Solid Propulsion for Space Application an Updated Road Map

Acta Astronautica 66 (2010) 201 -- 219

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Acta Astronautica

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Solid propulsion for space applications: An updated roadmap

Jean-Francois Guerya,∗, I-Shih Changb, Toru Shimadac, Marilyn Glickd, Didier Bourye, Eric Robertf ,John Napiorg, Robert Wardleh, Christian Péruta, Max Calabroi, Robert Glickd, Hiroto Habuc,Nobuhiro Sekinoj, Gilles Vigierk, Bruno d'Andreal

aSNPE Materiaux Energetiques, FrancebThe Aerospace Corporation, USAcJAXA, JapandConsultant, USAeSNECMA Propulsion Solide, FrancefCNES, FrancegAerojet, USAhATK, USAiConsultant, FrancejIHI Aerospace, JapankAstrium ST, FrancelAvio Spa, Italy

A R T I C L E I N F O A B S T R A C T

Article history:Received 6 February 2009Received in revised form25 May 2009Accepted 26 May 2009Available online 28 July 2009

Keywords:Solid propulsionRoadmapSolid propellantNozzleSpace propulsionComposite case

For the last 50 years solid propulsion has successfully created a multitude of small launch-ers and many first stages or boosters for heavy launchers with low risk, high performance,competitive cost, superb storability, and “instant” readiness in many countries. Technicalsupport for these successes arose from simple designs, very high thrust levels, and lowdevelopment and operation costs/risks. The first solid propulsion roadmap based on thesefoundations and rational projections was published in 2000 [A. Davenas, D. Boury, M. Cal-abro, B. D'Andrea, A. McDonald, Solid propulsion for space applications: a roadmap, in: 51stInternational Astronautical Congress, paper IAA-00-IAA.3.3.02, October 2000]. Moreover,subsequent information supports its enabling technologies (high strength composite cases,energetic material processing based on continuous mixing, low density insulation, reducedactuator energy requirements, and advanced detailed simulations) and applications (firststages, strap-on, add-ons, small launchers, and niche space applications). Missions currentlydevoted to solid propulsion and plans for present and future launchers and explorationmission developments in the USA, Japan, and Europe are sketched and targeted improve-ments, and potential breakthroughs are discussed.

© 2009 Elsevier Ltd. All rights reserved.

1. History

1.1. Roadmap2000: recapitulation

The first roadmap for solid propulsion space applica-tions [1] (Roadmap2000 herein) presented solid propulsion's

∗ Corresponding author. Present address: Centre de Recherche duBouchet, 91710 Vert le Petit, France.

E-mail address: [email protected] (J.-F. Guery).

0094-5765/$ - see front matter © 2009 Elsevier Ltd. All rights reserved.doi:10.1016/j.actaastro.2009.05.028

major advantages of simplicity, cost effectiveness (impulsevs. total cost), and reliability (relative to liquid rockets [2](LRs)) and identified space applications as first stages, strap-ons, add-ons, small launchers, and niches i.e. Mars PathfinderLander's retro motors and bag inflators. Solid propulsion'scost effective, low risk delivery of very high thrust levelsenables launchers with solid booster(s) (and liquid rocketcore's) and mixed solid and liquid propulsion upper stagearchitectures that offer adequate performance, missionflexibility, and reduced cost and risk. Enabling technologies

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202 J.-F. Guery et al. / Acta Astronautica 66 (2010) 201 -- 219

Nomenclature

ACS attitude control systemADN ammonium dinitramideAMMO poly(3-azidomethyl-3-methyloxetane)AP ammonium perchlorateBAMO poly(3,3-bisazidomethyloxetane)HNIW hexanitro hexaaza isowurtzitane (CL-20)HTPE hydroxyl terminated polyetherHTPB hydroxyl terminated polybutadiene

IHPRPT Integrated High-Payoff Rocket PropulsionTechnology

HMX cyclotetramethylene tetramineHNF hydrazinium nitroformateGAP glycidyl azide polymerNMMO poly(3-nitratomethyl-3-methyloxetane)PolyGLYN polyglycidyl nitrateRDX hexahydro-1,3,5-trinitro-1,3,5-triazineSRM solid rocket motor

were identified as high strength composite cases, signifi-cantly reduced actuation loads (and electric actuators), lowdensity insulation, robust propellant processing/motor cast-ing simulations, and robust continuous mixing propellantprocessing. Efficient production was identified as crucial forreduced cost and improved reliability and performance withsynergisms among controlled mixing/casting, robust predic-tions of cast grain characteristics and burn back simulations,and minimal insulation requirements. Higher energy propel-lants, examined in the context of the ∼20 years necessary toachieve industrial level production, hexogen (RDX) supple-ments for ammonium perchlorate (AP) with glycidyl azidepolymer (GAP) binder. Observed cost, hazards, and safetycharacteristics of both ingredients and propellant would bedecisive.

A rough calendar for possible demonstrations was pro-vided:

• In 2005 employing circa 2000's proven technologies.• In 2015 employing higher strength case materials, im-

proved propellant processing and simulation technologies,and RDX, GAP augmented propellants.

• In 2025 employing very high strength materials, matureprocessing and simulation technologies, and advancedpropellants.

Finally, international cooperation could materially aidprogress toward (i) new energetic materials and propellants,(ii) hazards, safety characteristics, and related standards/regulations, and (iii) cooperative technology demonstra-tions.

2. Solid vs. liquid propulsion

2.1. General

Current and projected space launch systems typically em-ploy architectures that combine solid and liquid rockets todeliver payloads to orbit reliably and cost effectively. There-fore, these mixed mode architectures' performance, reliabil-ity, etc. depends on both modes' attributes. Since chemicalpropulsion's primitive process is chemical energy depositionin propellant products, liquid rockets' separation of reac-tants until post injection and solid rocket (SR) propellants'well mixed reactants at 0.1–600�m length scales definethese system's general characteristics, potentials, and chal-lenges e.g. the set of “natural” liquid propellant is muchlarger than solid propellants' (inter-ingredient compatibility

is not required), the performance of solid rocket propel-lants is more dependent on innovative chemical synthesesand their large scale industrialization. In addition, the liquidrocket's visible mechanical complexities and a solid rocket'sinvisible propellant complexities have unique consequencesi.e. solid propellants' critical dependence on propellant for-mulation, adequate characterization, subsequent processinginto loaded motors, and instant readiness.

2.2. Reliability

Since the reliability ofmixedmode launcher architecturesdepend on the reliability of both modes, historical informa-tion circa July 31, 2008 were examined to reassess solid andliquid rocket reliability. For these data there were 416 fail-ures and 4506 successes in worldwide orbital space launchessince 1957. Although root causes for the 416 failures aredifficult to determine (and categorize), launch failures since1980 have been investigated, and data compilations showthere were 140 failures and 2497 successes for worldwidespace launches between January 1, 1980 and July 31, 2008.

A space launch failure can usually be attributed toproblems associated with a functional subsystem, such ascommand and control, environmental protection, electrical,guidance-navigation and control, ground support equip-ment, propulsion, separation, structures, telemetry, thrustvectoring and attitude control, and tracking and flight safety.In some cases failure is ascribed to unknown causes, whensubsystem failure information is not available. Propulsionsubsystem problems are presented in Table 1. The 82 of the140 worldwide launch failures in 1980–2008 were failuresof the propulsion subsystem. The 18 of the 33 US failures and41 of the 74 CIS/USSR failures in 1980–2008 were failuresof the propulsion subsystem. The propulsion subsystem isthe heaviest and largest subsystem of a launch vehicle, andits failure can be divided into failures in solid rockets (SR)and liquid rockets (LR). Out of the 82 propulsion failures in1980–2008, 15 were SR and 67 were LR propulsion subsys-tems. There were 662 launches with SRs and 2462 launcheswith LRs. Therefore, the success rate is 97.73% for SR and97.28% for LR propulsion subsystems. In Table 1, the sum ofnumber of launches with SRs (662) and with LRs (2462) isgreater than the total number of worldwide space launches2637 in 1980–2008, because some hybrid launchers useboth solid rocket motor (SRM) and LRE for the same launch.Clearly, success rates for SR and LR subsystems for the last29 years (1980–2008) were essentially identical i.e. SR andLR subsystem reliabilities have effectively converged.

J.-F. Guery et al. / Acta Astronautica 66 (2010) 201 -- 219 203

Table 1Success rate of solid and liquid propulsion subsystems in space launches.

Country Propulsion Non-propulsion Total

Solid Liquid

Succ Fail Succ Fail Succ Fail Succ Fail

USA 462 6 518 12 606 15 588 33CIS/USSR 6 0 1557 41 1605 31 1563 73Europe 101 0 173 9 182 1 173 10China 0 2 95 3 100 3 95 8Japan 54 2 37 2 58 1 54 5India 19 1 15 0 20 5 19 6Israel 5 1 0 0 6 1 5 2Brazil 0 2 0 0 2 0 0 2N. Korea 0 1 0 0 1 0 0 1

Total 647 15 2395 67 2580 57 2497 140

80-08 (%) 97.73 97.28 97.84 94.69

2.3. Summary

Solid propulsion offers cost effective, large thrust capa-bilities, but operating times limited to 2min. This technologyis well fitted for zero stage and boosters. Liquid propulsionoffers high specific impulse and restart capabilities, whichis well fitted for upper stages. In mixed mode launcher ar-chitectures solid rockets' very large thrust capabilities andliquid rocket's superior Isv and smaller thrust levels offerunique optimization opportunities absent from mono-modelaunchers. Therefore since mixed mode launcher architec-tures have demonstrated effectiveness in the past, theyshould become more effective in the future as the modes'reliability converges to the higher bound and the solidrocket mode's performance and cost effectiveness increaseswith improved case materials, propellant processing, anddetailed simulation.

3. SRM boosted launch vehicles

Most existing and “in development” launchers employSRM (Table 2). Moreover, their solid rockets' technologiesare effectively identical and improvements have conformedto Roadmap2000.

3.1. Existing launchers improvements

3.1.1. Ariane 5 improvementsA5 mid-evolution (ME) could be operational at the end of

the next decade (2018) if the decision to develop is decidedfavorably by 12/2011.

Since this solution will fully optimize Ariane 5's stagingfor the upper stage, improved performance would then re-quire modification of Ariane's lower stage and studies havecreated interest for these improvements. The “P80”, Vega'scurrent first stage, was initially a solid rocket booster demon-strator focused on monolithic composite case, electricalactuators, low couple nozzle, etc. Therefore, P80 was ademonstrator of a possible MPS evolution e.g. MPS2—successfully static tested in 2006 and 2007 in French Guyana.

There are multiple MPS2 projects and they offer a widerange of evolutions. Relative to performance, the compositecase alone offers a mass reduction that exceeds 40%.

The performance of SR boosters, even when the basicdesign is fixed, can be enhanced by processing improve-ments that include continuous propellant mixing, reducedpyrotechnic mass, application of “classical” industrial means,and detailed simulations that optimize their impacts. There-fore, these aspects will be considered.

3.1.2. H-II improvementsAmong Japan launch vehicles, the H-IIA has been support-

ing satellite launch missions as a major large-scale launchvehicle with superb reliability.

The H-IIB launch vehicle is an upgraded version of thecurrent H-IIA's launch capacity and is expected to enablefuture missions that include cargo transport to the Interna-tional Space Station (ISS) and to the Moon.

The H-IIB launch vehicle has two major functions. Oneis to launch the H-II Transfer Vehicle (HTV) to the ISS.The HTV will carry necessary daily commodities for thecrew astronauts and experimental devices, samples, spareparts, and other necessary research items for the ISS. Theother major function is to respond to broader launch needsthrough adroit utilization of H-IIA and H-IIB launch vehiclesin concert. Moreover, H-IIB's larger launch capacity enablessimultaneous launches of multiple satellites per missionthereby significantly reducing satellite launch costs. Theseadvances will enhance the Japanese space industry's vitality(Fig. 1).

The H-IIB launch vehicle is a two-stage rocket that em-ploys two liquid rocket engines (LE-7A) in the first-stage (onefor the H-IIA) and four strap-on solid rocket boosters (SRB-A) grained with hydroxyl terminated polybutadiene (HTPB)propellant (the standard H-IIA has two SRB-A strap-ons). Inaddition, the H-IIB's first-stage body has been expanded to adiameter of 5.2m (4m for the H-II). It also extends the firststage's total length by 1m from the H-IIA's. As a result ofthese enhancements, the H-IIB loads 1.7 times more propel-lant than the standard H-IIA.


Table 2SRM boosted launch vehicles in the world.

Status Country/launchers

Existing or ongoing development USA: STS, Delta IV, Atlas V, Delta II, Taurus, Minotaur I, Pegasus XLEurope: Ariane 5, VegaJapan: H-IIA, H-IIBOthers: Start-1, CZ-2C/SD, KT-1, GSLV, PSLV, Shavit-2, Shavit-1, VLS-1, TPD-1

Next launchers mid-term (2020) USA: Ares I, Ares V, Minotaur IVJapan: small launcherOthers: GSLV-MkIII

Fig. 1. H-IIB launch vehicle.

The H-IIA, H-IIB development strategy of clustering en-gines of demonstrated reliability reduces development timeand cost without performance penalties.

H-IIB development began FY2003 and the first launchwill occur FY2009.

3.1.3. Vega launcher improvementsVega's expected maiden flight date is 2009. However,

since its actual performance near 1.5 t is low, different evolu-tions are under study to improve performance. Among theseevolutions, one would increase the P80 first stage's 88 t ofpropellant to the P100's 100 t of propellant at constant di-ameter or extend the Z23 solid rocket motor's 24 t of pro-pellant to the Z40's 40 t of propellant by a diameter changefrom 1.9 to 2.6m.

3.2. New generation launchers (NGL)

3.2.1. The building block alternative in USAInterest in a building block launcher (BBL) is easy to un-

derstand e.g. limited development time and low risk. How-ever, this interest is conditioned by launcher suitability fora required mission. A perfect example of a BBL approachis the Ares I and Ares 5 launch vehicle in the USA. For thenext launchers for the Constellation program, it was decidedto use existing propulsive technologies, and in this framethe RSRM is the basis for the Ares 1 first stage, and theAres 5 boosters (coupled with a five (or six) RS68 cryogenicstage).

Compared to existing RSRM, several motor design mod-ifications are required to meet Ares 1 requirements and inparticular: ballistic performance, operability improvements,enhanced reliability, regulatory compliance, and replace-ment of obsolete materials and processes:

• design features in the motor,• propellant grain (one additional segment, grain shape and

propellant burn rate evolution to meet thrust and pressurelaws requirements, . . . ),

• nozzle throat and exit cone designs modifications, and• replacement of materials used in the manufacture of the

internal insulation, the case bond liner, and the O-ringsused to seal the joints between motor segment (use ofasbestos free material, new lower-temperature materialsin the O-rings, . . . ).

To validate them, the new materials and processes willbe first applied on subscale specimen or RSRM for groundtests. Then four DM tests are planed for the design validation(Fig. 2).

3.2.2. The building block alternative in EuropeIn Europe the BBL approach has been studied for two

main reasons: elaboration of a complementary launcher (forAriane 5) and replacement of Ariane 5 to better fit a differ-ent market (definition of a single payload launcher for GTO,called 1

2 Ariane 5).Five (5) stages developed for Ariane 5 are available:

• EAP boosters (240 t of solid propellant, metallic casing),• core stage EPC (170 t of cryo-propellant) with Vulcain 2

engine (135t thrust),• storable upper stage EPS (10 t of N2O4/MMH propellant)

with Aestus engine (3 t),• cryogenic upper stage ESCA (14 t of cryo-propellant) with

HM7 (7 t of thrust) engine derived from Ariane 4, and• Vinci demonstrator fit for a future high performance upper

stage (18 t of thrust).

In addition, Vega's development brings additional stages:

• The P80, partially an EAP demonstrator for advancedboosters, and the first stage of Vega (88 t of solid propel-lant, filament wound (FW) casing).

• The Zefiro 23 and Zefiro 9 (24 and 10 t of solid propellant).


Fig. 2. Ares launch vehicles.

In the BBL approach, several solutions have been com-pared: small Ariane 5 with replacement of the EAP by P80,2-stage configuration with double Vulcain, etc. The sim-pler, more cost effective solution has always been the “SolidBBL”, three stages, composed with an EAP (P240) as firststage, a P80 as second stage, and an upper cryogenic stage(Fig. 3).

In studies paralleling the above, designs to reach 5 t with-out supplementary booster(s) and 8 t (for max commercialpayload) by adding SRBs have been examined. Numerouspractical solutions exist e.g.

• Re-use existing projects to increase the performance ofthe building blocks P240 and P80.

• Add small boosters to the first stage.• Derive the building blocks and optimize them for the

BBL, even if A5 compatibility is lost (if it is a replacementlauncher there is no need to maintain compatibility)(Fig. 4).

3.2.3. Future heavy launchersDifferent concepts are considered: liquid stages, reuse of

four A5 SRM . . . In each configuration, solid rocket boostersmay be employed.

An original design is composed of a big SR first stage.Simulations have always shown this configuration presentsa recurring cost increase when compared with hydrogen ormethane configurations. Therefore, the goal here is to re-duce development costs (prohibitive for a huge monolithicstage) by exploiting either innovative technologies and pro-duction processes or an intermediate building block con-figuration (reuse existing production facilities and slowlymodify and optimize building blocks to reach requirementsby either retaining a three stage configuration or introducinga two stage using multiple segment boosters). A 3-segmentfilament wound booster with 435 t of propellant would beequivalent in performance to an optimized 370 t monolithicbooster.

Fig. 3. BBL.

3.2.4. Small launch vehicle (Japan)Scientific missions using small satellites are being pro-

posed for the next several years. Their functions includespace observation, Earth monitoring, and lunar and plan-etary exploration. Moreover, applications that reinforcetechnical foundations by demonstrating components, space-craft, design for flight capabilities, etc. are of interest. Thesemissions require a variety of orbits: low-Earth, polar, highlyelliptical, and transfer for lunar and planetary missions.Therefore, a versatile launcher with these capabilities isdesirable and an all solid system's mission capabilities,


Fig. 4. BBL with six strap on boosters (P30).

simplicity, readiness, and cost effectiveness recommendit over liquid and mixed mode systems. Consequently, anew solid rocket launch vehicle designed to realize variousmissions with frequent, timely, short lead-time launches atlow-cost is being considered in Japan.

An important aspect of this vehicle's development is tomaintain and improve solid rocket and related technolo-gies. Japanese system technology has been fostered throughsolid rocket developments from the 1955 pencil rocketto the present M-V launch vehicle. Moreover, this systemtechnology covers the vehicles' entire life-cycle e.g. design,manufacture, integration, assembly, and launch operations.This total system viewpoint is necessary to achieve cost-effective, highly reliable, and optimally performing solid-rocket system technologies of the future. Moreover, thisapproach strengthens the solid propulsion community's fun-damental technology bases as well as develops solid rocketmotors for sub-booster, first stages, upper stages, niche ap-plication, etc. and all solid vehicles for many applications(Fig. 5).

Currently, the new all solid launch vehicle has threestages and can launch a satellite weighing 1.2 t into LEOand 0.6 t into the transfer orbit to SSO. The first stage of therocket is the SRB-A employed as sub-booster of Japan's flag-ship launch vehicle H-IIA. Although the SRB-A's thrust is lowfor this new first-stage application, it is extremely cost ef-ficient. In contrast, the launch vehicle's upper stage motorsare new designs based on the M-V's upper-stage-motors.To achieve high-performance and low-cost simultaneously,each stage's size and performance is optimized to maximizethe orbiting satellite's mass. Fig. 6 presents an artist's imageof the new all solid launcher: it is about 24m high, 2.5mdiameter, and weighs about 91 t.

From viewpoints of responsiveness and operability,manufacturing and preparation time should be minimizedwithin adequacy constraints. Ideas to realize this includeimproving launch operation efficiency for rocket assemblyand checkups with compact ground inspection and test fa-cilities. The key to this concept is avionics as well as a newlydesigned rocket structure that enables easy rocket opera-tions. For the new rocket, networked avionics and a more“intelligent” rocket will enable autonomous checkups prior

to launch. In future extensions with avionics of enhanced“intelligence,” it is expected that launch control can bedrastically simplified.

Currently, micro-satellites that weigh less than 100kg arelaunched as piggy-back payloads. In this approach opportu-nities and launch windows are strictly limited because theirlaunch priority is very low. Therefore, it is currently difficultto place micro-satellites in their ideal orbits. Consequently,availability of a small, low-cost solid launch vehicle to launchmicro-satellites is desirable. Moreover, air launching or seabased launching enhances orbit flexibility and responsive-ness. Furthermore, the simplicity of ground equipment re-quired for launch operation identifies a solid rocket systemas an excellent candidate.

3.2.5. Airborne micro-launchersAnalyses of classical micro-launchers' inadequacies found

three key reasons: program organization, scale effects, andground installations (including launch pad). Therefore, air-borne launchers with reduced ground dependencies haveadditional performance potentials e.g.

• Significant initial velocity ( ≈ 150–200m/s for most cases;with supersonic aircraft 300–800m/s).

• Reduced gravity losses when altitude of separation fromaircraft is large.

• Improved nozzle performance because atmospheric pres-sure at separation from aircraft> 1bar enabling increasednozzle expansion ratio.

• Reduced losses from safety concerns i.e. flight launch overuncritical areas, etc.

• Reduced ground environment impacts.

3.2.6. Landers, Jettison motorsFor exploration missions (Mars or Moon insertion; soft

landing; ascent vehicle) liquid propulsion seems to be moresuitable due to versatility and adaptability. However solidpropulsion may be used, with a combination with atti-tude control system (ACS) if needed. The main advantagefor this technology is simplicity, superb storability, en-ergy density and high thrust capability. Thrust magnitude


Fig. 5. Japan small launch vehicle.

Fig. 6. MLA Trimaran (Snecma and Dassault Aviation).

control, limitation of scatterings, etc. would enhance itspotentials.

NASA's Vision for Space Exploration has multiple solidpropulsion elements that are currently in production. Aero-

Fig. 7. Aerojet's Orion Jettison Motor hot fire demonstration test in Sacra-mento, California.

jet has completed two successful hotfire demonstrations ofthe full scale Orion Jettison Motor that is being designed toseparate the spacecraft's launch abort system from the crewmodule during launch. These demonstration tests serve aspathfinders for the delivery of the rocket motor that will beused for the first full-scale test of the launch abort systemat the US Army's White Sands Missile Range in New Mexico(Fig. 7).

In addition to the work being performed on the OrionLaunch Abort System and ARES I and V launch vehicle, theNASA Constellation program has multiple opportunities forsolid rocket motor developments within the next severalyears.

4. Launch vehicle sensitivity analysis

4.1. Composite material strength

The Tsiokolvsky equation �V = g0Isv ln(Mi/Mf ) revealslaunch vehicle velocity increment �V depends on a stage'sdead mass Mi and its propellant's delivered specific impulseIsv. Therefore, since advanced solid rocket motor technolo-gies have converged on filament wound cases, movable(flex-seal) nozzle(s), and HTPB propellants, their domainof application is limited to small launch vehicles, strap-onboosters, and niches. For these applications their structuralindex and their ability to deliver high thrust at low totalcost provide a decisive advantage relative to classical liquidrockets. Composite filament cases, successfully employed inapplications for more than 40 years, have created a smallrevolution in SRM design e.g. the case's performance crite-rion pressure∗internal volume/solid propellant case mass


Table 3Comparison between current Ariane 5 MPS and composite case upgradedversion.

Current MPS Upgraded MPS

Mp 237.7 247.8Mi 30.8 27.5Isv 275.3 280.0Ae 6.99 6.99Tb 128 125Pmax 6.1 8.5

0.450

0.0160.165

0.314

0.055 0.048 StructureIgniterTh InsNozzleMiscTVC

Fig. 8. Mass breakdown for an advanced 90 t class SRM.

for a composite case is 5 times a metallic one's. Conse-quently, since this enables increased operational pressure,Mi (case mass) reduction is synergistically combined withincreased Isv per the Tsiokolvsky equation above.

An EADS (Astrium today) [3] study, whose results aretabulated below (Table 3), reveals the performance increaseassociated with composite cases.

With these tabular values, the Ariane 5 ECB version's per-formance gain is ∼30%. Consequently, new launch vehiclesshould naturally migrate toward high strength compositecases.

The Fig. 8 present a typical mass breakdown for a SRM(90 t class) with a filament (carbon ofmediumperformances)wound case. Therefore, three major potentials exist that cansignificantly reduce inert mass Mi: (i) market availabilityof higher strength fibers i.e. carbon reinforced nanotubes,(ii) availability of lower density insulations, and (iii) mini-mizing internal thermal protection demands via improvedprocessing, grain regression simulation, and optimization.For example, the Vega Launcher with fiber strength in-creased by a factor of 3 would save 1600kg on the firststage, 400 on the second, and 320 on the third for a 23%increase of payload mass.

4.2. Propellant

Very high strength composite cases also enable lowerdensity propellant i.e. density impulse's significance isreduced—thereby increasing propellant alternatives. For ex-ample, replace the Vega Launcher's HTPB propellant familywith one that provides an Isv increase of 10 s and a densitydecrease (1600 instead of 1800kg/m3). For this scenario, themass penalty will be of 600kg for the P80, 205kg for theZ23, and 100kg for the Z9 (estimated with a current stateof the art design code). The effect of the Isv increase aloneis a payload increase of 440kg (35%) and with the densitydecrease 300kg (resulting mainly from the third stage's

mass penality). Moreover, with a new generation of strongfibers this mass penalty will be marginalized. This example,voluntarily pessimistic in term of density, illustrates the po-tential of Isv increases achieved with alternative propellantsthat present Isv increases and reduced density impulse.

Solid propellants with �Isv ∼50 s appear possible in mid-term (+10years). The figure below reveals dramatic perfor-mance gains from �Isv alone for the basic four stages VegaLauncher examined above.

Moreover, these significant impacts are sensitive toVega's architecture e.g. a three stage configuration witha LOX/methane upper stage replacing the third stage andAVUM would double its performance—but with significantlyincreased overall cost (Fig. 9).

5. Casing

In a solid rocket motor, the casing is devoted not onlyto combustion chamber pressure containment, but alsoto carrying general loads delivered by the motor to thelauncher. For the latter function, in addition to static loads,additional constraints often come from the launcher's dy-namic behavior where case stiffness is usually an importantparameter.

A segmented case is mandatory when the propellantgrain is too large to be cast all at once (monolithic propel-lant grain). Currently, 100–150 t grains are commonly cast,meaning internal volumes in the range of 100m3. Sinceit would be possible to manufacture larger cases if neces-sary, the size of monolithic motors is currently constrainedby propellant processing/casting technology. Therefore, asnoted by Roadmap2000 these technologies enable improvedlaunchers.

Segmentations main drawbacks relative to a monolithiccase are inert mass, cost, and risk penalties that accompanyinter-segment joints.

• When the case is metallic, these penalties are limited butnot negligible e.g. labor associated with joining segmentsis not trivial. Clevis/tang or simple bolted flange assemblydesigns are commonly employed.

• When the case is a composite material, these penalties aresignificantly larger because an intermediatemetallic framebetween the composite cylindrical parts and the joininginterface is necessary i.e. joining is accomplished with themetallic frame [4] per Fig. 10.

Solid rocket motor cases can be made from metallic orcomposite materials. However, metal cases (for the pres-sure vessel function) are increasingly replaced in modernSRM by lighter composites e.g. DELTA 2, TITAN, and H-II.Moreover, expected improvements in the specific rupturestrength of metallic materials cannot reverse this trend.Therefore, the major long term advantage of metallic casesis their reusability (provided some over-thickness is intro-duced in the design and recovery/refurbishment costs areacceptable).

The post 1990s trend toward composite cases is also sup-ported by commercial EELV ATLAS 5 and DELTA 4 launcherswith architectures based on clusters of large, composite


3.02.82.62.42.22.01.81.61.41.21.0

0 10 20 30 40 50 60Isv increase (s)

Per

form

ance

ratio

VEGA Launcher: Effect of Isv increase

1+2+3

1+2

1st Stage

Fig. 9. Multiplication factor for payload mass vs. �Isv(s).

Composite case

Factory pin

Insulation

PropellantClevis S-curve Slag barrier Primary O-ring seal

Tang

Secondary O-ring sealField joint pin

Case-to-Case Field Joint

Fig. 10. View of SRMU inter-segment joining principle.

cased, strap-on boosters (overall mass in the range of 30–40 tof propellant). In Europe, the P80 FW demonstrator wasdeveloped with the objective of preparing the way towarda new generation of large, composite cased solid rocket mo-tors. Moreover, in Japan the MuV second stage evolved froma steel case (M24) to a carbon case (M25) in 2003. In addition,an upgrade version of the SHUTTLE system with a filamentwound motor was under development in the mid-1980sbefore the Challenger accident and two motors were suc-cessfully tested. Although risk benefits from the “heritage”principle constrains Ares 1 to a five segments, metal casedRSRM motor, a composite cased Ares 5 design is officiallyretained by NASA for growth potential.

5.1. Composite vs. metallic: cost and process

From cost and process standpoints metallic hardwaredevelopment seems to be in the asymptotic stage of itsevolution while composite hardware costs are in a ear-lier stage where significant improvements remain for thefuture i.e. the overall performance-to-price ratio of car-bon fibers is still increasing. Moreover, efficient carbonfibers are now available at competitive prices and futureprice decreases are predicted. From a manufacturing pro-

cess viewpoint filament winding techniques are mature.However, several new filament processes could bring costreductions e.g. infusion of dry preform or thermoplasticsprepreg.

For either of these case design options, there are no re-alistic diameter or length limitations. In the 1960s, a 6.60msteel case was successfully manufactured for Aerojet. More-over, recent development of very large aircraft e.g. the fullbarrel airframe of Boeing's 787—demonstrate the feasibilityof large composite structures.

The Aerojet Atlas V lightweight composite case solidrocket motor produces over 1,160 kN vacuum thrust andweighs slightly over 45 t. Over 18 motors have successfullyflown on the Atlas V with 100% mission success since thefirst flight in January of 2003. Future large rocket motor de-velopment can realize even more weight saving by makinghigher performance composites more affordable.

Non-destructive control techniques to detect defects inboth metallic and composite cases are mature.

5.2. Composite vs. metallic: design and performance

A comprehensive optimization analysis was reported inRoadmap2000 [1]. For composite cases, carbon fibers for a


Table 4Optimum pressure vs. case design parameter.

Case design Rangeofpressure

(MPa)

International SRM examples

Metallic 6–7 RSRM, MPSSegmented composite 8–9 SRMUMonolithic composite 9–11 Castor 120, M25, SRB-A, P80

monolithic filament wound booster lead to a PV/Mg perfor-mance factor (pressure×volume/mass×gravity), expressed inkm, of 45, to be compared with a 30km value for Kevlarand a 20km value for glass fiber that was used to decadesago. Therefore, FW carbon structures lead to large gains instructural ratios compared to other materials—particularlysteel.

From a purely material perspective, FW carbon/epoxymaterial exhibits large advantages over steel for mono-lithic structures. However, for segmented structures “joints”(a joining region with metallic interfaces—see Fig. 10) intro-duce mass, cost, and risk penalties associated with igniter,nozzle, launcher attachments, segments, etc. Moreover, spe-cific stiffness requirements (propellant bonding, limitationin axial motor elongation or bending . . . ) can intrude. Forlarge SRM, the case mass for FW carbon/epoxy technol-ogy is roughly half that for high strength steel technology.Although this advantage is most important for upper stages,it can produce significant payload improvements for firststage and booster applications (see launch vehicle sensitiv-ity analysis, composite material strength).

Internal and external interfaces are easily managed withmetallic cases. However, for composite designs, polar bossto roving joining and skirts to roving joining, even if theyalways remain difficult points, are now mastered technol-ogies.

The increased specific strength capability of compositemotor casings leads to increased optimal operating pressurerelative to metallic casings. Typically the optimum pressurefor a metallic case is 6–7MPa. For FW carbon fiber compos-ite, this optimal value is 8–9MPa for segmented cases and9–10MPa for monolithic cases (see Table 4).

For a stage operating at ground level with a limited di-ameter exit cone, increased combustion pressure is the onlypath to significantly improved delivered specific impulse i.e.a factor of paramount importance for overall launcher per-formance. For the Ariane 5 MPS with a 3m nozzle exit di-ameter, a 3MPa pressure increase produces an Isp gain ofabout 10 s and a payload increase of nearly 10%.

Case design is increasingly constrained by axial stiff-ness requirements and/or thrust transmission through theskirt. Therefore, case design is increasingly coupled to flightperformance aspects (grain regression/insulation, pressureoscillations, vibration environment, etc).

In summary, since the extra-mass of a full-diameterinter-segment connection presents a significant penalty,it is necessary to minimize the number of connections inorder to optimize the mass reduction due to FW carbon

Motor P80 case P425 project

Size factor

Case Diameter

3 m 4.5 m 1.5

Case Length 9 m 18 m 2.0

Fig. 11. Comparison views of the P80 motor and P425 Project.

structures. Moreover, this consideration augmented withaxial stiffness requirements, flight's acceleration and vibra-tion environment, etc. complicate the determination of theoptimum combustion pressure. Therefore, a thorough anal-ysis that considers technological and practical constraints isnecessary.

5.3. General trends and potential breakthroughs

The major trend is toward light, composite materialcases with the possibility of monolithic motor designs withvery large dimensions e.g. the recent P80 development.Preliminary design of a 425 t monolithic motor was recentlyperformed in France and it revealed “show-stoppers” wereabsent from the case itself [5] (see Fig. 11). The followingtable contrasts its case size with P80 and reveals the extrap-olation is reasonable and within US large motor experience.

If casting constraints require segmentation, a compositeinter-segment joining without metallic parts is necessary tooptimize composite material benefits. Therefore, the poten-tial of in-situ wrapping of thermoplastic composite tape tocreate a composite material joint is important. Although thisjoining would not be dismountable, history demonstratesthere has been no real need for this function.

The enormous strength and large electrical conductivityof carbon nanotube filaments enables very long, low losselectric power transmission with minimal supports and lightand efficient motors and generators. Because potential eco-nomic benefits from these applications are enormous, it willdrive this technology forward if possible. Therefore, its highstrength fiber benefits will naturally fall to solid propulsionif the basic technology matures.

Nano-technologies are showing great promise in enhanc-ing the mechanical properties of structural materials—bothmetallic and composite. Nano-fillers toughen compositestructures such as rocket motor cases and other pres-sure vessels by improving fracture toughness propertiesand allow these systems to be more robust. Additionally,nano-fillers improve composite thermal conductivity overstate-of-the-art composite systems allowing better thermalcontrol in space applications.

6. Nozzles

This section treats a solid rocket motor nozzle's mainfunctions: converting the combustion chamber's chemicalenergy deposition into thrust and adroitly vectoring thisthrust for vehicle guidance and control. Since the nozzle's


Fig. 12. View of C/C nose and throat parts (SPS).

function is to convert hot, high pressure products of thepropellants' chemical reactions to directional thrust byacceleration through a converging-throat-diverging geom-etry, the nozzle's internal surfaces are subjected to a harshenvironment: hot gases and liquid/solid phase alumina.Currently, there is no material that can withstand this envi-ronment throughout a motor burn without erosion/ablationand surface recession. Since convective heating typicallyreaches a relative maximum at the throat, heat resistantmaterials e.g. carbon–carbon composite (C/C) thermo-structures, graphite, and graphite phenolic composite mate-rials are employed. For the exit cone's diverging geometry,where heating is reduced, ablative materials e.g. carbonfiber reinforced plastic (CFRP) or silica fiber reinforcedplastic are employed.

6.1. Thermo-structural composite materials

Three dimensional (3D) needled or braided and 4D rein-forcement are available for carbon/carbon throat elementsand the nozzle's “nose” when required by the nozzle designand its aero-thermal flow field. Fig. 12 presents a nozzle en-trance/throat with C/C nose and throat. C/C material densi-ties ranging from 1.65 to 1.9 g/cm3 (resistance to ablationand cost increase with density) can be selected. In the future,very high temperature metallic or ceramic coatings and car-bon nanotube reinforcement (or fillers) will be investigated.

Because recession rate differs for C/C and CFRP ma-terials, wetted surface discontinuities (steps) can form atmaterial interfaces during motor operation. The backwardstep that forms after the C/C throat disturbs the subsequentflowfield and can produce a groove downstream and ulti-mately failure there (particularly significant in high-pressuremotors). Therefore, skillful contour design and material se-lection are necessary. Computational fluid dynamic (CFD)calculations fully coupled to construction material modelsare employed to optimize contour and materials to mini-mize deleterious effects while maximizing thrust.

For a solidmotor of short duration, a ceramicmatrix com-posite (CMC) nozzle can be a candidate (see Fig. 13). CMCnozzles are manufactured from carbon and carbon-silicon

Fig. 13. View of a CMC combustion chamber and exit cone (IHI Aerospace).

Fig. 14. MAGE apogee boost motor with C–C throat and C–C nozzleextension (SPS).

fibers with a carbon-silicon matrix: a heat resistant struc-tural material with low thermal conductivity.

Carbon/carbon structures present an alternative andFig. 14 presents a C/C exit cone. The IUS ORBUS 21 and theMAGE (illustrated) motors were successful applications ofthis technology.

With current technology, large C/C exit cones are readilymanufactured, reliable, and cost competitive with classicalphenolic based designs when integrated for additional per-formance. Indeed, deployable designs can be implementedto minimize overall motor length prior to operation. Fig. 15presents C/C extensions for solid rocket motors. The nozzleextension on the left is fixed and the nozzle extension onthe right is extendable (two cones) and illustrated fully ex-tended during a hot firing test at altitude.

6.2. Thermo-ablative composite materials

The ideal ablative material retains its shape with minimalrecession post charring and sufficient thermally protection.A future material candidate is 3D-CFRP. In 3D-CFRP, carbon


Fig. 15. Example of C–C nozzle extension.

fibers also cross the cloth laminations to reinforce the trans-verse direction and enable strong, hard char formation.

Carbon phenolic insulators are typically manufactured byautomated wrapping of (ex-rayon) carbon/phenolic prepregtape and can be classified as CFRP. Although this materialperforms in highly erosive environments, it generally re-quires mechanical support to withstand mechanical loads.Post firing tests of these materials often reveal large crackscreated by cool down thermal contractions that, unfortu-nately, have complicated hardware design and materialmodels. Moreover, after charring, the char's inter-laminarstrength is very low. Therefore, ply separation and pocket-ing can complicate applications. Consequently, successfulapplications require “deep knowledge” of material process-ing techniques and its sensitivities and design technology.Clearly, part designs must adequately account for all ofthese behaviors to be successful.

The new generation of 3D reinforced materials providesmore homogeneous material and improved mechanicalproperties (particularly in the inter-ply direction) for vir-gin and charred material states. This capability eliminatesnumerous issues related to delamination and designs forlarge self-standing parts without metallic supports are nowsuccessful. Moreover, 3D reinforcement enables low costex-PAN carbon fiber material replacements of high costex-RAYON materials. RTM process are also accessible forphenolic resin injection thereby avoiding cost and technicalissues related to traditional prepreg tapes. Fig. 16 illustratesa 3D CFRP part.

6.3. Thrust vectoring

When thrust vectoring is necessary, the “universal” solu-tion is currently flex-seal and external actuators. However,alternatives e.g. Socket-Ball, Tech-Roll Joints or Liquid Injec-tion in the exit cone, are employed for special applications.

6.3.1. Flex-sealThe flex-seal concept is based on a sequential stack of

elastomeric pads and structural shims that conform to a

Fig. 16. View of the P80 flex-seal insulator (SPS).

spherical shape. This design allows ready omni-axis nozzlevectoring of ∼5◦–6◦. However, special designs can achieveomni-axis vectoring of 15◦–20◦.

Within the flex-seal concept two improvements are in-creasingly employed for space applications:

• Self protection of the flex-seal to avoid complex thermal pro-tection systems. This is accomplished by increasing shimthickness. Although this technology was originally devel-oped for defense applications that required very compactdesigns, it is now sufficiently mature for low risk use inlarge space program applications.

• Low torque, low power flex-seals employ synthetic rubbers(rather than natural rubber) to easily achieve 50% torqueand TVC power level reductions. In the future “near zerotorque” designs should further decrease TVC power re-quirements to very low levels.

6.3.2. ActuatorsThe evolution of low torque TVC is important for stage

level applications because it reduces power necessary forrequired steering angles or angular velocity thereby en-abling electro-mechanical actuators (EMA). EMA's elim-inate hydraulic power issues e.g. cleaning, leakage, and


pressurization phase lags. Moreover, EMA's sourced byLithium-ion battery power packs contribute to SRM “instantreadiness.” Furthermore, future evolutions of “near zerotorque” flex-seal designs should decrease TVC power re-quirements to levels compatible with super-capacitor en-ergy sources and further reducing TVC system cost.

7. Propellants

Solid propellants for space application are typically basedon polybutadiene binder, ammonium perchlorate oxidizer,and aluminum (Al) additive. New launchers under develop-ment (Vega, Ares I and V, GSLV MkIII, . . . ) reuse the sametechnology. These developments follow Roadmap2000s pre-dictions.

The overview of SRM boosted launch vehicles done pre-viously shows that for the next decades, SRMs for launcherswill use current propellants—with formulation adjustmentsto meet specific requirements.

7.1. General trends in solid propellants evolution

Research and development of new solid propellants areperformed for military applications to improve responsesto the stimuli of IM-tests (insensitive munitions) and toincrease performance. For space application, developmentefforts focus on reducing cost and increasing performancewhile maintaining or improving other characteristics such asmechanical and ballistic properties and safety characteris-tics. All studies and measurements performed during the re-cent decades have demonstrated the environmental impactof launchers is very small and rather negligible comparedto other anthropogenic sources [6]. However. propellant-manufacturers continue to address technologies that can re-duce adverse environmental effects.

In the USA, the integrated high-payoff rocket propulsiontechnology (IHPRPT) program was initiated in 1996 to im-prove rocket propulsion systems [7]. For solid propellantmotors the goal is to improve overall performance by 8% [7].In Roadmap2000, the recommended approach is to progressstep by step to qualify new technologies with a series ofrelevant demonstration motors [1]. The European perspec-tive was described in 2004. Among the key solid propellantstechnologies that have been mentioned are continuous mix-ing/casting, high-energy propellants and green propellants[8].

7.2. Energetic compounds

As noted in the previous section, a solid propellant nor-mally is composed of an oxidizer, a fuel and a polymericbinder. Each of these three components individually havebeen the subject of considerable research in recent yearsand there has been a veritable explosion of new compoundsavailable to the formulator. However, each has unique char-acteristics, advantages and disadvantages and the evolutionof solid propulsion will be driven by the development ofthese new molecules and their availability at industrial lev-els in the ∼20+ year future.

On the other hand, a significant proportion of solid pro-pellants (based on total tons of propellants manufactured

around the world) are formulated with a polybutadienebinder plasticized with one (or more) inert and commer-cially available plasticizers. A co-polymer consisting ofpolyethylene oxide and polytetrahydrofuran (HTPE) hasbeen developed in the USA by ATK to satisfy military IM-test requirements. A low-energy plasticizer is combinedwith it. Other polymers such as polyethyleneglycol (PEG),polycaprolactone (PCP) and polyglycidyl adipate (PGA) areused to incorporate higher percentages of high-energyplasticizers.

Energetic polymers have been developed and are cur-rently being studied. Glycidyl azide polymer has gainedcomparatively wide acceptance and is used in commer-cial applications e.g. gas generators [5]. It is producedcommercially in USA (3M) and in France (EURENCO).Other polymers, polyglycidyl nitrate (PolyGLYN), poly(3-azidométhyl-3-methyloxetane) (pAMMO), poly(3,3-bisazidomethyloxetane) (pBAMO) and poly(3-nitratomethyl-3-methyl-oxetane) (pNMMO) are currently being studied.Generally, they are plasticized with energetic products[9,10]. The newer polymer, poly(methylvinyltetrazole)(PMVT), is quoted as under development in Russia [11].

Energetic fillers involved in propellant formulationand described in the open literature are mainly AP, RDX,cyclotetramethylene tetramine (HMX), hexanitrohexaaza-isowurtzitane (HNIW), ammonium dinitramide (ADN), andhydrazinium nitroformate (HNF). Ammonium perchlorate(AP, NH4ClO4) is used in the largest number of propellantsmanufactured across the world and will be used in propel-lants for new SRM developments. AP has been and will con-tinue to be selected because it offers low sensitivity, goodthermal stability and high density and its oxygen balanceand enthalpy of formation lead to outstanding deliveredenergy. However, it contains chlorine, HCl is generated inits reaction products and its comparatively high molecularweight has a negative impact on performance. RDX, HMX,and HNIW lead to increased performance relative to AP.HNIW is of particular interest in energy terms because ofits high density. Moreover, HNF and ADN offer an excellentcompromise between enthalpy of formation and oxygenbalance. An examination of oxygen balance values showsthat RDX, HMX and HNIW are monopropellants or slightlyoxidized. In contrast, HNF and ADN are oxidizers in the truesense of the term and are chlorine-free. Unfortunately, allsensitize propellants to shock sensitivity relative to AP inthe gap test when they are used at levels above 15–20%.

7.3. Future trends for solid propellants

The pacing item in the development of a new propellantcan be considered through its maturity and availability of itsrelevant new raw materials.

7.3.1. Short termOnly raw materials that are well known, at least at labo-

ratory scale, may be considered for short-term applications.Replacement of part of the AP with an energetic material

could be the first step of development of new materials. Thiscould allow the parent propellant's ballistic properties andsensitivity characteristics to be preserved while increasing


specific impulse. As indicated above, cost is a very impor-tant parameter. A study of the comparative cost of raw ma-terials, conducted in Roadmap2000 concluded AP and RDXcosts are low, the cost of HMX, ADN, and HNF are moder-ately high, and that of HNIW is high. Taking into accountthe level of maturity, RDX is a good higher energy candidateand is the solution demonstrated by ATK in 2000 under theIHPRPT program (a demonstrator booster was successfullystatic tested [12]). The specific impulse was increased andthe oxidizer-to-fuel ratio was lowered. The latter reducesthe erosive of carbon-based material in the nozzle. More-over, CSD/SEP (now SPS) experiments with motors using a90% total solids aluminized propellant demonstrated a 40%decrease of the nozzle erosion with a formulation containing12% HMX in comparison with that formulation without HMX[13]. The replacement of a small part of the AP by RDX orHMX leaves the mechanical and ballistic properties largelyunchanged with a small burning rate reduction. Safety char-acteristics are similar. Finally, replacement of part of the APby another filler reduces the amount of hydrochloric acidproduced.

Classically, solid rocket motor grains for defense andspace applications are produced with batch processing.However, during the period 1985–1995 efforts were madein the USA and at SNPE in France to develop a continu-ous process for composite propellant production. Althoughthe USA effort was terminated, SNPE's R&D was successful[14]. Moreover, SNPE has recently proposed the construc-tion of a continuous mixing facility to manufacture Ariane5 segments—segments that will be more consistent thancurrent batch processing can produce and offer temporalformulation control [15]. This accomplishment represents amajor step forward along Roadmap2000.

7.3.2. Mid-termFor mid-term applications, chlorine-free oxidizers are in-

triguing candidates and ADN and HNF are leading entrants.Hydrazinium nitroformate (HNF, CH5O6N5) is produced byAerospace Production Products (APP) in the Netherlands.The product is friction-sensitive and its impact sensitivity isintermediate between ADN's and HNIW's. Since the crudeproduct has a needle-like shape, morphology alterationthrough crystallization or other processing is necessarybefore propellant formulation. Crystallization studies haveimproved particle morphology; however, since the particlesare not yet spherical, progress remains [16]. HNF's thermalstability is lower than its competitors'. In summary, progressat the raw material level is necessary before this productcan be a viable candidate for propellant development.

Ammonium dinitramide (ADN, NH4N(NO2)2) was firstmanufactured in the USSR during the 1970s. ADN exhibitssensitivity similar to RDX. ADN-based propellants wereused in the former USSR for strategic missile propulsionbut details have not been published [9]. Most of the resultsavailable in the open literature come from studies con-ducted by defense agencies with non-aluminized propel-lants for tactical missiles. Although the specific impulse ofaluminized propellants increases as ADN replaces AP, pro-pellant density decreases and the density-impulse gain issmall. On the other hand, high strength composite cases de-

emphasize density-impulse and could make attractive thesesolutions.

High energy propellants based on nitrate ester plasticizedpoly-ether binders and with a high nitramine content showspecific impulse increases near 10 s compared to conven-tional formulations. They have been designed for defenceapplications, are sensitive to shock, and typically receive a1.1 hazard classification rating.

A mid-term evolution of composite propellants would bereached with the use of Alane (AlH3 for instance) instead ofaluminum, with an impulse increase (but density decrease),if stable AlH3 can be produced.

7.3.3. Far termDetailed simulation now guides the development of en-

ergetic molecules, and will accelerate the advent of new can-didates. The search continues with the same objectives asbefore e.g. increase enthalpy of formation and density with-out overly increasing hazards or decreasing stability. An ap-proach, being widely investigated at the time of this writing,is to increase the enthalpy of formation by building nitrogen-rich molecules and incorporating cyclic and cage structures.Material density typically increases in parallel to these pa-rameters. This approach led in the past to HNIW. Newermolecules of interest may be, for example, furazanes, furox-anes, tetrazines, tetrazoles etc.

This trend sees its outcome with the “high energy den-sity materials” or HEDM. A definition could be: “Materialsreleasing high energy per unit mass when decomposing orreacting”. The poly-nitrogen compounds are good potentialcandidates because of the high energy in themolecular nitro-gen triple bond and weak energies of the double and singlebond. Some of themore exotic candidates are still theoreticali.e. un-synthesized anywhere. In these cases, an estimateddensity and enthalpy of formation are calculated. Some ofthe calculated values are far greater than for the moleculesmentioned before and would lead to very high specific im-pulses. For example, octaazacubane (N8) with a predictedenthalpy of formation of 2200kJ/mole corresponds to a spe-cific impulse of 529 s [17].

The most immediate propellant approach is to envisionthese molecules as comparatively low percentage dopingagents. For example, a composition with 10 HTPB/30 AP/60N8 shows a predicted specific impulse (standard conditions)of 353 s, a volumetric specific impulse of 704 s g cm−3 and apredicted combustion temperature near 5000K. Significantefforts will need to be devoted to this synthesis. On thatpath, successes are reported for salts of N+

5 [18]. The use ofthese products will certainly require new materials for ther-mal protection such as insulation and liner and especiallyin the nozzle. This concept illustrates the great potential interm of performance for these exotic candidate materialsand they offer a long term breakthrough in solid propellantsperformance.

8. Controllable solid propulsion (CSP)

Controllable solid propulsion represents a breakthroughfor SRM use since it transforms the predetermined thrustlaw in a flexible one.


Solid Propellant

Pintle Nozzle Assembly

Actuator Assembly

Fig. 17. Aerojet's CSP rocket landing motors enhance astronaut and payloadsafety.

Aerojet is evaluating applying this technology to criticallanding rocket motors for both the crew return capsule androbotic landers.

Controllable solid propulsion technology is applicable forprecision landing rockets to decelerate and stabilize the crewcapsule or lander just prior to touchdown. An improvementover existing technology and other concepts is desirablefrom the perspective of astronaut and payload safety. Aero-jet's pintle-controlled motors with three-axis vehicle con-trol could provide enhanced safety and packaging features.Fig. 17 is one CAD concept of a CSP landing rocket that couldbe packaged around the perimeter of a crew capsule to pro-vide variable thrust controllability.

For crewed missions, the atmospheric conditions inwhich a crew capsule lands on Earth have a certain degreeof unpredictability. Factors, such as high winds and stronggusts, can introduce rapid and unusual vehicle responsesthat endanger the crew during landing. CSP technology is amature, low-risk approach to provide precision control tothe crew capsule over a large envelop of operating scenarios.

9. Reliability and simulation

Solid rocket motor quality assurance is based on de-tailed examinations of specific units with non-destructive in-spection (NDI) technologies (X-ray and ultrasound), qualityassurancemotor tests, and total quality control of motor pro-cessing. Therefore, because all motors cannot be inspectedindividually without prohibitive cost, and propellant pro-duction/processing aspects of motor manufacture have notbeen amendable to local inspection at adequately small spa-tial and temporal scales, overall process control has beenemployed. Consequently, the final product's quality is guar-anteed by process and trend control, that is, the quality isindirectly guaranteed with the complement of direct prod-uct inspections.

A way for improving reliability is to understand allphysical phenomena that occur in the manufacturing pro-cess and SRM operation. Consequently, there will be greatand growing demand for improved propellant and materialcharacterization (input information), motor testing (systemperformance information), and the development of nu-merical simulation technology that can connect adequatematerial characterization information with product perfor-mance. It is this synthesis operating in direct and inverse

modes and synergistically integrated with necessary diag-nostic developments, education, and diligence that leadsinescapably to superior product reliability.

SRM numerical simulation research has addressed a va-riety of aspects:

• SRM internal ballistics evaluation with burn-back simula-tion [19,20] integrated with casting process effects [21,22].

• Simulation of random packings [23] and deflagration wavepropagation through heterogeneous solid propellants[24–28] containing fine/ultrafine aluminum fuel [29,30].

• Multi-disperse, multi-phase flow simulations that includealuminum/alumina droplets [31–33], aluminum agglom-eration [31,34–36], and the slag mass accumulation [37].

• Simulation of vortex-shedding [38] and thrust oscillation[39] with the view point of the adaptive control [40], ofthe effect of burning aluminum droplets [41], of the noz-zle cavity effect [42], of the wall and the inhibitor effect[43,44], and of the large solid rocket boosters [45–49].

• Simulation of the internal flow with respect to the nozzleablation [50–52] and to the roll-torque generation [53].

• Simulation of combustion stability [54].• Assessment of the acoustic, vibration, and shock environ-

ments of an SRM firing [55], the assessment of the atten-uation of the radio frequency signals transiting the SRMplume [56,57], and so on.

In order to improve the reliability of SRM simulations,it is of critical importance to validate the accuracy of nu-merical simulations and model refinements of each physicalphenomenon with adequate (i) propellant characterization(flowfield and condensed phase boundary conditions) and(ii) motor test results. An excellent example of this processis research on thrust oscillations observed during the sec-ond half of the P230 motor's (Ariane-5's booster) burn. Thisresearch employed detailed numerical simulations to clar-ify the role of vortex shedding from inhibitors, the propel-lant grain's edges, and the burning surface, on the acousticpressure [38,41,42,45,47,48] in concert with cold flow ex-periments with adequate diagnostics, andwell instrumentedscale motor tests.

Although propellant characterization [58–60] and inter-nal ballistics theory are currently based on homogenizedpropellant boundary conditions, heterogeneous propellantsand their innately stochastic, poly-disperse chemically dis-crete morphology dominate applications [61]. Therefore,theory's smooth burning surface topography, determinis-tic, and spatially uniform injection boundary conditions(irrotational for isobaric flow in quiescent environments) arerobust for neither flowfield nor condensed phase becauseheterogeneity information has been purged (see Price's [62]seminal criticism and Fig. 18). Moreover, Massa, Jackson,and Buckmaster [63] prove heterogeneity information re-moved from boundary conditions must be appropriatelyrestored to the governing equations if results are to be robustfor the heterogeneous propellant. In addition, George andDavidson's [64] demonstration that asymptotic turbulentflows are sensitive to their source's space, time character-istics (large eddy structures appear to propagate this infor-mation through the flowfield) implies Fig. 18's phenomena


Fig. 18. Full-field instantaneous temperature contours for AP/HTPB propellant (left) and slices from the surface out to 1.5mm (right) show jet-like structurespersist far downstream of the combustion zone (Courtesty Dr. T.L. Jackson, CSAR/UIUC).

can alter its deflagration sourced flowfield's large eddy andturbulence structures. Furthermore, flowfield sensitivity toheterogeneity is supported by the empirical minimizationof RS Maverick's omni-present pressure oscillation for iso-burning rate, composition, and grain/motor geometry con-straints by adroit heterogeneity change [65]. Finally, Glickand Hessler [66] prove classical acoustic stability theory isnot robust for heterogeneous propellants.

Another example is lessons learned from the failure ofthe nozzle-liner of the H-IIA launch vehicle's SRB-A motorfrom localized ablation [67]. Of course, in this case, propermaterial selection for ablative parts and appropriate designof the nozzle contour are crucial. Detailed numerical simu-lations of the three-dimensional internal flow was utilizedin the return-to-flight activities of SRB-A.

An example of current and future interests is roll-torquegeneration due to SRM operation. It is an old problem[68,69], but evaluation of roll torque due to firing is nota simple task [53]. When a booster is employed as thefirst-stage motor (the USA's ARES-1 and Japan's next solidrocket) and a new launch system is made from new SRMs(the European VEGA LV), roll torque evaluation is necessaryin the design phase. Unfortunately, this is difficult theoreti-cally: recent theoretical results by Buckmaster, et al. implyinnate flowfield asymmetries (casting induced local burningrate variations, etc.) can create roll torques, and Shimadahas shown that swirling flows can be created in SRM undercertain conditions.

SRM physical phenomena are complex. Therefore, de-tailed numerical simulation developments have proceeded[46,63,70–72] and become another significant trend in SRMtechnical development. In order to integrate simulationsof different disciplines associated with SRM, the computerscience developments have been pursued that enables si-multaneous treatment of necessary phenomena and theirindividual spatio-temporal characteristic scales. Significantimprovements of this technology are expected and crucialto solid rocketry's future.

Consider a possible example of the future's multi-disciplinary simulation. First, simulation of the propellant

slurry cast into the motor chamber is coupled with sim-ulation of its random packing and its flow's rheology.Therefore, differences of local packing characteristic dueto the slurry flow parameters (stress, velocity) can be pre-dicted. Moreover, because small samples can define theirpacking's structure [73] and deflagration phenomena, thesimulation's material predictions can be validated and re-lated to local deflagration phenomena experimentally andtheoretically. Consequently, propellant formulation andits processing into a loaded motor can be related to thatgrain's local regression rate characteristics and boundaryconditions for condensed phase and flowfield (see Fig. 18).Ergo, fully coupled integration of the equations governingcondensed phase and flowfield can now define this motor'sdetailed behavior in space and time from ignition to burnupfrom formulation, processing, motor geometry, and flightenvironment information. This enables detailed validationof these predictions with adequate test data. Furthermore,inverse calculations from the motor test data can addressinternal process details—and subsequent cost effective, de-tailed characterizations at small scale, can validate bothdirect (from the detailed deflagration model) and inversetheoretical characteristics. Finally, detailed simulations canmentor the development of new diagnostics (and computa-tionally fast models) and optimize their parameters.

The above reveals synergistic interactions among de-tailed simulation, propellant characterization, andmotor testtechnologies are critical. Therefore, because ballistic char-acterization and motor testing are currently based onhomogenized propellant while heterogeneous propellantsdominate applications, improved ballistic characterization(see Fig. 18) and motor testing are necessary.

It is clear detailed simulation with adequate experi-mental data enables a “bootstrapping” approach to “deepunderstanding” of solid rocket processes and operationalimplementation in motor design/development. Indeed, thispath is similar to the one that enabled homogeneous pro-pellant theory, empirical data, and skilled, experienced, anddedicated personnel to develop modern solid propulsion.Therefore, the importance of detailed simulation and its


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Fig. 19. Time evolution of the energy.

experimental propellant/motor characterization compan-ions, its universal significance, and its connection to educa-tion, suggests their development as a path for internationalcollaboration and restoration of strong Academy/SRM prac-tice interactions.

10. The future

Solid propulsion continues to offer reliable, low cost,high thrust propulsion for booster applications to all launch-ers, upper stages of small launchers, and niche applications.Moreover, Roadmap2000s main conclusions are still validand remarkably prescient e.g. History, Roadmap2000: reca-pitulation. Its identification of efficient production as crucialto further cost reductions and reliability and performanceimprovements, and continuous propellant mixing/casting,composite cases as key ingredient has been enhanced in thelast 10 years.

Although solid rocketry currently rests on these basis,skilled and experienced (hard won during real motor de-velopments [74–76]) personnel have mid-wifed successful“births” of grained motors. Therefore, the shortage of newmotor development programs [77], the increasing retire-ment rate (exacerbated by “baby boomer” demographics) ofskilled and experienced personnel, presents potentially dele-terious shortages of these critically important personnel inthe future. The “building block” strategy is interesting froman economical point of view, but restraint the training ofsuch generations.

Fig. 19 is a plot of the evolution of the energy mas-tered along the humanity. It clearly demonstrates that thelife-time of an innovation shortens and shortens. It tookmillenniums to go from stone, bow and arrow to chemi-cal energy. It took centuries to go from black powders tocomposite propellants. It took decades to go from RDX toCL20. Each breakthrough or innovation offers a jump in per-formance, becomes the state-of-art, and is little improveduntil the next innovation supplants it. In the field of ener-getic materials, solid propulsion is still having long cyclesof use of existing technologies, whether because the costconstraints on space launchers developments condemnsto building block initiative, or because existing technologystay the best. As a consequence, solid propellants used forspace applications are today stuck to composite propellants,which represent the best compromise in performance, sen-sitivity and cost. On the contrary, the current growth ofknowledge and the acceleration of computing initiativesin chemistry could lead in the mid-term to discovery ofnew ingredients that will revolutionize solid propulsion(e.g. HEDM).

Breakthrough technologies could also be explored forspecific missions (e.g. Moon or Mars base) or space launch-ers with cryogenic solid propellants, ALICE or refrigeratedsolid propellants [1,78,79]. These breakthroughs enable theuse of liquid ingredients in solid propellant formulation, re-quiring an operation temperature below the melting point,and enlarge the variety of ingredients that can be used forreaching high Isp.


Solid propulsion's immediate future is secured by its cur-rent capabilities, future potentials, and the ability of person-nel to deliver them at low cost and risk. Moreover, adroitapplication of this technology provides a solid basis for nearfuture developments.

Acknowledgments

This paper was initiated by Marcel Pouliquen on behalf ofIAA Advanced Propulsion Working Group. An internationalteam was set up, and got from the beginning the supportof the main institutions and companies working on solidpropulsion in Europe, Japan and in the USA.

The group was leaded by Jean-Francois Guery, SNPEMateriaux Energetiques (Europe and overall coordination),I-Shih Chang, Aerospace Corporation (USA coordination)and Toru Shimada, JAXA, (Japan coordination).

The following people joined the working group for coor-dination meetings and/or contributed to the paper:

• Bruno d'Andrea, Avio Spa, Italy• Bernard Broquere SNECMA Propulsion Solide, France• Didier Boury, SNECMA Propulsion Solide, France• Max Calabro, Consultant, France• I-Shih Chang, The Aerospace Corporation, USA• Marilyn Glick, Consultant, USA• Robert Glick, Consultant, USA• Jean-Francois Guery, SNPEMateriaux Energetiques, France• Mark Kaufman, Aerojet , USA• Hiroto Habu, JAXA, Japan• Francesca Lillo, Avio Spa, Italy• John Napior, Aerojet, USA• Christian Pérut, SNPE Materiaux Energetiques, France• Eric Robert, CNES, France• Nobuhiro Sekino, IHI Aerospace, Japan• Toru Shimada, JAXA, Japan• Jean Thepenier, SNPE Materiaux Energetiques, France• Robert Wardle, ATK, USA• Gilles Vigier, Astrium ST, France

Three meetings took place: in Cincinnati, USA, on July 11,2007 (kick-off meeting during the AIAA JPC), in Hyderabad,India, on September 26, 2007 (progress meeting during theIAC), in Hartford on July 23, 2008 (progress meeting duringthe AIAA JPC). Three other meetings took place among thecoordinators.

A special thank to Bob and Marilyn Glick for rewriting inpure English the international contributions.

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Solid Propulsion for Space Application an Updated Road Map

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