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https://ntrs.nasa.gov/search.jsp?R=19930013818 2018-05-19T04:22:01+00:00Z
Space Transfer Concepts and Analysesfor Exploration Missions
NASA Contract NAS8-37857
Lunar Transportation FamilyImplementation Plan and Element Description Document
Boeing Aerospace and ElectronicsHuntsville, Alabama
" G.R. W'obdcdck "r "%STCAEM
Project ManagerBoeing Aerospace and Electronics
D615-10026-6 1
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Space Transfer Concepts and Analysesfor Exploration Missions
NASA Contract NAS8-37857
Lunar Transportation FamilyImplementation Plan and Element Description Document
Boeing Aerospace and ElectronicsHuntsville, Alabama
Docamaentation Set:
D615-10026-1 IP and ED Volume 1: Major Trades, Books 1 and 2D615-10026-2 IP and ED Volume 2: Cryogcnic/Acrobrakc Vehicle
D615-10(T26-3 IP and F_X)Volume 3: Nuclear Thermal Rocket VehicleD615-10026-4 IP and El9 Volume 4: Solar Electric Propulsion Vehicle
D615-10026-5 IP and ED Volume 5: Nuclear Electric Propulsion VehicleD615-10026-6 IP and ED Volume 6: Lunar Transportation Family
D615-10026-6 2
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m
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Lunar Transportation FamilyImplementation Plan and Element Description Document
Table of Contents
Section Paee
Cover Sheet .................................................................................... 1
Title Page ....................................................................................... 2
Table of Contents .............................................................................. 3
Symbols, Abbreviations and Acronyms .................................................... 4
II
II.
IVQ
V.
Evolution of the Concept ............................................................ 11
A. Reference Concept Development .......................................... 11B. Alternate Concepts .......................................................... 51C. Architecture Mawix .......................................................... 65
Requirements, Guidelines and Assumptions ..................................... 101Ao Level I, H and HI Requirements .......................................... 101B. Derived Requirements ....... :. ............................................. 109C. Assumptions ................................................................. 115
Mission Operations .................................................................. 119A. Mission Analysis and Performance Parameu'ics ........................ 119
B. Operating Modes and Performance ....................................... 139
Element Descriptions ................................................................ 153A. LTV/LEV Components ..................................................... 153B. Habitation Modules ......................................................... 163C. ACRV Modifications ....................................................... 205
D. Aerobrake .................................................................... 209E. Cost ........................................................................... 223
Commonality/Evolution ............................................................. 269A. System Commonality ....................................................... 269B. SmaUfMecLium/Large Scale Evolution .................................... 297
D615-10026-6 3
Symbols,
ACRVACSAFE
A&I
AI
ALARA
ALSALSPE
am
AR
ARGPER
ARSart-gasc
ASE
AU
BITBITEBLAPBFOBlVIR
CCABCAD/C.AMCAP
C_CELSS
CHCCG
cLctn
c/InCM
c/oCofF
conjCOSFAR
CO2
CryoC3
C&T
CTV
d
DDT&EDE
dogdcsc
Abbreviations and Acronyms
Advanced crew recov_ vchickAttitude control systemAcmbrake Flight _ntAttachment and inzgrationAluminum
As low as reasonably achievableAdvanced Launch SystemAnomalously large solar proton eventAtomic mass (unit)Arcaratio
Argument of p_g_
Atmosphcric revitalizationsystemArtificial gravityAscent
Advanced space engineAstronomical Unit (=149.6 million kin)
Built-in test
Buih-in t_st equipmentBoundary Layer Analysis ProgramBlood-forming organsBody mounted radiatm"
Dcgr_s Celsius
Cryogcaic/aezobraimComptcr-aide.dd_sign/computcr-aidcdmanufacturing
Cryogenic all-propulsive
Drag cocffaciemClosed Environmental L£fc Support SystemCrew health care
Center of gravityLift coefficientCentimeter = 0.01 meterOcw moduleC.cnter of massCheck out
Cost of facih'tics
ConjunctionCommisvc on Space Rcscm'ch of theIntomafionalC.ouncilof ScientificUnionsCarbon dioxid_
CryogenicHyperbolic excess velocitysquatw.d(inkm21s2)
Communications and Tclcn_try
Cargo Transport Vehicle (operates in Earth orbit)
yssign, dcvclopmcnt, testing, and evaluation
Dose _uivakmtDcgrccsDescent
"V
D615-10026-6,¢
DMS
dV
EAEarrEcECCVECWSECLSSEPESA
¢.S.O.ETETOEVA
FD&DFewFEL
FfFraFiFIFnFo
FrsFSE
FsFss
FuFvFY88
gGCNRGCRGEOGN2GN&C
GPS
Gy
habHDHEIHLLVhrs
Data mana_mcnt systemVdocity change (AV)
Earth arrivalEarth arrival
Modulus of elasticity in compressionEarth crew capture vehicleElement control work station
Environment controland lifesupportsystem
ElcctricpropulsionEm'opean Space Agency
Engine startopportunityExternalTankEarth-to-orbit
Extra-vehicularactivity
CkculationefficiencyfactorFire Detectionand Differentiation
Lifesupportweight factorFirstclement launch
Specificfloorcount factorSpecificfloorareafactor
Acrobrakc i.nmgrafion factorSpecificlengthfactorNormalized spatial unit count factorPath options factorUsefulperimeterfactorPartscount factor
Proximityconveniencefactor
Plan aspectratio factorSectionaspectratiofactor
FlightsupportequipmentVaultfactor
Safe-haven splitfactor
Spatialunitnumber factorVolume rangefactor
FiscalYear 1988 (=October I,1987 toScptcmbcr 30, 1988.
otheryears)
Accelcrauon inEarth gravities(=acceleration/9.80665m/s2)Gas core nuclear rocket
Galacticcosmic raysGcosynchronous Earth Orbit
Gascous nitrogen
Guidance, navigation,and controlGlobal PositioningSystem
Gray (SIunitof absorbed radiationenergy = 104 crg/gm)
Habitation
High DensityHuman Exploration Initiative (obsolete for SEDHeavy lift launch vehicleHours
Similarlyfor
D615-10026-6 5
hyg wHZEH2
H_
ICRP
IMLEOin.
inb_&ED
mgD
IspISRU
JEMJSC
kkeV
kgldbldbfkmKM
KM/SceKM/SECksi
LCCLIDLD
LDMLEO
LETLEV
LEVCM
Level IILH2
laOH
LLOLM
LOR
LOXLS
LTV
LTVCM
1.2
m
[M_Cnam[MARSlNMASEMAV
Hygeinc wamr
High atomicnumber and energyparticleHyarogenWater
InternationalCommission on RadiationProtectionInitialmass inlow Earthorbit
InchesInbound
ImplementationPlanand Element I_scriptionIndcl_ndant research and devclopmmatSpecific impulse (=thrust/mass flow rote)In-situresourceutilization
Japan Experiment Module (of SSF)Johnson Space C,cntcr
klb
Thousand e_ volt
Kilograms
Kilopounds (thousandsof pounds. Conversion to SI units=4448 N/klb)
Kilopound forceKilometers
Kilometers
Kilometerspea"second
Kilometersper second
Kilopounds per squareinch
Lifecyclecost
Lift-to-drag ratio
Low densityLong duralionmissionLOw Earth orbit
Linear energy transferLunar _cL_ion _hicleLunar excursionvehicle crew module
Space Explcranon Initiative project office, :Johnson Space CenmrLiquid hydrogenLi_um hydroxideLow Lunar orbitLunar ModukLunar orbit rendezvous
Liquid oxygenLunar surfaceLunar transf_ vehicleLunar transfer vehicle crow module
Lagrange point 2. A point _hind the Moon as sccn from the Earth whichhas the same orbitalperiodasthemoon.
Memrs
Western Union interplanetarytelegram]
Martianpornography]
Mission analysisand systems engineering(same asLevel IIq.v.)Mars ascentvehicle
D615-10026-6 6
M/CDAMCRVme_OP
IV_V
MOC
MOI
modM&P
MPS
MR
ml_MSFC
mt
natMTBF
_e
m 3
N
NASANCRPNEPNERVANTPNSONTRN204
OSEOTISoutb02
PBRPcPEEK
PEGAP/LPOTV
pot wPPU
proppsiPV
Ballistic coefficient (mass / drag coefficient times area)Modified crew recovery vehicleMass of electron
Maximum expected operating pressureMillion electron voltMars excursion vehicle
Muiti-layer insulationMillimeter (=0.001 m_ter)MonomethylhydrazineManned Mars vehicle
Mars orbitcaptureMars orbitinsertion
Module
Materials and processesMain propulsion systemMixtme ratio
Meters per secondMarshall Space Flight CenterMillionpounds per square inchMelric tons (thousands of kilograms)Metric tonsMean time between failuresMars transfer vehicle
Megawatts electricCubic Meters
Newton. Kilogram-meters per second squaredNot applicableNationalAeronauticsand Space AdministrationNationalCouncilon RadiationProtection
Nuclear-electricpropulsion
Nuclearengineforrocketvehicleapplication
Nuclear thermalpropulsion(same as NTR)Nuclear safeorbitNuclear thermal rocket
Nitrogen tetroxide
Orbital support equipmentOptimal Trajectoriesby ImplicitSimulation programOutbound
Oxygen
Particle bed re.actor
Chamber pressu_Polyether--cthcr ketonePowered Earth gravity assist
PayloadPersonnel orbital transfer vehiclePotable water
Power processing unitPropellantPounds per square inchPhotovoltaic
D615-10026-6 7
RAANRCSReRFRMLEOROIRPMRWAR&D
SAASAICSEISEPSISiCSMAsolSPESRBSSFSSMESTCAEM
stgsurfSv$1$2$3
t*
TBDTcTCSTEITEISt.f.THCTMITMISTPSTr&c
T/W
UN-W/25Re
VABVCSVinf
Heat flux (Joules per square centimeter)Radiation quality factor
Right ascension of ascending nod_Reaction control systmnReynolds numberRadiofrcquc yResupply mass in low Earth orbitReturn on invcsmmnt
Rcvolutiom per minut_Relative wind angleResearch and DevelopmentRendezvous and dock
South Atlantic AnomalyScience Applications Inttmmtional CorporationSpace Exploration InitiativeSolar-electric propulsionInternational system of units (tmtric systcm)Silicon carbide
Semimajor axisSolar day (24.6 hours for Mars)Soalr proton eventsSolid Rocket Boosm"
Space Station FreedomSpace Shuttle Main EngineSpace Transfer Concepts and Analysis for Exploration MissionsStageSurface
Sievicrt (SI unitof dose equivalent= Gy x Q)
Distance along aca'obralm surface forward of the stagnation pointDistancealongaea'obralmsurface afz of the stagnationpoint
Distance along aerohra_ surface starboard of the stagnation point
Metrictons(1000kg)To be demmined
Chamber _
Thermal controlsystem
Trans-Ea_ injectionTram-Earth injection stage
Tank weight factorTemperatu_ and humidity control
Trans-Mars injectionTram-Mars injectionstage
Thermal pmu=cfion s_r_znTracking,mlcmctry, and control
Thrust to weight ratio
Uranium nitride-Tungsten/25% Rhenium reactorfucl
VehicleAssembly Building
Vapor coolledshieldVelocityatinfinity
D615-I0026-6 8
WBc2C/B4CWMS
W/OWP-01
w/_l cm
Tungsmn be_llium cabide/Boron cabid_ composit_Waste managcn_.at sys_mWithout
Work package 1 (of SSF)
Warts l_r square ce_tin_t_r (should be Wcm -2)
Z
ze_og
Atomic number
An unaccchwated fram$ of mfcmncc, free-fall
[order:. numbvrs followed by greek lcttea's]
100K7n7_k
.-cAV
$
_tg
_<100,000 particles per cubic mcmr larger than 0.5 micron in diamcmrWhere n=(0,2-6): Boeing Company jct transport moc_l numi_rsKelvin (K)Positive charge equal to charge on electronChar_ on electronChangc invelocity
Standard_viation
Microgravity ( also call_ z_r_-gravity)
D615-10026-6 9
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D615-10026-6 10
I. Evolution of the Concept
A. Reference Concept Development
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D615-10026-6 12
I. Evolution of the Concept
During the 90 Day Study, NASA's two Office of Space Flight (Code M) Space Transfer Vehicle
(STV) contractors supported development of SEI lunar transportation concepts. This work treated
lunar SEI missions (and evolution to the support of later Mars missions) as the far end of a more
near-term STV program, most of whose missions were satellite delivery and servicing
requirements derived from Civil Needs Data Base (CNDB) projections. STCAEM's contribution
to that effort focused mainly on crew system design, since this was recognized as offering potential
for commonality with crew cab design for Mars excursion vehicles (MEVs).
Later, STCAEM began to address the complete design of a lunar transportation system.
Because of our Mars concept experience, our perspective was particularly sensitive to evolutionary
systems; the approach of looking back from a Mars mission perspective is thus complementary to
that of the parallel NASA studies. Our effort was guided by attention to two broad drivers. Fast
were precisely those technical requirements whose resolution had proved so intractable for earlier
concepts:
1) State-of-the-art understanding of constraints imposed by the detailed geometry of
aerobraking upon Earth return: non-symmetrical relative wind configurations for lifting flight
profiles; off-axis placement of composite mass-center (CM); and changing mass-balance conditions
due to sequential propellant expulsion.
2) The need to accommodate "mixed" payloads in a reasonable lunar exploration program:
versatility in the delivery of a wide variety of heavy-cargo payload manifests, rarely if ever mass-
split evenly; cargo processing and loading requirements in LEO; cargo exchange between transfer
vehicles and excursion vehicles; cargo offloading on the surface of the Moon; cargo placement on
manned flights; and shirtsleeve (IVA) exchange of crew between transfer and excursion vehicles.
3) Provision for transfer of cryogenic propellants: a typical scenario is supplying
I-2"I2, brought from Earth by a transfer vehicle, to a reusable lander based in low lunar orbit (LLO).
Cryogenic propellants are baselined, of course, because of the requirement for high-thrust
propulsion for planetary landing and ascent. (The use of nuclear thermal propulsion for lunar
transfer is potentially attractive, but still involves cryogenic propellant management.)
4) Potential for full system reusability: designs which drop tanks are better for limiting
aerobrake size, but have negative cost implications for advanced cryogenic storage technology, and
D615-10026-6 13
negative operational implications via the accumulation of empty tanks in cis-lunar space and on the
lunar surface. Mission modes which posit multiple annual flights for several decades drive us to
consider full reusability.
Second, we recognized that over several decades of lunar operations, many mission modes
should be accommodated. What is n_dea:i is not so much a single vehicle or pair of vehicles, but
rather an evolving lineage of vehicles, fabricated on long-lived production lines, which can be
adapted gracefully and economically to handle contemporary requirements. Two observations
keyed this investigation:
1) Lunar flight hardware decisions will probably be made before final site selection decisions.
This means that the lunar wansponation architecture should be careful not to constrain site selection
to less than potentially global access. Many possible mission modes must be preserved by the
arch/tecture.
2) An early lunar surface operations capability can be obtained by using a tandem-direct flight
mode, in which one lunar transfer vehicle (LTV) boosts another, "campsite" LTV, to a fractional-
orbit, direct landing on the Moon. The crew would be sent separately on an identical profile,
lemming directly to Earth's surface in their heat-shielded crew capsule. No LLO operations, no
aerobrake, no LTV recovery, and no space station rendezvous upon return would be needed, nor
would a specialized lunar lander (LEV).
What resulted was a lunar mmsportation family (LTF) concept, consisting of various
"models" of two basic, cryogenic vehicle "chassis": an LTV with 110 t propellant capacity, and
an LEV with 25 t capacity. Particular vehicle combinations from this evolutionary family can
handle 11 distinct mission modes, to provide versatile, flexible service for decades as mission
requirements evolve. For instance, the addition of an aerobrake would permit unmanned recovery
of the boost-stage LTV, prodding invaluable flight qualification experience for later man-rating.
(Such an aerobrake can be essentially the symmetrical central core of the asymmetrical Mars-class
aerobrakes discussed later, since the L/D requirement is only about 0.25 for lunar missions;
aerobrake technology evolution is then enhanced.) A heavy-cargo lander would be a modest
upgrade to the campsite vehicle design. If the scale of the exploration architecture justified the
more efficient lunar-orbit-rendezvous (I.OR) mode, a dedicated lunar excursion vehicle (LEV)
could be introduced. For fully reusable operations, a version of the campsite habitation module
D615-10026-6 14
would provide crew support during the ex_nded-walt required by remm orbit phasing. If electric
propulsion Mars missions were operated efficiently through a lunar libration-point, LTVs could
support these as weU as lunar operations. And LTVs could supply the f'mal capture propeUant to
an NTR vehicle returning from Mars. Lunar transportation operations can be upgraded to the use
of lunar-derived oxygen (LLOX) with this family of vehicles also. All combinations of crew and
cargo manifests identi_fied so far for lunar support, and all lunar-related SEI missions identified so
far, can be accommodated by the LTF.
D615-10026-6 15
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Lunar NEP
This section contains a preliminary parametric analysis of Lunar NEP performance
capabilities. The section contains the following
• IMLEO vs Isp for various powex levels
• PropeUant Mass vs Isp for various power levels
• Trip Time vs Isp for various power levels
• Payload Fraction vs Isp for various power levels
• Breakdown of IMLEO for various power levels - 10 kg/kW
• Breakdown of IMLEO for various power levels - 15 kg/kW
A power level of 3 MW (@ 5,000 sec Isp) will transfer 100 t of payload in less
than 6 months. The IMLEO of this vehicle is -150 L The analysis assumes constant bum
dme and uses fundamental electric propulsion equations.
4
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D615-10026-_ 56
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133 dayst_
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184 days
_248 days
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A. Level I, II and Ill Requirements
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Lunar IP&ED Text
[Editorial note: viewfoil charts are referred to as 'charts'. If the convention in the
IP&ED documents is to refer to them as 'figures' or 'tables',do a find and replace
operation (_g H in Microsoft Word 4.0)]
II. Requirements and Assumptions
II.A. Levied Requirements
There is not a conn'olled baseline (system specification and configuration) for this study.This sectionincludesour bestunderstandingof what NASA would defineasrequirements
iftheyhad tobe baselinedatthistime.Italsoincludeslower levelrequirementsand
assumptionswe derivedor made respectively.
Level Irequirements(firstchart)arcconcernedwith overallprogram schedule,mission,
funding,and interfacetootherprograms. Level 11requirements(charts2 through4) aregrouped by systems:Earth-to-orbit(ETO) wansponafion,ETO supportfacilities,space
wanspormtion vehicles,crew u'ansfcrmodule, and Lunar surfacesystem interface.
The levelIrequirementofa fin'stcargolandingin2000 leadtolevelH requirementsfor
ETO n'anspormtiontestflightand Space StationFreedom (SSF) supportreadinessin 1999.The Level I requirement to use SSF leads to specific SSF accommodations requirements intheLevel IIrequirements.
Level IT[requirementsincludevehiclespecificrequimmenr,s.The fifthchartlists
requirementsleviedon the spacewanspormtionvehicles.These are.propulsioncharacteristics:cryogenicpropellantwithan aerobrake;designmargins;and operationalcharacteristics:boiloffofcryogens,propeUantu'ansfercapability,and baselineorbit.
II.B. Derived Requirements
Inthecourseofthe study,derivedre+quimmentswere developedfi'omthelevied
requirementsand theanalyses.In some caseschanges to theleviedrequirementsarcrecommended. Issuesaddressedby thederived requirementsinthemissionareaincludethereferencemission,basing,on-orbitassembly,modularity,and life(chartsix).Other
areasaddressedinderivedrequirementsarehabitationmodule, propulsion,and wansfcr
vehiclerequirements(chartseven),and excursionvehicleand aerobrakerequirements
(charteight).Excursionvehiclerequirementsincludeperformancerequirementsand design
requirements.
II.C. Assumptions
Assumptions have tobc made intheinitialcycleofanalysisbecauseallthedam requiredis
not yetavailable.Inlateranalysiscycles,theseinitialassumptionswilleitherbe validatedor replacedby more correctvalues.The subjectsforwhich assumptionshave been
recordedarecrew size,cargo capacity,acrobrakecharacteristics,ETO vehiclecapacity,
missionmode, and engineoutcapability.
HI. Mission Operation
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III. Mission Operations
A. Mission Analysis and Performance Parametrics
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HI. Mission Operations
The mission operations section includes data on mission analysis studies and performancepammetrics as well as the operating modes and performance evaluations which include theSTCAEM recommendations.
A. Most of the lunar miss/on analysis and performance data was genera_'.zi during the90-day study time-frame. Included in this document is data on timing of translunartrajectories, the lunar orbit insertion AV for different arrival asymptotes, transit time, and
opportunity for lunar transit leaving fi'om the vicinity of Space Station Freedom.
B. Initially,we identifiedseven lunar vehiclemodes which could be implemented with apairof vehicles;an LTV-like and an LEV likevehicle.During the courseof configuring,
sizing and generating performance data on these vehicles,11 mission modes were
identifiedwhich can be implemented with 5 major elements;a 110 tpropulsionstage,a 25 t
propulsion stage,a transferhab, a crew cab and an aerobrake. Implementation of thevariousmissionmodes isaccomplished by "piecing"togethertherequiredelements. The
resultisa Lunar TransportationFamily (LIT) thatis flexibleand can evolve to meet
growing missionneeds and changing missionmodes.
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A. LTV/LEV Components
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IV. Element Descriptions
Element descriptions for the lunar transportation family included in this document arc alisting of the LTV/LEV components, trade studies and mass analyses of the transfer andexcursion modules, ACRV (MCRV) modifications required to fulfill lunar operations, theaerobrake shape and L/D to be used, and some costing methods and results.
A. Component listings, assumptions and sizing criteria ar_ included for the LTV, LEV,ACRV and the service module (Apollo command module derivative). This information isprovided to give an overview of the major components of the lunar transportation familyand their related subsystems.
B. An LTV/LEV habitat trade module study was conducted to size crew modules forvarying crew sizes and mission durations. Two types of u'ansfer modules were evaluated,an acrobrakcd module and a direct entry (Apollo-type) module, as weU as a single moduleconcept for transfer and excursion (direct entry at Earth) and excursion modules. Crewsizes of 2, 4, 6, and 8 for transfers of 24 days and surface stays of 1, 14, 28 and 42 days.Sizes for these 36 modules were generated from historical spacecraft data, and massstatements were generated from SSF and STCAEM estimates. Results of this u'ade studyprovide good estimates as to the size and mass of lunar crew modules.
A trade study was also performed to determine what point it becomes more mass efficientto have a separate surface hab along with an excursion module, if a base is not availableand missions of the excursion/exploration class are being performed. Results show that 3-8 days is the crossover point.
C. The SSF ACRV was originally believed to be easily adaptable to small scale lunarmission, which would allow the use of "existing" hardware. For the small scale program,the ACRV was to be used for the reentry phase back at Earth. However, the ACRVcurrently envisioned for SSF will not fulfzU lunar mission needs because of its size. Theinternal volume is extremely Limited since it is designed for a 6 hour mission rather than 7to 24 day missions. In order to provide sufficient volume for crew operations andequipment storage, the interior volume would need to be increased approximately 250%(giving the same amount of free volume as the Apollo Command Module). Increasing thevolume requires major structural modifications and the resulting craft would be unlike thecurrent ACRV. Therefore, a more appropriate use for the ACRV is as a crew module forthe LTV in a direct-to-surface (tandem direct) mission scenario. Prior to Earth reentry, thecrew transfers to the ACRV and separates from the LTV, which is expended.
D. The aerobrake to be used in the lunar transportation family is envisioned as being an
early version of the Mars low L/D (L/D = 0.5) shape. The idea is to take the symmetricalcenter portion of the hyperboloid shape and fly at an L/D of about 0.25. This symmetricalportion of the Mars shape is about the right size to accommodate the LTF without having todrop tanks. Another variation to this concept is to have standard "additions" that can beadded to the outer rim of the brake to accommodate growing mission needs. When Mars
comes into the picture, an "addition" can be fabricated to accommodate these missions.
Heating analysis was performed on this shape, and it was found that the ballistic coefficientis very low in comparison with Mars and the heating temperatures only reach about 1870°Kat the stagnation point.
E. Costing for the lunar transportation family is being calculated for both hardware cost,using the Boeing Parametric Cost Model (PCM), and life-cycle cost, using a model
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Lunar Transfer Crew Module
24 day duration
Cre w SizeHabitable Volume (m3)
i
ECLSS Total
Atmosphere Revitalization System (ARS)ACS (tanks & I/'2 nec. SSF equip.)
Atmos. CompositionMonitorAssemblyThermal Control/Temp. & Humidity
Control (I/2SSF- av.airequip.)
PotableWater and StorageSysmm
Fire Detectionand SuppressionSystemStructureTotal
End cones(2)
Berthingring/mechanism(I)
Berthinginterfaceplate(I)
Cylinderprimarystructure
Cylindersecondarystmctu_Standoff/u_lilities/distribution
I-law_es(2)
W'mdows (4)
Couches/sleepersComman_ControL/Power total
ECWS (1/2 SSF)DMS/andio - visualFault detection and isolanon
Power system(solararrays,batteries,
onboard equipment)
Lights(1/2SSF)i
Man.Systems l oralWMSlwastc storage
EVA suits/spaceclosureballs
Medicalequipment(1/2surf.e.quip,mass)_LonstimaOleS l'Otmi
Food and packagingAtmospheric mak=-up and 3 r_pmsses
(20% reserve)Other (clothes,hygiene equip.,etc.)
Potablewater
uther Iota|
Personneland effects
Equipment sparesTools
Radiationshelter
mass tarowtn t ota_ (J__/o)i
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274125
479
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D615-10026-6 176
Lunar Transfer Crew ModuleDirect Entry
24 day duration
SOlTdUOJ I_l_u
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1,766
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Note: All weights areinkilograms
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I day and 14 day surface stays
Crew Size
Habitable Volume (m3)**
ECLSS
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Command/Conu'ol/Power| i
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1,272
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** Volumes sized for 21 day nominal case
Note: All weights arc inIdlograms
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D615-10026-6 181
Lunar Transfer/Excursion Crew ModuleDirect Entry
35 day duration*
28 day surface stay
Crew Size 2
Habitable Volume (m3) 66
!ECLSS 1.293
Su'uculm 2,801
Comm.-d/Con_rollPower 1,703
Man-S_-ms 417Consumabl_ 893
Personnel and Et_ Spares, e¢. 305p.H_._n, SheJ_r 778
Emh En_ 7 Heat Shied 2.600
Earth R_ov¢_ P_uipm_nt 647
Mass Growth (15%) I_376
Total Module Mass 12,813
4 6 8
132 198 264
1.835 2.692 3,549
3.901 4.838 6,045
2.476 3.232 4,021
457 497 537
1.718 2,578 3.436
489 673 923
1,202 1,626 2,050
4,212 5,501 6,581
977 1,298 1,629
2,190 2+878 3_581
19,457 25,813 32_352
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7 day round trip time
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49 day duration*
42 day surface stay
Crew Size
Habitable Volume (m3)
ECLSS
Strucna'e
Command/Control/Power
2 4 6
84 168 . 252
1,693 2,403 3,546
3,145 4,586 6,002
1,703 2,476 3,232
457 497
2,289 3,435
489 673
1,202 1,626
4,943 6,456
1.131 1,528
2,511 3,359
22,487 30.,354
Man-Systems 417Consumables 1,146
Personnel and Effects, Spares, etc. 305Radiation Shelter 778
Earth Entry Heat Shield 3,094
Earth Recovery Equipment 737
Mass Growth (1_%) 1,673Total Module Mass 14,691
8
336
4,687
7,442
4,021
537
4,579
923
2,050
7,757
1,919
4,181
38_096
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7 day round trip time
Note: All weights are in kilogramsLunar sample material mass not included
1?RE ARY
D615-10026-6 183
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Lunar Family
Programmatics
The objectives of the Programmadcs task during the current phase of the study were: (1)
realistic initial schedules that include initial critical path program elements; (2) initial
descriptions of new or unique facilities requirements; (3) development of a stable, clear,
responsive work breakdown structure (WBS) and WBS dictionary; (4) initial realistic
estimates of vehicle, mission and program costs, cost uncertainties, and funding profile
requirements; (5) initial risk analysis, and (6) early and continuing infusion of
programmatics data into other study tasks to drive rexluirements/design/ffad= decisions.
The issues addressed during the study to dam included: (1) capturing all potential long-lead
program items such as precursor missions, technology advancement and advanced
development, related infrastructure development, support systems and new or modified
facility construction, since these are as important as cost and funding in assessing goal
achievability; (2) incorporating sufficient operating margin in schedules to obtain high
probability of making the relatively brief Mars launch windows; (3) the work breakdown
structure must support key study goals such as commonality and (4) cost estimating
accuracy and uncertainty are recurring issues in concept definition studies.
Introduction
The study flow, as required by MSFC's statement of work, began with a set of strawman
concepts, introduced others as appropriate, conducted "neckdowns", and concluded with a
resulting set of concepts and associated mconanendations.
As the study progressed, much discussion among the SEI community centered on
"architecua'es". In this study, architectures were more or less synonymous with concepts,
since the statement of work required that each concept be fully developed including
operations, support, technology, and so forth.
We started with ten concepts as shown in the "Overall Study Flow" chart. After the
"neckdown" was completed, significant effort was put into programmatics.
D615-10026-6
PREOe_-'I.-_CNGPAGE BLANK NO.,"r "" MF..D 225
As was indicatedearlier,we establishedthr_ levelsof activityto cvaluat_ in-space
transportationoptions. The minimum was justenough tomeet the President'sobjectives;
in fact "returnto the Moon to stay" was interpretedas permanent facilitiesbut not
permanent human presence.The minimum program had onlythr_ missionstoMars. The
median (fullscience)program aimed atsatisfyingmost of thepublishedscienceobjectives
forLunar and Mars exploration.The maximum program aimed forindustrializationofthe
Moon, forreturnof practicalbenefitsto Earth,and forthe beginnings of colonizationof
Mars. The range of activitylevels,as measured by people and materieldeliveredto
planetarysurfaces,was about a factorof 10.The range ofEarth-to-orbitlaunchrams was
less,sincewe adopted resultsof preliminarytradestudies,selectingmore advanced in
space transportationtechnologiesas baselinesforgreateracdvitylevels.The high level
schedules developed for these three levelsof activityarc shown in the "Minimum
Program", '_FuliScienceProgram" and "Industrializationand SettlementProgram" charts
and a comparison of them for both Lunar and Mars is shown in the "Lunar Program
Comparison" and "Mars Program Comparison" charts.
Schedule/Network Development Methodology
A PC sysmm calledOpen Plan by WST Corporationwas used,which allowsdirectcontrol
and lower costover alarger(mainfl'amc)system. The network was purposelykeptsimple.
Summary activitieswere used indevelopment of thenetworks. When detailedtoa lower
level,some activitieswillrequirea differentcalendarthan we used. One calendarwith a
five day work week - no holiday was used. Utilizingmulticalendarson a summary
network could confuse the development. The PreliminaryWBS Structtn_Level 7 was
followed for selectionof work to bc detailed.An example of Level 7 is: MEV Ascent
Vehicle Sn-ucture/Mechanisms.We thendeveloped a genericlogicstringof activitieswith
standarddurationsforlikeactivities.This logicwas thenappliedagainsteach WBS Level
7 element. To establishinterfacetiesbetween logicstringsand determinationof major
events,we used the Upper Level Summary Schedule and Summary I.avelTechnology
Schedule.
D615-10026-6
226
Goals/Purpose
There were two goals for the schedule/network development These were:
a. Guidelines for Future Dcvelopn_nt. The schedules are a preliminary road map to
follow in the development program.
b. Layout Basis Framework forNetwork. The networks can be used forfuturedetail
network development. This development can be in phasesretainingunattended logicfor
areaswhich can bc be derailed.
Status
Six preliminarynetworks have been d_velopeA They are:
- Lunar minimum
- Lunar fullscience
- Lunar indusmaliz_on
- Mars missions
- Mars fullscience
- Mars serdemcm
These networks willbc fltrtherdeveloped asinformationbecomes availableThe technology
developmem plan schedulesarc shown inthe Schedulessectionof thistext;an example of
the standard 6 year program phase C/D schedule isshown in the "Reference 6 yr.Full
ScaleDevelopment Schedule" chart. The network schedulesdeveloped duringthe study
am availableintheFinalReport CostDam Book.
Facilities
The facility requirements and approaches are discussed in the Facilities section of this text.
D615-10026-6
227
Work Breakdown Structure
The approach to d_veloping a W'BS tree and dictionary was to use the Space Station
Freedom Work Package One WBS as a point of departure to capture commonality,
modularity and evolution potentials. We worke, d with MSFC to evolve the WBS illustrated
in the six WBS charts given in this section. The WBS dictionary _mils are provided with
the WBS u_ in a separa_ dclivvrablc document.
Cost Data
Overall Approach
Space transfer concept cost estimates were developed through parametric and detail
estimating techniques using program/scenario plans and hardware and software
descriptions combined with NASA and subcontractor data. Our estimating approach
simulates the aerospace development and production env/ronment. It also re_ program
options not typical of aerospace programs. This flexibility allows assessment of innovative
program planning concepts.
Several tools were emPloyed in thisanalysis. For developing estimatesthe Boeing
ParametricCost Model (PCM) designed specificallyforadvanced system estimatingwas
used. Itutilizesa company-wide, uniform computerized dam base containinghistorical
dam compiled since1969. The second major toolisa Boeing developed LifeCycle Cost
Mod_l. The thirdtoolistheBoeing developedRcun'n on Investment('ROD Analyses.
The approach to cost estimating was to use the PCM to establishDDT&E and
manufacturing costof major hardware components orto use otherestirnatcs,(e.g.Nuclear
Working Croup estimator)fftheywere consideredsuperiorand thenfce._ithem totheLCC
model. Variationson c,quipment hardware or mission aitcmativescan be run through the
LCC and thencompared forareturnon investment.This flow isillustratedinthe"Costing
Methodology Flow" charts. We were able to investigatealternativeconcepts quickly,
giving system designersmore data for evolving scenario/missionresponsive concepts.
Transportationconcepts, tradestudies,and "neckdown" effortswere supported by this
approach.
D615-10026-6
228
Parametric Cost Model
PCM develops costfrom the subsystem leveland buildsupward to obtaintotalprogram
cost. Costs are estimated from physical hardware descriptions(e.g.,weights and
complexities)and program parameters (e.g.,quantifies,learningcurves,and integration
levels). Known costs are input directlyinto the estimate when available;the model
assessesthe necessarysystem engineeringand system testeffortsneeded for integration
intotheprogram. The PCM working unitisman-hours, which allowsrelationshipsthattie
physicalhardware descriptionsfirsttodesignengineeringor basicfactorylabor,and then
through the organizationalstructureto pick up functional areas such as systems
engineering, test,and development shop. Using man-hours instead of dollarsfor
estimatingrelationshipsenablesmore reliableestimates.The PCM features,main inputs,
and resultsarc shown in the "Boeing Parametric Cost Model (PCM)" chart. The
applicablePCM results,in constant1990 dollars,are then put intothe Life Cycle Cost
Model to obtaincostspreads forthe variousmissions/programs. The varioushardware
components costed for the threedifferentmissions/programs arc shown in the "LCCM
Hardware Assignments" chart.
The development of space hardwa_m and components needed to accomplish the three
differentLunar/Mars missions wcrc identified.These components are grouped intothree
differentcategoriesdefinedbelow.
HLLV(I-leavy LiftLaunch Vehicle)istheboosterrequiredto liftpersonnel,cargo and
fuelsintoLEO and supporttheLEO node operations.
Prouulsion Includesthe space propulsionsystem requiredto transferpeople,cargo and
equipment out of LEO and intospace. Space means Lunar, Mars and Earth destinations.
PropulsionSystems alsoincludean all-propulsivecryogenicTrans Mars InjectionSystem
(TMIS) for the Minimum Mission, the Nuclear Electric Propulsion Stage for the
Settlement/IndustrialMissions.
Modules Include the space systems thatare required to transferpeople, cargo and
equipment from LEO toLunar and Mars orbit;tode-orbitand sustainlifeand operationson
theLunar and Mars Surface;and,finally,toreturnpersonneland equipment toLEO.
D615-10026-6
229
CostBuildups
ThePCM cost Model can be used directly to obtain complete DDT&E cost, including
productionof major testarticles,by enteringintothemanufacturing sectiontheequivalent
numbers of unitsforeach item,includingthefirstflightarticle.However, when opcxatedin
thisway, PCM does not give the fn'stunitcost.To save time,we ol:_1"atedPCM so as to
givefirstunitcost,which we _ forlifecyclecostanalyses,and used the firstunitcost
tomanually estimatethetesthardware contentof theDDT&E program. The "wrap factors"
shown in the costbuildup sheetswe,re derived from the PCM runs as the factorthatis
appliedtodesign engineeringcostto obtaincomplete design and development costs,e.g.
includingnon-recurringitemssuch as sysmms engineeringand toolingdevelopment.
v
Life Cycle Cost Model
The LCCM costdataisa composite of HLLV costs,launch base facilitiescostestimate
based on $1sq.ft.and parametricestimatesderivedfrom the ParametricCost MOd_I. The
principalsource of informationisfrom the PCM. All hardware costestimates,with the
exceptionof HLLV, have been developed withthismodel.
The LCCM consistsof threeindividualmodels. One model isforthe Minimum Program
Scale;the second isfor theFull Science Program Scale;while the thirdmodel isfor the
Scttlemcnt/tnduswializationProgram Scale. The Minimum Program meets thePresident's
Space ExplorationInitiative(SEI)objectives.These capabilitiesincludepermanent Lunar
facilitiesbut not permanent human prcseneeand thr_ missionstoMars. The FullScience
program not only meets the President'sSEI objectivesbut alsoprovides for long term
bases for far-ranging surface exploration. The Setflement/Induswializationprogram
accomplishes the objectivesof the Minimum and Full Science program scalesand
additionallyreturnspracticalbenefitsto Earth. These models were developed using the
thr_ architecturelevelsdescribedinthe Boeing manifestworksheets. Totalcostforeach
system aretabulatedby yearand each year'stotalsfeedintoa summary sheetthatcalculams
the totalprogram costforeach level.Since theLCCM resultsam missionr_la_L notjust
vehiclerelated,theyarc not provided herebut arcavailablein theFinalReport Cost Data
Book. The LCCM was developed using MicrosoftExcel version2.2 for theMacintosh
computer. Any Macintosh equipped withExcel 2.2can be used toexecutethemodel
v
D615-10026-6
230
Return On Investment
One of the principal uses of the LCCM is to develop trades and return on investment for
technology options.As shown in the "CostingMethodology Flow" chart,two separate
lifecyclecostmodels (which includeDDT&E and productiohcostdataderivedfrom the
parametriccostmodels ) must be developed for each ROI case;a reference,and a case
utilizinga technologyoption. The two lifecyclecoststreamsarcseparatelyentered,and
the ROI model isexecuted.The flow alsoillustratesthatnot allof thedataenteredintothe
lifecyclecostmodel isderivedfrom availablecostingsoftware.Technical analysismust
accompany thisdata. For example, the number of unitswhich must be produced for the
DDT&E program must be determined. This isdone at the subsystem levelbased on
knowledge of past programs, and proposed system/subsystem tests. Since the ROI
analysisismission related,not justvehiclerelated,the dataisnot presentedhere but is
availableintheFinalReportCost Data Book.
Results
A summary of the costdata produced by the PCM for the lunarfamily of vehiclesare
given in the PCM summaries included in thissection.The PCM program was used to
produce DDT&E and productioncostestimatesforeach of our referenceMars and lunar
vehiclestothe subsystem level.The DDT&E costsgeneratedby thePCM do not include
allof the necessaryhardware forthefirstmission vehicle.Hence allnecessaryadditional
units(prototypes,testunits,lab units,etc.)were added intothe vehiclecost buildups as
shown in the "Lunar Cost Buildup" charts.The totalDDT&E includesadditionalcosts
(e.g..additionalunitsintheDDT&E program), contractorfeesand theengineeringwrap
factor.The totalDDT&E from the costbuildup and the unitcostfrom the PCM are the
primary vehicle cost inputs to the LCC model
D615-10026-6
231
Risk Analyses
Risk analyses were conducted to develop an initial risk assessment for the various
architectures. This presentation of risk analysis results considqrs development risk, man-
rating requirements, and several aspects of mission and operations risk.
Development Risk
All of the architectures and technologies investigated in this study incur some degree of
development risk; none are comprised entirely of fully developed technology.
Development risks are correlated directly with technological uncertainties. We identified
the following principal risks:
Cryogenics - High-performance insulation systems involve a great many layers of multi-
layer insulation (MLI), and one or more vapor-cooled shields. Analyses and experiments
have indicated the efficacy of these, but demonstration that such insulation systems can be
fabricated at light weight, capable of surviving launch g and acoustics loads, remains to be
accomplished. In addition, there are issues associated with propellant transfer and zero-g
gauging. These, however, can be avoided for early lunar systems by proper choice of
configuration and operations, e.g. the tandem-direct system recommended elsewhere in this
report. This presents the opportunity to evolve these technologies with operations of initial
flight systems.
Engines - There is little risk of being able to provide some sort of cryogenic engine for
lunar and Mars missions. The RL- 10 could be modified to serve with little risk; deep
throttling of this engine has already been demonstrated on the test stand. The risk of
developing more advanced engines is also minimal. An advanced development program in
this area serves mainly to reduce development cost by pioneering the critical features prior
to full-scale development.
Aerocapture and aerobraking - There are six potential functions, given here in approximate
.ascending order of development risk: aero descent and landing of crew capsules returning
from the Moon, aerocapture to low Earth orbit of returning reusable lunar vehicles, landing
of Mars excursion vehicles from Mars orbit, aero descent and landing of crew capsules
returning from Mars, aerocapture to low Earth orbit of returning Mars vehicles, and
D615-10026-6
232
acrocapun'eto Mars orbit of Mars excursion and Mars transfer vehicles. The "Development
Risk Assessment for Aerobraking by Function" chart provides a qualitative development
risk comparison for these six functions.
Aerocapture of vehicles requires large aerobrakes. For these to be efficient, low mass per
unit area is required, demanding dfficient structures made from very high performance
materials as well as efficient, low mass thermal protection materials. By comparison, the
crew capsules benefit much less from high performance structures and "rPS.
Launch packaging and on-orbit assembly of large aerobrakes presents a significant
development risk that has not yet been solved even in a conceptual design sense. Existing
concepts package poorly or arc difficult to assemble or both. While the design challenge
can probably be met, aerobrake assembly is a difficult design and development challenge,
representing an important area of risk.
)
Nuclear thermal rockets - The basic technology of nuclear thermal rockets was developed
and demonstrated during the 1960s and early 1970s. The development risk to reproduce
this technology is minimal, except in testing as described below. Current studies arc
recommending advances in engine performance, both in specific impulse (higher reactor
temperature) and in thrust-to-weight ratio (higher reactor power density). The risks in
achieving these are modest inasmuch as performance targets can be adjusted to technology
performance.
Reactor and enginetestsduringthe 1960s jettedhot,slightlyradioactivehydrogen directly
intothe atmosphere. Stricterenvironmentalcontrolssincethattime prohibitdischargeof
nuclearengine effluentintothe atmosphere. Design and development of fullcontainment
testfacilitiespresentsa greaterdevelopment riskthan obtainingthe needed performance
from nuclear reactors and engines. Full- containment facilities will be required to contain all
the hydrogen effluent, presumably oxidize it to water, and remove the radioactivity.
Electric Propulsion Power Management and Thrusters - Power management and thrusters
are common to any electric propulsion power source (nuclear, solar, or beamed power).
Unique power management development needs for electric propulsion are (1) minimum
mass and long life, (2) high power compared to space experience, i.e. megawatts instead of
kilowatts, (3) fast arc suppression for protection of thrusters. Minimizing mass of power
distribution leads to high distribution voltage and potential problems with plasma losses,
D615-10(Y26-6
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arcing, and EMI. Thus while pow_" management is a mam._ technology, the unique
rcqui_m_nts of electricpropulsionintroducea number ofdevelopment risksbeyond those
usuallyex1_riencedin spacepower systems.
Electricthrustertechnology has be,cn under development sincethe beginningof the space
program. Small tb.rustcrsarc now operational,such as the r_sistance-heat-augmented
hydrazinethrusterson certaincommunications spacecraft.Small arc and ion thrustersarc
nearingoperationaluse forsatcUitcstationke,cping.
Space transfer demands on electric propulsion performance place a premium on high power
in the jet per unit mass of electric propulsion system. This in turn places a premium on
thrusterefficiency;power in the jet,not electricalpower, propels spaceships. Space
transferelectricpropulsion alsorequiresspecificimpulse in the range 5000 to 10,000
seconds. Only ion thrustersand magnetoplasmadynamic (MPD) arcthrusterscan deliver
thisperformance. Ion thrustershave acccptabl_efficiencybutrelativelylow power perunit
of ion beam emittingarea. MPD thrustertechnologycan cl_liverthe needed Isp withhigh
power per thruster,but has not yetreached cfflciencicsof interest.Circularion thrusters
have been builtup to50 cm diameter,with sphericalsegment ion beam grids.These can
absorb on the order of 50 kWe each. A 10 MWe system would need 200 operating
thrusters.The development alternativesallhave significantrisk:(I)Advance the smm of
the artof MPD thrustersto achievehigh efficiency;(2)Develop propulsionsystems with
largenumbers of thrustersand controlsystems;or (3)Advance the stateof the artof ion
thrusterstomuch largersizeper thruster.
(
Nuclear power for electricpropulsion - Space power reactor technology now under
development (SP-100) may bc adequate;needed advances arc modest. Advanced power
conversion systems arc requiredto obtainpower-to-mass ratiosof interest.The SP-100
baselineisthermoelectric,which has no hope of meeting propulsionsystem performance
needs. The most likelycandidatesarcthe closedBrayton (gas)cycle and the potassium
Rankine (hquid/vapor)cycle. (Potassium provides the bestmatch of liquid/vaporfluid
propertiesto desired cycle tcmperatu_s.) Stiflingcycle,thermionics, and a high-
temperaturethermally-drivenfuelcellarcpossibilities.The basictechnology forBrayton
and Rankine cycles arc mature; both arc in widespread industrialuse. Prototype space
power Brayton and Rankine turbineshave run successfullyfor thousands of hours in
laboratories.The development riskhcrcisthatthesearcvery complex systems;thereisno
experiencebase forcoupling a space power reactortoa dynamic power conversioncycle;
D615-I0026-6
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there is no space power experience base at the power levels needed; and these systems, at
power levels of interest for SEI space transfer application, are large enough to require in-
space assembly and checkout. Space welding will be required for fluid systems assembly.
Solar power for space transfer propuls, ion - Solar power systems for space propulsion must
attain much higher power-to-mass ratios than heretofore achieved. This implies a
combination of advanced solar cells, probably muiti-band-gap, and lightweight structural
support systems. Required array areas arc very large. Low-cost arrays, e.g. $100/watt,
arc necessary for affordable system costs, and automated construction of the large area
structures, arrays, and power distribution systems appears also necessary. Where the
nuclear electric systems arc high development risk because of complexity and the lack of
experience base at relevant power levels and with the space power conversion technologies,
most of the solar power risk appears as technology advancement risk. If the technology
advancements can be demonstrated, development risk appears moderate.
Avionics and software - Avionics and software requirements for space transfer systems are
generally within the state of the art. New capability needs are mainly in the area of vehicle
and subsystem health monitoring. This is in part an integration problem, but new
techniques such as expert and neural systems are likely to play an important role.
An important factor in avionics and software development is that several vehicle elements
having similar requirements will be developed, some concurrently. A major reduction in
cost and integration risk for avionics can be achieved by advanced development of a
"standard" avionics and software suite, from which all vehicle elements would depart.
Further significant cost savings are expected from advancements in software development
methods and environments.
Environmental Control and Life Support (ECI_) - The main development risk in ECLS is
for the Mars transfer habitat system. Other SEI space transfer systems have short enough
operating durations that shuttle and Space Station Fre_om ECLS system derivatives will
be adequate. The Mars transfer requirement is for a highly closed physio-chemical system
capable of 3 years' safe and dependable operation without resupply from Earth. The
development risk arises from the necessity to demonstrate long life operation with high
confidence; this may be expensive in cost and development schedule.
D615-10026-6
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Man-Rating Approach
Man-rating includes thr_ elements: (1) Design of systems to manned flight failure tolerance
standards, (2) Qualification of subsystems according to normal man-rating requirements,
and (3) Flight demonstration of critical performance capabilities and functions prior to
placing crews at risk. Several briefing charts follow: the first summarizes a recommended
approach and lists the subsystems and elements for which man-rating is needed;
subsequent charts present recommended man-rating plans.
Mission and Operations Risk
These risk categories include Earth launch, space assembly and orbital launch, launch
windows, mission risk, and mitigation of ionizing radiation and zero-g risk.
Earth launch - The Earth launch risk to in-space transportation is the risk of losing a
payload because of a launch failure. Assembly sequences are arranged to minimiz_ the
impact of a loss, and schedules include allowances for one make-up launch each mission
opportunity.
Assembly and Orbital Launch Operations - Four sub-areas arc covered: assembly, test and
on-orbit checkout, debris, and inadvertent re-cnu'y.
Assembly operationsriskisreduced by verifyinginterfaceson the ground priortolaunch
of elements. Assembly operationsequipment such as robot arms and manipulators will
undergo space testingatthe node toqualifycriticalcapabilitiesand performance priorto
initiatingassembly operationson an actualvehicle.
Assembly risk varies widely with space transfer technology. Nuclear thermal rocket
vehicles appear to pose minimum assembly risk; eryo/aerobraking are intermediate, and
nuclear and solar electric systems pose the highest risk.
Test and on-orbit checkout must deal with consequences of test failures and equipment
failures. This risk is difficult to quantify with the present state of knowledge. Indications
are: (1) large space transfer systems will experience several failures or anomalies per day.
Dealing with failuresand anomalies must be a routine,not exceptional,part of the
operationsor theoperationswillnot bc abletolaunchspace transfersystemsfrom orbit;(2)
D615-10026-6
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vehiclesmust have highly capable self-test systems and must be designed for repair,
remove and replace by robotics where possible and for ease of repair by people where
robotics cannot do the job; (3) test and on-orbit checkout will run concurrently with
propellant loading and launch countdowns. These cannot take place on Space Station
Freedom. Since the most difficult part of the assembly, test and checkout job must take
place off Space Station Freedom the rest of the job probably should also.
Orbitaldebrispresentsriskto on-orbitoperations.Probabilitiesof collisionarclargefor
SEI-classspace transfersystemsin low Earth orbitfortypicaldurationsof a yearor more.
Shieldingismandatory. The shieldingshould be designed tobe removed beforeorbital
launch and used againon the nextassembly project.
Creationof dc'brismust alsobe dealtwith. This means that(I)debrisshieldingshould be
designed tominimize creationof addiuonaldebris,especiallyparticlesof dangerous size,
and (2)operationsneed tobe rigorouslycontrolledtopreventan inadvertentlossof tools
and equipment thatwillbecome a debrishazard.
Inadvertentre-entryisa low but possiblerisk. Some of the systems, especiallyelectric
propulsionsystems,can have verylow ballisticcoefficientand thereforerapidorbitaldecay
ram. Any of the SEI space transfersystems willhave moderately low ballisticcoefficient
when not loaded withpropellant.While designdetailsam not farenough along tomake a
quantitativeassessment,partsofthesevehicleswould probably survivereentrytobecome
ground impact hazards in case of inadvertentreentry. For nuclear systems, itwillbc
necessarytoprovide specialsupportsystems and infrastructuretodrivethe probabilityof
inadvertentreentrytoextrernclylow levels.
Launch Windows -Launch windows forsingle-bum high-thrustdeparturesfrom low Earth
orbitare no more than a few days be.causeregressionof the parking orbitlineof nodes
causes relativelyrapid misalignmcnt of the orbitplane and _partu_ vector. For lunar
missions,windows recuratabout 9-day intervals.
For Mars, the recurrence is less frequent, and the interplanetary window only lasts 30 to 60
days. It is important to enable Mars launch from orbit during the entire interplanetary
window. Three-impulse Mars depamn'es make this possible; a plane change at apogee of
the intermediate parking orbit provides alignment with the departure vector. Further
D615-10026-6
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analysis of the tlu'_-burn scheme is needed to assess penalties and identify circumstances
where it does not work.
Launch window problems are generally minimal for low-thrust (electric propulsion)
systems.
Mission Risk - Comparative mission risk was analyzed by building risk trees and
performing semi-quantitative analysis. The next chart presents a comparison of several
mission modes; after that are the risk trees for these modes.
Ionizing Radiations and Zero G - The threat fi'om ionizing radiations is presented elsewhere
in this document. Presented here are the mitigating strategies for ionizing radiations and
zero g.
Nuclear systems operations present little risk to flight crews. Studies by University of
Texas at Austin showed that radiation dose to a space station crew f_om departing nuclear
vehicles is very small provided that sensible launch and flight strategies are used. On-
board crews are protected by suitable shielding and by arrangement of the vehicle, i.e.
hardware and propellant between reactors and the crew and adequate separation distances.
After nuclear engines are shut off, radiation levels drop rapidly so that maneuvers such as
departure or return of a Mars excursion vehicle are not a problem. On-orbit operations
around a returned nuclear vehicle are deferred until a month or two after shutdown, by
which time radioactivity of the engine is greatly reduced.
Reactordisposalhas not been completelystudied.Options includesolarsystemescapeand
parkinginstableheliocentricorbitsbetween Earthand Venus.
Crew radiationdose abatement employs "stormshelters"forsolarflares,and eitheradded
shieldingof the entirevehicleor fasttransfers(or both) to reduce galacticcosmic ray
exposure.Assessments axe inprogress;tradeoffsof shieldingversusfastnips have yetto
be completed. Expected impact for lunar missions isnegligibleand for Mars missions,
modest.
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Commonality/Evolution
A. System Commonality
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V. Commonality/Evolution
A. Commonality - Having identified commonality as having high cost-effectiveness leverage,
ST_ d_veloped an evolutionary strategy for facilitating it and benefitting from it. The greamst
commonality challenge is between early systems designed for lunar missions and later systems
d_signcd for Mars missions. Because of similar flight regimes, crew systems requirements and
durations, excursion vehicle crew cabins for lunar and Mars uses proved most amenable to strict
commonality; for conceptual purposes the LEV and MEV crew cabs can be considered just sl.ighdy
different "models" of the same element.
An effort was made to extend the commonality approach to entire excursion vehicle designs
for the Moon and Mars. This extreme degree of commonality was found to appear conceptually
feasible only for a particular size class of vehicle (a 25 t propellant-capacity vehicle typical of
LEV designs is comparable to a so-called "mini-MEV", which would take 3 crew and 1 t
payload to the surface of Mars). Except for this special case, the differing gravity levels of the
Moon and Mars, and configuration complications arising from aerobraking upon descent at Mars,
tend to drive LEV and MEV designs apart. The LEV and MEV are likely to come on line years
apart in any case, and we found a more productive way to introduce commonality.
We found it useful to consider commonality at three program levels: (1) mission design,
using the same mission design to accomplish different programmatic architectures (a problematic
category because mission designs are by clef'tuition tied to unique requirements); (2)functional
element, using end-items from the same production line to f'dl different roles within a given
mission design (the most appropriate example we developed is the evolutionary LTF described
earlier); and (3) performing subsystem, using system assemblies or components from the same
production line in different functional elements (a sensible way to standardize industry, get
predictable performance and facilitate product longevity). This latter approach, applied to engines,
sensors, processors, some structural components and modular life support hardware, shows great
promise for cost-effectiveness and preserving program resiliency. At the component level,
extensive commonality can be worked into the fabric of SEI.
Another application of this subsystem/component commonality-for-evolution approach is
the potential use of hardware systems developed mostly for lunar transportation, augmented by
long-ditration crew systems, for early Mars missions staged out of high-orbit node locations Like
Earth-Moon L2. TMI AV for this mission mode is such that the need for a large TMIS is obviated
altogether, as is the need for a large cryogenic space engine. The use of chemically-propelled
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V. Commonality/Evolution
B. Program Scale Evolution - The lunar programs are classified into three scaleswhich utilize the transportation architectures in varying degrees. Each scale has a definedobjective and is limited to accomplishing the goal by varying assumptions. The smallscale scenario is a man-tended scientific outpost. Crews of four will remain on the lunarsurface overnight in a small "campsite". The medium scale develops a permanent basefor a crew of six. In addition to the scientific nature of the base, lunar oxygen wK! beproduced by the year 2010. The most ambitious program, the large scale, eventuallyleads to permanent settlement. The base continuously grows to accommodate 18 to 24
people by the year 2025. Along with the lunar oxygen plant, industrialization is veryimportant. Without specifying the type of the product, the industrial capabilities areequivalent to a I GWe Helium-3 plant.
After the program scales were established, a determination on required mass on the surfacewas determined along with cargo sizes and volumes for manifesting purposes. Eachprogram scale's mass-to-the-surface requirements vary depending on the amount of ISRU,number of inhabitants, mission durations, and level of industrialization. The mass for themedium scale program in the first 10 years is almost equivalent to that needed by the largescale program. The difference is seen in the subsequent 15 years of the two programs.The medium scale program continues without major modifications to the base or number ofinhabitants while the large scale program is always growing towards permanent settlementwith increasing personnel.
The number of cargo flights must be based upon the mass required by the program scale.Understanding the crew rotation before scheduling the cargo flights is very important. Inorder to build-up the base personnel, some flights will return to the Earth without all 6 crewmembers on-board. The small scale program requires 13 flights of cargo and cargo in 21years to accomplish its goals. The medium scale must have 56 flights and the large scalerequires 113 flights, both within 25 years. The cargo flights are manifested as tandemLTVs since they are capable of delivering 55 t in one mission. If LOR flights with lunaroxygen are used, the number of flights increases dramatically. These flights are capable ofdelivering 25 t of cargo when the LEV is already in orbit around the Moon, and 15 t if theLEV is delivered from Earth.
The end product of this analysis was a fairly detailed manifesting analysis. For the smallscale program, the flight rate of 13 flights in 21 years does not indicate the actual number ofETO launches and tankers that are needed to support the lunar program. For the mediumscale, approximately 700 t of cargo is required on the lunar surface by 2010. After thisdate, the tonnage is much less. By the end of the program, 56 flights of cargo and crewhave been made. At the end of the large scale program, more than 7,500 t of cargo is
placed on the lunar surface. Based on the amount of material manufactured at the base, 113flights are necessary to deliver the remaining cargo and crews. The HLLV (10 x 30 mshroud) flights are primarily based on the component masses. The aerobrake is volume-Limited and thus a separate flight. The LTV is designed to be launched intact in one flight.
The "campsite" and its associated LTV are also launched in a single flight. The propeLlantis launched separately and is transferred to the L'IWs while in LEO.
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