national space transportation system overview 1988
TRANSCRIPT
I
NationalSpace TransportationSystem
Overview
September 1988
N/_SA
CONTENTS
SPACE SHUTTLE PROGRAM ......................................... 1SPACE SHUTTLE REQUIREMENTS ................................... 1LAUNCH SITES .................................................... 3BACKGROUND AND STATUS ........................................ 5MISSION PROFILE ................................................. 8
Aborts ........................................................ 10Return to Launch Site ............................................ 11
Transatlantic Landing Abort ........................................ 12Abort to Orbit .................................................. 12Abort Once Around ............................................. 13
Contingency Abort .............................................. 13ORBITER GROUND TURNAROUND .................................. 13OPERATIONAL IMPROVEMENTS AND MODIFICATIONS ................ 18
Orbiter ........................................................ 18OMS and RCS System AC-motor-operated valves ......................... 18Primary Reaction Control System Thrusters .............................. 18Fuel Cell Power Plants ............................................ 18
Auxiliary Power Units ............................................. 18Main Landing Gear .............................................. 19Nose Wheel Steering ............................................. 19Thermal Protection System ......................................... 20Wing Modification .............................................. 20Mid-Fuselage Modifications ........................................ 20General Purpose Computers ........................................ 20Inertial Measurement Units ......................................... 20Crew Escape System ............................................. 21Emergency Egress Slide ........................................... 2117-inch Orbiter/Extemal Tank Disconnects .............................. 21
Space Shuttle Main Engine Margin Improvement Program ............... 22SSME Hight Program ............................................ 22
Solid Rocket Motor Redesign ...................................... 23Original versus Redesigned SRM Field Joint ............................ 23Original versus Redesigned SRM Case-to-Nozzle Joint ..................... 23Nozzle ...................................................... 23Factory Joint ................................................. 26Propellant .................................................... 2 6Ignition System ................................................. 26Ground Support Equipment ........................................ 26Design Analysis Summary ......................................... 26Veilfication/Certification Test ....................................... 26Non-Destructive Evaluation ........................................ 26
Contingency Planning ............................................ 29Independent Oversight ............................................ 29
NSTS PROGRAM MANAGEMENT .................................... 29NSTS Organization .............................................. 30
Launch Constraing Procedures ..................................... 31Launch Decision Process ......................................... 31Management Communications ...................................... 32
CHRONOLOGY .................................................... 33
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SPACE TRANSPORTATION SYSTEM
SPACE SHUTTLE PROGRAM
The Space Shuttle is developed by theNational Aeronautics and Space Administration.NASA coordinates and manages the SpaceTransportation System (NASA's name for theoverall Shuttle program), includingintergovernmental agency requirements andinternational and joint projects. NASA alsooversees the launch and space flight requirementsfor civilian and commercial use.
The Space Shuttle system consists of fourprimary elements: an orbiter spacecraft, two SolidRocket Boosters (SRB), an external tank to housefuel and oxidizer and three Space Shuttle mainengines.
The orbiter is built by Rockwell Intemational'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 SpaceCenter 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 thedirection of NASA's George C. Marshall Space
Flight Center, Huntsville, Ala.
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 orbiterand the two solid rocket boosters b_ reusable.
Other features of the Shuttle:
The orbiter has carried a flight crew of up toeight persons. A total of 10 persons could becarried under emergency conditions The basicmission is 7 days in space. The crew compartmenthas a shirtsleeve environment, and the acceleration
load is never greater than 3 Gs. In its return toEarth, the orbiter has a cross-range maneuveringcapability of 1,100 nautical miles (1,265 statutemiles).
The Space Shuttle is launched in an uprightposition, with thrust provided by the three SpaceShuttle engines and the two SRB. After about 2minutes, the two boosters are spent and areseparated from the extemal tank. They fall into theocean at predetermined points and are recoveredfor reuse.
Length
Height
Wingspan
Approximateweight* Grosslift-off, whichwig
varydependingonpayloadweightandonboardconsumables
• Nominalend-of-missionlandingwithpayload,whichwiflvarydependingonpayloadreturnweight
Thrust(sealevel}• Solidrocketboosters
• Orbitermainengines
Cargobay• Length• Diameter
Overall Shuttle Orbiter
184.2 feet
76.6 feet
4.5 millionpounds
3,300,000 poundsofthrusteachinvacuum
122.17 feet
56.67 feet
78.06 feet
230,000 pounds
393,800 poundsof thrusteachat sealevldat104 percent
60 feet15 feet
Space Shuttle Program
The Space Shuttle main engines continuefiring for about 8 minutes. They shut down justbefore the craft is inserted into orbit. The external
tank is then separated from the orbiter. It followsa ballistic trajectory into a remote area of the oceanbut is not recovered.
There are 38 primary Reaction Control System(RCS) engines and six vemier RCS engineslocated on the orbiter. The first use of selected
primary reaction control system engines occurs atorbiter/external tank separation. The selected
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primaryreactioncontrolsystemenginesareusedin theseparationsequenceto provideanattitudeholdfor separation.Thentheymovetheorbiterawayfrom the externaltank to ensureorbiterclearancefromtheareof therotatingexternaltank.Finally,theyreturnto anattitudeholdpriorto theinitiationof thefiring of theOrbitalManeuveringSystem(OMS)enginesto placetheorbiterintoorbit.
Theprimaryand/orvemierRCSenginesareusednormallyonorbit to provide attitude pitch,roll and yaw maneuvers as well as translationmaneuvers.
orbit, and only one thrusting sequence is used fordeorbit.
The orbiter's velocity on orbit is
approximately 25,405 feet per second (17,322statute miles per hour). The deorbit maneuverdecreases this velocity approximately 300 fps(205 mph) for reentry.
In some missions, only one OMS thrustingsequence is used to place the orbiter on orbit. Thisis referred to as direct insertion. Direct insertion is
a technique used in some missions where there arehigh-performance requirements, such as a heavy
SolidRocketBooster12.17 FeetDiameter
ExternalTank _ _ |
:: '-++-'-- " I
122.17 Feet - [
Orbiter Tank-Orbiter
Attachments
76.6 Feet
149.16 Feet
154.2 Feet
184.2 Feet
Space Shuttle Statistics
OrbiterWeightinPounds(Approximate}
OrbiterVehicle TotalDryWeightWith TotalDryWeight(OV} ThreeSpaceShuttle WithoutThreeSpecs
MainEngines ShuttleMainEngines
OV-102 Columbia 178,289 157,289
OV-103 Discovery 171,419 151,419
OV-104Atlantis 171,205 151,205
SolidRocketBoosterWeightsinPounds(Approximate)
1,300,000, Eachat Launch(PropellantWeight1,100,000, Each).InertWeight192,000, Each.
ExternalTankWeightinPounds(Approximate)
1,655,600 WithPropellants.InertWeight66,000.
56.8!
ool_ _!
The two OMS engines are used to place theorbiter on orbit, for major velocity maneuvers onorbit and to slow the orbiter for reentry, called thedeorbit maneuver. Normally, two OMS enginethrusting sequences are used to place the orbiter on
payload or a high orbital altitude. This techniqueuses the Space Shuttle main engines to achieve thedesired apogee (high point in an orbit) altitude,thus conserving orbital maneuvering systempropellants. Following jettison of the external
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tank, only one OMS thrusting sequence is requiredto establish the desired orbit altitude.
For deorbit, the orbiter is rotated tail first inthe direction of the velocity by the primary reactioncontrol system engines. Then the OMS enginesare used to decrease the orbiter's velocity.
During the initial entry sequence, selectedprimary RCS engines are used to control theorbiter's attitude (pitch, roll and yaw). As
aerodynamic pressure builds up, the orbiter flightcontrol surfaces become active and the primaryreaction control system engines are inhibited.
During entry, the thermal protection systemcovering the entire orbiter provides the protectionfor the orbiter to survive the extremely hightemperatures encountered during entry. Thethermal protection system is reusable (it does notbum offor ablate during entry).
The unpowered orbiter glides to Earth andlands on a runway like an airplane. Nominaltouchdown speed varies from 184 to 196 knots(213 to 225 miles per hour).
The main landing gear wheels have a brakingsystem for stopping the orbiter on the runway, andthe nose wheel is steerable, again similar to aconventional airplane.
There are two launch sites for the SpaceShuttle. Kennedy Space Center (KSC) in Floridais used for launches to place the orbiter inequatorial orbits (around the equator), andVandenberg Air Force Base launch site inCalifornia will be used for launches that place fileorbiter in polar orbit missions.
Landing sites are located at the KSC and
Vandenberg. Additional landing sites are providedat Edwards Air Force Base in California and White
Sands, N.M. Contingency landing sites are alsoprovided in the event the orbiter must return toEarth in an emergency.
LAUNCH SITES
Space Shuttles destined for equatorial orbitsare launched from the KSC, and those requiringpolar orbital planes will be launched fromVandenberg.
Orbital mechanics and the complexities ofmission requirements, plus safety and thepossibility of infringement on foreign air and landspace, prohibit polar orbit launches from the KSC.
Kennedy Space Center launches have anallowable path no less than 35 degrees northeastand no greater than 120 degrees southeast. Theseare azimuth degree readings based on due eastfrom 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 trajectoriesaround the Earth will never exceed an Earth
latitude higher or lower than 57 degrees north orsouth of the equator.
A launch path from KSC at an azimuth of 120degrees will place the spacecraft in an orbitalinclination of 39 degrees (it will be above or below39 degrees north or south of the equator).
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Space-Shuttle Launch Sites
These two azimuths - 35 and 120 degrees -represent the launch limits from the KSC. Anyazimuth angles further north or south wouldlaunch a spacecraft over a habitable land mass,adversely affect safety provisions for abort orvehicle separ_itibn conditions, or present theundesirable possibility that the SRB or externaltank could land on foreign land or sea space.
Launches from Vandenberg have an allowablelaunch path suitable for polar insertions south,southwest and southeast.
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Thelaunchlimits atVandenbergare201and158degrees.At a201-degreelaunchazimuth,thespacecraftwould be orbiting at a 104-degreeinclination.Zerodegreeswouldbeduenorthofthelaunchsite,andtheorbitaltrajectorywouldbewithin 14degreeseastor westof thenorth-southpole meridian. At a launchazimuthof 158degrees,thespacecraftwouldbeorbitingata70-degreeinclination,andthe trajectorywouldbewithin 20 degreeseast or westof the polarmeridian.Like KSC,Vandenberghasallowablelaunchazimuthsthatdonot passoverhabitableareasor involve safety,abort,separationandpoliticalconsiderations.
TheEarthrotatesfromwestto eastat a speedofapproximately 900 nautical miles per hour (1,035mph). A launch to the east uses the Earth's
Attempting to launch and place a spacecraft inpolar orbit from KSC to avoid habitable land masswould be uneconomical because the Shuttle's
payload would be reduced severely-down toapproximately 17,000 pounds. A northerly launchinto polar orbit of 8 to 20 degrees azimuth wouldnecessitate a path over a land mass; and mostsafety, abort, and political constraints would haveto be waived. This prohibits polar orbit launchesfrom the KSC.
Mission requirements and payload weightpenalties also are major factors in selecting alaunch site.
NASA's latest assessment of orbiter ascent
and landing weights incorporates currently
LzKpmql
MECO - Main Eo_ne CutoffOMS - OrbitalManeuveringSystem
NominalMissionor ATO
On-OrbitOMS Maneuvers
MECOISeparotion
MECO MECO/ N
Separation _
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IIIIItII
I"ET Impact
Danrbit
To Landing
,,1Abort and Normal Mission Profile
The Earth rotates from west to east at a speedof approximately 900 nautical miles per hour(1,035 mph). A launch to the east uses the Earth'srotation somewhat as a springboard. The Earth'srotational rate a/so is the reason the orbiter has a
approved modifications to all vehicle elements,including crew escape provisions, and assumes amaximum Space Shuttle main engine throttlesetting of 104 percent. It is noted that theresumption of Space Shuttle flights initially
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resumptionof Space Shuttle flights initiallyrequires more conservative flight design criteriaand additional instrumentation, which reduces the
following basic capabilities by approximately1,600 pounds:
•Kennedy Space Center Eastern Space andMissile Center (ESMC) satellite deploymissions. The basic cargo-lift capability for adue east (28.5 degrees) launch is 55,000pounds to a l l0-nautical-mile (126-statute-mile) orbit using OV-103 (Discovery) or OV-104 (Atlantis) to support a 4-day satellitedeploy mission. This capability will bereduced approximately 100 pounds for eachadditional nautical mile of altitude desired bythe customer.
The payload capability for the same satellitedeploy mission with a 57-degree inclination is41,000 pounds.
The performance for intermediate inclinationscan be estimated by allowing 500 pounds perdegree of plane change between 28.5 and 57degrees.
If OV-102 (Columbia) is used, the cargo-liftweight capability must be decreased byapproximately 8,400 pounds. This weightdifference is attributed to an approximately 7,150-pound difference in inert weight, 850 pounds oforbiter experiments, 300 pounds of additionalthermal protection system and 100 pounds toaccommodate a fifth cryogenic liquid oxygen andliquid hydrogen tank set for the power reactantstorage and distribution system.
•Vandenberg Air Force Base Western Spaceand Missile Center (WSMC) satellite deploymissions. Using OV-103 (Discovery) or OV-104 (Atlantis), the cargo-lift weight capabilityis 29,600 pounds for a 98-degree launchinclination and 110-nautical-mile (126-statute-mile) polar orbit. Again, an increase inaltitude costs approximately 100 pounds pernautical mile. NASA assumes also that the
advanced solid rocket motor will replace thefilament-wound solid rocket motor case
previously used for western test rangeassessments. The same mission at 68 degreesinclination (minimum western test rangeinclination based on range safety limitations) is49,600 pounds. Performance for intermediate
inclinations can be estimated by allowing 660pounds for each degree of plane changebetween inclinations of 68 and 98 degrees.
•Landing weight limits. All the Space Shuttleorbiters are currently limited to a total vehiclelanding weight of 240,000 pounds for abortlandings and 230,000 pounds for nominalend-of-mission landings. It is noted that eachadditional crew person beyond the five-personstandard is chargeable to the cargo weightallocation and reduces the payload capabilityby approximately 500 pounds. (This is anincrease of 450 pounds to account for the crewescape equipment.)
BACKGROUND AND STATUS
On July 26, 1972, NASA selected Rockwell'sSpace Transportation Systems Division inDowney, Calif., as the industrial contractor for thedesign, development, test and evaluation of theorbiter. The contract called for fabrication and
testing of two orbiters, a full-scale structural testarticle, and a main propulsion test article. Theaward followed years of NASA and Air Forcestudies to define and assess the feasibility of areusable space transportation system.
NASA previously (March 31, 1972) hadselected Rockwell's Rocketdyne Division todesign and develop the Space Shuttle mainengines. Contracts followed to Martin Marietta forthe external tank (Aug. 16, 1973) and MortonThiokors Wasatch Division for the solid rocket
boosters (June 27, 1974).
In addition to the orbiter DDT&E contract,
Rockwell's Space Transportation SystemsDivision was given contractual responsibility assystem integrater for the overall Shuttle system.
Rockwell's Launch Operations, part of the
Space Transportation Systems Division, wasunder contract to NASA's Kennedy Space Center
for turnaround, processing, prelaunch testing, andlaunch and recovery operations from STS-1through the STS- 11 mission.
On Oct. 1, 1983, the Lockheed SpaceOperations Co. was awarded the Space Shuttleprocessing contract at KSC for turnaroundprocessing, prelaunch testing, and launch andrecovery operations.
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Thefirst orbiterspacecraft,Enterprise(OV-101),wasrolledoutonSept.17,1976.OnJan.31, 1977,it wastransported38milesoverlandfrom Rockwelrsassemblyfacility at Palmdale,Calif., to NASA's Dryden Flight Research Facility
at Edwards Air Force Base for the Approach andLanding Test (ALT) program.
The 9-month-long ALT program wasconducted from February through November 1977
at Dryden and demonstrated the orbiter could fly inthe atmosphere and land like an airplane exceptwithout power, a gliding flight.
The ALT program involved ground tests andflight tests.
The ground tests included taxi tests of the 747Shuttle Carrier Aircraft (SCA) with the Enterprisemated atop the SCA to determine structural loadsand responses and assess the mated capability inground handling and control characteristics up toflight takeoff speed. The taxi tests also validated747 steering and braking with the orbiter attached.A ground test of orbiter systems followed theunmanned captive tests. All orbiter systems wereactivated as they would be in atmospheric flight.This was the final preparation for the mannedcaptive-flight phase.
Five captive flights of the Enterprise mountedatop the SCA with the Enterprise unmanned andEnterprise systems inert were conducted to assess
the structural integrity and performance-handlingqualities of the mated craft.
Three manned captive flights that followed thefive unmanned captive flights included anastronaut crew aboard the orbiter operating itsflight control systems while the orbiter remained
perched atop the SCA. These flights weredesigned to exercise and evaluate all systems in theflight environment in preparation for the orbiterrelease (free) flights. They included flutter tests ofthe mated craft at low and high speed, a separationtrajectory test and a dress rehearsal for the firstorbiter free flight.
In the five free flights the astronaut crewseparated the spacecraft from the SCA andmaneuvered to a landing at Edwards Air ForceBase. In the first four such flights the landingswere on a dry lake bed; in the fifth, the landingwas on Edwards' main concrete runway under
conditions simulating a return from space. Thelast two free flights were made without the tailcone, which is the spacecraft's configurationduring an actual landing from Earth orbit. These
flights verified the orbiter's pilot-guided approachand landing capability; demonstrated the orbiter'ssubsonic terminal area energy managementautoland approach capability; and verified theorbiter's subsonic airworthiness, integrated
system operations and selected subsystems inpreparation for the first manned orbital flight. Theflights demonstrated the orbiters ability toapproach and land safely with a minimum grossweight and using several center-of-gravity
For all of the captive flights and the first threefree flights, the orbiter was outfitted with a tailcone covering its aft section to reduce aerodynamicdrag and turbulence. The final two free flightswere without the tail cone, and the three simulated
Space Shuttle main engines and two orbitalmaneuvering system engines were exposedaerodynamically.
The final phase of the ALT program preparedthe spacecraft for four ferry flights. Fluid systemswere drained and purged, the tail cone wasreinstalled and elevon locks were installed.
The forward attachment strut was replaced tolower the orbiter's cant from 6 to 3 degrees. Thisreduces drag to the mated vehicles during the ferryflights.
After the ferry flight tests, OV-101 wasreturned to the NASA hangar at Dryden andmodified for vertical ground vibration tests atNASA's Marshall Space Flight Center,Huntsville, Ala.
On March 13, 1978, the Enterprise was ferriedatop the SCA to MSFC. At Marshall, Enterprisewas mated with the external tank and SRB and
subjected to a series of vertical ground vibrationtests. These tested the mated conflguration'scritical structural dynamic response modes, whichwere assessed against analytical math models usedto design the various element interfaces.
These were completed in March 1979. OnApril 10, 1979 the Enterprise was ferried toKennedy Space Center, mated with the externaltank and SRB and transported via the mobile
launcher platform to Launch Complex 39-A. At
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LaunchComplex39-A,theEnterpriseservedasapracticeandlaunchcomplexfit-checkverificationtoolrepresentingtheflightvehicles.
It wasferriedbacktoDrydenatEdwardsAFBinCaliforniaonAug.16,1979,andthenreturnedoverlandto Rockwell'sPalmdalefinal assemblyfacilityonOct. 30, 1979. Certaincomponentswererefurbishedfor useonflight vehiclesbeingassembledatPalmdale.TheEnterprisewasthenreturnedoverlandtoDrydenonSept.6, 1981.
DuringexhibitionatthePads,May andJune1983,EnterprisewasferriedtoFrancefortheAirShowaswell asto Germany,Italy,EnglandandCanadabeforereturningtoDryden.
FromApril to October1984,Enterprisewasferriedto VandenbergAFB 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 toVandenberg and used as a practice and fit-checkverification tool. On May 24, 1985, Enterprisewas ferried from Vandenberg to Dryden.
On Sept. 20, 1985, Enterprise was ferriedfrom Dryden Flight Research Facility to KSC. OnNov. 18, 1985, Enterprise was ferried from KSCto Dulles Airport, Washington, D.C., and becamethe property of the Smithsonian Institution. TheEnterprise was built as a test vehicle and is notequipped for space flight.
The second orbiter, Columbia (OV-102), wasthe first to fly into space, it was transportedoverland on March 8, 1979, from Palmdale to
Dryden for mating atop the SCA and ferried toKSC. It arrived on March 25, 1979, to beginpreparations for the first flight into space.
The structural test article, after I I months of
extensive testing at Lockheed's facility inPalmdale, was retumed to Rockwell's Palmdale
facility for modification to become the secondorbiter available for operational missions, it wasredesignated OV-099, the Challenger.
The main propulsion test article (MPTS-098)consisted of an orbiter aft fuselage, a trussarrangement that simulated the orbiter's mid-fuselage and the Shuttle main propulsion system(three Space Shuttle main engines and the extemai
tank). This test structure is at the Stennis SpaceCenter in Mississippi. A series of static firingswas conducted from 1978 through 1981 in supportof the first flight into space.
On Jan. 29, 1979, NASA contracted withRockwell to manufacture two additional orbiters,
OV-103 and OV-104 (Discovery and Atlantis),convert the structural test article to space flightconfiguration (Challenger) and modify Columbiafrom its development configuration to that requiredfor operational flights.
NASA named the first four orbiter spacecraftafter famous exploration sailing ships. In theorder they became operational, they are: Columbia(OV-102), after a sailing frigate launched in 1836,one of the first Navy ships to circumnavigate theglobe. Columbia also was the name of the Apollo11 command module that carried Neil Armstrong,Michael Collins and Edward (Buzz) Aldrin on thefirst 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 aNavy ship, which from 1872 to 1876 made aprolonged exploration of the Atlantic and Pacificoceans. It also was used in the Apollo programfor the Apollo 17 lunar module. Challenger wasdelivered to DSC on July 5, 1982. Discovery(OV-103), after two ships, the vessel in whichHenry Hudson in 1610-11 attempted to search fora northwest passage between the Atlantic andPacific oceans and instead discovered Hudson Bayand the ship in which Capt. Cook discovered theHawaiian Islands and explored southern Alaskaand western Canada. Discovery was delivered toKSC on Nov. 9, 1983. Atlantis (OV-104), after atwo-masted ketch operated for the Woods HoleOceanographic Institute from 1930 to 1966,which traveled more than half a million miles inocean research. Atlantis was delivered to KSC on
April 3, 1985.
In April 1983, under contract to NASA,Rockwell's Space Transportation SystemsDivision, Downey, Calif., began the constructionof structural spares for completion in 1987. The
structural spares program consisted of an aftfuselage, crew compartment, forward reaction
control system, lower and upper forward fuselage,mid-fuselage, wings (elevons), payload baydoors, vertical stabilizer (rudder/speed brake),
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body flap andoneset of orbital maneuveringsystem/reaction control system pods.
On Sept. 12, 1985, Rockwell International'sShuttle Operations Co., Houston, Texas, wasawarded the Space Transportation Systemoperation contract at NASA's Johnson SpaceCenter, consolidating work previously performedunder 22 contracts by 16 different contractors.
On July 31, 1987, NASA awardedRockwell's Space Transportation SystemsDivision, Downey, Calif., a contract to build areplacement Space Shuttle orbiter using thestructural spares. The replacement orbiter will beassembled at Rockwell's Palmdale, Calif.,
assembly facility and is scheduled for completionin 1991. This orbiter is designated OV-105.
MISSION PROFILE
In the launch configuration, the orbiter andtwo 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 thelaunch pad up to 30 seconds before liftoff is byslidewire. There are seven 1,200-foot-longslidewires, each with one basket. Each basket is
designed to carry three persons. The baskets, 5feet in diameter and 42 inches deep, are suspendedbeneath the slide mechanism by four cables. Theslidewires carry the baskets to ground level.
Upon departing the basket at ground level, theflight crew progresses to a bunker that is designedto protect it from an explosion on the launch pad.
At launch, the three Space Shuttle mainengines - fed liquid hydrogen fuel and liquidoxygen oxidizer from the extemal tank - areignited first. When it has been verified that the
engines are operating at the proper thrust level, asignal is sent to ignite the SRB. At the properthrust-to-weight ratio, initiators (small explosives)at eight hold-down bolts on the SRB are fired torelease the Space Shuttle for liftoff. All this takesonly a few seconds.
Maximum dynamic pressure is reached earlyin the ascent, nominally approximately 60 seconds
after liftoff. Approximately 1 minute later (2minutes into the ascent phase), the two SRB have
consumed their propellant and are jettisoned fromthe extemal tank. This is triggered by a separationsignal from the orbiter.
The boosters briefly continue to ascend, whilesmall motors fire to carry them away from theSpace Shuttle. The boosters then turn anddescend, and at a predetermined altitude,parachutes are deployed to decelerate them for asafe splashdown in the ocean. Splashdown occursapproximately 141 nautical miles (162 statutemiles) from the launch site. The boosters arerecovered and reused.
Meanwhile, the orbiter and external tank
continue to ascend, using the thrust of the threeSpace Shuttle main engines. Approximately 8minutes after launch and just short of orbitalvelocity, the three Space Shuttle engines are shutdown (main engine cutoff), and the external tankis jettisoned on command from the orbiter.
The forward and aft reaction control systemengines provide attitude (pitch, yaw and roll) andthe translation of the orbiter away from the externaltank at separation and retum to attitude hold priorto the orbital maneuvering system thrustingmaneuver.
The external tank continues on a ballistic
trajectory and enters the atmosphere, where it
disintegrates. Its projected impact is in the IndianOcean (except for 57-degree inclinations) in thecase of equatorial orbits KSC launch) and in theextreme southem Pacific Ocean in the case of a
Vandenberg launch.
Normally, two thrusting maneuvers using thetwo OMS engines at the aft end of the orbiter areused in a two-step thrusting sequence: to completeinsertion into Earth orbit and to circularize the
spacecraft's orbit. The OMS engines are also usedon orbit for any major velocity changes.
In the event of a direct-insertion mission, onlyone OMS thrusting sequence is used.
The orbital altitude of a mission is dependentupon that mission. The nominal altitude can varybetween 100 to 217 nautical miles (115 to 250statute miles).
The forward and aft RCS thrusters (engines)provide attitude control of the orbiter as well as
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anyminortranslationmaneuversalonga givenaxisonorbit.
At thecompletionof orbitaloperations,theorbiteris orientedin a tail first attitudeby thereactioncontrolsystem.ThetwoOMSenginesarecommandedtoslowtheorbiterfordeorbit.
Thereactioncontrolsystemturnstheorbiter'snoseforward for entry. The reactioncontrolsystemcontrolsthe orbiter until atmosphericdensity is sufficient for the pitch and rollaerodynamiccontrolsurfacestobecome effective.
Entry interface is considered to occur at400,000 feet altitude approximately 4,400 nauticalmiles (5,063 statute miles) from the landing siteand at approximately 25,000 feet per secondvelocity.
At 400,000 feet altitude, the orbiter ismaneuvered to zero degrees roll and yaw (wingslevel) and at a predetermined angle of attack forentry. The angle of attack is 40 degrees. Theflight control system issues the commands to roll,pitch and yaw reaction control system jets for ratedamping.
The forward RCS engines are inhibited priorto entry interface, and the aft reaction controlsystem engines maneuver the spacecraft until adynamic pressure of 10 pounds per square foot issensed, which is when the orbiter's ailerons
become effective. The aft RCS roll engines arethen deactivated. At a dynamic pressure of 20pounds per square foot, the orbiter's elevatorsbecome active, and the aft RCS pitch engines aredeactivated. The orbiter's speed brake is usedbelow Mach 10 to induce a more positivedownward elevator trim deflection. At
approximately Math 3.5, the rudder becomesactivated, and the aft reaction control system yawengines are deactivated at 45,000 feet.
Entry guidance must dissipate the tremendousamount of energy the orbiter possesses when itenters the Earth's atmosphere to assure that theorbiter does not either bum up (entry angle toosteep) or skip out of the atmosphere (entry angletoo shallow) and that the orbiter is properlypositioned to reach the desired touchdown point.
During entry, energy is dissipated by theatmospheric drag on the orbiter's surface. Higher
atmospheric drag levels enable faster energydissipation with a steeper trajectory. Normally,the angle of attack and roll angle enable theatmospheric drag of any flight vehicle to becontrolled. However, for the orbiter, angle ofattack was rejected because it creates surfacetemperatures above the design specification. Theangle of attack scheduled during entry is loadedinto the orbiter computers as a function of relativevelocity, leaving roll angle for energy control.Increasing the roll angle decreases the verticalcomponent of lift, causing a higher sink rate andenergy dissipation rate. Increasing the roll ratedoes raise the surface temperature of the orbiter,but not nearly as drastically as an equal angle ofattack command.
If the orbiter is low on energy (current range-to-go much greater than nominal at currentvelocity), entry guidance will command lower thannominal drag levels. If the orbiter has too much
energy (current range-to-go much less thannominal at the current velocity), entry guidancewill 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 planecontaining the orbiter's position vector and theheading alignment cylinder tangency point and theplane containing the orbiter's position vector andvelocity vector. When the azimuth error exceeds acomputer-loaded number, the orbiter's roll angle isreversed.
Thus, descent rate and down ranging arecontrolled by bank angle. The steeper the bankangle, the greater the descent rate and the greaterthe drag. Conversely, the minimum drag attitudeis wings level. Cross range is controlled by bankreversals.
The entry thermal control phase is designed tokeep the backface temperatures within the designlimits. A constant heating rate is established untilbelow 19,000 feet per second.
The equilibrium glide phase shifts the orbiterfrom the rapidly increasing drag levels of thetemperature control phase to the constant draglevel of the constant drag phase. The equilibriumglide flight is defined as flight in which the flight
path angle, the angle between the local horizontaland the local velocity vector, remains constant.
-9-
Equilibriumglideflight providesthemaximumdownrangecapability. It lastsuntil the dragaccelerationreaches33feetpersecondsquared.
Theconstantdragphasebeginsat thatpoint.Theangleof attackis initially 40degrees,but itbegins to ramp down in this phase toapproximately36degreesbytheendof thisphase.
In thetransitionphase,the angleof attackcontinues to ramp down, reaching theapproximately14-degreeangleof attackat theentryTerminalAreaEnergyManagement(TAEM)interface,at approximately83,000feetaltitude,2,500feetpersecond,Mach2.5and52nauticalmiles(59statutemiles)fromthelandingrunway.ControlisthentransferredtoTAEMguidance.
During the entry phasesdescribed,theorbiter'sroll commandskeepthe orbiteron thedragprofileandcontrolcrossrange.
TAEM guidancesteersthe orbiter to thenearestof twoheadingalignmentcylinders,whoseradiiareapproximately18,000feetandwhicharelocatedtangentto andoneithersideof therunwaycenterlineon the approachend. In TAEMguidance,excessenergyisdissipatedwithanS-turn;andthespeedbrakecanbeusedto modifydrag,lift-to-dragratio andflight pathangleinhigh-energyconditions.Thisincreasesthegroundtrack rangeastheorbiterturnsawayfrom thenearestHeadingAlignmentCircle(HAC) untilsufficientenergyis dissipatedto allowa normalapproachand landingguidancephasecapture,whichbeginsat 10,000feetaltitude.Theorbiteralsocanbeflownnearthevelocityfor maximumlift over drag or wings level for the range stretchcase. The spacecraft slows to subsonic velocity atapproximately 49,000 feet altitude, about 22nautical miles (25.3 statute miles) from the landingsite.
At TAEM acquisition, the orbiter is turneduntil it is aimed at a point tangent to the nearestHAC and continues until it reaches way point 1.At WP-1, the TAEM heading alignment phasebegins. The HAC is followed until landingrunway alignment, plus or minus 20 degrees, hasbeen achieved. In the TAEM pre-final phase, Iheorbiter leaves the HAC; pitches down to acquirethe steep glide slope, increases airspeed; banks toacquire the runway centerline and continues untilon the runway centerline, on the outer glide slope
and on airspeed. The approach and landingguidance phase begins with the completion of theTAEM pre-final phase and ends when thespacecraft comes to a complete stop on the
runway.
The approach and landing trajectory capturephase begins at the TAEM interface and continues
to guidance lock-on to the steep outer glide slope.The approach and landing phase begins at about10,000 feet altitude at an equivalent airspeed of290, plus or minus 12, knots 6.9 nautical miles(7.9 statute miles) from touchdow_n. Autolandguidance is initiated at this point to guide theorbiter to the minus 19- to 17-degree glide slope(which is over seven times that of a commercialairliner's approach) aimed at a target 0.86 nautical
mile (1 statute mile) in front of the runway. Thespacecraft's speed brake is positioned to hold theproper velocity. The descent rate in the laterportion of TAEM and approach and landing isgreater than 10,000 feet per minute (a rate ofdescent approximately 20 times higher than acommercial airliner's standard 3-degree instrumentapproach angle).
At 1,750 feet above ground level, a pre-flaremaneuver is started to position the spacecraft for a1.5-degree glide slope in preparation for landingwith the speed brake positioned as required. Theflight crew deploys the landing gear at this point.
The final phase reduces the sink rate of thespacecraft to less than 9 feet per second.Touchdown occurs approximately 2,500 feet pastthe runway threshold at a speed of 184 to 196knots (213 to 226 mph).
ABORTS. Selection of an ascent abort
mode may become necessary if there is a failurethat affects vehicle performance, such as the failureof a Space Shuttle main engine or an orbitalmaneuvering system. Other failures requiringearly termination of a flight, such as a cabin leak,might require the selection of an abort mode.
There are two basic types of ascent abortmodes for Space Shuttle missions: intact abortsand contingency aborts. Intact aborts are designedto provide a safe return of the orbiter to a plannedlanding site. Contingency aborts are designed topermit flight crew survival following more severfailures when an intact abort is not possible. A
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contingencyabortwouldgenerallyresultin a ditchoperation.
There are four types of intact aborts: Abort toOrbit (ATO), Abort Once Around (AOA),
Transatlantic Landing (TAL) and Return to LaunchSite (RTLS).
The ATO mode is designed to allow thevehicle to achieve a temporary orbit that is lowerthan the nominal orbit. This mode requires lessperformance and allows time to evaluate problemsand then choose either an early deorbit maneuveror an orbital maneuvering system thrustingmaneuver to raise the orbit and continue themission.
The AOA is designed to allow the vehicle tofly once around the Earth and make a normal entryand landing. This mode generally involves twoorbital maneuvering system thrusting sequences,with the second sequence being a deorbitmaneuver. The entry sequence would be similar toa normal entry.
The TAL mode is designed to permit an intactlanding on the other side of the Atlantic Ocean.This mode results in a ballistic trajectory, whichdoes not require an orbital maneuvering systemmaneuver.
The RTLS mode involves flying downrange todissipate propellant and then turning around underpower to retum directly to a landing at or near thelaunch site.
There is a definite order of preference for thevarious abort modes. The type of failure and thetime of the failure determine which type of abort isselected. In cases where performance loss is theonly factor, the preferred modes would be ATO,AOA, TAL and RTLS, in that order. The mode
chosen is the highest one that can be completedwith the remaining vehicle performance. In thecase of some support system failures, such ascabin leaks or vehicle cooling problems, thepreferred mode might be the one that will end themission most quickly. In these cases, TAL orRTLS might be preferable to AOA or ATO. Acontingency abort is never chosen if another abortoption exists.
The Mission Control Center-Houston is prime
for calling these aborts because it has a more
precise knowledge of the orbiter's position thanthe crew can obtain from onboard systems.Before main engine cutoff, Mission Control makesperiodic calls to the crew to tell them which abortmode is (or is not) available. If groundcommunications are lost, the flight crew hasonboard methods, such as cue cards, dedicated
displays and display information, to determine thecurrent abort region.
Which abort mode is selected depends on thecause and timing of the failure causing the abortand which mode is safest or improves missionsuccess. If the problem is a Space Shuttle main
engine failure, the flight crew and Mission ControlCenter select the best option available at the time a
space shuttle main engine fails.
If the problem is a system failure thatjeopardizes the vehicle, the fastest abort mode thatresults in the earliest vehicle landing is chosen.RTLS and TAL are the quickest options (35minutes), whereas an AOA requires approximately90 minutes. Which of these is elected depends on
the time of the failure with three good SpaceShuttle main engines.
The flight crew selects the abort mode by
positioning an abort mode switch and depressingan abort push button.
RETURN TO LAUNCH SITE. The
RTLS abort mode is designed to allow the returnof the orbiter, crew, and payload to the launch site,
Kennedy Space Center, approximately 25 minutesafter lift-off. The RTLS profile is designed toaccommodate the loss of thrust from one spaceshuttle main engine between liftoff and
approximately four minutes 20 seconds, at whichtime not enough main propulsion systempropellant remains to return to the launch site.
An RTLS can be considered to consist of three
stages -- a powered stage, during which the mainengines are still thrusting; an ET separation phase;and the glide phase, during which the orbiterglides to a landing at the KSC. The poweredRTLS phase begins with the crew selection of theRTLS abort, which is done after SRB separation.The crew selects the abort mode by positioning theabort rotary switch to RTLS and depressing theabort push button. The time at which the RTLS isselected depends on the reason for the abort. Forexample, a three-engine RTLS is selected at the
-11-
last moment, approximately 3 minutes, 34 secondsinto the mission; whereas an RTLS chosen due toan engine out at liftoff is selected at the earliesttime, approximately two minutes 20 seconds intothe mission (after SOR separation).
After RTLS is selected, the vehicle continues
downrange to dissipate excess main propulsionsystem propellant. The goal is to leave onlyenough main propulsion system propellant to beable to turn the vehicle around, fly back towardsKSC and achieve the proper main engine cutoffconditions so the vehicle can glide to the KSCafter external tank separation. During thedownrange phase, a pitch-around maneuver isinitiated (the time depends in part on the time of amain engine failure) to orient the orbiter/externaltank configuration to a heads up attitude, pointingtoward the launch site. At this time, the vehicle is
still moving away from the launch site, but themain engines are now thrusting to null thedownrange velocity. In addition, excess orbitalmaneuvering system and reaction control systempropellants are dumped by continuous orbitalmaneuvering system and reaction control systemengine thrustings to improve the orbiter weightand center of gravity for the glide phase andlanding.
The vehicle will reach the desired main enginecutoff point with less than 2 percent excesspropellant remaining in the external tank. At mainengine cutoff minus 20 seconds, a pitch-downmaneuver (called powered pitch-down) takes themated vehicle to the required external tankseparation attitude and pitch rate. After mainengine cutoff has been commanded, the externaltank separation sequence begins, including areaction control system translation that ensures thatthe orbiter does not recontact the external tank and
that the orbiter has achieved the necessary pitchattitude to begin the glide phase of the RTLS.
After the reaction control system translationmaneuver has been completed, the glide phase ofthe RTLS begins. From then on, the RTLS ishandled similarly to a normal entry.
TRANSAII.ANTIC LANDING ABORT.
The TAL abort mode was developed to improvethe options available when a main engine fails afterthe last RTLS opportunity but before the first timethat an AOA can be accomplished with only twomain engines or when a major orbiter system
failure, for example, a large cabin pressure leak orcooling system failure, occurs after the last RTLSopportunity, making it imperative to land asquickly as possible.
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Transatlantic Landing Abort Option
In a TAL abort, the vehicle continues on a
ballistic trajectory across the Atlantic Ocean to landat a predetermined runway. Landing occursapproximately 45 minutes after launch. Thelanding site is selected near the nominal ascentground track of the orbiter in order to make themost efficient use of space shuttle main enginepropellant. The landing site also must have thenecessary runway length, weather conditions andU.S. State Department approval. Currently, thethree landing sites that have been identified for adue east launch are Moron, Spain; Banjul, TheGambia; and Ben Guerir, Morocco.
To select the TAL abort mode, the crew mustplace the abort rotary switch in the TAL/AOAposition and depress the abort push button beforemain engine cutoff. (Depressing it after mainengine cutoff selects the AOA abort mode.) TheTAL abort mode begins sending commands tosteer the vehicle toward the plane of the landingsite. It also rolls the vehicle heads up before mainengine cutoff and sends commands to begin anorbital maneuvering system propellant dump (byburning the propellants through the orbitalmaneuvering system engines and the reactioncontrol system engines). This dump is necessaryto increase vehicle performance (by decreasingweight), to place the center of gravity in the properplace for vehicle control, and to decrease thevehicle's landing weight. TAL is handled like anominal entry.
ABORT TO ORBIT. An ATO is an abortmode used to boost the orbiter to a safe orbital
altitude when performance has been lost and it isimpossible to reach the planned orbital altitude. If
- 12-
aSpaceShuttlemainenginefailsin aregionthatresultsin amainenginecutoffunderspeed,theMissionControlCenterwill determinethat anabortmodeisnecessaryandwill informthecrew.Theorbitalmaneuveringsystemengineswouldbeusedtoplacetheorbiterinacircularorbit.
ABORT ONCE AROUND. The AOAabort mode is used in cases in which vehicle
performance has been lost to such an extent thateither it is impossible to achieve a viable orbit ornot enough Orbital Maneuvering System (OMS)propellant is available to accomplish the OMSthrusting maneuver to place the orbiter on orbitand the deorbit thrusting maneuver. In addition,an AOA is used in cases in which a major systemsproblem (cabin leak, loss of cooling) makes itnecessary to land quickly. In the AOA abortmode, one OMS thrusting sequence is made toadjust the post-main engine cutoff orbit so asecond orbital maneuvering system thrustingsequence will result in the vehicle deorbiting andlanding at the AOA landing site (White Sands,N.M.; Edwards AFB; or KSC). Thus, an AOAresults in the orbiter circling the Earth once andlanding approximately 90 minutes after liftoff.
After the deorbit thrusting sequence has beenexecuted, the flight crew flies to a landing at theplanned site much as it would for a nominal entry.
CONTINGENCY ABORT. Contingencyaborts are caused by loss of more than one mainengine or failures in other systems. Loss of onemain engine while another is stuck at alow thrust
setting may also necessitate a contingency abort.Such an abort would maintain orbiter integrity forin-flight crew escape if a landing cannot beachieved at a suitable landing fidd.
Contingency aborts due to system failuresother than those involving the main engines wouldnormally result in an intact recoveryof vehicle andcrew. Loss of more than one main engine may,depending on engine failure times, result in a saferunway landing. However, in most three-engine-out cases during ascent, the orbiter would have tobe ditched. The in-flight crew escape systemwould be used before ditching the orbiter.
ORBITER GROUND TURNAROUND
Spacecraft recovery operations at the nominalend-of-mission landing site are supported by
approximately 160 Space Shuttle launch operationsteam members. Ground team members wearingself-contained atmospheric protective ensemblesuits that protect them from toxic chemicalsapproach the spacecraft as soon as it stops rolling.The ground team members take sensormeasurements to ensure the atmosphere in thevicinity of the spacecraft is not explosive. In the
event of propellant leaks, a wind machine truckcarrying a large fan will be moved into the area tocreate a turbulent airflow that will break up gasconcentrations and reduce the potential for anexplosion.
A ground support equipment air-conditioningpurge unit is attached to the right-hand orbiter T-0umbilical so cool air can be directed through theorbiter's aft fuselage, payload bay, forwardfuselage, wings, vertical stabilizer, and orbitalmaneuvering system/reaction control system podsto dissipate the heat of entry.
A second ground support equipment groundcooling unit is connected to the left-hand orbiter T-O umbilical spacecraft Freon Coolant loops toprovide cooling for the flight crew and avionicsduring the postlanding and system checks. Thespacecraft fuel cells remain powered up at thistime. The flight crew will then exit the spacecraft,and a ground crew will power down thespacecraft.
AT KSC, the orbiter and ground supportequipment convoy move from the runway to theOrbiter Processing Facility.
If the spacecraft lands at Edwards, the same
procedures and ground support equipment areused as at the KSC after the orbiter has stopped onthe runway. The orbiter and ground supportequipment convoy move from the runway to the
orbiter mate and demate facility at Edwards. Afterdetailed inspection, the spacecraft is prepared to beferried atop the Shuttle carrier aircraft fromEdwards to KSC. For ferrying, a tail cone isinstalled over the aft section of the orbiter.
In the event of a landing at an alternate site, acrew of about eight team members will move tothe landing site to assist the astronaut crew in
preparing the orbiter for loading aboard the Shuttlecarder aircraft for transport back to the KSC. Forlandings outside the United States, personnel at thecontingency landing sites will be provided
- 13-
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- 14-
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-15-
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- 16-
minimum training on safe handling of the orbiterwith emphasis on crash rescue training, how totow the orbiter to a safe area, and prevention of
propellant conflagration.
Upon its return to the Orbiter ProcessingFacility (OPF) at KSC, the orbiter is safed(ordnance devices safed), the payload (if any) isremoved, and the orbiter payload bay is
reconfigured from the previous mission for thenext mission. Any required maintenance andinspections are also performed while the orbiter isin the OPF. A payload for the orbiter's nextmission may be installed in the orbiter's payloadbay in the OPF or may be installed in the payloadbay when the orbiter is at the launch pad.
The spacecraft is then towed to the VehicleAssembly Building and mated to the external tank.The extema/ tank and solid rocket boosters are
stacked and mated on the mobile launcher platformwhile the orbiter is being refurbished. SpaceShuttle orbiter connections are made and the
integrated vehicle is checked and ordnance isinstalled.
The mobile launcher platform moves the entirespace shuttle system on four crawlers to the launchpad, where connections are made and servicingand checkout activities begin. If the payload wasnot installed in the OPF, it will be installed at the
launch pad followed by prelaunch activities.
Space Shuttle launches from Vandenberg willuse the Vandenberg Launch Facility (SL6), whichwas built but never used for the manned orbital
laboratory program. This facility was modifiedfor Space Transportation System use.
The runway at Vandenberg was strengthenedand lengthened from 8,000 feet to 12,000 feet toaccommodate the orbiter returning from space.
When the orbiter lands at Vandenberg, thesame procedures and ground support equipmentand convoy are used as at KSC after the orbiterstops on the runway. The orbiter and groundsupport equipment are moved from the runway tothe Orbiter Maintenance and Checkout Facility at
Vandenberg. The orbiter processing proceduresused at this facility are similar to those used at theOPF at the KSC.
Space Shuttle buildup at Vandenberg differsfrom that of the KSC in that the vehicle is
integrated on the launch pad. The orbiter is towedoverland from the Orbiter Maintenance and
Checkout Facility at Vandenberg to launch facilitySL6.
SL6 includes the launch mount, access tower,mobile service tower, launch control tower,
payload preparation room, payload changeoutroom, solid rocket booster refurbishment facility,solid rocket booster disassembly facility, andliquid hydrogen and liquid oxygen storage tankfacilities.
The SRB start the on-the-launch-pad buildupfollowed by the external tank. The orbiter is thenmated 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 andlaunch control ground activities until the SpaceShuttle vehicle clears the launch pad tower.Responsibility is then turned over to MissionControl Center-Houston. The Mission Control
Center's responsibility includes ascent, on-orbitoperations, entry, approach and landing untillanding runout completion, at which time theorbiter is handed over to the postlandingoperations at the landing site for turnaround andre-launch. At the launch site the SRBs and
external tank are processed for launch and theSRBs are recycled for reuse.
- 17-
OPERATIONAL IMPROVEMENTS ANDMODIFICATIONS
Many of the changes and upgrades in theSpace Shuttle systems and components were underway before the 51-L accident as part of NASA'scontinual improvement and upgrade program.However, NASA has taken advantage of the SpaceShuttle program downtime since the accident toaccelerate the testing and integration of theseimprovements and upgrades as well as fixesrequired as a result of the accident.
ORBITER. The following identifies themajor improvements or modifications of theorbiter. Approximately 190 other modificationsand improvements were also made.
ORBITAL MANEUVERING SYSTEMAND REACTION CONTROL SYSTEMAC-MOTOR-OPERATED VALVES. The
64 valves operated by AC-motors in the OMS andRCS were modified to incorporate a "sniff' linefor each valve to permit monitoring of nitrogentetroxide or monomethyl hydrazine in the electricalportion of the valves during ground operations.This new line reduces the probability of floatingparticles in the electrical microswitch portion ofeach valve, which could affect the operation of themicroswitch position indicators for onboarddisplays and telemetry. It also reduces theprobability of nitrogen tetroxide or monomethyihydrazine 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 waswrapped around each of the 38 primary RCSthrust chambers to remove electrical power fromthese valves in the event of a primary RCS thrusterinstability.
FUEL CELL POWER PLANTS. End-
cell heaters on each fuel cell power plant weredeleted because of potential electrical failures andreplaced with Freon coolant loop passages tomaintain uniform temperature throughout thepower plants. In addition, the hydrogen pump andwater separator of each fuel cell power plant wereimproved to minimize excessive hydrogen gasentrained in the power plant product water. Acurrent 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 foreach fuel cell power plant was modified to preventoverheating and loss of heater elements. A stackinlet temperature measurement was added to eachfuel cell power plant for full visibility of thermalconditions.
The product water from all three fuel cellpower plants flows to a single water relief controlpanel. The water can be directed from the singlepanel to the Environmental Control and LifeSupport System (ECLSS) potable water tank A orto 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 allthree water relief panel lines causing them to freezeand preventing water discharge.
The product water lines from all three fuel cellpower plants were modified to incorporate aparallel (redundant) path of product water toECLSS potable water tank B in the event of afreeze-up in the single water relief panel. If the
single water relief panel freezes up, pressurewould build up and discharge through theredundant paths to water tank B.
A water purity sensor (pH) was added at thecommon product water outlet of the water reliefpanel to provide a redundant measurement of waterpurity (a single measurement of water purity ineach fuel cell power plant was providedpreviously). If the fuel cell power plant pH sensorfailed in the past, the flight crew had to sample thepotable water.
AUXILIARY POWER UNITS. TheAPUs that have been in use to date have a limitedlife. Each unit was refurbished after 25 hours of
operation because of cracks in the turbine housing,degradation of the gas generator catalyst (whichvaried up to approximately 30 hours of operation)and operation of the gas generator valve module(which also varied up to approximately 30 hoursof operation). The remaining parts of the APUwere qualified for 40 hours of operation.
Improved APUs are scheduled for delivery inlate 1988. A new turbine housing increases thelife of the housing to 75 hours of operation (50
-18-
missions);anewgasgeneratorincreasesitslife to75 hours;a new standoffdesignof the gasgeneratorvalvemoduleandfuelpumpdeletestherequirementfor a waterspraysystemthatwasrequiredpreviouslyforeachAPUuponshutdownafterthe first OMSthrustingperiodor orbitalcheckout;andtheadditionof athirdsealin themiddleof thetwoexistingsealsfortheshaftofthefuelpump/lubeoil system(previouslyonly twosealswerelocatedon theshaft,oneon thefuelpumpsideandoneon thegearboxlubeoil side)reducestheprobabilityof hydrazineleakingintothelubeoil system.
Thedeletionof thewaterspraysystemforthegasgeneratorvalvemoduleandfuelpumpforeachAPU results in a weight reduction ofapproximately150poundsfor eachorbiter.Uponthedeliveryof theimprovedunits,thelife-limitedAPUs 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 anelectrical short occurs within the valve electrical
coil, additional protection is provided to preventoverheating of the fuel isolation valves.
MAIN LANDING GEAR. The followingmodifications were made to improve the
performance of the main landing gear elements:
•The thickness of the main landing gear axlewas increased to provide a stiffer configurationthat reduces brake-to-axle deflections and
precludes brake damage experienced inprevious landings. The thicker axle shouldalso minimize tire wear.
•Orifices were added to hydraulic passages inthe brake's piston housing to prevent pressuresurges and brake damage caused by awobble/pump effect.
•The electronic brake control boxes were
modified to balance hydraulic pressurebetween adjacent brakes and equalize energyapplications. The anti-skid circuitrypreviously used to reduce brake pressure tothe opposite wheel if a flat tire was detectedhas now been removed.
•The carbon-lined beryllium stator discs ineach main landing gear brake were replaced
with thicker discs to increase braking energy
significantly.
•A long-term structural carbon brake programis in progress to replace the carbon-linedberyllium stator discs with a carbonconfiguration that provides higher brakingcapacity by increasing maximum energyabsorption.
•Strain gauges were added to each nose andmain landing gear wheel to monitor tirepressure before launch, deorbit and landing.
Other studies involve arresting barriers at theend of landing site runways (except lakebedrunways), installing a skid on the landing gear thatcould preclude the potential for a second blowntire on the same gear after the first tire has blown,providing "roll on rim" for a predictable roll if bothtires are lost on a single or multiple gear andadding a drag chute.
Studies of landing gear tire improvements arebeing conducted to determine how best to decreasetire wear observed after previous KSC landingsand how to improve crosswind landing capability.
Modifications were made to the KSC Shuttle
Landing Facility runway. The full 300-foot widthof 3,500-foot sections at both ends of the runwaywere ground to smooth the runway surface textureand remove cross grooves. The modifiedcorduroy ridges are smaller than those theyreplaced and run the length of the runway ratherthan across its width. The existing landing zonelight fixtures were also modified, and the markingsof the entire runway and overruns were repaintcd.The primary purpose of the modifications is toenhance safety by reducing tire wear duringlanding.
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-speedengagement of the nose wheel steering system andprovides positive lateral directional control of theorbiter during rollout in the presence of highcrosswinds and blown tires.
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THERMAL PROTECTION SYSTEM.The area aft of the reinforced carbon-carbon nose
cap to the nose landing gear doors has sustaineddamage (tile slumping) during flight operationsfrom impact during ascent and overheating duringreentry. This area, which previously was coveredwith high-temperature reusable surface insulationtiles, will now be covered with reinforced carbon-carbon.
The low-temperature thermal protectionsystem tiles on Columbia's midbody, payload baydoors and vertical tail were replaced with advancedFlexible Reusable Surface ilnsulation (FRSI)blankets.
Because of evidence of plasma flow on thelower wing trailing edge and elevon landing edgetiles (wing/elevon cove) at the outboard elevon tipand inboard elevon, the low-temperature tiles arebeing replaced with Fibrous Refractory CompositeInsulation (FRCI-12) and High-Temperature
(HRSI-22) tiles along with gap fillers onDiscovery and Atlantis. On Columbia only gapfiUers are installed in this area.
WING MODIFICATION. Before the
wings for Discovery and Atlantis weremanufactured, a weight reduction program wasinstituted that resulted in a redesign of certain areasof the wing structure. An assessment of wing airloads from actual flight data indicated greater loadson the wing structure than predicted. To maintainpositive margins of safety during ascent, structuralmodifications were incorporated into certain areasof the wings.
MID-FUSELAGE MODIFICATIONS.
Because of additional detailed analysis of actualflight data concerning descent-stress thermal-gradient loads, torsional straps were added to tieall the lower mid-fuselage stringers in bays 1through 11 together in a manner similar to a boxsection. This eliminates rotational (torsional)capabilities to provide positive margins of safety.
Also, because of the detailed analysis of actual
descent flight data, room-temperature vulcanizingsilicone rubber material was bonded to the lower
mid-fuselage from bays 4 through 11 to act as aheat sink, distributing temperatures evenly acrossthe bottom of the mid-fuselage, reducing thermalgradients and ensuring positive margins of safety.
GENERAL PURPOSE COMPUTERS.
New upgraded General Purpose Computers(GPC), IBM AP-101S, will replace the existingGPCs aboard the Space Shuttle orbiters in late1988 or early 1989. The upgraded computersallow NASA to incorporate more capabilities intothe orbiters and apply advanced computertechnologies that were not available when theorbiter was first designed. The new computerdesign began in January 1984, whereas the olderdesign began in January 1972. The upgradedGPCs provide two-and-a-half times the existingmemory capacity and up to three times the existingprocessor speed with minimum impact on flightsoftware. They are half the size, weighapproximately half as much, and require lesspower to operate.
INERTIAL MEASUREMENT UNITS.
The new High-Accuracy Inertial NavigationSystem (HAINS) will be phased in in 1988-89 toaugment the present KT-70 inertial measurement
- 20 -
units. ThesenewInertial MeasurementUnits(IMUs)will resultin lowerprogramcostsover thenext decade, ongoing production support,improved performance, lower failure rates andreduced size and weight. The HAINS IMUs alsocontain an internal dedicated microprocessor withmemory for processing and storing compensationand scale factor data from the IMU manufacturer's
calibration, thereby reducing the need for extensiveinitial load data for the orbiter's computers. TheHAINS is both physically and functionallyinterchangeable with the KT-70 IMU.
CREW ESCAPE SYSTEM. The in-flightcrew escape system is provided for use only whenthe orbiter is in controlled gliding flight and unableto reach a runway. This would normally lead toditching. The crew escape system provides theflight crew with an alternative to water ditching orto landing on terrain other than a landing site. Theprobability of the flight crew surviving a ditchingis very small.
The hardware changes required to the orbiterswould enable the flight crew to equalize thepressurized crew compartment with the outsidepressure via a depressurization valve opened bypyrotechnics in the crew compartment aft bulkheadthat would be manually activated by a flight crewmember in the middeck of the crew compartment;pyrotechnically jettison the crew ingress/egressside hatch in the middeck of the crewcompartment; and bail out from the middeck of the
orbiter through the ingress/egress side hatchopening after manually deploying the escape polethrough, outside and down from the side hatch
opening. One by one, each crew member attachesa lanyard hook assembly, which surrounds filedeployed escape pole, to his parachute harness andegresses through the side hatch opening. Attachedto the escape pole, the crew member slides downthe pole and off the end. The escape pole providesa trajectory that takes the crew members below theorbiter's left wing.
Changes were also made in the software of theorbiter's general purpose computers. Thesoftware changes were required for the primaryavionics software system and the backup flightsystem for transatlantic-landing and glide-return-to-launch-site aborts. The changes: provide theorbiter with an automatic-mode input by the flightcrew through keyboards on the commander's
and/or pilot's panel C3, which provides the orbiterwith an automatic stable flight for crew bailout.
The side hatch jettison feature also could beused in a landing emergency.
EMERGENCY EGRESS SLIDE. The
emergency egress slide provides orbiter flight crewmembers with a means for rapid and safe exitthrough the orbiter middeck ingress/egress sidehatch after a normal opening of the side hatch orafter jettisoning the side hatch at the nominal end-of-mission landing site or at a remote oremergency landing site.
The emergency egress slide replaces theemergency egress side hatch bar, which requiredthe flight crew members to drop approximately10.5 feet to the ground. The previous arrangementcould have injured crew members or prevented analready-injured crew member from evacuating andmoving a safe distance from the orbiter.
I"/-INCH ORBIIER/EXTERNAL TANK
DISCONNECIS. Each mated pair of 17-inchdisconnects contains two flapper valves: one onthe orbiter side and one on the external tank side.
Both valves in each disconnect pair are opened topermit propellant flow between the orbiter and theexternal tank. Prior to separation from the externaltank, both valves in each mated pair of disconnectsare commanded closed by pneumatic (helium)pressure from the main propulsion system. Theclosure of both valves in each disconnect pairprevents propellant discharge from the externaltank or orbiter at external tank separation. Valveclosure on the orbiter side of each disconnect also
prevents contamination of the orbiter mainpropulsion system during landing and groundoperations.
Inadvertent closure of either valve in a 17-inch
disconnect during main engine thrusting wouldstop propellant flow from the external tank to allthree main engines. Catastrophic failure of themain engines and external tank feed lines wouldresult.
To prevent inadvertent closure of the 17-inchdisconnect valves during the Space Shuttle mainengine thrusting period, a latch mechanism wasadded in each orbiter half of the disconnect. The
latch mechanism provides a mechanical backup tothe normal fluid-induced-open forces. The latch is
-21 -
mountedon ashaft in the flowstream so that it
overlaps both flappers and obstructs closure forany reason.
In preparation for external tank separation,both valves in each 17-inch disconnect are
commanded closed. Pneumatic pressure from themain propulsion system causes the latch actuatorto rotate the shaft in each orbiter 17-inch
disconnect 90 degrees, thus freeing the flappervalves to close as required for external tankseparation.
A backup mechanical separation capability isprovided in case a latch pneumatic actuatormalfunctions. When the orbiter umbilical initiallymoves away from the ET umbilical, the mechanicallatch disengages from the ET flapper valve andpermits the orbiter disconnect flapper to toggle thelatch. This action permits both flappers to close.
SPACE SHUTTLE MAIN ENGINEMARGIN IMPROVEMENT PROGRAM
Improvements to the Space Shuttle MainEngines (SSMEs) for increased margin anddurability began with a formal Phase II program in1983. Phase II focused on turbo-machinery toextend the time between high-pressure turbopumpoverhauls by reducing the operating temperature inthe high-pressure fuel turbopump and byincorporating margin improvements to the HighPressure Fuel Turbopump (HPFT) rotordynamics (whirl), turbine blade and HPFTbeatings. Phase II certification was completed in1985, and all the changes have been incorporatedinto the SSMEs for the STS-26 mission.
In addition to the Phase II improvements,additional changes in the SSME have beenincorporated to further extend the engines' marginand durability. The main changes were to thehigh-pressure turbo-machinery, main combustionchamber, hydraulic actuators and high-pressureturbine discharge temperature sensors. Changeswere also made in the controller software to
improve engine control.
Minor high-pressure turbo-machinery designchanges resulted in margin improvements to the
turbine blades, thereby extending the operating lifeof the turbopumps. These changes includedapplying surface texture to important parts of the
fuel turbine blades to improve the materialproperties in the pressure of hydrogen andincorporating a damper into the high-pressureoxidizer turbine blades to reduce vibration.
Main combustion chamber life has been
increased by plating a welded outlet manifold withnickel. Margin improvements have also beenmade to five hydraulic actuators to preclude a lossin redundancy on the launch pad. Improvementsin quality have been incorporated into the servo-component coil design along with modifications toincrease margin. To address a temperature sensorin-flight anomaly, the sensor has been redesignedand extensively tested without problems.
To certify the improvements to the SSMEs anddemonstrate their reliability through margin (orlimit testing), an aggressive ground test programwas initiated in December 1986. From December
1986 to December 1987, 151 tests and 52.363
seconds of operation (equivalent to 100 Shuttlemissions) were performed. The SSMEs haveexceeded 300,000 seconds total test time, theequivalent of 615 Space Shuttle missions. Thesehot-fire ground tests are performed at the single-engine test stands NASA's Stennis Space Centerin Mississippi and at Rockwell International'sRocketdyne Division's Santa Susana FieldLaboratory 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. Oneengine (serial number 2012) has been flown 10times; 10 other engines have flown between fiveand nine times. Two off-nominal conditions were
experienced on the launch pad and one duringflight. Two fail-safe shutdowns occurred on thelaunch pad during engine start but before SRBignition. In each case, the controller detected aloss of redundancy in the hydraulic actuatorsystem and commanded engine shutdown inkeeping with the launch commit criteria. Anotherloss of redundancy occurred in flight with a lossof a red-line temperature sensor and its backup.The engine was commanded to shut down, but theother two engines safely delivered the SpaceShuttle to orbit. A major upgrade of thesecomponents was implemented to prevent a
-22-
recurrenceof theseconditions and willincorporated for STS-26.
10
9
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SSME Flight Experience DemonstratesReliability (Through February 1, 1986)
SOLID ROCKET MOTOR REDESIGN
be
On June 13, 1986, President Reagan directedNASA to implement, as soon as possible, therecommendations of the "Presidential Commission
on the Space Shuttle Challenger Accident." NASAdeveloped a plan to provide a Redesigned SolidRocket Motor (RSRM). The primary objective ofthe redesign effort was to provide an SRM that issafe to fly. A secondary objective was to
minimize impact on the schedule by using existinghardware, to the extent practical, withoutcompromising safety. A joint redesign team wasestablished that included participation fromMarshall Space Flight Center, Morton Thiokol andother NASA centers as well as individuals fromoutside NASA.
An "SRM Redesign Project Plan" wasdeveloped to formalize the methodology for SRMredesign and requalification. The plan provided anoverview of the organizational responsibilities andrelationships, the design objectives, criteria andprocess; the verification approach and process; anda master schedule. The companion "Developmentand Verification Plan" defined the test program andanalyses required to verify the redesign and theunchanged components of the SRM.
All aspects of the existing SRM wereassessed, and design changes were required in thefield joint, case-to-nozzle joint, nozzle, factoryjoint, propellant grain shape, ignition system andground support equipment. No changes were
made in the propellant, liner or castable inhibitorformulations. Design criteria were established foreach component to ensure a safe design with anadequate margin of safety. These criteria focusedon loads, environments, performance,redundancy, margins of safety and verificationphilosophy.
The criteria were converted into specificdesign requirements during the PreliminaryRequirements Reviews held in July and August1986. The design developed from theserequirements was assessed at the PreliminaryDesign Review held in September 1986 andbaselined in October 1986. The final design wasapproved at the Critical Design Review held inOctober 1987. Manufacture of the RSRM test
hardware and the first flight hardware began priorto the Preliminary Design Review (PDR) andcontinued in parallel with the hardware certificationprogram. The Design Certification Review willreview the analyses and test results versus theprogram and design requirements to certify theredesigned SRM is ready to fly.
ORIGINAL VERSUS REDESIGNED
SRM FIELD JOINT• The SRM field-jointmetal parts, internal case insulation and seals wereredesigned and a weather protection system wasadded.
In the STS 51-L design, the application ofactuating pressure to the upstream face of the O-ring was essential for proper joint sealingperformance because large sealing gaps werecreated by pressure-induced deflections,compounded by significantly reduced O-ringsealing performance at low temperature. Themajor change in the motor case is the new tangcapture feature to provide a positive metal-to-metalinterference fit around the circumference of the
tang and clevis ends of the mating segments. Theinterference fit limits the deflection between the
tang and clevis O-ring sealing surfaces caused bymotor pressure and structural loads. The joints aredesigned so that the seals will not leak under twicethe 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-ringsealing gap dimension. The sealing gap and theO-ring seals are designed so that a positive
- 23 -
ORIGINAL VERSUS REDESIGNED SRM FIELD JOINT
FactoryJoint Insulation
'_ Nozz;e-to-C__
Equ_ment /"_AssemblyAids Nozzle/
Solid Rocket Booster Redesignand Reassessment
compression (squeeze) is always on the O-rings.The minimum and maximum squeeze requirementsinclude the effects of temperature, O-ringresiliency and compression set, and pressure. Theclevis O-ring groove dimension has been increasedso that the O-ring never fills more than 90 percentof the O-ring groove and pressure actuation isenhanced.
The new field joint design also includes a newO-ring in the capture feature and an additional leak
check port to ensure that the primary O-ring ispositioned in the proper sealing direction atignition. This new or third O-ring also serves as athermal barrier in case the sealed insulation isbreached.
The field joint internal case insulation was
modified to be sealed with a pressure-actuated flapcalled a J-seal, rather than with putty as in the STS51-L configuration.
Longer field-joint-case mating pins, with areconfigured retainer band, were added to improvethe shear strength of the pins and increase the
metal parts' joint margin of safety. The jointsafety margins, both thermal and structural, are
being demonstrated over the full ranges of ambienttemperature, storage compression, grease effect,assembly stresses and other environments.
Extemal heaters with integral weather seals wereincorporated to maintain the joint and O-ringtemperature at a minimum of 75 F. The weatherseal also prevents water intrusion into the joint.
Sindleritim
Forward ForwerdlP'_ -Vent Port
Tang CaptureFeatureTang
_:]:[eak CheckPortl_l_¢'_A_d_C----Sec_ondaryO-Ring_ , z CaptureFeatureO-Ring
_ Zinc ChromateJt_IT_z4r_IE%==__ Joint Heater
_--"-'-- Pins__ _T" Pressure'Sensit_ve
Clevisis "_ "Custom Shims
. '"longer Pins_New Pin RetainerBand
Aft
Odgin,,I Pdidesiped SolidField Joint Design Pocket Motor Imprevements
Field Joint Comparison
ORIGINAL VERSUS REDESIGNEDSRM CASE-TO-NOZZLE JOINT. The
SRM case-to nozzle joint, which experiencedseveral instances of O-ring erosion in flight, hasbeen redesigned to satisfy the same requirementsimposed upon the case field joint. Similar to thefield joint, cast-to-nozzle joint modifications havebeen made in the metal parts, internal insulationand 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. Thethird O-ring serves as a dam or wiper in front ofthe primary O-ring to prevent the polysulfideadhesive from being extruded into the primary O-ring groove. It also serves as a thermal barrier in
case the polysulfide adhesive is breached. Thepolysulfide adhesive replaces the putty used in the51-L joint. Also, an additional leak check port
was added to reduce the amount of trapped air inthe joint during the nozzle installation process andto aid in the leak check procedure.
NOZZLE. The internal joints of the nozzle
metal parts have been redesigned to incorporateredundant and verifiable O-rings at each joint. Thenozzle steel fixed housing part has beenredesigned to permit the incorporation of the 100radial bolts that attach the fixed housing to the
case's aft dome. Improved bonding techniquesare being used for the nozzle nose inlet, cowl/bootand aft exit cone assemblies. The distortion of the
nose inlet assembly's metal-part-to-ablative-partsbond line has been eliminated by increasing thethickness of the aluminum nose inlet housing and
- 24 -
improvingthebondingprocess.Thetape-wrapangleof the carbon cloth fabric in the areas of thenose inlet and throat assembly parts was changedto improve the ablative insulation erosiontolerance. Some of these ply-angle changes werein progress prior to STS 51-L. The cowl andouter boot ring has additional structural supportwith increased thickness and contour changes toincrease their margins of safety. Additionally, theouter boot ring ply configuration was altered.
GrNze Bead CorkImulation EA 934 AdhesivezOrigial FkdklJmt Ihmtgn
RedundantVent Valve(m 45 Nd 135 Degreed _ ExtrudedCork
\ \ Adhe_m / W'ahTl_lmlTlpe
W'_h TeflonTape AMation..._,Cmlel_Wm
or Adhesive
Adhe_ve i ! Cork
Red_igmKI Solid II_lketMetsr Field Joiad
Field Joint Protection System
Strm_el'mf FlapS _ /,. in Aft Dome .,,.
_\\ tl _Y" _ J _[ /L_" _W'q_r firing
......:..I I-_ L''Primary O'Riog_Vent Port
I-_ s_ o.ai__ 143D_ru
Axiol Bolts--'-="_[_ \ "S,ot-D-SIol=Lock CheckPort(90 Oe_'a Locotion)
0rigiMI Cue-to-Nozzle Jelnt Oml_pz P,_luilmd Cmm-to-IIo,lz_Joint Design
Case-to-Nozzle Joint Design
3 2 4 5
Joint 1 - Forwardto Aft Exit Coao
Jo_ 2 - Nose-inletAssemblyto Flex Bem'_Joint 3 - Nose-inletAssemSlyto Throat AssemSlyJoint 4 - Throet Assemblyto Fm'werdExit CramJoint 5 - Ftzed HousingAssemblyto Flex Bearing
Redesigned Solid Rocket MotorNozzle Internal Seals
MoistureSealAdded
LongerPinsRetentionBandAdded
j InsulationRy Layup
and Thickness
Changed
O-R_ Size
a_i GrooveChanged
Redesigned Factory Joint Nozzle-to-Case Joint
n
- 25 -
FACTORY JOINT. Minor modifications
were made in the case factory joints by increasingthe insulation thickness and lay-up to increase themargin of safety on the internal insulation. Longerpins were also added, along wit a reconfiguredretainer band and new weather seal to improvefactory joint performance and increase the marginof safety. Additionally, the O-ring and O-ringgroove size was changed to be consistent with thefield joint.
PROPELLANT. The motor propellantforward transition region was recontoured toreduce the stress fields between the star and
cylindrical portions of the propellant grain.
IGNITION SYSTEM. Several minor
modifications were incorporated into the ignitionsystem. The aft end of the igniter steel case,which contains the igniter nozzle insert, wasthickened to eliminate a localized weakness. The
igniter internal case insulation was tapered toimprove the manufacturing process. Finally,although vacuum putty is still being used at thejoint of the igniter and case forward dome, it waschanged to eliminate asbestos as one of itsconstituents.
GROUND SUPPORT EQUIPMENT.
The ground support equipment has beenredesigned to (1) minimize the case distortionduring handling at the launch site; (2) improve thesegment tang and clevis joint measurement systemfor more accurate reading of case diameters tofacilitate stacking; (3) minimize the risk of O-ring
Round.Out/
Ring
Round.Out
Ring_
Initial
Engagement
i--
CaplureTang CompleteEngagement Engagemenl
i
Ground Support Equipment Assembly Aids
damage during joint mating; and (4) improve leaktesting of the igniter, case and nozzle field joints.A Ground Support Equipment (GSE) assembly aidguides the segment tang into the clevis and roundsthe two parts with each other. Other GSEmodifications include transportation monitoringequipment and lifting beam.
DESIGN ANALYSIS SUMMARY.
Improved, state-of-the-art, analyses related tostructural strength, loads, stress, dynamics,fracture mechanics, gas and thermal dynamics, andmaterial characterization and behavior were
performed to aid the field joint, nozzle-to-case jointand other designs. Continuing these analyses willensure that the design integrity and system
compatibility adhere to design requirements andoperational use. These analyses will be verified bytests, whose results will be correlated with pretestpredictions.
VERIFICATION/CERTIFICATION
TEST. The verification program demonstratesthat the RSRM meets all design and performancerequirements, and that failure modes and hazardshave been eliminated or controlled. The
verification program encompasses the followingprogram phases: development, certification,acceptance, preflight checkout, flight andpostflight.
Redesigned SRM certification is based on
formally documented results of developmentmotor tests; qualification motor tests and othertests and analyses. The certification tests areconducted under strict control of environments,
including thermal and structural loads; assembly,inspection and test procedures; and safety,reliability, maintainability and quality assurancesurveillance to verify that flight hardware meetsthe specified performance and design
requirements. The "Development and VerificationPlan" stipulates the test program, which follows arigorous sequence wherein successive tests buildon the results of previous tests leading to formal
The test activities include laboratory andcomponent tests, subscale tests, full-scalesimulation and full-scale motor static test firings.Laboratory and component tests are used todetermine component properties and
characteristics. Subscale motor firings are used tosimulate gas dynamics and thermal conditions forcomponents and subsystem design. Full-scale
- 26 -
hardwaresimulatorsareusedto verifyanalyticalmodels; determine hardware assemblycharacteristics; determine joint deflectioncharacteristics;determinejoint performanceundershort-durationhot-gastests,includingjoint flawsand flight loads; and determineredesignedhardwarestructuralcharacteristics.
Fourteen full-scale joint assemblydemonstrationverticalmate/dematetests,witheightinterspersedhydroteststo simulateflighthardware refurbishment procedures,werecompletedearlyfortheredesignedcapture-featurehardware.Assemblyloadswereasexpected,andthe casegrowth was as predictedwith nomeasurableincreaseafterthreehydro-prooftests.
Hight-configurationaft andcentersegmentswerefabricated,loadedwith live propellant,andused for assembly test article stackingdemonstrationtestsat KennedySpaceCenter.Thesetestswerepathfinderdemonstrationsfortheassembly of flight hardware using newlydevelopedgroundsupportequipment.
In a long-termstacktest,a full-scalecastingsegment,with live propellant,hasbeenmatedverticallywithaJ-sealinsulationsegmentandisundergoingtemperaturecycling. This willdeterminethecompressionsetof theJ-seal,agingeffectsandlong-termpropellantslumpingeffects.
The Structural Test Article (STA-3),consistingof flight-typeforwardandaft motorsegmentsand forward and aft skirts, wassubjectedtoextensivestaticanddynamicstructuraltesting,includingmaximumprelaunch,liftoff andflight (maximumdynamicpressure)structuralloads.
RedesignedSRMcertificationincludestestingtheactualflightconfigurationoverthefull rangeofoperatingenvironmentsandconditions.Thejointenvironmentsimulator,transient pressure testarticle, and the nozzle joint environment simulatortest programs all utilize full-scale flight designhardware and subject the RSRM design features tothe maximum expected operating pressure,maximum pressure rise rate and temperatureextremes during ignition tests. Additionally, theTransient Pressure Test Article (TPTA) is
subjected to ignition and liftoff loads as well asmaximum dynamic pressure structural loads.
_es m-L:
Purge and Quench ._-
•. .... . Pr_.-_t.
Iltotat Cold}
ExternalTankAttachLoads
3:i I-i_!
I
[
1-MJion-PoundDeadweight Simulated
f VehicleLoad
j Mndir_d ForwardSkirt
_- t0nit_ Joint
FieldJoint* RedesignedSolid Rocket
Motor (RSRM)
_,i Field Joint• RSRM CaseHardware
and InsulationDesign
Coce-to-NozzleJoint
• BaselineRSRMCase
__ end Insulation
_ _t and Relief
Transient Pressure Test Article Configuration
Four TPTA tests have been completed tosubject the redesigned case field and case-to-nozzlejoints to the above-described conditions. The fieldand case-to-nozzle joints were temperature-conditioned to 75 F. and contained various typesof flaws in the joints so that the primary andsecondary O-rings could be pressure-actuated,joint rotation and O-ring performance could beevaluated and the redesigned joints could bedemonstrated as fail safe.
Six of the seven Joint Environment Simulators
(JES) tests have been completed. The JES testprogram initially used the STS 51-L configurationhardware to evaluate the joint performance withprefabricated blowholes through the putty. TheJES-1 test series, which consisted of two tests,established a structural and performance data basefor the STS 51-L configuration with and without areplicated joint failure. The JES-2 series, twotests, also used theSTS 51-L case metal-part jointbut with a bonded labyrinth and U-seal insulationthat was an early design variation of the J-seal.Tests were conducted with and without flaws built
into the U-seal joint insulation; neither jointshowed O-ring erosion or blow-by. The JES-3series, three tests, uses almost exact flightconfiguration hardware, case field-joint capturefeature with interference fit and J-seal insulation.
- 27 -
Freed Cartrid_ Onsi_ _1ThinPropeflantSlabs
F'mklJoint
• RedesignedSolidRocketMotor CaseHardwareend InsulationDesign
#
• Redesigned_ Rocket L_
Motor CaseHardware _ iand InsulationDesign \'_
_ght)Ig_w[With OuenchPort)
ForwardDome
CylinderLightweightWith Captme Future
PropellantCartridge
Grain
Propellant
Cy}_JerLightweightWith CaptureFeature
Propellent
Attach Segment
Segment
Port Dome
ClosedVent WithVent onCommandCapability
Joint Environment Simulation Configuration
Four of five nozzle JES tests have been
successfully conducted. The STS 51-L hardwareconfiguration hydro test confirmed predicted case-to-nozzle-joint deflection. The other three testsused the radially bolted RSRM configuration.
RadiallyBoltedCase.to.NozzleJe_t,
PTopeflantSlobs(BothDomes and
FixedHousing)
t_
vA ,
..--Top Assembly
/Piston Assembly
/Vents [4)
i Fixed Housing
..-Sleeve Assembly
...--Aft DomeAssembly
....-Standard Igniter
ForwardDomeAssembly
QuenchInjector
Nozzle Joint Environment Simulator Configuration
Seven full-scale, full-duration motor static
tests are being conducted to verify the integratedRSRM performance. These include oneengineering test motor used to (1) provide a database for STS 5l-L-type field joints; (2) evaluate
new seal material; (3) evaluate the ply-anglechange in the nozzle parts,; (4) evaluate theeffectiveness of graphite composite stiffener ringsto reduce joint rotation; and (5) evaluate field-jointheaters. There were two development motor testsand three qualification motor tests for final flightconfiguration and performance certification. Therewill be one flight Production Verification Motorthat contains intentionally induced defects in thejoints to demonstrate joint performance underextreme worse case conditions. The QM-7 andQM-8 motors were subjected to liftoff andmaximum dynamic pressure structural loads, QM-7 was temperature-conditioned to 90 F., and QM-8was temperature-conditioned to 40 F.
An assessment was conducted to determine the
full-duration static firing test attitude necessary tocertify the design changes completely. The
assessment included establishing test objectives,defining and quantifying attitude-sensitiveparameters, and evaluating attitude options. Bothhorizontal and vertical (nozzle up and down) testattitudes were assessed. In all three options,consideration was given to testing with andwithout extemally applied loads. This assessmentdetermined that the conditions influencing the jointand insulation behavior could best be tested to
design extremes in the horizontal attitude. Inconjunction with the horizontal attitude for theRSRM full-scale testing, it was decided toincorporate externally applied loads. A secondhorizontal test stand for certification of the RSRM
was constructed at Morton Thiokol. This new
stand, designated as the T-97 Large Motor StaticTest Facility, is being used to simulateenvironmental stresses, loads and temperaturesexperienced during an actual Shuttle launch andascent. The new test stand also providesredundancy 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 majorreassessments of each vehicle's design,processing, inspection and operations. While theShuttle SRM insulation/propellant integrity wasnot implicated in the 51-L failure, the intent is topreclude a failure similar to that experienced byTitan. The RSRM field joint is quite tolerant ofunbonded insulation. It has sealed insulation to
prevent hot combustion products from reachingthe insulation-to-case bond line. The bondingprocesses have been improved to reduce
contaminationpotential,andthenewgeometryOfthetangcapturefeatureinherentlyprovidesmoreisolation of the edge insulation area fromcontaminatingagents.A greatlyenhancedNon-DestructiveEvaluationprogramfortheRSRMhasbeenincorporated.Theenhancednon-destructivetesting includes ultrasonic inspection andmechanicaltestingof propellantandinsulationbondedsurfaces.All segmentswill againbeX-rayedfor thefirst flightandnear-termsubsequentflights.
CONTINGENCY PLANNING. To
provide additional program confidence, both near-and long-term contingency planning wasimplemented. Alternative designs, which might beincorporated into the flight program at discretedecision points, include field-joint graphite-composite overwrap bands and alternative sealsfor the field joint and case-to-nozzle joint.Alternative designs for the nozzle include adifferent composite lay-up technique and a steelnose inlet housing.
Alternative designs with long-lead-timeimplications were also developed. These designsfocus on the field joint and cast-to-nozzle joint.
Since fabrication of the large steel componentsdictates the schedule, long-lead procurement ofmaximum-size steel ingots was initiated. Thisallowed machining of case joints to either the newbaseline or to an alternative design configuration.Ingot processing continued through forging andheat treating. At that time, the final design wasselected. A principal consideration in thisconfiguration decision was the result ofverification testing on the baseline configuration.
INDEPENDENT OVERSIGHT. Asrecommended in the "Presidential Commission
Report" and at the request of the NASAadministrator, the National Research Council
established an Independent Oversight Panelchaired by Dr. H. Guyford Stever, who reportsdirectly to the NASA Administrator. Initially, thepanel was given introductory briefings on theShuttle system requirements, implementation andcontrol, the original design and manufacturing ofthe SRM, Mission 51-L accident analyses andpreliminary plans for the redesign. The panel hasmet with major SRM manufacturers and vendors,and has visited some of their facilities. The panelfrequently reviewed the RSRM design criteria,engineering analyses and design, and certification
program planning. Panel members continuouslyreview the design and testing for safe operation,selection and specifications for material, andquality assurance and control. The panel hascontinued to review the design as it progressesthrough certification and review the manufacturingand assembly of the first flight RSRM. Panelmembers have participated in major programmilestones, project requirements review, andpreliminary design review; they also willparticipate in future review. Six written reportshave been provided by the panel to the NASAadministrator.
In addition to the NRC, the redesign team hasa design review group of 12 expert seniorengineers from NASA and the aerospace industry.They have advised on major program decisionsand serve as a "sounding board" for the program.
Additionally, NASA requested the four othermajor SRM companies -- Aerojet StrategicPropulsion Co., Atlantic Research Corp.,Hercules Inc. and United Technologies Corp.'sChemical Systems Division -- to participate in theredesign efforts by critiquing the design approachand providing experience on alternative design
approaches.
NSTS PROGRAM MANAGEMENT
The Space Shuttle program is the majorsegment of NASA's National SpaceTransportation System (NSTS) managed by theOffice of Space Flight (OSF) at NASAHeadquarters in Washington, D. C. The office is
headed by an Associate Administrator who reportsdirectly to the NASA Administrator and is chargedwith providing executive leadership, overalldirection and effective accomplishment of theSpace: Shuttle and associated programs, includingunmanned launch vehicles.
The Associate Administrator for Space Flightexercises institutional management authority overthe activities of the NASA field organizations
whose primary functions are related to the NSTSprogram. These are Johnson Space Center (JSC),Houston, Texas; Kennedy Space Center (KSC),
Fla.; Marshall Space Flight Center (MSFC),Huntsville, Ala.; and Stennis Space Center(formerly National Space TechnologyLaboratories), Bay St. Louis, Miss.
- 29 -
DIRECTOR I "__ LevelINSTS I _'_
ASSISTANTQDIR ECTOR J
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............ ...... ...... ...............\ ..........
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NSTS Organization and Relationships
The directors of these organizations, alongwith the Associate Administrator for Space Flight,also are members of the Office of Space FlightManagement Council. This group meets regularlyto review Shuttle program progress and to providean independent and objective assessment of thestatus of the overall program.
NSTS ORGANIZATION. Within the
OSF, centralized management authority for theSpace Shuttle program is charged to the Director,
NSTS. This individual is the program's generalmanager and has full responsibility and authorityfor the operation and conduct of the Shuttle
program. These responsibilities include programcontrol, budget planning and preparation,scheduling and the maintenance of a balancedprogram. The NSTS director reports to theAssociate Administrator for Space Flight.
Organizational elements of the NSTS officeare located at NASA Headquarters, JSC, KSC,MSFC and at the Vandenberg Launch Site (VLS);in Califomia.
The NSTS office has two deputies who areresponsible for the day-to-day management andoperation of the Shuttle program. They are:Deputy Director, NSTS Program, a NASAHeadquarters employee whose duty station is atJSC, and the Deputy Director, NSTS Operations,also a NASA Headquarters employee, whose dutystation is at KSC. Both individuals report directlyto the NSTS director.
Specific major responsibilities of the DeputyDirector, NSTS Program, include the following:
•Establishing policy and providing continuousdirection to all elements engaged in Shuttleprogram activities.
•Establishing and controlling the Level II(two) requirements baseline that provides thedetailed requirements that supplement andimplement the Level I (one) requirements.
•Detailed program planning, budgeting,scheduling, system configuration managementand program direction.
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•Systemengineeringandintegrationof theflight vehicle, ground systems and facilities.
•Integration of payloads with orbiter.oMission planning and integration.
There are five organizational elements underthe Deputy Director, NSTS Program, charged withaccomplishing the management responsibilities ofthe program. They are: NSTS EngineeringIntegration, NSTS Management Integration,NSTS Program Control and NSTS Integration andOperations -- all of which are located at JSC. Thefifth division is the Shuttle Projects Office, locatedat MSFC, which has overall management andcoordination of the MSFC elements -- the solid
rocket boosters, external tank and main engines --involved in the Shuttle program.
The Deputy Director, NSTS Operations, onthe other hand, is specifically charged with thefollowing major functions:
•Formulating policy, program plans andbudget requirements in support of Shuttleoperations at KSC, JSC, Edwards AFB and
Vandenberg AFB, as well as other programoperations facilities including the worldwidecontingency landing sites.
•Final vehicle preparation, mission executionand return of the orbiter for processing for itsnext flight.
•Management of the presentation andscheduling of the Flight Readiness Review(FRR).
•Chairing and management of the MissionManagement Team (MMT).
The duties of the NSTS Deputy Director,Operations, are carried out by three OperationsIntegration offices located at JSC, KSC andMSFC.
Management relationships in the centralizedNSTS organization are configured into four basicmanagement levels which are designed to reducethe potential for conflict between the programorganizations and NASA institutionalorganizations.
The NSTS Director serves as the Level I
manager and is responsible for the overall programrequirements, budgets and schedules.
The NSTS Deputy Directors are Level IImanagers and responsible for management andintegration of all elements of the program. Thisincludes integrated flight and ground systemrequirements, schedules and budgets.
NSTS project managers located at JSC, KSCand MSFC are classified as Level III (three)managers and are responsible for managingdesign, qualification and manufacturing of Shuttlecomponents, as well as a launch and landingoperations.
NSTS design authority personnel andcontractors are Level IV managers and areresponsible for the design, development,manufacturing, test and qualification of Shuttlesystems.
lAUNCH CONS'IRAINT PROCEDURES.
As part of the FRR, a launch constraints list isestablished and approved by the AssociateAdministrator. The Deputy Director, NSTSOperations, has the responsibility to tract each of
the constraints and to assure that they are properlyclosed out prior to the L-2 day MMT review. TheAssociate Administrator or his designee (Director,NSTS) has the final closeout authority.
LAUNCH DECISION PROCESS.
Major decision-making meetings leading to adecision to launch are Flight Readiness Reviewsand Mission Magament Team reviews.
The FRR is usually held 2 weeks before ascheduled launch. Its chairman is the Associate
Administrator for Space Flight. Present at thereview are all senior program and fieldorganization management officials and supportcontractor representatives.
During the review, each manager must assesshis readiness for launch based on hardware status,
problems encountered during launch processing,launch constraints and open items. Each NASAproject manager and major Shuttle componentsupport contractor representative is required tosign a Certificate of Flight Readiness.
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TheMMT,madeupof program/projectlevelmanagers,andchairedby theDeputyDirector,NSTSOperations,providesaforumfor resolvingproblemsandissuesoutsidethe guidelinesandconstraintsestablishedfor theLaunchandFlightDirectors.
TheMMT will be activated at launch minus 2
days (L-2) for a launch countdown status briefing.The objective of the L-2 day meeting is a assessany deltas to flight readiness since the FRR and togive a "go/no-go" to continue the countdown.
The MMT will remain active during the f'malcountdown and will develop recommendations onvehicle anomalies and required changes topreviously agreed to launch commit criteria. TheMMT chairman will give the Launch Director a"go" for coming out of the L-9 minute hold and isresponsible for the final "go/no-go" decision.
MAN AGt_ENT LY/dMUNICATIONS. I n
the area of management communications, weeklyintegrated program schedules are published whichprovide detailed data from each project element andthe NSTS Engineering Office. These widelydistributed schedules are designed to createmanagement awareness of the interrelated tasksand critical program paths needed to meetimportant program milestones.
Additionally, all project and programmanagement personnel meet monthly at theProgram Director Management Review to brief theprogram status and to resolve any program issuesand concerns. This review is followed by ameeting of the Management Council to also reviewstatus and to resolve any issues brought forwardby the Director, NSTS.
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CHRONOLOGY
1972
Jan. 5
March 15
Aug. 9
President Nixon proposes development of a reusable space transportation system, theSpace Shuttle.
NASA selects the three-part configuration for the Space Shuttle -- reusable orbiter,partly reusable SRB and an expendable external tank.
Rockwell receives NASA contract for construction of the Space Shuttle orbiter.
IOct. 17
Sept. 17
1975
First Space Shuttle main engine tested at the National Space Technology Laboratories,Miss.
Rollout of orbiter Enterprise (OV-101).
IJuly 18
Aug. 12
Sept. 13
1976
Thiokol conducts 2-minute firing of an SRB at Brigham City, Utah.
First free flight Approach and Landing Test (ALT) of orbiter Enterprise from Shuttlecarrier aircraft at Dryden Flight Research Center, Calif. Flight duration: 5 minutes, 21seconds. Landing occurred on Runway 17.
Second Enterprise ALT flight of 5 minutes, 28 seconds; landing on Runway 15. (Threemore ALT flights were flown by Enterprise on Sept. 23 Oct. 12 and Oct. 25.)
IJan. 18
1978
Thiokol conducts second test firing of an SRB.
IMarch 8
March 20-24
June 15
1979 [i ii
Orbiter Columbia (OV-102) transported 38 miles overland from Palmdale to DrydenFlight Research Center.
Columbia flown on Shuttle carrier aircraft to Kennedy Space Center with overnightstops at E1 Paso and San Antonio, Texas, and Eglin AFB, Fla.
First SRB qualification test firing; 122 seconds.
INov. 26
Dec. 29
1980 _ I
Columbia mated to SRBs and external tank at Vehicle Assembly Building (VAB) forSTS-1 mission.
Space Shuttle vehicle moved from VAB to Launch Complex 39A for STS-Imission. 1981
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CHRONOLOGY
IFeb. 20
April 20-21
Aug. 4
Aug. 26
Nov. 12-14
Nov. 24-25
Dec. U
1980 continued
Hight readiness firing of Columbia's main engines; 20 seconds.
Columbia returned to KSC by Shuttle carrier aircraft via Tinker AFB, Okla.
Columbia mated with SRBs and external tank for STS-2 mission.
Space Shuttle vehicle moved to Launch Complex 39A for STS-2 mission.
STS-2, first flight of an orbiter previously flown in space
Columbia transported back to KSC via Bergstrom AFB, Texas.
Spacelab 1 arrives at KSC.
1982
Feb. 3
Feb. 16
March 22-30
April 6
May 16
May 25
June 27-July 4
June 30
July 1
July 4-5
July 14-15
Sept. 9
Sept. 21
Nov. 11-16
Nov. 21-22
Nov. 23
Nov. 30
Dec. 18
Columbia moved to VAB for mating in preparation for STS-3 mission.
Assembled Space Shuttle vehicle moved from VAB to launch pad for STS-3 mission.
STS-3 mission; landing at White Sands, N.M.
Columbia returned to KSC from White Sands.
Columbia moved to VAB for mating in preparation for STS-4.
STS-4 vehicle moved to launch pad.
STS-4 mission flown; first concrete runway landing at Edwards AFB.
Orbiter Challenger (OV-099) rolled out at Palmdale.
Challenger moved overland to Dryden.
Challenger flown to KSC via Ellington AFB, Texas.
Columbia flown to KSC via Dyess AFB, Texas.
Columbia mated with SRBs and external tank in preparation for STS-5.
STS-5 vehicle moved to launch pad.
$TS-5 mission; landing at Edwards AFB.
Columbia returned to KSC via Kelly AFB, Texas
Challenger moved to VAB and mated for STS-6.
STS-6 vehicle moved to launch pad.
Flight readiness firing of Challenger's main engines; 20 seconds.
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CHRONOLOGY
1983
Jan. 22
April 4-9
May 21
May 26
June 18-24
July 26
June 28-29
Aug. 2
Aug. 30-Sept. 5
Sep. 9
Sept. 23
Sept. 28
Oct. 17
Oct. 19
Nov. 5
Nov. 6
Nov. 8
Nov. 8-9
Nov. 28-Dec. 8
Dec. 14-15
Second flight readiness firing of Challenger's main engines; 22 seconds.
STS-6 mission, first flight of Challenger.
Challenger moved to VAB for mating in preparation for STS-7 mission.
Challenger moved to launch pad for STS-7.
STS-7 mission flown with landing at Edwards AFB.
Challenger moved to VAB for mating in preparation for STS-8.
Challenger flown back to KSC via Kelly AFB.
STS-8 vehicle moved to launch pad.
STS-8 mission; first night launch and landing at Edwards AFB.
Challenger returned to KSC via Sheppard AFB, Texas.
Columbia moved to VAB for mating in preparation for STS-9.
STS-9 vehicle moved to launch pad.
STS-9 launch vehicle moved back to VAB from pad because of SRB nozzle problem.
Columbia moved to Orbiter Processing Facility.
Orbiter Discovery (OV-103) moved overland to Dryden.
Discovery transported to Vandenberg AFB, Calif.
STS-9 vehicle again moved to launch pad.
Discovery flown from Vandenberg AFB to KSC via Carswell AFB, Texas.
STS-9 mission; landing at Edwards AFB.
Columbia flown to KSC via El Paso, Kelly AFB and Eglin AFB.
1984
Jan. 6
Jan. 11
Feb. 3-11
March 14
March 19
April 6-13
I
Challenger moved to VAB for mating in preparation of STS 41 B mission.
STS 41-B vehicle moved to launch pad.
STS 41-B mission; first landing at KSC.
Challenger moved to VAB for mating in preparation for STS 41-C mission.
STS 41-C vehicle moved to launch pad.
STS 41-C mission; landing at Edwards AFB.
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CHRONOLOGY
IApril 17-18
May 12
May 19
June 2
June 25
June 26
July 14
Aug. 9
Aug. 30-Sept. 5
Sept. 8
Sept. 9-10
Sept. 13
Oct. 5-13
Oct. 18
Oct. 23
Nov. 7
Nov. 8-16
1984 continuedi
Challenger flown back to KSC via Kelly AFB.
Discovery moved to VAB for mating in preparation for STS 41-D.
STS 41-D vehicle moved to launch pad.
Flight readiness firing of Discovery's main engines.
STS 41-D launch attempt scrubbed because of computer problem.
STS 41-D launch attempt scrubbed following main engine shutdown at T minus 4seconds.
STS 4 I-D vehicle moved back to VAB for remanifest of payloads.
STS 41-D vehicle again moved out to the launch pad.
STS 41-D mission; first flight of Discovery;landing at Edwards AFB.
Challenger moved to VAB for mating in preparation for STS 4I-G mission.
Discovery returned to KSC via Altus AFB, Okla.
STS 41-G launch vehicle moved to launch pad.
STS 41-G mission; landing at KSC.
Discovery moved to VAB for mating in preparation for STS 51-A mission.
STS 51-A launch vehicle moved to launch pad.
STS 51-A launch scrubbed because of high shear winds.
STS 51-A mission; landing at KSC.
1985
Jan. 5
Jan. 24-27
Feb. I0
Feb. 15
March 4
March 23
March 28
April 6
Discovery moved to launch pad for STS 51-C mission.
STS 51-C mission landing at KSC.
Challenger moved to VAB for mating in preparation for STS 5 I-E mission.
STS 51-E vehicle moved to launch pad.
STS 51-E vehicle rolled back to VAB; mission cancelled; payloads combined with STS51-B.
Discovery moved to VAB for mating in preparation for STS 51-D mission.
STS 51-D vehicle moved to launch pad.
Atlantis (OV-104) rollout at Palmdale.
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CHRONOLOGY
IApril 10
April 12-19
April 13
April 15
April 29-May 6
May l0
May 28
June 4
June 17-24
June 24
June 28
June 29
July 11
July 12
July 14
July 29-Aug. 6
July 30
Aug. 6
Aug. 10-11
Aug. 24
Aug. 25
Aug. 27-Sept. 3
August 29
Sept. 7-8
Sept. 12
Oct. 3-7
Oct. U
Oct. 12
1985 continued
Challenger moved to VAB for mating in preparation for STS 51-B mission.
STS 51-D mission; landing at KSC.
Atlantis ferried to KSC via Ellington AFB, Texas.
Challenger moved to launch pad for 51-B missing.
STS 51-B mission; landing at Edwards AFB.
Challenger transported back to KSC via Kelly AFB.
Discovery moved to VAB for mating in preparation for STS 51-G.
STS 51-G vehicle moved to the launch pad.
STS 51-G mission; landing Edwards AFB.
Challenger moved to VAB for mating in preparation for STS 51-F.
Discovery ferried back to KSC via Bergstrom AFB, Texas.
STS 51-F vehicle moved to the launch pad.
Refurbished Columbia moved overland from Palmdale to Dryden.
STS 51-F launch scrubbed at T-minus 3 seconds because of main engine shutdown.
Columbia returned to KSC via Offutt AFB, Neb.
STS 51-F mission landing at Edwards AFB.
Discovery moved to VAB for mating in preparation for STS 51-1 mission.
STS 51-I vehicle moved to the launch pad.
Challenger flown to KSC via Davis-Monthan AFB, Ariz.; Kelly AFB; and Eglin AFB.
STS 51-I mission scrubbed at T minus 5 minutes because of bad weather.
STS 51-1 mission scrubbed at T-minus 9 minutes because of an onboard computerproblem.
STS 51-1 mission; landing at Edwards AFB.
Atlantis moved to launch pad for the 51-J mission.
Discovery flown back to KSC via Kelly AFB.
Hight readiness firing of Atlantis' main engines; 20 seconds.
STS 51-J mission; landing at Edwards AFB.
Atlantis retumed to KSC via Kelly AFB.
Challenger moved to VAB for mating in preparation for the STS 61-A mission.
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CHRONOLOGY
1985 continuedI l
Oct. 16
Oct. 30-Nov. 6Nov. 8
Nov. 10-11
Nov. 12
Nov. 18
Nov. 22
Nov. 26-Dec. 3
Dec. l
Dec. 7
Dec. 16
Dec. 19
Dec. 22
Challenger vehicle moved to the launch pad for STS 61-A mission.
STS 61-A mission; landing at Edwards AFB.Atlantis moved to VAB for mating in preparation for the STS 61-B.
Challenger flown back to KSC via Davis-Monthan AFB, Kelly AFB and Eglin AFB.
STS 61-B vehicle moved to the launch pad.
Enterprise (OV-101) flown from KSC to Dulles Airport, Washington, D.C., andturned over to the Smithsonian Institution.
Columbia moved to the VAB for mating in preparation STS 61-C.
STS 61-B mission landing at Edwards AFB.
STS 61-C vehicle moved to launch pad.
Atlantis retumed to KSC via Kelly AFB.
Challenger moved to VAB for mating in preparation for the STS 51-L mission.
STS 61-C mission scrubbed at T minus 13 seconds because of SRB auxiliary power
unit problem.
STS 51-L vehicle moved to Launch Pad 39B.
1986
Jan. 6
Jan. 7
Jan. 10
Jan. 12-18
Jan. 22-23
Jan. 27-28
Feb. 3
March 24
May 12
July 8
STS 61-C mission scrubbed at T minus 31 seconds because of liquid oxygen valveproblem on pad.
STS 61-C mission scrubbed at T minus 9 minutes because of weather problems atcontingency landing sites.
STS 61-C mission scrubbed T minus 9 minutes because of bad weather at KSC.
STS 61-C mission; landing at Edwards AFB.
Columbia returned to KSC via Davis-Monthan AFB, Kelly AFB and Eglin AFB.
STS 51-L launched from Pad B. Vehicle exploded 1 minute, 13 seconds after liftoffresulting loss of seven crew members.
President Reagan announced the formation of the Presidential Commission on theSpace Shuttle Challenger Accident, headed by William P. Rogers, former Secretary ofState.
NASA publishes "Strategy for Safely Returning the Space Shuttle to Flight Status."
President Reagan appoints Dr. James C. Fletcher NASA Administrator.
NASA establishes Safety, Reliability Maintainability, and Quality Assurance Office.
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CHRONOLOGY
1986 continued
July 14
Aug. 15
Aug. 22
Sept. 5
Sept. 10
Oct. 2
Oct. 30
Nov. 6
NASA's plan to implement the recommendations of the Rogers commission wassubmitted to President Reagan.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 commercialsatellites, except for those which are Shuttle-unique or have national security or foreignpolicy implications.
NASA announced the beginning of a series of tests designed to verify the ignitionpressure dynamics of the Space Shuttle solid rocket motor field joint.
Study contracts were awarded to five aerospace firms for conceptual designs of analternative or Block II Space Shuttle solid rocket motor.
Astronaut Bryan O'Connor was named chairman of Space Flight Safety Panel. Thispanel, with oversight responsibility for all NASA manned space program activities,reports to the Associate Administrator for Safety, Reliability, Maintainability andQuality Assurance.
After an intensive study, NASA announced the decision to test fire the redesigned solidrocket motor in a horizontal attitude to best simulate the critical conditions on the field
joint which failed during the 51-L mission.
Discovery moved to OPF where more than 200 modifications are accomplished forSTS-26 mission.
Office of the Director, National Space Transportation System, established in the NASAHeadquarters Office of Space Flight.
July 31
Aug. 3
1987
Rockwell Intemational awarded contract to build a fifth orbiter to replace theChallenger.
Discovery in the Orbital Processing Facility is powered up for STS-26 mission.
IMid-Jan.
March 28
May 28
June 10
June 14
June 21
July 4
Aug. 10
1988
Main engines are installed in Discovery.
Stacking of Discovery's SRBs gets underway.
Stacking of Discovery's SRBs completed.
SRBs and External Tank are mated.
The fourth full-duration test firing of the redesigned SRB motor is carried out.
Discovery rolls over from OPF tp the VAB.
Discovery moved to Launch Pad 39B for STS-26 mission.
Flight Readiness Firing of Discovery's main engines is conducted successfully.
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