CASE FILE SPACE SHUTTLE AUXILIARY … 7 3 - 19827 \ msc 04423-1 mdc e0674 case file copy final report space shuttle auxiliary propulsion system design study executive summary /vfcdo/v/vtl_l-

$Page 1: CASE FILE SPACE SHUTTLE AUXILIARY … 7 3 - 19827 \ msc 04423-1 mdc e0674 case file copy final report space shuttle auxiliary propulsion system design study executive summary /vfcdo/v/vtl_l-$
N 7 3 - 1 9 8 2 7 \MSC 04423-1

MDC E0674

CASE FILECOPY FINAL REPORT

SPACE SHUTTLE

AUXILIARY PROPULSION SYSTEM

DESIGN STUDY

EXECUTIVE SUMMARY

/VfCDO/V/Vtl_l- DOUGLAS ASTRONAUTICS CO/VTFVt/VV -EAST

/VTCOO/V/VHI.L DOUGLAS

COHPOftATIOM

https://ntrs.nasa.gov/search.jsp?R=19730011100 2018-06-19T02:52:19+00:00Z

MSC (M23-1

COPY NO.

SPACE SHUTTLE

AUXILIARY PROPULSION SYSTEM

DESIGN STUDY

EXECUTIVE SUMMARY

29 DECEMBER 1972 REPORT MDC E0674

FINAL REPORT

Prepared by: PJ. KellyT.F. Schweickert

Approved by: L.F. KohrsManager, Propulsion

Contract No. NAS 9-12013

MCDONNELL. DOUGLAS ASTRONAUTICS COMPANY - EAST

Saint Louis. Missouri63166 (314) 232-0232

MCDONNELL DOUGLAS

CORPORATION

APS STUDY MDC E0674Executive Summary 29 December 1972

ABSTRACT

This report summarizes the effort conducted in support of the SpaceShuttle Auxiliary Propulsion System Design Study and defines the pre-liminary design of the final selected subsystems. The study was per-formed for the National Aeronautics and Space Administration underContract NAS 9-12013.

/VffCDOJVMELL DOUGLAS ASTRONAUTICS COMPANY ' EAST

APS STUDY MDC E0674

Executive Summary 29 December 1972

TABLE OF CONTENTS

Section Page

1. Introduction .. 1-1

2. Study Approach 2-1

3. Oxygen/Hydrogen Studies 3-1

3.1 Requirements Definition - Phase A 3-13.2 Candidate RCS Comparisons - Phase B . 3-13.3 RCS/OMS Integration Study Phase C 3-53.4 Special RCS Studies - Phase D ...3-5

4. Earth Storable Propellent Studies 4-14.1 Requirements 4-24.2 Concepts Description 4-24.3 Design Definition 4-74.4 Concepts Comparison 4-174.5 Reusability Effects 4-22

5. Conclusions 5-1

6. References 6-1

EFFECTIVE PAGES

Title Page

ii and iii

1-1 and 1-2

2-1 through 2-4

3-1 through 3-9

4-1 through 4-25

5-1 through 5-3

6-1

MCDONNELL. DOUGLAS ASTRONAUTICS COMPANY • EAST

APS STUDYExecutive Summary

MDC E067429 December 1972

1. INTRODUCTION

To provide the technology base necessaryfor design of the Space Shuttle, NASA hassponsored a number of technology programsrelated to Auxiliary Propulsion Systems(APS). Among such programs has been aseries of design studies intended to providethe system design data necessary for selectionof preferred system concepts, and to delineaterequirements for complementing componentdesign and test programs. The first of thesesystem study programs considered a broadspectrum of system concepts but, because ofhigh vehicle impulse requirements coupledwith safety, reuse, and logistics considera-tions, only cryogenic oxygen and hydrogenwere considered as a propellant combination.Additionally, unknowns in thruster pulsemode ignition and concern over thedistribution of cryogenic liquids served toeliminate liquid-liquid feed systems from thelist of candidate concepts. Therefore, onlysystems which delivered propellants to thethrusters in a gaseous state were consideredfor the Reaction Control System (RCS). Theresults of these initial studies, reported inReferences A through D, indicated thatamong the many options for design of agaseous oxygen/hydrogen system, an ap-proach using heat exchangers to thermallycondition the propellants and turbopumps toprovide system operating pressure would bestsatisfy requirements for a fully reusableSpace Shuttle. These initial studies focusedattention on this general system type but didnot examine in depth several viableapproaches for turbopump system design andcontrol. To fill this need, NASA contractedwith McDonnell Douglas Astronautics Com-pany-East (MDAC-E) in July 1971 for addi-tional study of Space Shuttle AuxiliaryPropulsion Systems. This contract (NAS9-12013), titled "Space Shuttle AuxiliaryPropulsion System Design Study," was underthe technical direction of Mr. Darrell Ken-drick, Propulsion and Power Division,Manned Spacecraft Center, Houston, Texas.

The objective of this study was to develop

design and programmatic data for candidateSpace Shuttle RCS in sufficient detail that avalid selection could later be made betweenthe various concepts. As originally defined,the study considered only oxygen and hydro-gen propellants. The program was dividedinto the five phases listed below:

1. Phase A - Requirements Definition2. Phase B - Candidate RCS Concept

Comparisons3. Phase C - RCS/OMS Integration4. Phase D - Special RCS Studies5. Phase E - System Dynamic Performance

AnalysisPhase A defined all design and operating

requirements for the APS. The result of thisphase showed that requirements for thebooster and orbiter stages were sufficientlysimilar to allow concentration of all design ef-fort on the orbiter stage, with the resultsbeing applicable to the fly-back-type boosterstage as well. These results are documented inReference E. In Phase B, very detailed designand control analysis for the three mostattractive oxygen/hydrogen RCS conceptswere conducted. Reference F documents thePhase B results. Phase C evaluated thepotential for integration of the RCS with theOrbit Maneuvering System (OMS). ReferenceG provides documentation of the Phase Coxygen/hydrogen efforts. In Phase D twoalternates to gaseous oxygen/hydrogenturbopump RCS were investigated. The re-sults of this phase are documented in Refer-ence H.

According to the initial program plan,Phase E was to analyze the oxygen/hydrogenRCS dynamic response characteristics. How-ever, concurrent with this study, vehiclestudies showed that smaller Shuttle orbiterswith external, expendable main engine tank-age would provide a more cost effectiveapproach than the larger vehicle used togenerate baseline requirements in Phase A.This change in vehicle design resulted in asignificant reduction in the APS impulse re-quirements. This effect, together with a

1-1

DOUGLAS ASTRONAUTICS COMPANY • E AST


companion Shuttle program decision to allowscheduled system refurbishment, allowedconsideration of earth storable propellantsystems for auxiliary propulsion. Thus, inNovember 1971, NASA issued a contractchange order to consider monopropellant andbipropellant storable systems. This resultedin a redirection of Phases C and E of thestudy. Phase C was expanded to include eval-uation of candidate storable propellant RCS,with varying degrees of integration with theOMS and the Auxiliary Power Unit (APU).In the redirected Phase E, performance of themost promising storable RCS/OMS/APUoptions was analyzed. Reference I documentsthe results of the storable propellant Phase Cand E efforts.

Results from the oxygen/hydrogen portionof the study indicated that, for the largerorbiter with internal main tanks, maximumpayload weight is obtained with a liquid-liquid oxygen/hydrogen RCS installedintegrally within the vehicle. For the smaller


orbiter with external main tanks, installationof the oxygen/hydrogen RCS is practical, butsufficient volume is not available for anoxygen/hydrogen OMS. Thus, the OMS mustutilize higher density storable propellants.The high costs associated with the develop-ment of oxygen/hydrogen technology werenot considered justifiable without applicationto both the RCS and OMS. Therefore, the ox-ygen/hydrogen propellants were eliminatedfrom further consideration after the orbiterdesign change. Consequently, althoughoriginally intended to study oxygen/hydro-gen systems, the main portion of the study ef-fort concentrated on earth storable propel-lants.

This report summarizes the tasks con-ducted in Phases A through E of theAuxiliary Propulsion System Design Studyand defines in summary form the final config-urations for the alternate propellant config-urations, with primary emphasis on earthstorable propellants.

1-2



NIDC E067429 December 1972

2. STUDY APPROACH

The Space Shuttle Auxiliary PropulsionSystem Design Study was conducted in fivephases. Reference J provides a detailedprogram plan for the complete study, and de-fines the task objectives and their relation-ship to the overall study. The tasks for theoxygen/hydrogen and earth storable propel-lants studies are shown in Figure 2-1 in flowchart form. Figure 2-2 defines the programschedule.

Phases A, B, C (RCS/OMS Integration),and D considered oxygen and hydrogenpropellants. In Phase A, vehicle requirementswere defined. Specifically, engine thrustlevels, locations, and number of engines wereestablished. Thruster minimum impulse-bitwas determined, based on historic engine dataand, in conjunction with deadband require-ments, used to establish total impulse forlimit cycle operation. Combined withmaneuver requirements, this information wasused to establish total impulse requirementsfor the oxygen/hydrogen studies.

Phase B was a continuation of the initialoxygen/hydrogen studies which had identi-fied the following three high value concepts:(1) a series-upstream turbine concept whichused the combustion products from a singlegas generator to first power the turbopumpand then thermally condition the propellants,(2) a series-downstream turbine concept inwhich the order of gas generator exhaust flowthrough the heat exchanger and turbine wasreversed from that above and, (3) a parallelRCS concept which employed separate gasgenerators to power the turbopump and heatexchanger. In Phase B, detailed analyses wereperformed to define preferred controls andoptimum system design points for the threeRCS concepts. Additionally, steady-state andtransient operational characteristics and theeffects of system malfunctions were evalu-ated. Finally, the concepts were compared onthe basis of pertinent selection criteria.

The Phase C RCS/OMS integration effortevaluated the integration potential betweenthe oxygen/hydrogen RCS and the Orbit

Maneuvering System (OMS). Integration op-tions ranged from a fully integrated system toa separate system in which only propellantstorage was common. Preferred methods wereselected and design points were developed fortwo fully integrated systems, one partiallyintegrated system, and one separate system.

Phase D explored the potential of the twoalternate, oxygen/hydrogen RCS conceptslisted below:

(1) Gaseous oxygen/hydrogen systems,with conditioners similar to those of thePhase B candidates, but using alternatemeans of providing system pressure,e.g., electric or hydraulic motor drivenpumps or pneumatic bellows pumps.

(2) Liquid oxygen/hydrogen systems,which eliminated conditioning equip-ment entirely and delivered the propel-lants to the engines in a liquid ratherthan a gaseous state.

For these two system concepts, the opera-tional characteristics of electric and hydraulicmotor driven pumps were evaluated to definesystem performance. Additionally, a detailedthermal analysis of the liquid system wasperformed to determine the feasibility of thisconcept.

In the Phase C earth storable propellantsstudy, effort focused on providing sufficientcomparative data on alternate storable pro-pellant concepts to allow selection of the bestapproaches for the Phase E System Perform-ance Analysis. RCS/OMS/APU integrationoptions were evaluated to determine theproper compromise between performance andoperating requirements. Both monopropellant(hydrazine) and bipropellant (nitrogen tetrox-ide/monomethylhydrazine) concepts wereconsidered. Preliminary baseline designs, re-flecting various levels of system integration,served as reference points for detailed designand installation studies and for concurrentstudies of APU implementation and advancedpressurization and tankage concepts. Fromthese studies, design concepts were updatedand six systems were selected for the Phase E

2-1

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APS STUDY MDC E0674


final performance analysis. In this phase, Propulsion Design Study. The following sec-system designs and performance were refined, tions provide a summary of the alternateand system reuse, maintenance, safety, and oxygen/hydrogen and earth storable systems,operational criteria were established. Also presented is a description of the designs

This report summarizes the tasks conduct- selected, their operation, and the rationaleed in Phases A through E of the Auxiliary behind their selection.

2-4

MCDONNELL DOUGLAS ASTRONAUTICS COMfAtW - EAST



3. OXYGEN/HYDROGEN STUDIES

The oxygen/hydrogen effort performed inthis study was divided into three areas: (1) afurther evaluation of turbopump RCS config-urations initially studied under Contract NAS8-26248, (2) an assessment of the integrationpotential of the RCS and QMS, and (3) aninvestigation of RCS concepts utilizing powersources other than turbopumps. The sectionsthat follow provide a summary of the resultsof these investigations.

3.1 Requirements Definition - Phase A -RCS thrust level and total impulse were es-tablished for the vehicle configuration, designcharacteristics, and acceleration requirementsdefined in the Space Shuttle Vehicle Descrip-tion and Requirements Document(SSVDRD).The distinguishing feature of this orbiter wasthat the main engine propellant tanks wereinternal to the vehicle, resulting in a largeorbiter stage.

A detailed study was conducted to defineAPS thrust level, total impulse and number ofthrusters. Options available within the con-straints imposed by the vehicle accelerationrequirements and vehicle configuration werecompared to establish the installation andthruster characteristics which would provideminimum RCS weight. A thrust level of 1150Ibf was selected for the RCS thrusters. At thislevel, 33 RCS thrusters were required. It wasalso required that the system be capable ofsustaining the firing of five thrusters (5750 Ibfequivalent thrust). This corresponds to theuse of four thrusters for translation and theequivalent of one additional thruster for ve-hicle attitude control during the maneuver.The total RCS impulse, including bothattitude control and vernier translation man-euvers of less than 20 ft/sec was 2.23 millionIbf-sec. Axial translation maneuvers in excessof 20 ft/sec were allocated to the OrbitManeuvering System (OMS). Impulse re-quirements for three typical missions were de-fined. Figure 3-1 summarizes the RCS andOMS design requirements.

Several general requirements which apply

to both RCS and OMS design include minimalmaintenance with ease of component removaland replacement, a minimum service life of100 mission cycles over a ten year period withcost effective refurbishment, and seven daysof self-sustaining life for each mission.Additionally, the RCS and OMS must pro-vide control for crew safety after the failure ofany two critical components, except in theevent of an abort caused by a main enginefailure. In this case, the OMS must only oper-ate after a single failure since the main engineloss constitutes the first system failure.

RCS NUMBER OF THRUSTERSTHRUSTER THRUST (LB)

NUMBER OF CONDITIONERS

SYSTEM THRUST (LB)

TOTAL IMPULSE (LB-SEC)

RESUPPLY

EASTERLY LAUNCHSOUTH POLAR

OMS NUMBER OF ENGINESENGINE THRUST (LB)

SYSTEM THRUST (LB)

TOTAL IMPULSE (LB-SEC)

RESUPPLY

EASTERLY LAUNCH

SOUTH POLAR

ORBITER

331,150

35,750

2.23 x 106

2.23 x 106

2.15 x 106

DESIGNED FOR

ORBITALOPERATIONS

10.34 x 106

3.72 x 106

12.87 x 106

BOOSTER

241,150

49,200

500,000

DESIGNED FORABORT

3

12,00024,000

FIGURE 3-1 RCS/OMS DESIGNREQUIREMENTS

3.2 Candidate RCS Comparisons - Phase BThree alternate gaseous oxygen/hydrogenRCS configurations were considered in thisstudy phase. These differ in the arrangementof the gas,'generators, turbopumps, and pro-pellant conditioners, as shown in Figure 3-2.The first concept places the turbine upstreamfrom the conditioner in the gas generator ex-haust flow. The second concept reverses theorder of the turbine and conditioner. The thirdconcept uses two parallel gas generators, onefeeding the turbine and one feeding the con-ditioner.

3-1

/tfCDOJV/VE'f-f. DOUGLAS ASTRONAUTICS COMfANY - EAST



SYSTEMWEIGHT, LBN1TOTAL IMPULSE, LBF-SECSPECIFIC IMPULSE, SECMIXTURE RATIO

THRUSTERTHRUST LEVEL, LBFSPECIFIC IMPULSE, SECMIXTURE RATIOCHAMBER PRESSURE, LBF 'IN2A

GAS GENERATOR

COMBUSTION TEMPERATURE, °RCHAMBER PRESSURE, LBF 'IN2AMIXTURE RATIO

HEAT EXCHANGERHOT SIDE EXIT TEMPERATURE, °RCOLD SIDE EXIT TEMPERATURE, °R

TURBOPUMPFLOW RATE, LBM/SECDISCHARGE PRESSURE, LBF 1N2ATURBINE PRESSURE RATIO

ACCUMULATORSNUMBER OF CYCLESVOLUME, FT3

MINIMUM PRESSURE, LBF 'IN2ASWITCH PRESSURE, LBF/IN2AMAXIMUM PRESSURE, LBF/IN2A

SERIES TURBINEUPSTREAM

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FIGURE 3-2 RCS DESIGN SUMMARY

As an example of system operation, con-sider the turbine upstream concept, shownschematically in Figure 3-3. During systemoperation, propellants from the cryogenicstorage tanks are pumped to high pressureusing turbopumps, thermally conditioned tosuperheated vapors in heat exchangers, andthen stored in accumulators. Gaseous propel-lants are supplied to the film cooled thrusterassemblies through pressure regulators. Theenergy for propellant pressure and tempera-ture conditioning is supplied by combustion

products from bipropellant gas generators.Oxygen and hydrogen propellants are sup-plied to the gas generators from the accum-ulators. The accumulators operate in ablowdown mode with accumulator pressuredecaying from a maximum value to aswitching value. At the switching value,accumulator resupply is initiated. The gasgenerators are ignited, providing energy topower the turbopump and heat exchanger.Accumulator pressure continues to decay to aminimum value during the conditioner

3-2

/VffCOO/V/VEt-f- DOUGLAS ASTRONAUTICS COMfANY - EAST


assembly start transient, and then begins toincrease as steady state resupply flowrate isachieved. Resupply flow is maintained by theconditioner assembly until accumulatorpressure rebuilds to its maximum value,whereupon propellant flow to the gas gen-erators is terminated. The accumulator

LH


blowdown/recharge cycle is repeated asmany times as necessary to satisfy missiontotal impulse requirements. Although theaccumulators operate over a wide pressurerange, propellant flow is regulated to constantsupply pressure downstream of the accumu-lators, maintaining a constant inlet pressureto the thrusters and gas generators.

VENT

GO,

GASGENERATOR

FIGURE 3-3 BASIC RCS CONCEPT

Design points for the three systems are de-fined in Figure 3-2. As shown, the two seriesconcepts are nearly identical in systemperformance. A power balance of these sys-tems requires high hot side heat exchangerflow rate, and at this flow rate, pump powerrequirements are satisfied with low turbinepressure ratios. As a result, vent pressuresare relatively high in the two series concepts.However, in the parallel RCS, low turbineflow rates and corresponding high turbinepressure ratios are required to efficiently uti-lize the available thermal energy from the gasgenerator combustion products. Even withthe high turbine pressure ratios, the enthalpyof the exhaust gas is high for the parallelRCS, resulting in a lower system specific im-pulse.

All three RCS concepts can be controlledwithin tight operational limits. Figure 3-4summarizes the optional control concepts forthe three candidate configurations. Gas gen-erator combustion temperature control wasnecessary to avoid excessive turbine/heat

exchanger gas inlet temperatures, and wasbest achieved through modulation of the gasgenerator oxygen valve. This controlprovided a large system weight reduction andreduced the operating bands of other criticalsystem parameters, such as conditioned pro-pellant temperature and pump dischargepressure (flow rate). Control of hydrogen con-ditioned temperature in the two seriesconcepts and both hydrogen and oxygenconditioned temperature in the parallel RCSwas selected to provide additional systemweight reductions. This control was bestachieved by modulating the amount of heatexchanger cold side bypass flow. Thisbypass was required in the series-upstreamturbine RCS (hydrogen conditioner) and par-allel RCS (both hydrogen and oxygen condi-tioners) to preclude r^O condensation andicing on the hot side tube walls. It was in-corporated in the series-downstream turbineRCS (hydrogen conditioner) for the purposeof conditioned temperature control. The final

3-3

MCDONNELL. DOUGLAS ASTRONAUTICS COM fANY • EAST


system control provided for modulation of thegas generator H2 valve in response to pumpdischarge pressure. This control providedonly a modest system weight benefit, but itprovided excellent control of heat exchangercold side inlet conditions and turbopump


power, minimizing the development risk as-sociated with these assemblies. Heat ex-changer cold side flow instability has beenencountered in previous development pro-grams, and its potential for occurrence is re-duced with tight control of inlet conditions.

Series RCS(Turbine Upstream)

Series RCS(Turbine Downstream)

Parallel RCS

ACTIVE CONTROL POINT

1. GAS GENERATOROXYGEN VALVE

2. GAS GENERATORHYDROGEN VALVE

3. HEAT EXCHANGER COLDSIDE BYPASS VALVE(H2 ONLY FORSERIES CONCEPTS)

INTERMEDIATE CONTROL FUNCTIONS CONTROLLED PARAMETER

( GAS GENERATOR"V COMBUSTION TEMPERATURE

TURBINE INLET TEMPERATURE,FLOW RATE, PRESSURE RATIOAND POWER

HEAT EXCHANGER COLDSIDE FLOW RATE

PUMP DISCHARGEPRESSURE

CONDITIONEDPROPELLANTTEMPERATURE

FIGURE 3-4 CONTROL POINT SUMMARY

The three RCS concepts were compared onthe basis of weight, complexity, flexibility,reliability, and technology considerations.The two series concepts rated even, rankinghigher overall than the parallel RCS. A choicebetween the two series concepts depends onthe Shuttle development philosophy. Theupstream turbine RCS affords the lowestweight and volume, but the system require-ments must be firmly established at the pro-gram outset. Attempts to improve the system

performance at a later date for a second gen-eration vehicle would lead to a complete con-ditioner redesign due to the turbine inlet tem-perature restrictions. The series-downstreamturbine RCS can be uprated by increasing gasgenerator combustion temperature andreconfiguring only the heat exchanger for thecorrespondingly higher hot side inlet temper-ature.

Several technology concerns were identifiedduring the course of the study. Among these

3-4

MCDO/VtVELL. DOUGLAS ASTRONAUTICS COMFANY - EAST


was the high conditioner cycle life require-ment of 5000 cycles (50 cycles per mission for100 missions) which is a significant extensionover the demonstrated life capabilities of cur-rent turbopump and heat exchanger designs.In addition, transient conditioner startupanalyses showed that turbopump shaft accel-erations on the order of 165,000-200,000RPM/sec can be expected. Since currentexperience with propellant-cooled bearings isapproximately 40,000 RPM/sec, pump bear-ing design must be regarded as a criticaltechnology area.

3.3 RCS/OMS Integration Study-Phase C-During this portion of the APS Design Study,all viable oxygen/hydrogen RCS/OMSintegration configurations were compared.The RCS baseline configuration for this studyphase was the parallel flow concept. The OMSuses liquid propellant engines fed from a pro-pellant storage assembly. The degree ofintegration varied from a fully integratedRCS/OMS with common turbomachinery,gas generators, and heat exchangers to a sep-arate system with common propellant tank-age only.

Preliminary screening resulted in theselection of four candidate concepts. Theselected configurations were two fullyintegrated systems, one partially integratedsystem, and a separate system. Fullyintegrated systems utilize common turbo-pumps to supply liquid propellants to theOMS engines, and gaseous propellants to theRCS thrusters via the heat exchangers andaccumulators. In the partially integrated sys-tem, the RCS and OMS are provided dedi-cated gas generators and turbopumps, whichare powered by oxygen and hydrogen propel-lants stored in common accumulators. In theseparate system, the propellant tankage is theonly RCS/OMS interface.

The fully integrated system is attractivedue to hardware commonality, However, as-sociated problems include OMS and RCSmixture ratio differences, RCS accumulatorresupply during OMS operation, and controlsfor propellant sequencing to the OMSengines. Mixture ratio differences were re-solved by using either an extra oxygen pump


for the OMS operation or by utilizing bileveloperating pumps. The remaining problemsare alleviated at reduced levels of integration.

Hardware commonality can be implement-ed in the partially integrated system by usinga common-design turbopump capable of bi-level operation. The bilevel turbopumps aredesigned for the OMS requirements andoperate off-design for the RCS. For both fullyand partially integrated systems, therecommended method of providing accum-ulator makeup propellant during OMSoperation is with a separate, small heat ex-changer/gas generator unit sized to conditiononly the accumulator makeup gas. Thisoption requires the development of anadditional component, but is the preferredapproach since it minimizes the number ofRCS conditioner cycles.

Separating the RCS and OMS completelyexcept for common propellant storage elimi-nates integration concerns, but does requiredifferent component designs. In the separateRCS/OMS, a staged combustion cycle OMSengine is recommended.

At the lower velocity requirements of theeasterly launch mission, the candidateRCS/OMS systems were all weight competi-tive. At the higher velocity increment (2000fps), the separate RCS/OMS system is themost attractive.

3.4 Special RCS Studies - Phase D - Twoalternatives to the turbopump gaseousoxygen/hydrogen systems discussed inSection 3.2 were evaluated. The first alterna-tive maintained the same basic approach asthe turbopump systems but used a highpressure liquid accumulator with a smallpump, thereby reducing pump power require-ments to levels where power sources otherthan hot gas turbines could be considered.Eliminating the turbopump avoided the tech-nology concerns associated with turbopumpbearing life in an environment requiring manyrapid startups per mission.

The second alternative also used a liquidaccumulator in conjunction with a low powerpump but considered distribution of oxygenand hydrogen as liquids to engines which hadthe capability for liquid propellant ignition.

3-5



The "liquid" concept thereby eliminated theneed for gaseous accumulators and thermalconditioning equipment. The following para-graphs summarize the evaluation of these twoalternatives.

3.4.1 Alternate Gaseous Oxygen/HydrogenSystems - The gaseous systems evaluated asalternates to the turbopump systems are il-lustrated in Figure 3-5. The most funda-mental approach is a fully pressurized sys-tem. However, the weight penalty for fullpressurization of both the hydrogen and oxy-gen is severe (5000 Ibm).

The use of hydraulic or electric pumpsreduces the weight penalties that occur withfull pressurization. The number of pump

NIDC E067429 December 1972

cycles are minimized through the use of liquidaccumulators installed downstream of thepumps. Power for pump drive is provided bythe APU. The motor driven pump systemsare weight competitive with the referenceturbopump system. Their principal drawbackis their dependence on APU operation and theincreased number of APU starts incurred.Additionally, power requirements are inexcess of Other APU requirements and highercapacity APU's would be required.

System power requirements can be signifi-cantly reduced at constant system weightthrough the use of a hydraulic hybrid system.The oxygen side of this system is fully pres-surized, whereas a hydraulic motor pump -liquid accumulator configuration is employed

FULLY PRESSURIZED

ELECTRIC PUMP

I

or? :HYDRAULIC PUMP

FIGURE 3-5 GASEOUS 02/H2 CONCEPTS

3-6

MCDONNELL DOUGLAS ASTRONAUTICS COMPANY • EAST


for the hydrogen. The use of fully pressurizedoxygen results in some weight penalty, butsince oxygen accounts for less than 20% ofthe total propellant volume, this penalty issmall, and allows simplification of system de-sign and development.

Pump dependence on APU operation can becompletely eliminated with the use of bellowspumps. System controls are such that whenone liquid accumulator is depleted, itspressurant is vented and refill is providedunder the storage tank pressure head. The useof heated helium pressurant to minimize ventlosses results in a weight competitive system.

3.4.2 Liquid Oxygen/Hydrogen Systems -Liquid oxygen/hydrogen RCS concepts wereevaluated to determine the feasibility of dis-tributing liquid propellants to the engines.This approach would allow removal of the gasgenerators, propellant heat exchangers, andgaseous accumulators, as depicted in Figure3-6. The particular configuration illustrated isbut one of a generic series derived from thebasic turbopump RCS. In addition toadvantages in system simplicity, the liquidconcept offers a large system weightreduction by eliminating the heavy gaseousaccumulators and by avoiding gaseouspropellant conditioning losses.

APU

I

FIGURE 3-6 LIQUID - LIQUID CONCEPT

Liquid cryogenic systems were not con-sidered in previous studies due to concernsassociated with propellant heating in a large,relatively complex distribution system andbecause of concerns with engine pulse modeignition. Figure 3-7 illustrates these concerns.Although the ignition temperature limits are


significantly lower than those previously con-sidered, a review of ignition phenomena withengine manufacturers showed liquid ignitionto be feasible. A thorough investigation ofpropellant heating was also performed toevaluate the feasibility of liquid propellantdistribution.

FEED SYSTEM DESIGN -COULD A PRACTICAL DISTRIBUTION SYSTEM BEDESIGNED AND WOULD THAT DESIGN PROVIDESUFFICIENTLY LOW PROPELLANT HEATINGLEVELS FOR ENGINE OPERATION?

PRESSURE(PS I A)

800600400200

60 80 100 120

TEMPERATURE - °R

ENGINE IGNITION-WAS LIQUID IGNITION FEASIBLE IN APULSE MODE ENGINE?

200°R

<tcc.

GASEOUSSYSTEMDESIGNREGIME

LIQUID SYSTEMDESIGN REGIME

ENTROPY

FIGURES-? PREDOMINANT LIQUIDSYSTEM QUESTIONS

3-7

MCDONNELL. DOUGLAS ASTRONAUTICS COM f ANY • EAST


The primary consideration in the distribu-tion of liquid cryogens is the minimization ofheat input to preclude propellant vaporizationand large density changes. For this study, thehydrogen was delivered supercritically,thereby ensuring single phase flow. Varia-tions in hydrogen density were found to bemore critical than oxygen density variations,and a maximum hydrogen temperature of65 °R was established to maintain acceptableengine performance. A system thermal modelwas assembled and an analysis performed todetermine the implications of this hydrogentemperature constraint. The model incor-porated vacuum jacketed lines with highperformance insulation (HPI) between theinner and outer lines, and thruster thermalstandoffs, similar to those commonly em-ployed for hydrazine engines.

Propellant utilization in limit cycleoperation serves to balance heat input andoutput, and steady state temperatures belowthe 65 °R temperature limit are achieved forvehicle deadbands of 8 degrees or less. Thisanalysis verified that thermal management ofliquid propellants in the APS distributionsystem is feasible if proper attention is givento thermal insulation and isolation of majorheat inputs, such as thruster heat soakback.

Having established the feasibility of theliquid system concept, the remainder of theliquid study effort was devoted to design andsizing considerations. Based on the gaseousstudies described in Section 3.4.1, a hybridsystem, using fully pressurized oxygen andpumped hydrogen, was selected as thebaseline concept. Trade studies wereconducted to determine the most attractivepump and pump driver concepts. Thepreferred approach, based on considerationsof weight and simplicity, was determined tobe a centrifugal pump driven by a hot gas


turbine of the same general type used for thegaseous systems. A liquid system turbopumpwould present considerably less technologyrisk than a gaseous system turbopump since,with liquids, the accumulator can be sized toprovide acceptable shaft acceleration and in-creased bearing life for a minimal weight in-crease. In addition, the liquid system can ac-commodate a wider range of pumpperformance, and it would be much easier tointegrate the pump into the system than in agaseous system where pump performance canaffect operation of the propellant heat ex-changers.

The results of this study suggest two ap-proaches to the evolution of a high perform-ance system. Figure 3-8 shows theseapproaches. One approach starts with agaseous hydrogen-liquid oxygen system. Thesystem uses liquid oxygen because the weightpenalties associated with avoidance of oxygenpumps are small and distribution of liquidoxygen was found to be feasible. Gaseoushydrogen is used because the engine ignitionrequirements are state-of-the-art, similar tothe Pratt and Whitney RL-10 engine. Thisgaseous system could be upgraded by eitherincreasing the gas generator operatingtemperature, thereby increasing systemefficiency, or by utilizing a liquid turbopumpsystem which eliminates the gas generatorand gaseous accumulator. The decision as towhich means of improvement was most at-tractive could be made on the basis of therelative status of technology demonstrationprograms in the areas of liquid ignition andhigh temperature heat exchangers.

The second approach shown in Figure 3-8starts with a simple, fully pressurized liquid-liquid system; later a turbopump and liquidaccumulator could be added to the hydrogenside providing a high performance design.

3-8

DOUGLAS ASTRONAUTICS COMfAfW • EAST



Gaseous Concepts Liquid Concepts

2000°RMARK II CONCEPTS 2000°R

FIGURE 3-8 HYDROGEN/OXYGEN APS CONCEPT SUMMARY

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MCDONNELL DOUGLAS ASTRONAUTICS COMPANY - EAST



4. EARTH STORABLE PROPELLANT STUDIES

During the storable propellant portion ofthe auxiliary propulsion system study, vari-ous RCS/OMS/APU systems were consider-ed to evaluate relative system performance,weight, complexity, flexibility, and vehicleinterface characteristics. Concepts consideredincluded various levels of RCS/OMS/APUintegration. Both modular concepts andconcepts installed integrally within thevehicle were evaluated. Propellant candidateswere monopropellant hydrazine and hyper-golic bipropellants (NTO/MMH).

The basic earth storable Reaction ControlSystems are shown schematically in Figure4-1 for both monopropellant and bipropellantconfigurations. During system operation,

Monopropellant

N2H4

liquid propellents are supplied at highpressure to the thrusters. Propellant tankpressures are maintained by regulatedgaseous helium, and propellant acquisition isprovided by surface tension screens.Component redundancy is consistent with afail-safe, fail-safe philosophy.

The Phase C and E earth storable studywas conducted for the purpose of providingdesign data sufficient to allow resolution ofthe following options:

1. Choice of propellents2. Method of installation (modular vs inte-

gral)3. Degree of OMS/RCS interaction4. Degree of APU integration

Bipropellant

NTO

FIGURE 4-1 EARTH STORABLE RCS STUDY CONCEPTS

4-1

lOfJXUGLAS ASTnaOKJAUTICS COMP*AKIV • EAST


To fulfill these objectives, the effort wasdivided into two phases, RCS/OMS/APU In-tegration Study (Phase C), and SystemPerformance Analysis (Phase E). The PhaseC preliminary system analyses were conduct-ed to establish nominal design points andestablish system sizing data. These baselinedesign points then served as references fordetailed design and installation studies andfor concurrent studies of APU implementa-tion, propellant utilization, and advancedpressurization and tankage concepts. InPhase E, the results from the Phase C studiesand a system reuse study, conducted in paral-lel, were used to compare the selected con-cepts on the basis of safety in flight andground operations, ease of maintenance, reus-ability forecasts, and complexity of flight andground operations. Selected systems wereupdated and refined to incorporate installedsystem considerations, revisions to thecomponent models, and revisions resultingfrom advanced technology studies. Systemanalyses were then repeated to establish thedesign points and thus define final systemweights, volumes, and component require-ments.

4.1 Requirements - The orbiter vehicle con-sidered in this study differed from the oxy-gen/hydrogen study vehicle in that it


contains no inboard main engine tankage. In-stead, the main engine tanks are attached tothe underside of the vehicle and are jettisonedafter orbit insertion. This results in a consid-erably smaller orbiter than the one used in theoxygen/hydrogen studies. Figure 4-2 sum-marizes the changes in requirements from theO2/H2 studies to the storable propellantstudies. Mission requirements are common toboth the O2/H2 and earth storable studiesand were discussed previously in Section 3.1.

Vehicle configuration studies defined threegeneral installation arrangements, as shownin Figure 4-3. In the modular RCS concept,two wing tip pods and a nose pod house theRCS. The modular RCS(OMS) incorporates anose pod, and two fuselage mounted side podscontaining RCS thrusters which are alsocapable of performing OMS maneuvers. Inthe integral RCS concept, the tankage andthrusters are installed integrally throughoutthe vehicle. Impulse requirements vary forthese three concepts due to differences inthruster locations. Figure 4-4 summarizes theimpulse requirements for these candidateinstallation concepts.

4.2 Systems Description - Based on thePhase C preliminary studies, the six systemslisted below were selected for the Phase Esystem performance analysis.

02/H2 STUDIES EARTH STORABLE STUDIES

VEHICLE

ORBIT INSERTION WEIGHTSYSTEM STUDIESREDUNDANCY CRITERIAOMS OPERATION

FULLY REUSABLE

330,000 LBMRCS AND OMS

FAIL-OPERATIONAL/FAIL-SAFEASCENT ABORT; ON-ORBIT AVTANKAGE FOR 2000 FT/SEC

PARTIALLY REUSABLE(EXPENDABLEBOOST PROPELLANT TANKS)265,000 LBMRCS, OMS, AND APUFAIL-SAFE/FAIL-SAFEON-ORBIT AVTANKAGE FOR 1000 FT/SEC(ADD-ON TANKS IN PAYLOAD BAY)

FIGURE 4-2 REVISED REQUIREMENTS FOR EARTH STORABLEPROPELLANT STUDIES

4-2

K/3CSSIOKIKIEILIL. CQKrtff'ARIV • EAST


N1DC E067429 December 1972

Modular (Wing-Tip) Modular (Fuselage)

APU

PCS(QMS)

Integral

APU

^-RCS AND QMS

FIGURE 4-3 CANDIDATE VEHICLE INSTALLATIONS

ON-ORBIT TRANSLATIONS

ATTITUDE MANEUVERS

ON-ORBIT LIMIT CYCLE

RCS DISTURBANCE

REENTRY - YAW-ROLL- PITCH

TOTAL

IMPULSE REQUIREMENT, LB-SEC

INTEGRALRCS

1,133,470

167,321

35,315

48,260

333,712123,00058,600

1,899,678

MODULAR RCS

NOSE POO

136,995

26,320

6,420

21,445

302,330

493,510

1,83

EACHWING POD

498,845

64,225

24,830

26,540

14,43040,280

669,150

1,810

MODULAR RCS (QMS)

NOSE POD

119,500

69,000

20,575

25,675

302,330

537,080

9,805

EACHFUSELAGE POD

4,449,674(QMS = 3,915,924)(RCS = 533,750)

52,250

15,222

30,470

47,00039,500

4,634,116,313

FIGURE 4-4 IMPULSE REQUIREMENTS

4-3

MCDONNELL. DOUGLAS ASTRONAUTICS COI*tr»AIVY - EAST


1. Modularized monopropellant RCS2. Modularized bipropellant RCS3. Modularized bipropellant RCS perform-

ing all maneuvers4. Integral bipropellant RCS sharing

common tankage with the OMS5. Integral monopropellant RCS sharing

common tankage with the APU6. Modularized monopropellant APU

Modularized Monopropellant RCS - Theinstallation of this concept is illustrated inFigure 4-5. In this baseline design, the twowing tip pods and nose pod are used for allon-orbit RCS functions. Reentry yaw isprovided entirely by the nose pod. The for-ward-firing thrusters of the wing tip pods are


protected against the high reentry heatingrates and heat loads by thermal protectiondoors. As shown, the doors and door hydrau-lic actuation mechanisms are attached to thewing, thus facilitating pod installation and re-moval by eliminating the need for a hydraulicinterface between the pod and wing.

Modularized Bipropellant RCS - The basicinstallation features of the monopropellantand bipropellant modular RCS differ verylittle. A typical bipropellant nose pod instal-lation is shown in Figure 4-6. The thrusters inthe nose pod are canted to provide, in con-junction with the wing tip thrusters, up-downand left-right translational maneuvers. A•separate fuselage mounted OMS is used toperform high (>20 ft/sec)AV maneuvers.

DOOR ACTIVATORr-& HINGE

WING TIP POD

WING MOUNTED-THRUSTER DOOR-ECLS COMPONENTS-NOSE POD

PROPELLANT DISTRIBUTION& THRUSTER VALVES

/-PRESSURANT TANK

PROPELLANTTANK

PRESSURANTCONTROLS

SEPARATIONPLANE

TPSJ SEPARATIONPLANE

FIGURE 4-5 WING TIP RCS POD INSTALLATION

4-4

fHCDO/V/VELL DOUGLAS ASTRONAUTICS COMfAKIY - EAST



PROPELLANTTANKS(4)

Section A-ASEPARATIONPLANE

PRESSURANT 'CONTROLS SEPARATION

PLANE

FIGURE 4-6 NOSE RCS POD INSTALLATION

4-5

MCDONNELL DOUGLAS ASTRONAUTICS COMPANY - EAST


Modularized Bipropellant RCS (OMS) -Figure 4-7 illustrates the general arrangementand pod installation of the ModularRCS(OMS) configuration. In this concept,forty-eight RCS thrusters are used to performall maneuvers, thereby eliminating the needfor a dedicated OMS engine. The nose podarrangement for this design is similar to thenose installation for the modular RCS case(Figure 4-6). One of the principle designfeatures of the fuselage-mounted side pods isthat they are shielded by the wings duringreentry. The pod location and shape are tai-lored to preclude any interference with thepayload bay door. Landing center-of-gravityproblems are minimized in the ModularRCS(OMS) by extending the side pods for-


ward of the aft payload bulkhead >and byplacing the oxidizer tanks in the most forwardportions of the pods.

Integrated Bipropellant RCS/OMS -Figure 4-8 depicts the installation of theintegrated RCS/OMS. As. illustrated, theentire system is installed integrally within thevehicle. The design incorporates thirty-seven600 Ibf RCS engines and two 6000 Ibf OMSengines which are served by commontankage. The two fuel and two oxidizer tanksare mounted directly below the payload bayto minimize axial center-of-gravity changesand to preclude the need for a propellantdump during launch aborts. Vertical center-of-gravity travel is accommodated by gim-balling the OMS engines.

_MMH l 099:

' ''''

FIGURE 4-7 MODULARIZED RCS (OMS) SIDE PODS

4-6

MCDONNELL. DOUGLAS ASTRONAUTICS COIHfAIMY- EAST



THRUSTERS600 LB THRUST

PRESSURIZATION BOTTLES

FIGURE 4-8 INTEGRATED RCS AND QMS INSTALLATION

Modular Monopropellant APU and Inte-grated Monopropellant RCS/APU - The APUis installed modularly within the rearfuselage. The design incorporates two mono-propellant tanks and four APU's. In normaloperation, two of the APU's are active, one isidle, and one dormant. In the integrated RCS/APU concept, the propellant is supplied tothe APU's and RCS thrusters from commontankage installed integrally below the pay-load bay. APU propellant pressure is raised inthis case from tank pressure to a higherchamber pressure by an APU-driven boostpump.

4.3 Design Definition - The schematic forthe modular bipropellant RCS and modularbipropellant RCS(OMS), shown in Figure 4-9,is typical of all candidate concepts. Propellanttank operating pressure is maintained by the

use of pressure regulators, and regulationredundancy is provided by utilizing threeparallel regulator branches. Equality in pro-pellant tank pressures (bipropellant systems)is effected by the pressure equalizing valvelocated downstream of the oxidizer heliumregulator. On-orbit propellant acquisition isaccomplished by cylindrical surface tensionscreens. Because reentry accelerations willcause screen breakdown, a false bottom isincorporated in the tanks to isolate sufficientpropellant in the lower compartment for entrymaneuvers. Thrusters are grouped in sets oftwo or three, and in the event of a malfunc-tion, can be isolated either individually or ingroups. Upon completion of the mission, ahelium purge downstream of the thrusterisolation valves is accomplished usingresidual pressurant.

4-7

AffCOO/V/WfLf. DOUGLAS ASTRONAUTICS COMPANY - EAST



PURGE

I-H

PURGE

FIGURE 4-9 MODULAR PCS - NTO/MMHMODULAR RCS (QMS) - NTO/MMH

Figure 4-10 summarizes the designconditions for the six candidate Phase E con-figurations. These design configurations haveevolved based on the preliminary (Phase C)studies and the alternate configuration anal-yses. The following paragraphs provide asummary of the alternate design approachesconsidered for the thrusters, tankage, pres-surization, and thermal control. Also pro-vided is a description of the designs selected,their operation, and the analysis and rationaledetermining their selection.

4.3.1 Thruster Assemblies - The analyticalmodel for the monopropellant thruster wasdefined by the Aerojet Liquid Rocket

Company (ALRC). A schematic drawing ofthe monopropellant thruster assembly withthe associated pressure budget, performance,and weights is shown in Figure 4-11. Designthrust is 600 Ibf at a chamber pressure of 150Ibf/in^. The injector, fabricated from 304Lstainless steel, supplies fuel to the catalystbed at low velocities. The Shell 405 catalystgranules are retained by two layers of screenand a cylindrical, perforated tube retainer.The catalyst is contained within a compart-ment which provides lateral and columnarsupport to the catalyst granules. All parts ofthe catalyst compartment and nozzle arefabricated from Hastelloy B.

4-8

MCDONNELL DOUGLAS ASTHOIVAUTICS COIHIfANY' EAST



MODULAR RCS N2H4

- NTO 'NIMH

WING AND NOSE MODULES UTILIZING HELIUM PRESSURIZATION; TITANIUM TANKAGE:

SURFACE TENSION PROPELLANT EXPULSION; CONVENTIONAL NOZZLE THRUSTERS;

ELECTRIC HEATER/HEAT PIPE THERMAL CONTROL; NO PROPELLANT INTERCONNECTS

BETWEEN MODULES

WING AND NOSE MODULES CONTAINING HELIUM PRESSURIZATION; ULLAGE PRESSUREEQUALIZATION; TITANIUM TANKAGE; SURFACE TENSION PROPELLANT EXPULSION;

FILM-COOLED THRUSTERS; ELECTRIC HEATER THERMAL CONTROL; NO INTRA-MODULE

INTERCONNECTS

MODULAR RCS (OMS) - NTO/MMH BIPROPELLANT FUSELAGE AND NOSE MODULES CONTAINING HELIUM PRESSURIZATION;

ULLAGE PRESSURE EQUALIZATION; TITANIUM TANKAGE; SURFACE TENSION PROPELLANT

EXPULSION; FILMCOOLED THRUSTERS; ELECTRIC HEATER THERMAL CONTROL; NO INTRA-

MODULE INTERCONNECTS

INTEGRATED RCS/OMS - NTO/MMH BIPROPELLANT SYSTEM WITH COMMON, INTEGRATED TANKAGE; HELIUM PRESSURIZATION; ULLAGE

PRESSURE EQUALIZATION; TITANIUM TANKS; SURFACE TENSION PROPELLANT EXPULSION;

FILM-COOLED RCS THRUSTERS; REGEN-COOLED OMS ENGINES (2); ELECTRIC HEATER THERMALCONTROL

INTEGRATED RCS/APU -N2H4 MONOPROPELLANT SYSTEM WITH COMMON, INTEGRATED, TITANIUM TANKAGE; HELIUM PRESSURIZA-

TION/APU BOOST PUMP; SURFACE TENSION PROPELLANT EXPULSION; CONVENTIONAL

NOZZLE THRUSTERS; WATER-COOLED APU; ELECTRIC HEATER/HEAT PIPE THRUSTER THERMALCONTROL

MODULAR APU -N2H4 MONOPROPELLANT SYSTEM WITH ACTIVE-ACTIVE-IDLE-DORMANT REDUNDANCY; MANIFOLDED

TITANIUM TANKAGE; HELIUM PRESSURIZATION; SURFACE TENSION PROPELLANT

EXPULSION; THERMAL BED GAS GENERATOR; MODULATED, WATER-COOLED HYDRAULIC SYSTEM,

CONDUCTIVE-COOLED ALTERNATORS

FIGURE 4-10 PHASE E SYSTEM STUDIES

PRESSURES (PSIA)266 VALVE INLET250 INJECTOR INLET190 UPSTREAM BED150 CHAMBER

PERFORMANCE (e = 40:1)239.7 SPECIFIC IMPULSE, SEC1.78 THRUST COEFFICIENT4250 CHARACTERISTIC VELOCITY

FT/SEC

WEIGHT (LB)18.8 INJECTOR, CHAMBER & NOZZLE2.7 CATALYST3.6 VALVE

25.1 TOTAL

FIGURE 4-11 600 LBF MONOPROPELLANT THRUSTER ASSEMBLY

4-9

MCDONNELL DOUGLAS ASTRONAUTICS COMfANY • EAST



It is estimated that monopropellantthruster catalyst beds will require replace-ment every 5 to 10 flights. Due to this antici-pated high repair frequency, interest has beenfocused on monopropellant thruster main-tenance. In the thruster installation conceptdepicted in Figure 4-12, the thruster andthruster valve are separately mounted tosupport structure; gland seals between thetwo components permit the thruster to be re-moved without disturbing the valve(s) ornecessitating system drain and decontamina-tion. The series thruster valves adequatelyprotect ground personnel from the toxic pro-pellant. Once removed, the entire unit wouldbe transferred to the supplying facility forservicing. Catalyst pack replacement would

be accomplished by cutting open the thrustchamber body, replacing the bed, andrewelding the chamber. Flight acceptancetests would be performed at the same facility.

The bipropellant thruster model employedin this study was fuel film cooled, and con-sisted of a stainless steel parallel platelet in-jector and an integral thrust chamber andnozzle of silicide coated columbium. Figure4-13 presents the thruster performance sensi-tivities.

Two methods of maximizing the RCS(OMS)thruster performance have been implementedin this analysis:

1. Use of statistically determined higherperformance thrusters for the -X function.

2. Reduced thruster film cooling.

ETHYLENE PROPYLENE0-RINGS

BAYONET SEAL

INLET\VALVE

(REMAINSIN PLACE)

FIGURE 4-12 MONOPROPELLANT THRUSTER INSTALLATION

4-10

MCDONNELL DOUGLAS ASTRONAUTICS CO/frffVl/VV• EAST



DESIGN POINTF = 600 LBFPC=200PSIAe = 40:1MR= 1.65

WEIGHT BREAKDOWN - LBMINJECTOR 3.6CHAMBER AND NOZZLE 3.5VALVE 4.4

PRESSURE BUDGET - LBF/IN/

VALVE INLET 300INJECTOR INLET 270CHAMBER 200

11.5

o oiwLU

1LU

£ 300

o

c3Sj 296£

O

> 2924(

/

**f^f

^^

)0 600 800 100 150 200 40THRUST - LBF CHAMBER PRESSURE - LBF/IN.

60 80

EXPANSION RATIO

100

FIGURE 4-13 BIPROPELLANT RCS THRUSTER PERFORMANCE

Although RCS weight is relatively insensitiveto RCS : performance (21 Ibm / sec spec-ific impulse), improvements in the -X trans-lational performance result in significantweight savings (103 Ibm/sec). To take ad-vantage of this potential weight savings, astatistical procedure for selecting high per-formance thrusters is proposed. In thismethod, illustrated in Figure 4-14, thrusterperformance data from injector tests and/orthruster flight acceptance tests is used toidentify the higher performing injectors. Theaverage increase in selected thruster per-formance relative to the shipset nominal valueis dependent upon the ratio of the number of-X thrusters required to the number ofthrusters per shipset. For the RCS(OMS),where 12 out of 48 thrusters are required, theaverage performance gain is three seconds.Concurrent with the -X thruster performancegain is a one second performance degradationin the remaining 36 thrusters of the shipset.

The second method of improving thrusterperformance is to design for a shorter servicelife. The data presented in Figure 4-13 is

based on a nominal wall temperature of2200°F, and corresponds to 22% fuel filmcooling. Since the primary life constraint isthe number of thruster cold starts, thrusterreplacement rates are established by the RCSthrusters. 2200°F is consistent with a 100mission life. A decrease in the percent filmcooling sufficient to yield a one second per-formance gain results in a thruster replace-ment every 50 missions.

The implementation of these two modifi-cations on the modular RCS(OMS) thrusterresults in a -X translational thruster specificimpulse of 306.2, or four seconds greater thanthe nominal performance. This value was usedfor the RCS(OMS) sizing analysis.

4.3.2 Tankage and Propellant AcquisitionAssembly - The baseline propellant storageconcept for this study consisted of 6A1-4VTitanium tankage with a surface tensionacquisition device. This design was the resultof an evaluation of candidate expulsiondevices and a review of materials compatibil-ity.

4-11



APPROACH: THRUSTER PERFORMANCE DATA FROM INJECTOR TESTS AND/ORTHRUSTER FLIGHT ACCEPTANCE TESTS WILL BE USED TO SCREEN HIGH-PERFORMANCE THRUSTER S/N'S FOR THE AXIAL TRANSLATION FUNCTIONS

MARQUARDT R4D PERFORMANCE DISTRIBUTION434 THRUSTER SAMPLES

'SP_0

AVG 0 0.4 0.6 0.8NUMBER OF -X TRANSLATION THRUSTERS

1.0

NUMBER OF THRUSTERS PER SHIPSET

FIGURE 4-14 PERFORMANCE SCREENING OF RCS (QMS) THRUSTERS

Shuttle reusability requirements havelimited consideration of propellant acquisitionconcepts to nonmetallic bladders/diaphragms,metallic bellows, pistons and surface tensionpositive expulsion devices. Figure 4-15 sum-marizes the relative merits of these concepts.

Surface tension devices were chosen as thebaseline propellant acquisition method forthis study. Devices of this type are passive,thus providing high reliability and unlimitedreuse capability. The design consists of screenchannels located around the tank circum-ference, plus a single enclosed collector mani-fold which connects each channel to the outletsump.

The acquisition device will selectively passliquid to the feed system as long as there iscontact with the liquid mass. The wall orient-ed nature of the device ensures that thiscontact will be made. Screen mesh and flowpassage dimensions are selected so that the

pressure drop across the screen vapor/liquidinterface never exceeds the screen bubblepoint prior to reentry. During reentry, decel-eration forces result in channel draining;however, these same forces will orient thepropellant at the outlet of the tank forcontinued propellant use.

4.3.3 Pressurization Assembly - A regu-lated ambient temperature storage heliumpressurization system served as the referencefor this study. This system employs gaseoushelium stored at 4500 lbf/in.2 in titaniumpressure bottles. For bipropellant systems,the fuel and oxidizer have separate pressur-ization systems to preclude the possibility ofpropellant vapor mixing and reaction withinthe pressurization system. Propellant tankoperating pressure is maintained by the use ofpressure regulators, and redundancy isprovided with three parallel regulatorbranches.

4-12

MCDONNELL. DOUGLAS ASTRONAUTICS COMr*A*IY • EAST



CYCLE LIFE

RELATIVE WEIGHT

PERMEATION 'LEAKAGE

SENSITIVITY TO

DYNAMIC ENVIRONMENT

DEVELOPMENT STATUS

DEVELOPMENT RISK

MAJOR ADVANTAGES

MAJOR DISADVANTAGES

BLADDER 'DIAPHRAGM

150 (ELASTOMER)

15-30 (TEFLON)

1.0

HIGH PERMEATION

SUSCEPTIBLE TO TEARS

AND CLAMP-UP FAILURES

DURING SLOSH

GOOD

LOWEST RISK AND COST

SIMPLE. LIGHTWEIGHT

DEVICE

POOR CYCLE LIFE:

PERMEABLE MATERIALS:BLADDER ADSORBS

PROPELLANT

METALLIC BELLOWS

500

3.0

NEGLIGIBLE

PRONE TO CONVOLUTION

WEAR AND IMPACT

DAMAGE

GOOD

MODERATE: QUESTIONABLE

AVAILABILITY OF LARGE

DIAMETER. SEAMLESS

TUBING

POSITIVE SEPARATION OF

PROPELLANT AND PRES-

SURANT: CONTROL OF

FAILURE MODE BY SELEC-TION OF CORE SPRING

CONSTANT: GOOD PROPEL-

LANT STORABILITY

HEAVY AND DIFFICULT

TO CLEANCORROSION)

PISTON

1000

2.7

LIQUID FILM ON WALL

FOLLOWING EXPULSION

PISTON COCKING

FAIR

MODERATE:MAJOR

DEVELOPMENT EFFORT

ASSOCIATED WITH BACK-UP ROLLING DIAPHRAGM

SIMPLE DEVICE WITH

GOOD CYCLE LIFE

RESIDUAL LIQUIDFILM: EXPOSURE OF

DYNAMIC SEALS TOPROPELLANTS

SURFACE TENSION

UNLIMITED

1.3

SATURATED PROPELLANTS

CHANGE IN EFFECTIVE

PORE SIZE

GOOD -*• FAIR; SMALL

TANKS OR SUMPS

FAIR -»-POOR; LARGE TANKS

MODERATE TO HIGH:

UNABLE TO GROUND TEST

LARGE TANKS

PASSIVE DEVICE WITH

POTENTIAL FOR

UNLIMITED LIFE

DIFFICULT TO TEST

AND TO VERIFYINTEGRITY: DUTYCYCLE SENSITIVE

FIGURE 4-15 SUMMARY OF CANDIDATE EXPULSION DEVICES

In depth studies were conducted to eval-uate the weight savings potential offered byadvanced pressurization concepts. Pump fed,volatile liquid, and hydrazine decomposition(monopropellant systems only) pressurizationsystems were compared to the referenceregulated helium system from the viewpointsof weight and complexity. The alternatesystems are shown conceptually in Figure4-16.

A comparison of the primary considera-tions for the four concepts is presented inFigure 4-17. The significant conclusionsdrawn from these comparisons are:

1. For monopropellant systems, hydrazine

decomposition pressurization does showa weight savings over a regulated heliumsystem but at the expense of increasedcomplexity.

2. A pump feed system is lighter than itsregulated helium counterpart, again withincreased system complexity. Addition-ally, this system requires liquid pressureregulators, when used in bipropellantsystems; to avoid large mixture ratioexcursions.

3. Volatile liquid pressurization, althoughattractive from a reusable-refillable mod-ule aspect, is not weight competitivewith any of the other systems.

4-13

/frfCOO/V/VEf-i- DOUGLAS ASTKOMAUTICS COMPANY - EAST


Helium


HydrazineDecomposition

FILL

VENT

He PRESSUREPAD

FILL

FILL

VENT

VENTTO PROPELLANT

TANKTO MAIN

N2H4 TANK

Volatile Liquid

VENT FILL

PROPELLANTTANK

TOTHRUSTERS &OJFILL

TO THRUSTERS

FIGURE 4-16 ALTERNATE PRESSURIZATION CONCEPTS

RELATIVE WEIGHT-APU

-MONOPROPELLANT RCS- BIPROPELLANT RCS (OMSi

RELATIVE VOLUME

PRESSURE BAND

DESIGN OPERATING TEMPERATURE

SENSITIVITY TO MISSION DUTYCYCLE

PROPELLANT UTILIZATIONUNBALANCE ERROR CONTROLBUDGET

MAJOR ADVANTAGES

MAJOR DISADVANTAGES

REGULATED HELIUM

+ 137+ 375+ 700

5.0

* 25°.

AMBIENT

FAIRLY INSENSITIVE

SMALL

WIDE APPLICATION.MINIMAL DEVELOPMENT

HEAVY: HIGH REGULATOR FAILURE RATE

N2H4

DECOMPOSITION

+ 20REF

1.0

+ 21°,

200° F

FAIRLY INSENSITIVE

SMALL

LIGHTWEIGHT. REDUCESTHERMAL CONTROL

COMPLEX: TIGHT PRESSURE DEAOBANO NECESSITATES MANY PUMP CYCLES:REQUIRES HEAT EX

CHANGER

PUMP FEED

REF+ 100REF

2.0

+ 100*

- 0 '

AMBIENT

ACCUMULATOR SIZELIMITS RCSAV

LARGE 0/F ERRORFOR ACCUMULATORSOUT OF PHASE

LIGHTWEIGHT: PRESSURANT LEAKAGENOT CRITICAL

WIDE VARIATION INTHRUSTER INLETPRESSURE OR HIGHPUMP POWER:

VOLATILE LIQUID

+ 43+ 1075

+ 4400

2.2

+ 85*- 0 '

125 - 165°F

HEATER POWER LEVELLIMITS TOTAL RCS THRUST

LARGE ERRORS DUE TOMODULE TEMPERATUREDIFFERENCES

RELIABLE NO-NIOVINGPARTS. NO FILL VENTREQUIREMENT FOR RECYCLE

HEAVY: WIDE PRESSUREBAND: HIGH POWERREQUIREMENT

•"WITHOUT LIQUID PRESSURE REGULATION

FIGURE 4-17 COMPARISON OF CANDIDATE PRESSURIZATION CONCEPTS

4-14



4.3.4 Thermal Control - Analysis was per-formed to evaluate the technical complexityand to define the weight implications associ-ated with the thermal control of the alternateBCS configurations.Module thermal control isrequired primarily to protect the system fromthe extreme environments in space, as well asthose induced during entry. Additionally,monopropellant thruster injectors requirecooling to preclude explosive decompositionof the propellant under certain malfunctionconditions. The primary thermal constraintsare associated with the propellants and thethrusters. Allowable propellant temperatureranges have been established to be 40 to125°F for bipropellants (NTO/MMH) and 50to 125 °F for monopropellant hydrazine. Thesteady state and transient thermal responseof the wing tip RCS modules have beenexamined using a two-dimensional thermalmodel. These calculations indicate that themaximum steady state uncontrolled temper-ature range is -110 to 165°F. Minimum tem-peratures, which occur with continuouslyshaded pods, require heaters to preventpropellant freezing. Heaters are sized toprovide a maximum power of 303 watts forthe monopropellant system (including 10watts per thruster to maintain a minimum150°F catalyst temperature), and 161 wattsfor a bipropellant system. Correspondingmaximum energy requirements are 36.8 kwh(monopropellant system) and 17.3 kwh (bi-propellant system). The maximum tempera-ture of 165 °F is somewhat above the desiredmaximum temperature, and thermal controlis required to prevent propellant overheating.

Thruster thermal requirements have beendefined in order to provide adequate thrusterlife and reliability. Figure 4-18 summarizesthe thermal limitations associated withmonopropellant thruster start up, operation,heat soakback, and nonoperation. Thecounteracting constraints on minimumcatalyst bed temperature and maximum in-jector soakback temperature are of primarysignificance. The restriction on minimumcatalyst temperature arises from the poorstructural properties of the spontaneouscatalyst (Shell 405) and its tendency to


generate "fines" under repeated cold thrusterstarts. The restriction on injector temperatureis based on propellant thermal stability con-siderations, i.e., the maximum injectortemperature is kept sufficiently low so as topreclude explosive detonation of the pro-pellant under certain malfunction conditions.

Thermal control of bipropellant thrusters isnot as restrictive. The primary concern forbipropellant thrusters is with vacuum-igni-tion pressure spiking. During pulsingoperation, energy-rich detonatable chemicalresidues (mostly monomethylhydrazine ni-trate) can accumulate and, in sufficientquantity, can produce high-magnitude igni-tion overpressures. To alleviate this problemon the Apollo CSM and LM RCS, the thrusterinjectors were maintained in excess of 70°F topromote rapid vaporization of the fuel.Meeting this same criteria with 600 Ibfthrusters will require a maximum power inputof 5.4 watts/thruster.

The basic aspect of monopropellant thrust-er thermal control is the conduction pathbetween the thruster and the surroundingstructure. To minimize the injector and valveseat temperature, it would be desirable toattach the injector and valve to massivestructure with a high heat capacity. However,such a connection would provide a substantialheat short during periods of nonoperation,and would thus conflict with the goal of mini-mizing the heater power required to maintainminimum catalyst temperature. Of the alter-nate thermal connection concepts which wereevaluated, the most attractive was a heatpipe.

In order to determine the operational con-siderations of a heat pipe system, a thermalmodel of the module was constructed. Thismodel was used to determine the steady stateheater requirements necessary to maintainthruster minimum temperatures, to deter-mine the heat delivered to the EnvironmentalControl and Life Support (ECLS) duringthruster soakback, and the system transientresponse. The thruster-ECLS model was usedto examine the response of the thruster, themounting plate, and the module during a heatsoakback condition. For the single firing case,

4-15

fwcin><a>RiKiEii.iL BS<a>ajGiLA&



o

LU

OO

OQ«£

15

10

Detonation Potential(304L SS LINES) o

ocLUQ.

DETONATABLE

DECOMPOSITION

•EXOTHERM

Q»

\\

o \§ >

\Bs

I

Catalyst Loss• 10 STARTS

0 40°FD 65°F

A 120°FkX

X

9.v-&.i i i

PROPELLANTPROPELLANTPROPELLANT

A

01 1

300 400 500PROPELLANT LINE TEMPERATURE -°F

50 100 150 200 250 300

REF: AFRPL TR 71-41

INJECTORTEMPERATURE

1500°F

INITIAL CATALYST TEMPERATURE - °F

REF: AFRPL TR 71-103

VALVE SEATTEMPERATURE f 200°F

/-BED TEMPERATURE2 150°F

MAXIMUM EXTERNALTEMPERATURE <600°F

600

FIGURE 4-18 MONOPROPELLANT THRUSTER THERMAL CONSTRAINTS

the injector temperature rises to 500°F whenuncontrolled. With the heat pipe, the injectortemperature rise is reduced and it is cooledmore rapidly. The problem is more servere,however, when multiple firings occur. For thiscase (Figure 4-19), soakback continued for2000 seconds. At that time, it was assumedthat a second pulse occurred in which thethruster and catalyst temperatures wereelevated to the steady state firing conditions.With a heat pipe system, the injector tem-perature does not exceed the 500 °F maximuminjector temperature. For the uncontrolledsystem using only the capacitance of themodule and the thruster mounting plate,however, the injector temperature rises toabout 560°F, well above the allowable limit.These thruster-module-ECLS calculations in-dicate that satisfactory operation can beachieved by linking the module to the ECLSwith heat pipes.

INJECTOR TEMPERATUREHEAT PIPE

4000

FIGURE 4-19 EFFECT OF PULSE

OPERATION ON THRUSTER

TEMPERATURE TRANSIENTS

4-16

MCDONNELL. DOUGLAS ASTKOMAUTICS COMPANY • EAST


KIDC E067429 December 1972

Steady state and transient thermalresponses have been examined for the fuse-lage mounted module. Maximum uncontroll-ed propellant temperatures (115°F) weresomewhat less than for the wing tip podbecause of additional communication with thevehicle. However, heater power levelsrequired to maintain minimum temperatureswere substantially higher, due to theincreased tank size and reduced thermalcommunication with earth. Tank structureand support transients have been examined toevaluate techniques for reducing the principalleaks. Tank support was provided by alumi-num structure cantilevered from the fuselageside. In this configuration there is no con-duction heat transfer from the tank to thethruster enclosure, and the thruster enclosureserves as a radiation shield between the tanksand space. The results show that to maintain40°F conditions for the four tanks in a modulesubjected to a cold environment requires an

10.5-

10.4-

10.3-

10.2-

o CHAMBER PRESSURE = 200 LBF/IN.2Ao ATTITUDE CONTROL THRUSTERS

EXPANSION RATIO = 40o X-TRANSLATIONAL THRUSTERS

EXPANSION RATIO = 40

input of 330 watts.

4.4 Concepts Comparison - System sizinganalyses were performed for the candidateconfigurations. Optimal design points weredetermined by generating system weightsensitivities to chamber pressure, expansionratio, and mixture ratio (for bipropellantsystems), as presented in Figure 4-20 for themodular bipropellant RCS. As shown, theexpansion ratios of the -X translationalthrusters have been optimized as an inde-pendent parameter. This results in a signifi-cant weight savings for the RCS(OMS); thesavings realized by the remaining systems areminimal and would not warrant the use of adifferent expansion ratio. Pod structure andthermal protection weight drives theoptimum modular RCS design points to lowexpansion ratio and high chamber pressure(both favoring smaller thrusters and thereforesmaller pods).

10.5-

10.4-

10.3-

10.2-

10.1-

o MIXTURE RATIO = 1.65o ATTITUDE CONTROL

THRUSTERS EXPANSIONRATIO = 40

o X-TRANSLATIONALEXPANSION RATIO = 40

1.0 1.2 1.4 1.6 1.8 2.0MIXTURE RATIO

o CHAMBER PRESSURE =

50 100 150 200CHAMBER PRESSURE - LBF/IN/A

10,300-

10,250-

UJ3= 10,200-

10,150-

200 LBF/IN.?Ao MIXTURE RATIO = 1.65 EX-TRANS= 20

eX-TRANS = 80

EX-TRANS= 60

eX-TRANS= 4°

20 40 60ATTITUDE CONTROL THRUSTERS

EXPANSION RATIO

80

FIGURE 4-20 SYSTEM WEIGHT SENSITIVITIESModular RCS-NTO/MMH

4-17

23><O><L/(G/LX5lS ASTDSORJAQJITIHCS EAST


Figure 4-21 summarizes the optimal designparameters and system weights. In order toprovide a common ground for weight com-parison, a total propulsion system weightcomprised of the applicable RCS, OMS, andAPU weight is included. The evaluation of adedicated OMS was not a part of this study.However, in order to properly compare thealternate concepts, a generic OMS was neces-sary. The OMS weight was derived from theOrbit Maneuvering System Trade Studies(Contract NAS 9-12755).

When comparing system weights, it isnecessary to differentiate between systemexpendables weight, which has a 1:1 tradeoffwith payload, and system inert weight, whichreduces payload by 1.4 Ib for each pound in-


crease. Thus, the proper method of comparingsystems is on. the basis of payload penalty.Comparisons on the basis of payload magnifythe weight penalty associated with modu-larized system concepts. Figure 4-22 presentsthe relative payload weights. Comparison ofthe candidate configurations reveals thefollowing:

1) The lightest system approach is realizedwith an integral, bipropellant RCS/OMSand a modular monopropellant APU.

2) The payload penalty for modularizing thebipropellant RCS and OMS is 2184 Ibm.

3) The modularized, bipropellant RCS(OMS)is almost 1300 Ibm heavier (on a payloadbasis) than the combined weight of a mod-ularized RCS and modularized OMS.

SYSTEM

MODULAR RCS

(MONOPROPELLANT)

MODULAR RCS(BIPROPELLANT)

MODULAR

RCS (OMS)

(BIPROPELLANT)

INTEGRATED

RCS/OMS(BIPROPELLANT)

INTEGRATED

RCS/APU(MONOPROPELLANT)

MODULAR APU(MONOPROPELLANT)

DESIGN SUMMARY

WING AND NOSE MODULES

HELIUM PRESSURIZATION

TITANIUM TANKAGESURFACE TENSION POSITIVE EXPULSION

CONVENTIONAL NOZZLE THRUSTERS

WING AND NOSE MODULES



FILM COOLED THRUSTERS

FUSELAGE AND NOSE MODULES




COMMON TANKAGE LOCATED BELOW

PAYLOAD BAY

HELIUM PRESSURIZATIONTITANIUM TANKAGESURFACE TENSION POSITIVE EXPULSION


COMMON TANKAGE LOCATED BELOW

PAYLOAD BAYHELIUM PRESSURIZATION (RCSlBOOST PUMP PRESSURIZATION (APU)

SURFACE TENSION POSITIVE EXPULSION

CONVENTIONAL NOZZLE THRUSTERS

ACTIVE. ACTIVE. IDLE. DORMANT APU

REDUNDANCY

WATER COOLED APU

AR FUSELAGE MODULESHELIUM PRESSURIZATION

TITANIUM TANKAGE

SURFACE TENSION POSITIVE EXPULSION

ACTIVE. ACTIVE. IDLE. DORMANT APU

REDUNDANCY

WATER COOLED APU

OPTIMAL DESIGN PARAMETERS

RCS IMPULSE

1.831.810

1.831.810

1.973.464

(RCS)

7.831.849

(OMS)

1.899.678

(RCS)

7.841.338

(OMS)

1.899.678

F

600

600

600

600

600

PTANK

312

368

250

250

272

766

PC

150

200

150

150

(RCSl

150

(OMS)

125(RCS)1000

(APU)

500

t

20(ATT CONT)

20(=X)

40(ATT CONT)

40

( • X )

60(ATT CONT)

80

( •X)

60(ATT CONTl

60( • X )

40(ATT CONT)

40( • X )

MR

1.65

1.65

1.65

'sp

228.8(ATT CONT)

228.8(:X)

296.1

(ATT CONT)

296.1

( • X )

299.1

(ATT CONT)

306.2(-X)

299.1

(ATT CONT)299.1

( X)

315.9

(OMSl

235.4(ATT CONT)

235.4

(• XI

WEIGHT

12.889 (12,889 - RCS)( 3,295 - APU)

(28,790 - OMS)*

44,974

10.133 (10,133 -RCS)

( 3,295 - APU)(28,790 - OMS)'

42.218

40,155 (40.155-RCS (OMS)

( 3,295 - APU)

43.410

37,360 (37, 360-RCS/OMS)( 3.295 - APU)

40,655

14.586 (14.586- RCS APU)

(28.790 - OMS)'

43.376

3.295

*OMS SPECIFIC IMPULSE = 313 SEC

FIGURE 4-21 DESIGN POINT SUMMARIES

4-18

DOUGLAS ASTRONAUTICS COMPANY • EAST



APU

CO

MONOPROP APU

FIGURE 4-22 RELATIVE PAYLOAD WEIGHTS

4) The modularized monopropellant RCShas a reduced payload of 3130 Ibm whencompared with the modularized bipropel-lant RCS.

Each configuration is the result of an individ-ual optimization; tankage and thruster loca-tions have been separately established, anddesign points defined consistent with theparticular requirements of each system. Thesefinal comparisons are therefore considered tobe realistic evaluations of the alternate con-figurations. One significant advantageafforded by integral systems is weight, butthe maintenance complications associatedwith integral systems are unacceptable for areusable vehicle. Consequently, only modularsystems can be seriously considered for use onShuttle.

As discussed in Section 1, the objective ofthis study was to develop design and pro-grammatic data for competitive reactioncontrol systems in sufficient detail that aselection can be made between the variousconcepts. In keeping with this objective, theconcluding effort in this study was an assess-ment of the weight implications of selectedconfiguration changes on the design pointweights. The effects of changes in pres-surization concept, type of tank expulsion,and tank material are shown in Figure 4-23 forthe modular bipropellant RCS. Weightsavings are possible in the area of pressur-ization, with the largest savings available forthe high impulse configurations (modularRCS(OMS) and integrated RCS/OMS). In

4-19

/VfCDO/V/Vff-f. DOUGLAS ASTRONAUTICS CO IMF*A IVY • EAST


general, the pump fed pressurization conceptoffers the largest savings; however, its adop-tion results in increased system complexity.By contrast, the savings afforded by compos-ite pressurant tanks reflect no decrease in sys-tem reliability. Additionally, composite tanksgenerally provide a leakage failure mode


rather than fracture. The weight penalties as-sociated with system redundancy are also pre-sented to allow evaluation of the weightpenalty associated with the fail safe/fail saferedundancy philosophy. The weight savingsshown represent the elimination of all com-ponents except those necessary for comple-tion of a failure free mission.

SYSTEM

BASELINE SYSTEM: TWO WING TIP AND ONE NOSE-MOUNTED MODULES CONTAINING A TOTALOF FORTY 600 LBF FILM COOLED THRUSTERS, HELIUM PRESSURIZATION,TITANIUM TANKS.AND SURFACE TENSION POSITIVE EXPULSION.

RELATIVE SYSTEM WEIGHT-600 -400 -200 0 +200 + 400 + 600

PRESSURIZATION- VOLATILE LIQUID-PUMPREDUNDANT EXPULSION TANKS- POSITIVE EXPULSION DEVICE (SCREEN TANK)- BLADDER- BELLOWS- PISTON

TANK MATERIAL-301 CRYOFORM- COMPOSITE PRESSURANT TANKS

NON-REDUNDANT CONFIGURATION-VALVES, REGULATORS, ETC.- THRUSTERS

-59-339

-121

J+642

100 +1707

J *230+1399

REFERENCE WEIGHT = 10,133 LBM

FIGURE 4-23 WEIGHT RELATIVE TO BASELINE MODULAR RCS (N204/MMH)

As originally defined, Phase E was to be afinal performance analysis of the six systemsdescribed above. However, prior to thecompletion of this evaluation, NorthAmerican Rockwell (NR) was awarded theSpace Shuttle Orbiter prime contract byNASA. The NR shuttle configurationemploys a dedicated bipropellant OMS and amonopropellant RCS installed in fuselage andnose modules. The RCS utilizes 40 thrustersof 1000 Ibf each. Common size propellant andpressurant tanks are used in the nose andfuselage modules. In order to maximize theutility of the APS study results, additional

analysis was performed to further study thefuselage module options. Five alternatefuselage configurations were evaluated:

1. Dedicated OMS, common RCS-OMStankage

2. Common RCS-OMS thrusters, commontankage (the RCS(OMS) of Phase E)

3. Dedicated OMS, dedicated tankage (bi-propellant RCS)

4. Common RCS-OMS thrusters, dedicatedtankage

5. Dedicated OMS, dedicated tankage(monopropellant RCS)

4-20

MCDONNELL DOUGLAS ASTITOn/Al/TICS COMPANY • EAST


The design points, shown in Figure 4-24, forthese alternate fuselage configurations wereestablished by analogy to the Phase Esystems. Common tankage contains RCS andOMS propellants jointly, whereas dedicatedtankage provides separate tankage for thetwo functions. The use of dedicated tankagefor the OMS function profits from the factthat full-tank surface tension acquisition isnot required for the large tanks, since a smallscreen trap is sufficient.

A comparison of relative payload penaltiesfor the five concepts is presented in Figure4-25. This figure reflects a 1:1 tradeoffbetween system expendable weight andpayload decrease, and a 1:1.4 tradeoffbetween system inert weight and payloaddecrease. Comparison of these systemsreveals the following:

1. Minimum vehicle weight is provided bythe concept employing a dedicated OMSand dedicated tankage.

2. A 2700 Ibm payload penalty is associatedwith the use of a monopropellant RCS, asopposed to a bipropellant RCS.

IWDC E067429 December 1972

3. The use of RCS thrusters for allmaneuvers results in a 750 Ibm payloadpenalty, referenced to the minimumweight system.

4. Dedicated tankage is the preferred choicefor the RCS(OMS) configuration, sinceweight differences are minimal.

The final comparison of interest concernsthe fuselage mounted bipropellant RCSconcept (configuration No. 3) describedabove, and the wing mounted bipropellantmodular RCS concept. The wing moduleconfiguration is approximately 300 Ibmlighter. This difference is minimal, and there-fore definition of the more attractive conceptmust certainly consider additional para-meters, such as maintainability.

Additional evaluation of the candidateconfigurations with regard to operational,maintenance, and safety considerations wasperformed. The objective was to determinewhat advantages or disadvantages are asso-ciated with various classes of systems, there-by allowing general comparisons to be made,e.g., monopropellant vs bipropellant, integralvs modular.

POD DESIGN CHARACTERISTICS

RCS 'OIYIS

TANKAGE

COMMON

COMMON

DEDICATED

DEDICATED

DEDICATED

PROPELLANTS

RCS

N204 'MMH

N204 'MMH

N20, ViMH

N204'MMH

N2H4

OMS

N204 'MMH

N204 'MMH

N204 'MMH

N204.'MMH

N204 WMH

RCS 'OMS

ENGINES

DEDICATED

COMMON

DEDICATED

COMMON

DEDICATED

SYSTEM DESIGN POINTS

THRUST

RCS

600

600

GOO

600

600

OMS

6000

600

6000

600

6000

CHAMBER

PRESSURE

RCS

150

ISO

175

150

125

OMS

150

150

125

150

125

TANK

PRESSURE

RCS

249

249

285

249

235

OMS

249

249

205

249

205

EXPANSION

RATIO

RCS

60

60

60

60

40

+XRCS

60

90

60

90

40

OMS

90

-

75

-

75

MIXTURE

RATIO

RCS'

1.65

1.65

1.65

1.65

-

OMS

1.65

1.65

1.65

1.65

1.65

SPECIFIC IMPULSE

RCS(')

299.1

299.1

299.3

299.1

235.3

X R C S

299.1

306.2

299.3

306.2

235.3

OMS

315.9

-

313.5

-

313.5

(1) USE 76'. lsp (SS) FOR PULSING

FIGURE 4-24 DESIGN POINT SUMMARY

4-21

/MODO/V/Vf«.(. DOUGLAS ASTRONAUTICS COMPANY * EAST



3000

i>-

2000

O_

O

O

* 1000LU

2701

745

REFERENCE

701

169

TANKAGE

ENGINES DEDICATED

PROPELLANT

DEDICATED COMMON

COMMON DEDICATED DEDICATED COMMON

NTO/MMH NTO/MMH

FIGURE 4-25 COMPARATIVE PAYLOAD PENALTIES FORCANDIDATE TANK/ENGINE SYSTEMS

4.5 Reusability Effects - Propellant handl-ing considerations have a considerableinfluence on earth storable system designs.Due to the toxicity, corrosiveness and, in thecase of bipropellants, hypergolic nature of thepropellants, safety considerations dictate thatonly those personnel directly involved in RCSservicing be allowed in the proximity of thesystem during these operations. For a systemthat is installed integrally within the vehicle,t,his constraint would force vehicle mainte-nance operations to be conducted serially, andwould extend the vehicle turnaround time byapproximately two days. To meet the Shuttleobjective of a two week turnaround, attentionhas focused on the use of removable, self-con-tained modules. The modules would beremoved from the vehicle after landing andtaken to a remote facility suitable for safemaintenance and filling operations. Vehiclemaintenance could then proceed without

elaborate precautions. The following dis-cussion defines the anticipated maintenanceprocedures, assuming a modularized propul-sion system.

The major portion of the system will remain"wet" but it is considered necessary to purgedry the thruster assemblies for safety andreuse. System safing will begin during re-entry, following vehicle transition to airplanemode of flight. At this time, propellant isola-tion valves will be closed and the thruster as-semblies purged with residual helium pres-surant.

Figure 4-26 summarizes the ground safingand servicing requirements. After vehiclelanding and cooldown, system depressuriza-tion will be verified and a nitrogen purge ofthe thrusters will be performed to assure thatall propellants have been cleared. Systempower will then be removed and thrusterthroat plugs will be installed.

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Safing Area

L J >^x

OPERATION• PURGE THRUSTERS• DEPRESSURIZE SYSTEM• INSERT THRUSTER

THROAT SEALS. REMOVE SYSTEM DATA TAPES• REMOVE MODULES Hypergolic Building

NORMAL OPERATION• VISUALLY INSPECT. PERFORM THRUSTER LEAK TESTS. FUNCTIONALLY TEST VALVE

DRIVER CIRCUITS• FUNCTIONALLY TEST HEATER

SYSTEM• LOAD PROPELLANTS• MOVE MODULE TO VABMAINTENANCE• DRAIN & DECONTAMINATE SYSTEM. PURGE SYSTEM WITH HOT GN2

. REMOVE & REPLACE FAULTY COMPONENTS

. FUNCTIONALLY TEST REPAIRS

VAB and Launch Area

OPERATION. INSTALL MODULES IN VEHICLE• VALIDATE ELECTRICAL CONNECTIONS. REMOVE THRUSTER THROAT SEALS• SERVICE HELIUM BOTTLES• PRESSURIZE PROPELLANT TANKS

FIGURE 4-26 GSE REQUIREMENTS

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MCDOftlNELL. DOUGLAS ASTRONAUTICS CO/tfffVl/VV • EAST


Propulsion modules will be removed to aremote facility for servicing. Normal servicingwill include such operations as testing valvedriver circuits and the heater system and per-forming leak checks. The tanks and controlcomponents will be maintained wet to themaximum extent possible. Gravity fill proce-dures will be employed, and propellant quan-tities determined by weight (modules re-moved) or by overfilling and metering off therequired ullage volume (modules installed).

Maintenance operations will be performedbased on inflight checkout intelligence data.Failure probability analyses show that the re-quired system repair frequency will be high.As shown in Figure 4-27, estimates vary froma propellant system failure every 1 to 3 flightsfor a bipropellant RCS and a propellant sys-tem failure every 1 to 5 flights for a monopro-pellant RCS. These numbers illustrate theimportance of component accessibility in re-ducing maintenance downtime. A review ofcomponent failure data from previously flownspacecraft has indicated that the primarycause of failure is contamination, and thecomponents most susceptible to contamina-tion are the pressurant check valves and thepropellant valves. System reliability and re-usability would be benefited by maintainingthe propellant feed system in a wetted condi-tion.


The replacement of propellant systemcomponents other than thrusters and pres-surization components would require eithercomplete or partial system draining and flush-ing to remove residual propellants from thesystem in order to assure a safe environmentfor maintenance personnel. Of the variouscleaning methods available, the single-flushmethod and the vapor-phase flush method, inwhich the solvent is introduced into the sys-tem in its vapor phase, appear to be the mostpromising approaches for decontamination ofthe shuttle RCS when necessary. A review ofavailable solvents has identified Isopropanoland Freon TF as the most attractive solventsfor fuel and oxidizer systems, respectively.

Several conclusions regarding reuse areapplicable regardless of the configurationchosen. The successful implementation of amulti-mission vehicle will require thoroughconsideration of reusability throughoutsystem design, including the establishment ofthermal control requirements consistent withreusability, and in the definition of servicing,safing and maintenance operating procedures.The safety and reuse criteria identified in thisstudy have been so categorized, and are sum-marized in Figure 4-28. Reuse considerationsnecessitate added care in the selection ofcomponent types and arrangement to mini-mize the generation and effects of contami-nants on system operation.

SUBSYSTEM ASSEMBLY

TOTAL RCS

TOTAL RCS LESS INSTRUMENTATION

RCS INSTRUMENTATION

PRESSURIZATION COMPONENTS

PROPELLANT STORAGE AND CONTROL

PROPELLANT TANKS & FILL VALVES

THRUSTER ASSEMBLIES

MONOPROPELLANTMISSIONS

1.0 -4.7

1.4 -6.6

4.1 -16.5

10.4 — -57.3

6.5 -52.0

19.1— * 187.0

2.1 -8.8

BIPROPELLANT

MISSIONS

0.7 -2.9

0.9 -3.7

3.5 -14.3

5.0 — -25.3

4.1 -31.8

23.4 — 224.0

1.4 -5.0

FIGURE 4-27 MEAN TIME BETWEEN COMPONENT FAILUREModular RCS

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APS STUDY MDC E0674


Design• PROVIDE SYSTEM ACCESS WITH VEHICLE IN EITHER HORIZONTAL OR VERTICAL (LAUNCH) ATTITUDE• EMPLOY INTERLOCKS OR OTHER SAFEGUARDS ON MANUAL VALVES TO ASSURE VALVES ARE IN FLIGHT

POSITION PRIOR TO LIFT-OFF• USE SEPARATE PRESSURANT SUPPLIES FOR FUEL AND OXIDIZER• DESIGN FOR FAIL-SAFE, FAIL-SAFE REDUNDANCY OR BACK-UP CAPABILITY ON ALL ACTIVE COMPONENTS• USE COMPOSITE (OVER WRAP) PRESSURANT TANKS TO ASSURE TANK FAILURE IS BY LEAKAGE RATHER

THAN FRACTURE. EMPLOY FLEXIBLE PROOF TEST FACTORS ON PRESSURE VESSELS, ADJUSTING FOR TANK MATERIAL

ENVIRONMENT, HOOP LOADS AND REQUIRED LIFE. PROOF TEST WITH LIQUID NITROGEN TO REDUCEREQUIRED PROOF PRESSURE LEVELS AND/OR TO VERIFY GREATER CYCLE LIFE FOR GIVEN DESIGNSAFETY FACTOR

• USE MATERIALS THAT ARE COMPATIBLE WITH PROPELLANTS AND RESIDUES FORMED BY PROPELLANTREACTION WITH THIRD AGENTS, I.E., H20, C02, SOLVENTS, ETC

• PROVIDE REPLACEABLE COMPONENT CARTRIDGES FOR HIGH FAILURE RATE ITEMS• USE BREAKABLE SEALS BETWEEN MONOPROPELLANT THRUSTER AND THRUSTER VALVES TO FACILITATE

CATALYST REPLACEMENT

Thermal Control• THERMALLY CONTROL TO FOLLOWING TEMPERATURE CONSTRAINTS:

(°F)N2H4 THRUSTER CATALYST> 150 (CATALYST LIFE)

INJECTOR < 500 (DETONATION POTENTIAL)VALVE < 200 (SEAT LIFE)

N204/MMH THRUSTER INJECTOR > 70 (IGNITION PRESSURE SPIKING)VALVE <200 (SEAT LIFE)

PROPELLANTS N2H4 > 50 (FREEZING)N204/MMH > 40 (IGNITION PRESSURE SPIKING/N204 FREEZING)

Sating and Maintenance• PERFORM POST-DEACTIVATION FLIGHT PURGE OF THRUSTER ASSEMBLIES. DEPRESSURIZE SYSTEM FOR GROUND SERVICING AND/OR POD TRANSPORT• INSERT THRUSTER THROAT SEALS AFTER LANDING• MAINTAIN WET TANK AND CONTROL COMPONENTS TO MAXIMUM PRACTICAL EXTENT. WHEN NECESSARY,

FLUSH SYSTEM WITH VAPORIZED SOLVENTS (FREON TF-NTO; ISOPROPANOL - MMH, N2H4) FOLLOWED BYHOT GN2 PURGE. PULSATE FLOW OF GASIFIED SOLVENTS TO SCAVENGE PROPELLANT VAPORS).

• PROVIDE CLOSED-VENT SYSTEM FOR PROPELLANT DUMP• AVOID AIR DRY OF EPT RUBBER EXPULSION BLADDERS

Servicing• UTILIZE MOLECULAR SIEVES TO REMOVE SOLUBLE IRON FROM N204 PROPELLANT DURING FILL OPERA-

TIONS (MINIMIZE RHEOPEXY POTENTIAL)• EMPLOY GRAVITY FILL PROCEDURES FOR PROPELLANT SERVICING. AVOID BACKFILL THROUGH SCREEN SURFACE TENSION TANKS TO AVOID SCREEN LOADING IN UNSUPPORTED

DIRECTION. AVOID PROPELLANT TEMPERATURE CONDITIONING DURING FILL. VERIFY PROPELLANT LOAD BY COMPLETELY

FILLING TANKS AND OFF-LOADING ULLAGE

FIGURE 4-28 SAFETY AND REUSE CRITERIA

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5. STUDY CONCLUSIONS

Candidate oxygen/hydrogen propellantauxiliary propulsion systems were comparedon the basis of weight, technology, simplicity,flexibility, and reusability. Three oxygen/hydrogen turbopump RCS configurationswere evaluated with respect to operationalcharacteristics and controls requirements. Allthree configurations were judged to becompetitive, with the series-downstream tur-bine RCS offering a slight advantage. Tech-nology concerns associated with this class ofsystem include a high conditioner cycle liferequirement and excessive turbopump shaftaccelerations.

On the basis of weight, the most attractivesystem was an oxygen/hydrogen system inwhich the propellents are delivered to thethrusters as liquids. In addition, this systemconcept escapes the conditioner life and turbo-pump shaft acceleration technology concernsassociated with the oxygen/hydrogen turbo-pump RCS.

A change in orbiter design during the studyreduced RCS impulse requirements and pro-moted renewed interest in earth storable pro-pellant systems. Figure 5-1 shows that, evenat smaller orbiter impulse levels, a storablesystem weighs on the order of 3,000 to 5,000

CX3

CO

16,000

14,000

12,000

10,000

8,0

6,000

4,000

STORABLESORBITER

IOXYGEN/HYDROGENORBITER

G02/GH2

SERIES(TURBINE DOWNSTREAM)

PUMPED LH2

PRESS'D L02

1.5 1.75 2.0 2.25

TOTAL IMPULSE - 106 LBF-SEC

2.5

FIGURE 5-1 COMPARISON OF MODULAR STORABLE RCS CONCEPTSWITH INTEGRAL CRYOGENIC RCS CONCEPTS

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MCDONNELL. DOUGLAS ASTKONAUTICS COMPANY • EAST



Ibm more than a liquid-liquid oxygen/hydro-gen system. However, the reduced volume ofthe smaller orbiter does not allow use of anoxygen/hydrogen OMS. This weight differ-ence could not justify the high cost and tech-nology risk associated with the developmentof an oxygen/hydrogen RCS alone. Thus,earth storable propellant systems were se-lected for the RCS and oxygen/hydrogenpropellants were eliminated from further con-sideration.

Three basic storable propellant conceptswere defined: a modular concept utilizingwing and nose modules, a modular conceptutilizing fuselage and nose modules, and anon-modular concept wherein the RCS was in-

tegral with the vehicle. For each concept,alternate configurations were defined byspecifying the propellants (monopropellant orbipropellant) and either common or dedicatedtankage and RCS thrusters/OMS engines.

Integral systems suffer, relative to modularsystems, in the areas of safety, ease of main-tenance, development flexibility, and growthpotential. Although attractive from a weightstandpoint, the above considerations are suf-ficient to eliminate integral systems fromcontention.

Figure 5-2 summarizes the relativeadvantages of wing and fuselage modularsystems. No clearcut preference is evident;weights are comparable, and no significant

WING TIP POD FUSELAGE POD

WEIGHT

CONTROL

10,133 LBM

[IMPULSE REQUIREMENTS MINIMIZED BY LARGE MOMENTARMS. (FINE CONTROL MORE DIFFICULT TO ACHIEVE).THRUSTER LOCATIONS MINIMIZE CROSS-COUPLING

10,400 LBM

BIAS DISTURBANCE TORQUES (YAW)

RESULT IN UNEQUAL PROPELLANTEXPENDITURE BETWEEN PODS FORON-ORBIT CONTROL

SAFETY/

MAINTENANCE

POD LOCATIONS FACILITATE REMOVAL/INSTALLATION

OPERATIONS. COMPONENT ACCESSIBILITY MOREDIFFICULT (PODS REMOVED) BECAUSE OF TOTAL

^ENCLOSURE - MORE ACCESS DOORS

POD ACCESS DIFFICULT BECAUSE OF POD

LOCATION OVER THE WING. COMPONENTSEASILY ACCESSIBLE WITH PODS REMOVED

THERMALCONTROL

UNCONTROLLED TEMPERATURE RANGE - 110°F TO

+ 165°F. MDAC-RSI AVERAGE UNIT WEIGHT = 2.34LBM; FT2

THRUSTER HEATERS - 5.4 WATTS/THRUSTERTANK HEATERS - 140 WATTS

DOORS REQUIRED OVER FWD-FACING THRUSTERS

UNCONTROLLED TEMPERATURE RANGE

-120°FTO+115°F. MDAC-RSI AVERAGEUNIT WEIGHT = 1.65 LBF/FT2

THRUSTER HEATERS - 5.4 WATTS/THRUSTERTANK HEATERS - 50 WATTS

AFFECT ONAERODYNAMICS

EFFECTS MINIMIZED BY. KEEPING PODS ON TOP SIDE OF WING (HYPERSONIC

STABILITY). MINIMIZING POD FRONTAL AREA (DRAG)TAPERING OF ELEVONS IS REQUIRED TO PROVIDESUFFICIENT WING STRUCTURE FOR POD SUPPORT

EFFECTS MINIMIZED BY

. MINIMIZING POD FRONTAL AREA (DRAG)

. POD BOATTAIL (BASE DRAG)

MINIMAL Jl EFFECTS ON CONTROL

SURFACES DUE TO AFT LOCATION OFCONTROL THRUSTERS

SEVERITY OF

DYNAMICENVIRONMENTS

UNSTEADY FLOW ON WING MAY INDUCE WING/PODRESPONSE THAT COULD RESULT IN FLUTTER HIGHACCELERATION AND PROPELLANT SLOSH LOADSLIKELY BECAUSE OF RCS LOCATION (REMOTE FROMVEHICLE CG)

MINIMAL EFFECT ON TANKAGE DUE TO

CLOSE PROXIMITY TO VEHICLE ROLLAXIS

FIGURE 5-2 WING-TIP/FUSELAGE POD COMPARISONBipropellant N204/MMH

5-2




technology concerns impact either concept.However, the wing modules do complicatewing design, and the forward firing thrusterprotection doors are unattractive. These con-siderations, coupled with the benefits asso-ciated with the design and development of aconsolidated propulsion system make thefuselage module concept the better choice.

Within a fuselage module concept, threeviable configurations remain: a dedicatedOMS,coupled with either a monopropellant ora bipropellant RCS, and a bipropellant RCSfor all maneuvers. For each system, dedicatedtankage is more attractive than commontankage due to lessened development risk.Based on the study criteria, the dedicatedOMS - bipropellant RCS is the most attrac-tive concept. However, cost considerations,

not included in this study could alter thisposition. The monopropellant RCS suffers asignificant weight penalty, but potentiallyoffers reduced development effort and,possibly, reduced maintenance requirements.Cost trades between reduced developmentcosts but increased operational costs (due tothe payload penalty) are necessary to definethe monopropellant RCS potential. TheRCS(OMS) is weight competitive with thededicated RCS-OMS configuration and, addi-tionally, would be less costly since it deletesthe costs associated with the OMS engine de-velopment. This concept suffers in compari-son to the dedicated RCS-OMS configurationsolely on the basis of its reduced flexibility tofuture increases in translational thrust re-quirements, e.g., potential future high thrustrequirements for ascent abort.

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6. REFERENCES

A. Kendall, A.S., McKee, H.B., and Orton,G.F., "Space Shuttle Low Pressure Auxil-iary Propulsion Subsystem Definition -Subtask A Report", McDonnell DouglasReport No. MDC E0303, 29 January 1971.

B. Green, W.M., and Patten, T.C., "SpaceShuttle Low Pressure Auxiliary Propul-sion Subsystem Definition - Subtask B Re-port", McDonnell Douglas Report No.MDC E0302, 29 January 1971.

C. Anglim, D.D., Baumann, T.L., and Ebbes-meyer, L.H., "Space Shuttle High Pres-sure Auxiliary Propulsion Subsystem De-finition Study - Subtask A Report",McDonnell Douglas Report No. MDCE0299, 12 February 1971.

D. Gaines, R.D., Goldford, A.I., and Kaem-ming, T.A., "Space Shuttle High PressureAuxiliary Propulsion Subsystem Defini-tion Study - Subtask B Report",McDonnell Douglas Report No. MDCE0298, 12 February 1971.

E. Orton, G.F. and Schweickert, T.F., "SpaceShuttle Auxiliary Propulsion System De-sign study - Phase A Requirements Defini-tion", McDonnell Douglas Report No.MDC E0603, 15 February 1972.

F. Orton, G.F., and Schweickert, T.F.,"Space Shuttle Auxiliary Propulsion Sys-

tem Design Study • Phase B Report,Candidate RCS Concept Comparisons",McDonnell Douglas Report No. E0567, 15February 1972.

G. Bruns, A.E., and Regnier, W.W., "SpaceShuttle Auxiliary Propulsion SystemDesign Study - Phase C Report, OxygenHydrogen RCS/OMS Integration Study",McDonnell Douglas Report No. MDCE0523, 15 June 1972.

H. Baumann, T.L., Patten, T.C., and McKee,H.B., "Space Shuttle Auxiliary PropulsionSystem Design Study - Phase D Report,Oxygen-Hydrogen Special RCS Studies",McDonnell Douglas Report No. MDCE0615, 15 June 1972.

I. Anglim, D.D., Bruns, A.E., Ferryman,D.C., and Weiland, D.L., "Space ShuttleAuxiliary Propulsion System DesignStudy - Phase C and E Report, StorablePropellants, RCS/OMS/APU IntegrationStudy", McDonnell Douglas Report No.MDC E0708, 29 December 1972.

J. Kelly, P.J., "Space Shuttle AuxiliaryPropulsion System Design Study - Pro-gram Plan", McDonnell Douglas ReportNo. MDC E0436, 15 July 1971, revised 6December 1971.

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