insulin crystal growth experiment in space are leading to a new understanding of diabetes -- a hormone deficiency disease. This has the potential to significantly reduce expensive treatments, since treatment of diabetes accounts for one-seventh of the nation's health care costs.
Sixteen million Americans suffer from hormone deficiency diseases such as diabetes, hepatic failure, hemophilia, Parkinson's and Huntington diseases. "The space-grown insulin crystals have provided us new, never-before-seen information," said Dr. G. David Smith, scientist at Hauptman-Woodward Medical Research Institute, Buffalo, NY. "As a result, we now have a much more detailed picture of insulin," Smith said. Because of the increase in crystal size, Smith's team is able to study in more detail the delicate balance of the insulin molecule. Natural insulin molecules hold together and gradually release into the human body.
With some of the new and unexpected findings, researchers may be able to improve how insulin is released from its inactive-stored state to its active state. This could greatly improve the quality-of-life of people on insulin therapy by cutting down on the number of injections they have to take. "This new information can be used in the development of a new therapeutic insulin treatment for the control of diabetes," said Smith. Hauptman-Woodward is partnering with the Center for Macromolecular Crystallography, a NASA Commercial Space Center in Birmingham, AL. "We are doing crystal growth experiments in the near- weightlessness of space that really tell the story of how insulin works and give us clues of how, in the long run, to defeat diabetes," said Dr. Marianna M. Long, associate director of the center located at the University of Alabama at Birmingham.
Insulin is one of the most important hormones in the human body because it regulates the body's blood sugar levels. In people with diabetes, insulin is not produced in sufficient quantity, nor regulated properly. This metabolism disorder impairs the body's ability to use digested food for growth and energy. Like many chemicals in the body, the three-dimensional structure of insulin is extremely complex. The intricate, blueprint-like arrangement of atoms within the insulin molecule determines how well the hormone interacts within the body. When grown on the ground, insulin crystals do not grow as large or as ordered as researchers desire -- obscuring the blueprint of the insulin molecules.
The center in Birmingham is one of NASA's 10 Commercial Space Centers managed by the Space Product Development Office within the Microgravity Research Program Office at NASA's Marshall Space Flight Center in Huntsville, AL. Each center represents a NASA partnership with industry and academia, pursuing product-oriented research in areas such as biotechnology, agriculture and materials. Unique research opportunities of the space environment are made available to encourage private industries to exploit the benefits of space-based research to develop new products or services. NASA research has furthered the understanding of many diseases, including AIDS, heart disease, cancer, respiratory syncytial virus, sickle cell anemia, hepatitis and rheumatoid arthritis.
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"NASA Forms Office In Russia In Preparation For ISS Operations (1 June, 1998)",670,0,0,0
NASA has formed the Office of Human Space Flight Programs, Russia, to oversee the transition from the Phase One Shuttle-Mir program to the assembly and operation of the new International Space Station. Astronaut Michael A. Baker (Captain, USN) will lead this office. Baker recently was named Assistant Director to Johnson Space Center Director George W. S. Abbey to supervise the transition of human space flight initiatives associated with this cooperative effort.
Baker will be NASA's lead representative to the Russian Space Agency and its contractors on operational issues as part of NASA's Human Exploration and Development of Space (HEDS) initiative. This places Russian liaison for all human space flight operations and initiatives under one office and consolidates preparations for the assembly of the International Space Station, including mission operations, crew training, logistics and technical liaison activities with Russian space organizations.
"It is critical to ensure a seamless transition takes place between the successful Phase One program and the start of construction of the International Space Station," Abbey said. "With this new office, the ISS program can take advantage of the knowledge and momentum gained from Phase One under the direction of Frank Culbertson.
"Baker's leadership and expertise provide the framework to coordinate activities with our Russian colleagues and provide him the unique opportunity to bring the operations team together for this project," Abbey said.
Baker has flown four shuttle missions, including his most recent flight as commander of STS-81 aboard Atlantis in January 1997, the fifth docking of a shuttle to the Mir Space Station. Prior to that flight, Baker served as the NASA Director of Operations at Star City from March to October 1995.
Other activities in Russia that fall under the leadership of Baker include the astronaut training responsibilities at the Gagarin Cosmonaut Training Center outside Moscow at Star City and all NASA mission operations functions at the Mission Control Center in Korolev.
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"International Space Station Adjusts Target Dates For First Launch (June 1, 1998)",671,0,0,0
Representatives of all nations involved in the International Space Station have agreed to officially target a November 1998 launch for the first station component and to revise launch target dates for the remainder of the 43-flight station assembly plan. In meetings of the Space Station Control Board and the Heads- of-Agency on May 30 and 31 at NASA's Kennedy Space Center, FL, all station partners agreed to target launch dates of Nov. 20, 1998, for the Control Module today named Zarya (Russian word for sunrise) and Dec. 3, 1998, for Shuttle mission STS-88 with Unity.
Changes in the construction schedule for the third station component, the Russian-provided Service Module, led the partners to reschedule the first assembly launches. The Service Module will house the first station occupants and the European Space Agency-provided Data Management System. Although the new dates move the launch of the first station component, Zarya, from June to November, the target dates agreed upon for many major station milestones during the latter portions of the five-year assembly plan are little changed.
In addition, several enhancements to the station's assembly have been made, including an exterior "warehouse" for spare parts and a Brazilian- provided carrier for exterior station components that are launched aboard the Space Shuttle. The International Space Station partners set an April 1999 target launch date for the Russian Service Module. The first station crew -- Commander Bill Shepherd, Soyuz Commander Yuri Gidzenko and Flight Engineer Sergei Krikalev -- will be launched aboard a Russian Soyuz spacecraft in summer 1999 to begin a five- month inaugural stay. Launch of the U.S. laboratory module is set for October 1999. Launches of other laboratory modules provided by Europe, Japan and Russia, will take place later in the assembly sequence.
The Canadian-provided station robotic arm, or Space Station Remote Manipulator System, will be launched in December 1999. Scientific research will commence aboard the station early in the year 2000. The expansion from a three-person crew to a six-person capability is planned in November 2002. The final launch in the assembly sequence is set for January 2004, only one month later than in the previous assembly plan.
Some issues in this assembly sequence remain under review and will be resolved at a Space Station Control Board meeting in September. NASA continues the development of an Interim Control Module (ICM) as a contingency against further delays in the Service Module and as a potential additional propellant capability for a more robust space station. A decision concerning the configuration of the ICM will be made later this year. During the Heads-of-Agency meeting, the Russian Space Agency (RSA) stated that the Russian government has made the International Space Station its number one civil space priority.
RSA noted that progress on the Service Module continues to meet a launch in April 1999. RSA also is working to deorbit Mir as early as safely possible, with a goal of developing a potential to deorbit by July 1999. The international partners expressed their concern with delays to the International Space Station program to date and brought to the attention of RSA that it is critical to all participating nations that the station program schedule is met.
The agencies' leaders also acknowledged the atmosphere of cooperation, the accomplishments and the successful achievements of the Shuttle-Mir Program (Phase I) and look forward to the smooth transition to Phases 2 and 3 of the International Space Station. In addition, they highlighted the ongoing International Space Station training currently under way for the first four station crews.
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"Spacecraft Images Capture Magnetic Energy Burst On Sun (May 29, 1998)",672,0,0,0
The first images from NASA's Transition Region and Coronal Explorer (TRACE) spacecraft reveal activity in the solar atmosphere in stunning detail and include the first detailed observations of a magnetic energy release, called a magnetic reconnection. The magnetic reconnection was observed on May 8, 1998, in a region of the solar atmosphere where two sets of perpendicular magnetic loops expanded into each other. Magnetic reconnection occurs when magnetic fields "snap" to a new, lower energy configuration, much like when a twisted rubber band unwinds or breaks.
A magnetic reconnection can release vast amounts of energy and is responsible for explosive events on the Sun, such as flares, that can cause communication and power system disruptions on Earth. High resolution movies of a relatively small but clear magnetic reconnection event and other spectacular solar activity observed by TRACE were presented today during the spring meeting of the American Geophysical Union in Boston. "The TRACE spacecraft is unique in that it has both high spatial and temporal resolution in the extreme ultraviolet, wavelengths of light that reveal the multimillion degree temperature of the Sun," said Dr. Alan Title, TRACE Principal Investigator from the Stanford Lockheed Institute for Scientific Research (SLISR) in Palo Alto, CA.
"We can image solar activity in finer detail than existing spacecraft, and we can take a new image once every few seconds. Both are necessary for our mission, which is to understand in great detail how energy is transported from the solar surface into the outer atmosphere. In the past, spacecraft of lower resolution were forced to average over much larger areas and periods of time. This made it difficult to get at the fundamental physics." "In our magnetic reconnection movie, we can distinguish the fine details of the magnetic fields and see how they change during time periods of about a minute.
TRACE has given us many surprises, and new ones occur nearly every observation. We found that even large areas of the Sun, some more than 60,000 miles long, can heat up or cool down significantly and thus appear and disappear in just a few minutes," said Title. The TRACE spacecraft, launched from Vandenberg AFB, CA, on April 1, 1998, joins a multinational fleet of International Solar Terrestrial Physics project spacecraft studying the Sun during a critical period when solar activity is beginning its rise to a peak early in the new millennium.
The Sun goes through an 11-year cycle from a period of numerous intense storms and sunspots to a period of relative calm and then back again. The coming months in the Sun's cycle will provide solar scientists with periods of intense solar activity interspersed with periods when the Sun is relatively passive and quiet. This will give TRACE the chance to study the full range of solar conditions, even in its relatively short planned lifetime.
TRACE is training its powerful telescope on the so-called "transition region" of the Sun's atmosphere, a dynamic region between the relatively cool surface and lower atmosphere regions of the Sun (about 10,000 degrees Fahrenheit) and the extremely hot upper atmosphere called the corona (up to three million degrees Fahrenheit). Using portions of the telescope sensitive to extreme-ultraviolet and ultraviolet wavelengths of light, TRACE is studying the detailed connections between the fine-scale surface features and the overlying, changing atmospheric structures of hot, electrically charged gas called plasma.
The surface features and atmospheric structures are linked by fine-scale solar magnetic fields. The solar atmosphere is constantly evolving because the magnetic fields that dominate the corona are continuously displaced by the convective motions in the outer layers of the Sun just below the photosphere. The TRACE science team also will study the evolution of events, such as massive flarings and huge eruptions, in the Sun's atmosphere. These events originate at the Sun's visible surface, the photosphere, and travel upward through its atmosphere (chromosphere and transition region), and then into its super-hot corona before speeding out into space, sometimes towards Earth.
The power of the TRACE telescope to do detailed studies of the solar atmosphere makes this observatory unique among the current group of spacecraft studying the Sun. The spacecraft has roughly 10 times the temporal resolution and five times the spatial resolution of previously launched solar spacecraft. A Sun-synchronous orbit is uninterrupted by Earth's shadow for eight months at a time, allowing the mission the greatest chance to observe the random processes which lead to flares and massive eruptions in the Sun's atmosphere. The TRACE core team consists scientists from Lockheed Martin Advanced Technology Center, Stanford University, NASA's Goddard Space Flight Center, the University of Chicago, Montana State University, and the Harvard-Smithsonian Center for Astrophysics.
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"Fire and Ice: Arctic Expedition Probes Roles of Clouds in Climate Change (May 29, 1998)",673,0,0,0
An ice-breaking ship, research airplanes, space satellites and an international team of scientists are converging in the Alaskan Arctic this month to learn more about global climate change through the study of clouds and radiation of the Sun during the spring and summer. The FIRE Arctic Cloud Experiment (FIRE/ACE) is studying a variety of cloud systems in a two-phase campaign April 7 through June 13 and July 6 through 30. FIRE (First International Satellite Cloud Climatology Project Regional Experiment) is led by NASA, in collaboration with other government and private organizations, and will take place in Alaska in the Beaufort Sea and in the skies over the coastal town of Barrow.
"We know very little about Arctic clouds and how they interact with the polar surface and atmosphere," said NASA Langley's Dr. Patrick Minnis, FIRE project scientist. "The data from FIRE/ACE will provide the opportunity to greatly expand our knowledge of the Arctic climate -- an important component in any global climate change scenario. "The ultimate goal is to learn enough to more accurately forecast global climate change," Minnis added. "The better we can understand it, the more we'll be able to determine the possible effects of potential global change, such as iceberg melting and coastline flooding."
The current campaign will use four aircraft, numerous Earth-bound instruments and six orbiting satellites to take measurements of clouds from high above them, inside them, and below them. A ER-2 airplane, based at NASA's Dryden Flight Research Center, Edwards, CA, is scheduled to arrive in Alaska in mid-May. During missions the aircraft will fly at about 70,000 feet with a suite of remote sensors to study clouds that form in the vicinity of "leads" -- large cracks that expose water -- and melt ponds. Three other aircraft will participate as well.
A University of Washington CV-580, a National Center for Atmospheric Research C-130, and a Canada National Research Council Convair CV-580 will use a variety of instruments to take measurements "in situ," or inside the clouds. Additionally the Canadian Coast Guard icebreaker vessel, the "Des Groseilliers," has been frozen into the Arctic ice pack for a year as part of a climate-modeling project sponsored by the National Science Foundation and Office of Naval Research. Scientists making measurements from the ice will share data with scientists making measurements on this spring's series of overflights. NASA is the lead agency for FIRE.
In Canada, the lead agency is the Atmospheric Environment Service of Environment Canada. Other supporting agencies are the National Research Council of Canada, and in the U.S., the National Science Foundation, Office of Naval Research, Department of Energy, National Oceanic and Atmospheric Administration, and Department of Defense. Participating will be more than 80 researchers from eight NASA centers, five U.S. agencies, 13 U.S. universities and educational consortia, and three private U.S. companies, as well as scientists from Canada, Great Britain, and the Netherlands.
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"Hubble Finds a Runaway World (May 28, 1998)",674,0,0,0
NASA's Hubble Space Telescope has given astronomers their first direct look at what is possibly a planet outside our solar system -- one apparently that has been ejected into deep space by its parent stars. The discovery, made by Susan Terebey of the Extrasolar Research Corporation in Pasadena, CA, and her team using Hubble's Near Infrared Camera and Multi-Object Spectrometer (NICMOS), further challenges conventional theories about the birth and evolution of planets, and offers new insights into the formation of our own Solar System.
Located in the sky within a star-forming region in the constellation Taurus, the object, called TMR-1C, appears to lie at the end of a strange filament of light that suggests it has apparently been flung away from the vicinity of a newly forming pair of binary stars. At a distance of 450 light-years, the same distance as the newly formed stars, the candidate protoplanet would be ten thousand times less luminous than the Sun. If the object is a few hundred thousand years old, the same age as the newly formed star system which appears to have ejected it, then it is estimated to be 2-3 times the mass of Jupiter, the largest gas giant planet in our Solar System.
Also possible is that the object is up to ten million years old, the same age as other young stars nearby, in which case it may be a giant protoplanet or a brown dwarf star. A brown dwarf star is a small star that has failed to sustain nuclear fusion. The candidate protoplanet is now 130 billion miles from the parent stars and predicted to be hurtling into interstellar space at speeds up to 20,000 miles per hour (10 kilometers/sec) -- destined to forever drift among the Milky Way's starry population.
Hubble researchers estimate the odds at two percent that the object is instead a chance background star. "If the results are confirmed, this discovery could be telling us gas giant planets are easy to build. It seems unlikely for us to happen to catch one flung out by the stars unless gas giant planets are common in young binary systems," said Terebey. "The results don't directly tell us about the presence of any terrestrial planets, like Earth," she adds. "However, we believe gas giants do influence the formation of much smaller rocky planets." Current models predict that very young giant planets are still warm from gravitational contraction and formation processes.
This makes them relatively bright in infrared light compared to old giant planets such as Jupiter. Even so, young planets are difficult to find in new solar systems because the glare of the central star drowns out their feeble glow. Young planets ejected from binary systems would therefore represent a unique opportunity to study extrasolar planets with current astronomical technology. The discovery also challenges conventional theories that predict gas giant planets take millions of years to coagulate from dust in space.
Instead, it favors more recent ideas that large, low-density planets may condense out of gas very quickly, at the same time their parent star does. "This observation pushes back the clock on planet formation and offers short time scales which allow us to see how things form. This provides valuable new clues to the origin of our Solar System," says Terebey. The candidate protoplanet was accidentally discovered by Terebey and colleagues while studying Hubble infrared images of newly formed protostars in a molecular cloud in Taurus.
The exquisite sensitivity and sharpness of NICMOS clearly revealed the object's pinpoint image. However, it might have been dismissed as a background star if not for the presence of a bizarre 130- billion-mile-long filamentary structure that bridges the space between the binary pair and the candidate protoplanet. "I said to myself, 'This is really weird, what in the world could it be?'" recalls Terebey. She speculates it could be a tunnel the runaway object burrowed through a dust cloud surrounding the stars. This created a "light tube" which channels light from the stars deep inside their dusty cocoon - like a light beam traveling through a length of fiber optic cable.
This brought Terebey to the tantalizing possibility that the planet had been flung into deep space by a gravitational "slingshot" effect from its parent stars. This could have happened if the planet's orbit allowed it to rob momentum from the stars and pick up so much speed that it escaped the system, similar to the way spacecraft perform gravitational "slingshot" maneuvers to pick up speed by flying close by a planet. "We know that many triple star systems eventually toss out the lowest mass star. And we can predict the speed at which the object should be moving, based on the separation of the binary stars," said Terebey.
Future observations call for images taken at a later date, to confirm the object's predicted movement across the sky. In addition, the spectrum of the object will tell whether the object is a background star, brown dwarf, or something whose spectrum is less easy to predict, such as a giant protoplanet. "We will just have to wait and see if future observations confirm this picture," said Terebey. "However it turns out, we have come to appreciate that protoplanet ejection by young binary stars ought to happen, and it offers a new way to search for giant planets." "These future observations will be critical in verifying that this object is truly a planet and not a brown dwarf," said Dr. Ed Weiler, Director of the Origins Program at NASA Headquarters, Washington, DC.
"We are sharing this preliminary data with the public at a very early stage in the research process because of its potential importance and because of the compelling nature of the image. If the planet interpretation stands up to the careful scrutiny of future observations, it could turn out to be the most important discovery by Hubble in its 8 year history".
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"Surveyor Data Reveal More Evidence of Mars' Past (May 27, 1998)",675,0,0,0
New mineralogical and topographic evidence suggesting that Mars had abundant water and thermal activity in its early history is emerging from data gleaned by NASA's Mars Global Surveyor spacecraft. Scientists are getting more glimpses of this warmer, wetter past on Mars while Global Surveyor circles the planet in a temporary 11.6-hour elliptical orbit. Findings from data gathered during the early portions of this hiatus in the mission's orbital aerobraking campaign are being presented today at the spring meeting of the American Geophysical Union in Boston.
Among many results, the Thermal Emission Spectrometer instrument team, led by Dr. Philip Christensen of Arizona State University, Tempe, has discovered the first clear evidence of an ancient hydrothermal system. This finding implies that water was stable at or near the surface and that a thicker atmosphere existed in Mars' early history. Measurements from the spectrometer show a remarkable accumulation of the mineral hematite, well-crystallized grains of ferric (iron) oxide that typically originate from thermal activity and standing bodies of water.
This deposit is localized near the Martian equator, in an area approximately 300 miles (500 kilometers) in diameter. Fine-grained hematite, with tiny particles no larger than specks of dust, generally forms by the weathering of iron-bearing minerals during oxidation, or rusting, which can occur in an atmosphere at low temperatures. The material has been previously detected on Mars in more dispersed concentrations and is widely thought to be an important component of the materials that give Mars its red color. The presence of a singular deposit of hematite on Mars is intriguing, however, because it typically forms by crystal growth from hot, iron-rich fluids.
Meanwhile, the Mars Orbiter Laser Altimeter instrument is giving mission scientists their first three-dimensional views of the planet's north polar ice cap. Principal Investigator Dr. David Smith of NASA's Goddard Space Flight Center, Greenbelt, MD, and his team have been using the laser altimeter to obtain more than 50,000 measurements of the topography of the polar cap in order to calculate its thickness, and learn more about related seasonal and climatic changes.
These initial profiles have revealed an often striking surface topology of canyons and spiral troughs in the water and carbon dioxide ice that can reach depths as great as 3,600 feet below the surface. Many of the larger and deeper troughs display a staircase structure, which may ultimately be correlated with seasonal layering of ice and dust observed by NASA's Viking mission orbiters in the late 1970s. The laser data also have shown that large areas of the ice cap are extremely smooth, with elevations that vary only a few feet over many miles. At 86.3 degrees north, the highest latitude yet sampled, the cap achieves an elevation of 6,600 to 7,900 feet (1.25 to 1.5 miles or 2-2.5 kilometers) over the surrounding terrain.
The laser measurements are accurate to approximately one foot (30 centimeters) in the vertical dimension. In June, the ice cap's thickness will reach a maximum during the peak of the northern winter season. Thickness measurements from April will be compared to those that will be taken in June, contributing to a greater understanding of the Martian polar cap's formation and evolution.
In addition, the Global Surveyor accelerometer team, led by Dr. Gerald Keating of George Washington University, Washington, DC, has discovered two enormous bulges in the upper atmosphere of Mars in the northern hemisphere, on opposite sides of the planet near 90 degrees east latitude and 90 degrees west longitude. These bulges rotate with the planet, causing variations of nearly a factor of two in atmospheric pressure, and systematic variations in the altitude of a given constant pressure of about 12,000 feet (four kilometers).
After a month-long period during which the Sun was between Earth and Mars and thus degraded communications with Global Surveyor, the spacecraft has resumed taking scientific data in its temporary elliptical orbit. In September, it will once again begin dipping into the upper atmosphere of Mars each orbit in a process called aerobraking. The drag from this procedure will allow the spacecraft to reach a low circular orbit and begin its primary two-year global mapping mission starting in March 1999.
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"Solar Flare Leaves Sun Quaking (May 27, 1998)",676,0,0,0
Scientists have shown for the first time that solar flares produce seismic waves in the Sun's interior that resemble those created by earthquakes: They observed a flare-generated solar quake that contained about 40,000 times the energy released in the great 1906 earthquake that devastated San Francisco. (The amount of energy released was enough to power the United States for 20 years at its current level of consumption, and was equivalent to an 11.3 magnitude earthquake.)Dr. Alexander G. Kosovichev, a senior research scientist from Stanford University, and Dr. Valentina V. Zharkova from Glasgow (United Kingdom) University found the tell-tale seismic signature in data on the Sun's surface.
The data were collected by the Michelson Doppler Imager onboard the Solar and Heliospheric Observatory (SOHO) spacecraft immediately following a moderate-sized flare on July 9, 1996. "Although the flare was a moderate one, it still released an immense amount of energy," said Dr. Craig Deforest, a researcher with the SOHO project. "The energy released is equal to completely covering the Earth's continents with a yard of dynamite and detonating it all at once. "SOHO is a joint project of the European Space Agency and NASA. The finding is reported in the May 28 issue of the journal Nature, and will be the subject of a press conference at the spring meeting of the American Geophysical Union in Boston, MA, May 27 at 9 a.m. EDT.
The solar quake that the science team recorded looks much like ripples spreading from a rock dropped into a pool of water. But over the course of an hour, the solar waves traveled for a distance equal to 10 Earth diameters before fading into the fiery background of the Sun's photosphere. Unlike water ripples that travel outward at a constant velocity, the solar waves accelerated from an initial speed of 22,000 miles per hour to a maximum of 250,000 miles per hour before disappearing. "People have looked for evidence of seismic waves from flares before, but they didn't have a theory so they didn't know where to look," says Kosovichev.
Several years ago Kosovichev and Zharkova developed a theory that can explain how a flare, which explodes in space above the Sun's surface, can generate a major seismic wave in the Sun's interior. According to the currently accepted model of solar flares, the primary explosion creates high-energy electrons (electrically charged subatomic particles). These are funneled down into a magnetic flux tube, an invisible tube of magnetic energy, and produce X-rays, microwaves and a shock wave that heats the solar surface. Kosovichev and Zharkova developed a theory that predicts the nature and magnitude of the shock waves that this beam of energetic electrons should create when they slam down into the solar atmosphere.
Although their theory directed them to the right area to search for the seismic waves, the waves that they found were 10 times stronger than they had predicted. "They were so strong that you can see them in the raw data," Kosovichev says. The solar seismic waves appear to be compression waves like the "P" waves generated by an earthquake. They travel throughout the Sun's interior. In fact, the waves should recombine on the opposite side of the Sun from the location of the flare to create a faint duplicate of the original ripple pattern, Kosovichev predicts.
Now that they know how to find them, the SOHO scientists say that the seismic waves generated by solar flares should allow them to verify independently some of the conditions in the solar interior that they have inferred from studying the pattern of waves that are continually ruffling the Sun's surface. SOHO is part of the International Solar-Terrestrial Physics (ISTP) program, a global effort to observe and understand our star and its effects on our environment. The ISTP mission includes more than 20 satellites, coupled with ground-based observatories and modeling centers, that allow scientists to study the Sun, the Earth, and the space between them in unprecedented detail. ISTP is a joint program of NASA, ESA, Japan's Institute for Astronautical Science, and Russia's Space Research Institute.
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"Strongest Stellar Magnetic Field Yet Observed Confirms Existence of Magnetars (May 20, 1998)",677,0,0,0
A neutron star, located 40,000 light years from Earth, is generating the most intense magnetic field yet observed in the Universe, according to an international team of astronomers led by scientists at NASA's Marshall Space Flight Center in Huntsville, AL.
The discovery confirms the existence of a special class of neutron stars dubbed "magnetars." Magnetars have a magnetic field estimated to be one thousand trillion times the strength of Earth's magnetic field. A neutron star is a burned-out star roughly equal in mass to the Sun that has collapsed through gravitational forces to be only about 10 miles across. Magnetars have a magnetic field that is about 100 times stronger than the typical neutron star.
The discovery, to be published in the May 21 issue of the journal Nature, was made by a team of astronomers at the Marshall Space Flight Center led by Dr. Chryssa Kouveliotou of the Universities Space Research Association, working with Dr. Stefan Dieters of the University of Alabama in Huntsville (UAH), Professor Jan van Paradijs of UAH and the University of Amsterdam, and Dr. Tod Strohmayer of NASA's Goddard Space Flight Center in Greenbelt, MD.
"This finding should help us better calculate the rate at which stars die and create the heavier elements that later become planets and other stars," Kouveliotou said.
Kouveliotou and her team determined the strength of the magnetic field by combining data gathered by NASA's Rossi X-Ray Timing Explorer satellite with data from the Advanced Satellite for Cosmology and Astrophysics, a collaborative mission between Japan and the United States.
"The magnetic field generated by this star is truly incredible," Kouveliotou said. "It is so intense that it heats the surface to 18 million degrees Fahrenheit. Periodically, the field drifts through the crust of the neutron star, exerting such colossal forces that it causes a 'starquake.' The 'starquake' energy is then released as an intense burst of low-energy gamma rays."
Since these bursts happen quite often and the bulk of their energy is in low-energy (soft) gamma rays, the objects associated with them had been named Soft Gamma Repeaters. When bursting, Soft Gamma Repeaters are among the brightest objects in the sky, giving off as much energy in a single second as the Sun does in an entire year. The magnetar in question, called SGR 1806-20 by astronomers, was first discovered when it emitted soft gamma ray bursts.
Astronomers have debated the origin of Soft Gamma Repeaters since they were first observed in 1979. With this discovery, however, researchers believe the origin of Soft Gamma Repeaters lies in the 'starquake' phenomena of magnetars. The magnetar theory was first proposed in 1992 by astrophysicists Dr. Robert Duncan of the University of Texas at Austin and Dr. Christopher Thompson of the University of North Carolina at Chapel Hill.
Astronomers believe that at least 10 percent of neutron stars are born with magnetic fields that are strong enough to be considered magnetars. Neutron stars are created in supernovae explosions and they spin rapidly, at rates up to hundreds of revolutions per second.
The magnetar SGR 1806-20 is observed to be spinning once every 7.5 seconds and is slowing down roughly three milliseconds per year. Superstrong magnetic fields cause a neutron star to 'brake' and 'cool down,' making it practically impossible to observe them in radio waves or X-rays. This means there could be thousands or even millions of these dark relics scattered throughout our Milky Way galaxy. This could account for the large number of observed supernovae remnants without detectable neutron stars at their centers.
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"Discovery Launch To Mir on Mission STS-91 Set For June 2 (May 20, 1998)",678,0,0,0
Space Shuttle managers selected June 2 as the official date for the launch of Shuttle Discovery on the ninth planned docking mission with the Russian Space Station Mir.
The flight, designated STS-91, will deliver logistics and supplies to Mir and bring home NASA Astronaut Andrew Thomas, the seventh and final NASA astronaut to serve as a Mir crew member. Thomas has been on the orbiting station since late January.
Discovery will launch from Kennedy Space Center Launch Complex 39A. The current launch time of 6:10 p.m. EDT may vary slightly because of calculations of Mir's precise location in space at the time of lift-off due to Shuttle rendezvous phasing requirements. The STS-91 mission is scheduled to last 9 days, 19 hours, 53 minutes. An on-time launch and nominal mission duration would have Discovery landing back at Kennedy on June 12 at 2:03 p.m. EDT
"The nine joint Shuttle-Mir docking missions and the seven astronauts who served as station crew members have provided us with a wealth of insight and experience to be used as we begin construction of the International Space Station later this year," said George Abbey, Director of the Johnson Space Center, who chaired the review.
The launch team is evaluating a minor overboard water leak from the fuel cell No. 3 relief valve to determine its acceptability for flight. The leak was first seen during Monday's super lightweight tank test when the fuel cell was brought on line to support tanking test operations.
The launch date decision follows completion of the Flight Readiness Review at Kennedy by Shuttle managers from NASA and prime contractor United Space Alliance. STS-91 will be Discovery's 24th mission into space and the 91st Space Shuttle flight in the program's history.
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"Scientists Report TRMM Data Exceeding Expectations (May 19, 1998)",679,0,0,0
The world's first spaceborne rain radar -- aboard the Tropical Rainfall Measuring Mission (TRMM), a joint U.S.-Japanese mission -- is exceeding expectations for accuracy and resolution, and the spacecraft is providing unprecedented insights into rainfall producing cloud systems over tropical land masses and oceans.
"We're extremely excited about these new images and the quality and quantity of the data we're receiving. In several instances, the data resolution is much better than we had anticipated," said Dr. Christian Kummerow, TRMM Project Scientist, at NASA's Goddard Space Flight Center, Greenbelt, MD. "Previously, it was not possible to gather radar precipitation data over the oceans and TRMM has changed all that."
TRMM is NASA's first mission dedicated to observing and understanding tropical rainfall, which comprises more than two-thirds of all rainfall, and how it affects the global climate.
Global rainfall is the primary distributor of heat through atmospheric circulation. The recent El Nino serves as a perfect example of the atmospheric circulation changes that can result from a displacement of the normal precipitation patterns in the central Pacific. More precise information about this rainfall and its variability is crucial to understanding and predicting global climate and climate change.
The Precipitation Radar aboard TRMM is the first rain radar ever launched into space. It measures precipitation distribution over both land and sea areas. Some of the most dramatic Precipitation Radar data was received on March 9 over Melbourne, FL, during the passage of a line of very severe thunderstorms. In comparing the TRMM radar data of the storm with that taken by ground-based radars, the three dimensional TRMM radar showed better vertical resolution of the storm structure. The vertical structure is critical for determining a storm's overall intensity as well as determining the height at which the heat release associated with precipitation is occurring.
Another image released today shows TRMM's radar-derived view of a severe thunderstorm over Houston, TX. The TRMM radar demonstrated significantly better capability to define ambiguities, or occasional "false readings," associated with ground-based radars.
The TRMM spacecraft fills an enormous void in the ability to calculate world-wide precipitation because so little of the planet is covered by ground-based radars. Presently, only two percent of the area covered by TRMM is covered by ground-based radars.
"Since rainfall represents energy conversion, hurricane researchers are eager to use the rainfall data as input to hurricane forecast models," notes Jerry Jarrell, director, National Hurricane Center.
Also aboard TRMM is the Microwave Imager, providing exceptional resolution of storm systems. TRMM's Microwave Imager has better spatial resolution and a new lower frequency channel than previous instruments, according to Kummerow.
An interesting preliminary finding from the Lightning Imaging Sensor (LIS), another instrument on the TRMM satellite, is that its data indicate little lightning over the oceans and 90 percent of lightning occurring over land. Researchers believe that the greater lightning activity over land is primarily due to a larger convection -- or heat -- effect associated with land. This results in greater ice production and, consequently, more lightning. "The beauty of TRMM is that with the Precipitation Radar and the microwave imager, we can test this hypothesis time and again," said LIS Principal Investigator Hugh Christian, at the Global Hydrology and Climate Center at the Marshall Space Flight Center, Huntsville, AL. "TRMM will enable us to gain fundamental insights into the properties of these convective storms and thus better estimate the effects on global weather patterns."
The Clouds and the Earth's Radiant Energy System (CERES) instrument aboard the spacecraft measures how much sunlight the planet's atmosphere, surface and clouds reflect and how much energy it radiates to space from its store of heat energy. "CERES achieved new levels of calibration that we've never reached before in looking at the Earth," said Dr. Bruce Barkstrom, a scientist at NASA Langley Research Center, Hampton, VA, which manages CERES. "Those new levels will help us reduce the uncertainty of how the Earth uses the energy from the Sun to drive the climate system."
By studying rainfall regionally and globally, and the difference in ocean and land-based storms, TRMM is providing scientists the most detailed information to date on the processes of these powerful storms, leading to new insights on how they affect global climate patterns. TRMM's complement of state-of-the-art instruments will provide extremely accurate measurements of the distribution and variability of tropical rain and lightning, and the balance of solar radiation absorbed and reflected by Earth's atmosphere.
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"Hubble Provides Views of How To Feed A Black Hole (May 14, 1998)",680,0,0,0
Astronomers have obtained an unprecedented look at the nearest example of galactic cannibalism -- a massive black hole hidden at the center of a nearby giant galaxy that is feeding on a smaller galaxy in a spectacular collision. Such fireworks were common in the early universe, as galaxies formed and evolved, but are rare today.
Although the cause-and-effect relationships are not yet clear, the views provided by complementary images from two instruments aboard NASA's Hubble Space Telescope are giving astronomers new insights into the powerful forces being exerted in this complex maelstrom. Researchers believe these forces may even have shifted the axis of the massive black hole from its expected orientation.
The Hubble wide-field camera visible image of the merged Centaurus A galaxy, also called NGC 5128, shows in sharp clarity a dramatic dark lane of dust girdling the galaxy. Blue clusters of newborn stars are clearly resolved, and silhouettes of dust filaments are interspersed with blazing orange-glowing gas. Located only 10 million light-years away, this peculiar-looking galaxy contains the closest active galactic nucleus to Earth and has long been considered an example of an elliptical galaxy disrupted by a recent collision with a smaller companion spiral galaxy.
Using the infrared vision of Hubble, astronomers have penetrated this wall of dust for the first time to see a twisted disk of hot gas swept up in the black hole's gravitational whirlpool. The suspected black hole is so dense it contains the mass of perhaps a billion stars, compacted into a small region of space not much larger than our Solar System.
Resolving features as small as seven light-years across, Hubble has shown astronomers that the hot gas disk is tilted in a different direction from the black hole's axis -- like a wobbly wheel around an axle. The black hole's axis is identified by the orientation of a high-speed jet of material, glowing in X-rays and radio frequencies, blasted from the black hole at 1/100th the speed of light.
This gas disk presumably fueling the black hole may have formed so recently it is not yet aligned to the black hole's spin axis, or it may simply be influenced more by the galaxy's gravitational tug than by the black hole's.
"This black hole is doing its own thing. Aside from receiving fresh fuel from a devoured galaxy, it may be oblivious to the rest of the galaxy and the collision," said Ethan Schreier of the Space Telescope Science Institute, Baltimore, MD. Schreier and an international team of co-investigators used Hubble's Near Infrared Camera and Multi-Object Spectrometer to probe deeper into the galaxy's mysterious heart than anyone has before.
The hot gas disk viewed by Hubble investigators is perpendicular to the galaxy's outer dust belt, while the black hole's own internal accretion disk of superhot gas falling into it is tilted approximately diagonally to these axes.
"We have found a complicated situation of a disk within a disk within a disk, all pointing in different directions," Schreier said. It is not clear if the black hole was always present in the host galaxy or belonged to the spiral galaxy that fell into the core, or if it is the product of the merger of a pair of smaller black holes that lived in the two once-separate galaxies.
Having an active galaxy just 10 million light-years away from Earth rather than hundreds of millions or billions of light-years distant offers astronomers a unique laboratory for understanding the elusive details of the behavior of supermassive black holes as fueled by galaxy collisions.
"Though Hubble has seen hot gas disks around black holes in other galaxies, the infrared camera has for the first time allowed us to peer at this relatively nearby, very active, but obscured black hole region," Schreier added.
The team of astronomers is awaiting further Hubble data to continue its study of the disk, as well as ground-based spectroscopic observations to measure the velocity of entrapped material around the black hole. This will allow the astronomers to better calculate the black hole's mass. The current results are scheduled to appear in the June 1, 1998 issue of Astrophysical Journal Letters.
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"Most Powerful Explosion Since the Big Bang Challenges Gamma Ray Burst Theories (6 May, 1998)",681,0,0,0
A recently detected cosmic gamma ray burst released a hundred times more energy than previously theorized, making it the most powerful explosion since the creation of the universe in the Big Bang.
"For about one or two seconds, this burst was as luminous as all the rest of the entire universe," said Caltech professor George Djorgovski, one of the two principal investigators on the team from the California Institute of Technology, Pasadena, CA.
The team measured the distance to a faint galaxy from which the burst originated at about 12 billion light years from the Earth. The observed brightness of the burst despite this great distance implies an enormous energy release. The team's findings appear in the May 7 issue of the journal Nature.
The burst was detected on Dec. 14, 1997, by the Italian/Dutch BeppoSAX satellite and NASA's Compton Gamma Ray Observatory satellite. The Compton observatory provided detailed measurements of the total brightness of the burst, designated GRB 971214, while BeppoSAX provided its precise location, enabling follow-up observations with ground-based telescopes and NASA's Hubble Space Telescope.
"The energy released by this burst in its first few seconds staggers the imagination," said Caltech professor Shrinivas Kulkarni, the other principal investigator on the team.
The burst appears to have released several hundred times more energy than an exploding star, called a supernova, until now the most energetic known phenomenon in the universe. Finding such a large energy release over such a brief period of time is unprecedented in astronomy, except for the Big Bang itself.
"In a region about a hundred miles across, the burst created conditions like those in the early universe, about one millisecond (1/1,000 of a second) after the Big Bang," said Djorgovski.
This large amount of energy was a surprise to astronomers. "Most of the theoretical models proposed to explain these bursts cannot explain this much energy," said Kulkarni. "However, there are recent models, involving rotating black holes, which can work. On the other hand, this is such an extreme phenomenon that it is possible we are dealing with something completely unanticipated and even more exotic."
Gamma-ray bursts are mysterious flashes of high-energy radiation that appear from random directions in space and typically last a few seconds. They were first discovered by U.S. Air Force Vela satellites in the 1960s. Since then, numerous theories of their origin have been proposed, but the causes of gamma-ray bursts remain unknown. The Compton observatory has detected several thousand bursts so far.
The principal limitation in understanding the bursts was the difficulty in pinpointing their direction on the sky. Unlike visible light, gamma rays are exceedingly difficult to observe with a telescope, and the bursts' short duration exacerbates the problem. With BeppoSAX, scientists now have a tool to localize the bursts on the celestial sphere with sufficient precision to permit follow-up observations with the world's most powerful ground-based telescopes.
This breakthrough led to the discovery of long-lived "afterglows" of bursts in X-rays, visible and infrared light, and radio waves. While gamma-ray bursts last only a few seconds, their afterglows can be studied for several months. Study of the afterglows indicated that the bursts do not originate within our own galaxy, the Milky Way, but rather are associated with extremely distant galaxies.
Both BeppoSAX and NASA's Rossi X-ray Timing Explorer spacecraft detected an X-ray afterglow. BeppoSAX precision led to the detection of a visible light afterglow, found by a team from Columbia University, New York, NY, and Dartmouth College, Hanover, NH, including Professors Jules Halpern, David Helfand, John Torstensen, and their collaborators, using a 2.4-meter telescope at Kitt Peak, AZ, but no distance could be measured from these observations.
As the visible light from the burst afterglow faded, the Caltech team detected an extremely faint galaxy at its location, using one of the world's largest telescopes, the 10-meter Keck II telescope at Mauna Kea, Hawaii. The galaxy is about as faint as an ordinary 100 watt light bulb would be as seen from a distance of a million miles.
Subsequent images taken with the Hubble Space Telescope confirmed the association of the burst afterglow with this faint galaxy and provided a more detailed image of the host galaxy.
The Caltech team succeeded in measuring the distance to this galaxy, using the light-gathering power of the Keck II telescope. The galaxy is at a redshift of z=3.4, or about 12 billion light years distant (assuming the universe to be about 14 billion years old).
From the distance and the observed brightness of the burst, astronomers derived the amount of energy released in the flash. Although the burst lasted approximately 50 seconds, the energy released was hundreds of times larger than the energy given out in supernova explosions, and it is about equal to the amount of energy radiated by our entire Galaxy over a period of a couple of centuries. Scientists say it is possible that other forms of radiation from the burst, such as neutrinos or gravity waves, which are extremely difficult to detect, carried a hundred times more energy than that.
NASA is planning two missions to further investigate gamma-ray bursts: the High Energy Transient Experiment II (HETE II), scheduled to launch in the fall of 1999, and the Gamma Ray Large Area Space Telescope (GLAST), scheduled to launch in 2005. HETE II will be able to precisely locate gamma-ray bursts in near real-time and quickly transmit their locations to ground-based observatories, permitting rapid follow-up studies. GLAST will detect only those gamma-ray bursts that emit the highest energy gamma rays, and will be able to locate them with sufficient precision to permit coordinated observations from the ground. Because not much is known about the bursts at these high energies, the observations may permit researchers to choose among competing theories for the origin of gamma-ray bursts.
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"Preflight Briefings For Final Shuttle-Mir Mission Set For May 11 (5 May, 1998)",682,0,0,0
A series of background press briefings on the STS-91 mission, the final Shuttle flight to dock with the Mir space station, will be held on Monday, May 11, at the Johnson Space Center, Houston, TX, beginning at 9 a.m. EDT.
The major objective of the mission is the return of Andy Thomas from four months of research on the Mir as the seventh and final U.S. astronaut to live and work on the Russian complex. Thomas' departure will mark the end of more than two years of a continuous U.S. presence in space.
The briefings will begin with an overview of the STS-91 mission followed at 10 a.m. with a briefing on the Shuttle-Mir Phase One program. After taking a break for the daily NASA Video File at noon, there will be a briefing on the cargo being carried in the Spacehab module at 12:30 p.m., followed by a briefing on the Alpha Magnetic Spectrometer (AMS) payload at 1 p.m. The STS-91 astronauts will hold their preflight press conference beginning at 2:30 p.m. All of the briefings will be carried live on NASA Television.
NASA Television is available through the GE-2 satellite, transponder 9C located at 85 degrees West longitude, vertical polarization, with a frequency of 3880 MHz, and audio at 6.8 MHz.
STS-91 PREFLIGHT BRIEFINGS
Monday, May 11, 1998
(All times shown are EDT)
9 a.m. -- MISSION OVERVIEW
Paul Dye, STS-91 Lead Flight Director
10 a.m. -- PHASE ONE OVERVIEW
Frank Culbertson, Director, Shuttle-Mir Phase One Program
John Uri, Shuttle-Mir Mission Scientist
Noon -- NASA VIDEO FILE
12:30 p.m. -- SPACEHAB BRIEFING
Mike Bain, Shuttle-Mir Program Manager, Spacehab
Carolyn Overmyer, SHUCS experiment
1 p.m. -- ALPHA MAGNETIC SPECTROMETER (AMS) BRIEFING
Mark Sistilli, Program Manager, AMS, NASA
John O'Fallon, Director, High Energy Physics, Dept. of Energy
2:30 p.m. -- STS-91 CREW PRESS CONFERENCE
Charles Precourt, Commander
Dominic Gorie, Pilot
Franklin Chang-Diaz, Mission Specialist-1, Payload Commander
Wendy Lawrence, Mission Specialist-2
Janet Kavandi, Mission Specialist-3
Valeri Ryumin, Mission Specialist-4
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"Astronomers Find Planet Construction Zone Around Nearby Star (21 April, 1998)",683,0,0,0
NASA astronomers using the new Keck II telescope in Hawaii have discovered what appears to be the clearest evidence yet of a budding solar system around a nearby star. Scientists released an image of the probable site of planet formation around a star known as HR 4796, about 220 light-years from Earth in the constellation Centaurus.
The image, taken with a sensitive infrared camera developed at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, shows a swirling disk of dust around the star. Within the disk is a telltale empty region that may have been swept clean when material was pulled into newly formed planetary bodies, the scientists said. "This may be what our solar system looked like at the end of its main planetary formation phase," said Dr. Michael Werner of JPL, who co-discovered the region, along with Drs. David Koerner and Michael Ressler, also of JPL, and Dana Backman of Franklin and Marshall College, Lancaster, PA.
"Comets may be forming right now in the disk's outer portion from remaining debris." The discovery was made on March 16 from the giant 33-foot (10-meter) Keck II telescope atop Mauna Kea, Hawaii. Keck II and its twin, Keck I, are the world's largest optical and infrared telescopes. Attached to the Keck II for this observation was the mid-infrared camera, developed by Ressler at JPL and designed to measure heat radiation. The four scientists reported their discovery in a submission to The Astrophysical Journal Letters. The disk was discovered independently and contemporaneously at the Cerro Tololo Observatory in Chile by another team of scientists, led by Ray Jayawardhana of the Harvard-Smithsonian Center for Astrophysics (CfA), Cambridge, MA, and Dr. Charles Telesco of the University of Florida, Gainesville.
Koerner of JPL said the finding represents a "missing link" in the study of how planetary systems are born and evolve. "In a sense, we've already peeked into the stellar family album and seen baby pictures and middle-aged photos," Koerner said. "With HR 4796, we're seeing a picture of a young adult star starting its own family of planets. This is the link between disks around very young stars and disks around mature stars, many with planets already orbiting them." "This is the first infrared image where an entire inner planetary disk is clearly visible," Werner said.
"The planet- forming disk around the star Beta Pictoris was discovered in 1983 by the Infrared Astronomical Satellite (IRAS), and also later imaged with the Hubble Space Telescope, but glaring light from the star partially obscured its disk." The apparent diameter of the dust disk around HR 4796 is about 200 astronomical units (one astronomical unit is the distance from Earth to the Sun). The diameter of the cleared inner region is about 100 astronomical units, slightly larger than our own solar system. HR 4796 was originally identified as an interesting object for further study by Dr. Michael Jura, an astronomy professor at the University of California, Los Angeles.
The star, HR 4796, is about 10 million years old and is difficult to see in the continental United States, but is visible to telescopes in Hawaii and the southern hemisphere. The discovery of the HR 4796 disk was made in just one hour of observing time at Keck, but the JPL team plans to return to Hawaii in June for further studies. They hope to learn more about the structure, composition and size of this disk, and to determine how disks around stars in our galaxy produce planets. They plan to study several other stars as well, including Vega, which was featured prominently in the movie, "Contact."
The Harvard/Florida research team that also found the HR 4796 disk included Drs. Lee Hartmann and Giovanni Fazio of Harvard- Smithsonian Center for Astrophysics, and Scott Fisher and Dr. Robert Pina of the University of Florida. JPL's use of the Keck telescope is supported by NASA's Origins program, a series of missions to study the formation of galaxies, stars, planets and life, and to search for Earth-like planets around other stars that might have the right conditions for life. The W. M. Keck Observatory is owned and operated by the California Association for Research in Astronomy, a joint venture between the University of California, California Institute of Technology (Caltech), Pasadena, CA, and NASA.
Use of the Keck Observatory for Origins research is managed by JPL for NASA's Office of Space Science, Washington, DC. JPL is a division of Caltech. The research of both teams was supported in large part by the NASA Origins Program, with additional support to the CfA- Florida team from the National Science Foundation, the National Optical Astronomy Observatories, and the Smithsonian Institution; and with additional NASA support for the CalTech/JPL-Franklin & Marshall team, including use of the Keck Observatory.
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"Deep Space 1 Launch Rescheduled To October (17 April, 1998)",684,0,0,0
The planned July 1998 launch of NASA's Deep Space 1 technology validation mission from Cape Canaveral, FL, has been rescheduled for October. The delay is due to a combination of late delivery of the spacecraft's power electronics system and an ambitious flight software development schedule, which together leave insufficient time to test the spacecraft thoroughly for a July launch.
The power electronics system regulates and distributes power produced by not only the solar concentrator array, a pair of experimental solar panels composed of 720 cylindrical Fresnel lenses, but also by an on-board battery. Among many other functions, it helps the solar array to operate at peak efficiency, and ensures that the battery is able to cover temporary surges in power needed so that the ion propulsion system (which needs electricity for its basic operations) receives a steady power supply.
"With a new launch date for this bold mission, we can be more confident that we will be ready to fully exercise our payload of important technologies," explained Chief Mission Engineer Dr. Marc Rayman of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. "The entire DS1 team looks forward to this opportunity to make a significant contribution to science missions of the future through the capabilities we are testing on DS1."
Deep Space 1 is the first launch in NASA's New Millennium program, a series of missions designed to test new technologies so that they can be confidently used on science missions of the 21st century. Among the 12 technologies the mission is designed to validate are ion propulsion, autonomous optical navigation, a solar power concentrator array and an integrated camera and imaging spectrometer.
The earlier July launch period for DS1 allowed it to fly a trajectory encompassing flybys of an asteroid, Mars and a comet. By the end of May, the mission design team is scheduled to finalize new target bodies in the Solar System for DS1 to encounter based on an October launch date. The New Millennium Program and Deep Space 1 are managed by JPL for NASA's Office of Space Science, Washington, DC. JPL is a division of the California Institute of Technology.
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"Space Science Update: 'Planets Under Construction' (16 April, 1998)",685,0,0,0
The next Space Science Update (SSU), set for 11 a.m. EDT, Tuesday, April 21, 1998, at NASA Headquarters, will feature the discovery astronomers are calling the clearest evidence yet of a solar system forming around a nearby star. Two independent teams of astronomers used telescopes in Hawaii and Chile to image the region near the star where it appears planets may be forming, or may recently have formed, a 'missing link' of planetary formation between dust disks and systems with planets already formed.
Panelists will be: * Dr. Michael Werner, NASA Jet Propulsion Laboratory/Caltech, Pasadena, CA. * Dr. David Koerner, JPL/University of Pennsylvania, Philadelphia. * Dr. Michael Jura, University of California, Los Angeles. * Dr. Lee Hartmann, Harvard-Smithsonian Center for Astrophysics, Cambridge, MA. * Dr. Ed Weiler, Director of NASA's Origins Program, Washington, DC, panel moderator. The SSU will originate from NASA Headquarters Auditorium, 300 E St., S.W., Washington, DC, and will be carried live on NASA TV with two-way question-and-answer capability for reporters covering the event from participating NASA centers.
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"New Sunspot Cycle To Be Bigger Than Average (13 April, 1998)",686,0,0,0
The current sunspot cycle will be above average but no record setter, according to scientists at NASA's Marshall Space Flight Center. "It's like saying we're going to have a mild or cold winter," said Dr. David Hathaway. "We're in a similar state in predicting what the sun's climate is going to do."
Hathaway is a co-author with Robert M. Wilson and Edwin J. Reichmann, also at NASA/Marshall, of "An Estimate for the Size of Cycle 23 Based on Near Minimum Conditions." It will appear in the May 1998 issue of the \IJournal of Geophysical Research (Space Physics)\i.
The sun now is on the upswing of its 23rd activity cycle, a numbering scheme that dates from the mid-19th century, following introduction of the "relative sunspot number" by Rudolf Wolf of the Zurich Observatory in 1848. Wolf's sunspot number (now called the International sunspot number or the Zurich number) represents a blend of actual numbers of individual spots and numbers of groups of spots on the sun.
On average, this number varies from a minimum through a maximum to the next minimum in about 11 years. Because the solar magnetic fields reverse at the peak of each 11-year cycle, solar activity cycle actually spans a 22-year "Hale cycle." Cycle 23 is the last half of the current Hale cycle (composed of Cycles 22 and 23) that began in 1986.
"The consensus [among solar physicists] is that this cycle will be above average in size and probably a fast riser," Wilson said. "Sunspot maximum should not be perceived as the top of the cycle curve, but instead it should be thought of as an interval of peak activity which usually spans about 2 to 4 years and includes the actual maximum in sunspot number." For Cycle 23, the peak interval starts in 1999.
Predicting the solar cycle is more than a matter of scientific curiosity. An active sun can cause geomagnetic storms that endanger satellites and disrupt communications and power systems on Earth. It also heats the Earth's outer atmosphere so that spacecraft are exposed to more atmospheric drag and to greater erosion by atomic oxygen.
We even have tantalizing hints that the Earth's climate may be linked to sunspots. The "Little Ice Age" corresponded with a 70-year period, 1645-1715, when sunspots were sparse in number, the Maunder minimum. Also, there are strong statistical associations linking current trends in climate (surface temperatures) to trends in solar activity, as outlined in another paper by Wilson for the \IJournal of Geophysical Research (Atmospheres)\i.
Still, with almost 250 years of observations - of which only the last 150 years are considered truly reliable- predictions are akin to the Farmer's Almanac, Hathaway said.
"There's no real physics involved," he explained. "It's all statistical inferences."
The reason is that while scientists believe that sunspots are driven by the dynamo that lies hidden beneath the photosphere, they are unsure of what controls the dynamo.
So what they do is study measurements of several activities related to the sun and look for patterns that proceed the sunspot cycle. Hathaway and Wilson said these include variations in geomagnetic indices (like the aa and Ap indices, which respond to the changing conditions in the solar wind), the occurrences of high-latitude spots, the inferred strengths of the sun's polar fields, and the number of geomagnetically disturbed days (those days when the Ap index exceeds 25) over the course of the preceding cycle.
A part of the \Iaa\i (or \Iap\i) index is directly proportional to sunspot number. This component varies directly with the current solar cycle and is related to sporadic solar events like flares and disappearing filaments and prominences. On the other hand, "the remaining portion seems to be associated with the next solar cycle" Hathaway said, "although there's no conclusion on what produces this connection." This latter component seems to be related to recurrent features on the sun.
"It's probably related to the polar fields," Wilson said. "Polar coronal holes tend to reach maximum size near cycle minimum, and the maximum in the geomagnetic indices usually occurs within a couple of years or so before minimum."
In turn, this serves as a good indicator of where the sunspot cycle is headed, but it's not an absolute measure. Hathaway has several graphs showing different prediction methods applied to past sunspot cycles. That is, they used the data from early in a cycle to "predict" what the rest of the cycle would do. Most predictions match reality fairly close, except for Cycle 19, the current record cycle, which peaked at 190 in sunspot number in 1957, just as the first space physics satellites were being launched.
Based on the various precursor techniques, Hathaway, Wilson, and Reichmann predict that Cycle 23 will rise faster than normal to its peak, attaining maximum amplitude sometime during the latter half of 1999 to the first half of 2000, and that it will measure about 170 plus or minus 20 units (yearly sunspot number). They expect Cycle 23 to continue until sometime in 2006 when the next cycle, Cycle 24, should begin.
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"President Clinton To Visit Johnson Space Center (13 April, 1998)",687,0,0,0
President Bill Clinton will visit NASA's Johnson Space Center on Tuesday, April 14 for briefings on various NASA programs. During the visit, the President will meet with Senator John Glenn, now in training for an October space flight, and get a briefing on Shuttle mission STS-90, scheduled for launch on Thursday. The President will place a phone call to the STS-90 crew members at the Florida launch site, and address Johnson Space Center employees and invited guests.
The President's call to the STS-90 crew and his address to employees will be carried live on NASA Television beginning at approximately 12:30 p.m. EDT. NASA Administrator Daniel S. Goldin and Johnson Director George W.S. Abbey will escort the President during his tour.
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"Science Team Chosen for Mission to Explore the Subsurface of Mars (8 April, 1998)",688,0,0,0
Nine researchers have been selected to be the Science Team for the Mars Microprobes, a technology validation mission that will hitchhike to the red planet aboard NASA's 1998 Mars Polar Lander mission. Two identical probes will be carried as a secondary payload on the lander, due for launch in January 1999. Following an 11-month cruise, the Microprobes will separate from the lander before it enters the Martian atmosphere, and then hit the ground at approximately 400 mph.
During the impact, each microprobe will separate into two sections: the forebody and its instruments will penetrate up to six feet (two meters) below the surface, while the aftbody will remain near the surface to communicate with a radio relay on NASA's Mars Global Surveyor orbiter while making meteorological measurements. The nine selected scientists are: * David Catling, NASA Ames Research Center, Moffett Field, CA * Ralph Lorenz, University of Arizona, Tucson * Julio Magalhaes, NASA Ames Research Center * Jeffrey Moersch, NASA Ames Research Center * Paul Morgan, Northern Arizona Univ., Flagstaff * James Murphy, NASA Ames Research Center * Bruce Murray, California Institute of Technology, Pasadena * Marsha Presley, Arizona State Univ., Phoenix * Aaron Zent, NASA Ames Research Center.
The scientific objectives of the Mars Microprobes include searching for the presence of water ice in the soil and characterizing its thermal and physical properties. A small drill will bring a soil sample inside the probe, heat it, and look for the presence of water vapor using a tunable diode laser. An impact accelerometer will measure the rate at which the probes come to rest, giving an indication of the hardness of the soil and any layers present. Temperature sensors will estimate how well the Martian soil conducts heat, a property sensitive to different soil properties such as grain size and water content.
A sensor at the surface will measure atmospheric pressure in tandem with a sensor on the Mars Polar Lander. The Mars Microprobes mission, also known as Deep Space-2 (DS- 2), is scheduled to be the second launch in NASA's New Millennium Program of technology validation flights, designed to enable advanced science missions in the 21st century. "I'm delighted with the selection of this excellent group of investigators. The Mars Microprobe will give us a glimpse of the subsurface of Mars, which in many ways is a window into the planet's history," said Dr. Suzanne Smrekar, the DS-2 project scientist at NASA's Jet Propulsion Laboratory, Pasadena, CA.
"The region of Mars we will explore is similar to Earth's polar regions in that it is believed to collect ice and dust over many millions of years. By studying the history of Mars and its climate, we are likely to better understand the more complex system on our own planet." In addition to the miniaturized science instruments capable of surviving high velocity impact, technologies to be tested on DS-2 include a non-erosive, lightweight, single-stage atmospheric entry system or aeroshell; power microelectronics with mixed digital/analog advanced integrated circuits; an ultra-low temperature lithium battery; an advanced three-dimensional microcontroller; and flexible interconnects for system cabling.
"The combination of a single-stage entry vehicle with electronics and instrumentation that can survive very high impact loads will enable us to design a whole new class of very small, rugged spacecraft for the in-situ exploration of the planets," explained Sarah Gavit, DS-2 project manager at JPL. "Slamming high-precision science instruments into the surface of Mars at 400 mph is very challenging, no doubt about it! But once this type of technology is demonstrated, we can envision future missions that could sample numerous regions on Mars or make network measurements of global weather and possible Marsquakes," said DS-2 program scientist Dr. Michael Meyer of NASA Headquarters, Washington, DC.
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"Increasing Greenhouse Gases May Be Worsening Arctic Ozone Depletion (8 April, 1998)",689,0,0,0
In late 1997, larger levels of ozone depletion were observed over the Arctic than in any previous year on record. Now, using climate models, a team of scientists reports why this may be related to greenhouse gases, according to a paper published in the April 9 issue of Nature. The study suggests the increase in greenhouse gas emissions is one possible cause of the observed trends in Arctic ozone losses and that this may delay recovery of the ozone layer.
The research team, consisting of researchers from NASA's Goddard Institute for Space Studies (GISS) and Columbia University, New York, investigated the response of ozone to projected future emissions of greenhouse gases and ozone-depleting halogens over time, using the GISS climate model. This is the first time ever that the interaction between ozone chemistry and the gradual buildup of greenhouse gases has been studied in a climate model.
"Buildup of greenhouse gases leads to global warming at the Earth's surface, but cools the stratosphere. Since ozone chemistry is very sensitive to temperature, this cooling results in more ozone depletion in the polar regions," said Dr. Drew Shindell of Columbia University, the lead author of the study. NASA will continue research in this area to determine if these models are accurate. The "greenhouse effect" is defined as the warming of climate that results when the atmosphere traps heat radiating from Earth toward space.
Certain gases in the atmosphere -- such as water vapor, carbon dioxide, nitrous oxides and chlorofluorocarbons -- act like glass in a greenhouse, allowing sunlight to pass into the "greenhouse," but blocking Earth's heat from escaping into space. Ozone, a molecule made up of three atoms of oxygen, comprises a thin layer of the upper atmosphere which absorbs harmful ultraviolet radiation from the Sun and protects people, animals and plants from too much ultraviolet sunlight. Distribution and concentration of stratospheric ozone are influenced in two ways by human-driven activity in addition to natural, seasonal variations.
Of first importance is the direct impact of industrially produced chlorofluorocarbons. Although ozone levels around the globe are expected to continue to decline over the next several years, NASA is now detecting decreasing growth rates of ozone-depleting compounds in the upper part of the atmosphere, indicating that international treaties to protect the ozone layer are working. The second influence on stratospheric ozone levels is the indirect impact of "greenhouse gases" on atmospheric temperatures. Ozone destruction is quite sensitive to temperature increases in the atmosphere.
Since upper atmospheric temperatures in the Northern Hemisphere during winter and spring generally are warmer than those in the Southern Hemisphere, ozone depletion over the Arctic has been much smaller than over the Antarctic during the 1980s and early 1990s. The Arctic stratosphere, however, gradually has cooled over the past few decades resulting in very large ozone depletion, especially during 1996-97. In the simulations performed by Shindell and his team, temperature and wind changes, induced by increasing greenhouse gases, clearly alter the dynamics of the atmosphere. According to this model, as the abundance of greenhouse gases gradually increases, the frequency of Northern Hemisphere sudden stratospheric warming is reduced, leading to significantly colder lower stratospheric temperatures.
If proven correct, this dynamic effect would add to the greenhouse cooling of the stratosphere. "Results suggest that the combination of these two cooling effects causes dramatically increased ozone depletion so that ozone loss in the Arctic by the year 2020 roughly is double what it would be without greenhouse gas increases," said Dr. David Rind of GISS, a co-author of the paper. Increasing greenhouse gases therefore may be at least partially responsible for the very large Arctic ozone losses in recent winters.
The authors caution, however, that though the model predicts a general trend towards increasing ozone depletion, the year-to-year variability is quite large, especially in the Arctic. For example, several years in the late 1990s and early 2000s show very little Arctic ozone depletion, while others show record losses. In fact, the 1997-98 winter that just occurred was characterized by significantly less ozone loss than the preceding six winters. A factor that should be considered, however, is the consistency in model predictions, i.e. whether the same results can be reproduced by other models.
According to this model, the severity and duration of the Antarctic ozone depletion also may increase due to greenhouse gas-induced stratospheric cooling over the coming decades. However, ozone in the Antarctic is already so depleted that any additional losses may be relatively small, Rind added. The research was conducted by scientists at GISS, The Center for Climate Systems Research, Columbia University, and Science Systems and Applications Inc., New York. The GISS research is part of NASA's Earth Science Enterprise, a long- term coordinated research effort to study the Earth as a global system.
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"New Lab Studies Death of Stars, Origin of Planets (6 April, 1998)",690,0,0,0
Almost a century ago, American physicist Robert Millikan watched an oil droplet rise and fall as he gently applied an electrical charge. He was pursuing a fundamental measurement of nature, the size of the electron-volt. In the process, he simulated part of what happens when dust condenses into planets.
Now scientists at NASA's Marshall Space Flight Center are running a sophisticated variation of Millikan's experiment as they try to understand dust behavior in space.
"The main purpose is to study the microphysics of a charged particle when it's exposed to a plasma [electrified gas]," said Dr. James Spann of the Space Sciences Laboratory at NASA/Marshall.
Spann is principal investigator for the new Dusty Plasma Laboratory which he and graduate student Catherine Venturini have been developing over the last two years. Today and Wednesday (April 6, 8) they present their initial work to the Seventh Workshop on the Physics of Dusty Plasmas at the University of Colorado in Boulder.
"I call this the three-dimensional Millikan experiment," Spann said. Millikan's original experiment - which won him the Nobel Prize in physics in 1923 - worked in two dimensions as he watched an oil droplet rise and fall as X-rays ionized it.
Space is not a pristine vacuum. It is peppered with cosmic rays, atoms, molecules, and asteroids, and so on up to planets and stars. And dust. Lots of dust. Indeed, dust is the stuff from which planets are made. A dying star disgorges heavy atoms that eventually form dust grains that, in turn, coalesce into ever-larger debris until planets are formed around new stars.
What has not been given much study is the fact that dust is readily pushed around by sunlight and by stellar winds, and easily electrified by exposure to ultraviolet light or even friction with other dust particles. When electrified, they will repel each other, like your hair standing on end after brushing on a cold, dry day. That repulsion can alter how dust grains gather when the cloud is so loose that has a weak gravity field.
Scientists have run experiments on quantities of dust, but Spann and Venturini are making the first systematic investigation on the behavior individual particles. "We're approaching it from a fundamental perspective," Spann explained. "Here's one particle. How does it change under various conditions, as the electrical charge changes, with different masses, with different shapes and materials?"
The answers should have many uses.
"Applications range from atmospheric aerosols [droplets and dust in the air] to stellar formation," Spann said. "Dust is everywhere, in planetary atmospheres, the zodiacal light [sunlight reflected back from deep space], in planetary origins."
Spann and Venturini simulate dust in space with a table top apparatus designed to isolate a single bit of dust in a vacuum and manipulate it with an electrical charge. Spann started developing the Dusty Plasma Laboratory started in 1995 under the Center Director Discretionary Fund. Venturini, then a senior in physics at Loyola Marymount University in Los Angeles, Calif., was recruited through the NASA Academy program.
"On my first day, Jim said, 'OK, we have to put this thing together'," Venturini said. She continues her work in the Dusty Plasma Lab as she finishes her master's degree in physics at UAH. Her current work is funded by the Alabama Space Grant Consortium
"It's not a complicated set-up," Spann said. "But it has a lot of parts that have to work together. We have a lot of things that are normally used in laboratories, but not together."
The system comprises several tubes of stainless steel welded together so they intersect at the center of the table. Atop the intersection is a small, windowed chamber which, in turn, is crowned by the dust injector and an electron gun.
Venturini explained that most of the tubing is just a manifold connected to vacuum pumps to evacuate the chamber after the dust is injected. All of the action takes place in the windowed chamber, which encloses a quadrupole trap: two hemispherical electrodes at top and bottom and a ring electrode at the middle (Picture 1). These form hilltops surrounding an unusual valley. When the electrodes are charged, the center becomes the point of zero potential. An electrified dust mote will sort of "fall downhill," repelled on all sides by the electrodes.
Then they watch the behavior of the dust particle through a TV camera coupled to a long-distance microscope (at left is a typical image). A helium neon laser selected because its red light will not impart an electrical charge to the dust illuminates. Blue light can do so; ultraviolet light almost certainly will. While that is an important factor for dust behavior in space, it would be a complication for experiments where Spann and Venturini are trying to deduce the basics of dusty plasma behavior.
For now, they use table salt, or store it. Household dust won't do since that usually comes from clothing, skin, animal fur, dead insects, and other items not found in space.
Sodium chloride is dissolved in pure water, and a tiny droplet is squeezed through a syringe into the test chamber, which still contains air. The quadrupole trap captures the droplet and suspends it in midair.
The water evaporates, leaving just a tiny crystal of salt about 3 to 5 microns wide (that's less than a 5/10,000ths of an inch). Next, the air is evacuated ever so gently so no breeze dislodges the dust grain. This takes about an hour, and uses a sapphire-coated needle valve to allow the tiniest wisps of air to escape.
Once conditions are right, the grain is bombarded with electrons and the grain's behavior is measured. Spann said the grain would charge until it either repels new electrons, or it accepts them while shedding other electrons like newcomers bumping people out of a crowded room.
Where Millikan tweaked his experiment at just a few volts, Spann and Venturini apply voltages of 100 to 1,000 electron-volts. It's tougher than it sounds. "The difficult part of an experiment is to make it simple," Spann said. He and Venturini are still controlling the apparatus by hand, and Venturini is writing a computer program to measure the grain's motions both to take scientific data and to nudge the grain back into position.
"What we're trying to establish in the lab is an area of expertise that could be used by experimental and modeling groups," Spann said.
More complex experiments could examine the behavior of dust grains when they absorb then re-emit infrared radiation. This could help scientists understand what they are seeing in images of so-called planetary nebula, immense dust clouds surrounding young or dying stars.
It can also help in understanding how the view is obscured in other wavelengths.
"The mid-infrared background that telescopes have to contend with is dominated by interplanetary and interstellar dust particles," Spann said, "so understanding that is very important." For example, telescopes looking into space directly opposite the sun see a dull, glowing patch of space. This is the gegenschein (also called the zodiacal light), sunlight scattered back by dust.
An effect more familiar to most people is the beautiful tail a comet forms as it approaches the sun. The tail comprises dust, ice, and gas energized by sunlight and the solar wind. The Dusty Plasma Laboratory will help in understanding the mechanisms that make the tail spread as it does.
Once they have the system mastered, they will experiment with a range of different dust motes that conduct, store, or insulate against electricity. Then they will be ready to start investigating the cosmos on a microscale.
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"Neurolab Shuttle Mission To Launch April 16 (3 April, 1998)",691,0,0,0
Space Shuttle program managers today affirmed April 16 as the launch date for NASA's second Shuttle mission of 1998 -- a two week life sciences research flight that will focus on the most complex and least understood part of the human body, the nervous system. The Flight Readiness Review held at NASA's Kennedy Space Center, FL, yesterday is the final major review by all Shuttle project offices to evaluate the readiness of the flight crew and vehicle, along with launch and mission control flight teams, to support the launch of Space Shuttle Columbia on the STS-90 Neurolab mission.
Columbia is scheduled for launch on April 16, 1998 from NASA's Kennedy Space Center Launch Complex 39-B. The 2 1/2 hour available launch window opens at 2:19 p.m. EDT. The STS-90 mission is scheduled to last 15 days, 21 hours, 50 minutes. However, mission managers are reserving an option of extending the flight one additional day for science operations if Shuttle electrical power margins permit.
A launch on April 16, and a 16- or 17- day nominal mission would have Columbia landing at Kennedy on May 2 or 3. The STS-90 Mission Commander is Richard A. Searfoss. Pilot for the flight is Scott D. Altman. There are three mission specialists assigned to this mission -- Richard M. Linnehan, who is also serving as the Payload Commander; Kathryn P. (Kay) Hire; and Dafydd (Dave) Rhys Williams from the Canadian Space Agency. Two payload specialists -- Jay Clark Buckey, Jr., and James A. (Jim) Pawelczyk -- round out the seven member STS-90 crew.
STS-90 will be the 25th flight of Columbia and the 90th mission flown since the start of the Space Shuttle program in April 1981.
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"Solar Images To Be Made By Unique X-ray Telescopes (2 April, 1998)",692,0,0,0
A unique cluster of telescopes that make X-rays take a U-turn has been selected
for a fourth flight to capture "multicolored" images that will help us understand why the sun's outer atmosphere is so hot.
"One of the major objectives is to follow up on something we saw on the first flight 10 years ago," said Dr. Arthur B.C. Walker II of Stanford University, the principal investigator for the Chromospheric/Corona Spectroheliograph telescope. It will actually be a bundle of up to 19 telescopes, each taking pictures of the sun in a slightly different X-ray energy. The array is an upgrade of the Multi-Spectral Solar Telescope Array (MSSTA) which flew on October 23, 1987, May 13, 1991, and November 3, 1994.
The 1987 flight - which also made the September 30, 1988 cover of \IScience\i magazine - returned pictures that showed where the sun's atmosphere was as hot as 1 million deg. K (about 1.8 million deg. F) also showed spectral lines that indicated temperatures of about 700,000 deg. K (1.26 million deg F).
"We were mystified by this," Walker said. "We are now convinced that there is material at about 700,000 degrees K in the transition region and which contributes to coronal heating."
NASA recently selected the Chromospheric/Corona Spectroheliograph under the solar physics research program. Richard Hoover of NASA's Marshall Space Flight Center and Troy W. Barbee of Lawrence Livermore National Laboratory are co-investigators with Walker. Their project is entitled Investigation of the Corona/Chromosphere Interface.
This is the same region that will be studied by the Transition Region and Coronal Explorer (TRACE) scheduled for launch Thursday evening from California. The Chromospheric/Corona Spectroheliograph will complement TRACE by providing images of solar gases at temperatures as high as 5 million degrees K (9 million deg. F).
While the sun is more than 99.9 percent hydrogen and helium, it carries significant quantities of carbon, iron, calcium, silicon, and other elements. Heavier elements have more protons (carbon is 6, iron is 26) in their nuclei than do lighter elements (hydrogen is 1, helium is 2). That means that as electrons are stripped from heavier atoms, the charge of the larger number of protons is devoted to the few remaining electrons. It takes ever more energy to strip off another electron.
As a result, light from energetic atoms acts like a tracer that reveals where the sun is hot and at what temperatures. This is important to dissecting activities from the sun's corona - its outer atmosphere - through the transition region and to the chromosphere and photosphere - the visible "surface."
The challenge is that the X-ray emissions are so energetic that they pass through materials rather than being reflected as visible light would be. The usual trick to making X-ray images is called grazing incidence reflection. Just as light will reflect off clear glass (or a rock will skip on a pond) if it strikes at a shallow angle, X-rays will reflect - and be focused - if they, too, strike at an even shallower angle.
Several X-ray telescopes, such as, the Advanced X-ray Astrophysics Facility use this.
The MSSTA works by a different effect. Its multi-layer mirrors comprise an ultrasmooth mirror coated by up to 100 layers of heavy elements like tungsten spaced by layers of lightweight elements like carbon. In effect, the layers work like a Bragg crystal, which will reflect X-rays. Everything is extremely smooth, on the order of 0.1 nm (a 10 billionth of a meter, or 1/250 millionth of an inch).
These reflect a little bit of the X-rays at the surface of each layer pair. The choice of materials and the thickness of the layers determine precisely which wavelength is making the X-rays interfere with each other reflection. In this way, the scientists can fine tune a telescope to observe in a narrow band of wavelengths (a spectral band) or even one wavelength. That makes it possible to measure the temperature of the solar atmosphere. To observe the sun in several wavelengths at once, several telescopes must be flown together.
This unique approach makes it possible to use conventional optical layouts - like the Hubble Space Telescope's Ritchey Chretien design - and get a much larger collecting area and brighter images than are possible grazing incidence optics of the same size. The design was invented by Barbee (and separately by scientists at IBM) and pioneered by Barbee, Walker, and Hoover for use in telescopes.
The MSSTA carries up to 19 telescopes of various sizes, each with a filter designed to admit only radiation of a specific wavelength or wavelength band, each corresponding to a specific temperature in the sun's atmosphere. Even though each image is taken in black-and-white, each represents a different wavelength and a different temperature in the solar atmosphere. To help in studying them, scientists often give them false colors to distinguish one from the other. This is similar to a color print that is really made from four black-and-white negatives, each to print a different color.
On its fourth flight, the array will include a telescope that can see FE XVII; iron stripped of 9 of its 26 electrons. That takes temperatures up to 5 million deg. K.
"It would be a better indicator of the distribution of high-temperature gases in the solar atmosphere," Walker said. This may also reveal small flares that may be one source of energy being pumped into the corona.
For the C/CS flight, expected by early 2000 near the around the time of solar maximum. MSSTA will be upgraded and some new telescopes and detectors installed. As with its first two flights, the telescope will be boosted by a Terrier Black Brant IX launched from the White Sands Missile Range, N.M. The C/CS payload will be boosted to an altitude of 230 km (144 mi) and fall then parachute back to Earth for recovery. During the coast above Earth's atmosphere, the telescope array will be pointed precisely at the sun for about 6 minutes. Each telescope will take 10 to 15 full-disk images. Ground-based observatories will take pictures at the same time in white light and H-alpha, and with telescopes equipped to map magnetic fields.
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"NASA Program Spawns New Safety Software For Pilots (31 March, 1998)",693,0,0,0
Two new software packages enabling pilots to use laptops to avoid hazardous terrain and find their place on maps are the latest success stories of a NASA program bringing together entrepreneurs and space engineers. Pilots of small planes, for whom such tools have been largely unavailable until now due to cost and the sheer size of bulky hardware, may soon be able to carry onboard the personal computer equivalent of collision-avoidance systems now used by the military and commercial airlines.
"TerrAvoid" and "Position Integrity" combine Global Positioning Satellite (GPS) data with high-resolution maps of the Earth's topography. Dubbs and Severino, Inc., based in Irvine, CA, has developed software that allows the system to be run on a battery-powered laptop in the cockpit. The packages, designed primarily for military sponsors and now positioned to hit the consumer market in coming months, came about as the result of the Technology Affiliates Program at NASA's Jet Propulsion Laboratory's (JPL), Pasadena, CA.
Intended to give American industry assistance from NASA experts and to facilitate business use of intellectual property developed for the space program, the Technology Affiliates Program introduced the start-up company of Dubbs and Severino to JPL's Dr. Nevin Bryant four years ago. Dubbs and Severino had an idea for mapping software to help private airplane pilots, inspired in part by the fatal crash of a pilot friend of company president Bob Severino.
The twist: the package was to be completely software-driven, instead of requiring expensive hardware, as was the norm up to that time. Bryant's Cartographic Applications Group at JPL had developed GeoTIFF, an architecture standard providing geo-location tools for mapping applications. GeoTIFF proved to be the crucial key that the start-up company needed to bring the idea to fruition, allowing the firm to develop low-cost software packages.
GeoTIFF is now in the public domain, and its use for commercial product development has evolved into an industry standard over the last year. Through the Technology Affiliates Program, Dubbs and Severino obtained JPL's assistance early on and thus gained a jump-start in adapting the architecture for their products' specific needs. "JPL gave us a demonstration and opened up the red carpet. It was a match made in heaven," says Severino. Merle McKenzie, manager of JPL's Commercial Technology Program, said that Dubbs and Severino's ability to utilize technology originally developed for NASA provides a strong example of the many advantages of technology transfer programs.
"This is a win-win partnership through which yet another American business gets a boost from the space program," McKenzie said. "TerrAvoid" is a terrain avoidance system that graphically shows pilots if they are flying dangerously close to mountains: safe sections can be seen in green, while hazardous sections show up in red, with those proportions changing in real time as the pilot moves through hilly terrain. In a sense, the system "looks" out over a plane's flight path, sweeping 360 degrees, warning the pilot if there are any upcoming hazards.
The software integrates GPS tracking data with maps on CD-ROM, and is approximately 1/20th the cost of its nearest competitor. "Position Integrity," which also co-registers real-time GPS data with local maps on CD-ROM, is a moving map detailing the exact position of the pilot. Because of the unique features of GeoTIFF, this software can be adapted to operate with any map, chart or photo image in the world, while comparable versions are limited solely to either military, scientific or commercial maps. GeoTIFF also enables the package to feature four windows at once, a useful and unique option for pilots who need to work simultaneously with maps, charts, photo images and sketches at different scales and zoom levels.
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"Earth Dragging Space and Time as it Rotates (27 March, 1998)",694,0,0,0
An international team of NASA and university researchers has found the first direct evidence of a phenomenon predicted 80 years ago using Einstein's theory of general relativity -- that the Earth is dragging space and time around itself as it rotates. Researchers believe they have detected the effect by precisely measuring shifts in the orbits of two Earth-orbiting laser-ranging satellites, the Laser Geodynamics Satellite I (LAGEOS I), a NASA spacecraft, and LAGEOS II, a joint NASA/Italian Space Agency (ASI) spacecraft.
The research, which is reported in the current edition of the journal Science, is the first direct measurement of a bizarre effect called "frame dragging." The team was led by Dr. Ignazio Ciufolini of the National Research Council of Italy and the Aerospace Department of the University of Rome, and Dr. Erricos Pavlis of the Joint Center for Earth System Technology, a research collaboration between NASA's Goddard Space Flight Center, Greenbelt, MD, and the University of Maryland at Baltimore County.
"General relativity predicts that massive rotating objects should drag space-time around themselves as they rotate," said Pavlis. "Frame dragging is like what happens if a bowling ball spins in a thick fluid such as molasses. As the ball spins, it pulls the molasses around itself. Anything stuck in the molasses will also move around the ball. Similarly, as the Earth rotates, it pulls space-time in its vicinity around itself. This will shift the orbits of satellites near the Earth.
"We found that the plane of the orbits of LAGEOS I and II were shifted about six feet (two meters) per year in the direction of the Earth's rotation," Pavlis said. "This is about 10 percent greater than what is predicted by general relativity, which is within our margin of error of plus or minus 20 percent. Later measurements by Gravity Probe B, a NASA spacecraft scheduled to be launched in 2000, should reduce this error margin to less than one percent. This promises to tell us much more about the physics involved."
Einstein's theory of general relativity has been highly successful at explaining how matter and light behave in strong gravitational fields, and has been successfully tested using a wide variety of astrophysical observations. The frame-dragging effect was first derived using general relativity by Austrian physicists Joseph Lense and Hans Thirring in 1918. Known as the Lense-Thirring effect, it was previously observed by the team of Ciufolini using the LAGEOS satellites and has recently been observed around distant celestial objects with intense gravitational fields, such as black holes and neutron stars.
The new research around Earth is the first direct detection and measurement of this phenomenon. The team analyzed a four-year period of data from the LAGEOS satellites from 1993 to 1996, using a method devised by Ciufolini three years ago. The other team members are Dr. Federico Chieppa of Scuola d'Ingegneria Aerospaziale of the University of Rome, and Drs. Eduardo Fernandes and Juan Perez-Mercader of Laboratorio de Astrofisica Espacial y Fisica Fundamental (LAEFF) in Madrid.
The measurements required the use of an extremely accurate model of the Earth's gravitational field, called the Earth Gravity Model 96, which became available only recently due to the collaborative work of the Laboratory for Terrestrial Physics at Goddard, the National Imagery and Mapping Agency (formerly the Defense Mapping Agency), Fairfax, VA, and the Ohio State University, Columbus, OH. It was developed over a four-year period using tracking data from approximately 40 spacecraft.
Dr. John Ries, an expert in satellite geodesy at the University of Texas at Austin, cautions that it is very challenging to remove the much larger effects of tidal changes and small zonal influences in the Earth's gravitational field, so that estimating the possible errors in the measurement of the Lense-Thirring effect is itself uncertain. "The relativistic effect being sought is about ten million times smaller than classical Newtonian disturbances on the plane of the LAGEOS orbits, requiring an enormously accurate treatment of background effects," said Dr. Alan Bunner, science program director for the Structure and Evolution of the Universe in the Office of Space Science at NASA headquarters, Washington, DC. LAGEOS II, launched in 1992, and its predecessor, LAGEOS I, launched in 1976, are passive satellites dedicated exclusively to laser ranging, which involves sending laser pulses to the satellite from ranging stations on Earth and then recording the round-trip travel time.
Given the well-known value for the speed of light, this measurement enables scientists to determine precisely the distances between laser ranging stations on Earth and the satellite. LAGEOS is designed primarily to provide a reference point for experiments that monitor the motion of the Earth's crust, measure and understand the "wobble" in the Earth's axis of rotation, and collect information on the Earth's size, shape, and gravitational field. Such research is part of NASA's Earth Science enterprise, a coordinated research program that studies the Earth's land, oceans, ice, atmosphere and life as a total system.
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"Mars Global Surveyor to Attempt Imaging of Features of Public Interest (26 March, 1998)",695,0,0,0
NASA's Mars Global Surveyor spacecraft is about to begin a summer-long set of scientific observations of the red planet from an interim elliptical orbit, including several attempts to take images of features of public interest ranging from the Mars Pathfinder and Viking mission landing sites to the Cydonia region. The spacecraft will turn on its payload of science instruments on March 27, about 12 hours after it suspends "aerobraking," a technique that lowers the spacecraft╣s orbit by using atmospheric drag each time it passes close to the planet on each looping orbit.
Aerobraking will resume in September and continue until March 1999, when the spacecraft will be in a final, circular orbit for its prime mapping mission. It will not be possible to predict on which orbit the spacecraft will pass closest to specific features on Mars until Global Surveyor has established a stable orbit and flight controllers are able to project its ground track. This process should be completed in the next few days.
The exact time of observations and the schedule for the subsequent availability of photographs on the World Wide Web are expected to be announced early next week. "Global Surveyor will have three opportunities in the next month to see each of the sites, including the Cydonia region, location of the so-called 'Face on Mars,' " said Glenn E. Cunningham, Mars Global Surveyor project manager at NASA's Jet Propulsion Laboratory, Pasadena, CA.
"The sites will be visible about once every eight days, and we'll have a 30- to- 50-percent chance of capturing images of the sites each time." Several factors limit the chances of obtaining images of specific features with the high-resolution mode of the camera on any one pass. These factors are related primarily to uncertainties both in the spacecraft╣s pointing and the knowledge of the spacecraft╣s ground track from its navigation data.
In addition, current maps of Mars are derived from Viking data taken more than 20 years ago. Data obtained by Global Surveyor╣s laser altimeter and camera during the last few months have indicated that our knowledge of specific locations on the surface is uncertain by 0.6 to 1.2 miles (1 to 2 kilometers). As a result, the locations of the landing sites and specific features in the Cydonia region are not precisely known.
In addition, the Mars Pathfinder and Viking landers are very small targets to image, even at the closest distance possible, because they are the smallest objects that the camera can see. The Cydonia features, on the other hand, are hundreds to thousands of times larger and the camera should be able to capture some of the features in that area. Global Surveyor╣s observations of the Viking and Pathfinder landing sites will provide scientists with important information from which to tie together surface observations and orbital measurements of the planet.
Data from landing sites provide "ground truth" for observations of the planet made from space. As for the "Face on Mars" feature, "most scientists believe that everything we've seen on Mars is of natural origin," said Dr. Carl Pilcher, acting science director for Solar System Exploration in NASA's Office of Space Science, Washington, DC. "However, we also believe it is appropriate to seek to resolve speculation about features in the Cydonia region by obtaining images when it is possible to do so."
During the aerobraking hiatus, the spacecraft will be orbiting Mars about once every 11.6 hours, passing about 106 miles (170 kilometers) above the surface at closest approach and about 11,100 miles (17,864 kilometers) at its farthest distance from the planet. The pause in aerobraking allows the spacecraft to achieve a final orbit with lighting conditions that are optimal for science observations. Mars Global Surveyor is part of a sustained program of Mars exploration, managed by JPL for NASA's Office of Space Science, Washington, DC. Lockheed Martin Astronautics, Denver, CO, which built and operates the spacecraft, is JPL's industrial partner in the mission. Malin Space Science Systems, Inc., San Diego, CA, built and operates the spacecraft camera. JPL is a division of the California Institute of Technology, Pasadena, CA.
A NASA scientist has found a new puzzle in the sky, an X-ray pulsar that appears to be in a lopsided orbit that makes it burst twice every "year" rather than once.
In a paper accepted for publication in The Astrophysical Journal, Colleen Wilson of NASA's Marshall Space Flight Center describes the discovery of an accreting X-ray pulsar which has no visible component.
When it is active, GRO J2058+42 appears to burst in X-rays twice each time it circles its primary star. "It's an unusual behavior," she said. "It's not like what we see with other accretion-powered pulsars." These are pulsars that burst in X-rays and gamma rays as they gobble gas in a disk from a larger parent star.
Wilson made her discovery with the Burst and Transient Source Experiment aboard the Compton Gamma Ray Observatory, and built on it with additional observations by the Rossi X-ray Timing Explorer. This one of 12 known transient accreting X-ray pulsars with no visible companion.
Pulsars - rotating neutron stars - are among the most intriguing objects in the sky. They were found in 1965 when radio astronomers discovered several objects that emitted radio waves with clock-like precision.
The sources soon were identified as rapidly rotating neutron stars with intense magnetic fields. Where radio pulsars have the regularity of a Swiss watch, accretion pulsars are like cheap alarm clocks that easily gain and lose time - and go off when you least expect it.
Since the April 1991 launch of BATSE, a NASA/Marshall instrument, astrophysicists have detected 20 of 45 known pulsars and discovered another 5 new X-ray pulsars. These accreting pulsars are not the same as the mysterious gamma-ray bursts that appear to be coming from the edge of the universe. They are energetic enough to be recorded by BATSE, but don't always get noticed right away.
So, as Wilson reviewed BATSE data in September 1995, she found a burst that registered 140 milliCrabs (or, 140/1,000ths the brightness of the Crab Nebula, which astrophysicists use as a standard candle). By using a computer to fold the data on itself, Wilson found that the source repeated every 198 seconds, an indication of a massive, compact object spinning at high speed.
"The way we can tell it was something new is that the ones that aren't real don't repeat from day to day," Wilson said.
GRO J2058+42 (the numbers indicate its approximate position, 20 hrs, 58 minutes along the celestial equator, and 42 degrees north) had signaled the start of a giant burst spanning a 46-day period. The burst peaked at 140 milliCrab (140/1,000ths the brightness of the Crab Nebula). BATSE saw another 500 days of weaker bursts.
BATSE and the Rossi all-sky monitor showed bursts every 54 days, but BATSE saw the bursts alternate in brightness in a 110-day cycle. Clearly, Wilson was on to something new.
In the other accretion-powered pulsars, the emissions come only once an orbit when they do occur; sometimes they fade for weeks or months. J2058+42's twice-an-orbit bursts are new.
Wilson said one possibility is that J2058+42 is a binary star system composed of a type Be star (a type B star with emission lines), about 8 to 15 times the mass of our sun, with a neutron star in a lopsided orbit (Picture 2).
An excretion disk of gas, expelled by the Be star in an unusual stellar wind, orbits the Be star about even with the equator. But the neutron star is in an elliptical orbit inclined to the equator, so it sails in and out of the excretion disk twice every 110 days.
As a result, J2058+42 is quiet most of the time, and emits X-rays twice every orbit as the dense, magnetically intense neutron star plows through the cloud of gas. When the pulsar is near periastron (closest approach to the star), the bursts are more intense, now around 1.4 percent the brightness of the Crab Nebula. When the pulsar is near apastron (most distant from the star), the bursts are weaker, down to less than 1 percent of the Crab, almost at the threshold of the BATSE.
A more likely explanation, Wilson said, is that somehow the pulsar winds up the accretion disk at periastron and accelerates its consumption of matter, and at apastron it's taking a lighter meal from the stellar wind.
The identification is based on J2058+42's behavior matching that of other Be-pulsar binaries.
"About half the known pulsars in high-mass X-ray binaries are thought to be with Be stars," she said. The high mass here refers to the mass of the companion star, not the pulsar.
Unfortunately, no one has been able to get a look at the object for other observations that might confirm an optical identification. Wilson said this is the sixth time BATSE has found such an object for which no visible component can be found.
"Part of the problem is that the error box is so big," she said. Because BATSE is designed to watch the entire sky, it cannot locate a source with the same precision of a telescope zeroing in on a star. Other data from before September 1995 from BASTE and from Roentgensatellit (a European Space Agency satellite) also don't show anything that stands out as a possible source. Neither does a survey of optical images, possibly because interstellar dust simply dims the star before its light arrives here.
Even though observations with Rossi have narrowed the position estimate to 4 arc-minutes (about 1/8th the apparent diameter of the Moon), the error box is still quite large for an optical telescope to search. Wilson estimates the distance to J2058+42 at about 23,000 to almost 50,000 light years (almost a quarter of the way across the galaxy).
Meanwhile, J2058+42 has become too quiet for BATSE to detect. Wilson will keep checking the data, and also looks forward to observations from the Rossi Explorer which should be made this week.
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"NASA Starts Work on New Space Infrared Telescope Facility (25 March, 1998)",697,0,0,0
NASA Administrator Daniel S. Goldin today authorized the start of work on the Space Infrared Telescope Facility (SIRTF), an advanced orbiting observatory that will give astronomers unprecedented views of phenomena in the universe that are invisible to other types of telescopes. The authorization signals the start of the design and development phase of the SIRTF project, which is managed by NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA. Scheduled for launch in December 2001 on a Delta 7920-H rocket from Cape Canaveral, FL, SIRTF represents the culmination of more than a decade of planning and design to develop an infrared space telescope with high sensitivity, low cost and long lifetime of at least two-and-a-half to as many as five years.
"The Space Infrared Telescope Facility will do for infrared astronomy what the Hubble Space Telescope has done in its unveiling of the visible universe, and it will do it faster, better and cheaper than its predecessors," said Dr. Wesley Huntress, NASA's associate administrator for space science. "By sensing the heat given off by objects in space, this new observatory will see behind the cosmic curtains of dust particles that obscure much of the visible universe. We will be able to study fetal stars, detect other solar systems and study the most ancient, distant galaxies at the edge of the universe."
Conventional optical telescopes can study stars and other objects that glow brightly enough to emit light in the visible portion of the electromagnetic spectrum. However, many objects, such as planets and unignited stars, do not "shine" in visible or ultraviolet light. Others that may burn brightly are veiled from Earth's view behind the vast clouds of dust and gas that populate the universe. Some of the most fascinating objects and processes in the universe may exist behind these cosmic curtains of dust and gas, such as black holes, quasars, regions where stars are forming in galaxies and regions where planets are forming around stars.
Most of these concealed attractions are detectable only with infrared telescopes, whose unique capability lies in their ability to sense the heat of dark, faint or hidden objects. Infrared telescopes also provide the means to study the most distant objects at the edge of the expanding universe. Optical and ultraviolet light emitted from stars, galaxies and quasars since the birth of the universe has shifted, over time and distance, into the infrared portion of the spectrum. SIRTF will provide important insights into when and how the first galaxies and stars formed. SIRTF, whose design and development is cost-capped at $458 million, will be one of astronomy's most advanced telescopes.
Its unconventional approach uses new technologies, an innovative mission design and small launch vehicle. It is being developed on a quick schedule that closely integrates the work of the contractor and academic teams responsible for development and delivery. Its design promises high sensitivity and observing capability along with efficiency of operations and long lifetime. SIRTF is the fourth and final element in NASA's family of complementary spaceborne "Great Observatories" that includes the Hubble Space Telescope, the Compton Gamma Ray Observatory and the Advanced X-ray Telescope Facility.
The project also represents a bridge to NASA's new Origins program, which seeks to answer fundamental questions about the birth and evolution of the universe. SIRTF will lay the groundwork for many investigations fundamental to the Origins program, such as studies of the birth and evolution of galaxies, their stars, and searches for planets that orbit some of those stars. Lockheed Martin Missiles and Space, Sunnyvale, CA, is responsible for the spacecraft and for the SIRTF system integration and testing. Ball Aerospace and Technology Corp., Boulder, CO, is responsible for the cryogenic telescope assembly.
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"NASA Set to Study the Sun's Turbulent Upper Atmosphere (19 March, 1998)",698,0,0,0
NASA's Transition Region and Coronal Explorer (TRACE) mission, scheduled for launch at 9:40 p.m. EST (6:40 p.m. PST) March 30, 1998, will greatly improve understanding of events in the Sun's atmosphere, including intense storms and flares, which can have an impact on power and communications systems on Earth. The TRACE mission will join a fleet of spacecraft studying the Sun during a critical period when solar activity is beginning its rise to a peak early in the new millennium.
The Sun goes through an 11-year cycle from a period of numerous intense storms and sunspots to a period of relative calm and then back again. The coming months in the Sun's cycle will provide solar scientists with periods of strong solar activity interspersed with periods when the Sun is relatively passive and quiet. This will give TRACE the chance to study the full range of solar conditions, even in its relatively short planned lifetime. TRACE will train its powerful telescope on the dynamic so- called 'transition region' of the Sun's atmosphere, between the relatively cool surface and lower atmosphere of the Sun where temperatures are about 6,000 degrees Fahrenheit, and the extremely hot upper atmosphere called the corona, where temperatures are up to 16 million degrees Fahrenheit.
Using instruments sensitive to extreme-ultraviolet and ultraviolet wavelengths of light, TRACE will study the detailed connections between the fine-scale surface features and the overlying, changing atmospheric structures of hot, ionized gas, called plasma. The surface features and atmospheric structures are linked by fine-scale solar magnetic fields. The power of the TRACE telescope to do detailed studies of the solar atmosphere makes this observatory unique among the current group of spacecraft studying the Sun.
"The spacecraft has roughly ten times the temporal resolution and five times the spatial resolution of previously launched solar spacecraft. Its findings are eagerly awaited by the solar science community," said Dr. Alan Title, TRACE principal investigator from the Stanford Lockheed Institute for Scientific Research in Palo Alto, CA. "We can expect to resolve some present mysteries of the Sun's atmospheric dynamics as well as discover new and exciting phenomena." TRACE will be launched into a polar orbit to enable virtually continuous observations of the Sun, uninterrupted by the Earth's shadow for months at a time.
This orbit will give the mission the greatest chance of observing the random processes which lead to flares and massive eruptions in the Sun's atmosphere. The TRACE telescope is really four telescopes in one. Its 30-centimeter (12-inch) primary and six-centimeter (2-inch) secondary super-polished mirrors are individually coated in four distinct quadrants to allow light from different bandwidths (colors) to be reflected and analyzed.
An electronic detector collects images over a 231,000-by-231,000- mile field of view, nearly 25 percent of the Sun's disk. A powerful data handling computer enables very flexible use of the detector array including adaptive target selection, data compression and image stabilization. "TRACE was completed on time, under budget, and met all performance goals," said Jim Watzin, Small Explorer project manager, NASA Goddard Space Flight Center, Greenbelt, MD. "I'm really proud of this team. They have produced a magnificent observatory in a manner that saved NASA nearly $9.7 million over the initial cost estimate."
TRACE, which costs $49 million, is the third launch in the Small Explorer series of small, quickly developed, relatively low-cost missions. TRACE will be launched on an Orbital Sciences Corp., Dulles, VA, Pegasus-XL rocket released from an L-1011 jet aircraft at the Western Range, Vandenberg Air Force Base, CA. The launch window is open for 10 minutes. TRACE will be the first space science mission with an open data policy. All data obtained by TRACE will be available to other scientists, students and the general public shortly after the information becomes available to the primary science team.
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"New Global Surveyor Data Reveals More (13 March, 1998)",699,0,0,0
For the first time in Mars exploration, a spacecraft has captured the full evolution of a Martian dust storm. NASA's Mars Global Surveyor mission also has returned new insights into the deeply layered terrain and mineral composition of the Martian surface, and to highly magnetized crustal features that provide important clues about the planet's interior.
These findings are among the early results from the Mars- orbiting mission being reported in today's issue of Science magazine. This first set of formal results comes from data obtained in October and November 1997, while the spacecraft was just beginning to use the drag of Mars' upper atmosphere to lower and circularize its highly elliptical orbit in a process called aerobraking. At the time, a dust storm was brewing on Mars and had grown to about the size of the South Atlantic Ocean.
The Global Surveyor data suggest that the event began as a set of small dust storms along the edge of the planet's southern polar cap, according to Dr. Arden Albee of the California Institute of Technology, Pasadena, CA, the Mars Global Surveyor mission scientist. By Thanksgiving, it had expanded into a large regional dust storm in Noachis Terra that covered almost 180 degrees longitude, while spanning 20 degrees south latitude to nearly the tip of the Martian equator.
"As this storm obscured the Martian landscape, we followed it in detail using several instruments onboard Mars Global Surveyor," Albee said. "The Thermal Emission Spectrometer mapped the temperature and opacity of the atmosphere while the camera followed the visual effects. The effects of the storm extended to great heights of about 80 miles (130 kilometers) and resulted in great increases in both atmospheric density and variability from orbit to orbit. These atmospheric measurements have great significance to future Mars missions that will be using aerobraking techniques too."
Before the storm, atmospheric dust was generally distributed very uniformly, Albee said. Observations of the limb of the planet in the northern hemisphere revealed both low-lying dust hazes and detached water-ice clouds at altitudes of up to 34 miles (55 kilometers). Movement of these clouds was tracked by the spectrometer as the planet rotated. Atmospheric turbulence disrupted these cloud patterns as the small storms began to rise and kick more dust into the air. As the storm began to abate, small local storms began to crop up again along the edges of the south polar cap, and ice clouds formed in depressions as the carbon dioxide cap continued to retreat.
In addition to these unprecedented observations of a full- blown Martian dust storm, measurements from the spacecraft's Magnetometer and Electron Reflectometer have yielded new findings about Mars' strong, localized magnetic fields. These patches of the crust, which register high levels of magnetism, are beginning to unlock some of the mysteries surrounding Mars' internal dynamo and when it died, said Dr. Mario Acuna of NASA's Goddard Space Flight Center, Greenbelt, MD. "These locally magnetized areas on Mars could not form without the presence of an overall global magnetic field that was perhaps as strong as Earth's is today," says Acuna.
"Since the internal dynamo that powered the global field is extinct, these local magnetic fields act as fossils, preserving a record of the geologic history and thermal evolution of Mars." Magnetic fields are created by the movement of electrically conducting fluids, and a planet can generate a global magnetic field if its interior consists of molten metal hot enough to undergo convective motion, similar to the churning motion seen in boiling water. "The small size and highly magnetic nature of these crustal features, which measure on the order of 30 miles (50 kilometers), are found within the ancient cratered terrain rather than within the younger volcanic terrain," Acuna said. "By correlating crustal age with magnetization, we have a perfect window on Mars' past, which will help us to determine when Mars' internal dynamo ceased operating."
High-resolution images of dunes, sandsheets and drifts are also helping to reveal earlier chapters of Martian history. Landforms shaped by erosion are almost everywhere, according to Albee, and many bear a striking resemblance to Colorado's Rocky Mountains. Rocky ridges poke through the Martian dust just as the jagged edges of cliffs pierce through a blanket of snow in the Rockies. Martian dust appears to have spilled down the sides of ridges just as fresh snow slides down a ski slope.
"One almost expects to see ski tracks crisscrossing the area," Albee added. "These images present a sharp contrast to the images of boulder-strewn deserts found at the Viking and Pathfinder landing sites."
The Martian crust also exhibits much more layering at great depth than was expected. The steep walls of canyons, valleys and craters show the Martian crust to be stratified at scales of a few tens of yards, which is an exciting discovery, Albee noted. "At this point we simply do not know whether these layers represent piles of volcanic flows or sedimentary rocks that might have formed in a standing body of water," he said.
The Thermal Emission Spectrometer, led by principal investigator Dr. Philip Christensen of Arizona State University, is beginning to obtain a few infrared emission spectra of the surface, although it is still too cold on the surface for the best results. The best spectra clearly indicate the presence of pyroxene and plagioclase, minerals which are common in volcanic rocks, with a variable amount of dust component. No evidence was found for carbonate minerals, clay minerals or quartz. If present in these rocks, their abundance must be less than about ten percent.
Their absence indicates that carbonates are not widespread over the surface of the planet, but they may still be found in specific locations that either favored their initial deposition or their subsequent preservation. This finding could have important implications for identifying areas that may preserve signs of ancient life on Mars, since carbonate minerals are commonly formed in biological processes, Albee said.
Striking results also have been obtained from Global Surveyor's laser altimeter over Mars' northern hemisphere, which is exceptionally flat with slopes and surface roughness increasing toward the equator, according to principal investigator Dr. David Smith of Goddard. The initial data for this region helps scientists interpret a variety of landforms, including the northern polar cap, gigantic canyons, ridges, craters of all sizes and shield volcanoes. Most surprising are views of extraordinarily mundane regions -- as flat as the Bonneville Salt Flats in Utah - that extend over vast northern regions of the planet.
Mars Global Surveyor will complete the first phase of its two-part aerobraking strategy at the end of March, at which time the science instruments will be turned on again for most of the next six months. Over this period, the spacecraft will stay in an 11 1/2-hour orbit and collect an additional bounty of data at a closest approach of about 106 miles (170 kilometers) above the surface, much closer than the spacecraft will pass over the planet once it has reached its formal mapping orbit in March 1999. This closer orbit will allow the science teams to take more detailed measurements of the Martian atmosphere and surface without magnetic interference from the solar wind.
"When we decided to slow the pace of aerobraking to reduce the force on the solar panel that was damaged after launch, we knew we would get a bonus -- the ability to collect much more science data closer to the planet than will be possible during the prime mapping mission," said Glenn E. Cunningham, Mars Global Surveyor project manager at NASA's Jet Propulsion Laboratory, Pasadena, CA.
"Additionally, the six-month period between the end of March and early September will yield an extraordinary opportunity as the lowest point of the orbit migrates over the northern polar cap. All of this information that is coming back now is really icing on the cake, a spectacular precursor to the global mapping data expected to start flowing next year." Mars Global Surveyor is part of a sustained program of Mars exploration known as the Mars Surveyor Program. The mission is managed by the Jet Propulsion Laboratory for NASA's Office of Space Science, Washington, DC. JPL's industrial partner is Lockheed Martin Astronautics, Denver, CO, which developed and operates the spacecraft. JPL is a division of the California Institute of Technology.
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"More Evidence Points To Impact As Dinosaur Killer (12 March, 1998)",700,0,0,0
Note: This electronic version incorporates a correction that was made after the original news release was issued.
Two new sites in Belize and Mexico add further evidence to the hypothesis that an asteroid or comet collided with Earth about 65 million years ago, subsequently killing off the dinosaurs and many other species on the planet. Researchers Adriana Ocampo of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, and Kevin Pope of Geo Eco Arc Research, La Canada-Flintridge, CA, led an international team that discovered the two new sites during a recent expedition sponsored by NASA's Exobiology Program and The Planetary Society, Pasadena, CA.
"We discovered an important new site in Alvaro Obregon, Mexico, about 140 miles (230 kilometers) from the rim of the Chicxulub crater. This crater was formed when a 6-to-8-mile diameter (10-to-14-kilometer diameter) asteroid or comet collided with Earth," Ocampo said. "The site contains two layers of material, or ejecta, thrown out by the impact that flowed across the surface like a thick fluid, known as fluidized ejecta lobes," added Pope.
"This is the closest surface exposure of ejecta to the Chicxulub crater that has yet been found and the best example known on Earth from a really big impact crater." Centered on the coast of Yucatan, Mexico, the Chicxulub crater is estimated to be about 120 miles (200 kilometers) in diameter. The impact 65 million years ago kicked up a global cloud of dust and sulfur gases that blocked sunlight from penetrating through the atmosphere and sent Earth into a decade of near-freezing temperatures.
The drop in temperature and related environmental effects are thought to have brought about the demise of the dinosaurs and about 75 percent of the other species on Earth. The Earth orbits the Sun in a swarm of so-called near-Earth objects, whether they are comets or asteroids, yet the science of detecting and tracking them is still relatively young. Only a handful of astronomers around the world search for these objects, and they estimate that currently only about one-tenth of the population of near-Earth objects has been detected.
Chicxulub is the only impact event that has been correlated with mass extinctions to date. The site has been dated geologically to the boundary between the Cretaceous and Tertiary periods, also known as the K/T boundary. Local geologist Brian Holland of Punta Gorda, Belize, guided the expedition to another new ejecta site about 290 miles (480 kilometers) from the crater rim. This Belize site contains tiny spheres of altered green glass, called tektites. Tektites are rocks that have been melted to glass by the severe heat of an impact.
Expedition member Jan Smit of Free University, Amsterdam, noted that the Belize tektites were similar to those found in Haiti and northern Mexico. This finding links the stratigraphy of the Belize sites to the more distant Caribbean and Mexican ejecta sites. Alfred Fischer of the University of Southern California, Michael Gibson of the University of Tennessee at Martin, and Jaime Urrutia and Francisco Vega of the National Autonomous University of Mexico helped the team collect 900 pounds (400 kilograms) of samples, including drill cores, for paleomagnetic studies. They also collected fossils from the site to help date the deposits and add new pieces to the puzzle of what happened at Chicxulub 65 million years ago.
Impact ejecta is very rare on Earth, but covers much of the surface of Mars because Mars' surface has remained stable and unchanged for billions of years, thus preserving debris from these rare impact events. Also, such fluidized ejecta lobes have never been observed directly on Earth before and can serve as an excellent laboratory for studying the ejecta lobes surrounding many Martian craters. "The discovery of these new ejecta sites is very exciting," said team co-leader Ocampo. "It is like seeing a bit of Mars on Earth." The exact nature of these ejecta lobes on Mars remains a mystery, Ocampo noted. Some scientists think they were created by an abundance of water in the Martian crust, which turned the ejecta into a muddy, molasses-like material.
Others suggest the fluidized ejecta lobes were enabled by a much thicker atmosphere in Mars' early history. As flying ejecta from an impact event flew through the Martian atmosphere, it was reduced by friction to a very dense, turbulent cloud of debris, which also flowed like water. Study of the Chicxulub fluidized ejecta may help settle this debate and shed new light on theories that the Martian surface may once have been more hospitable for life.
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"Astronomers Track Down Asteroids in Hubble Archive (9 March, 1998)",701,0,0,0
Astronomers have stumbled on an unusual asteroid hunting ground: the thousands of images stored in the Hubble Space Telescope archive. The hunt, by Robin Evans and Karl Stapelfeldt of NASA's Jet Propulsion Laboratory, Pasadena, CA, has yielded a sizable catch of small asteroids -- about 100. Their preliminary analysis suggests that a total population of 300,000 small asteroids -- essentially rocks just over half a mile to two miles wide (1-3 kilometers) -- are orbiting between Mars and Jupiter in a band of space debris known as the main belt. Currently, there are 8,319 confirmed main-belt asteroids whose orbits have been measured, and about the same number have been sighted but not confirmed.
Most astronomers stalk the Hubble archive for bigger game, such as quasars, distant galaxies, and supernovae, but Evans and Stapelfeldt have discovered that the pursuit of smaller prey such as asteroids can be equally successful. Over a three-year period, the two astronomers and their collaborators have searched through more than 28,000 Wide Field and Planetary Camera 2 (WFPC2) images, looking for wide, looping streaks of light, the telescope's tell-tale signatures of asteroids. Most of the ones they found are too faint to be observed by current ground-based search programs.
Hubble captures their images purely by accident: Nearby asteroids inevitably wander across the telescope's field of view while other, higher priority targets are being observed. "The archive images are distributed fairly evenly across the sky, so we find asteroids according to both their position in the sky and their number," Evans said. "As expected, we see the asteroids concentrated towards the ecliptic plane and we see small asteroids because they are the most numerous.
Small main-belt asteroids such as these are the ones most likely to evolve into Earth-crossing asteroids due to encounters with their larger neighbors. Some of the asteroids in our survey could eventually migrate toward Earth." The Hubble archives represent a newly tapped information resource which could help scientists more precisely estimate the risks the asteroids pose to Earth. According to Evans and Stapelfeldt, the Hubble archival data also strongly limit the number of small comets that could be passing very near Earth.
Last year, Dr. Louis A. Frank of the University of Iowa in Iowa City, using data from NASA's Polar spacecraft, reported he found evidence that about a dozen small comets strike Earth's upper atmosphere each minute. Evans and Stapelfeldt estimate the such small comets should be bright enough to produce thousands of detectable trails in the Hubble archival images, but these were not seen. The Hubble images capture an asteroid as a long trail produced by its motion across the camera's field of view. The trails appear like the streaks of light found on photos taken at night of speeding cars with their headlights on.
Finding asteroids isn't what the two astronomers originally had in mind. As members of the WFPC2 science team, Evans and Stapelfeldt were examining test images of distant stars and galaxies to ensure that the new camera was functioning properly. These were among the first images taken with WFPC2, which had restored sharp focus to Hubble's images when it was installed in late 1993.
Stapelfeldt's wife, Deborah Padgett (also an astronomer), pinpointed the first asteroid in 1994 while looking at images on the couple's home computer. Intrigued, Evans and Stapelfeldt began combing through more than 1,600 of the science team's survey photos, finding 12 more asteroids. This discovery prompted their large-scale search, by eye, of two years worth of Hubble archival images.
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"Lunar Prospector Finds Evidence of Ice at Moon's Poles (5 March, 1998)",702,0,0,0
There is a high probability that water ice exists at both the north and south poles of the Moon, according to initial scientific data returned by NASA's Lunar Prospector. The Discovery Program mission also has produced the first operational gravity map of the entire lunar surface, which should serve as a fundamental reference for all future lunar exploration missions, project scientists announced today at NASA's Ames Research Center, Moffett Field, CA.
Just two months after the launch of the cylindrical spacecraft, mission scientists have solid evidence of the existence of lunar water ice, including estimates of its volume, location and distribution. "We are elated at the performance of the spacecraft and its scientific payload, as well as the resulting quality and magnitude of information about the Moon that we already have been able to extract," said Dr. Alan Binder, Lunar Prospector Principal Investigator from the Lunar Research Institute, Gilroy, CA.
The presence of water ice at both lunar poles is strongly indicated by data from the spacecraft's neutron spectrometer instrument, according to mission scientists. Graphs of data ratios from the neutron spectrometer "reveal distinctive 3.4 percent and 2.2 percent dips in the relevant curves over the northern and southern polar regions, respectively," Binder said. "This is the kind of data 'signature' one would expect to find if water ice is present."
However, the Moon's water ice is not concentrated in polar ice sheets, mission scientists cautioned. "While the evidence of water ice is quite strong, the water 'signal' itself is relatively weak," said Dr. William Feldman, co-investigator and spectrometer specialist at the Department of Energy's Los Alamos National Laboratory, NM.
"Our data are consistent with the presence of water ice in very low concentrations across a significant number of craters." Using models based on other Lunar Prospector data, Binder and Feldman predict that water ice is confined to the polar regions and exists at only a 0.3 percent to 1 percent mixing ratio in combination with the Moon's rocky soil, or regolith. How much lunar water ice has been detected? Assuming a water ice depth of about a foot and a half (.5 meters) -- the depth to which the neutron spectrometer's signal can penetrate -- Binder and Feldman estimate that the data are equivalent to an overall range of 11 million to 330 million tons (10-300 million metric tons) of lunar water ice, depending upon the assumptions of the model used.
This quantity is dispersed over 3,600 to 18,000 square miles (10,000-50,000 square kilometers) of water ice- bearing deposits across the northern pole, and an additional 1,800 to 7,200 square miles (5,000-20,000 square kilometers) across the southern polar region. Furthermore, twice as much of the water ice mixture was detected by Lunar Prospector at the Moon's north pole as at the south.
Dr. Jim Arnold of the University of California at San Diego previously has estimated that the most water ice that could conceivably be present on the Moon as a result of meteoritic and cometary impacts and other processes is 11 billion to 110 billion tons. The amount of lunar regolith that could have been "gardened" by all impacts in the past 2 billion years extends to a depth of about 6.5 feet (2 meters), he found. On that basis, Lunar Prospector's estimate of water ice would have to be increased by a factor of up to four, to the range of 44 million to 1.3 billion tons (40 million to 1.2 billion metric tons).
In actuality, Binder and Feldman caution that, due to the inadequacy of existing lunar models, their current estimates "could be off by a factor of ten in either direction." The earlier joint Defense Department-NASA Clementine mission to the Moon used a radar-based technique that detected ice deposits in permanently shadowed regions of the lunar south pole. It is not possible to directly compare the results from Lunar Prospector to Clementine because of their fundamentally different sensors, measurement "footprints," and analysis techniques.
However, members of the Clementine science team concluded that its radar signal detected from 110 million to 1.1 billion tons (100 million to 1 billion metric tons) of water ice, over an upper area limit of 5,500 square miles (15,500 square kilometers) of south pole terrain. There are various ways to estimate the economic potential of the detected lunar water ice as a supporting resource for future human exploration of the Moon. One way is to estimate the cost of transporting that same volume of water ice from Earth to orbit.
Currently, it costs about $10,000 to put one pound of material into orbit. NASA is conducting technology research with the goal of reducing that figure by a factor of 10, to only $1,000 per pound. Using an estimate of 33 million tons from the lower range detected by Lunar Prospector, it would cost $60 trillion to transport this volume of water to space at that rate, with unknown additional cost of transport to the Moon's surface.
From another perspective, a typical person consumes an estimated 100 gallons of water per day for drinking, food preparation, bathing and washing. At that rate, the same estimate of 33 million tons of water (7.2 billion gallons) could support a community of 1,000 two-person households for well over a century on the lunar surface, without recycling. "This finding by Lunar Prospector is primarily of scientific interest at this time, with implications for the rate and importance of cometary impacts in the history and evolution of the Solar System," said Dr. Wesley Huntress, NASA Associate Administrator for Space Science.
"A cost-effective method to mine the water crystals from within this large volume of soil would have to be developed if it were to become a real resource for drinking water or as the basic components of rocket fuel to support any future human explorers." Before the Lunar Prospector mission, historical tracking data from various NASA Lunar Orbiter and Apollo missions had provided evidence that the lunar gravity field is not uniform. Mass concentrations caused by lava which filled the Moon's huge craters are known to be the cause of the anomalies.
However, precise maps of lunar mass concentrations covering the moon's equatorial nearside region were the only ones available. Lunar Prospector has dramatically improved this situation, according to co-investigator Dr. Alex Konopliv of NASA's Jet Propulsion Laboratory, Pasadena, CA. Telemetry data from Lunar Prospector has been analyzed to produce a full gravity map of both the near and far side of the moon.
Konopliv also has identified two new mass concentrations on the Moon's nearside that will be used to enhance geophysical modeling of the lunar interior. This work has produced the first-ever complete engineering-quality gravity map of the moon, a key to the operational safety and fuel-efficiency of future lunar missions. "This spacecraft has performed beyond all reasonable expectations," said NASA's Lunar Prospector mission manager Scott Hubbard of Ames.
"The findings announced today are just the tip of the iceberg compared to the wealth of information forthcoming in the months and years ahead." Lunar Prospector is scheduled to continue its current primary data gathering mission at an altitude of 62 miles (100 kilometers) for a period of ten more months. At that time, the spacecraft will be put into an orbit as low as six miles (10 kilometers) so that its suite of science instruments can collect data at much finer resolution in support of more detailed scientific studies.
In addition, surface composition and structure information developed from data returned by the spacecraft's Gamma Ray Spectrometer instrument will be a crucial aspect of additional analysis of the polar water ice finding over the coming months. The third launch in NASA's Discovery Program of lower cost, highly focused planetary science missions, Lunar Prospector is being implemented for NASA by Lockheed Martin, Sunnyvale, CA, with mission management by NASA Ames. The total cost to NASA of the mission is $63 million.
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"Collins Named First Female Shuttle Commander (5 March, 1998)",703,0,0,0
Astronaut Eileen Collins (Lt. Col., USAF) will become the first woman to command a Space Shuttle when Columbia launches on the STS-93 mission in December 1998. First Lady Hillary Rodham Clinton made the announcement today in the Roosevelt Room at the White House. Collins will be joined on the flight deck by Pilot Jeffrey S. Ashby (Cmdr., USN) and Mission Specialists Steven A. Hawley, Ph.D., and Catherine G. "Cady" Coleman, Ph.D (Major, USAF). CNES Astronaut Michel Tognini (Col., French Air Force) was named to the crew on November 12.
Selected as an astronaut in 1990, Collins has served as a pilot on her two previous space flights. Her first space flight was STS-63 in February 1995 as Discovery approached to within 30 feet of Mir, in a dress rehearsal for the first Shuttle/Mir docking. In May 1997, she visited the Mir space station as pilot on board Atlantis for the sixth Shuttle/Mir docking mission, delivering Astronaut Mike Foale and returning Jerry Linenger to Earth. STS-93 will be the first flight for Ashby.
Hawley will be making his fifth space flight during STS-93, having flown previously on STS-41D in 1984, STS-61C in 1986, STS-31 in 1990 and STS-82 in 1997. Coleman has one previous space flight to her credit, having flown on STS-73, the second United States Microgravity Laboratory mission in October/November 1995.
Tognini, who spent 14 days on the Mir space station in 1992, will be making his first Shuttle flight on STS-93. During the five-day mission, the crew will deploy the Advanced X-ray Astrophysics Facility Imaging System (AXAF), which will conduct comprehensive studies of the universe. AXAF will be the most advanced X-ray telescope ever flown.
When scientists begin using AXAF next year, they will finally be able to unlock the secrets of some of the most distant, powerful and violent objects known to exist in the universe. They will study such exotic phenomena as exploding stars called supernovae, strange powerful objects called quasars, and mysterious black holes which are so massive that everything near them is pulled inside causing an explosion of X-rays that AXAF can study.
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"Background on Europa Data From The Galileo Mission To Jupiter (2 March, 1998)",704,0,0,0
At today's press briefing, Brown and NASA scientists will show the most detailed images ever seen of the Jupiter moon Europa. Recently transmitted from the Galileo spacecraft, the images provide three key pieces of evidence supporting the idea that water may lurk beneath Europa's surface.
\BBackground \b
Water or ice? Liquid or slushy or frozen solid? Ever since the Voyager spacecraft missions flew through the Jupiter system in 1979, planetary scientists have wondered about the layer of ice surrounding the planet. Europa's blindingly bright ice surface makes it one of the brightest objects in our solar system. Recent Galileo spacecraft images have provided evidence that Europa had a liquid ocean underneath the frozen crust sometime in its history, but it is not clear if this ocean still exists. Of the various explanations proposed by scientists, most scenarios of Europa's evolution have the water layer freezing solid earlier in its history.
The moon's surface is -260░ F, which could freeze an ocean over several million years. But some scientists are beginning to think that the warming caused by a tidal tug of war with Jupiter and neighboring moons could be keeping large parts of the ocean liquid.
\BKey images \b
New stereo and very high resolution images of Europa just transmitted to Earth from the Galileo Europa Mission fly-by in December 1997 may help support the theory that water or slush may slosh beneath Europa's frozen crust. Detailed enough to see a truck-sized object on the surface, the new images are hundreds of times higher resolution than the best Voyager images and three to 20 times higher than earlier Galileo pictures.
The Brown and NASA scientists point to three key pieces of evidence from the detailed images:
\B╖\b The subdued topography of the young impact crater Pwyll, whose rays cover a significant part of the surface of Europa;
\B╖\b Large plates of ice and iceberg-like structures called "chaos terrain"; and
\B╖\b Gaps between plates of ice known as "wedges" where new crust appears to have formed recently.
\BOceans and life\b
"Together, the craters, chaos and wedges support the hypothesis that in Europa's most recent history, liquid or at least partially liquid water existed at shallow depths below the surface of Europa in several different places," says James Head, Brown University professor of geological sciences. "These and other data lend support to the hypothesis that Europa is warm and active today and potentially characterized by a global subsurface water layer or ocean. Europa, like Mars and the Saturn moon Titan, is a laboratory for the study of conditions that might have led to the formation and evolution of life. The combination of interior heat, liquid water, and infall of organic material from comets and meteorites means that Europa has the key ingredients for life, and it represents an exciting environment that is worthy of further detailed exploration."
\BCrater evidence\b
Rays and debris from the impact that formed Pwyll Crater radiate over a large part of the moon's surface. Galileo took pictures of the impact crater from two perspectives to determine the three-dimensional shape of the crater. Colleagues at the DLR (German Aerospace Research Establishment) converted these images into a colored map showing the depth of the crater and the height of its peaks.
Unlike most young, deep impact craters, the floor of Pwyll is at the same level as the exterior, says Brown graduate student Geoffrey Collins. The central peaks of the crater are more than 2,000 feet high - four times higher than the Washington Monument - and higher than the crater rim. This means that this young crater was warm and weak and collapsed during or very shortly after the meteorite impact, in contrast to craters formed in cold, stiff material. Debris that flowed from the violent impact is dark, suggesting excavation of different material from below the surface. All this suggests that water just beneath the surface was warm enough to be slushy in the moon's recent history.
\BChaos evidence\b
The new images from Galileo help answer some questions about other areas of Europa that are littered with fractured and rotated blocks of crust the size of several city blocks (dubbed chaos terrain). These fractured ice chunks appeared to be either sliding on soft glacier-like ice below the surface or floating like icebergs in a more fluid material. The new images show that the material between the cracked and separated plates of crust is rough and swirly, says Robert Pappalardo, a postdoctoral research scientist at Brown. The pieces are immersed in what appears to be a slush that is now frozen solid. The very low temperatures at the surface of Europa (-260░ F) mean that any water exposed at the surface would freeze immediately and might create this kind of texture. The rough chaos terrain, as well as the movement and rotation of the blocks, suggest that the crust was at least partially liquid at shallow depths.
\BWedges Evidence\b
Other images are helping unravel more mysteries. Pieces of the moon's glaringly white crust are separated by wedged-shaped pieces of darker, newer crust, welling up from below, freezing and cracking. The separated pieces of white crust would fit back together like a jig-saw puzzle, suggesting that plate tectonic-like activity might be occurring on Europa to form the wedges. Composed of a set of narrow linear ridges and parallel grooves, the dark wedge has many similarities to new crust formed at mid-ocean ridges on the Earth's sea floor, says Brown graduate student Louise Prockter, who has studied high-resolution sonar images of the Mid-Atlantic Ridge and has visited the Pacific Ocean floor in the research submersible vehicle Alvin. Like Earth, new crust seems to be welling up, separating, and replacing older crust. On Europa, the molten material solidifying on the surface was likely slushy ice or liquid water.
\BNext Step\b
To confirm the existence of such a layer, determine its depth and investigate its nature and global extent, further observations are planned for the Galileo Europa Mission, and other experiments are planned for a Europa Orbiter Mission to be launched in 2003, says Michael J. S. Belton of the National Optical Astronomy Observatory in Tucson, Ariz., and team leader for the solid state imaging system.
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"Detailed Images From Europa Point To Slush Below Surface (2 March, 1998)",705,0,0,0
The latest, most detailed pictures of the Jupiter moon Europa lend more support to the theory that slush or even liquid water lurks beneath the moon's surface. Those pictures were presented and discussed by scientists from Brown University and NASA during a press briefing today on the Brown campus.
Slightly smaller than Earth's moon but many times brighter, Europa's icy surface has intrigued scientists ever since the Voyager spacecraft missions flew through the Jupiter system in 1979. At -260░ F, the moon's surface temperature could deep-freeze an ocean over several million years, but some scientists are beginning to think that warmth from a tidal tug of war with Jupiter and neighboring moons could be keeping large parts of Europa's ocean liquid.
The latest images released today were taken in December 1997 by the Galileo spacecraft and just received on Earth. The new images provide three key pieces of evidence showing that Europa may be slushy just beneath the icy crust and possibly even warmer at greater depths. The evidence includes a strangely shallow impact crater, chunky textured surfaces like icebergs, and gaps where new icy crust seems to have formed between continent-sized plates of ice.
Some of the new images focus on the shallow center of the impact crater known as Pwyll. Impact rays and debris scattered over a large part of the moon show that a meteorite slammed into Europa relatively recently, about 10-100 million years ago. The darker debris around the crater suggests the impact excavated deeply buried material. But the crater's shallow basin and high set of mountain peaks may mean that subsurface ice was warm enough to collapse and fill in the deep hole, says Brown graduate student Geoffrey Collins, a member of the Galileo research team.
A subsurface ocean warm enough to be slushy also may explain the origins of an area littered with fractured and rotated blocks of crust the size of several city blocks, called "chaos" terrain. The new images show rough and swirly material between the fractured chunks, which may have been suspended in slush that froze at the very low surface temperatures, says Robert Pappalardo, a postdoctoral research scientist at Brown and a member of the Galileo research team.
On a larger scale, large plates of ice seem to be sliding over a warm interior on Europa, much like Earth's continental plates move around on our planet's partly molten interior.
The new images of Europa show that the darker wedge-shaped gaps between the plates of ice have many similarities to new crust formed at mid-ocean ridges on the Earth's sea floor, says Brown graduate student Louise Prockter, a member of the Galileo research team who has studied high-resolution sonar images of the Mid-Atlantic Ridge and has visited the Pacific Ocean floor in the research submersible vehicle Alvin. The new crust welling up between the separating plates on Europa was likely initially slushy ice or possibly liquid water that has frozen and fractured, Prockter says.
"Together, the evidence supports the hypothesis that in Europa's most recent history, liquid or at least partially liquid water existed at shallow depths below the surface of Europa in several different places," says James Head, Brown University professor of geological sciences and a group leader of the Galileo research team.
"The combination of interior heat, liquid water, and infall of organic material from comets and meteorites means that Europa has the key ingredients for life," Head says. "Europa, like Mars and the Saturn moon Titan, is a laboratory for the study of conditions that might have led to the formation of life in the solar system."
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"NASA Radar Reveals Hidden Remains At Ancient Angkor (12 February, 1998)",706,0,0,0
New evidence of a prehistoric civilization and remnants of ancient temples in Angkor, Cambodia, have been discovered by researchers using highly detailed maps produced with data from an airborne imaging radar instrument created by NASA. Experts say the findings, made possible by the Airborne Synthetic Aperture Radar (AIRSAR) developed by NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, may revolutionize the way archaeologists view the ancient city's development.
Angkor is a vast complex of some 1,000 temples covering about 100 square miles of northern Cambodia. Little is known of the prehistoric occupation of this fertile flood plain, but at its height the city housed an estimated population of one million people. The famous temples were built from the eighth to thirteenth century AD and were accompanied by a massive hydrological system of reservoirs and canals.
Today, much of the civilization of Angkor is hidden beneath a dense forest canopy and is inaccessible due to poor roads, land mines and political instability. "The radar data have enabled us to detect a distribution of circular 'prehistoric' mounds and undocumented temples far to the northwest of Angkor," said Dr. Elizabeth Moore, Head of the Art and Archaeology Department at the School of Oriental and African Studies at the University of London. "The site's topography is highlighted by the radar, focusing our attention on previously neglected features, some at the very heart of the city.
"The radar maps not only bring into question traditional concepts of the urban evolution of Angkor, but reveal evidence of temples and earlier civilization either absent or incorrect on modern topographic maps and in early twentieth century archaeological reports," she said. "The radar images make apparent many features that are not readily identifiable on the ground," said Dr. Anthony Freeman, a radar scientist at JPL who has collaborated with Moore for the past three years studying the use of radar on the Angkor site.
"We can see differences in vegetation structure and some features that are obscured by vegetation cover." In December 1997, Moore surveyed a small mound on the perimeter of the famous 12th century AD temple, Angkor Wat, that Freeman had first noticed in the radar image. "Previous archaeological accounts from 1904 and 1911 note only two temples and make no mention of the distinct circular form of the mound. We found four to six temple remains, including pre-Angkorean structures," Moore said. "This suggests occupation of the 12th century site some 300 years earlier, radically changing accepted chronologies of Angkor." Angkor's beauty is seen in its temples, but the greatness of the Khmer city lies in the multitude of water-related constructions, according to Moore.
The Khmer kings nominally dedicated temples to Hindu and Buddhist deities, but the underlying significance was veneration of ancestral spirits, ensuring fertility of the land. Management of water was essential, both for control during the monsoon rains and conservation during the dry season and involved the construction of moats, dikes, canals, tanks, and reservoirs. The largest of these reservoirs, dated to the 12th century AD, is five miles long and its function remains a matter of archaeological debate.
"These new detailed topographic maps have shown us many more hydrological features and highlighted how they function in the rituals and daily life of the Khmer people," Moore explained. "Using a technique known as radar interferometry, which combines two images to create a three-dimensional topographic map, we can construct a map of the area surrounding Angkor that is more accurate than most maps we have of the United States," said Dr. Scott Hensley, a radar engineer at JPL.
"This map lets us see both natural and human-made water management features at the site with great clarity." "Angkor is situated on the edge of the Tonle Sap lake, a unique body of water that doubles in size during the rainy season. These maps give us new insights into the human impact on this ecosystem, from the ancient Khmer to the present day, and are of importance in the study of our changing Earth," Freeman continued.
The Angkor radar images were taken in late 1996 as part of the AIRSAR Pacific Rim Deployment and were a follow-up to the 1994 study of Angkor with data collected by the Spaceborne Imaging Radar-C/X-band Synthetic Aperture Radar (SIR-C/X-SAR) that flew on NASA's Space Shuttle Endeavour. Like SIR-C/X-SAR, AIRSAR transmits and receives three radar frequencies in both horizontal and vertical polarizations.
While both systems use C-band and L-band wavelengths, AIRSAR has the added benefit of P-band, a longer wavelength that can penetrate below the forest canopy. In addition, AIRSAR can be flown in a mode called TOPSAR that allows it to measure topography and create three-dimensional images of the surface.
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"Scientists View First Close-Ups of Strange, Layered Polar Terrain (11 February, 1998)",707,0,0,0
Swirling bands of eroded, layered rock, reminiscent of the edges of Alaskan ice sheets, and an array of light and dark mottled patterns blanket the frigid floor of Mars' south pole, where NASA's newly named Mars Polar Lander will touch down in late 1999.
The new images of the landing zone for the Mars Polar Lander, taken by the camera aboard NASA's Mars Global Surveyor, confirm that this strange, layered terrain in the south polar region represents a dramatic departure from the now-familiar Martian landscapes observed by the Viking landers and Mars Pathfinder. In December 1999, the next lander in a steady series begun by Pathfinder will set down in this uncharted territory to dig for traces of frozen, subsurface water.
"Despite ground fog that obscures part of the surface in these images, we can see much more surface detail than we've ever seen before, which suggests that the 75-degree south latitude landing zone is quite a bit more rugged and geologically diverse than we had previously thought," said Dr. Michael Malin of Malin Space Science Systems, Inc., San Diego, CA. Malin is principal investigator for the Global Surveyor camera and the cameras on the 1998 missions, the Mars Polar Lander and its newly named partner, the Mars Climate Orbiter.
In the current images from Mars Global Surveyor, obtained during an aerobraking orbit from about 1,700 miles above the planet's surface, objects about 48 feet across can be resolved. Once the spacecraft reaches its final mapping orbit early next year, at an average of 234 miles above the surface, the camera will be able to resolve ground features as small as seven to nine feet across. This greater clarity will enable views of objects as small as boulders or as subtle as sand dunes.
Over the next year, the Global Surveyor images will be used in concert with other spacecraft data such as that obtained by its thermal emission spectrometer to better characterize the geology of the Martian south pole. After Global Surveyor has reached its mapping orbit, data from the spacecraft's laser altimeter, which measures the height and roughness of Martian surface features, will be combined with the imaging data to aid the final choice of landing sites.
"We have a wonderful opportunity in the next year to study this region with data from Mars Global Surveyor, which underscores the true advantage of conducting a continuing program of Mars exploration," said Dr. John McNamee, Mars Surveyor '98 project manager at NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA.
"We will be able to characterize the geology of the whole region and find the best spot to land, one that presents a balance between lander safety and scientific interest. This process does not have to be finalized until June 1999, five months after the lander has been launched and six months before it lands."
The spacecraft are currently being prepared for transfer to the Lockheed Martin environmental test chambers to ensure that they can survive and operate in the extreme conditions at Mars. At the completion of this testing, the spacecraft will be flown separately to NASA's Kennedy Space Center, FL, for integration with their launch vehicles.
The 1998 Mars lander and orbiter missions are designed to learn more about the history of Mars' climate and the behavior of related Martian volatiles, such as water vapor and ground ice. The orbiter, scheduled for launch on Dec. 10, will conduct a two-year primary mission to profile the Martian atmosphere and map its surface. The lander, scheduled for liftoff on Jan. 3, 1999, will carry out a three-month mission to search for traces of subsurface water in this frozen, layered terrain and any evidence of a physical record of climate change.
To meet these scientific objectives, the orbiter will carry a rebuilt version of the Pressure Modulated Infrared Radiometer (PMIRR) that was lost with Mars Observer in 1993. This atmospheric sounder will observe the global distribution and time variation of temperature, dust, water vapor and condensates in the Martian atmosphere. PMIRR is a collaboration between JPL, Oxford University and Russia's Space Research Institute.
Like Mars Global Surveyor, the Mars Climate Orbiter carries a dual camera system, contained in an amazingly compact package about the size of a pair of binoculars. The Mars color imager's one-pound wide-angle camera will return daily low-resolution global views of the planet's atmosphere and surface, while its medium-angle camera will provide higher resolution (30 feet per pixel) images.
The medium-angle camera will build global and regional maps of Mars in multiple colors over the course of the mission. These maps will be used to characterize surface properties and changes in the distribution of dust.
The 1998 lander carries three scientific packages: the Mars descent imager, provided by Malin, which will view the landing site at increasingly higher resolution as the lander descends to the surface of Mars; the atmospheric lidar experiment, provided by the Russia space institute, which will monitor the presence and height of atmospheric hazes, coupled with a miniature microphone furnished by The Planetary Society, Pasadena, CA, to record the sounds of Mars; and the Mars Volatile and Climate Surveyor (MVACS) package, led by principal investigator Dr. David Paige of the University of California, Los Angeles.
MVACS includes a surface stereo imager based on the Mars Pathfinder camera, both built at the University of Arizona; a meteorology package, built at JPL; a robotic arm, also built at JPL, to acquire soil samples and close-up images of the surface and subsurface; and the thermal and evolved gas analysis experiment, built at the University of Arizona. JPL will oversee mission operations with the spacecraft team at Lockheed Martin Astronautics and the instrument teams located at their home institutions during the lander and orbiter missions.
"MVACS and the other science experiments are tailor-made for the exploration of Mars' south pole," said Dr. Richard Zurek, project scientist at JPL. "The robotic arm, which is reminiscent of the Viking arm and scoop that were used to carry out biology experiments in the mid-1970s, is, in fact, much more versatile. It can reach farther out, dig up to three feet below the surface and then place soil samples in a miniature oven, called the evolved gas analysis experiment, where the samples are 'cooked' and analyzed for chemical and gas content."
Piggybacking on the Mars Polar Lander are two small 4.5-pound microprobes provided by NASA's New Millennium technology validation program. Deployed before landing, they will penetrate and embed themselves beneath the Martian surface to study subsurface materials.
The Mars Polar Lander and the Mars Climate Orbiter are the second set of launches in a long-term NASA program of Mars exploration known as the Mars Surveyor Program. The missions are managed by JPL for NASA's Office of Space Science, Washington, DC. Lockheed Martin Astronautics, Denver, CO, is NASA's industry partner in the mission.
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"Shock Wave Sheds New Light on Fading Supernova (10 February, 1998)",708,0,0,0
NASA's Hubble Space Telescope is giving astronomers a ringside seat to a never before seen titanic collision of an onrushing stellar shock wave with an eerie glowing gas ring encircling a nearby stellar explosion, called supernova 1987A. Though the star's self-destruction was first seen nearly 11 years ago on Feb. 23, 1987, astronomers are just now beginning to witness its tidal wave of energy reaching the "shoreline" of the immense light-year wide ring. Shocked by the 40-million mile per hour sledgehammer blow, a 100-billion mile diameter knot of gas in a piece of the ring has already begun to "light up," as its temperature surges from a few thousand degrees to a million degrees Fahrenheit.
"We are beginning to see the signature of the collision, the hammer hitting the bell. This event will allow us to validate ideas we have built up over the past ten years of observation," says Robert Kirshner of the Harvard-Smithsonian Center for Astrophysics (CfA) in Cambridge, MA. "By lighting up the ring, the supernova is exposing its own past." Astronomers predict it's only a matter of years before the complete ring becomes ablaze with light as it absorbs the full force of the crash. Illuminating the surrounding space like a flashlight in a smoky room, the glowing ring is expected to literally shed a brilliant new light on many unanswered mysteries of the supernova: What was the progenitor star? Was it a single star or binary system? Are a pair of bizarre outer rings attached to an invisible envelope of gas connecting the entire system?
"We have a unique opportunity to probe structure around the supernova and uncover new clues to the final years of the progenitor star before it exploded," adds Richard McCray of the University of Colorado in Boulder, CO. "The initial supernova flash only lit up a small part of the gas that surrounds the supernova. Most of it is still invisible. But the light from the crash will give us a chance to see this invisible matter for the first time, and then perhaps we can unravel the mystery of the outer rings."
Though scientists will never solve the paradox of what happens when an irresistible force meets an immovable object, the supernova collision is the closest real-world example yet. "This supernova gives us an unprecedented opportunity to directly witness new physics of shock interactions," says McCray. "Though astronomers have measured shock effects from the expanding debris of many supernovae which are centuries-old, their impact velocities are at least ten times slower than the ones we see today in supernova 1987A."
The ring was formed 20,000 years before the star exploded. One theory is that it resulted from stellar material flung off into space as the progenitor star devoured a stellar companion. The ring's presence was given away when it was heated by the intense burst of light from the 1987 explosion. The ring has been slowly fading ever since then as the gas cools. Several years ago radio waves and X-rays were detected as the fastest moving explosion debris slammed into cooler invisible gas inside the ring.
In spring of 1997 the newly installed Space Telescope Imaging Spectrograph (STIS) first measured the speed of the supernova debris pushing along the shock wave. "The STIS lets you see the invisible stuff," says George Sonneborn of Goddard Space Flight Center in Greenbelt, MD. "We see the shock happening everywhere around the ring." In July, Hubble Wide Field and Planetary Camera-2 images taken by Robert Kirshner and co- investigators showed that a compact region near on the ring lit up like a sparking diamond set in an engagement ring.
Supernova 1987A is the brightest stellar explosion seen since Johannes Kepler observed a supernova in the year 1604. It is located about 167,000 light-years from Earth in the Large Magellanic Cloud. The Space Telescope Science Institute is operated by the Association of Universities for Research in Astronomy, Inc., for NASA, under contract with the Goddard Space Flight Center, Greenbelt, MD. The Hubble Space Telescope is a project of international cooperation between NASA and the European Space Agency.
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"Space Station Agreements To Be Signed In Washington (29 January, 1998)",709,0,0,0
Today marks an important milestone for the International Space Station as senior government officials from 15 countries meet in Washington to sign agreements to establish the framework for cooperation among the partners on the design, development, operation and utilization of the Space Station. Acting Secretary of State Strobe Talbott will sign the 1998 Intergovernmental Agreement on Space Station Cooperation, along with representatives of Russia, Japan, Canada and participating countries of the European Space Agency (Belgium, Denmark, France, Germany, Italy, the Netherlands, Norway, Spain, Sweden, Switzerland and the United Kingdom).
The signing will be held at the U.S. State Department's Dean Acheson auditorium at 4 p.m. EST. Three bilateral memoranda of understanding also will be signed by NASA Administrator Daniel S. Goldin separately with his counterparts: Russian Space Agency General Director Yuri Koptev, ESA Director General Antonio Rodota and Canadian Space Agency President William (Mac) Evans.
The memorandum of understanding between NASA and the government of Japan will be signed at a later date. Today's new agreements supersede previous Space Station agreements among the U.S., Europe, Japan and Canada signed in 1988. These new agreements reflect changes to the Space Station program resulting from significant Russian participation in the program and program design changes undertaken by the original partnership in 1993.
Led by the U.S., the International Space Station will be the largest, most complex international cooperative science and engineering program ever attempted. Taking advantage of the technical expertise from participating countries, the International Space Station will bring together scientists, engineers and researchers from around the globe to assemble a premier research facility in orbit. Beginning in June 1998, with the launch of the first Space Station element, the partnership will ultimately assemble more than 100 components in low Earth orbit over the next five years, using approximately 45 assembly flights.
When completed, the Station will provide access for researchers around the world to permanent, state-of-the-art laboratories in weightlessness. As currently envisioned, the International Space Station will support a crew of up to seven and include five complete pressurized laboratories and attached external sites for research. The Station will provide a focal point for space operations among the partners well into the next century, and will serve as a stepping-stone for human exploration of the solar system. On the football-field-sized station, permanent crews will perform long-duration research in a variety of scientific disciplines advancing the world's understanding of life sciences, earth sciences and materials processing, while fostering commercial research activities in space.
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"Space Pioneers Recall First U.S. Satellite Launch Upon 40th Anniversary (28 January, 1998)",710,0,0,0
Forty years ago this week, a team of scientists and engineers successfully launched Explorer 1, the first U.S. satellite to orbit the Earth. This historic accomplishment marked the nation's debut in the Cold War-era space race and set the stage for the establishment of the civilian space agency that would become NASA.
NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, was still operated as a research laboratory for the U.S. Army when it was selected in November 1957 to develop the first U.S. satellite, including its science package, its communications system, and the high-speed upper stages for the Army's Redstone rocket that would guide the tiny, 20-pound Explorer 1 into the great unknown.
JPL and the Army completed the assignment and successfully launched the satellite in less than three months. JPL and the Army Ballistic Missile Agency, based in Huntsville, AL, joined in firing the satellite toward space from the missile test center at Cape Canaveral, FL, on Jan. 31, 1958.
The scientific experiment onboard, a cosmic ray detector built by Dr. James Van Allen of the University of Iowa, soon returned one of the most important findings of the space program: the discovery of what are now known as the Van Allen Radiation Belts around the Earth. Explorer 1 went on to operate for three months.
Following the Soviet Union's launch of Sputnik on Oct. 4, 1957, "there was a lot of pressure to get a satellite in orbit as quickly as possible," said Dr. William Pickering, then JPL's director and the orchestrator of the Explorer 1 effort at JPL.
The intensive effort was accomplished by a team of experts from U.S. academia and the military, along with top World War II German rocket scientists such as Dr. Wernher von Braun, who emigrated to the United States in the post-war years to help lead the development of American rocket capabilities.
A globally linked telecommunications system developed by JPL tracked Explorer 1 and received its scientific data as it circled Earth. Amateur radio operators around the world were invited to listen in on Explorer 1's radio communications, including one key amateur radio shack operated largely by JPL ham radio operators at the Los Angeles County Sheriff's substation in Temple City, near JPL.
The most difficult technical challenge, said Pickering, "was getting the three rocket stages to work consistently, to get it all to go in the right direction, with no guidance system." Considering the telecommunications and computing capability of the Explorer 1 era versus that available for last summer's Mars Pathfinder mission, Pickering said, "it's astonishing to think what has happened over 40 years."
Van Allen, still an active planetary and space physics researcher, recalled that, the morning after the historic Explorer 1 launch, "a big press conference had been called at the Great Hall of the National Academy of Sciences in Washington, DC, and although it was 1:30 in the morning, there was still a huge crowd of reporters waiting around."
Donna Shirley, Mars Exploration program manager at JPL, was in high school when the news hit that Explorer 1 had been launched. "It was a terrific emotional moment," she recalled. "It seemed like a scary thing that the Soviet Union was so powerful that they could launch Sputnik.
When Explorer went up, it was, 'Rah, rah, our team!'" she said. "It seemed to be framed in 'us versus them' rather than focused on the real technical and scientific achievement. But the dawn of the Space Age affected my life a lot."I don't think the 'right stuff' to work in the space program has really changed all that much" since the days of Explorer 1, said Shirley. "You don't have cigar-smoking guys with slide rules anymore, but I think the 'right stuff' is still the same: dedication and competence."
In late 1958, JPL was reassigned from the U.S. Army to NASA when the civilian space agency was created, and has helped lead the world's exploration of space with robotic spacecraft since then. Operated as a division of the California Institute of Technology, JPL has sent spacecraft to all of the known planets except Pluto, and this year will launch major astronomy and planetary exploration missions to comets, asteroids and Mars, along with many Earth-observing efforts.
As the size of NASA's space missions takes advantage of miniaturized electronics to shrink to fit the new "faster, better, cheaper" mold, some complete space science instrument packages are about the size of that on tiny Explorer 1, Shirley said.
"Miniaturization is allowing us to shrink down the brains of our spacecraft but still allow us to do more with them than we used to. The challenge now is to shrink the rest of the spacecraft down."
Considering the future of space science, Van Allen observed that "there is no shortage of great ideas on what we'd like to do. 'Faster, better, cheaper' is NASA's mantra, and the recent successful launch of the Lunar Prospector spacecraft is the best example of that. But the Hubble Space Telescope is a good example of big projects that will continue to be conducted. I think we have a very bright future in space science in all areas. There is good public support," he said. "There is virtually no limit to what can be investigated in interplanetary science and astronomy."
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"Fast-Spinning Pulsar Discovery Provides Evolutionary Link (26 January, 1998)",711,0,0,0
Scientists have announced the discovery of a superdense star spinning at more than 60 times per second, and calculate it could have been spinning as fast as 150 times per second or more when it formed some 4,000 years ago. Most astronomers had not previously believed this class of star, called a pulsar, could form with such a rapid spin. "This shatters the glass ceiling," said astrophysicist John Middleditch of the U.S. Department of Energy's Los Alamos National Laboratory in New Mexico.
"This is the fastest high-energy pulsar of its type we know about." "The pulsar is spinning twice as fast as any young pulsar that we have seen before," adds Dr. Frank Marshall of NASA's Goddard Space Flight Center, Greenbelt, MD, who led the team making the discovery. "To put it in perspective, this pulsar is spinning more than 6 million times as rapidly as the Earth." The newly discovered pulsar establishes a link between fast-spinning pulsars with relatively weak magnetic fields and slow-spinning ones with strong fields, suggesting there may be a natural continuum between the two known types.
The pulsar was found by Dr. Marshall and his colleagues Drs. William Zhang and Eric Gotthelf of Goddard, and Middleditch, by examining X-ray emissions recorded by NASA's Rossi X-ray Timing Explorer spacecraft in 1996, and confirmed with observations using the joint Japanese/U.S. Advanced Satellite for Cosmology and Astrophysics (ASCA) spacecraft. Pulsars get their name because their emissions appear to turn on and off, or pulse, very rapidly.
Astronomers believe the stars channel some of their energy into a beam of radiation, and as the star spins the beam sweeps through space like a lighthouse beacon. By counting how rapidly the beam flashes at Earth, scientists can calculate a pulsar's rate of spin. When a star explodes as a supernova it leaves behind a lingering core about 15 miles across but packed with as much matter as in Earth's Sun. The star is so dense that neutrons are the only form of matter that exist in the star, thus earning the name "neutron star." Those whose rapid spin can be observed are called "pulsars."
The team identified the pulsar as most likely being associated with the remnant of a supernova (catalogued N157B by astronomers) that exploded in the Large Magellanic Cloud, a companion to our Milky Way galaxy, about 4,000 years ago. (The age estimate comes from other X-ray and visible observations of the spreading, tattered gas cloud from the supernova blast and is in agreement with that predicted by theoretical models.) Data from both the Rossi and ASCA satellites were used to calculate the rate at which the pulsar's spin is slowing, which in turn provides an estimate of its age: 5,000 years old, a close match to the age estimate for the supernova remnant.
The other well-known high energy pulsar, in the Crab Nebula, spins just under 30 times per second, and is generally thought to have been spinning at only 60 times a second at its birth in 1054 AD. Since the Crab pulsar's discovery in 1968, astronomers have spotted pulsars spinning as fast as hundreds of times per second. These so-called "millisecond pulsars" (because their spin periods are only a few thousandths of a second) have magnetic fields a thousand times weaker than the Crab pulsar. Most astronomers believe that the weak-field, millisecond pulsars were born with a slow spin and were "spun up" after sucking in gaseous material from an orbiting stellar companion, but astronomers have not located enough suitable binary star systems to account for the large numbers of millisecond pulsars being discovered.
The pulsar found in N157B, whose magnetic field is only a few times weaker than the Crab pulsar's, suggests an evolutionary link between the strong-field, slower-spinning energetic pulsars and the weak-field millisecond pulsars. Its discovery confirms a prediction published by Gotthelf and Dr. Q. Daniel Wang of Northwestern University. "This is a fantastic confirmation of our hypotheses; that the central source of X-ray light from N157B is a fast pulsar associated with a supernova remnant, like that seen in the Crab nebula," commented Gotthelf.
"Now, clearly, it seems that the weaker the magnetic field, the faster the pulsar will spin at birth -- possibly all the way down to one- or two-millisecond periods (corresponding to spin rates of 1,000 to 500 times per second) for fields of the strength measured for the weak-field pulsars," Middleditch said. Astronomers continue to search for a pulsar at the heart of SN1987A, a supernova that appeared in the southern skies Feb. 23, 1987.
Most astronomers who study this supernova expect that a rapidly spinning, weak-field pulsar should eventually reveal itself for observation, which would provide another link in theories of how fast pulsars are born. Marshall and his team encouraged other researchers to study N157B at other regions of the spectrum to see if its pulsations are observable there, too. The team announced their discovery last week through a circular distributed by the International Astronomical Union.
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"Additional Experiments Selected For Mars 2001 Missions (22 January, 1998)",712,0,0,0
NASA has selected additional instruments for the Mars Surveyor 2001 missions, which will study Mars' environment. The Mars Surveyor 2001 missions will follow two other robotic Mars missions to be launched in late 1998 and early 1999. All are part of NASA's long-term, systematic exploration of Mars in which two missions are launched to the planet approximately every 26 months.
"In a sense, these missions allow virtual presence by humans and provide precursor data and subsequent infrastructure for possible human missions in the 21st century," said Arnauld Nicogossian, Associate Administrator of NASA's Office of Life and Microgravity Sciences and Applications. "By adding capability to missions already planned, this near term effort will result in cost effective, tangible progress in carrying out the Human Exploration and Development of Space strategy and contribute to the Origins program of NASA's Office of Space Science."
NASA's Office of Life and Microgravity Sciences and Applications has selected the following investigations for the Mars 2001 Orbiter, due for launch in March of that year, and the Mars 2001 Lander/Rover, due for launch in April 2001:
ò The Martian Radiation Environment Experiment will characterize the radiation environment in the orbit and on the surface of Mars simultaneously. This experiment will consist of radiation spectrometers on both the Mars 2001 Orbiter and on the Mars 2001 Lander. Dr. Guatam Badhwar from NASA's Johnson Space Center, Houston, TX, is the principal investigator.
ò The Mars Environmental Compatibility Assessment will characterize Martian dust and soil to identify potential undesirable and harmful interactions with human explorers and associated hardware, and to evaluate properties of the soil related to its use as a construction material.
Dr. Thomas Meloy from West Virginia State University is the principal investigator. A team consisting of NASA's Jet Propulsion Laboratory (JPL), Pasadena, CA, and Lockheed Martin Astronautics, Denver, CO, will develop the missions, led by JPL. The radiation and dust investigations were selected from 39 proposals submitted to NASA in August 1997.
Both of the 2001 missions are part of an ongoing NASA series of robotic Mars exploration spacecraft that began with the launches of the Mars Global Surveyor in November 1996. The 2001 missions represent the first step in a NASA initiative to integrate the requirements for Space Science and the Human Exploration and Development of Space program into a single robotic exploration program.
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"Earth Swingby Puts NEAR Spacecraft on Final Approach to Eros (20 January, 1998)",713,0,0,0
NASA's Near Earth Asteroid Rendezvous (NEAR) spacecraft, built by The Johns Hopkins University Applied Physics Laboratory (APL) in Laurel, MD, will become the first interplanetary spacecraft that can possibly be seen with the naked eye when it swings by Earth Jan. 22-23. The spacecraft's solar panels will reflect the Sun's rays onto the Earth in a greeting as it flies by for an adjustment of its trajectory to correctly align the spacecraft for a rendezvous with asteroid 433 Eros, its mission target.
Launched Feb. 17, 1996, NEAR completed a flyby of the asteroid Mathilde in June 1997 and is now on its way back to Earth. Late Thursday, Jan. 22, the spacecraft will approach Earth over the Pacific Ocean traveling at about 20,000 mph. Because the United States will be in darkness as NEAR approaches, if there is no cloud cover, several geographic areas will be able to see the Sun reflecting off the spacecraft's solar panels, which will act as large mirrors. These sunglints will be visible on the East Coast, Friday, Jan. 23, at about 1:30 a.m. EST and the West Coast at about 1:45 a.m. EST (Thursday, 10:45 p.m. PST).
The spacecraft then swings around the Aleutian Islands and over Siberia before reaching its closest point to Earth, about 336 miles above Ahvaz in southwest Iran, Friday, Jan. 23, at 11:23 a.m. local time (2:23 a.m. EST), traveling at about 29,000 mph, its fastest speed for the swingby. Although NEAR will be close to Earth at this time, daylight may obscure its image. The spacecraft then swings over Africa and on to Antarctica before pulling away from the Earth at a speed of about 15,000 mph.
The swingby will have changed NEAR's trajectory to approximately 11 degrees south of the Earth's ecliptic plane, the orbital path the Earth takes as it circles the Sun, and put the spacecraft on target for its Jan. 10, 1999, rendezvous with Eros. NEAR scientists and engineers are using the swingby as an opportunity to test performance and calibration of the spacecraft's six instruments and to practice coordinated multi- instrument observations of the type that will be used at Eros. The Multispectral Imager, a visible light camera that will help determine the physical characteristics of Eros, and the NEAR- Infrared Spectrograph, used to study surface minerals, will be calibrated by comparing their readings of geological features with proven measurements of the same areas.
These instruments will also be used to take images of the Earth along the spacecraft's path. NEAR's Magnetometer will be calibrated by comparing swingby data with known measurements of the Earth's magnetic field. Other activities during the swingby will include using the X-Ray/Gamma-Ray Spectrometer to observe celestial gamma ray bursts and to collect data on gamma ray and X-ray backgrounds. These data are needed so scientists can better remove background impurities from the measurements to be made at Eros.
NEAR is expected to capture its first images of Eros, a 25- mile-long near-Earth asteroid, a few months prior to the 100th anniversary of the asteroid's discovery on Aug. 13, 1898. After reaching Eros, the spacecraft will start its orbit about 600 miles above the asteroid's surface, descending to 200 miles by February and coming as close as 10 miles during its yearlong study. Scientists will thoroughly map Eros and will examine its surface composition and physical properties.
On Feb. 6, 2000, the mission is expected to end with a controlled descent onto the asteroid, sending dozens of high-resolution pictures as the spacecraft closes in on Eros. The NEAR mission will be the first close-up study of an asteroid.
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"First Station Element to be Shipped to Russian Launch Site (16 January, 1998)",714,0,0,0
The International Space Station will complete a major milestone toward its first launch as the first station piece, a U.S.-funded and Russian-built control module, is shipped from a Moscow factory next week to its Russian Space Agency launch site in Baikonur, Kazahkstan. In advance of the shipment of the control module, formerly called the Functional Cargo Block and designated by the Russian acronym FGB, a rollout ceremony and press conference will be held at the Khrunichev State Research and Production Center in Moscow at 11 a.m. Moscow time on Saturday, Jan. 17.
Highlights of the rollout ceremony will be broadcast, tape-delayed, on NASA Television at 3 p.m. EST Saturday, with a repeat airing at 6 p.m. EST. The actual shipping of the control module is scheduled to begin on Thursday, Jan. 22. The 20-ton module is targeted for a late June launch to begin the five-year, 45-flight orbital assembly of the new space station. It will be launched on a Russian Proton rocket from the Baikonur Cosmodrome in Kazahkstan. The control module was built by the Khrunichev factory, under contract to The Boeing Company, the prime contractor to NASA for the International Space Station. It will depart Khrunichev via a special rail car late next week to begin the 1,200-mile, five- day train journey to Baikonur, where it will begin five months of launch preparations and final testing.
"When the control module arrives at Baikonur, all of the elements for our first two launches will be undergoing final launch processing," International Space Station program manager Randy Brinkley said. "The year of the International Space Station is 1998. This is something that all of us have looked forward to for a very long time. We have a lot of exciting and challenging activities ahead as we begin our assembly in orbit.
The incredible efforts of a worldwide engineering and development team will be coming to fruition, and a new, unprecedented phase of space construction will begin." Shortly after the control module is launched from Russia, Endeavour will launch on Space Shuttle mission STS-88 from the Kennedy Space Center, FL, with the second piece of the station, a connecting module called Node-1, built by Boeing at NASA's Marshall Space Flight Center, Huntsville, AL. The node was shipped to Kennedy to begin a year of launch preparations and final testing in June 1997. Two mating adapters have since been shipped to Kennedy from California and are being attached to the node prior to its launch.
Endeavour's crew will dock the control module to the node and perform three spacewalks to make final connections between the two components during the 11-day flight. The station will then await the launch of the Russian-built Service Module, a component that will become the early living quarters, targeted for December. The first crew of the new station is planned for launch on a mission in early 1999.
The 20-ton control module will provide early power and propulsion for the station as well as the capability to remotely rendezvous and dock with the Service Module. Construction began on the control module at Khrunichev in December 1994. NASA Television is available in the continental United States and is carried on GE-2, Transponder 9C, 85 degrees West longitude, vertical polarization, with a frequency of 3880 Mhz, and audio of 6.8 Mhz.
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"Astronomers Discover an Infrared Background Glow in the Universe (9 January, 1998)",715,0,0,0
Astronomers have assembled the first definitive detection of a background infrared glow across the sky produced by dust warmed by all the stars that have existed since the beginning of time. For scientists, the discovery of this "fossil radiation" is akin to turning out all the lights in a bedroom only to find the walls, floor and ceiling aglow with an eerie luminescence. The telltale infrared radiation puts a limit on the total amount of energy released by all the stars in the universe.
Astronomers say this will greatly improve development of models explaining how stars and galaxies were born and evolved after the Big Bang. The discovery reveals a surprisingly large amount of starlight in the universe cannot be seen directly by today's optical telescopes, perhaps due to stars being hidden in dust, or being too faint or far away to be seen. The discovery culminates several years of meticulous data analysis from the Diffuse Infrared Background Experiment aboard NASA's Cosmic Background Explorer (COBE), which was launched in 1989.
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"Old Faithful Black Holes Ejects Mass Equal to an Asteroid (7 January, 1998)",716,0,0,0
Scientists observing a disk of matter surrounding a black hole in our galaxy have discovered that the disk is periodically disrupted and hurled outward in opposite directions from the black hole, in jets moving at nearly the speed of light. The black hole replenishes the disk by pulling hot gas from the surface of a nearby "companion" star, and then undergoes another disruption, repeating the sequence at half-hour intervals.
The researchers represent teams at the California Institute of Technology, Pasadena, CA, the Massachusetts Institute of Technology (MIT), Boston, and NASA's Goddard Space Flight Center, Greenbelt, MD, which all worked to correlate the disappearance of X-ray emitting hot gas in the disk with the appearance, shortly thereafter, of rapidly expanding jets.
Dr. Ronald Remillard of MIT and Dr. Jean Swank of Goddard are presenting X-ray results, obtained with NASA's Rossi X-ray Timing Explorer (RXTE), which show the disappearing disk. Dr. Stephen Eikenberry of Caltech is presenting new infrared observations which demonstrate that when the X-rays from the disk vanish, the jets suddenly appear. The observations will be the subject of a press conference by three of the researchers involved, to be held Jan. 7 during the winter meeting of the American Astronomical Society in Washington, DC.
The disks of hot gas, known as accretion disks, are commonly observed around black holes with orbiting stellar companions, but the near simultaneous disappearance of the disk and formation of the jet has never been seen before. It promises to shed light on the origin of the enigmatic jets, also commonly observed near accreting black holes, but poorly understood.
"The system behaves like the celestial version of Old Faithful," notes Dr. Craig Markwardt, a researcher working with Swank at Goddard. "At fairly regular intervals, the accretion disk is disrupted and a fast moving jet is produced."
"This jet is staggeringly more powerful than a geyser," adds Swank. "Every half-hour, the black hole, in the constellation Aquila, throws off the mass equal to that of a 100 trillion ton asteroid at nearly the speed of light (approximately 650 million miles per hour). This process clearly requires a lot of energy -- each cycle is equivalent to six trillion times the annual energy consumption of the entire United States."
"What is even more amazing is that we are seeing the first clues to the source of matter ejected in the jets -- the correlations we discovered indicate that the jet material must come from the inner disk. For years theorists have hypothesized that the jets come from somewhere close to the black hole, but no one had ever actually seen that direct link until now" said Eikenberry.
Black holes are very massive objects with gravitational fields so intense that near them, nothing, not even light, can escape their pull. While this prevents anyone from observing black holes directly, their presence can be inferred from effects on nearby matter. Many of the known or suspected black holes are orbiting a close "companion" star.
