Stanley K. Borowski

Glenn Research Center, Cleveland, Ohio

David R. McCurdy

Vantage Partners, LLC, Brook Park, Ohio

Laura M. Burke

Glenn Research Center, Cleveland, Ohio

The Nuclear Thermal Propulsion Stage (NTPS):

A Key Space Asset for Human Exploration and 

Commercial Missions to the Moon

NASA/TM—2014-218105

October 2014

AIAA–2013–5465
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Stanley K. Borowski

Glenn Research Center, Cleveland, Ohio

David R. McCurdy

Vantage Partners, LLC, Brook Park, Ohio

Laura M. Burke

Glenn Research Center, Cleveland, Ohio

The Nuclear Thermal Propulsion Stage (NTPS):

A Key Space Asset for Human Exploration and 

Commercial Missions to the Moon

NASA/TM—2014-218105

October 2014

AIAA–2013–5465

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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Acknowledgments

The authors express their thanks to Chris Moore and John Warren from NASA Headquarters, Mike Houts from NASA Marshall 

Space Flight Center, along with John Taylor and Gary Kelm from NASA Glenn Research Center for their continued encouragement 

and strong support of this work. The author Stanley K. Borowski also expresses his thanks to Bob Sauls (bob.sauls@xp4d.com) 

for his exceptional digital images depicting the vehicles and mission concepts discussed in this paper.

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2014-218105 1 

The Nuclear Thermal Propulsion Stage (NTPS): A Key Space Asset 

for Human Exploration and Commercial Missions to the Moon 

Stanley K. Borowski 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

David R. McCurdy 

Vantage Partners, LLC 

Brook Park, Ohio 44142 

Laura M. Burke 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The nuclear thermal rocket (NTR) has frequently been discussed as a key space asset that can bridge 

the gap between a sustained human presence on the Moon and the eventual human exploration of Mars. 

Recently, a human mission to a near Earth asteroid (NEA) has also been included as a “deep space 

precursor” to an orbital mission of Mars before a landing is attempted. In his “post-Apollo” Integrated 

Space Program Plan (1970 to 1990), Wernher von Braun, proposed a reusable Nuclear Thermal 

Propulsion Stage (NTPS) to deliver cargo and crew to the Moon to establish a lunar base initially before 

sending human missions to Mars. The NTR was selected because it was a proven technology capable of 

generating both high thrust and high specific impulse (Isp ~900 s)—twice that of today’s best chemical 

rockets. During the Rover and NERVA programs, 20 rocket reactors were designed, built and 

successfully ground tested. These tests demonstrated the (1) thrust levels; (2) high fuel temperatures; 

(3) sustained operation; (4) accumulated lifetime; and (5) restart capability needed for an affordable inspace transportation system. In NASA’s Mars Design Reference Architecture (DRA) 5.0 study, the 

“Copernicus” crewed NTR Mars transfer vehicle used three 25 klbf “Pewee” engines—the smallest and 

highest performing engine tested in the Rover program. Smaller lunar transfer vehicles—consisting of a 

NTPS with three ~16.7 klbf “SNRE-class” engines, an in-line propellant tank, plus the payload—can be 

delivered to LEO using a 70 t to LEO upgraded SLS, and can support reusable cargo delivery and crewed 

lunar landing missions. The NTPS can play an important role in returning humans to the Moon to stay by 

providing an affordable in-space transportation system that can allow initial lunar outposts to evolve into 

settlements capable of supporting commercial activities. Over the next decade collaborative efforts 

between NASA and private industry could open up new exploration and commercial opportunities for 

both organizations. With efficient NTP, commercial habitation and crew delivery systems, a “mobile 

cislunar research station” can transport crews to small NEAs delivered to the E-ML2 point. Also possible 

are week-long “lunar tourism” missions that can carry passengers into lunar orbit for sightseeing (and 

plenty of picture taking), then return them to Earth orbit where they would re-enter and land using a small 

reusable lifting body based on NASA’s HL-20 design. Mission descriptions, key vehicle features and 

operational characteristics are described and presented.

Nomenclature 

E-ML2 Earth-Moon L2 Lagrange point 

K temperature (degrees Kelvin) 

klbf thrust (1000’s of pounds force) 
NASA/TM—2014-218105 2 

LEO Low Earth Orbit (= 407 km circular) 

LOX/LH2 Liquid Oxygen/Liquid Hydrogen propellant 

NERVA Nuclear Engine for Rocket Vehicle Applications 

SLS/HLV Space Launch System/Heavy Lift Vehicle 

SNRE Small Nuclear Rocket Engine 

t metric ton (1 t = 1000 kg) 

V velocity change increment (km/s) 

1.0 Introduction and Background 

Less than a month after the successful landing of Apollo 11 on the Moon, Wernher von Braun, then 

director of the Marshall Space Flight Center (MSFC), outlined NASA’s Integrated Space Program Plan 

(1970 to 1990) to both the President’s Space Task Group (Ref. 1) and the Senate Committee on Aeronautics 

and Space Science (Ref. 2). The plan envisioned the development of a space shuttle and orbiting space 

station, along with a reusable NTR propulsion stage that would function as a “workhorse” space asset 

delivering cargo and crew to the Moon for construction of a lunar base, and then for a human landing on 

Mars in 1982. The utilization of NTP for lunar mission applications was evident in the fact that the 

operating characteristics and requirements for the NERVA flight engine were to be based on a nonoptimum, eight-burn, crewed mission to lunar polar orbit (LPO) and return (Ref. 3). Also considered was a 

four-burn, reusable cargo delivery mission as well. 

Despite the successes of the Rover/NERVA programs and the technological triumph of Apollo, the 

public’s interest in space waned and Apollo flights 18, 19, and 20 were cancelled, along with NASA’s postApollo plans for a lunar base and a human mission to Mars. After the final Apollo 17 mission to TaurusLittrow in December 1972 and cancellation of the Rover/NERVA programs in January 1973, short of flight 

demonstration, interest in human Moon/Mars missions and NTP development remained relatively dormant 

for more than a decade. 

On July 20, 1989, the 20th anniversary of Apollo 11, President Bush proposed a Space Exploration 

Initiative (SEI) for the United States, which called for a return to the Moon “to stay”, followed by a journey 

to Mars (Ref. 4). From 1989 to 1993, NASA conducted and funded both in-house and industry studies 

(Refs. 5 to 7) that outlined a campaign of human exploration that included the establishment of a 

transportation node in LEO, a permanent base on the Moon, then human missions to Mars. NASA’s 

baseline lunar transportation system (LTS) used LOX/LH2 engines for both the lunar transfer vehicle (LTV) 

and the single stage lunar descent/ascent vehicle (LDAV). The LTV core stage utilized two sets of 

propellant drop tanks—the first jettisoned after trans-lunar injection (TLI) and the second after lunar orbit 

insertion (LOI). The LDAV was expended in lunar orbit and the LTV utilized aeroassist for LEO capture 

(Ref. 5). 

During this same time period, Glenn Research Center (GRC) re-introduced NTP as a viable LTS option 

and quantified its benefits for a variety of lunar mission applications (Ref. 8). With its high thrust and high 

specific impulse capability, NTP enabled a fully reusable LTS with both the NTR-powered LTV and the 

LDAV returned to LEO for refueling, refurbishment and reuse (Fig. 1(a)). For a comparable initial mass in 

low Earth orbit (IMLEO), the fully reusable NTP system had a “return payload” mass fraction of ~23 

percent—twice that of the partially reusable aeroassisted chemical system. Another important consideration 

was the “g-loading” on the crew during Earth return. For the aeroassisted chemical LTS, the g-loading was 

~5 to 7 gE versus 0.5 to 0.7 gE for the NTP system at the beginning and end of the Earth orbit capture (EOC) 

burn—an order of magnitude reduction (Ref. 8). 

In NASA’s First Lunar Outpost (FLO) study conducted in 1992, an expendable NTP TLI stage powered 

by three 25 klbf “Pewee-class” engines (Fig. 1(b)) was compared against a chemical injection stage for 

sending a large (~96 t) integrated lander and ascent stage to the Moon (Ref. 9). The chemical injection stage 

had a total mass of ~155 t compared to ~105 t for the NTP system—a savings of ~50 t for this single TLI 
NASA/TM—2014-218105 3 

Expendable TLI Stage 

for “First Lunar Outpost” 

Mission uses 3 - 25 klbf

 

NTR Engines – Fast 

Track Study (1992) 

Reusable Lunar Transfer 

Vehicle uses Single 75 klbf

 

NTR Engine – SEI (1990 - 91) 

Reusable Lunar Commuter 

Shuttle uses 3 -15 klbf

 

LANTR Engines – (1997) 

Reusable Lunar Cargo Transports Using 

Clustered 15 - 25 klbf

 NTR Engines – (2012) 

Reusable Crewed Lunar Landing Mission Using 

Clustered 15 - 25 klbf

 NTR Engines – (2012) 

(a) (b) 

(c) 

(d) (e) 

Figure 1.—Sampling of crewed and cargo lunar transfer vehicles designed by GRC over the past two decades show 

a transition away from single large to multiple smaller engines. 

maneuver. The FLO study also marked the first use of smaller clustered NTR engines to help improve 

packaging and overall mission reliability. 

Over the next 5 years, with its industry partner Aerojet, GRC quantified the operational characteristics 

and performance potential of an enhanced NTR—known as the LOX-Augmented NTR (LANTR)—in an 

evolutionary LTS architecture (Ref. 10). The LANTR concept adds an oxygen (O2) “afterburner” nozzle and 

feed system to the basic NTR allowing O2 injection and supersonic combustion in the engine’s hot H2

exhaust downstream of the nozzle throat. By controlling the O2-to-H2 mixture ratio, the LANTR engine can 

operate over a wide range of thrust and Isp levels while its reactor core power level remains relatively 

constant. Also, by refueling in low lunar orbit (LLO) with LOX and LH2 produced from lunar regolith 

(ilmenite or iron-rich volcanic glass) or polar ice deposits, a reusable bipropellant LANTR LTS can deliver 

significant payload on each round trip and ultimately enable a rapid “commuter” shuttle (Fig. 1(c)) 

capable of 24 hr “one-way” flights to and from the Moon (Ref. 11). 

In 2004, NASA’s Constellation program began and by 2005 the Exploration Systems Architecture 

Study (ESAS) had outlined the basic building blocks needed to land a four-person crew on the Moon by 

December 2019 (Ref. 12). ESAS examined four different lunar mission options with various size cargo 

HLVs. Included were: (1) a 2-Launch lunar orbit rendezvous (LOR); (2) a 2-Launch Earth orbit 

rendezvous (EOR)-LOR; (3) a 2-Launch EOR-Direct Return approach similar to NASA’s earlier FLO 

study; and (4) a 1.5-Launch EOR-LOR option which NASA eventually selected for use in its 

Constellation program. The basic transportation elements are shown in Figure 2. A cargo HLV delivered 

a dual use upper stage, called the Earth departure stage (EDS), plus an expendable 2-stage LDAV called 
NASA/TM—2014-218105 4 

Earth Departure Stage with Altair LDAV and 

Orion Multipurpose Crew Vehicle (MPCV) 

Altair LDAV 

Figure 2.—Lunar Space Transportation System elements for the Constellation “1.5 launch” architecture. 

Altair to LEO. A much smaller launch vehicle then delivered the Orion capsule, service module and crew 

to LEO for rendezvous and docking with Altair. 

The EDS then performed the TLI burn to send the combined stack on its way to the Moon in much 

the same way as the Saturn V’s S-IVB upper stage did during Apollo. However, compared to the Apollo 

Command and Service Module, the Orion’s service module was less capable carrying only enough 

propellant to return itself and its crew to Earth. As a result the job of LOI for the combined Orion-Altair 

stack was assigned to Altair’s LOX/LH2 descent stage already tasked with landing on the lunar surface 

with a storable ascent stage and a crew of four—twice that carried by the Apollo lunar module (LM). As a 

result, the mass of Altair mushroomed to ~46 t—three times the mass of the LM. The design features of 

the Altair LDAV, expendable with a storable ascent stage, also seemed somewhat inconsistent with 

landing and locating an outpost near one of the Moon’s poles where deposits of lunar polar ice have been 

detected and could be processed to supply LOX/LH2 propellants for a reusable LDAV. Unfortunately, 

after a 5 year effort, the Constellation program was cancelled by the Obama administration in favor of 

pursuing technologies needed for a crewed mission to a near Earth asteroid (NEA) in the late 2020’s 

followed by an orbital mission to Mars by the mid-2030’s (Ref. 13). 

Despite this apparent setback to lunar exploration, interest in the Moon continues. In its Global 

Exploration Roadmap (GER) (Ref. 14), the International Space Exploration Coordination Group, with 

participants from NASA and 13 other space agencies, identified two possible pathways for future human 

missions after ISS utilization. These pathways have been designated the “Moon Next” and “Asteroid 

Next” scenarios. Both approaches utilize a stepwise development and demonstration of capabilities that 

are required for the eventual human exploration of Mars. 

The “Moon Next” pathway has a strong appeal to many who would like to see humans again walk on 

its surface and to whom the Apollo program has become a distant memory. Located just 3 days from 

Earth, the Moon is an entire world awaiting exploration, future settlement and potential commercialization. It is also an ideal location to test and demonstrate key technologies and systems (e.g., surface 

habitation, long-range pressurized rovers, surface power and resource extraction systems) that will allow 

people to explore, work, and live self-sufficiently on another planetary surface. Crewed NEA missions 

would follow and demonstrate the additional in-space capabilities needed to reach Mars such as advanced 

propulsion. Efficient propulsion and an affordable in-space transportation system with reuse capability 

will also be essential if initial lunar outposts are to evolve into eventual settlements capable of supporting 

commercial activities. 
NASA/TM—2014-218105 5 

In FY’12, GRC quantified the benefits of using NTP for human missions to two candidate 

NEAs—2000 SG344 and Apophis (Refs. 15 and 16) using scaled-down versions of the “Copernicus” 

crewed Mars transfer vehicle (MTV) design developed during NASA’s Mars DRA 5.0 study (Refs. 17 

and 18). These smaller Asteroid Survey Vehicles (ASVs) had a lower IMLEO than their chemical and 

SEP-chemical counterparts and were also reusable. The same two key components used in the Copernicus 

MTV and the smaller ASVs—the “core” nuclear thermal propulsion stage (NTPS) and the integrated 

“saddle truss” and LH2 propellant drop tank assembly connecting the NTPS to the payload element—

were also configured for reusable lunar cargo delivery (Fig. 1(d)), and crewed lunar landing missions 

(Fig. 1(e)) (Ref. 16) as envisioned by von Braun. 

Recently, industry has stepped up its efforts to develop and demonstrate technologies and systems 

that can support potential near term and future commercial space ventures. Examples include SpaceX, 

Sierra Nevada Corporation (SNC) and Boeing’s involvement in commercial cargo and crew delivery to 

the ISS, development of inflatable space habitats by Bigelow Aerospace, as well as, plans for propellant 

depots by United Launch Alliance (ULA), asteroid mining by Planetary Resources Inc. and Deep Space 

Industries, even commercial human flights to the surface of the Moon by the Golden Spike Company. In 

addition to its significant Commercial Crew and Cargo Program (Ref. 19), NASA’s Advanced 

Exploration Systems (AES) program will also test out inflatable space habitat technology aboard the ISS 

using the Bigelow Expandable Activity Module (BEAM) (Ref. 20) scheduled for launch in 2015. 

NASA and Bigelow Aerospace have also partnered, under a Space Act Agreement signed in March, 

to study potential public-private collaborations in space (Ref. 21). Following discussions with 20 

companies and international space agencies, Bigelow presented the first of two reports to NASA (Ref. 22) 

that focused on near-term opportunities in LEO, in lunar orbit and on the Moon, and indicated a 

significant private sector interest in lunar activities (Ref. 22). Even NASA’s plans for a crewed mission to 

a NEA by 2025 are being reexamined in favor of a less ambitious mission that would capture a small 

asteroid and return it to E-ML2 (Ref. 23). The SLS and Orion MPCV would then launch and transport 

two astronauts there to collect samples and study the asteroid up close. 

Over the past 4 years, NASA’s Human Architecture Team has been pursuing a strategy, referred to as 

a Capability Driven Framework (CDF), which assumes the development and utilization of evolving 

technologies and systems to perform a series of ever more challenging missions first to a NEA, then the 

Moon and finally Mars. With flat or reduced budgets for NASA expected in the future, a CDF may be 

short-sighted and jeopardize the Agency’s ability to orbit Mars by 2035 by diverting scarce resources 

away from proven technologies like NTP towards less capable systems that are large, operationally 

complex to use and unlikely to support fast transit missions to Mars. 

This paper proposes a multi-prong approach to human exploration that involves collaboration 

between NASA and private industry and can support either the “Moon Next” or “Asteroid Next” 

pathways outlined in the GER. It assumes a “Technology Driven Framework” with NASA and other 

space agencies working together to develop the high-leverage technologies and systems needed for Mars 

(like a heavy lift version of SLS, NTP and reverse turbo-Brayton refrigeration for zero-boiloff LH2 

storage) albeit with scaled-down systems to start with. Also, by exploiting the technology synergies that 

exist between NASA’s SLS (e.g., large aluminum/lithium (Al/Li) LH2 tanks) and existing flight-tested 

chemical rocket hardware (e.g., LH2 turbopumps, regenerative- and radiation-cooled nozzles and skirt 

extensions) substantial savings in the time and cost to develop a NTPS are to be expected. 

Industry would develop, mature then provide the commercial services—including crew delivery, 

space habitats, propellant depots and lunar landers—necessary to sustain an affordable in-space 

transportation system. Forty years after Apollo 17, as we look to the future, NASA, industry and the 

country must consider the following pressing question: Can we afford to continue human space 

exploration while operating systems in a throwaway mode? With a reusable “workhorse” NTPS, valuable 

in-space assets can be returned to Earth orbit for refurbishment and reuse, thereby reducing the cost of 

space travel not only for NASA but for future private sector endeavors as well. 
NASA/TM—2014-218105 6 

This paper examines the use of NTP for a variety of reusable lunar mission applications of potential 

interest to both NASA and industry. The missions include exploration lunar cargo delivery and crewed 

lunar landing missions to both equatorial and lunar polar orbits, a crewed survey mission to a small 

asteroid returned to E-ML2, and a week-long, round trip “lunar tourism” mission. Individual vehicle 

components are limited to 70 t and it is assumed that an “upgraded” SLS, with a capable upper stage, is 

available to deliver them to LEO. The paper covers the following topic areas. First, the operational 

principles and performance characteristics of the baseline ~16.7 klbf SNRE used in this analysis are 

presented followed by a brief discussion of NASA’s current NTP development activities funded under its 

AES program. Mission and transportation system ground rules and assumptions used in the analysis are 

then presented. The paper’s focus then turns to a discussion of the candidate lunar missions and includes a 

description of various mission scenarios, key vehicle features and engine operational requirements. The 

paper ends with a summary of our findings and some concluding remarks. 

2.0 NTR System Description, Performance Characteristics and 

Development Status 

The NTR uses a compact fission reactor core containing 93 percent “enriched” Uranium (U)-235 fuel 

to generate 100’s of megawatts of thermal power (MWt) required to heat the LH2 propellant to high 

exhaust temperatures for rocket thrust. In an “expander cycle” Rover/NERVA-type engine (Fig. 3), high 

pressure LH2 flowing from either a single or twin turbopump assembly (TPA) is split into two paths with 

the first cooling the engine’s nozzle, pressure vessel, neutron reflector, and control drums, and the second 

path cooling the engine’s tie-tube assemblies. The flows are then merged and the heated H2 gas is used to 

drive the turbine. The hydrogen turbine exhaust is then routed back into the reactor pressure vessel and 

through the internal radiation shield and core support structure before entering the coolant channels in the 

reactor core’s fuel elements. Here it absorbs energy produced from the fission of U-235 atoms, is 

superheated to high exhaust temperatures (Tex ~2550 to 2950 K depending on fuel type and uranium 

loading), then expanded out a high area ratio nozzle (~300:1) for thrust generation. 

Controlling the NTR during its various operational phases (startup, full thrust and shutdown) is 

accomplished by matching the TPA-supplied LH2 flow to the reactor power level. Multiple control drums, 

located in the reflector region surrounding the reactor core, regulate the neutron population and reactor 

power level over the NTR’s operational lifetime. The internal neutron and gamma radiation shield, 

located within the engine’s pressure vessel, contains its own interior coolant channels. It is placed 

between the reactor core and key engine components to prevent excessive radiation heating and material 

damage. 

Figure 3.—Schematic of “expander cycle” NTR engine with dual LH2 turbopumps. 
NASA/TM—2014-218105 7 

Figure 4.—Coated particle and composite SNRE fuel element and tie tube arrangement. 

The fuel elements (FE) tested in the Rover/NERVA program (Ref. 24) consisted of a “graphite 

matrix” material that contained the U-235 fuel in the form of either coated particles of uranium carbide 

(UC2) or as a dispersion of uranium and zirconium carbide (UC-ZrC) referred to as “composite” fuel. 

Each FE (see Fig. 4) had a hexagonal cross section (~0.75 in. across the flats) and 19 axial coolant 

channels that were coated with niobium carbide (NbC) initially, then with zirconium carbide (ZrC) using 

a chemical vapor deposition (CVD) process. This protective coating, applied to the exterior FE surfaces 

as well, helped reduce hydrogen erosion of the graphite. Individual elements were 52 in. in length and 

produced ~1 MWt. 

This basic FE shape was introduced in the KIWI-B4E reactor and became the standard used in the 

75 klbf Phoebus-1B, 250 klbf Phoebus-2A, 25 klbf Pewee and the 55 klbf NERVA NRX series of engines 

(Ref. 24). Also included in the engine’s reactor core were cooled coaxial tie tube (TT) elements that 

provided structural support for the FEs, as well as a source of energy for turbine drive power. The TTs 

also included a sleeve of zirconium hydride (ZrH) moderator material to help raise neutron reactivity 

(shown in Fig. 4). In the larger size engines tested in Rover/NERVA, a “sparse” FE—TT arrangement 

was used with each FE having two adjacent TTs and four adjacent FEs comprising its six surrounding 

elements (Ref. 25). In this sparse pattern, the FE to TT ratio is ~3 to 1. 

A Small Nuclear Rocket Engine (SNRE) design producing ~16.4 klbf of thrust was analyzed by 

Los Alamos National Laboratory near the end of the Rover/NERVA program (Ref. 26). The FE had the 

same hexagonal cross section and coolant channel number, but was 35 in. long, used composite fuel, and 

produced ~0.65 MWt. To help increase core reactivity, the “SNRE” FE—TT pattern increases the number 

of TTs with each FE having three adjacent TTs and three adjacent FEs that surround it (shown in Fig. 4). 

With the SNRE pattern, the FE to TT ratio is ~2 to 1. An important feature common to both the sparse 

and SNRE FE—TT patterns is that each tie tube is surrounded by and provides mechanical support for six 

fuel elements. Also, the same hexagonal cross-section for the FE and TT elements can be maintained 

across a range of different thrust levels although shorter length elements are likely for lower thrust 

engines. Additional analysis by Schnitzler (Refs. 25, 27, and 28) has shown that the SNRE design can be 

scaled down to even smaller thrust levels (~7.5 klbf) or up to the 25 klbf Pewee-class engine used in Mars 

DRA 5.0. 
NASA/TM—2014-218105 8 

The rationale for considering the NTR for lunar missions is simple—it is a proven, high thrust 

(10’s of klbf) propulsion technology with a specific impulse that is 100 percent higher than today’s best 

chemical rockets. During the Rover/NERVA programs(1955 to 1972), a technology readiness level 

(TRL ~5 to 6) was achieved (Ref. 24). Twenty rocket reactors were designed, built and ground tested in 

integrated reactor/engine tests that demonstrated: (1) a wide range of thrust levels (~25, 50, 75, and 

250 klbf); (2) high temperature carbide-based nuclear fuels that provided hydrogen exhaust temperatures 

up to 2550 K (achieved in the Pewee engine); (3) sustained engine operation (over 62 min for a single 

burn on the NRX-A6); as well as; (4) accumulated lifetime at full-power; and (5) restart capability (>2 hr 

with 28 startup and shutdown cycles on the NRX-XE experimental engine)—all the requirements needed 

for a viable lunar space transportation system, also for human NEA and Mars exploration missions. 

Lastly, it is important to note that NTP requires no large scale-ups in size or performance that are needed 

with other non-chemical propulsion options. In fact, the smallest and highest performance engine tested 

during the Rover/NERVA programs—the 25 klbf “Pewee” engine (Ref. 24) is sufficient for a human 

mission to Mars when used in a clustered engine arrangement. Even smaller, lower thrust engines are 

sufficient for lunar mission applications. 

The NTR engine baselined in this paper is an ~16.675 klbf “SNRE-class” expander cycle engine with 

the following performance parameters: Tex ~2726 K, chamber pressure ~450 psia,  ~300:1, and Isp~900 s. 

The LH2 flow rate is ~8.40 kg/s and the engine thrust-to-weight ratio is ~3.06. The overall engine length 

is ~6.1 m including the radiation-cooled nozzle skirt extension, and the nozzle exit diameter is ~2.31 m. 

The engine’s reactor core contains 564 FEs and 241 TTs each ~0.89 m (35 in.) long. The core power level 

and fuel matrix power density are ~367 and ~3.44 MWt/liter, respectively. The U-235 fuel loading is 

~0.60 g/cm3 and the inventory of 93 percent enriched U-235 in the core is just under 60 kg. 

High temperature ZrC-coated UC-ZrC in graphite “composite” fuel was selected as the primary fuel 

form in this analysis. Composite FEs were first tested in the “Nuclear Furnace” element test reactor

(Ref. 24) and withstood peak power densities of ~4500 to 5000 MWt/m3. They also demonstrated better 

corrosion resistance than the standard coated particle graphite matrix fuel element used in the previous 

Rover/NERVA reactor tests. Composite fuel’s improved corrosion resistance is attributed to its higher 

coefficient of thermal expansion that more closely matches that of the protective ZrC coating, thereby 

helping to reduce coating cracking. 

Electrical-heated composite fuel elements were also tested by Westinghouse in hot hydrogen at 

2700 K for ~600 min—equivalent to ten 1-hr cycles. At the end of Rover/NERVA, composite fuel 

performance projections(Ref. 29) were estimated at ~2 to 6 hr at full power for hydrogen exhaust 

temperatures of ~2500 to 2800 K and fuel loadings in the range of ~0.60 to 0.45 g/cm3

. A ceramic-metal 

or “cermet” fuel consisting of uranium dioxide (UO2) in a tungsten (W) metal matrix was developed in 

the General Electric (GE)-710 and Argonne National Laboratory (ANL) nuclear rocket programs 

(Refs. 30 and 31) as a backup to the Rover/NERVA fuel. While no integrated reactor/engine tests were 

conducted, a large number of fuel specimens were produced and exposed to non-nuclear hot H2 and 

irradiation testing with promising results. Both fuel options are currently under development. 

A NTP technology development effort was started by NASA in FY’11 under the Advanced In-Space 

Propulsion (AISP) component of its Exploration Technology Development and Demonstration (ETDD) 

program. The effort included two key tracks: “Foundational Technology Development” followed by 

“Technology Demonstration” projects. The Foundational Technology Development component continues 

today under NASA’s AES program and the Nuclear Cryogenic Propulsion Stage (NCPS) project 

(Ref. 32) which began in FY’12. This 3-year project is a collaboration between NASA and DOE and 

includes five key tasks and objectives: 

Task 1. Mission Analysis, Engine/Stage System Characterization and Requirements Definition is 

aimed at establishing performance goals for fuel development, testing and concept designs for both small, 

scalable demonstration engines and higher thrust-class engines needed for future human Moon, NEA and 

Mars missions; 
NASA/TM—2014-218105 9 

Task 2. NTP Fuel Development and Assessment is focused on recapturing fabrication techniques, 

testing and maturing fuel, coating and cladding materials, then selecting between the two primary fuel 

forms previously identified by the DOE and NASA—NERVA composite and UO2 in tungsten cermet fuel 

(Ref. 33). Partial, then full-length fuel elements, based on “heritage” designs, will be fabricated and tested 

in the NTR Element Environmental Simulator (NTREES)(Ref. 34) at the MSFC using up to ~1.2 MW of 

RF heating to simulate the NTP thermal environment including exposure to hot H2. Fuel element 

performance at required temperature, duration, and with cycling to mimic engine restart, will be evaluated 

before final fuel selections are made and irradiation testing begins; 

Task 3. Engine Design, Analysis and Modeling is aimed at developing conceptual designs for small 

(~7.5 klbf) demonstration engines and higher thrust-class (~25 klbf) engines utilizing the candidate fuels 

discussed above. State-of-the-art numerical models are being used to determine reactor core criticality, 

detailed energy deposition and control rod worth within the reactor subsystem(Refs. 25, 27, and 35), 

provide thermal, fluid and stress analysis of reactor fuel elements and core components (Refs. 36 and 37), 

and predict engine operating characteristics and overall mass(Refs. 38 and 39); 

Task 4. Demonstration of Affordable Ground Testing is focused on maturing equipment requirements 

and cost estimates for a “proof-of-concept” validation test of the Subsurface Active Filtration of Exhaust 

(SAFE) (Refs. 40 and 41) concept (also known as the “bore-hole” option) at the Nevada Test Site (NTS). 

Non-nuclear, subscale hot gas injection tests, some with a radioactive tracer gas, would be conducted in a 

vertical bore-hole to obtain test data on the effectiveness of the porous rock (alluvium) to capture, holdup 

and filter the engine exhaust. The data would also help calibrate design codes needed by DOE to design 

the SAFE test facility and support infrastructure needed for the small engine ground technology 

demonstration tests and larger thrust class engine tests to follow; and 

Task 5. Formulation of an Affordable and Sustainable NTP Development Strategy is aimed at 

outlining the content of an affordable development plan that utilizes separate effects tests (e.g., NTREES 

and irradiation tests), existing assets and innovative SAFE testing at the NTS, and small scalable engines 

for ground and flight technology demonstrations. The above tasks, successfully carried out by the end of 

FY’14, would provide the basis for a NCPS Phase II effort that could culminate with ground technology 

demonstration tests at the NTS in the early 2020’s, followed by a flight technology demonstration mission 

in 2025 (Ref. 42). 

3.0 Mission, Payload and Transportation System Ground Rules 

and Assumptions 

Specific mission and payload ground rules and assumptions used in this paper are summarized in 

Table 1. It provides information about the different lunar mission scenarios, and the assumed parking 

orbits at Earth, the Moon and at the Earth-Moon L2 Lagrange point. Specific trajectory details and V 

budgets are provided within the appropriate sections of the paper. In addition to the large V 

requirements for the primary propulsion maneuvers, like trans-lunar injection (TLI), smaller V 

maneuvers are needed for rendezvous and docking (R&D) of vehicle components during the LEO 

assembly phase, vehicle mid-course correction (MCC) maneuvers, and for spacecraft attitude control 

during in-space transit. 

Cargo delivery and crewed lunar landing missions are considered first in this paper. On cargo flights, 

an integrated habitat lander with surface mobility (shown in Fig. 5(a)) is delivered to lunar orbit by a 

reusable NTR transport in a manner reminiscent of von Braun’s reusable lunar NTR shuttle. On the 

crewed landing missions, a forward mounted saddle truss connects the payload elements to the transfer 

vehicle’s in-line tank. The truss is open on its underside and its forward adaptor ring provides a docking 

interface between the Orion MPCV and a single stage LOX/LH2 LDAV (shown in Fig. 5(b)). The LDAV 

is a “heritage” design (Ref. 7) analyzed in considerable detail during SEI. It carries a crew of four plus 5 t 

of surface payload stored in two “swing-down” containers mounted on each side of the crew cab. The 

LDAV mass breakdown with propellant loadings for a range of landed payload is shown in Table 1. 
NASA/TM—2014-218105 10 

TABLE 1.—MISSION AND PAYLOAD GROUND RULES AND ASSUMPTIONS 

Pre-deployed Wheeled

 Lunar Habitat Modules 

Orion MPCV docking to 

Lunar Descent Ascent Vehicle

 via connecting Saddle Truss 

Inflatable Habitat Module with attached

 Commercial Crew HL-20 Lifting Body

 and MMSEV 

MMSEV 

(a) 

LDAV 

MPCV (b) 

(c) 

Figure 5.—Payload elements carried by NTP cargo and crewed lunar transfer vehicles. 
NASA/TM—2014-218105 11 

For crewed science missions to small NEA(s) delivered to E-ML2, or commercial orbital tourist 

flights around the Moon, an inflatable habitation module, like TransHab or Bigelow Aerospace’s BA-330 

module, can be carried to accommodate larger crew/passenger sizes along with the extra life support, 

consumables and accommodations needed to sustain them. A seven-passenger “mini-Shuttle” based on 

NASA’s HL-20 lifting body concept is another option for commercial crew delivery that is being 

developed by Sierra Nevada Corporation (Ref. 43). A small auxiliary multi-mission space excursion 

vehicle (MMSEV) can also be carried on NEA science missions. The MMSEV provides a livable volume 

for a crew of two for up to two weeks (Ref. 44) and would be attached to the back end of the TransHab 

module. The above elements are shown in Figure 5(c). The MMSEV provides ~200 to 

300 m/s of translational V, suitports for EVA sorties, and remote manipulation capability for sample 

collection. For the lunar landing and small NEA science missions analyzed here, it is assumed that the 

crews collect and return ~100 kg of samples. 

Table 2 lists the key ground rules and assumptions used in the NTP lunar transportation system 

consisting of the core NTPS plus an in-line LH2 propellant tank (shown in Fig. 6). The NTPS uses a 

three-engine cluster of ~16.7 klbf SNRE-class engines discussed previously in Section 2.0. The engines 

use composite fuel with a U-235 fuel loading of 0.6 g/cm3

. With a hydrogen exhaust temperature (Tex) of 

~2726 K, chamber pressure of ~450 psia, and a nozzle area expansion ratio of ~300:1, the Isp is ~900 s. 

The total mission LH2 propellant loading consists of the usable propellant plus performance reserve, 

post-burn engine cooldown, and tank-trapped residuals. For the smaller auxiliary maneuvers, a storable 

bipropellant RCS system is used. All transfer vehicle configurations utilize a split RCS with 

approximately half the AMBR thrusters and bipropellant mass located on the rear NTPS and the other 

half located at the front end of the in-line tank or saddle truss adaptor section just behind the missionspecific payload. 

TABLE 2.—NTR TRANSPORTATION SYSTEM GROUND RULES AND ASSUMPTIONS 
NASA/TM—2014-218105 12 

Figure 6.—Key NTP lunar transfer system elements—the NTPS and in-line LH2 tank. 

The LH2 propellant carried in the NTPS and its in-line tank is stored in the same “state-of-the-art” 

Al/Li LH2 propellant tank being developed for the SLS/HLV that will support future human exploration 

missions. For this analysis, tank sizing assumes a 30 psi ullage pressure, 5 gE axial/2.5 gE lateral launch 

loads, and a safety factor of 1.5. A 3 percent ullage factor is also assumed. All tanks use a combination 

foam/multilayer insulation (MLI) system for passive thermal protection. A zero boil-off (ZBO) “reverse 

turbo-Brayton” cryocooler system is used on the NTPS to eliminate boil-off during LEO assembly and 

over the course of the mission. It is not used on the in-line LH2 tank since it is drained during the TLI 

maneuver. The propellant tank heat load is largest in LEO and sizes the ZBO cryocooler system. Solar 

photovoltaic arrays are baselined for supplying the primary electrical power needed for all key transfer 

vehicle subsystems. 

Table 2 also provides the assumed “dry weight contingency” (DWC) factors, along with the 

requirements for delivered mass to LEO and the shroud cylindrical payload envelope for the upgraded 

SLS/HLV. A 30 percent DWC is used on the NTR system and advanced composite structures (e.g., stage 

adaptors, trusses) and 15 percent on heritage systems (e.g., Al/Li tanks, RCS, etc.). The NTPS determines 

the upgraded SLS lift capability (~70 t) and the usable PL volume within the shroud (~7.6 m OD and 

~26.8 m in length). The combined saddle truss (~13.7 m) and LDAV (~9.6 m) with its attached cargo 

containers used on the crewed landing mission (Fig. 5(b)) has this same approximate length. For crewed 

NEA science missions to E-ML2, the PL includes a “packaged” TransHab module with PVA power 

system, a short saddle truss and a MMSEV (Fig. 5(c)). The overall length of this element is ~22.8 m. 

Options for crew delivery include NASA’s Orion MPCV, SpaceX’s Dragon and Boeing’s CST-100

capsules, and Sierra Nevada Corporation’s “Dream Chaser” lifting body. The commercial crew delivery 

systems would utilize man-rated versions of ULA’s Atlas V or SpaceX’s Falcon 9 launch vehicles. 

4.0 NTP Space Transportation System for Lunar Cargo and Crewed 

Landing Missions 

As discussed in the Introduction, NTP can play an important role in returning humans to the Moon by 

providing an affordable in-space lunar transportation system with reuse capability that could allow initial 

lunar outposts to evolve into eventual settlements capable of supporting commercial activities. Utilization 

of efficient NTP for lunar cargo delivery and crewed lunar landing missions is also consistent with the 

“Asteroid Next” pathway that includes human missions to the Moon to test out key surface systems 

(e.g., habitats, power systems, and long-range pressurized rovers) needed for an eventual human landing 

on Mars. 
NASA/TM—2014-218105 13 

The NTPS is the “workhorse” element of the LTS. It uses a three-engine cluster of 16.675 klbf

composite fuel SNRE-class engines and also carries external radiation shield mass for crew protection. 

The NTPS uses an Al/Li LH2 tank which has a diameter (D) and length (L) of 7.6 m D15.7 m L. The 

tank’s LH2 propellant capacity is ~39.7 t. The NTPS also carries avionics, RCS, auxiliary battery and 

PVA power, docking and a reverse turbo-Brayton ZBO refrigeration systems located in the forward 

cylindrical adaptor section. To remove ~42 W of heat penetrating the 60 layer MLI system in LEO 

(where the highest tank heat flux occurs), the 2-stage cryocooler system requires ~5.3 kWe for operation. 

Twin circular PVAs provide the electrical power needs for the NTPS. 

The second major element is an “in-line” Al/Li propellant tank that connects the NTPS to the forward 

PL element. It has the same 7.6 m D and supplies the LH2 propellant needed for the “2-perigee burn” TLI 

maneuver. Depending on the particular mission and the PL carried, the tank length can vary from ~15.7 m 

(same length as in the NTPS) to ~18.7 m and capable of holding ~49.0 t LH2 propellant. The in-line tank 

element also includes forward and aft cylindrical adaptor sections that house quick connect/disconnect 

propellant feed lines, electrical connections, a RCS along with docking and payload adaptors. The total 

length of the in-line element varies from ~20.7 to 23.7 m. 

For the reusable cargo delivery mission, three upgraded SLS launches are used to deliver the vehicle 

and payload elements. Here they are assembled via autonomous rendezvous and docking (AR&D). The 

NTP cargo transport then departs from LEO (C3 ~ –1.678 km2

/s2, VTLI ~3.214 km/s including a g-loss of 

~117 m/s) and captures into a 300 km circular LLO (arrival Vinf ~1.151 km2

/s2 and VLOC ~906 m/s 

including g-loss) approximately 72 hr later. Key phases of the cargo delivery mission are illustrated in 

Figure 7. Once in orbit, the habitat lander separates from the lunar NTR (LNTR) transport and descends 

to the surface, landing autonomously at a predetermined location on the Moon. It is assumed that the 

habitat landers use LOX/LH2 chemical engines and are also equipped with either deployable wheels 

(shown in Fig. 5(a)) or articulated landing gear allowing movement in both the vertical and horizontal 

directions so that the landers can either “drive or walk” short distances from the landing site. Connecting 

several “functionally different” lander modules together (for habitation, science, equipment servicing) 

would form a large contiguous pressurized volume for the crew and also provide a “building block” 

approach to establishing an initial lunar base. 

Docked Habitat 

Landers 

Habitat Lander Descent 

to the Lunar Surface 

Delivery of Habitat 

Lander to LLO (300 km) 

NTP Lunar Cargo Transports 

Departing from LEO (407 km) 

Figure 7.—Reusable NTP lunar cargo delivery mission phases. 
NASA/TM—2014-218105 14 

After payload separation and a day in LLO, the LNTR cargo transport performs a trans-Earth 

injection (TEI) burn (C3 ~ 0.945 km2

/s2, VTEI ~857 m/s including g-loss) and returns to Earth 72 hr later. 

On final approach, it performs a braking burn (arrival Vinf ~ –1.755 km2

/s2, and VEOC ~366 m/s) and 

captures into a 24-hr EEO. Post burn engine cool-down thrust is then used to assist in orbit lowering. An 

auxiliary tanker vehicle, operating from a LEO servicing node/propellant depot, supplies the additional 

LH2 needed by the cargo transport for final orbit lowering and rendezvous with the LEO transportation 

node where it is refurbished and resupplied before its next mission. 

The key phases of the crewed NTR landing mission are illustrated in Figure 8. Again, as with the 

cargo mission, three upgraded SLS launches are used to deliver the two NTR vehicle elements and the 

payload element—consisting of an integrating saddle truss assembly (STA) and a LDAV with its surface 

cargo containers—to a 407 km orbital altitude where they are assembled via AR&D. In addition to a front 

and rear docking capability, the STA’s forward adaptor ring also carries twin PVAs and a RCS. Once 

assembled, the Orion MPCV and crew are launched and rendezvous with the LNTR vehicle positioning 

itself inside the STA and docking with the LDAV using a docking port and transfer tunnel mounted to the 

STA’s forward adaptor ring (shown in Figs. 5(b) and (8)). 

NTP Transfer Vehicle 

Insertion into 300 km LLO 

Crewed Lunar Transfer Vehicle 

Departure from 407 km LEO 

Orion MPCV R&D 

with LDAV 

Figure 8.—Reusable NTP crewed lunar landing mission—outbound mission leg. 
NASA/TM—2014-218105 15 

After the “2-perigee burn” TLI maneuver (C3 ~ –1.516 km2/s2, VTLI ~3.214 km/s including a g-loss 

of ~110 m/s), the crew begins its 3-day coast to the Moon. Because the crewed LNTR transport carries a 

significant amount of payload mass (the STA, MPCV, and LDAV) back from the Moon, it uses a longer 

(~18.7 m) in-line tank to supply the required amount of LH2 propellant needed for this reusable mission. 

After its 72-hr transit, the LNTR vehicle performs the LOC burn (arrival Vinf ~1.217 km2

/s2 and VLOC

~913 m/s including g-loss) inserting itself and its payload into LLO. 

Once in LLO, the crew then enters the LDAV, separates from the LNTR transport and prepares to 

land (shown in Fig. 9). The LDAV has a “wet” mass of ~35.3 t (Table 1) that includes the crew cab 

(~2.5 t), the descent/ascent stage (~6.1 t) and its LOX/LH2 propellant (~20.9 t), surface payload (~5 t), 

plus the four crew and their suits (~0.8 t). After separating from the LNTR, the LDAV’s two payload 

containers are rotated 180 and lowered into their landing position in preparation for descent to the lunar 

surface. The V budget for the LDAV includes the following (Ref. 7) Vdes ~2.115 km/s and Vasc

~1.985 km/s. The LDAV uses five RL 10A-4 engines that operate with a Isp ~450 s consistent with the 

Martin Marietta design (Ref. 7). It expends ~13.4 t of LOX/LH2 propellant during the descent to the 

surface (shown in Fig. 9). After lunar touchdown, the crew can operate out of the LDAV for ~3 to 

14 days using its surface landed payload or longer (~180 days) using the pre-deployed habitat landers. 

As the “exploration and surface systems testing” phase of the mission nears its completion, the crew 

prepares the LDAV for departure. At liftoff, the LDAV mass is ~15.1 and ~5.5 t of propellant is used 

during the ascent to LLO. The LDAV, with 100 kg of lunar samples, then rendezvous with the LNTR 

vehicle and preparations for the TEI maneuver begin. After completing the departure burn 

(C3 ~ 0.949 km2

/s2, VTEI ~856 m/s with g-loss), the crew spends the next 3 days in transit readying the 

LNTR for the final phase of the mission—capture into a 24-hr EEO with the MPCV and LDAV payload 

(depicted in Fig. 9) followed by MPCV separation and capsule re-entry of the crew. 

Trans - Earth Injection 

& Coast Back to Earth 

LDAV with Payload in 

Pre-descent Position 

Crewed LDAV Landing 

Near Habitat Landers 

Propulsive Capture 

into a 24-hr EEO 

Figure 9.—Reusable NTP crewed lunar landing mission—landing and return mission leg. 
NASA/TM—2014-218105 16 

4.1 Features and Characteristics of LNTR Cargo Delivery and Crewed Landing 

Vehicles 

The key features, component lengths, and launch mass requirements for the lunar cargo and crewed 

landing missions are shown in Figure 10. Also shown is a reusable Asteroid Survey Vehicle (ASV) for a 

crewed mission to NEA 2000 SG344 in 2028 (Ref. 16) for comparison. The ASV uses the same NTPS 

used for the various lunar transfer vehicles discussed in this paper. It has an IMLEO of ~179.6 t that 

includes the NTPS (~69.5 t), a saddle truss and drop tank assembly (~54.8 t) and the crew PL section 

(~55.3 t). The overall vehicle length is ~78.3 m including the 8.9 m long Orion MPCV. The LH2 tank 

lengths for the NTPS and the drop tank are identical at ~15.7 m with each tank carrying ~38.9 t of LH2

propellant (~98 percent of tank’s maximum capacity of ~39.7 t). Using the saddle truss/drop tank 

assembly has advantages for higher V NEA and Mars missions because it allows the drained LH2 drop 

tank to be jettisoned after the Earth departure burn thereby reducing propellant consumption on 

subsequent mission maneuvers. Though made of composite material, the saddle truss still has 

considerable mass that must be injected from LEO along with the drop tank. For the fully reusable lunar 

missions considered here, use of an in-line tank is the preferred approach allowing a higher delivered 

payload for a lower total launch mass. 

The LNTR cargo transport shown in Figures 7 and 10 has an IMLEO of ~186.7 t consisting of the 

NTPS (~70 t), the in-line tank element (~52.6 t), and the habitat lander (~61.1 t) with its connecting 

structure (~3.0 t). The overall vehicle length is ~60.4 m. Like the ASV, the cargo transport also uses 

~15.7 m long tanks in the NTPS and in-line element with each tank carrying ~39.7 t of LH2 propellant. 

By focusing on smaller size engines and maximizing the use of common hardware elements (e.g., same 

size NTPS, propellant tanks) for a variety of mission applications it should be possible to reduce vehicle 

development and recurring costs while also improving overall affordability. 

6.1 m 

26.8 m 23.1 m 21.2 m 

12.9 m 

4 crew 

TransHab 

Primary

 PVAs (4) 

Small PVA 

 (1 of 2) 

In-Line LH2 Tank 

7.6 m 

Common NTPS 

Communications 

Antenna (1 of 2) 

3 – 16.675 klbf 

 SNREs 

26.8 m 

Lunar Habitat 

Lander 

Orion

 MPCV 

Short Saddle Truss, 

Transfer Tunnel & MMSEV 

Orion MPCV with LDAV 

26.8 m 

26.8 m 

23.7 m 

20.7 m 

ASV: 2000 SG344 (2028)

(LEO – NEA – 6-hr EEO)

• 4 crew 

• PL + MPCV ~55.3 t 

• IMLEO ~179.6 t 

• Max Lift ~69.5 t (NTPS) 

• Total Mission Burn 

 Time: 54.5 min 

Lunar Cargo Delivery:

(LEO – LLO – 24-hr EEO)

• Habitat Lander ~61.1 t 

• IMLEO ~186.7 t 

• Max Lift ~70 t (NTPS) 

• Total Mission Burn 

 Time: 49.2 min 

Crewed Lunar Landing:

(LEO – LLO – 24-hr EEO)

• 4 crew 

• LDAV + MPCV ~48.9 t 

• IMLEO ~188.6 t 

• Max Lift ~70 t (NTPS) 

• Total Mission Burn 

 Time: 55 min 

Long Saddle Truss 

and LH2 Drop Tank 

Figure 10.—Reusable NTP vehicles for NEA, lunar cargo and crewed landing missions. 
NASA/TM—2014-218105 17 

For the reusable cargo delivery mission, the five primary engine burns use ~74.5 t of LH2 propellant. 

With 50 klbf of total thrust and Isp ~900 s, the total engine burn time is ~49.2 min. The first of the two 

TLI perigee burns is the longest at ~21.4 min. The duration of the remaining burns are as follows: second 

perigee burn (~15.5 min), LOC (~8.0 min), TEI (~3.1 min) and EEO capture (~1.2 min). These 

performance requirements are well below those demonstrated in the NERVA program that included a 

62 min maximum single burn demonstrated by the NRX-A6, and ~2 hr of accumulated burn time with 

27 restarts demonstrated by the NRX-XE (Ref. 24). Furthermore, given projected full power operational 

lifetimes of ~6 to 10 hr for NERVA-derived engines using composite fuel (Ref. 29), cargo transport 

vehicles with multi-mission capability appear viable. 

The cargo transport’s delivery capability to LPO has also been analyzed for a “8-burn” Earth-to-LPO 

round trip mission reminiscent of that considered previously in establishing the operational characteristics 

and requirements for the NERVA flight engine (Ref. 3). The cargo vehicle again uses a 

2-perigee burn departure from LEO and after a 3-day transit captures initially into an elliptical lunar orbit 

with a perilune of ~100 km and apolune of ~27,700 km. Two subsequent burns provide the plane change 

and circularization maneuvers that put the cargo transport into a 100 km circular LPO. Following a 

similar departure sequence for TEI and a 3-day return to Earth, the cargo vehicle again captures into a 

24-hr EEO. Despite the higher total V requirements for LOC and TEI, the cargo transport still delivers 

~51.2 t to LPO. The IMLEO is ~177.0 t and the total engine burn time is ~48.9 min. 

The crewed lunar landing mission has an IMLEO of ~188.6 t that includes the NTPS (~70.0 t), the 

in-line tank assembly (~63.3 t), the STA (~6.4 t), the wet LDAV (~29.5 t) with its surface payload (~5 t), 

the Orion MPCV (~13.5 t), consumables (~0.1 t), and four crewmembers (~0.8 t includes lunar EVA 

suits). The overall vehicle length is ~77.3 m. For the crewed mission, the LH2 propellant loads in the 

NTPS and in-line tank are at their maximum capacity of ~39.7 and ~49.0 t for their specified tank lengths 

of ~15.7 and ~18.7 m, respectively. 

The crewed landing mission also requires five primary burns. With 50 klbf of total thrust, Isp ~900 s, 

and ~83.2 t of LH2 propellant used during the mission, the total engine burn time is ~55 min, again well 

under the capabilities demonstrated in the NERVA program. The first perigee burn is the longest at ~20.9 

min and the remaining maneuvers having the following burn durations: second perigee burn (~16.2 min), 

LOC (~8.2 min), TEI (~6.9 min), and EOC (~2.8 min). 

5.0 Crewed NTP Science Mission to Small Asteroid Returned to E-ML2 

Lagrange Point 

As mentioned in the Introduction, NASA’s plans for a crewed mission to a NEA in the 2025 to 2030 

timeframe are being reexamined in favor of a less ambitious robotic mission that would use a xenonfueled, 40 kWe solar electric propulsion (SEP) system to capture a small (~7 m diameter and ~500 t) NEA 

and return it to the E-ML2 point (Ref. 23). The SLS and Orion MPCV would then launch and transport 

the crew there to collect samples and study the asteroid up close. A year earlier, in 2012, NASA 

considered proposing the establishment of a modest, crew-tended outpost at E-ML2, referred to as the 

“Waypoint Station”. Orion crews would visit the outpost periodically conducting telerobotic science on 

the lunar far side, as well as radio astronomy and scientific observations of the Sun and Earth in the quiet 

zone behind the Moon. Returning a small asteroid to the E-ML2 point would provide an additional target 

of opportunity for future scientific and commercial activities in translunar space. 

A reusable ASV using efficient NTP could function as a versatile “mobile research station” that can 

transport crew and support scientific activities at multiple locations including E-M Lagrange points, 

equatorial and lunar polar orbits, and select NEAs like 2000 SG344. The crewed “Prospector” ASV 

(shown below in Fig. 11) can support a month-long E-ML2 small asteroid science mission as well as a 

reusable 327-day round trip mission to 2000 SG344 in 2028 carrying a crew of four. Operated in an 

expendable mode, the round trip time can be reduced to 178 days. 
NASA/TM—2014-218105 18 

Use of the Prospector ASV for a small asteroid science mission at E-ML2 is depicted in Figure 11. 

The ASV carries a crew of four and the mission duration is ~33 days which includes an ~9.8 day 

outbound transit, an ~5.8 day stay at E-ML2, and an ~17.4 day inbound transit. Shorter round trip times of 

~2 to 3 weeks are also possible. The ASV departs from LEO and captures into a 4-hr EEO at the end of 

the mission. The total mission V is ~5.15 km/s and includes gravity losses plus lunar flyby impulsive 

burns performed during the outbound and return mission legs. 

Rather than having tethered astronauts operating out of the close confines of the Orion capsule, the 

Prospector ASV carries an inflatable habitation module to accommodate larger crew sizes (possibly with 

combined NASA and industry personnel), along with the extra life support, consumables and scientific 

equipment needed to analyze retrieved samples on site. A 2-person MMSEV, attached to the rear of the 

habitation module (shown in Fig. 11), is also part of the payload allowing close-up inspection and sample 

gathering from the asteroid during the mission. 

Like the cargo transport, Prospector (shown in Fig. 12) uses common ~15.7 m long tanks in both the 

NTPS and in-line element with each tank carrying ~39.7 t of LH2 propellant. It has an IMLEO of ~170.8 t 

consisting of the NTPS (~68.3 t), the in-line tank element (~52.3 t), and the crewed payload (~50.2 t). The 

PL element includes the habitation module (~22.7 t), a short saddle truss with RCS and PVAs (~8.03 t), 

the MMSEV (~6.7 t), the Far Side Voyager lifting body (~11.65 t), the crew and their suits (~0.8 t) plus 

consumables (~0.32 t). The overall vehicle length is ~79.3 m which includes the ~9 m long Far Side 

Voyager. 

Prospector ASV prior 

to Earth Departure 

Prospector ASV departs LEO 

for E-ML2 Lagrange point 

MMSEV Close-up Inspection 

of Small Asteroid at E-ML2

Figure 11.—Reusable NTP crewed science mission to small NEA at E-ML2 Lagrange point. 
NASA/TM—2014-218105 19 

26.8 m 20.7 m 22.8 m 

4 crew 

TransHab 

 Primary 

 PVAs (4) 

7.6 m 

NTPS 

3 – 16.675 klbf

 SNREs Far Side Voyager

Lifting Body Short Saddle Truss, 

Transfer Tunnel 

and MMSEV 

In-Line LH2 Tank 

Communication 

Dishes (2) 

Circular PVAs

 (1 of 2) 

Cryocooler 

and RCS 

Forward

RCS 

Figure 12.—Key ASV features and dimensions for crewed science mission to small NEA at E-ML2. 

The small asteroid science mission requires six primary burns and uses ~73 t of LH2 propellant. With 

50 klbf of total thrust and Isp ~900 s, the total engine burn time is ~48.3 min. The first perigee burn is the 

longest at ~19.8 min. The remaining maneuvers having the following burn durations: second perigee burn 

(~14.4 min), outbound lunar flyby burn (~1.4 min), L2 insertion/departure burns (~0.4 min), inbound 

lunar flyby burn (~1.6 min), and EOC (~10.7 min). 

6.0 Commercial Orbital “Tourism” Missions to the Moon Using NTP 

The lunar missions discussed above are expected to be of interest primarily to NASA although there 

could be participation with industry in cargo delivery and possible small asteroid science missions to 

E-ML2 in the future. Within a decade, it is conceivable that an orbital “lunar tourism” industry could 

develop providing a commercial opportunity for an industry consortium to operate. Interest in private 

spaceflight does exist, albeit only for the wealthy at present. Virgin Galactic (Ref. 45) has reportedly 

booked over 625 paying passengers (at $200,000 per person including a $20,000 deposit) for suborbital 

flights on its new SpaceShip 2 spacecraft. Each flight is expected to carry six passengers and two pilots, 

last ~2.5 hr and provide ~6 min of weightlessness. Bigelow Aerospace (Refs. 46 and 47) is planning on 

launching and operating its “Alpha Station” in LEO for a wide range of commercial activities and with 

stay times lasting anywhere from 10 to 60 days. Single, per seat rates will be either $26.25 or 

$36.75 million depending on the transportation provider selected by the client. With these activities 

progressing forward, how long will it be before tourists in LEO point to the Moon and say, “I'd like to go 

there, can you take me?” 

Many people are familiar with some of the prominent surface features and places on the Moon’s Near 

Side, like the great rayed-craters, Tycho and Copernicus, and the Apollo 11 landing site in the “Sea of 

Tranquility”, but only 27 humans on the 9 Apollo lunar flights have seen some of the Moon’s Far Side 

features (shown in Fig. 13) up close and personal. With an efficient, reusable LNTR transportation 

system, using proven NTR technology, week-long tourism missions to the Moon can carry passengers 

into lunar orbit for close-up examination (and plenty of picture taking) then return them to Earth orbit 

where they would re-enter and land using a reusable mini-Shuttle like SNC’s Dream Chaser or the Far 

Side Voyager (FSV) depicted in this paper. 

Key phases of the outbound portion of the orbital lunar tourism mission are illustrated in Figure 14. 

The mission begins with the launch of seven passengers and crew in the FSV aboard a human-rated 

version of a commercial launch vehicle like ULA’s Atlas V. Once in LEO, the FSV rendezvous with 

Lunar Vistas Unlimited’s (LVU’s) commercial passenger transport (CPT) after which the CPT initiates a 

“2 perigee burn” departure maneuver (C3 ~ –1.678 km2

/s2, VTLI ~3.206 km/s including a g-loss of 

~109 m/s). The CPT carries with it a spacious habitation and observation module plus the FSV to the 

Moon. During the 3-day transit to the Moon, the passengers aboard the CPT will be able to experience the 
Figure 13.—Sampling of noted surface features on the far side of the Moon.

Figure 14.—Commercial orbital mission to the Moon—launch and Earth departure phases.

NASA/TM—2014-218105 20
NASA/TM—2014-218105 21 

0-gE environment of space as they view a shrinking Earth and expanding Moon from the four large 

viewing ports on the habitation/observation module as the CPT gradually rotates about its longitudinal 

axis oriented perpendicular to its flight path. 

On the final lunar approach, the passengers return to their seats in the FSV and the CPT performs the 

orbit insertion burn (arrival Vinf ~1.151 km2

/s2 and VLOC ~911 m/s including a 1 percent g-loss) and 

captures into a 300 km circular LLO (shown in Fig. 15). Over the next 24 hr, the passengers on the CPT 

will orbit the Moon ~10.5 times experiencing both the Moon’s Sunlit Far Side and Earthlit darkened Near 

Side, and viewing surface features few people on Earth will ever see. After a memorable day of sight 

seeing in lunar orbit, the CPT performs the TEI burn (C3 ~ 0.945 km2

/s2, VTEI ~865 m/s including a 

1 percent g-loss) and begins the 3-day voyage back to Earth. On final approach, the CPT performs a 

braking burn (arrival Vinf ~ –1.755 km2

/s2, VEOC ~370 m/s including a 1 percent g-loss) and captures into 

a 24-hr EEO. On either the first or second perigee pass, the passengers and crew return to the FSV, detach 

from the CPT and perform a small de-orbit burn, landing back at the Kennedy Space Center (KSC) or 

other commercial landing site (shown in Fig. 15). The CPT’s post burn engine cool-down thrust is used to 

lower its orbit until it can rendezvous with the tanker vehicle that will supply it with the LH2 propellant 

needed for final orbit lowering and return to the LEO transportation node. Here it will be refurbished and 

resupplied in preparation for its next lunar mission. 

Like the cargo transport and Prospector ASV, the CPT (shown in Fig. 16) uses the same common 

~15.7 m long tank in both the NTPS and in-line element. It has an IMLEO of ~169.5 t consisting of the 

NTPS (~70 t), the in-line element (~53.4 t), and the crewed payload element (~46.1 t). This last element 

includes the habitation/observation module with its PVAs (~32.9 t), the seven passengers and crew with 

their suits (~1.4 t), consumables (~0.15 t), plus the FSV mini-Shuttle (~11.65 t). The overall vehicle 

length is ~71.3 m including the FSV. 

“Experience Lunar Orbit Insertion..... 

....and Return to KSC all in 1-Week’s Time”

...and Earthlit Lunar Night Side.... 

....the Moon’s Sunlit Far Side... 

Sea of Moscow 

Figure 15.—Commercial orbital mission to the Moon—insertion, Lunar observation and Earth return. 
NASA/TM—2014-218105 22 

26.8 m 20.7 m 14.8 m 

Passenger Habitation 

/ Observation Module 

Primary 

 PVAs (4) 

7.6 m 

NTPS 

3 – 16.675 klbf

 SNREs 

Far Side Voyager

Lifting Body In-Line LH2 Tank 

Communication 

Dishes (2) 

Circular PVAs

 (1 of 2) 

Cryocooler 

and RCS 

Forward

 RCS 

Figure 16.—Key features and dimensions for the commercial lunar passenger transport. 

The CPT mission requires five primary burns. With 50 klbf of total thrust, Isp ~900 s, and ~74.7 t of 

LH2 propellant used during the round trip mission, the total engine burn time is ~49.4 min. The first of the 

two TLI perigee burns is again the longest at ~18.8 min. The duration of the remaining burns are as 

follows: second perigee burn (~14.5 min), LOC (~7.3 min), TEI (~6.3 min), and 24-hr EEO capture 

(~2.5 min). 

With a fixed NTPS and in-line tank mass, and total propellant loading, it is also possible to return the 

CPT deeper into Earth’s gravity well (e.g., into a 6-hr rather than 24-hr EEO) at the end of the mission in 

order to reduce the amount of propellant that the tanker needs to deliver to the CPT for final orbit 

lowering. Doing so increases the capture V requirements at Earth by ~300 percent (arrival Vinf ~ 

–1.755 km2

/s2, VEOC ~1097 m/s including a 1 percent g-loss) and the propellant needed for capture from 

~3.8 to ~9.8 t, thereby decreasing the mass of the “outfitted” habitat/observation module that can be 

transported to the Moon and back. 

For this particular case, the CPT’s IMLEO decreases from ~169.5 to ~156.0 t with the following mass 

breakdown: NTPS (~69.9 t), the in-line tank (~53.3 t), the habitation/observation module with its PVAs 

(~19.6 t), the seven passengers and crew with suits (~1.4 t), consumables (~0.15 t), plus the FSV

(~11.65 t). The total engine operating time for the mission remains roughly the same at ~49.7 min with 

the individual burn durations as follows: first perigee burn (~17.3 min), second perigee burn (~13.4 min), 

LOC (~6.7 min), TEI (~5.8 min), and 6-hr EEO capture (~6.5 min). For the round trip mission, the burnup of U-235 fuel in each engine is ~15.3 g (assuming ~1.2 g consumed per megawatt-day of operation) or 

~0.025 percent of the total amount of U-235 contained in each SNRE. It is therefore quite apparent that 

fuel burn-up in the engines of the CPT is not an issue and that significant reuse capability exists with 

these systems. 

With its high performance NTPS, the CPT can also fly an 8-burn LEO-LPO-24-hr EEO round trip 

trajectory carrying an ~24 t habitat/observation module with its seven passenger and crew and the FSV. 

The CPT uses a 2-perigee burn departure from LEO and after a 3-day transit captures into an elliptical 

lunar orbit with a perilune of ~100 km and apolune of ~27,700 km. Two subsequent burns provide the 

plane change and circularization maneuvers that place the CPT into a 100 km circular LPO (depicted in 

Fig. 17). Following a similar departure sequence for TEI and a 3-day return to Earth, the CPT again 

captures into a 24-hr EEO. The IMLEO is ~162.7 t, the total engine burn time is ~49.1 min, and the 

U-235 fuel burn-up is ~15 g (~0.025 percent of the total amount of U-235 contained in each engine). 
NASA/TM—2014-218105 23 

Figure 17.—Commercial passenger transport in polar orbit over the Moon’s southern hemisphere. 

7.0 Summary and Conclusions 

The NTR has frequently been discussed as a key space asset that can bridge the gap between a 

sustained human presence on the Moon and the eventual human exploration of Mars. Wernher von Braun 

himself envisioned a reusable “workhorse” NTPS delivering cargo and crew to the Moon for construction 

of a lunar base, then being outfitted with additional propellant tanks for human missions to Mars. The 

NERVA program also utilized a non-optimum 8-burn crewed mission to lunar polar orbit and back as its 

DRM for determining the operating characteristics and requirements for the NERVA flight engine. 

The superiority of NTP over conventional chemical propulsion was well documented in analyses 

conducted during NASA’s SEI. Its use enabled a fully reusable LTS architecture that included the ability 

to also return the LDAV to LEO for refurbishment, resupply and reuse. The addition of an oxygen 

afterburner nozzle to the NTR also offers the potential for bipropellant operation and refueling in LPO 

with LOX and LH2 propellants produced from lunar polar ice in the future. Yet despite NTP’s 

performance advantages and growth potential, it is often overlooked as a viable propulsion option by 

NASA mission planers. Their frequent response to the question, “Why don’t we consider NTP for the 

Moon?” is that it’s overkill or simply not needed since we did the Apollo program with chemical 

propulsion. It is true that chemical propulsion got us to the Moon but each mission started with an ~3000 t 

Saturn V on the launch pad at KSC and ended with the return of a charred capsule, the crew and ~63.5 kg 

of lunar samples (average of six landing missions). If humans are to return to the Moon and establish a 

permanent presence there, affordability, maintainability and reusability will be essential. Throwing away 

$100’s of millions of dollars of expensive hardware during each mission may not be a sustainable or 

acceptable mode of operation in the future. 

Despite cancellation of the Constellation program, interest in lunar exploration continues. The Global 

Exploration Roadmap continues to show the “Moon Next” pathway as one of its two possible options for 

future human exploration with the second option being the “Asteroid Next” path. Private industry (e.g., 

SpaceX, SNC and Bigelow Aerospace) has also stepped up its pursuit of commercial space business and 

the private sector has indicated a significant interest in lunar activities as evidenced by Golden Spike’s 

proposed plans to offer commercial human landing missions on the Moon by the early 2020’s. 

With its high thrust and specific impulse (100 percent higher than today’s best chemical rockets), 

proven NTP technology can provide the affordable “access through space” needed to support future 

exploration and commercial development activities on the Moon. It is also the only advanced propulsion 

system successfully ground tested over the full range of performance parameters required for tomorrow’s 
NASA/TM—2014-218105 24 

human missions to the Moon, NEAs, and Mars. For the lunar missions discussed in this paper, a threeengine cluster of SNREs are used on the NTPS. This size engine is smaller than the smallest and highest 

performing engine tested in the Rover program—the 25 klbf-class Pewee engine. Through the NCPS 

project, NASA and DOE are currently engaged in a number of key NTP task activities that include stateof-the-art engine modeling, mission design and requirements definition, fuel element fabrication and nonnuclear testing, and analysis of affordable options for nuclear ground testing and engine development. If 

successfully completed by the end of FY’14, a NCPS-Phase II effort would begin that could culminate in 

ground testing a small NTR engine at the NTS in the early 2020’s, followed by a flight demonstration 

mission several years later. 

Four different lunar mission applications—cargo delivery, crewed landing, small asteroid science at 

E-ML2, and orbital tourism—are examined using a reusable NTP transportation system consisting of two 

key elements. Each element is limited to 70 t in LEO. The NTPS uses three composite fuel SNRE-class 

engines that operate at 900 s specific impulse and produce ~50 klbf of total thrust. In one week’s time, 

using less than 75 t of LH2 propellant, the NTP cargo transport can deliver ~61 t of payload to LLO then 

return to a 24-hr EEO for tanker servicing and subsequent return to LEO for resupply before the next 

mission. In this same time period, the cargo transport can also deliver ~51.2 t of payload to LPO if so 

desired. 

For the crewed landing mission, the LNTR transfer vehicle uses the same NTPS, carries out ~53.6 t 

of payload to LLO, then returns to the same 24-hr EEO with ~28.7 t of payload mass including collected 

samples and the spent LDAV. It requires a slightly longer in-line tank to provide the extra propellant 

needed for the mission. The total engine burn times for the cargo and crewed landing missions are 

~50 and 55 min, respectively and both missions require five engine restarts, performance requirements 

well below those demonstrated in the NERVA program. The U-235 burn-up for the cargo and crewed 

landing missions is also quite small—no more 0.028 percent of the total U-235 inventory in each SNRE. 

The NTP crewed science mission to a small asteroid returned to E-ML2 using SEP is an interesting 

combination of technology utilization, system demonstration and mission destinations that has synergy 

with both the “Asteroid Next” and “Moon Next” pathways being considered in the GER. Capturing and 

returning a small asteroid to E-ML2 helps to validate the technology for a 40 kWe SEP system and brings 

together in translunar space two targets of opportunity that are of scientific and commercial interest to 

both NASA and private industry. It also provides a precursor demonstration mission for a crewed NTP 

spacecraft, Prospector, that can function as a mobile Waypoint Station initially, before being used for 

actual NEA missions lasting anywhere from ~6 months to a year in duration. The round trip time for the 

small asteroid science mission at E-ML2 can range from ~2 weeks to a month and can utilize a 

combination of NASA and commercial space assets. Like the cargo transport vehicle, Prospector also 

uses the same length LH2 tank in both the NTPS and in-line tank element in order to maximize 

component commonality and minimize recurring costs. 

A commercial passenger transport for orbital tourism is perhaps the most interesting of the NTP lunar 

mission applications. It allows 7-day round trip missions including 3-day transits to and from the Moon 

with a day in lunar orbit for some “out of this world” sightseeing. With the available propellant in the 

NTPS and in-line tank element, the CPT can trade habitation/observation module mass against different 

elliptical capture orbits back at Earth. The shorter the orbital period, the lower the mass transported out to 

the Moon and back. With a 6-hr EEO and IMLEO of ~156 t, the CPT’s NTPS and in-line elements can be 

delivered to LEO on two upgraded SLS launches while the habitat/observation module with its PVAs at 

~20 t can potentially use a smaller commercial launch vehicle. With its high performance NTPS, the CPT 

can also fly an 8-burn LEO-LPO-24-hr EEO round trip trajectory carrying an ~24 t habitat/observation 

module, seven passengers and crew plus the FSV mini-Shuttle. The CPT can also carry larger modules 

and more passengers on each round trip if the FSV is left at the LEO transportation node and used only for 

passenger delivery and pickup at the end of the mission. 

NTP is frequently identified as a propulsion option to be used primarily for Mars but it can also play 

an important role in future human exploration and commercial missions to the Moon by providing 

affordable transportation through cislunar and translunar space. With a sustained and credible funding 
NASA/TM—2014-218105 25 

profile, NASA, DOE and industry could have an operational NTR engine and stage available within a 

decade. After that, how long will it be before the travel section of your Sunday newspaper features an 

advertisement for week-long orbital missions to the Moon with the caption, “Come Fly with Us...Come 

Spend a Day in Lunar orbit”? 

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Expandable Module on Space Station,” January 16, 2013. 
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The Nuclear Thermal Propulsion Stage (NTPS): A Key Space Asset for Human

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The nuclear thermal rocket (NTR) has frequently been discussed as a key space asset that can bridge the gap between a sustained human presence on the Moon and the

eventual human exploration of Mars. Recently, a human mission to a near Earth asteroid (NEA) has also been included as a “deep space precursor” to an orbital mission of

Mars before a landing is attempted. In his “post-Apollo” Integrated Space Program Plan (1970 to 1990), Wernher von Braun, proposed a reusable Nuclear Thermal Propulsion

Stage (NTPS) to deliver cargo and crew to the Moon to establish a lunar base initially before sending human missions to Mars. The NTR was selected because it was a proven

technology capable of generating both high thrust and high specific impulse (Isp ~900 s)-twice that of today’s best chemical rockets. During the Rover and NERVA programs,

20 rocket reactors were designed, built and successfully ground tested. These tests demonstrated the (1) thrust levels; (2) high fuel temperatures; (3) sustained operation; (4)

accumulated lifetime; and (5) restart capability needed for an affordable in-space transportation system. In NASA’s Mars Design Reference Architecture (DRA) 5.0 study, the

“Copernicus” crewed NTR Mars transfer vehicle used three 25 klbf

 “Pewee” engines-the smallest and highest performing engine tested in the Rover program. Smaller lunar

transfer vehicles-consisting of a NTPS with three ~16.7 klbf

 “SNRE-class” engines, an in-line propellant tank, plus the payload-can be delivered to LEO using a 70 t to LEO

upgraded SLS, and can support reusable cargo delivery and crewed lunar landing missions. The NTPS can play an important role in returning humans to the Moon to stay by

providing an affordable in-space transportation system that can allow initial lunar outposts to evolve into settlements capable of supporting commercial activities. Over the next

decade collaborative efforts between NASA and private industry could open up new exploration and commercial opportunities for both organizations. With efficient NTP,

commercial habitation and crew delivery systems, a “mobile cislunar research station” can transport crews to small NEAs delivered to the E-ML2 point. Also possible are

week-long “lunar tourism” missions that can carry passengers into lunar orbit for sightseeing (and plenty of picture taking), then return them to Earth orbit where they would

re-enter and land using a small reusable lifting body based on NASA’s HL-20 design. Mission descriptions, key vehicle features and operational characteristics are described

and presented.

15. SUBJECT TERMS

Nuclear thermal rocket (NTR); Spacecraft; Lunar transportation system; Lunar tourism; Human exploration

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU

18. NUMBER

OF

PAGES

34

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) a. REPORT

U

b. ABSTRACT

U

c. THIS PAGE

U 19b. TELEPHONE NUMBER (include area code)

443-757-5802

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18