NASA Technical Memorandum 107295

AIAA-96-2810

High Leverage Space Transportation System

Technologies for Human Exploration Missionsto the Moon and Beyond

Stanley K. Borowski and Leonard A. DudzinskiLewis Research Center

Cleveland, Ohio

Prepared for the32nd Joint Propulsion Conference

cosponsored by AIAA, ASME, SAE, and ASEELake Buena Vista, Florida, July 1-3, 1996

J

National Aeronautics and

Space Administration

https://ntrs.nasa.gov/search.jsp?R=19960049765 2020-06-16T03:27:46+00:00Z

HIGH LEVERAGE SPACE TRANSPORTATION SYSTEM TECHNOLOGIESFOR HUMAN EXPLORATION MISSIONS TO THE MOON AND BEYOND

Stanley K. Borowski* and Leonard A. Dudzinski**NASA Lewis Research Center

Cleveland, OH 44135(216) 977-7091 and -7107

ABSTRACT

The feasibility of returning humans to the Moonby 2004, the 35 th anniversary of the Apollo 11landing, is examined assuming the use of existinglaunch vehicles (the Space Shuttle and Titan IVB),a near term, advanced technology space trans-portation system, and extraterrestrial propellant--specifically "lunar-derived" liquid oxygen orLUNOX. The lunar transportation system (LTS)elements consist of an expendable, nuclearthermal rocket (NTR)-powered translunar injection(TLI) stage and a combination lunar lander / Earthreturn vehicle (LERV) using cryogenic liquidoxygen and hydrogen (LOX/LH2)chemical

propulsion. The "wet" LERV, carrying a crew of 2,is configured to fit within the Shuttle orbiter cargobay and requires only modest assembly in lowEarth orbit. After Earth orbit rendezvous anddocking of the LERV with the Titan IVB-launchedNTR TLI stage, the initial mass in low Earth orbit(IMLEO) is ~40 t. To maximize mission performanceat minimum mass, the LERV carries no return LOXbut uses ~7 t of LUNOX to "reoxidize" itself for a"direct return" flight to Earth followed by an"Apollo-style" capsule recovery. Without LUNOX,mission capability is constrained and the total LTSmass approaches the combined Shuttle-Titan IVBIMLEO limit of ~45 t even with enhanced NTR andchemical engine performance. Key technologiesare discussed, lunar mission scenarios described,and LTS vehicle designs and characteristics arepresented. Mission versatility provided by using asmall "all LH2" NTR engine or a "LOX-augmented"derivative, either individually or in clusters, for outerplanet robotic orbiter, small Mars cargo, lunar"commuter", and human Mars exploration classmissions is also briefly discussed.

* Ph.D./Nuclear Engineering, Senior Member AIAA"'Aerospace Engineer, Member AIAA

INTRODUCTION

In January 1996, NASA issued its strategic planfor the Human Exploration and Development ofSpace (HEDS). 1 The HEDS enterprise envisionsan exciting future where humans travel routinelyand economically'to and through space" to theMoon initially, then on to Mars and other planetarybodies in our Solar System. Preceding a humanlunar return, robotic orbiter missions like LunarProspector2 will first map the chemical compositionof the lunar surface to help identify sites forpossible outposts and settlements that are rich inextraterrestrial resources. Teleoperated surfacelander experiments would then demonstrate theability to extract these resources and process theminto propellants and life support gases allowinghumans to "live off the land." In addition to "in-situ"resource utilization (ISRU), efficient surface powergeneration, cryofluid storage and transfer andadvanced propulsion technologies will also berequired if sustained human presence atdramatically lower costs are to be achieved. Thesesame technologies will also help open the spacefrontier for commercial and industrial development.

Planning future human missions to the Moonand Mars in today's environment is particularlychallenging. Despite an expected decline in itsfuture budgets, NASA and its field centers areexamining a variety of mission architectures and"high leverage" technology options that could bedeveloped to send humans "back to the Moon"before 2005 and then on to Mars no later than2018. Prospects for an early human lunar returnmission in 2001 for under $1 billion dollars3 ispresently being examined by engineers at theJohnson Space Center. Designing such a missionappears particularly challenging given that theInternational Space Station will still be in its finalassembly phase in the 2001-2002 timeframe andthat the Lunar Prospector Discovery mission isestimated to cost $ 63 million dollars. 2 Futhermore,

with decreasing budgets in its future, NASA mustinvest its scarce resources wisely and cannot affordto waste them on developing "short shelf fife"technologies to conduct a few "see we can do it "missions. The agency would then have to invest innew technologies to do the "Real Lunar Program,"as well as, initiating the follow-on "Mars Program" inthe post-2010 timeframe. The challenge to NASAtherefore, is to identify and develop the fewestnumber of high leverage technologies that canaccomplish most of NASA's planned missions(e.g., cargo and piloted, lunar and Mars, as well as,robotic science missions) and do it at an acceptablecost.

This paper examines the feasibility of returninghumans to the Moon as early as 2004, the 35 hanniversary of Apollo 11, using currently existingU.S. launch vehicles (the Space Shuttle and theTitan IVB), and a near term, advanced technologyLTS consisting of a small NTR-powered TLI stageand a cryogenic, integrated lunar lander/ Earthreturn vehicle (LERV) carrying a 2-person crew.The LERV is designed to utilize LUNOX propellantproduced earlier using facilities and lunar surfacevehicles operated telerobotically from Earth. In theApollo Program, the heavy lift "Saturn V" launchvehicle was used to place -140 t into LEOconsisting of the S-IVB TLI stage and its payload--the "3 person" Command and Service Module(CSM) and the Lunar Excursion Module (LEM)weighing ~32 t and 16 t, respectively (Figure 1). Asingle 200 thousand pounds thrust (200 klbf)LOX/LH2 J2 engine (specific impulse, Isp~425 s)

powered the S-IVB stage while storable, pressure-fed, hypergolic propellants were used in both theCSM (Isp~314 s) and the LEM (Isp-305 s). A "lunarorbit rendezvous" (LOR) mission profile was alsoadopted for Apollo because of its mass efficiency.

The LTS discussed here uses a single 20 t-class LERV with LOX/LH2 chemical engines and a

"lunar direct" mission profile which provides "globalaccess" to the Moon and an "anytime abort"capability for the crew once the LERV is refueled,or more appropriately, "reoxidized" with LUNOXpropellant. Because ~7 kilograms of mass isrequired in LEO for each kilogram of massdelivered to the lunar surtace, the utilization ofLUNOX from the outset enables a LERV withrobust payload / vehicle margin while maintainingsize and mass compatible with that of the SpaceShuttle orbiter. To deliver the LERV to therequired TLI conditions in a mass efficient manner,

a small (~15 klbf), "all LH2" NTR transfer stage, also

weighing ~20 t, is utilized. Although the principalfocus of this paper is on the LTS elements for theinitial human lunar return mission, the technologiesand systems described here provide the basis for aLTS that can evolve with time from an expendablearchitecture to one where reusable lunar landingand transfer vehicles (LLVs and LI-Vs) benefit fromthe production and utilization of LUNOX both onthe lunar surface and in low lunar orbit (LLO). Onceavailable in LLO, even conventional LH2 NTR

vehicles could benefit from LUNOX usage byoutfitting them with an oxygen propellant module,feed system and "LOX-afterburner" nozzleallowing bipropellant operation and revolutionaryperformance4 enhancements equivalent to thatfound in "gas-core" NTR engines.

This paper describes system and missionanalysis results performed by Lewis ResearchCenter's Advanced Space Analysis Office over thelast nine months in support of NASA's HumanLunar Return Study and the Space TransportationStrategic Plan. The paper first discusses the keytechnologies assumed in our study and describestheir characteristics. Mission and transportationsystem ground rules and assumptions are thenpresented along with a description of a referencelunar mission scenario. The lunar transportationsystem elements are then described and thebenefits of using LUNOX to increase deliveredpayload while decreasing vehicle size areillustrated through comparison with a system usingonly Earth-supplied LOX. The paper concludeswith a brief discussion of the applicability /evolvability of the small NTR engine and a "LOX-augmented" derivative for other NASA missions.

LUNOX. CRYOGENIC LUNAR LANDER AND NTRPROPULSION TECHNOLOGY DESCRIPTION

LUNOX Production and Utilization

Lunar - derived oxygen (LUNOX) has beenidentified5 as the most promising initial resource tobe developed on the lunar surface. A local sourceof LOX could replenish life support systems, andfuel cells used to power electric surface vehicles.Most importantly, the ability to provide the LERV's"oxygen-rich" chemical rocket engines (whichtypically operate at mixture ratios of ~5 to 6) with asource of return propellant oxidizer reduces thesize and mass of the LERV, as well as, the TLI

• Apollo Program- Saturn V .._

- Lunar Orbit "_,t_ l//Rendezvous = 7.6 m

Command Service 5-Module (-32 t)

Lunar Excursion

Module (-16 t)

6.6 m--_F,_

Saturn IV B

TLI Stage

J2 at 200K Ibf

_4.9 m

• Early Human Lunar Return- Shuttle and Titan iV B

- Lunar Direct

.4--- g.1 m---P-

4.0 m

CryogenicIntegrated Lander/Earth Return Vehicle

(LERV) (-20 t)

Small NTR

TLI Stage(- 20 t)

NTR at 15K Ibf

Fig. 1

IMLEO - 40 tIMLEO - 140 t

Lunar Transportation Systems Relative Size and Mass--Then and Now

stage which provides the LERV with its required

injection velocity. This "feedback effect" can cut

the required mission IMLEO by a factor of 2.

Oxygen is also attractive as a resource

because it is abundant in the lunar regolith (~45%

by mass) 5 and can be extracted using a variety o!techniques.6 Two of the more promising concepts

for oxygen production involve hydrogen reductionof ilmenite (FeTiO3) in high-titanium mare soil and

ferrous iron in volcanic glass. 7 Oxygen production

via hydrogen reduction is a two step process. First,the iron oxide (FeO) in ilmenite or volcanic glass is

reduced and oxygen is liberated to form water:

Fe-riO3(s) + H2(g) "-J" Fe(s) + TiO2(s) + H20(g)

orFeO(s) + H2(g) --_.. Fe(s) + H20(g)

Next, the water vapor is electrolyzed to regenerate

the hydrogen reactant and oxygen resource. The

hydrogen is recycled back to react with more lunarfeedstock while the oxygen is liquified and stored

in "well-insulated" storage tanks.

Reduction expe riments7 on samples of

high-titanium mare soil and iron-rich volcanic glasscollected during the Apollo 17 mission to the

Taurus-Littrow region of the Moon have produced

significant amounts of oxygen. Yields of~3.0 weight percent (wt %) have been measured

for ilmenite-rich, titanium soil at a reduction

temperature of ~1050 C after 3 hours. Using theiron-rich "orange" volcanic glass, a yield of

-5.1 wt % was achieved at ~1100 C over the same

time period. These experimental results suggestthat iron- and titanium-rich soils and iron-rich

glasses, in particular, would be attractive feed-stocks tor lunar oxygen production. In addition to

each existing in large quantities on the Moon, both

are fined grained and friable and could be usedwith little or no processing prior to reduction. A

4 to 5% extraction efficiency (~20 to 25 t of lunar

feedstock per ton of LUNOX) represents an order

of magnitude improvement in oxygen yield overearlier estimates. 8 This is very important because

reduced mining and beneficiation of bulk regolith

can lower the mass and power requirements of a

LUNOX production plant and its support vehicles.

Because the lunar mission architectureexamined here assumes that LUNOX is available tosupport the first piloted mission, cargo missions willbe required in advance to establish the necessarymining infrastructure on the lunar surface. It isenvisioned9 that initial cargo flights will deliver theLUNOX production plant and nuclear powersupply. A reactor sytem is preferred because itallows operation during the lunar night and is lessmassive than a photovoltaic system with energystorage. The reactor would be mounted on a smallteleoperated cart and transported a safe distanceaway from the LUNOX facility before powergeneration begins. Subsequent flights woulddeliver high duty-cycle, electric vehicles for loadingand handling regolith and for transportingLUNOX from the production site to the LERV(see Figure 2).

The mass and power requirements for ateleoperated lunar production facility providing~10.5 t of LUNOX per year (enough for 1 pilotedmission per year with a 50% reserve) are estimatedto be -4.8 t and ~35 kWe (~312 kg LUNOX / kWe ),

respectively. Additional system element massesfor the above production facility include the nuclear

reactor system at -3.0 t, 2 regolith loaders andhaulers at ~3.5 t and ~1.9 t, respectively, and 2LUNOX tanker vehicles at ~2.9 t.

Because the LUNOX scenario places suchheavy reliance on teleoperated facilities andsupport vehicles for mission success and crewsafety, it is important that such systems berigorously tested first on the ground and then indress rehearsal on the Moon before committing tothe actual piloted mission. NASA Lewis is studyingthe requirements for a demonstration test of asubscale teleoperated =LUNOX tanker" and itsmission functions under lunar vacuum and thermalconditions at its Plum Brook Space Power Facility(SPF), the world's largest space environment testchamber.lO Using a currently existing "quonsenthut" support structure outfitted with a cold wall andquartz heating lamp arrangement to simulate lunarday / night thermal conditions, the tanker demowould receive oxygen trom a LOX donor tank, thenstore, transport and transfer its LOX cargo to areceiver tank simulating the LERV (illustrated inFigure 3). These functions would be conductedduring lunar morning, noon and nighttime thermalconditions to determine insulation effectiveness,

Fig. 2 "LUNOX" Facility and Teleoperated Support Vehicles

4

I

/

r(SPl_

Fig.3

a)FLO Cryo LERV SEI Cryo LLV APolloLEM

Lunar Direct b) Lunar Direct ¢) LOR d) LORIMLEO ~93t IMLEO ~20t IMLEO-321 IMLEO ~16t

Fig. 4 Relative Size and Mass of Proposed Cryogenic LunarLander Vehicles and the Apollo Lunar Module

direct Earth entry at mission end. The FLO studyalso adopted a "lunar campsite" strategy in which apre-integrated, reusable habitat module wasdelivered to the Moon in advance ol the crew. Themass of the hab module needed to support thecrew for -45 days on the Moon (a lunar day, night,day cycle) was ~36 t which resulted in the need fora large cryogenic lander stage weighing -57 t.

In the LERV concept examined here(Figure 4b), features from both the '_J0-Day" andFLO lunar studies are adopted. A single stage,cryogenic lander is utilized to reduce the systemcomplexity and added mass of a second stage, andto achieve the high performance needed for the"lunar direct" mission profile assumed here. A 2person crew and short duration (-3 days) stay onthe Moon, reduces the requirements on / andmass of the crew capsule subsystems, and allowsfor cryogenic propellant usage with acceptableboiloff levels. A "common" crew module / Earth

return capsule allows direct reentry at mission endand reduces the size and mass of the LTS

elements, which would grow if reusability wasimposed on the scenario.

The technologies to support a cryogenic LERVare presently under active development in otherNASA and ESA programs. In the Delta Clipper-Experimental Advanced (DC-XA) program, 14technologies are being tested for future single-stage-to-orbit vehicle concepts. The "vertical take-off and landing" DC-XA vehicle has alreadydemonstrated feasibility for many of the systemsrequired for a LERV like a throttleable LOX/LH2

engine--the RL10A-5 provides the DC-XA with a

throttling capability from -30% to full thrust.Other advanced "weight-saving" technologiesinclude a Russian-built aluminum lithium (AI/Li)LOX tank, a graphite epoxy composite LH2 tank,composite intertank structure, and an integratedliquid-to-gas O2/I-12 RCS system to control the DC-

XA through its various airborne maneuvers in apropellant efficient way. It is possible that the DC-XA vehicle could serve as technology testbed for aLERV program in the near future. Lastly, ESA iscurrently developing a Crew Transfer Vehicle(CTV) and a smaller, unmanned AtmosphericReentry Demonstrator (ARD), to be flight tested in1997, which is derived from an "Apollo-style"command module, is The ARD will test lightweightconstruction materials, heat shield tiles andcoatings and demonstrate a parachute / recoverysystem for the 2.8 t ARD capsule. Many of thesetechnologies and possibly a "CTV/ARD-derivative"vehicle could form the basis for the LERV capsule.

Conventional and "LOX-Auamented" NuclearThermal Rocket (NTR! PmDulsion Technoloay

NTR propulsion is the key to providing "/owcost access through space" for human explorationmissions of the future. Unfortunately, NASA'sinvestment in this technology has been fleeting atbest--a Nuclear Propulsion Office existed briefly atLewis Research Center from 1991 through 1993--despite the NTR's proven performance. During theRover/NERVA (Nuclear Engine for Rocket VehicleApplication) nuclear rocket programs (1955-1973), 16 a total of twenty rocket reactors weredesigned, built and tested "open air" at the NuclearRocket Development Station at the Nevada Test

Site. These integrated reactor/engine tests,using

LH2 as both reactor coolant and propellant,

demonstrated a wide range of engine sizes (-50 to

250 klbf), high temperature graphite fuel providing

substantial hydrogen exhaust temperatures

(-2350-2550 K), sustained engine operation (over

60 minutes for a single burn) and restart capability--over 20 startups / shutdowns on the same engine.

What's new about NTR propulsion today thatwarrants a renewed investment in this technolgy ?

The answer lies in a reduced size, higher

performance engine that can be ground tested at

full power in a "contained facility" meeting current

environmental regulations. At present, LewisResearch Center is studying the benefits and

development costs of a small (10-15 klbf) NTR

engine that would use improved, high

temperature, tricarbide fuel. The fuel, developed in

the nuclear rocket program of the former Soviet

Union known today as the Commonwealth of

Independent States (CIS), 17 is capable of

producing hydrogen exhaust temperatures inexcess of 3100 K. Design studies, conducted by a

joint US/CIS industry team18,19 and funded by theNuclear Propulsion Office between 1992-1993,

produced a small advanced NTR engine with

impressive parameters: thrust -15 klbf, Isp~940-

960 s, engine thrust-to-weight~3.1, and enginefuel lifetime of -4.5 hours. Recent development

schedule projections and cost estimates by LewisResearch Center and Idaho National Engineering

Lab (INEL) indicate that a small 15 klbf NTR enginecan be developed, tested in INEL's "Contained

Test Facility" (CTF) with "scrubbed" H 2 exhaust,

and integrated into a small flight stage in -7 yearsat a cost of -$2.5 billion including the first flight

unit. Recurring stage cost for subsequent missionswas estimated to be -$150 million.

Although not the focus of this paper, an

enhanced version of the NTR, which leverages the

benefits of LUNOX, is being pursued at Lewiswhich combines conventional LH2-cooled NTR

and supersonic combustion ramjet (scramjet)

technologies to form a LOX-augmented NTR

(LANTR) engine. 4 The LANTR concept (Figure 5)

utilizes the large divergent section of the NTRnozzle as an "afterburner" into which LOX is

injected and supersonically combusted withreactor-heated hydrogen emerging from theLANTR's choked sonic throat--'scramjet propul-

Reactor ..-, Nozzle --I

Core _ Throat\_ I - _ / Supersonic

L _ LOX

Hydrogen _/_" '_ '_.-_""_ _'_'_--_- and ....._ _ -,,.j,_SupersomcCoolant/ _ _ _rsoni¢'_- _ T;'rust

Propellant _ _,,__ _ Augm_entation __-//___ ,ox ('r,x+ ->3,500K

j InjectionSubsonic

Hot H2

(Tex ~ 2,500 to 2,900 K)

Life (hrs)Tc (°K)

O/H MR = 0.0

1.0

3.05.07.0

Isp (sec)

5 10 30

2,900 2,800 2,600

941 925 891772 762 741647 642 631576 573 566514 512 508

Tankage

Fraction (%)

14.07.44.13.02.5

*For 15K Ibf LANTR with chamber pressure = 2,000 psia and _ = 500 to I

T/WengRatio

3.0*

4.8

8.211.013.1

Fig. 5 Schematic / Characteristics of LOX-Augmented NTR

7

Yeerly "1Direct Flight" Orb#or Mimaloem:_m (2.3 yean),

S_po Unto,- (U yore), Neptune (12.6 yore)

Titan 4B /ex. Ptuto Jupiter

Sckmce Miamions-_, Orb4Mr OrlMterM_M______$ _ ,, .Pluto

December 2004 June 2006 2018

Technology CTF at INEL B Right Tosl F_

Validation Available i _ .,.,,,cqo m,..,,_, .'rv

_.r'_elino

1997 1,98 1999 _2000 _ 2004_005 "_0 2015 _ 2020 _*_2025 -

2O30

(- 2000) (2003) (2010) (2016-18) {- 2¢Z5)"Early HumanLunar Return"

(2O04)

Fig. 6 NTR Development / Utilization Road Map

sion in reverse." By varying the LOX-to-LH2 mixture

ratio (MR), the LANTR engine can operate over awide range of thrust and Isp values while thereactor produces a relatively constant poweroutput. Thrust augmentation means that "bigengine" performance can be obtained usingsmaller, more affordable, easier to test NTRengines. The increased use of high-density LOX inplace of low-density LH2 also reduces hydrogentank volume. This feature provides importantflexibility to vehicle designers allowing smallLANTR-based transfer stages to be configured toaccommodate "mass- and/or volume-constrained"

launch vehicles.20 Finally, once LUNOX becomesavailable in LLO to "reoxidize" LANTR-basedtransfer stages, "low cost" lunar transportation willbecome commonplace along with a host of othermission possibilities. A possible "road map"detailing the near term development and ultimateutilization of both conventional and LOX-

augmented NTR systems is illustrated in Figure 6.

LUNAR MISSION / TRANSPORTATION SYSTEMGROUND RULES AND ASSUMPTIONS

The ground rules and assumptions for thereference and alternative lunar mission scenariosexamined in this study are summarized in Table 1.Provided are details on outbound and return

payloads, parking orbits, mission velocity change

(z_V) requirements and duration, and launchvehicle characteristics. In addition to the primaryAV maneuvers indicated, midcourse correctionmaneuvers (AV = 30 m/s) are also performed usingeither storable or gaseous 02/H2, bipropellant

RCS systems. Tables 2 and 3 detail assumptionsfor the NTR-powered TLI stage and the chemicalpropulsion LERV, respectively, on primary andsecondary propulsion, cryogenic tankage, thermalprotection and boiloff rates, primary structure, andcontingency factors used in this study.

An aluminum-lithium alloy "Weldalite"(Ftu = 111 ksi, p = 0.0976 Ibm/in3 = 2700 kg/m3) has

been used for the cryogenic LOX and LH2 tanks

and a graphite / epoxy composite IM7/977-2(Ftu =91 ksi, p = 0.057 Ibm/in3 = 1577 kg/m3) for

non-pressurized primary structures. Wallthicknesses for the LH 2/LOX tanks were calculated

based on a 35 / 50 psi internal pressure andincluded hydrostatic loads using a "4 g" load factoralong with a safety factor of 1.5. A 2.5% ullage wasalso assumed in this study.

A 1.5 inch helium-purged, multilayer insulation(MLI) system (at 50 layers per inch) is assumed forthermal protection of the TLI stage LH 2 tank while

2 inches are applied to the LERV's LOX and LH2tanks. These insulation thicknesses reduce boiloffto acceptable levels and also satisfy the "ground

Table1. ReferenceLunarMissionGroundRulesandAssumptions

PayloadOutbound: 3.40 - 4.50 t Crew Capsule0.45 t Crew (2) and Suits

0.05 - 0.30 t Surfacepayload

Payload Inbound: 3.40 - 4.50 t Crew Capsule0.45 t Crew (2) and Suits

0.01 - 0.05 t LunarSamples

• Parking Orbits: 185 km110-300 km

Circular (Earth departure)Circular (lunar arrival/departure)

• Trans-lunar injection AV: 3145 rn/s + g-losses• Lunarorbit capture / trans-Earth injection_V: 1050m/s• Lunar descent / lunar ascent 2_V:2000 m/s / 1900m/s• Direct lunar descent / direct ascent and Earth return AV: 2510 rn/s• NTR TLI stage disposal AV: 30 rn/s (lunar swingby and gravity assist)• Mission duration: 12-14 days (2-4 in LEO, 7 in transit, 3 days at/on Moon)• Launch vehicle type: Space Shuttle and Titan IVB• Payload Delivery Capability: 25.4 t and 21.6 t to 28.5 deg. inclination• LTS assembly scenario: Space Shuttle then Titan IVB launches with EOR & D (IMLEO: - 40 to 45 t)

Table 2. NTR TLI Stage System Assumptions

NTR System:

RCS System:

Thrust / WeightFuel / PropellantIspExternal Shield Mass

Flight ReserveTrapped ResidualsCooldown (effective)

PropellantIspTankage

Cryogenic MaterialTankage: Diameter

GeometryInsulation

LH 2Boiloff*

= 15 klbf / 4904 Ibm

= Tricarbide / Cryogenic LH2= 940 s (@ 2900 K) / 960 s (@ 3025 K)= 2.76 kg/MWt of reactor power= 1% of usable LH2 propellant= 1% of total tank capacity= 3% of usable LH2propellant

• Primary Structure:

• Contingency:

= N2OJMMH= 320 s= 5% of total RCS propellants

= "Weldalite" AI/Li alloy= 4.6m

= Cylindrical tank with 1'_2 domes= 1.5 inches MLI + micrometeoroid debris shield

= 1.75 kg/m2/month (LEO @ ~ 240 K)= 10.73 kg/day (for 12.4 m tank length)= 11.74 kg/day (for 13.6 m tank length)

Materials = "Weldalite" AI/Li and Graphite/Epoxy Composite

Engine, shield, and stage dry mass = 15%

" Assumes 3 x "Lockheed Eqn" heat flux estimates for MLI At < 2 inches

9

hold"thermalprotectionrequirementsfor "wet-launched"LH2tanksof 1.5inches.21Theinstalleddensityofthe 1.5inchMLIsystemis-2.05kg/m2,andtheresultingLH2boiloffratefromtheTLIstagewhilein LEOis -1.75kg/m2/month(basedonanestimatedheatfluxof -0.294W/m2ata LEOsinktemperatureof -240 K).FortheLERV,maximumboiloffratesforLOXandLH2havebeenestimatedassuminga lunarsurfacetemperatureof -394 K(equivalentto "early afternoon" on the Moon). Atthis temperature, the estimated heat flux is-1.1 W/m2 and the boiloff rates for LH 2 and LOX

are -6.51 and 13.26 kg/m2/month, respectively.Also shown in Table 3 are boiloff rates for the

LERV propellants in LEO, cislunar space and LLO.Lastly, a 0.25 mm thick sheet of aluminum(corresponding to -0.682 kg/m2) is included in thetotal tank weight estimates to provide protectionagainst micrometeoroids and regolith backscatterfrom engine exhaust during landing and takeoff.

LUNAR MISSION SCENARIO DESCRIPTION

The reference lunar scenario examined in thisstudy assumes a split mission architectureinvolving both cargo and piloted missions operatedinitially in an "expendable mode" in order tomaximize payload delivered to the Moon on eachmission, maintain a "two launch" mission capability(IMLEO limited to -40 - 45 t), and reduce the LTSsize and cost. The piloted mission employs a"lunar direct" flight profileg,22 and assumes theavailability of LUNOX for lander refill and Earthreturn. The mission flight profile is illustrated inFigure 7. The Space Shuttle and Titan IVB, with acombined payload delivery capability to LEO of-45 t, are used to deliver the LTS elements to the185 km (100 n. mi.) staging orbit. The LERV's"wet" lander stage and crew capsule with acombined mass of 20 t are delivered first in the

orbiter's payload bay "side-by-side." Once in orbitthe capsule is installed atop the lander and the

Table 3. LERV Transportation System Assumptions

• PrimaryPropulsion:

Total Thrust = 12.5 - 15.0 Idbf

Propellant = LOX/LH2Isp = 450 s (@ O/H MR = 5.2)

= 465 s (@ O/H MR = 6.0)Flight Reserve = 1% of usable propellantTrapped Residuals = 1% of total tank capacity

• RCS System: Propellants = N204/MMH and GO2/GHzIsp = 320 s and 400 sTankage = 5% of total propellants (storable)

• Cryogenic MaterialTankage: Geometries

Insulation

LH2/LOX Boiloff*

= "Weldalite" AI/Li alloy= Cylindrical with 1/'_2 domes/spheres= 2.0 inches MLI + micrometeoroid debris shield

= 1.31 / 2.44 kg/mZ/month (LEO @ - 240K)= 0.56 / 0.90 kg/m2/month (in-space @ - 172K)= 1.91 / 3.68 kg/m2/month (LLO @ - 272K)= 6.51/13.26 kg/m2/month (LS @ - 394K)

• Primary Structure: Materials = "Weldalite" AI/Li and Graphite/Epoxy Composite

• Contingency: Dry Stage Mass = 15%

*Assumes 3x "Lockheed Eqn" heat flux estimates for MLI t_t

LuaarSvdqbyAssist

NTR

DispeNd to

He_C Orbit

(~100,000 Years)

CrewCapsaleLERVSepaa,atJoa

LERV

NTR/LEBV1 STS:LERU

SeparaUon_ 11"_ NB:HR

Fig. 7 Reference "Lunar Direct" Mission Scenario Using NTR, LERV and LUNOX

_ntegrated LERV's systems are checked out. Next,

the NTR-powered TLI stage, also weighing -20 t, islaunched to LEO by the Titan IVB where Earth orbit

rendezvous and docking (EOR&D) with the LERV

results in the integrated spacecraft shown in

Figures 7 and 8.

Over the next 1 to 2 days, the integrated LTS

is checked out in preparation for the TLI launch

window. At the appropriate time the NTR engine is

powered up and performs a sustained 28 minuteTLI burn which sends the coupled spacecraft on a

trajectory to the Moon. Gravity losses for the"single burn" Earth departure scenario are

estimated to be -205 m/s. Following an

appropriate cooldown period (-5 hours) for the

NTR engine, the piloted LERV and TLI stage

separate with the LERV continuing on its nominal

mission while the NTR stage executes a

retargeting maneuver with its RCS system to

perform a "trailing edge" lunar swingby. The

11

resulting lunar gravity assist is used to deliver the

"spent" NTR stage to a long-lived (-105 year) helio-centric orbit with minimal risk of Earth reencounter.

Following a 3.5 day transit, the LERV uses its

four LOX/LH2 engines to propulsively capture into

a temporary parking orbit around the Moon.

Pausing here allows time for navigational updates

and phasing alignment over the desired landingsite, and most importantly, time to communicate

with the LUNOX production facility and its tanker

vehicle to verify that all is ready for the LERV's

arrival. Should equipment problems arise on thesurface, the LERV would abort the landing and

return to Earth using the remaining onboard

landing propellant (equivalent AV=2000 m/s).

At the appropriate time, the LERV, with its 2

person crew, deorbits and lands on the lunarsurface--an event which would be televised world-

wide by cameras onboard the waiting LUNOX

tanker. The LERV consists of a "state-of-the-art"Apollo-style capsule mounted atop a combinationservice module / lunar lander. It lands on the lunarsurface with its LOX tanks essentially "empty" butwithin a reasonable distance of the LUNOXproduction facilities predeployed on an earliercargo flight. The LERV does carry sufficient LH 2fuel for the return trip back to Earth however.Shortly after landing, the crew's first major activity isto remotely activate and direct the nearby surfacetanker to transport and refill the LERV's LOX tankwith -7 t of LUNOX supplied by the productionfacility. Once reoxidized, the LERV is ready toleave should an emergency situation arise over thenext 3 days. Because the LERV is only designedto support its crew for 3 days on the lunar surface,extended staytimes will require a surface habitationfacility be delivered on an earlier cargo mission.Near the end of the surface stay, the crew with itssamples, ascends to a temporary lunar phasingorbit, then performs the trans-Earth injection (TEl)burn for the trip back. Near Earth, the crew moduleseparates from the lander stage and performs adirect Earth entry while the lander stage isexpended in cislunar space via an Earth fly-by.

LTS VEHICLE DESCRIPTIONS / COMPARISONS

The relative size and mass of the NTR transfer

stage and chemical LERV with and without LUNOXusage is shown in Figure 8. Table 4 summarizesLERV system assumptions and mass breakdownboth in LEO and on the lunar surface (IMLS) priorto Earth return. The mass elements include thesuited crew, surface-landed payload (S/PL), theLERV's return crew capsule (LERC), and lunarlanding stage, and returned lunar samples orpayload (R/PL). The required amounts of Earth-supplied LOX and LH 2, as well as, quantities ofLUNOX for ascent and Earth return are also shown.

Figure 8a shows the reference piloted vehicleconfiguration for the "LUNOX scenario" discussedabove. After EOR&D, the total vehicle length andmass is -25 meters and -40 t, respectively. TheTLI stage itself is -18.7 m long but can easily beaccommodated by the Titan IVB booster with a76 foot (-23.2 m) long payload fairing. The NTR-powered TLI stage uses a single 15 klbf engine,dual turbopumps for improved reliability, andternary carbide fuel elements. At the hydrogenexhaust temperature and nozzle inlet pressure of2900 K and 2000 psia, respectively, the specificimpulse is -940 s using a nozzle expansion ratio of

12

300 to 1. Other elements of the NTR TLI stageinclude: (1) an external radiation shield for crewprotection; (2) a 4.6 m diameter by 12.4 m long LH2tank; (3) a forward cylindrical adaptor housing theRCS, avionics, auxiliary power, and dockingsystems; (4) forward and aft cylindrical band skirts;and (5) a conical thrust structure. The TLI stage"dry" mass is -7 t which includes -3.64 t for the15 klbf NTR and shield. The LH2 and RCS

propellant loads total -13 t. Included in this total ispropellant to perform a small (-30 m/s) "trailingedge" lunar swingby maneuver for stage disposalafter LERV separation. The stage is also providedwith an -4 kilowatt electric (kWe) fuel cell auxiliary

power system which can provide the stage with-2 kWe average power for up to 5.5 days. Table 5summarizes the mass properties of the TLI stageshown in Figures 7 and 8a.

The reference LERV uses four newthrottleable LOX/LH 2 engines (each at -6 klbf with

4:1 throttling) which operate at an O/H mixture ratioMR = 5.2 and Isp = 450 s. The LERV lander stageis set at 15 % of the "wet" LERV mass in LEO.Preliminary weight estimates of individualsubsystem elements show this assumption to beconservative and to provide ample mass margin. Intotal, the mass of the crew capsule and "dry" landerstage is -7.5 t or -37.5% of the LERV's initial massin LEO, and this weight can be redistributed asnecessary if dictated by future analysis. In additionto the crew, the LERV also transports -0.3 t ofequipment (-0.22 t for an "Apollo-style" lunar roverand -0.08 t of science equipment) to the lunarsurface. Power for the LERV during most of the 12day mission is supplied by two 4 kWe fuel cells

which provide -3 kW e average power, as well as,consumable water for the crew and coolant for theactive thermal control system. Fuel cell reactantstotaling -0.36 t, at an O/H ratio of 8 to 1, are storedwithin the LERV's propellant tanks. The LERVtransports -1.34 t of LH2 to the lunar surface (with

-40 kg being lost to boiloff over the 3 day lunarstay) and refills its LOX tank with -7 t of LUNOX toreturn the LERV with its -4.5 t crew capsule and-0.5 t of crew and lunar samples back to Earth. TheLOX, LH2 and storable RCS propellant loads are

-8.38, 2.95 and 0.34 t, respectively, for a totalIMLEO for the LERV of -20 t which includes-0.16 t for the LERV's TLI stage adaptor. Thesame LERV lander and TLI stage are also capableof delivering -6.6 t of lunar surface payload on"l-way" cargo missions. This is sufficient to

18.7In

a) "Lunar Direct" w/LUNOX

9.1 ¢

4.1[]

12.4m

b) "Direct Descent" w/o LUNOXE

3.8nl

19.9 m

_4,6

LH213.6 In

I ltil]l ,_

l. !15 klbl 4. m

oIMLEO:45t

Fig. 8 Relative Size/Mass of NTR Stage and LERV With / Without LUNOX

accommodate all of the envisioned

elements needed to support the piloted

mission scenario.

surfaceLUNOX

Without LUNOX, the LERV must carry its own

return LOX resulting in a substantial reduction in

mission capability. To stay within the IMLEO limit of

-45 t provided by the Shuttle and Titan IVB, a

"direct descent" trajectory profile23 must be

adopted (see Figure 9) to reduce the total mission

AV requirements (see Table l).While directdescent can improve performance in the cargo

flight mode, it is a higher risk trajectory profile for

piloted missions and does not provide for a "freereturn" abort. Lunar surface site accessibility and

lighting conditions flexibility is also degraded.

Propulsion system performance and vehicle sizemust also be increased (see Figure 8b). For the TLI

stage, the NTR fuel temperature is increased to

-3025 K to achieve a higher Isp of -960 s.

However, for a given reactor power level, thisincrease comes at the expense of a decrease in

thrust to -14.7 klbf. The TLI stage length and mass

also increase in order to accommodate the larger

LH 2 propellant load (-14.4t) needed to send a

heavier LERV (-23.4 t) to the Moon.

For the LERV's LOX/LH2 chemical engines, the

chamber pressure and mixture ratio are increased

to achieve a Isp of -465 s. A gaseous O2/H2 RCS

with Isp of -400 s is also assumed. This increase in

13

Table 4. LERV System Assumptions and Mass Characteristics

• Mission Profile:

• Propellant:• MR/Isp:• IMLEO/IMLS:

• Crew (2) & Suits:• MSIPL:• MtERc:

• Mm_,_o,:• MRCS prop:

• MLOX:- Outbound- Inbound- LS boiloff

• MLH2:

- Outbound- Inbound- LS boiloff

• IV_.u,ox:• MR_L:

Lunar Direct w/LUNOX

LOX/LHz5.2 1450 s

20.0 t / 16.25 t0.45 t0.30 t4.42 t

3.0t/0.16t0.19 t/0.15 t

8.38 t

1.61 t1.30 tO.04t6.78 t0.05 t

Direct Descent wo/LUNOX

LOX/LH26.0 1465s

23.56 t / 13.17 t0.45 t0.05 t3.36 t

3.5t/0.16 t0.18 t / 0.06 t

8.59 t4.88 t0.04t

1.43 t0.81 t0.04t

0.01 t

performance cannot compensate, however, forthe extra 5 t of return LOX and the heavier landerstage required to carry it. The result is a decreasein the available mass and therefore capability ofcrew capsule (from -4.4 to 3.4 t) , as well as, inpayload delivered to the lunar surface (seeTable 4). With no lunar rover and only -50 kg ofscience equipment available to the astronauts,EVA activities will be limited to short distances

Table 5. TLI Stage Mass Properties

Staqe Element Mass (ka)Structure 607Propellant Tank 1065Thermal Protection System 508Avionics and Power 293

Reaction Control System 227NTR Assembly

15 klbf NTR 2224External Shield 940Propellant Feed/TVC 171

Contingency (15%) 905Dry Mass 6940LH2 Propellant 12741RCS Propellant 100Wet TL I Stage Mass 19781

away from the LERV similar to that on the Apollo12 and 14 missions. Returned lunar samples willalso be limited to -10 kg, substantially less thanthat returned on the earlier Apollo missions.Although an initial 2004 human lunar returnmission appears doable without LUNOX,subsequent flights would greatly benefit from itsuse and help to restore lost mission versatility andvehicle robustness.

14

OTHER MISSION APPLICATIONS FOR THE NTR

Investing wisely in "high leverage" technologieswith growth capability and applications to moreambitious human missions beyond the 2004 lunarreturn discussed here is an important point not tobe overlooked. The small NTR stage is an excellentexample of a wise technology investment forNASA for with it a wide range of missions becomepossible (see Figure 10). In less than 6 monthsafter humans return to the Moon (see Figure 6),the same small NTR stage could be used todispatch a robotic orbiter mission to Pluto using asingle Titan IVB launch and a Jupiter gravity assistopportunity available in June 2006. Direct flight,"short transit time" orbiter missions to Saturn(2.3 years), Uranus (6.6 years), and Neptune(12.6 years) are also possible each year.24

DispOSedtOflei_ Orbit

(-100,000 Years)

Lu,' S.dmJy60,m Assist

CrewCaps

LEOAsSmaddyFlights:2

Oescest

y

I_ Stale

Trai_ry1 ST$:LERV

_ 1 TitanIVB:NTR

NTR/LEIIV

Separa'doa

Fig. 9 Alternative "Direct Descent" Mission Scenario Using NTR and LERV Only

Small NTR

4.5 m

Commercializstion

and Settlement

O/H MR*: 0 to 7

Thrust: 15 to 66K Ibf

Isp: 940 to 514 $ec

T/Weng: 3.0to 13.1

"For Lox-Augmenled NTR Option

Fig. 10

S_ngle Tlta_ 40

(IMLEO -20 t)

Storable Capture Stage

Comm(_ NTR

Inleclbon Stage

4.6 m

"m_3 c,wu.v

.__p_O_ '_F (IMLEO --40 I)

Deve_pment

4.6 m ' _ : --

_' . Mars Ca_

(IMLEO -35 t)

36 _ Lunar

Shuffle

Module

Ctustere(I

Small NTR's

t

- '_ ']!:;i' : = i (IMLEO -140 t)

Small NTR Mission Versatility: "One Size Fits All"

15

How far can we go with LANTR propulsion?

EaCh

.L

LH2and Lox H >'L__

Fig. 11 Human Expansion Possibilities With LOX-Augmented NTR (LANTR)

In preparation for the human exploration of Mars,small Mars cargo missions capable of delivering upto 10 t of surface payload can be carried out usingthe Titan IVB-launched small NTR stage for trans-Mars injection and the Space Shuttle for launchingthe Mars payload with its biconic-shaped aero-capture and descent system. Higher capacity cargoand Mars piloted missions would use the same15 klbf NTR engine in a clustered arrangement toprovide the optimum mission thrust level. A cargoversion of the current Space Shuttle (Shuttle C) oran "in-line" Shuttle-derived vehicle with a liftcapability of -70 t to LEO would probably beutilized for these missions to avoid the substantialLEO asembly requirements using "volume-limited"smaller capacity (-20 t-class) launch systems suchas the reusable launch vehicle (RLV).

With time, initial outposts on the Moon andMars will grow to centralized bases and settlementswith a substantial human presence. LUNOXproduction facilities on the lunar surface will supplypropellant depots in low lunar orbit and in tum a"LOX-augmented" NTR (LANTR)-powered LTSwhich will "revolutionize" cislunar space travel.Commuter flights2S to and from the Moon with"one-way" trips of 36 to 24 hours will becomecommonplace using LANTR-powered passengershuttles. On Mars, reusable LANTR-poweredlanders, operating from specially prepared landingsites, will transport significant quantities of payloadto and from a Phobos station and propellant depot.The Phobos depot would also provide reusableLANTR-powered Mars transfer vehicles with theirreturn propellant allowing them to transport morehigh value cargo to Mars instead of bulk propellantand expended lander/aerobrake hardware mass.

16

Beyond the Moon and Mars lies the asteroidbelt and the Jupiter system (see Figure 11) wherelarge quantities of water are believed to exist.Water ice has been detected on Ceres, the largestof the main-belt asteroids, and the Galileansatellites, Europa, Ganymede, and Callisto, areknown to possess large amounts of water ice ontheir surfaces from the Voyager missions. UsingISRU and LANTR technologies, extraterrestrialsources of LOX and LH2 can be developed and

utilized to facilitate human expansion into the SolarSystem. While such missions and capabilities arewondrous to imagine what is most exciting to theauthors is the fact that these technologies couldbe developed in the next 10 to 15 years!

ACKNOWLEDGEMENTS

The authors wish to express their thanks tomanagement (Joe Nieberding and Don Palac) fortheir support and encouragement during thecourse of this work. They would also like to expresstheir gratitude to the following individuals: KentJoosten (NASA JSC) for use of the artworkdepicted in Figure 2; Bud Mitchell (LeRC / PlumBrook Station) for the illustration in Figure 3; BobSefcik (NASA LeRC) for NTR schedule andcost projections; and Stephanie Black for herinvaluable assistance in preparing this paper.

RF_.£EBE_3E_

1. The Strategic Plan--NASA's Enterprise for the,Human Exploration and Development of Space,National Aeronautics and Space Administration,(January 1996).

2. L. David, "Moon Mission Has Hidden Goals,"

Space News, (April 8-14, 1996), p.31.

3. L. David, "Gotdin:Plan a Moon Trip for Less

Than a Billion Dollars," Space News, (November

27-December 3, 1995), p.10.

4. S.K. Borowski et al., "A Revolutionary Lunar

Space Transportation System Architecture UsingExtraterrestrial LOX-Augmented NTR Propulsion,"

AIAA-94-3343, American Institute of Aeronautics

and Astronautics (June 1994).

14. L. David, "NASA's Marshall Targets DC-XA

Flight Testing in April," Space News, (September

4-10,1995), p.18.

15. C. Covault, "Reentry Flight Test Set ForSecond Ariane 5," Aviation Week and Space

Technology, (March 25, 1996), pp.51-53.

16. D.R. Koenig, "Experience Gained from the

Space Nuclear Rocket Program (Rover),"LA-10062-H, Los Alamos National Laboratory

(May 1986).

5. T.S. Sullivan and D. S. McKay, Using Space

Resources, NASA Johnson Space Center (1991).

6. L.A. Taylor and W. D. Carrier Ill, 'q'he Feasibilityof Processes for the Production of Oxygen on the

Moon," in Engineering, Construction & Operations

in Space Ill, New York, 1992, pp.752-762.

7. C.C. Allen, R. V. Morris, and D. S. McKay,

"Experimental Reduction of Lunar Mare Soil and

Volcanic Glass," Journal of Geophysical Research,

99, 23, 173 (November 1994).

8. E.L. Christiansen, et al., "Conceptual Design

of a Lunar Oxygen Pilot Plant," Eagle Engineering

Report No. 88-182, Eagle Engineering, Inc.,

Houston, Texas (July 1, 1988).

9. B.K. Joosten and L. A. Guerra, "Early Lunar

Resource Utilization: A Key to Human Exploration,"AIAA-93-4784, American Institute of Aeronautics

and Astronautics (September 1993).

10. Lewis Research Center R & D Facilities, NASA

Lewis Research Center, Cleveland, Ohio.

11. A. Cohen, et al., Report of the 90-Day Study

on Human Exploration of the Moon and Mars,National Aeronautics and Space Administration

(November 1989).

17. J.S. Clark, M. C. Mcllwain, V. Smetanikov,

E. K. D'Yakov, and V. A. Pavshook, "US / CIS EyeJoint Nuclear Rocket Venture, Aerospace

America, Vol. 31, (July 1993).

18. D.W. Culver, V. Kolganov, and R. Rochow,

"Low Thrust, Deep Throttling, US/CIS Integrated

NTRE," 1 lth Symposium on Space Nuclear Power

Systems, Albuquerque, New Mexico, (January 9-

13, 1994).

19. S.K. Borowski, et al., "Nuclear Thermal

Rocket / Vehicle Design Options for Future NASAMissions to the Moon and Mars," AIAA-93-4170,

American Institute of Aeronautics and Astronautics

(September 1993).

20. S.K. Borowski, "Human Lunar Mission

Capabilities Using SSTO, ISRU and LOX-

Augmented NTR Technologies--A PreliminaryAssessment," AIAA-95-2631, American Institute ofAeronautics and Astronautics (July 1995).

21. R.H. Knoll, R. J. Stochl, and R. Sanabria, "A

Review of Candidate Multilayer Insulation Systems

for Potential Use of Wet-Launched LH2 Tankage

for the Space Exploration Initiative Lunar

Missions," AIAA-91-2176. American Institute ofAeronautics and Astronautics (June 1991).

12. Analysis of Synthesis Group Architectures:

Summary & Recommendations, ExPO Document

XE-92-004 (May 1991 ).

13. S.K. Borowski and S. W. Alexander, "'Fast

Track' NTR Systems Assessment for NASA's

First Lunar Outpost Scenario," AIAA-92-3812,American Institute of Aeronautics and Astronautics

(July 1992).

22. R.M. Zubrin, "The Desigh of Lunar and Mars

Transportation Systems Utilizing ExtraterrestrialResources." IAA-92-0161, 43rd Congress of the

International Astronautical Federation,

Washington, DC, (August28-September 5,1992).

23. P. Bialla, "Early Low Cost Lunar Missions,"AIAA-93-4219, American Institute of Aeronautics

and Astronautics (September 1993).

17

24. S.K. Borowski, "Robotic Planetary ScienceMissions Enabled With Small NTR Engine/StageTechnologies," in Pro(;:. l:_th S_vmDosium onSpa(;;:e Nuclear Power and Pro[}ulsion. AlP Conf.Proc. No. 324, January 1995, Vol. 1, pp. 311-319.

25. S.K. Borowski, D. W. Culver, and M. J.Bulman, "Human Exploration and Settlement ofthe Moon Using LUNOX-Augmented NTRPropulsion," in Proc. 12th SvmDosium on SpaceNv(;:le_r Power and ProDulsion, AlP Conf. Proc.No.324, January 1995, Vol. 1, pp.409-420.

18

I Form Aopro vedREPORT DOCUMENTATION PAGE OMBNo. 0704-0188

Public repotting burden for this _ of informationis estimated to .average 1 hour pet response, including the timefor reviewinginsmJctlons, searchingexisting da_ imurc_,ga_heringand mJdntab_ingthe data needed, and comp4e_ngand reviewkngthe ¢o41ectJeno4information. Send comments regardingthis burden eatmale or any othoraspect of this¢o_lenlionof information,incluCrmgsuggestionsfix reducing this burden, to WashingtonHeadquattem Services. Directoratelot Infor_ Operations and Repocts,1215 JeffenmnDavis Highway. Suite 1204, Arlington,VA 22202-4302. and to the Office ol Management and Budget. Paperwork ReductionProject(0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) !2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

August 1996 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

High Leverage Space Transportation System Technologies for Human ExplorationMissions to the Moon and Beyond

e. AUrr_P4s)

Stanley K. Borowski and Leonard A. Dudzinski

7. PERFORMINGORGANIZATIONNAME(S)AND ADDRESS(ES)

National Aeronautics and Space AdministrationLewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

WU-_I_I

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-10373

10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

NASA TM- 107295

AIAA-96-2810

11. SUPPLEMENTARYNOTES

Prepared for the 32rid Joint Propulsion Conference cosponsored by AIAA, ASME, SAE, and ASEE, Lake Buena Vista,

Florida, July 1-3, 1996. Responsible person, Stanley K. Borowski, organization code 6850, (216) 977-7091.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified -Unlimited

Subject Categories 16 and 20

This publication is available from the NASA Center for AeroSpace Information, (301) 621--0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

The feasibility of returning humans to the Moon by 2004, the 35th anniversary of the Apollo 11 landing, is examined

assuming the use of existing launch vehicles (the Space Shuttle and Titan IVB), a near term, advanced technology space

transportation system, and extraterrestrial propellant--specifically "lunar-derived" liquid oxygen or LUNOX. The lunar

transportation system (LTS) elements consist of an expendable, nuclear thermal rocket (NTR)-powered translunar

injection (TL1) stage and a combination lunar lander/Earth return vehicle (LERV) using cryogenic liquid oxygen and

hydrogen (LOX/LH2) chemical propulsion. The "wet" LERV, carrying a crew of 2, is configured to fit within the Shuttle

orbiter cargo bay and requires only modest assembly in low Earth orbit. After Earth orbit rendezvous and docking of theLERV with the Titan IVB-launched NTR TLI stage, the initial mass in low Earth orbit (IMLEO) is --40 t. To maximize

mission performance at minimum mass, the LERV carries no return LOX but uses -7 t of LUNOX to "reoxidize" itself for

a "direct return" flight to Earth followed by an "Apollo-style" capsule recovery. Without LUNOX, mission capability is

constrained and the total LTS mass approaches the combined Shuttle-Titan IVB IMLEO limit of .-45 t even with enhancedNTR and chemical engine performance. Key technologies are discussed, lunar mission scenarios described, and LTS

vehicle _.signs and characteristics are presented. Mission versatility provided by using a small "all LH2" NTR engine ora "LOX-augmented" derivative, either individually or in clusters, for outer planet robotic orbiter, small Mars cargo, lunar"commuter", and human Mars exploration class missions is also briefly discussed.

14. SUBJECT TERMS

Space shuttle; Titan IVB; Nuclear thermal rocket; NTR; LOX-augrnented NTR; LUNOX;Human lunar return; Cryogenic lander/Earth return vehicle; LERV; Lunar direct

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Unclassified

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