Space & Astronomy/Space and Astronomy (October 1993).iso/mac/TEXT/SHUTTLE/STS_OUT.TXT

Text File | 1993-02-05 | 115KB | 2,312 lines

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SPACE SHUTTLE OVERVIEW

        Flight hardware for the Space Shuttle is manufactured at many
locations around the United States by NASA prime contractors and
subcontractors.

        In the case of the orbiter, the prime contractor is Rockwell
International, Downey, Calif., and major components and subsystems
for it are assembled at the firm's production plant at Palmdale,
Calif.  After an orbiter is built, it is flown to NASA's Kennedy
Space Center (KSC) Fla., atop a specially-equipped Boeing-747
aircraft called the Shuttle Carrier Aircraft.

        The Space Shuttle main engines are produced by the Rocketdyne
Division of Rockwell International, Canoga Park, Calif.  The engines
are shipped to the KSC after they have undergone engine test firings
on stands at NASA's John C.  Stennis Space Center (formerly the
National Space Technology Laboratories) near Bay St.  Louis, Miss.

        The Shuttle's huge external tank is built at NASA's Michoud Assembly
Facility near New Orleans, La., by Martin Marietta Corp., Michoud
Aerospace.  The tanks are shipped to KSC by barge, arriving at the
center's turn basin canal in the Launch Complex 39 area, where they
are unloaded and moved to the Vehicle Assembly Building (VAB).

        Several aerospace firms components for the Shuttle's solid rocket
boosters (SRB).  The solid propellant motors are built by the Wasatch
Division of the Morton Thiokol Chemical Corp., Brigham City, Utah.

        All other SRB components are produced by United Space Boosters,
Inc., Huntsville, Ala.  Booster stacking -- assembly of the entire
solid rocket booster -- is performed by the Lockheed Space Operations
Co., Titusville, Fla., the Shuttle processing contractor at KSC.

        Processing flow procedures for new and reused Shuttle flight
hardware are essentially similar.  Differences in the procedures
occur early in the pre-integration activity.  For example, newly-
produced orbiters usually undergo a period of powered-down processing
to allow time to finish work that may not have been completed at the
manufacturing plant, or to make modifications ordered after the
orbiter leaves the plant.  Also, during the initial flow processing
of a new orbiter, the main engines and orbiter maneuvering system
pods undergo checkouts before being installed.

        Another requirement for new orbiters is that their main engines are
test fired on the launch pad.  Called the Flight Readiness Firing,
the purpose of the test is to verify that the main propulsion system
works the way it is designed to work.

        For orbiters that have already flown, turnaround processing
procedures include various post-flight deservicing and maintenance
functions which are carried out in parallel with payload removal and
the installation of equipment needed for the next mission.

        If new flight hardware is called for, additional pre-launch tests
are usually needed.  For example, if new auxiliary power units (APU)
are installed, they must undergo "hot firings" to verify their
operational readiness.

        If changes are made in external tank design, the tank usually will
usually require a tanking test in which it is loaded with liquid
oxygen and hydrogen just as it is before launch.  This is called a
"confidence check" and determines the tank's ability to withstand the
high pressures and super cold temperatures of the cyrogenics.

        After separate hardware checks and servicing of major flight
elements are completed -- a process called stand-alone processing --
actual Shuttle vehicle integration starts with stacking of the SRBs
on a Mobile Launcher Platform in one of the high bays of the VAB.
Next, the external tank is moved from its VAB location and is mated
with the SRBs.

        The orbiter, having completed its pre-launch processing and after
horizontally-integrated payloads have been installed, is towed from
the Orbiter Processing Facility (OPF) to the VAB and hoisted into
position alongside the SRBs and the external tank.  It then is then
attached to the external tank.  The mating is then complete.

        The mobile launcher concept was originally developed for the Apollo
program.  It permits the complete checkout of the vehicle in the
enclosed protection of the VAB before moving the vehicle to the
launch pad.  This provides greater protection of flight hardware from
the elements and allows for more systematic checkout processing using
computer techniques.  Thus, the Shuttle spends a relatively short
time on the launch pad.

        When the VAB pre-launch preparations are completed, the entire
system -- the assembled Space Shuttle and the Mobile Launcher
Platform -- is lifted by the Crawler Transporter and rolled slowly to
the launch pad.  The move takes about 6 hours.

        At the pad, vertically integrated payloads are loaded into the
payload bay.  Then, propellant servicing and needed ordnance tasks
are performed.  Finally, the countdown gets underway, launch
readiness is confirmed and launch takes place.

        Only minutes following the launch, recovery crews, on station in the
Atlantic Ocean off shore from the launch site, prepare to recover the
spent SRBs thus beginning the process of vehicle turnaround.  While
the Shuttle is carrying out its mission in orbit, back on Earth the
ground crews already are preparing for the next mission.


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SPACE TRANSPORTATION SYSTEM


SPACE SHUTTLE PROGRAM

        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 spacecraft, 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 integration of the overall space
transportation system.  Both orbiter and integration contracts are
under the direction of NASA's Johnson Space Center in Houston, Texas.

        The SRB motors are built by the Wasatch Division of Morton Thiokol
Corp., Brigham City, Utah, 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,
Huntsville, Ala.


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SPACE SHUTTLE REQUIREMENTS

        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 Florida is used for launches to place the orbiter in
equatorial orbits (around the equator), and Vandenberg Air Force Base
launch site in California 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 California
and White Sands, N.M.  Contingency landing sites are also provided in
the event the orbiter must return to Earth in an emergency.


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LAUNCH SITES

        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 azimuth degree readings based on due east from KSC as 90 degrees.

        A 35-degree azimuth launch places the spacecraft in an orbital
inclination of 57 degrees.  This means the spacecraft 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 azimuth of 120 degrees will place the
spacecraft 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 azimuth angles further north or south would
launch a spacecraft 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
azimuth, the spacecraft 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 azimuth of 158 degrees, the
spacecraft 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 spacecraft in polar orbit from KSC
to avoid habitable land mass 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 azimuth would necessitate a path over a land mass; 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 hydrogen 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.)


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BACKGROUND AND STATUS

        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 integrater 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 spacecraft, 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
Aircraft (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 astronaut 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 astronaut crew separated the spacecraft
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 spacecraft'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 spacecraft 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, Huntsville, 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 California 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 France for the Air Show as well as to Germany, Italy,
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 spacecraft after famous
exploration sailing ships.  In the order they became operational,
they are:  Columbia (OV-102), after a sailing frigate 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 Alaska 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, Texas, 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.


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MISSION PROFILE

        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
hydrogen 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
spacecraft'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 damping.

        The forward RCS engines are inhibited prior to entry interface, and
the aft reaction control system engines maneuver the spacecraft 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.  Azimuth 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 azimuth 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 tangent 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 spacecraft
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 tangent 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 spacecraft 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 spacecraft'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 spacecraft 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 spacecraft 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).

        ABORTS.  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.

        RETURN TO LAUNCH SITE.  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.

        TRANSATLANTIC   LANDING  ABORT.  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, Spain; 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.

        ABORT TO ORBIT.  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.

        ABORT ONCE AROUND.  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.

        CONTINGENCY ABORT.  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.

"6_2_4_2_8.TXT" (5835 bytes) was created on 01-03-89

ORBITER GROUND TURNAROUND

        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 spacecraft as soon as it stops rolling.  The
ground team members take sensor measurements to ensure the atmosphere
in the vicinity of the spacecraft 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 spacecraft Freon Coolant loops
to provide cooling for the flight crew and avionics during the
postlanding and system checks.  The spacecraft fuel cells remain
powered up at this time.  The flight crew will then exit the
spacecraft, and a ground crew will power down the spacecraft.

        AT KSC, the orbiter and ground support equipment convoy move from
the runway to the Orbiter Processing Facility.

        If the spacecraft 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 spacecraft is prepared to
be ferried atop the Shuttle carrier aircraft 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
astronaut crew in preparing the orbiter for loading aboard the
Shuttle carrier aircraft 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 spacecraft 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 hydrogen 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.


"6_2_4_2_9.TXT" (40951 bytes) was created on 01-03-89

OPERATIONAL IMPROVEMENTS AND MODIFICATIONS

        Many of the changes and upgrades in the Space Shuttle systems and
components were under way before the 51-L accident as part of NASA's
continual improvement and upgrade program.  However, NASA has taken
advantage of the Space Shuttle program downtime since the accident to
accelerate the testing and integration of these improvements and
upgrades as well as fixes required as a result of the accident.

        ORBITER.  The following identifies the major improvements or
modifications of the orbiter.  Approximately 190 other modifications
and improvements were also made.

        ORBITAL MANEUVERING SYSTEM AND REACTION CONTROL SYSTEM
AC-MOTOR-OPERATED VALVES.  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 nitrogen 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
nitrogen tetroxide or monomethyl hydrazine leakage into the bellows
of each ac-motor-operated valve.

        PRIMARY REACTION CONTROL SYSTEM THRUSTERS.  The wiring of the fuel
and oxidizer injector solenoid valves was wrapped around each of the
38 primary RCS thrust chambers to remove electrical power from these
valves in the event of a primary RCS thruster instability.

        FUEL CELL POWER PLANTS.  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 hydrogen
pump and water separator of each fuel cell power plant were improved
to minimize excessive hydrogen gas entrained in the power plant
product water.  A current measurement detector was added to monitor
the hydrogen pump of each fuel cell power plant and provide an early
indication of hydrogen 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 pH sensor failed in the past, the flight crew had to sample the
potable water.

        AUXILIARY POWER UNITS.  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 turbine 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
turbine 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.

        MAIN LANDING GEAR.  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 beryllium 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 beryllium 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.

        NOSE WHEEL STEERING.  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.

        THERMAL PROTECTION SYSTEM.  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 insulation 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 iInsulation (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 Insulation (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.

        WING MODIFICATION.  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.

        MID-FUSELAGE MODIFICATIONS.  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.

        GENERAL ,PURPOSE COMPUTERS.  New upgraded General Purpose Computers
(GPC), IBM 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.

        INERTIAL MEASUREMENT UNITS.  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 microprocessor with memory for
processing and storing compensation and scale factor data from the
IMU manufacturer's calibration, 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.

        CREW ESCAPE SYSTEM.  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 crew member 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 crew member 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.

        EMERGENCY EGRESS SLIDE.  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 crew members to drop approximately
10.5 feet to the ground.  The previous arrangement could have injured
crew members or prevented an already-injured crew member from
evacuating and moving a safe distance from the orbiter.

        17-INCH ORBITER/EXTERNAL TANK DISCONNECTS.  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.

SPACE SHUTTLE MAIN ENGINE MARGIN IMPROVEMENT PROGRAM

        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), turbine 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 turbine 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 turbine blades, thereby extending the
operating life of the turbopumps.  These changes included applying
surface texture to important parts of the fuel turbine blades to
improve the material properties in the pressure of hydrogen and
incorporating a damper into the high-pressure oxidizer turbine 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 California.

SSME FLIGHT PROGRAM

        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.

SOLID ROCKET MOTOR REDESIGN

        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.

        ORIGINAL VERSUS REDESIGNED SRM FIELD JOINT.  The SRM field-joint
metal parts, internal case insulation 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 circumference 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         ORIGINAL VERSUS REDESIGNED SRM FIELD JOINT

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 insulation is breached.

        The field joint internal case insulation 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.

        ORIGINAL VERSUS REDESIGNED SRM CASE-TO-NOZZLE 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 insulation and O-rings.  Radial bolts with
Stato-O-Seals were added to minimize the joint sealing gap opening.
The internal insulation 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.

        NOZZLE.  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 insulation 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.

        FACTORY JOINT.  Minor modifications were made in the case factory
joints by increasing the insulation thickness and lay-up to increase
the margin of safety on the internal insulation.  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.

        PROPELLANT.  The motor propellant forward transition region was
recontoured to reduce the stress fields between the star and
cylindrical portions of the propellant grain.

        IGNITION SYSTEM.  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 insulation 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 asbestos as one of its constituents.

        GROUND SUPPORT EQUIPMENT.  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.

        DESIGN ANALYSIS SUMMARY.  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 pretest predictions.

        VERIFICATION/CERTIFICATION TEST.  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 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 insulation
segment and is undergoing temperature cycling.  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 data base 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 insulation that was an early design variation of the J-seal.
Tests were conducted with and without flaws built into the U-seal
joint insulation; 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 insulation.

        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 engineering 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 graphite 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 insulation 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.

        NON-DESTRUCTIVE EVALUATION.  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 insulation.  It has sealed insulation
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 geometry of
the tang capture feature inherently provides more isolation of the
edge insulation 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
insulation bonded surfaces.  All segments will again be X-rayed for
the first flight and near-term subsequent flights.

        CONTINGENCY PLANNING.  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.

        INDEPENDENT OVERSIGHT.  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, engineering 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.


"6_2_4_2_11.TXT" (12433 bytes) was created on 01-03-89

CHRONOLOGY


l972

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,
Utah.

Aug. 12 First free flight Approach and Landing Test (ALT) of orbiter
Enterprise from Shuttle carrier aircraft 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 aircraft to Kennedy
Space Center with overnight stops at El Paso and San Antonio, Texas,
and Eglin AFB, Fla.

June 15 First SRB qualification test firing; 122 seconds.


1980

Nov. 26 Columbia mated to SRBs and external tank at Vehicle Assembly
Building (VAB) for STS-l mission.

Dec. 29 Space Shuttle vehicle moved from VAB to Launch Complex 39A
for STS-l mission.1981

1980 continued

Feb. 20 Flight readiness firing of Columbia's main engines; 20
seconds.

April 20-21     Columbia returned to KSC by Shuttle carrier aircraft 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, Texas.

Dec. ll Spacelab l arrives at KSC.


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, Texas.

July 14-15      Columbia flown to KSC via Dyess AFB, Texas.

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, Texas

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, Texas.

Sept. 23        Columbia moved to VAB for mating in preparation for STS-9.

Sept. 28        STS-9 vehicle moved to launch pad.

Oct. 17 STS-9 launch 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-9 vehicle again moved to launch pad.

Nov. 8-9        Discovery flown from Vandenberg AFB to KSC via Carswell AFB,
Texas.

Nov. 28-Dec. 8  STS-9 mission; 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.

l984 continued

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.

1985 continued

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, Texas.

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, Texas.

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. 1985 continued


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. 1986 continued

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        Astronaut 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 tp 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.