JOURNAL OF SPACECRAFT AND ROCKETSVol. 33, No. 2, March-April 1996

Investigation of Space Launch Vehicle Catastrophic Failures

I-ShihChang*The Aerospace Corporation, El Segundo, California 90245

The failures of the world's major space launch vehicles from Jan. 1,1984, through Dec. 31,1994, are reviewed.The cause of failure and corrective action for each failed vehicle are documented, the vulnerable areas of thevehicle are discussed, and the reliability of the major space launch vehicles from different countries in the worldis compared. The review provided some of the lessons learned during the last 11 years and revealed that thepropulsion subsystem is still the Achilles' heel of the space program. The intent is to pave the way for identifyingthe critical skills and processes needed to mitigate future space-related mission failures.

Introduction

N EXT to wars, nuclear reactor accidents, major transportationaccidents, and natural disasters, a space launch failure is one

of the most expensive losses in the national resources for a nation inpursuit of technological advancement. A space mission involving alarge launch vehicle and a sophisticated satellite can easily cost hun-dreds of millions of dollars. This cost does not include the expense,time, and effort spent during the recovery period and the damage tonational prestige. Therefore, it is very important that the recurrenceof space-launch failures be prevented.

In this paper, the space launch vehicle (SLY) failures over thelast 11 years are reviewed. This review was one of the tasks carriedout during the recovery program in the wake of the Titan IV K-l 1launch-vehicle failure. TTie objectives of this study can be stated asfollows: 1) provide information on the past space-related missionfailures from Jan. 1,1984, through Dec. 31,1994; 2) document thecause of failure and corrective action for each failed SLY; 3) focusattention on the vulnerable areas of the vehicle; and 4) identify thecritical skills and processes needed to mitigate future launch failures.The world space-launch data are collected first. A simple analysisthen is performed to compare the reliability of the major SLVs fromdifferent countries. The cause of failure and corrective action forevery U.S. vehicle are presented. Because of the lack of reportedinformation, the cause of failure and corrective action for the foreignspace launch vehicles are presented only selectively. Finally, lessonslearned from the past worldwide space-launch catastrophic failuresare summarized.

Classification of Subsystem FailuresThe failure of a launch vehicle is attributed to the problem asso-

ciated with one of the following subsystems in this study.Propulsion consists of the main propulsion components of rocket

motor, liquid engine, orbital maneuvering thruster, and attitude con-trol thruster; combustion chamber; nozzle; solid propellant; liquidpropellant; thrust vector actuator and gimbal mechanism; feed lines;control valves; turbopumps; ignitor; motor internal insulation andbondline.

Structures include solid rocket motor (SRM) core support struc-tures; motor case; ignitor housing; helium, oxidizer, and fuel stor-age tank structure; interstage structure; nose cone; payload fairing;support skirt.

Avionics contains onboard software; circuit board; gyro; flightcomputer; attitude sensors; load relief sensors; range safety provi-sions; navigation and guidance control equipment; inertial measure-ment unit; flight instrumentation and telemetry equipment.

Separation/staging comprises staging rockets; separation mech-anism; electrical connection or wiring for separation control.

Received Sept. 30,1994; revision received May 2,1995; accepted for pub-lication May 4,1995. Copyright © 1995 by I-Shih Chang. Published by theAmerican Institute of Aeronautics and Astronautics, Inc., with permission.

* Distinguished Engineer, Vehicle Systems Division. Senior MemberAIAA.

Electrical involves wire harnesses; electrical connectors; electri-cal power supply; electrical relay boxes; battery; solenoids.

Other relates to environment; communication-, and so forth.

U.S. and Foreign SLV FailuresThe SLV failures in the U.S. and in foreign countries during the

last 11 years (Jan. 1, 1984, through Dec. 31, 1994) are given inTables 1 and 2, respectively. These data are obtained from Refs. 1-4. The vehicle name, failure date, and payloads for each country arelisted. The failed subsystems and the causes of failure are identified,based on the information reported in the references. There were 5SLV failures in the U.S. Department of Defense (DOD) space pro-grams and 9 SLV failures in the U.S. non-DOD space programs. Incomparison, there were 29 SLV failures in foreign countries duringthe last 11 years.

Analysis of SLV FailuresThe success-failure records of the U.S. and foreign SLVs during

the last 11 years are given in Tables 3 and 4, respectively. Figure 1compares the numbers of launch-vehicle subsystem failures, andFig. 2 reveals the subsystem failure rate of the SLVs in the U.S. andin foreign countries during the last 11 years. Figure 3 shows the total

I

propulsion struct, avionics separa. electri. otherSubsystem

Fig. 1 SLV subsystem failures.

£4.

propulsion struct, avionics separa. electri.Subsystem

Fig. 2 SLV subsystem failure rate.

198

CHANG 199

Table 1 U.S. SLV failures

Vehicle

17.5. DODTitan IVTitan 34DAtlas-CentaurTitan 34DTitan 34DU.S. non-DODPegasus XLAtlas-CentaurAtlas-CentaurPegasusAtlas-CentaurTitan IIIDeltaSTS-ChallengerAtlas-Centaur

Flight

K-llD-3

AC-67D-9D-7

STEP-1AC-74AC-71

F-2AC-70CT-2178

51-LAC-62

Fail date

08/02/9309/02/8803/26/8704/18/8608/28/85

06/27/9403/25/9308/22/9207/17/9104/18/9103/14/9005/03/8601/28/8606/09/84

Satellite

Not announcedNot announcedNot announcedNot announcedNot announced

STEP-la

Navy UHF Follow-On21

Galaxy 1-R ComsatDARPA Microsatsa

BS-3H ComsatIntelsatVI-3NOAA-GOES 7TDRSS 2Intelsat V-9

Failed subsys.

PropulsionPropulsionOtherPropulsionPropulsion

AvionicsPropulsionPropulsionSeparationPropulsionSeparationElectricalPropulsionPropulsion

Cause of failure

Motor case burnthrough (inadequate restrictor repair)Transtage fuel-tank leakageVehicle launched in thunderstorm and struck by lightningMotor case burnthrough (insulation debond)Stage I engine propellant leakage and premature shutdown

Autopilot software used erroneous aerodynamic load coefficientsPower loss and premature shutdown of the first-stage engineCentaur C-l engine failed to achieve full thrustStage and payload separation anomaliesCentaur C-l engine failed to achieve full thrustSecond stage failed to separate because of incorrect interface wiringStage I relay box electrical shortHot gas leaked through O-ring in the SRM jointFuel line leaking in Centaur reaction control system

"Satellites were launched under commercial launch service contracts.

Table 2 Foreign SLV failures

Country Vehicle Fail date Satellite Failed subsys. Cause of failure

China

France

India

CIS

USSR

CZ-2ECZ-3CZ-3

Ariane 42PAriane 44LPAriane 44L

Ariane 2Ariane 3

PSLVASLVASLV

SL-14 TsyklonSL-12 ProtonSL-4 SoyuzSL-16 ZenitSL-16ZenitSL-16 Zenit

SL-8 KosmosSL-16 Zenit

SL-12 ProtonSL-6 MolniyaSL-12 ProtonSL-12 ProtonSL-12 ProtonSL-12 ProtonSL-12 Proton

SL-14 TsyklonSL-6 MolniyaSL-14 Tsyklon

12/21/9212/28/9101/29/8412/01/9401/24/9402/22/9005/31/8609/12/8509/20/9307/13/8803/24/8705/25/9405/27/9304/27/9312/25/9202/05/9208/30/9106/25/9110/04/9008/09/9006/21/9002/17/8801/18/8804/24/8701/30/8712/29/8610/15/8610/03/8611/27/84

Optus B2Comsat STTW 5

Exp. GEO ComsatPAS-3

Eutelsat 2F5, Turksat 1Japanese Superbird B

Intelsat VA-5Spacenet F-3, ECS-3

IRS-1ESRS2(Rohini)

SROSSKosmos 2281

Gorizont ComsatKosmos 2243Kosmos 2227

KosmosKosmosKosmosKosmosKosmos

Kosmos 2084Kosmos 1917,1918,1919

KosmosKosmos 1838, 1839, 1840Kosmos 1817, Ekran-16A

KosmosKosmos

Kosmos 1783Kosmos 1612

StructuresPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionAvionicsOtherElectricalSeparationPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionPropulsionAvionicsAvionicsAvionicsPropulsionPropulsionPropulsion

Structural flaw in the rocket's fairing caused explosionPremature third stage shutdownThird stage failed to restartBlockage or leakage of LOX on the third-stage engineThird-stage turbopump shutdown because of overheatingCloth clogging first-stage piping, vehicle explodedImproper ignition of third-stage engineFailure of third-stage hydrogen valveRocket veered off course during second-stage separationHigh winds and premature cutoff of strap-ons at 150 sStage 1 did not ignite because of short circuitSecond and third stages failed to separateState 2 and 3 engine failures because of fuel contaminationExplosion caused by residual propellants in the final stageSecond stage blew up shortly after orbital insertionSecond-stage malfunctionedSecond-stage failureSecond-stage failureVehicle exploded over launch pad because of engine failureStage 2 engine failure because of fuel line clogged by a piece of ragFourth-stage failureStage 4 engine failure because of fuel contaminationStage 3 engine failure because of fuel line destructionStage 4 control-system failure because of instrument defectStage 4 control-system failure because of relay component defectStage 3 control-system failure because of relay contact separationLaunch vehicle failureLaunch vehicle failureLaunch vehicle failed to shut down at perigee

Table 3 U.S. SLV success-failure record

U.S. DODa U.S. non-DOD U.S.Year

19841985198619871988198919901991199219931994TotalSuccess rate, %

Success

12645616141012128

10595.5

Failure

011110000105

Success

911235212715111895

91.3

Failure102000121119

Success2117681118261727232620093.5

Failure1131101212114

Includes all DOD-involved government space launches.

numbers of successes and failures, and Fig. 4 shows the success rateof the space launches for the U.S. and for foreign countries. Thelaunch success rate from Jan. 1, 1984, through Dec. 31, 1994, is95.5% for U.S. DOD and 91.5% for U.S. non-DOD. The DODmissions are subjected to detailed technical oversight and launch-readiness certifications and have a better launch success rate thanthat of commercial programs, even though the same launch systemis used for the two categories of launches. In comparison, the launchsuccess rate during the last 11 years is 100% for Israel and Japan,97.9% for the Commonwealth of Independent States (CIS)AJSSR,92.1% for France, 89.3% for China, and 50% for India. Note thatthere were only two space launches for Israel and six for India. Thisimplies that the rates shown here can vary significantly for thesetwo countries in the future.

It is important to point out that a decrease in the launch success rateusually translates to many millions of dollars lost to space-relatedmission failures.

200 CHANG

Table 4 Foreign SLV success-failure record

ChinaYear

19841985198619871988198919901991199219931994TotalSuccess rate, %

Success2122405031525

89.3

Failure100000011003

FranceSuccess

4322775877658

92.1

Failure011000100025

IndiaSuccess

000000001023

50.0

Failure000110000103

IsraelSuccess

000010100002

100.0

Failure000000000000

JapanSuccess

3223223211223

100.0

Failure000000000000

CIS/USSRSuccess

969890948974745953464882197.9

Failure

10322032221

18

« 800

o 700

2 600

400

300

200

100

0

200

I SUCC.II fail I

2 . 0 UUSA China France India Israel Japan CIS/USSR

Country

Fig. 3 Space launch log.

100 100 97.9

USA China France India Israel Japan CIS/USSRCountry

Fig. 4 Space launch success rate.

Propellant Cut,Pressurized and Burning

Fig. 5 Restrictor repair and propellant cut.

Fig. 6 Transtage fuel tank andoxidizer tank.

General observations that can be made from the analysis of thesedata are as follows: 1) the predominant root cause of the SLV failuresis in the propulsion subsystem (U.S. 64.3%; foreign 72.4%); 2) fewfailures are related to the basic design; 3) some failures are relatedto human errors, improper handling, poor procedures, management,workmanship, or judgment; and 4) fuel leakage accounts for manyof the liquid-engine failures in the propulsion subsystem.

Cause of Failure and Corrective Action for the U.S. SLVsThe cause of failure and corrective action for the SLVs in the U.S.

during the last 11 years are discussed briefly in this section.

U.S.DOD: Titan IV K-llLaunch site: Western Range.Fail time: 101.2s.

Cause of FailureThe propellant of the Titan IV SRM was cut approximately

0.25 in. deep and extended 34 in. in the radial direction from thebore during the restrictor repair shown in Fig. 5. The repair wasmore extensive than had ever been attempted on such a motor seg-ment. The specific cut made in the propellant of the motor segmentwas expected to be closed by internal pressure generated at ignition

from the results of mathematical computer modeling. But at mo-tor ignition, the face of the cut was pressurized and open, allowingflame to propagate along the cut. This propagation resulted in earlyexposure of the SRM side-wall insulation and eventual motor-caseburnthrough at 101.2 s. The designed (nominal) burn time of theSRMisl27s(Ref.5).

Corrective Action1) Define the hardware acceptance criterion and qualification pro-

gram for the SRM segments with repaired restrictor.2) Improve manufacturing process for SRM segments to avoid

restrictor repair.

U.S. DOD: Titan 34D-3Launch site: Eastern Range.Fail time: 289 s.

Cause of FailureRepair processing during prelaunch activities or shrapnel impact

during the payload fairing release event at 289 s resulted in damageto the upper portion of the transtage fuel tank and pressurization linesshown in Fig. 6. A substantial fuel leak of approximately 1340 Iboccurred during park orbit, and a large helium-tank gas leak occurred

CHANG 201

during the transtage first burn. Not enough helium was left in thesystem to allow start of the second burn.6

Corrective ActionThe fuel-tank manhole cover, seals, lines, and fittings of assem-

bled systems were removed and replaced using revalidated processesand procedures. All relevant inventories were retested and reac-cepted. Manufacturing and inspection procedures for new parts wereimproved.

U.S. DOD: Atlas-Centaur (AC-67)Launch site: Eastern Range.Fail time: 70.7 s.

Cause of FailureThe Atlas G vehicle was struck by lighting (see Fig. 7), which

resulted in an erroneous full-scale positive yaw command inducedby an electrical transient at 48.4 s of flight. The erroneous yawmaneuver resulted in the loss of vehicle control. The destruct signalwas sent at 70.7 s into flight.7

Corrective Action1) Improve training of managers on critical launch support

functions.2) Specify, duties and responsibilities of launch weather team.3) Revise launch-vehicle weather criteria and update to current

knowledge.4) Improve voice communication circuits.

U.S. DOD: Titan 34D-9Launch site: Western Range.Fail time: 8.7s.

Cause of FailureThe SRM case insulation unbonded, and the butt joint opened

under the ignition pressure. The combustion flame directly heatedthe steel case wall, shown in Fig. 8, and burnthrough of the motorcase occurred at 8.7 s after ignition. The nominal burn time for theTitan 34D SRM is 105 s (Ref. 8).

Corrective Action1) Tighten hardware quality control (material, dimension, toler-

ance).2) Improve manufacturing environment and process.3) Ensure insulation-to-case bondline integrity (visual, x-ray, ul-

trasonic, proof test).4) Revise SRM segment assembly procedure.

Fig. 7 Lightning strike caused AC-67failure.

debond

Fig. 9 Stage 1 oxidizer suctionsystem.

OxidizerFeed lineAssembly

_S/A-1 Flange s/A-2

• OX PumpiS/A-2

BoltedJoint

MarmanClamp

MarmanClamp

BoltedJoint

Fig. 10 Location of theMarman clamps.

Fig. 11 Pegasus XL aerody-namic control.

Pitch

Fig. 8 Motor case insulation debond.

U.S. DOD: Titan 34D-7Launch site: Western Range.Fail time: 213 s.

Cause of FailureThe stage 1 engine suffered three separate major anomalies. First,

during subassembly-2 (S/A-2) start transient (110 s) a large oxidizerleak of 165 Ib/s occurred in the engine oxidizer suction line, shownin Fig. 9. Second, at 213 s an internal fuel leak of 30 Ib/s occurredin subassembly 1 (S/A-1) downstream of the combustion chamberand created a vehicle side force. Third, the S/A-1 shut down at 213 sdue to failure of its turbopump assembly.9

Corrective ActionThe two Marman clamps on the oxidizer suction lines, shown in

Fig. 10, were redesigned. The design integrity and manufacturingprocesses were revisited, and improvements were incorporated toincrease their reliability.

U.S. Non-DOD: Pegasus XL STEP-1Launch site: Western Range.Fail time: 39 s.

Cause of FailureThe first-stage ignition for the Pegasus XL occurred at 5 s af-

ter drop from the L-1011 aircraft. At 32 s divergence in roll andyaw, shown in Fig. 11, began and increased exponentially to lossof control at 39 s. The vehicle continued planned course throughfirst-stage burnout. A command destruct signal was sent at 171.5 safter the second stage failed to ignite. Inaccurate estimation of the

202 CHANG

TurbopumpHigh PressurePropellant Ducts

Fig. 12 MA-5 booster.

Fig. 13 Oxidizer regulator.

DIAPH--RAGM

BLEED PLUG

vehicle aerodynamics and of their uncertainty levels has been iden-tified as the principal cause of the roll-yaw divergence. These aero-dynamic predictions were used to design the flight control system,which became unstable and allowed the vehicle to lose control in acoupled roll-yaw spiral divergence.1()

Corrective Action1) Develop an aerodynamic model through use of a combination

of wind-tunnel data and analytic techniques.2) Implement procedures to ensure the inspection and acceptance

requirements cover all aspects of aerodynamic shapes and massproperties.

3) Develop a comprehensive control uncertainty model for thenew aerodynamic model.

4) Incorporate yaw load relief into the lateral autopilot.

U.S. Non-DOD: Atlas-Centaur (AC-74)Launch site: Eastern Range.Fail time: 103 s.

Cause of FailureThe thrust from the MA-5 engine shown in Fig. 12 for the Atlas 1

first-stage booster dropped off to 65% of the nominal level 103 sinto the flight, because of problems with a regulator (see Fig. 13)used to control the throttle setting. Inefficient burning of fuel at thelower throttle setting used up propellant from the first-stage engine.This resulted in depletion shutdown of the Centaur upper stage 22 searly. The payload was placed in an incorrect orbit. The designednominal burn time for the MA-5 booster engine is 174 s (Ref. 11).

Corrective Action1) Reduce precision regulator control band by lengthening stem

in regulator.2) Use special tool to avoid loose adjustment screw.

U.S. Non-DOD: Atlas-Centaur (AC-70, AC-71)AC-71

Launch site: Eastern Range.Fail time: 178s.

AC-70Launch site: Eastern Range.Fail time: 360 s.

Cause of FailureThe Centaur C-l engine shown in Fig. 14 failed because of in-

gestion of air into the turbopump during the Atlas boost phase. Theair, entering through a stuck-open check valve, liquefied and frozein the C-l engine liquid hydrogen (LH2) pump and gearbox, which

c-2

Atlas LOX Tank

Fig. 14 Centaur engines.

T-4 K CentaurDisconnect Solenoid

Isolation

Fig. 15 Location of isolation and check valves.

Wlng ^ / FajrlngStage3

Stage 2Interstage

Stage 1

Fin Rockets^

Fig. 16 Pegasus launch vehicle: exploded view.

prevented the engine from achieving full thrust. The low perfor-mance of C-l engine produced an unbalanced thrust and resulted inthe loss of vehicle control and of the mission.12

Corrective Action1) Add a solenoid isolation valve (shown in Fig. 15) in the helium

chilldown system to prevent ingestion of air or nitrogen into theturbopumps.

2) Add a liquid-nitrogen-gaseous-helium (LN2-GHe) heat ex-changer for ground prechill, and increase the duration of prestartchilldown to maintain the turbopump temperature above the nitro-gen freezing point (-320°F).

3) Add a 10-/xm absolute filter downstream of the heat exchangerin the ground chilldown system to minimize contamination in theengine.

4) Delete the helium warmup purge and tank blowdown purgeto eliminate freezing of moisture from the tanks and ducts duringprelaunch operations.

U.S. Non-DOD: Pegasus F-2Launch site: Eastern Range.Fail time: 102.8 s.

Cause of FailureThe vehicle angled off in a wrong direction and placed the payload

in a significantly lower orbit than planned. The major anomaly was

CHANG 203

Aluminum SheathRDX Charge

NASA StandardDetonator Block Initiator Foam B|Mt

Attenuator^-——-

Linear Shaped Charge

\ Interstage Skin

Fig. 17 LSC-detonator block design.

LSCPiggyback

"LSC RetainerFig. 18 LSC retainer and piggyback.

Fig. 19 Titan III CT-2 payload separation system: BW, bridge wire;SFC, squib firing circuit; and I/F, interface.

the incomplete first-second-stage separation (Fig. 16) as the second-stage burn started in a nose-down attitude at 102.8 s. Subsequentanomalies were incomplete payload fairing separation at 214 s andejection of the fin rocket nozzle assembly because of overpressurecaused by unsuccessful expulsion of a weather seal.13

Corrective Action1) Increase linear shaped charge (LSC), add spacer protect LSC

under detonation block (Fig. 17), and add piggyback LSC (Fig. 18)to ensure simultaneous detonation for graphite motor case cuttingduring first-second-stage separation.

2) Increase strength of fairing hinges through heat treatment,restrain umbilicals, and replace shear pin with spring to precludefairing snag.

3) Redesign weather seal and requalify fin rockets.

U.S. Non-DOD: Titan III CT-2Launch site: Eastern Range.Fail time: 1499.5 s.

Cause of FailureThe payload separation system, shown in Fig. 19, was designed

for two satellites and had two discrete outputs from the missileguidance computer (MGC). For this mission it carried only a singlesatellite. The wiring team mis wired the harness, which connectedthe MGC payload separation discretes to the payload separation

Fuel Start TankOxidizer

Start TankRelayBoxGas

Generator

HeatExchanger

ThrustFrame

Turbo-pump

Main FuelValve

MainOxidizerValve

ThrustChamber

Fig. 20 RS-27 engine.

O-RingSeals

Zinc ChromateThermal Barrier

Leak Check PortJ |nsulation

Fig. 21 Original field joint.

Capture Feature Armand O-Ring (New)

r-T-SealThermal Barrier (New)

Heater (New) -* \- Q-Ring Seals

Fig. 22 Redesigned field joint.

device. The satellite never received the separation signal, and thebooster and payload failed to separate.14

Corrective Action1) Validate and verify that the wiring and discrete lists are com-

patible.2) Carry discrete identification in circuit nomenclature in check-

out drawing.3) Ensure that flight combined system test is functionally equiv-

alent to the mission-unique flight sequence.

U.S. Non-DOD: Delta 178Launch site: Eastern Range.Fail Time: 71 s.

Cause of FailureThe first-stage Rocketdyne RS-27 engine, shown in Fig. 20, was

shut down at 71 s because of an electrical short in the relay box.The vehicle was destroyed at 91 s by the range safety. The designed(nominal) burn time is 223 s (Ref. 15).

204 CHANG

j~tH _Interstage«; ''"Adapter

- Intelsat V

Nose Fairing

SplitBarrel

Centaur\—i \Stub>-C-X Adac

^PayloadAdapter

Adapter

Fig. 23 Interstage adapter for Atlas G vehicle.

s N NotchxBlast Shield

Fig. 24 LOX tank blast shield.

Corrective Action1) Add relays and lock-in circuits to the RS-27 engine relay box

for built-in redundancy.2) Add a second 28-V power source from the flight battery to the

engine relay box to overcome the effects of power surges.

U.S. Non-DOD: STS ChallengerLaunch site: Eastern Range.Fail time: 73.0 s.

Cause of FailureThe launch management waived the temperature-dependent

launch commit criteria and launched the STS Challenger at 38°Finstead of 53°F derived from the previous temperature experienceof the STS program. The rubber O-rings in the motor case joint,shown in Fig. 21, lost their resiliency in the cold. The combustionflame leaked through the O-rings and case joint and impinged onthe motor aft attach struts and external tank. Failure of the aft strutscaused the aft end of the motor to move outward and forced thenose of the SRM into the upper portion of the external tank. Theexplosion ensued at 73 s after ignition. The designed nominal burntimeis'120s(Ref. 16).

Corrective Action1) Incorporate a capture feature arm at the joint to limit joint gap

opening as shown in Fig. 22.2) Add a third O-ring and J-seal thermal barrier for improved

thermal protection.3) Widen O-ring groove to improve O-ring pressure actuation

capability.4) Add joint heaters to maintain O-ring resiliency.5) Add an external joint weather seal to prevent moisture and ice

accumulation.

U.S. Non-DOD: Atlas-Centaur (AC-62)Launch site: Eastern Range.Fail time: 288.8 s.

pressure and the collapse of the intermediate bulkhead during thecoast phase prior to the second burn.17

Corrective Action1) Increase clearance between interstage adapter and LO2 tank

blast shield, as shown in Fig. 24.2) Reduce pressure in Centaur LO2 and LH2 tanks at the time of

Atlas-Centaur separation and during Centaur-powered flight.3) Add three tubing clamps in the hydrazine system to improve

tubing support.

Cause of Failure and Corrective Actionfor the Foreign SLVs

The cause of failure and corrective action for some of the foreignSLVs during the last 11 years are discussed briefly in this section.

China: Long March CZ-2ELaunch site: Xichang, China.Fail time: 48 s.

Cause of Failure for the CZ-2E on December 21, 1992The Australian communications satellite Optus B2 exploded at

48 s into the flight from Xichang, China. Evidence points to a struc-tural flaw in the rocket's fairing, which probably imploded duringlaunch.18

France: Ariane 42PLaunch site: Kourou, French Guiana.Fail time: 900 s.

Cause of Failure for the Ariane 42P on December 1, 7994A blockage or a leak in the liquid oxygen (LOX) line to the gas

generator resulted in oxygen shortage and 30% reduction in the gasgenerator pressure, which cut the third-stage engine speed from anominal 62,000 rpm to about 50,000 rpm.18

Corrective Action1) Improve manufacturing and processing to reduce the risk of

contamination.2) Install a 400-/>tm filter at the LOX injection unit input to strain

out any particles.3) Conduct additional engine tests to certify new filter and engine

for flight.

France: Ariane 44LPLaunch site: Kourou, French Guiana.Fail time: 430 s.

Cause of Failure for the Ariane 44LP on January 24, 1994Delayed chilldown of the LOX turbopump bearing, because of

filter obstruction or on-line thermal blockage, caused overheatingof the bearing at 60 s after normal third-stage ignition. This resultedin the LOX circuit opening, interruption of the LOX supply to thechamber and gas generator, and shutdown of the third-stage engineat 80 s after third-stage ignition.19

Corrective Action1) Use ultradecontaminated filters at ground-onboard interfaces.2) Adopt a continuous flushing of LH2 and LOX purge interfaces

to prevent entry of air or moisture.3) Include a nickel-silver, self-lubricating coating to reduce fric-

tion of the turbopump bearing.

France: Ariane 3Launch site: Kourou, French Guiana.Fail time: 277 s.

Cause of FailureA significant leak occurred in the LO2 tank at the time of Atlas-

Centaur separation for the Atlas G vehicle shown in Fig. 23. Theleak resulted in the LO2 tank pressure falling below the LH2 tank

Cause of Failure for the Ariane 3 on September 12, 1985A solid-propellant, cartridge-type igniter in the combustion cham-

ber malfunctioned, or a problem in the propellant feed at ignition ofthe third stage occurred. The ignition was 0.4 s later than planned,

CHANG 205

after which the normal operating condition of the engine was notreached, and the engine stopped at 277 s after launch. The thirdstage uses LH2 and LOX propellants. Its normal burn time is 720 s(Ref. 15).

India: PSLVLaunch site: Sriharikota, India.Fail time: 246 s.

Cause of Failure for the PSLV on September 20, 1993The rocket veered off course as the second stage separated from

the vehicle. The third and fourth stages ignited but fell far shortof the planned altitude. The internal guidance system and shiftingliquid fiiel in the second stage of the rocket were the possible causesof the failure.18

CIS/USSR: ProtonLaunch site: Tyuratam, Kazakhstan.Fail time: 572 s.

Cause of Failure for the SL-12 Proton on May 27,1993A highly contaminated fuel mixture caused a burnthrough of the

rocket's cooling tubes in the combustion chamber of the second-stage engine at 205 s into the flight from the Baikonur Cosmodrome.This burnthrough caused a rapid consumption of fuel by the second-stage engine. The third stage ignited earlier and ran out of fuel tooearly. The rocket lost control at 572 s into the flight.18

Corrective ActionCopper, iron, zinc, and aluminum were found in the rocket's liquid

fuel (a mixture of nitrogen tetroxide and unsymmetrical dimethyl-hydrazine) in quantities exceeding several times the maximum ad-missible content. All rocket propellant at Baikonur will be analyzed,and fuel pipelines will be drained and flooded with new propellantand analyzed. A series of propellant quality checks will be insti-tuted at Baikonur, and better records will be kept of how the fuel isdelivered and tested.

SummaryThe review revealed that the propulsion subsystem is still the

Achilles' heel of the space program. To mitigate launch failureswith the present fleet of the SLVs, special attention needs to be paidto the following items in the propulsion subsystem: 1) solid-motormanufacturing and process control; 2) liquid-engine manufactur-ing and process control; 3) system-level liquid-engine testing andevaluation; and 4) robust design of liquid-engine and solid-motorcomponents.

It is a prudent practice to adhere to the following rules, and notto disturb Zeus' weather game: 1) do not launch a vehicle on a coldday (STS 51-L, United States); 2) do not launch a vehicle on a rainyday (AC-67, United States); and 3) do not launch a vehicle on awindy day (ASLV-6, India).

Space-related mission failures result in the waste of vast amountsof world resources. In hindsight, some of the failures could have

been prevented. This paper has briefly discussed the causes ofthe SLV failures and provided some of the lessons learned dur-ing the last 11 years. It is hoped that the information contained inthis paper will be of help in mitigating some SLV failures and inreducing the waste of useful world resources in the future.

AcknowledgmentsThis work was supported by the U.S. Air Force Space and

Missile Systems Center under Contract No. F04701-93-C-0094.The author wishes to thank M. Adams, J. D. Gilchrist, R. F.Johnson, and W. W. Wang for providing useful information to thisinvestigation.

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K.J. WeilmuensterAssociate Editor