1

ABSTRACTLarge Solid Propulsion has been used sincethe beginning of space launch activities and iscurrently used on numerous space launchers inthe world. Three main missions are identified :stage one of heavy launchers, strap onboosters for liquid core stage and propulsionsystem for small launchers.

This paper reviews the current situation oflarge solid propulsion for space launchactivities, giving an idea of what is thisbusiness and emphasizes the main identifiedinterests of this type of propulsion. Anoverview of the currently operational SRMs isthen presented giving some information on theproducts themselves. Some main trends ofSolid Propulsion are finally underlined to givea clear understanding of the technicalpotentials of this propulsion mode.

1 INTRODUCTION

Solid rocket motors are widely used in spacelaunchers. An overview of current 'westernlaunchers' shows that :

- they are used as stage one of heavylaunchers like the SHUTTLE, TITAN,H2 and ARIANE 5,- they assist the core stage of the newmedium launchers such as the DELTA4 or ATLAS 5 offering a wide range ofperformances, following the examplesof ARIANE 4 and DELTA 2- they lead to the best performance-to-price ratio for numerous commercialsmall launchers like PEGASUS,ATHENA, M5 or the VEGA projectof ESA

For all these systems, solid rocket motorseasily deliver a large thrust level for a limitedmotor volume. Various grain shape optionsallow to design motors delivering a thrust lawshape highly adapted to launcherrequirements. Reproducible performances can

be obtained thanks to a deep knowledge ofmaterials behavior.As they are static systems with a limitednumber of components, the reliability of solidrocket motor is very high. It is also generallyadmitted that, even if costs are always twohigh, the cost range of SRM is one of the mostattractive of large propulsion systems.

The first part of this paper is devoted to anoverview, during the last five years, of thetypical uses of solid rocket motors for spacelaunchers.The second part is focused on solid rocketmotors themselves, their technicalcharacteristics and some productioninformation data.The third part emphasizes some deep trendson this business and presents a short summaryof the technical possibilities of this propulsionmode.

2 THE "MARKET" OF LARGESOLID PROPULSION FOR SPACELAUNCH

2 1 RATIONALE FOR SOLIDPROPULSION

The interest of Solid Rocket Motors for SpaceLaunchers is supported by several fundamentalcharacteristics described here after.

high operabilityThe motor or the stage can be manufactured,assembled and stored several months beforeuse. During take off procedure they do notneed any difficult or risky operations at launchpad like for example tanks filling up. Thischaracteristic is reinforced by the trends to useElectro-Mechanical Actuators (EMA) forThrust Vectoring and Control function (TVC)avoiding some issues linked to high pressurehydraulic systems. A SRM powered stage is in

39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit20-23 July 2003, Huntsville, Alabama

AIAA 2003-4963

Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2

fact very close to the PLUG and PLAYconcept.For many years this high operability hasfavored the selection of solid propulsion formany deterrent force missile systems.

high reliabilitySeveral basic technical principles and designrules of SRM's are the basic sources of an highreliability level. Following major points can beunderlined:

- no movable parts except the flex-sealwhose design is generally 'robust'- rather simple mechanical loading casewith relevant acceptance tests ( caseproof test, flex-seal test…)- separation of thermal protectionfunction and mechanical functionexcept for the throat area

A general comparison with other rocketpropulsion modes is not easy to perform but itis more or less recognized that SRM presentsan overall high reliability.

high densityThe following table illustrates this naturaladvantage of solid propellant versus otherchemical propellants.

PROPELLANT HTPBSOLID

UDMHN2O4

LOXRP1

LOXLH2

Density 1.80 1.15 1.05 0.35

Theoretical ISPvacuum

310 s 340 s 360 s 460 s

Density x ISP 558 s 391 s 378 s 161 s

VOLUME RATIO REF. X 1.5 X 1.5 X 3

Figure 1 : Propellant characteristics

The 'Massic' Specific Impulse of solidpropellant is the lowest but the 'Volumic'Specific Impulse is the highest.

Compared to Solid Propellant a 'Storable'UDMH/N2O4 or 'Semi-Storable' LOX/RP1propellant will need 50 % more volume and a

cryogenic LOX/LH2 propulsion stage wouldrequire at least 3 times this volume.When large propellant mass are necessary, asfor a stage one mission, this parameter willlead to a more or less large volume for thestage and of course impact on the stage cost.For example a large propellant tank will implya large and probably more expensive inter-stage structure than a small propellant tank.

2 2 PAST 5 YEARS MARKETREPORT

The table in annex 1 reports the market ofsolid rocket motors for space launcher.All the Space Launches relying on Large SolidRocket Motors of the last five years, 1998 to2002, are 'tentatively' reported.A mass criteria of 7.5 tons is taken as lowerlimit of this analysis for large SRM definition.

For this market analysis two very simplecriteria are used :- the number of solid rocket motors whichhave flown- the flown propellant mass that is probablymore representative of the cost parameter dueto the large range of motor mass from 10 tonsto 500 tons

OVERALL MASS AND MOTORS

An average number of 110 large solid rocketmotors are flown each year.

YEARLY FLOWN LARGE SRM's

020406080

100120140160180

1998 1999 2000 2001 2002

Figure 2

This number is decreasing since 5 years inaccordance with the shrinking launch rate.

3

But the percentage of launcher using SRM ismore or less stable at about 50 % as it isillustrated in the following figure.

NUMBER OF SPACE LAUNCHES WITH LARGE SRM's

020406080

100

1998 1999 2000 2001 2002

TOTAL SPACELAUNCHESFLIGHT WITHLARGE SRM's

Figure 3

The overall propellant flown mass is close to8000 tons per year and relatively stable. Alarge amount, close to 50 %, of this mass iscoming from the RSRM/SHUTTLE flights.

PROPELLANT FLOWN MASS IN TONS

0

2000

4000

6000

8000

10000

12000

1998 1999 2000 2001 2002

Figure 4

BUSINESS LEADERSThe following table presents the situation ofthe four primary 'leaders' in terms of flownmass, the order of magnitude being in therange of 1000 tons per year or more.

1998 / 2002 FLOWN MASS

05000

1000015000

2000025000

30000

RSRM SRMU MPS GEM's OTHERS

TONS OF PROPELLANT

Figure 5The RSRM/SHUTTLE is the workhorse ofthis business with more than 50 % of theoverall propellant mass flown by year.

A second rank of wide uses is coming fromthe ARIANE 5/MPS and the TITAN/SRMU.It is to be noticed that the SRMU activity willsoon decline with the TITAN launcher end oflife.

The last important activity is linked to theGEM family cumulating the three products :GEM 40, 46 and 60. Of course the number offlown motors is much higher due to the strapon concept of small motors.

1998 / 2002 - FLOWN MOTORS UNITS

48 20 24

271196

0

100

200

300

RSRM SRMU MPS GEM's OTHERSMOTOR NAME

Figure 6

SECOND LEVEL (OTHERS)

The two following tables gives an idea of themarket situation for a 'second level' of businessactivity in space solid rocket motors. Theaverage flown mass for these motors is in therange of 100 tons per year.

It is to be noticed that CSD 7S, CASTOR 4Aand PAP motors are right now out of themarket due to the stop of launcher programs :TITAN 4A, ATLAS 2A and ARIANE 4.

1998 / 2002 FLOWN MOTORS

2

5248

13

48

0

10

20

30

40

50

60

CSD 7S CAST.4A PAP CAST 120 PSLV SRBA

Figure 7Castor 120 production remains small since thebeginning due to the astonished activity onsmall launcher segment in the US.

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SRBA market should grow up in the comingyear following the H2A launch rate.

1998 / 2002 FLOWN MASS (tons)

0

100

200

300400

500600

700

CSD 7S CAST.4A PAP CAST 120 PSLV SRBA

Figure 8

2 3 TYPICAL MISSIONS FOR SRM

Three typical uses of solid propulsion motorscan be identified from the current marketanalysis.

First StageSHUTTLE - TITAN - ARIANE 5 - H2AStrap-OnDELTA 2,3 and 4 and ATLAS 5Small LauncherATHENA, TAURUS, PEGASUS and M5.

The PSLV launcher being somewhat quite un-conventional, its motors are difficult toclassify. For the following chapters thefollowing classification is retained : S138 firststage, PSOM strap on, S7 small launcher.

First Stage

The following table gives an idea of thepercentage of the thrust delivered at take offby SRM's on heavy launchers.

LAUNCHER SHUTTLE TITAN4B

ARIANE5

H2A

S. R.M. RSRM SRMU MPS SRBA

% OF TAKEOFF THRUST

82 % 100 % 93 % ≈ 80 %

Figure 9These SRM's should really be granted of thefirst stage mission, avoiding the unfair 'stage 0'wording.

The following technical advantages of SRMcan be underlined to illustrates the interests ofsolid propulsion for such first stage missions :

- high thrust : Solid Rocket Motors easilydeliver the large thrust level required duringthe first two minutes of heavy launchers flight.There is a good fitting between the possiblepropellant burn rate and the optimum thrustduration of the first stage mission.

- adjustable thrust : the various grain shapeoptions allow to design motors with optimumthrust law shape. The thrust limitationnecessary to cope with the maximum dynamicpressure phase is obtained by the grain shapeitself with limited costs coming from castingtooling.As for numerous launchers the mission of thefirst stage is more or less always the same dueto system constraints, there is no realdisadvantage to offer a fixed predeterminedthrust shape.

- reproducible thrust : a deep materialbehavior knowledge and a raw material batchpolicy allow to obtain a satisfactory thrustshape reproducibility for the tandem designcommonly adopted on numerous heavylaunchers.

Strap-On

Several launcher architectures are based on acentral liquid core with additional Strap-Onboosters. It was the original design of theretired Ariane 4 and the DELTA 2. It has beenretained for the DELTA 4 and the ATLAS 5.

The following table illustrates the mainadvantage of these architecture. Largeperformances increases can be obtained withadditional boosters on the core stage operatingat take off or just after.

For the here after comparison, the selectedlauncher base versions are already equippedwith Strap-On in order to start from an

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efficient vehicle. The simple version of theselaunchers without Strap-On presents in factvery poor take off thrust and the resultingperformances are limited.

The addition of two solid Strap-On gives anextra performance ranging from 20 to 40 %. Itillustrates that the thrust level at take off is animportant driving parameter for performances.Of course such additional motors will generateextra costs but the overall performance to costratio will be significantly improved.

LAUNCHER ARIANE4

DELTA 4 ATLAS5

Basic version 42 L M+ (5,2) 521

GTO performance

% of increase

+21 % + 41 % + 27 %

Boosted version

Two more Strap-On

44 LP M+ (5,4) 541

Figure 10 : Impact of Strap-On

If the launcher system and the Strap-Ondesigns are highly compliant, it becomespossible to adapt the launcher performance tothe payload requirement by adjusting thenumber of strap on motors. Such a policyallows of course to fine tune the launcherconfiguration and to minimize the resultingprice.

Small Launchers

Commercial operational western smalllaunchers, PEGASUS - TAURUS - ATHENAand M5, are fully based on Solid RocketMotor technology. The current EuropeanVEGA program is also designed with threesolid propulsion stages.A general 'performance to cost' rationale iswidely recognized to solid propulsion modefor these missions.

From ballistic point of view SRM's appearswell adapted to the Low Earth Orbit trajectory

needs where high thrust level and shortburning time can be used efficiently.

A lot of Russian small launchers (KOSMOS,ROCKOT) are in fact reconfigured militarymissile and their available liquid propulsionmode is to be considered more as commercialopportunity than as definitely an absolute realadvantage.

3 S. R. M. "PRODUCTS" FORSPACE LAUNCHERS

Complying with the here-above analysis, threelarge classes of motors are defined in thischapter to present motors characteristics(design, technology and performances) andsome production data :

- large motors for heavy launchers stage 1- strap on motors for assisted take offlaunchers- motors for small launchers

SegmentationThe SRBA of the H2A that is currently thelargest monolithic motor with 65 tons ofpropellant. The very large motors aresegmented in order to limit the size ofmanufacturing means and the mass to behandled by crane. The current maximum massper segment is close to 150 tons.

There is no theoretical limitations to thisparameter. The main drivers are the cost andsize of the segment casting pit and the motorintegration building.

For example an 800 tons motor was designed,manufactured and ground tested by AEROJETin the 60's in Florida in the frame of the 260inch SRM program.

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3 1 OVERVIEW OF LARGE MOTORS FOR HEAVY LAUNCHERS

The three following tables present a summary of large solid rocket motors currently used onoperational heavy launchers. The PSLV S138 has been added to this table although if theGSLV/PSLV is not exactly an heavy launcher.

MOTORLAUNCHER

RSRMSHUTTLE

SRMUTITAN 4B

P230ARIANE 5

SRB-AH II A

S 138PSLV

Diameter 3,7 m 3,2 m 3,0 m 2,5 m 2,8 m

Motor length 38 m 31 m 27 m 12 m 20 m

Propellant mass 503 t 314 t 240 t 65 t 138 t

Binder PBAN HTPB HTPB HTPB HTPB

A.P. / Al. 70/16 69/19 68/18 68/18 -

Segment 4 3 3 1 5

MEOP pressure 70 b 86 b 69 b 118 b 59 b

Case material D6AC Carbon D6AC Carbon Maraging

Throat material Phenolic Phenolic C/C C/C PhenolicActuation Flex-seal

HydraulicFlex-sealHydraulic

Flex-sealHydraulic

Flex-sealE.M.A..

LiquidInjection

Motor Inert mass 68 t 29 t 29 t 6 t 18 t

Nozzle Σ 7,5 16 11 18 8

Figure 11 : Large Motors for Heavy Launcher - Design and Technology Data

MOTORLAUNCHER

RSRMSHUTTLE

SRMUTITAN 4B

P230ARIANE 5

SRB-AH II A

S 138PSLV

Propellant mass 503 t 314 t 240 t 65 t 138 t

Stage inert mass 88 t 36 t 36 t 11,5 t 30 t

Average thrust 11,8 MN 6,1 MN 5,0 MN 1,8 MN 3,6 MN

ISP – vacuum 267 s 284 s 275 s 280 s 269 s

Burning time 123 s 135 s 128 s 100 s 103 s

Vectoring ± 5° ± 6° ± 5° ± 5 ° ± 3°

Figure 12 : Large Motors for Heavy Launcher - Stage Propulsive Data

MOTORLAUNCHER

RSRMSHUTTLE

SRMUTITAN 4B

P230ARIANE 5

SRB-AH II A

S 138PSLV

Diameter 3,7 m 3,2 m 3,0 m 2,5 m 2,8 m

Motor length 38 m 31 m 27 m 12 m 20 m

Propellant mass 503 t 314 t 240 t 65 t 138 t

Industrial ATK ATK EUP IHI ISRONationality USA USA FRA/ITA Japan India

Ground tests 7 / 6 6 7 5 2

Qualification 80 / 88 1993 1995 2000 1991

Flown SRM ≈ 224 24 28 8 8

Flight Failure 1 - motor 0 0 0 0

Figure 13 : Large Motors for Heavy Launcher - Production Data - Fall 2002

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3 2 OVERVIEW OF SMALL LAUNCHER MOTORS

The three following tables present a short summary of solid rocket motors used on operational smalllaunchers. A criteria of a minimum mass of 7,5 tons was applied for this selection.Motors of small launchers derived from military missiles, START SHAVIT and MINOTAUR, arenot presented in this paper.

MOTOR CASTOR120

ORBUS21 D

ORION50 S (G)

ORION50SXL(G)

S7 HPM M14 M24 M34

Diameter 2,4 m 2,3 m 1,3 m 1,3 m 2,0 m 2,5 m 2,5 m 2,2 m

Motor length 9 m 3 m 8 m 10 m 3,5 m 14 m 7 m 3.5 m

Propellant mass 49 t 10 t 12 t 15 t 7,5 t 72 t 31 t 10 t

Binder HTPB HTPB HTPB HTPB HTPB HTPB HTPB HTPB

A.P. / Al. 69/19 68/18 88 % 88 % 86 % 68/20 68/20 68/20

Max pressure 100 b 55 b 59 b 75 b 62 b 59 b 59 b 59 b

Case material Carbon Carbon Carbon Carbon Aramid Maraging Maraging Carbon

Throat material C/C C/C - - - C/C Graphite Graphite

Actuation Flex-sealHydraulic

Tech-RollE.M.A.

Option(G)

Option (G) Flex-sealE.M.A.

Flex-sealE.M.A.

LITVC Flex-SealE.M.A.

Inert Mass 4,3 t 0,9 t 1,1 t 1,3 t 0,7 t 12 t 3,4 t 1,0 t

Nozzle Σ 17 / 24 64 40 40 69 11 31 Ext. 96

Figure 14 : Small Launcher Motors - Design & Technology Data

MOTOR CASTOR120

ORBUS21 D

ORION50 S (G)

ORION50SXL(G)

S7 HPM M14 M24 M34

Propellant mass 49 t 10 t 12 t 15 t 7,5 t 72 t 31 t 10 t

Inert mass 4,3 t 0,9 t 1,1 t 1,3 t 0,7 t 12 t 3,4 t 1,0 t

Average thrust 1600 KN 170 KN 450 KN 600 KN 190 KN 3800 KN 1250 KN 290 KN

ISP – vacuum 280/286 s 295 s 285 s 295 s 299 s 276 s 288 s 301 s

Burning time 83 s 158 s 73 s 69 s 110 s 50 s 71 s 102 s

Vectoring ± 5 ° ± 3 ° - - ± 4 ° ± 5° - -

Figure 15 : Small Launcher Motors - Propulsive data

MOTOR CASTOR120

ORBUS21 D

ORION50S (G)

ORION50SXL(G)

S7 HPM M14 M24 M34

Diameter 2,4 m 2,3 m 1,3 m 1,3 m 2,0 m 2,5 m 2,5 m 2,2 m

Motor length 9 m 3 m 8m 10 m 3,5 m 14 m 7 m 3,5 m

Propellant mass 49 t 10 t 12 t 15 t 7,5 t 72 t 31 t 10 t

Industrial ATK P & W ATK ATK ISRO IHI IHI IHINationality USA USA USA USA India Japan Japan Japan

Ground tests 2 X + 1 2 1 X + 3 3 3 3

Qualification 1993 1994 1990 1993 93/02 1995 1995 1995

Flown SRM 14 8 15 22 7 3 2 2

Flight Failure 1 - stage 0 0 0 0 1 - motor 0 0

Figure 16 : Small Launcher Motors - Production Data - Fall 2002

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3 3 OVERVIEW OF STRAP ON MOTORS

The three following tables present a short summary of solid rocket motors used as "strap on" onoperational launchers.

MOTOR GEM 40Ground

GEM 46Ground

GEM 60 AEROJETSRB

CASTOR4A XL

PSOM

Diameter 1,0 m 1,2 m 1,5 m 1,6 m 1,0 m 1,0 m

Motor length 11 m 12 m 13 m 17 m 12 m 11 m

Propellant mass 12 t 17 t 30 t 37 t 13 t 9 t

Binder HTPB HTPB HTPB HTPB HTPB HTPB

A.P. / Al. 88 % - - - 68/20 -

Max pressure 75 b 104 b 89 b - 54 b 44 b

Case material Carbon Carbon Carbon Carbon Steel Steel

Throat material C/C C/C C/C Phenolic Phenolic GraphiteActuation No Flex-seal

HydraulicFlex-sealHydraulic

No no No

Motor Inert mass 1,0 t 1,9 t 3,1 t - 1,9 t -

Nozzle Σ 11 14 11 - 9 8

Figure 17 : Strap On Motors - Design & Technology Data

MOTOR GEM 40Ground

GEM 46Ground

GEM 60 AEROJETSRB

CASTOR4A XL

PSOM

Propellant mass 12 t 17 t 30 t 37 t 13 t 9 t

Stage Inert mass 1,2 t 2,3 t 3,9 t 4,0 t 2,1 t 2,0 t

Average thrust 500 KN 600 KN 830 KN 1150 KN 600 KN 440 KN

ISP – vacuum 274 s 279 s 275 s 275 s 269 s 262 s

Burning time 63 s 77 s 92 s 94 s 59 s 44 s

Vectoring No ± 5 ° ± 5 ° No No No

Figure 18 : Strap On Motors - Propulsive Data

MOTOR GEM 40 GEM 46 GEM 60 AEROJETSRB

CASTOR4A XL

PSOM

Diameter 1,0 m 1,2 m 1,5 m 1,6 m 1,0 m 1,0 m

Motor length 11 m 12 m 13 m 17 m 12 m 11 m

Propellant mass 12 t 17 t 30 t 37 t 13 t 9 t

Industrial ATK ATK ATK AEROJET ATK ISRONationality USA USA USA USA USA India

Ground tests 7 3 3 4 3 -

Qualification 1990 1997 2000 2002 1993 1993

Flown SRM "660" 27 2 0 8 50

Flight Failure 1 - motor 0 0 0 0 0

Figure 19 : Strap On Motors - Production Data - Fall 2002

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4 GENERAL TRENDS

4 1 TECHNOLOGY

The following long term trends can be noticedfor large space solid rocket motortechnologies.

PropellantFormulation for the propellant is quite uniformwith a predominance of the followingcomposition:- binder : Hydroxyl-Terminated-Polybutadiene(HTPB)- reducer : Aluminum ( Al)- oxidizer :Ammonium Perchlorate (AP)

Only the RSRM shuttle motor whose originaldesign was performed in the beginning of the70's continue to use a PBAN (polybutadieneacrilic nitril) binder.

CasesLess and less motors are using metallic cases.Since 1995 all the newly developed motors arebased on carbon epoxy filament woundtechnology. Carbon case allows to increaseperformance through an inert mass decreaseor/and a specific impulse increase.The cost of carbon case is also on a decreasingtrend while the metallic based cases havereached some kind of asymptotic situation.

NozzlesFor actuated motors, there is a wide use of theflex-seal concept for nozzle steering and onlyvery few motors use the gas or liquid injectionprinciple, perhaps due to its limited vectoringangle capability lower than 3 °. Bothhydraulic an electric Thrust Vectoring andControl systems are used. Several Strap Onmotors presents fixed and canted nozzleswhen the core stage vectoring capability isstrong enough.

For throat parts several technical solutionsexists : graphite, graphite/phenolic,carbon/phenolic or carbon/carbon. All nozzlesinclude carbon phenolic insulators.

There is only one example of extendible exitcone for the M5 third stage.

4 2 PROPULSIVE PERFORMANCES

The following long term trends can be noticedfor large space solid rocket motor propulsiveperformances.

Specific impulse

From the ISP parameter point of view, twofamilies of motor can be identified dependingon the case design (excluding the PSOMmotor).

VACUUM ISP MINIMUMVALUE

MAXIMUMVALUE

METALLIC CASE 267 s 288 s

COMPOSITE CASE 274 s 301 s

Figure 20 : ISP range versus Case Techno.

The most recent motors present in fact thehighest ISP but the improvements remainsmall.

Inert mass

The inert mass ratio is defined hereafter asPROPELLANT MASS/TOTAL MASS.

Excluding the PSLV first stage motors(PSOM and S138), and in harmony with theISP parameters, two families of motor appear.

'STAGE' MASSRATIO

MINIMUMVALUE

MAXIMUMVALUE

METALLIC CASE 0.85 0.87

COMPOSITE CASE 0.89 0.91

Figure 21 : Mass Ratio versus Case Techno.

In fact three values of the total inert massexists, generating some misunderstanding :only the motor without the TVC, or including

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the TVC, or the full stage including also skirtsand lines.The here above figures are mostly related tothe stage values.

Propellant mass

The natural trend is an increase of propellantmass. It is illustrated by the following tablesshowing the evolution of Strap On motormass.

STRAP ON YEAR OFQUALIF.

PROPELLANTMASS

GEM 40 1990 12 t

CASTOR 4A XL 1993 13 t

GEM 46 1997 16 t

GEM 60 2000 30 t

AEROJET SRB 2002 37 t

Figure 22 : Strap on PropellantMass Growth Trend

It is the same type of evolution for the largemonolithic motors. The P80 motor is underdevelopment for the VEGA small launcher.

LARGEMONOLITHIC

YEAR OFQUALIF.

PROPELLANTMASS

CASTOR 120 1992 49 t

M 24 1995 31 t

SRBA 2000 65 t

P 80 2005 88 t

Figure 23 : Large Monolithic Propellant Mass Growth Trend

4 3 FAILURE ANALYSIS

Four motor failures are reported in the hereabove tables. This number is very lowcompared to the overall number of flownmotors reported close to 1100. The resultingreliability figure is higher than 0.996.

The CHALLENGER SHUTTLE accident wasinitiated by a leakage at an inter-segmentjoining favored by an 'icy' weather conditions

for launch. Unfortunately this burn-throughdamaged a strut leading to a large break of thecentral liquid core and the final explosion. Thedamaged booster itself continued its flightwith an enlarging lateral hole while the otherbooster acted quite nominally. The inter-segment joining zone was re-designed to avoidthe here-above issue.

The CASTOR 120 failure came from a fire inthe TVC oil exhaust duct that destroyed theTVC harness leading to a loss of vehiclecontrol.

For the M14 first stage, a vehicle loss wasgenerated by a failure of the graphite throat.The throat insert has been turn to C/Ccomposite to solve this issue.

On a DELTA 2 rocket one a crack appearedduring take off on one of the GEM 40 leadingto the vehicle explosion. This crack wascoming from an internal case damage after thehydro-proof test. Reinforced specificultrasonic inspections are now implemented ontheses motors.

4 4 DEVELOPMENT PHASE

Figure 24 indicates the number of groundfiring tests performed to qualify a motor sincethe 70's.

The number of firing tests for new motordevelopment is decreasing and is close to 4right now. It is the same for development ofup-rated versions of existing motors thatrequires only an average of two firing tests.

Such an evolution is possible due to a highlevel of modelisation : propellant combustion,internal ballistic flow (multi-phase), ignitiontransient, thermal analysis of heat transfers tointernal insulators , mechanical analysis of caseand nozzle structures.

11

LESS AND LESS TESTS AT BENCH

0

2

4

6

8

10

12

14

1965 1970 1975 1980 1985 1990 1995 2000 2005

SRM QUALIFICATION DATE

GR

OU

ND

FIR

ING

TES

TS NEW DESIGN

UPGRADES

TITAN 5S

7S

C4

HPM5,5S

SRM RSRM

SRMU

MPSC4A

C4B

G40

G60G46

SRBA

MPSE

PAP4

PAP3

H2

PSLVC120

C4AXL

OR50S

OR50SXL

SPACE SOLID ROCKET MOTORSMASS > 7,5 TONS

O21

O21D

AEROJET

Figure 24 : Ground firing Test Decrease Trend

6 CONCLUSIONS

Solid rocket propulsion mode is used on all the active commercial 'western and japanese' spacelaunchers mainly because high thrust level and adapted thrust law can be proposed at affordablecosts with a high level of reliability.Three main missions are currently identified : stage one of heavy launchers, strap on motors to assistliquid core stage and motors for small launchers.In a short term future, continuous but limited performances progress can be expected, throughreduced development programs. Large motors sizes appears also as the most attractive for SpaceLaunch Business.

REFERENCES

(1) PROSPECTIVE DEVELOPMENTS – SOLID PROPULSION - D. MUGNIER - CNES/DLA– Launcher propulsion towards the year 2010 -Bordeaux symposium 11 and 12 JUNE 9l(2) AIAA 91 – 3391 SOLID PROPULSION OPTIONS FOR NATIONAL LAUNCH SYSTEM.R.D. SAUVAGEAU – THIOKOL - BRIGHAM - UTAH(3) INTERNATIONAL REFERENCE GUIDE TO SPACE LAUNCH SYSTEMS- AIAA –S.J.ISAKOWITZ – SECOND EDITION -(4) SOLID ROCKET PROPULSION : STATUS AND EVOLUTION - AIAA ProfessionalDevelopment Course - 20/21 july 2000 huntsville ALABAMA.(5) Jane's Space Directory(6) DEVELOPMENT OF M-5 ROCKETMATSUO / KOHNO / ONODA - ISAS - IAF 97 V 10 8

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ANNEX 1 : 1998 / 2002 REPORT ON SPACE LAUNCHES WITH SRM

LAUNCHER MOTOR SRM SRM YEAR FLOWN FLOWNNAME MASS number 1998 1999 2000 2001 2002 MOTORS MASS

unit tonsSHUTTLE RSRM 500 2 5 3 5 6 5 48 24000ARIANE 42 P PAP 9,5 2 1 3 1 10 95ARIANE 44 LP PAP 9,5 2 2 3 3 1 18 171ARIANE 44 P PAP 9,5 4 2 1 2 20 190ARIANE 5 MPS 237 2 1 1 4 2 4 24 5688DELTA 2 7320 GEM 40 12 3 1 1 2 2 18 216DELTA 2 7420 GEM 40 12 4 3 6 1 1 44 528DELTA 2 7925 GEM 40 12 9 8 3 3 4 2 180 2160DELTA 3 GEM 46 17 9 1 1 1 27 459DELTA 4 M+ GEM 60 30 2 1 2 60ATLAS 2AS CASTOR 4 A 10 4 2 4 3 3 1 52 520ATLAS 5 AEROJET 41 5 0 0TITAN 4A CSD 7S 296 2 1 2 592TITAN 4B SRMU 313 2 1 3 2 3 1 20 6260ATHENA 1 CASTOR 120 49 1 1 1 2 98

ORBUS 21 D 10 1 1 1 2 20ATHENA 2 CASTOR 120 49 2 1 2 6 294

ORBUS 21 D 10 1 1 2 3 30PEGASUS ORION 50 SXL 15 1 6 3 2 1 12 180TAURUS CASTOR 120 49 1 2 1 1 1 5 245

ORION 50 SG 12 1 2 1 1 1 5 60PSLV S138 138 1 1 1 1 3 414

PSOM 9 6 1 1 1 18 162S7 7 1 1 1 1 3 21

GSLV S138 129 1 1 1 129H2 SRB 59 2 1 1 4 236H2 A 202 SRBA 66 2 1 1 4 264H2 A 2024 SRBA 66 2 2 4 264

CASTOR 4A XL 13 4 2 8 104J1 SRB 59 1 0 0

M 23 10 1 0 0M5 M14 71 1 1 1 2 142

M24 31 1 1 1 2 62M34 10 1 1 1 2 20

SHAVIT STAGE 1 13 1 1 1 2 26STAGE 2 9 1 1 1 2 18

START STAGE 1 23 1 1 1 2 46STAGE 2 11 1 1 1 2 22

FLIGTH WITH SRM 46 43 33 35 27 559 43796

TOTAL LAUNCHER FLIGHTS 82 78 85 60 65 TOTAL TOTALFLOWN FLOWN

PERCENTAGE OF FLIGHT WITH SRM 56% 55% 39% 58% 42% SRM MASS

1998 1999 2000 2001 2002YEAR

NUMBER OF FLOWN SRM 154 128 95 107 75 559

MASS OF FLOWN SRM 8674 7125 9149 10207 8641 43796