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

Text File | 1993-02-05 | 45KB | 865 lines

"6_2_3_3.TXT" (3127 bytes) was created on 12-12-88

ORBITER MANUFACTURING AND ASSEMBLY

The structures of the orbiter were manufactured at various companies
under contract to Rockwell International's Space Transportation
Systems Division, Downey, Calif.

The upper and lower forward fuselage, crew compartment, forward
reaction control system and aft fuselage were manufactured at
Rockwell's Space Transportation Systems Division facility in Downey
and were transported overland from Downey to Rockwell's Palmdale,
Calif., assembly facility.

The midfuselage was manufactured by General Dynamics, San Diego,
Calif., and transported overland to Rockwell's Palmdale assembly
facility.

The wings (including elevons) were manufactured by Grumman, Bethpage,
Long Island, N.Y., and transported by ship from New York via the
Panama Canal to Long Beach, Calif., and then transported overland to
Rockwell's Palmdale assembly facility.

The vertical tail (including rudder/speed brake) were manufactured by
Fairchild Republic, Farmingdale, Long Island, N.Y., and transported
overland to Rockwell's Palmdale assembly facility.

The payload bay doors were manufactured at Rockwell International's
Tulsa, Okla., facility and transported overland to Rockwell's Palmdale
assembly facility.

The body flap was manufactured at Rockwell International's Columbus,
Ohio, facility and transported overland to Rockwell's Palmdale
assembly facility.

The aft orbital maneuvering system/reaction control system pods were
manufactured by McDonnell Douglas, St.  Louis, Mo., and transported by
aircraft to Rockwell's Palmdale assembly facility.  They were also
transported by aircraft from Rockwell's Palmdale assembly facility to
the Kennedy Space Center.

Approximately 250 major subcontractors supplied various systems and
components to Rockwell's Palmdale assembly facility.

Rockwell's Palmdale assembly facility is where all the individual
parts, pieces and systems came together and were assembled and tested.
Upon completion, the spacecraft was turned over to NASA for transport
overland from Palmdale to Edwards Air Force Base, California.  NASA's
Dryden Flight Research Facility at Edwards Air Force Base is the site
of the mate-demate facility for mating or demating the spacecraft and
the shuttle carrier aircraft.

For mating atop the shuttle carrier aircraft, the orbiter is raised
horizontally in the mate facility until the shuttle carrier aircraft
can be towed under the orbiter.  The orbiter is then lowered and
attached to two aft and one forward attach points on the shuttle
carrier aircraft.  These attach points on the orbiter are the same
attach points where the external tank is attached to the orbiter.

For ferry flights of the orbiters after delivery from Palmdale, dummy
orbital maneuvering system/reaction control system pods were used
along with a tail cone installed over the aft section of the orbiter
to streamline airflow.

The space shuttle main engines were manufactured by Rockwell
International's Rocketdyne Division in Canoga Park, Calif.  They are
shipped separately from Rocketdyne to the National Space Technology
Laboratories, then to the Kennedy Space Center.


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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/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 RCS 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's
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 auxiliary power units 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 event 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:

1.  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.

2.  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.

3.  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.

4.  The carbon-lined beryllium stator discs in each main landing gear
brake were replaced with thicker discs to increase braking energy
significantly.

5.  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.

6.  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 lake bed 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
Kennedy Space Center landings and how to improve crosswind landing
capability.

Modifications were made to the Kennedy Space Center shuttle landing
facility runway.  The full 300-foot width of 3,500-foot sections at
both ends of the runway were ground to smooth the runway surface
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 re-entry.
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 insulation 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.

MIDFUSELAGE 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 midfuselage 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 midfuselage from bays 4 through 11 to act as a heat sink,
distributing temperatures evenly across the bottom of the midfuselage,
reducing thermal gradients and ensuring positive margins of safety.

GENERAL-PURPOSE COMPUTERS.

New upgraded general-purpose computers (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 2.5
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 will be phased in to
augment the present KT-70 inertial measurement units in 1988-89.
These new 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
vendor'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.

Note that the side hatch jettison feature 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.


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SPACE SHUTTLE MAIN ENGINE MARGIN IMPROVEMENT PROGRAM

Improvements to the SSMEs for increased margin and durability began
with a formal Phase II program in 1983.  Phase II focused on
turbomachinery 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 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
turbomachinery, 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 turbomachinery 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 at the NASA National Space Technology
Laboratories in Mississippi and at Rockwell International's Rocketdyne
Division's Santa Susana Field Laboratory in California.


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SSME FLIGHT PROGRAM

By January 1986, there had 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
solid rocket booster 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 redline
temperature sensor and its backup.  The engine was commanded to shut
down, but the other two engines safely delivered the space shuttle to
orbit.  A major upgrade of these components was implemented to prevent
a recurrence of these conditions and will be incorporated for STS-26.


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SOLID ROCKET MOTOR REDESIGN

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.