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CrewedLunarMissionsandArchitecturesEnabledbythe
NASASpaceLaunchSystems
BenjaminDonahueandSheldonSigmon
BoeingExplorationLaunchSystems,Huntsville,Alabama,35806
Abstract
TheNASASpaceLaunchSystems(SLS)outstandingcapabilitiesforlaunchingheavy,largediameterpayloads
willenablerobustHumanLunarandMarscampaigns.Lunararchitecturescurrentlybeinganalyzedinclude
bothmulti‐launchvehicleandSLSonlycampaigns.FortheLunarArchitecturetheNearRectilinearHaloOrbit
(NRHO)isus eda san agg regationnodeforth ela nde r,t ran sferstageandOrionelements.TheSLScapabilities
toLaunchLanderelementsisdiscussed,asisthestatusoftheSLS,andthenewlargeExplorationUpper
Stage,currentlyindevelopment,whichwilloptimizetheSLSCoreandBoosterStages.
I.
TheNASASpaceLaunchSystem
SLS consists of a series of increasingly
capable vehicles to incrementally
expand Beyond Earth Orbit (BEO)
exploration from lunar space and then
to Mars. The SLS Block 1 utilizes the
Interim Cryogenic Propulsion Stage
(ICPS) and the Block 1B features the
new,largeExplorationUpperStage
(EUS).TheICPSisaderivativeofthe
Delta‐IV upper stage; the EUS is in
development. Only the SLS can deliver
the 27.5 mt Orion Crew Vehicle to the
Moon; it delivers significantly more
payloadtoLEOandBEOdestinations
thananyotherexistingorplanned
launch system. Payload capabilities to
Trans‐Lunarinjection(TLI)are shownin
Fig.1fortheFalconHeavy(left),theSLS
Block1B,Block2andtheVulcan(right).
The Block 1B/EUS has a TLI capability
between39and43mt.Thelater2030’s
eraSLS Block 2would provide a53 mt
TLIcapability.Byenablinglargermargins
inthedesignofexplorationplatformsand
theabilitytosendmultiplecopiesofatmosphericandsurfaceprobes,higherresolutionspatialandtemporaldatacan
becollectedinasinglemission. Missionriskcanbereducedbyincreasing theredundancyofeachindividualsystem
andthearchitecturebyusingmultiplecopiesofthesamesystems.
II.
SLSCoreStageDevelopment
TheSLSBlock1isprogressingtowardaFY2020launch.MajorassemblesareinmanufactureandtestatNASAMichoud
AssemblyFacility(MAF)andacompleteCorestagewillentertestingatNASAStennisSpaceCenter(SSC)in2020.InFig
2theCoreStage’sforwardsection(Forwardskirt,LO2tank,andintertank)isshownjoinedtotheAftLH2Tank.InFig4
Figure1LaunchVehicleLunarPerformanceinMetricTons(mt)
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theCoreStageForwardSkirtisshown.TheCoreIntertankisshowninaNASAMSFCteststandinFig.5.
Figure2SLSCoreStageatMichoudAssemblyFacility(MAF)
Figure3InspectionofSLSBarrelSection
TheEngineSectionutilizesfourhighIspLO2/LH2RS‐25enginesprovidingover2millionpoundsofthrustatliftoff.The
fourRS‐25engineswillconsumeapproximately4,500lbofpropellantpersecond.
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Figure4SLSCoreStageForwardSkirtatMAF
Figure5SLSCoreStageIntertankStructureinTestatNASAMSFC
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III.
InterimCryogenicPropulsionStage(ICPS)
Theinitial Block 1 versionof the SLS usesthe Interim Cryogenic Propulsion (ICPS) Upper stage, whichis a close
derivativeoftheDelta‐IV5.0mdiameterUpperstage.TheICPS and Orion is supported by the Launch Vehicle
SpacecraftAdaptor(LVSA).TheICPSispoweredbyasingleRL10‐C1engine.ThefirstflightoftheBlock1isscheduled
forlate2020,andwillcarryanuncrewedOrionspacecraftaroundtheMoon.
Figure6SLSICPSUpperStageatNASAKSC
Fig.6SLSCoreStageLiquidOxygenTankStructuralTestArticleArrivingatMSFC
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IV.
ExplorationUpperStage(EUS)
TheEUSisthenextevolutioninthedevelopmentoftheSLS.TheEUSoptimizesthepowerfulSLSCoreandBooster
Stages and greatly increases the SLS injected payload capability. The EUS is a multi‐mission stage with
accommodationsformissionmodificationkitsandvariablepropellantloadingcapability;itisasuspended stage
poweredby4RL‐10enginesandincreasesthecapabilityofSLSvstheBlock1withtheICPS.TheEUSisroughlyfour
timesthemassandfourtimesthethrustofICPS.RelativesizesofupperstagesaregiveninTables1and2.TheEUS
isdesignedtoprovide43.0mttotheMoon,enablingrobust,sciencerichexplorationmissions.
Table1LargeUpperStagesComparison
Table2ICPSandEUSParameters
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TheEUSconsistsof7 major elements (Fig. 8), including theForwardSkirt,
LH2 Tank, Mid‐body, LOX Tank, Equipment Shelf, Thrust Structureand
Engines. These elements support an array of systems including avionics,
pressurization,RCS,mainpropulsionsystems(MPS)andsensors.
TheEUSForwardSkirtactsastheinterfacetotheIntegratedSpacecraftand
PayloadElementandtheEUSLH2Tank.Theforwardskirtstructureisan8.4
mdiameter,1.8mlongOrthogridbarrel.
TheEUSLH2Tankisusedtostoreuptoapproximately78,000gallonsofLH2
at cryogenic temperatures of −423˚ F (−253˚C). The primary structure is
madeoffrictionstirwelded,ellipticallyshapeddomesandOrthogridbarrel
section.Keepingthetankatcryogenictemperaturesandtopreventicing,the
tankiscoveredwithspray‐onfoaminsulation(SOFI).
TheEUSMid‐bodyiscomposedofanaftadapterandmetallicV‐strutswhich
providetheprimarystructuralconnectionbetweentheLH2andLOXtanks.
The Mid‐body is also the primary support for the vehicle pressurization
system,communicationshardwareandantenna.
The5.5 mdiaEUSLOX tankholdsup to 25,000gallonsof ‐297˚F(‐183°C)
LOX.Th eLOXtan kisalso theinterfacefo rthethruststru cturewhichattaches
tothedomeandanequipmentshelfwhichattachestotheaftflangeofthe
tank.
The EUS equipment shelf is a honeycomb panel,ortho‐grid plate and bracketed design to support the avionics and RCS
structuralneeds.TheshelfisattachedtotheLOXtankflangebystruts.Avionicscomponentsthataresupportedontheshelf
areflightcomputers,guidanceandnavigation,powerdistributionsystemandRL‐10enginecontrols.Propulsioncomponents
supportedincludetheRCSanditshydrazinepropellanttanks.
Figure8SLSExplorationUpperStage
Figure9SLSICPSandEUSTLICapabilityBreakdown
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TheEUS ThrustStructureis anI‐beamand strutconstructionattachedtothe LOXtankdome.Thestructureistheprimary
structuresupporting4RL10engineswhichprovideacombinedthrustcapabilityof99,000lbf.
TheInterstageattachestheEUStotheCoreandsupportstheEUSduringCoreflight.Thisstructuremustsupporttheweight
ofeverythingontopoftheCoreStage,whichcouldbeinexcess180mt.Theinter‐stageisanisogridpaneldesign;contained
withitistheseparationsystemthatutilizesapushersystemtoseparatetheEUSafterCoreStagemainenginecutoff(MECO).
Fig.10containsacomparisonofLaunchcapabilitiestoTLI.TheSLSprovidessignificantlymorepayloadtothemoonthanany
othervehicle. The Block 1BprovidesasignificantimprovementoverBlock1. The 2030’s era Block2willfacilitateCrewed
Marsmissions.InFig12SLSFairingsizesarelisted.InFig.13aLarge,8mdia.monolithicopticSpaceTelescopeisillustrated.
Figure10Trans‐LunarInjection(TLI)PayloadCapabilitiesforEightLaunchVehicles
Figure11Block1B:IllustrationofOrion/USA/EUSatIgnitionjustafterCoreStageJettison
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Figure12SLSBlock1Band2FairingSizesandVolumes
Figure13SLSEUSand8.0mdiaVeryLargeSpaceTelescopeafterFairingJettison
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V.
NASAHumanLandingSystem(HLS)Architecture
NASA’sreturntotheMoonplaninvolvesaggregatinglanderandtransferstage elements in a Near Rectilinear Halo
Orbit(NRHO).TheretheOrionwillhostthelunarlander.InsomeplanningscenariostheOrionissupportedbyaNRHO
Gatewayplatform (Fig. 14 middle); the Gateway would consist of (ata minimum) a habitation module, Powerand
Propulsion Element (PPE) and a Logistics Module. In addition to the Gateway, thisarchitecture would include four
transportationelements;theAscentElement(AE)(topleft,Fig14),theDescentElement(DE)(middleleft),theTransfer
Element(TE)(bottomleft)and,forlatermissions,aRefuelingElement(RE)(notshown).
In this scenario, the 4 elements are flown out separately and aggregatedinNRHO.Onceready,thecrew
transfersfrom the Orion to the Lander, and the3stage(AE/DE/TE)combinationtransfersfromNRHOto Low Lunar
Orbit(LLO)withtheTEprovidingthedelta‐Velocity(dV).OnceinLLO,theTEseparatesandreturnstoNRHO.Thecrew,
intheDE/AE,descendtothesurface.Afterasurfacestay,theAEascendsdirectlybacktoNRHO.TheREmaybeused
intwoways;first,toreplenish(topoff)theDEtanksinNRHO;andsecond,torefueltheAEandTEforreuse.Foreach
newlunarsortiemissionanewDEandREarelaunched.
IntheNASAdiagram(Fig.14),threetypes of launchers inject theAE,DE,TEandREintoTLI.Includingthe
Orion,thereare5transportationelementsthat operatefromNRHO inthisarchitecture.This scenariocarrieswith it
somedegreeofcomplexityand,toreducethecomplexity,asimplifiedversionofthisarchitecturewasdevelopedand
ispresentedhere.ThesimplificationkeepstheNRHOasastaginglocationandtheaggregationofelementsthere,but
reduces the number of transportation elements, the number of launch vehicle types and the number of launches
required.Forthissimplifiedarchitectureacrewedlunarlanderconceptwasdeveloped.
InFig.15,NHRO orbitisdepicted andmissiondVsarelisted foreachlegofthejourneyfromtheEarthtotheMoon.
AfterinjectionintoTLI,transferdVtotheNRHOis458m/s.Afteraggregationofelements,the(1daytransit)NRHO‐
to‐LLOtransferrequires750m/sdV.ThedescentfromLLOdVis2,100m/s.(Totalone‐way,post‐TLItosurface(through
NRHO)is3,308m/s).AEascenttoNHROrequires2,700m/s.TotaldVis6,008m/s,splitamongTE,DEandAEelements.
As mentioned, the HLS architecture depicted in Fig. 14 is a “3 launch” scenario. For this architecture, In NRHO,
Figure14NASAHLSArchitectureDiagram.TransportationElements(left),LaunchedSeparately,
areAggregatedinNearRectilinearHaloOrbit(center),ThenProceedtotheSurface(right)
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propellant transfer is required to
replenish (top off) the DE tanks.
Thisis because neither of thetwo
non‐SLS Launch vehicles (Fig. 14)
are capable of injecting a fully
fueled DE to NRHO. (Early
estimates of the DE mass place it
over16mtfullyfueled).Thusthe
DE (in this scenario) must be
launched with a partial propellant
off‐load to remain within the TLI
capabilityof thelauncher.Onceat
NRHO, the DE is and RE must
autonomously rendezvous, and
oncepropellant line interconnects
aremadeandverified,propellantis
transferred into the DE’s partially
filledtanksuntiltheyarefully
fueled.
VI.
SLS“2Launch”SimplifiedNASAHLSArchitecture
ThesignificantlyhigherTLIlaunchcapabilityoftheSLSwillenablethethreeprimaryHLSElements(AE,DEandTE)to
belaunchedin2launchesratherthan3,andwillallowtheDEtobelaunchedfullyfueled.This SLS“Twolaunch”scenari o
eliminatesoneLaunchperLunarSortiemission,andeliminatestheneedtotransferpropellanttotheDEatthegateway.
Itwouldalso,eliminate,or postpone, the necessity for aREfortheinitial series of missions. Once the Architecture
reachesits“sustainability”stage,theREwouldbebroughtintorefueltheAEandTEattheNRHOtoallowthemtobe
reused.In Fig. 16the EUS, Orionand SLS Lander are illustrated. InFig 17, a SLSlaunchedAE/DE lander conceptis
illustrated(left,inthe8.4mdiafairing).ThisAEconcepthasasingleenginelocateddirectlyunderneathitscylindrical
Ascent Cabin (F ig.17 ‐18).Similarly ,theSLSlau nchedDE conceptalsoutilizesasingleengine.Centrallypositioned,single
engineconfigurationsarelesscomplexthanmultipleenginedesigns;placingasingleengineatbottomcenterensures
thrustisdirectlybelowthevehicle’scenterofgravityatalltimes.Otherconfigurationsareunderstudy,includingthose
thatusemultipleengines;eitherasclusterscentrallypositioned,orindividualenginespositionedaroundtheperimeter
ofthestage.Forthisconcept,theAE’senginenozzleisembeddedinthetopoftheDE.AtthebottomcenteroftheDE
isitsdescentengine.Asshown,theSLSlaunchedAEconcepthasfoursphericalsidemountedtanksattacheddirectly
totheCrewCabin(Fig18).TheDE’s4tanksarepositionedtoallowtheaccommodationoftwoside‐mountedpayload
pallets(Fig16,17)whichallowforeaseofdownloading.Pallets,attachedatahingepointnearthebottom,arerotated
Figure15NASAHLSArchitecture
TrajectoryPhasesandMission
Delta‐Velocities
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downtoahorizontalposition,foreasyaccesstopayloads.Thelarge8.4mdiafairingprovidesthewidthtomakethis
sideplacementpossible;ifconstrainedto5.0mfairing,sidemountingcouldnotbereadilyachieved.Oneofthepallet
positionsmayb eoc cup ied bya rover.TheSL SBlock 1B“ Two Launch”conceptLanderisshowninFig.16(right).Because
neitherof the elements (AE, DE or TE) areconstrainedtofitwithinasmaller5mdiafairing,their geometry can be
optimized;thelanderforeaseofsurfacepayloadoff‐loadingandlowstageheight.ShowninFig.17aretheAEandDE;
theAECrewcabinisaverticalcylinderwithadockingportontop,anda singleengineonthebottom.TheAEengine
nozzlefitsdownintothetopportionoftheDE.TheDE’sbottomcenter engineissurroundedby 4 cylindricaltanks,
structureandpayloads.TheLanderisshown with landing legs stowed (Fig.17 left), and (right)deployed.TheAEis
illustratedinFig.18;itsengineisimmersedintothecabin.TheDEcanalsobeflowninacargoonlymode.Inthatcase
anadditionalpayloadispositionedontopinplaceoftheAE.Thatpayload might be a surfacehabitat,solarpower
station, nuclear surface power unit, insitu propellant plant, powerbeamingstation,‘FarSide’Astronomyassetor
mining/excavationmachinery.TheSLSBlock1BCargoversionTLIcapabilityis43.0mt;thiscapabilityissufficientto
launchHLSlanderelements,withexcesscapabilitytocoverweightgrowth,increasesinpayload,increasesincrewsize,
orstaytime.
Landersizinganalysisusedthefollowingassumptions:
1) AECabinsizedforCrewof3,butcarriescrewof2initially,5daysurfacemission,50kgsamples
2) Crewcabin3.25mt,plus0.45mtforcrew,consumablesandonboard(AE)scienceequip.
3) DEsurfacepayload0.5mt,SeparateAirlock,mass0.85mt(including2EVAsuitsandsupplies)
4) Propellantresidualsandflightperformancereserves(FPR)of4%,Drymassmargins15%
5) Cryocoolersarecarriedtopreventboiloffforthesoftcryogenoptions(LO2/LCH4)
6) Hepressuranttanks:CompositeOverwrapPressureVessel(COPV)4,000psia
7) OutboundtoNRHOis5days,NRHOtosurface1day,surfstay5days,NRHOreturn1day
8) AEthrustrequired7,500lbf,DEthrust19,200lbf,pervehicleT/Wof1.8(local)atignition
9) RCSstorablebipropellant290secIsp,AERCSthruster50lbf,DEthruster100lbf
10) SLSBlock1BCargoinjects43.0mttoTLI,Block1BCrewinjects12.5mttoTLIinadditiontoOrion.
TheTE(notillustrated)willrideasaco‐manifestedpayloadwiththeOrionintheCrewSLSlaunchandtravelswiththe
OriontotheNRHO.Theun‐crewedlanderisthenlaunched,andtransferstoNRHO.There itwillbejoinedtotheTE;
oncecheckedout,thecrewwillboard and the Lander/TE combination will separate. Lander propulsiontradestudy
Figure16SLSTwoLaunchArchitectureTransportationElements(TransferElement(TE)notshown)
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resultsaregiveninSectionXII,whereSLSlaunchedAE,DEandTEmassesareexplainedandlaunchmarginsaregiven
forseveralvariationsofthelander.TheApolloLunarExcursionModule(LEM)andthe2008NASAConstellationprogram
AltairlanderconceptarereviewedinSectionsVIIandVIII.CandidatelanderenginesarereviewedinSectionsIXandX.
TheseSectionsset the context forthe SLS lander trade studyforthesetofassumptionsandgroundrulesthatare
appropriateforaNRHObasedLunarsortiearchitecture.
Fig.17SLSTwoLaunchCrewedLunarLanderConcept.TheSLSwide8.4mdiaFairingallowstheDescent
ElementsGeometrytobeoptimizedforEaseofCargoOff‐loadingandLowStageHeight.
Figure18AscentElement(AE)Configuration
VerticalCylinder,withSingleImmersedEngine.MoreDetailsonpage20.
InternalVolumeVarieswithCrewSizeandStayTime.
TheSLS Lander is designedtothelarge8.4 m diafairingandtakesadvantageofthatwidediametertolocatethesurface
cargoattheouterperimeteroftheDE,eitherastwosidepallets(Fig.17)oras4smallerpayloadbays(notshown).Ineither
option,surfacecrews,atgroundlevelwouldhaveeasyaccessto thecontents.LimitingaLandertoasmaller 5.0mfairing
mightpreventsidemounting.ThetwincargopalletscontainequalmassesofpayloadtokeeptheLanderCenterofGravity
(Cg)centered.Payloadelementscouldbeloadedontothelanderin NRHO;oroff‐loaded fromtheLandertotheNRHOas
missionneedsmaydictate.ConfiguringtheDEforasingle,centerenginefacilitatesthedualpalletcargoapproach.
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VII.
ApolloLunarExcursionModule(LEM)
TheApolloCommandModule(CM)andLunarExcursionModule(LEM) areillustratedinFig.19; theLEMfortheApollo14
missionmassed15.3mt.TheLEMDEprovidedonlythedescentfromLLOtoSurfacedV.TheLEMwascapturedintoLLOby
theCM Service Module. Fig. 20 lists LEM DE, AE andRCSEngineinformation.Boththe LEM and CM ServiceModuleused
storablepropellants(NTO/AZ50).Isp’swere311secfortheLEMAscentElementand305secforDescent.
Figure19ApolloCommandModuleandLunarExcursionModule(LEM)
Figure20ApolloLEMEngineInformation
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VIII.
2008NASAConstellationProgramAltairLanderConcept
NASA’s2008‐09 Constellation Program Altair Lander Concept is illustrated inFig.21.TheAltairDEfeatureddeep cryogen
(LO2/LH2)propellants; it’sAEused storablepropellants.Due toverylowdensityLH2,coupledwiththefactthattheAltair
wasrequiredtodotheLunarOrbitInsertion(LOI)burn(capturingbothitselfandtheOrion),itsDEwasverylarge.Totalmass
oftheAltairwas45.6mt,3timestheLEM(Fig.22).Altair’sDEfeatured4RL‐10engines,8proptanks,ajoinedtrussstructure
andmassed38mt.Altairwouldhavestood3storieshighmakingcrewegressachallenge.TheCargoonlyversionlocatedits
payloadsontop;off‐loadingthesetothesurfacewouldhavebeendifficult.Boeingevaluatedalternativestothisconfiguration
andpublishedsomeofthefindingsinRef.2,3and4.
Figure21NASAAltairLanderConceptIllustration
Fig.22SizeComparisonAltairandApolloLunarExcursionModule(LEM).
TheLEMuseddenseStorablepropellant;Altairtheleastdenseofpropellants(LH2).AltaircapturedbothitselfandtheOrion
intoLowLunarOrbit(LLO)beforedescenttotheSurface;theLEMwascapturedintoLLObytheApolloServiceModule.
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IX.
PumpFedStorablePropellantLanderEngineCandidates
Aerojet‐RocketdyneXLR‐132andAestus‐IIenginedataarelistedinFig.23.Bothenginesuseddensestorablepropellant,but
unlike the LEM engines, they feature a higher Isp pump fed cycle.ARXLR‐132,demonstratedinflight,andAestusII,
demonstratedintest,bothachieved340Isp,asignificantimprovementvsthe315secpressurefedusedbyApollosystems.
Fig.23PumpFedStorableEngineCandidates(AestusIIupperleft,XLR‐132bottomright)
X.
MethanePropellantLanderEnginePerformance
TheAerojet‐RocketdyneRS‐18LO2/CH4(methane)pumpfedengineispicturedontheteststandinFig.24.Expandercycle
methaneenginesarecapableof375secIsp,anotableimprovementoverstorablesystems.Methane,asoftcryogen,isless
densethan mono‐methylhydrazine(MMH)but significantly moredensethanLH2. In Table3,variationof Isp withEngine
ChamberPressure(Pc)islisted.Isp’sof368,371and375scorrespondtoPcof600,750and1000psia,forfixednozzleexit
diameterof65inches,andthrustlevelsof16,000lbf.InthisanalysisaconservativeIspof365secwasused,corresponding
toaPcof600psia;arelativelylowPc,identicaltothePcoftheLO2/LH2RL‐10engine.
Table3PumpFed,ExpanderCycleMethaneEnginePerformance
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XI.
SLSLaunchedLanderConceptPropulsionTradeStudy
Asizing evaluation was done foraprospectiveSLSlaunchedCrewed Lunar lander and TE appropriate for aNRHOcrewed
lunarprogram.Thetradestudycomparesfivepropellant/enginecombinationsfortheAE,DEandTE.TheSLScargolaunch
(Block1B, with fairing)injectstheLander (integratedAEandDE)toTLI;theDEprovidesthedvfortheoutboundtransfer
NRHO.That launch is followed by a secondSLS Crewed Block 1B,which injects theOrion and theTE as a co‐manifested
payload,also to NRHO. The Orion’s Service Module (SM) provides dv for the combination on the outboundtrip.Allthree
transferelementswereevaluatedwitheachofthe5propellantcombinations.WhatistobedeterminedintheTradeStudy
arethemassesoftheAE,DEandTEasafunctionofthefivepropellantselections(Table4),aswellastheSLSlaunchmargins
foreachofthesecombinations.Twoofthefivearepressure‐fedoptions,withpropellanttankpressuresof270psia;threeof
thefivearepumpfedoptionswithtankpressuresof40psia.EngineSpecificImpulse(Isp)rangefrom320‐385sec.
Figure24MethaneEngineinTest2009
Table4FivePropellant/EngineCombinationsforLanderTradeStudy
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Forthemissionscenariosgiveninthisreport,theTEdoesthetransfertoLLOburnstoplacetheLanderintoLLO,butitdoes
notcarryanyreturn‐to‐NRHOpropellant;itoperatesinanexpendablemode.Inotheranalysis(notincludedinhere)theTE
carriesthemodestadditionalpropellantneededforittodepartLLO,byitselfandreturntoNRHO.
XII.
SLSLaunchedLanderMassasaFunctionofPropellantandEngineOptions
TheobjectiveofthelanderandTEevaluationistoestimatestagemassesandtodeterminehowmuchSLSlaunchmarginis
available.VehiclemassesarelistedinTable6.DataisincludedfortheAE(row10),theDE(row11),landersummary(sumof
AEandDE,row12)andTE(row16).Thesemassesincludeallpayloadsand airlockmass.Thesummed AEandDE(lander)
massesarecomparedtotheavailableSLS Block1BCargoTLIinjectionmass(43.0mt,row14).Thedifferenceinthesetwo
valuesandthepayloadattachfitting(PAF,adaptor)isthe“launchmargin,”i.e.thesurplusmassavailableinthesystemfor
launchingthelander.A negativelaunchmargin meansthelandermassisgreaterthantheamounttheCargoBlock1Bcan
injecttoTLI;marginsarelistedinRow14. AlsotheTE massiscomparedtotheavailableBlock1BOrionco‐manifestedTLI
injectedmass(12.5mt,bottomrow),thesemarginsaregiveninrow18.
Table6containsdataforsevenlanderandTEcombinations;thefirst5columnscorrespondtothe5propellantsdescribedin
Table4.(Ineachofthese5,propellantandenginechoicesarethesamefortheAE,DEandTE).Thesixthandseventhcolumns
haveAE’s and DE’s with differentpropellants(TEis same as DE).Column 6 data is forapressure‐fed storable propellant
(“storable”)AEcoupledtoapump‐fedmethane(CH4)DEandTE.Theseventhcolumnappliestoapump‐fedstorableAEand
amethaneDEandTE.Resultsindicatethatfor4ofthe7options,significantSLSlaunchmarginisavailable.Columns2,4,5
and7ofTable6showmarginsof8.4,9.1,12.1and8.7mtrespectively;landermassforthese4casesare33.7,32.9,29.99,
and33.2mt.Thesystemchosenasareferenceamongthese5optionsisthe340sIspstorablepump‐fedoption(Col.2Table
6).UsingXLR‐132demonstratedtechnologyanddensestorablepropellants,theSLSBlock1Bcanlaunchthisreference33.7
mtlanderwithalaunchmarginof8.4mt(Col.2,Fig.26‐26)).FortheSLSOrionlaunch,aStorable9.4mtTEiscarriedwitha
marginof2.76mt.Thehighestperformingofthe7optionsisthestagedcombustionmethaneoption;withalanderof29.99
Table5MissionandVehicleParameters:ApolloLEM,AltairandSLSLaunchedLanders
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andTEof7.75mt,availableSLSmarginsare12.1and4.3mt(Col.5). Another high performing option isexpander cycle
methane(Col. 4); which achieves a landerof32.9mt,alaunchmargin of 9.1 mt, and a TE mass of 8.86 mtwitha3.4mt
margin.Although providing slightly less margin than the two pump‐fedCH4 systems, the storable option would eliminate
boiloffconcerns.(Forallmethaneconcepts,activecryocoolersandadvancedcryogenicfluidmanagement(CFM)systemsare
carriedtominimizeboiloff).Fig.25listsmassforTE,DEandAEofthe340secIspStorablesystem.InFig.26theAEisshown
atliftoff; also listed is ‘as launched’ massfor the referencepump‐fed Storable option (Col.2), including surface payload,
airlock,inertandpropellantmasses.ThedataindicatestheBlock1B/EUSenablesthetwolaunchlunarmissions.
XIII.
ArchitecturalLevelFindings
The‘TwoLaunch’SLSGatewayarchitectureisrobust,i.e.hassignificantmarginformassgrowth,increasesinpayload,crew
size,staytimeorsomecombinationofthese,andsimplifiestheHLSarchitecturebyreducingthenumberoflaunchesfrom3
to2.ThisArchitectureisalsorobustinthesensethatitcandeliverallelementsfullyfueled.
Forthecaseofthe‘3Launch’architecture(Fig.14),theDEisflownseparatelyinasmallerlauncherthatiscapableof15mt
toTLI(Fig.1left).EarlyindicationsarethattheDEforthe‘3Launch’architecturewillmassmorethan15mt,andthustheDE
mustbelaunchedoff‐loaded,requiringalaterrendezvouswiththeRefuelingElement(RE),forpropellanttransfertotopoff
its tanks. Listing the total number of elements in the ‘3 launch’ architecturewould begin with the (at least 3) Gateway
elements,ifitisutilized–(habitat,PPEandlogisticsmodule);theOrion;andthe4separatelylaunchedTransferelements(if
theREisincluded),distributedamong3differentlaunchers.Thetotalamountsto8individualelementsandprocurements.
Tosummarize,theTwoLaunchSLSapproachwouldsimplifythearchitecturebyeliminatingtherequirementforpropellant
transfer,reducingthenumberof launches, while providing robustness to massgrowth, crew size, stay time, andmission
evolution.Eitherthepumpfedstorable(Col.2)orpumpfedexpandercycleMethaneoption(Col.4)wouldprovidesufficient
performance;thestagedcombustionmethane(Col.5)providesthelowestmassoftheoptionsconsideredhere.Thepump‐
fedstorablesystem(XLR‐132technology)maybepreferredasitwouldeliminatethedevelopmentofactiveCFM.
Table6MassSummariesforSLS‘TwoLaunch’ArchitectureLanderPropulsionTradeStudy
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Figure25TE,DE,AEmasses(lefttoRight):Propellant,Inert,CrewCabin(AE)andSurfacePayload(DE)
StorablePump‐FedPropulsion.These,alongwithOrion,arelaunchedwith2SLSBlock1Blaunchers.
Fig.26AscentLiftoff(left),andAE,DE,TEMassStatement(right).AECabinSizedfor3,Carries2forthismission.
StorablePump‐FedPropulsionOption.SLSLaunchMargins(bottomright)
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AlternateAscentStageConfigurations
FourconfigurationvariationsoftheAEareillustratedinFig.27.Thevariationsinvolveengine,tankagepositioningandcrew
cabincylinderorientation.Fig.27farleft,showsaverticalcylindercabin,withasingleengineand4verticallystackedtanks;
leftmiddle,ahorizontalcylindercabin,singleenginewith4horizontallyalignedtanks;middleright,ahorizontalcylindercab
withasingleimmersed(slotted)engineand4verticaltanks;farright,ahorizontalcylindercab,4sideengines(2arevisible)
and4horizontaltanks.TheReferenceAEshownearlier(Fig.17‐18,26)andinTable8,consistsofahorizontalcylindercab,
singleimmersedengineand4verticallystackedtanks.AEmassesarelistedinTable8.Thisisa‘minimum’Crewcabinasitis
sizedforacrewof3fora5daysurfacestaytime,initialmissionscarry2crew.TheAEascendsthroughatotalof2,890m/s
dVtoreachNRHO,andhaspropellantmargins(unusablesandflightperformancereserves)of4%.
Table
7
LanderandTransferElementPropulsionTradeStudySummary
Fig.27LanderCrewedAscentElement(AE)ConfigurationVariations
Table8
A
E:Sizedfor3,Carries2,5DaySurfaceStay,StorablePump‐fedPropellant,AscendstoNRHO
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Conclusions
TheNASASLSFamilyofLaunchVehiclesprovideagame‐changingcapabilityforBeyondEarthOrbitExplorationMissions.The
EUS Upper Stage is sized to optimize the powerful SLS Core and Booster stages, providing a significant improvement in
performanceovertheinterimICPSupperstage.WiththeEUS,theSLSBlock1Bcan providerobustcrewedlunarmissions.
TheSLSBlock1BcandeliverallprospectiveNASAHLSLanderelementsintwolaunchesratherthanthree,whilemaintaining
launchmarginformassgrowth,increasesinpayloads,andmissionevolution.
References
1. NASA,SpaceLaunchSystem(SLS)MissionPlanner’sGuide,ESD30000InitialBaseline,ReleaseDate:04/12/17
2. Benton,M.,Caplin,G.,Reiley,K.,Donahue,B.,Messinger,R.,Smith,D.,BoeingDesignTradesinSupportoftheNASA
AltairLunarLanderConceptDefinition,AIAA2008‐7798,AIAASpace2008Conference,SanDiego,CA,9Sept2008
3. Donahue,B.,Grayson,G.,Caplin,G.,Reiley,K.,Smith,D.,LunarLanderAscentModuleConfigurationandPropulsion
Study,AIAA2009‐6406,AIAASpace2009Conference&Exposition,Pasadena,CA,14Sept2009
4. Donahue,B.,Caplin,G.,Smith,D.,Maulsby,C.,Behrens,J.,LunarLanderConceptsforHumanExploration,AIAAJournal
ofSpacecraftandRockets,Vol.45,No.2,March‐April2008