2

Mars Transportation Assessment Study

(MTAS)

National Aeronautics and Space Administration 
3

ST U D Y SP O N S O R E D B Y

Tom Cremins

Associate Administrator for Strategic Engagement and Assessments

NASA Headquarters

M TA S T E A M

Dr. Ave K. Kludze 

Study Lead, NASA Headquarters

Mr. Leonard A. Dudzinski 

Deputy Study Lead, NASA Headquarters 

Mr. Richard Davis

Study Executive, NASA Headquarters

Name Affiliation 

MTAS Team 

Bob Collom HQ

Ell Bogat HQ

John Dankanich MSFC

Andy Presby GRC

Stephen Edwards MSFC

Reggie Alexander MSFC

John Scott JSC

Daniel (Dan) Dorney MSFC

Lee Mason GRC

Michael Houts MSFC

Harold Gerrish MSFC

Patrick Chai LARC

June Zakrajsek GRC

George Schmidt GRC

Brandon Klefman GRC

Michael Meyer GRC

James Cannon MSFC

Michelle Rucker JSC

Ronald Litchford HQ

David Jacobson GRC

Bret Drake JSC

Raymond Merrill LARC

Marc Gibson GRC

Thomas Brown MSFC

Chris Iannello KSC

Melissa Mcguire GRC

Kurt Hack GRC

Laura Burke GRC

Caroline Holmes GRC

Christopher Moore HQ

Jeffrey Hojnicki GRC

Guy Kemmerly LARC

Jason Hartwig GRC

Scott Martinelli HQ

Donald Pettit JSC

Adalberto Sierra KSC

Tony Beaver MSFC

Kevin Dickens GRC

Laura Hardin MSFC

Brian Beaver KSC

Alicia Dwyer Cianciolo LARC

Steve McCarty GRC

David A. Smith GRC

Brent Faller GRC

Christine L. Schmid GRC

James E. Fittje GRC

John Z. Gyekenyesi GRC

Anthony J. Colozza GRC

Christopher Heldman GRC

Lucia Tian GRC

Thomas Packard GRC

Thomas Parkey GRC

Natalie Weckesser GRC

W. Peter Simon GRC

Onoufrios 

Theofylaktos 

GRC

Elizabeth Turnbull GRC

Michelle Munk LARC

Craig Douglas HQ

Paul Schmitz GRC

John Yim GRC

Max Chaiken GRC

Caroline Austin GRC

Michael Martini GRC

Nicholas Lantz GRC

Bushara Dosa GRC

Pete Peterson GRC

Louis Pinero GRC

Kevin Higdon MSFC

Mitchell Rodriguez MSFC

Steven (Steve) Oleson GRC

Denton Gibson KSC

Kurt Polzin MSFC

Steve Bauer LARC

Laura Ulrich KSC

William C. Hill MSFC

DOE Team 

Sam Bays INL

Florent Heidet ANL

Adam Nelson ANL

Lou Qualls ORNL

Jorge Navarro ORNL

Mike Smith ORNL

Brian Ade ORNL

Dong Choe INL

Patrick McClure LANL

Dennis Tucker INL

Lise Charllot INL

Spencer Snow INL

Adrian Wagner INL

Douglas Burns INL

Robert O’Brien INL

Milos Atz ANL

Alisha Kasam-Griffith ANL

Paul Talbot INL
NASA 4

Acknowledgements

The MTAS Team would like to thank Walt Engelund (Deputy Associate Administrator/STMD), Jim Reuter 

(Associate Administrator/STMD), and Tom Cremins (Associate Administrator for Space Security Interests) for 

their steadfast support of this effort and guidance of the study through the Federated Board.

The MTAS Team also very much appreciates the members of the MTAS Red Team: Jeffrey George, John (Jay) 

Falker, Jeffrey Sheehy, John Connolly, Kurt Hack, Stanley Borowski, Michael Lapointe, and Hussein Khalid, for 

providing an independent internal technical review through each phase of the study period. This team has 

extensive experience in past nuclear studies and their reviews and inputs were invaluable.

An additional thank you to Brian Wessel and Tom Mcmurry at the Office of General Council who helped 

coordinate a major part of the approval process and were instrumental in getting this report released.

We are also very grateful to the External Review team: Roger Myers, Dasari V. Rao and Bobby Braun, 

for providing an independent external review of this document. Their very constructive comments and 

suggestions helped communicate the study findings more impactfully. We also thank Nikolai Joseph for 

coordinating the External Review; and Ellen Gersten, Deputy Associate Administrator for Office of Technology, 

Strategy, and Policy (OTSP) for reviewing this document and offering guidance.

A great deal of gratitude goes to Dr. Bhavya Lal, the Associate Administrator for Office of Technology, Policy 

and Strategy, (OTPS) who facilitated the External Review and personally provided comments and suggestions 

that enhanced this document.

Finally, the MTAS Team would like to thank Ms. Pam Melroy, Deputy NASA Administrator, for initiating the 

External Review, and for her support to make this an impactful study that advances the national knowledge on 

the NTP and NEP propulsion systems and vehicles.
NASA 5

Foreword

The Mars Transportation Assessment Study (MTAS) was initiated in 2019 as an agency-level, independent 

assessment of Nuclear Thermal and Nuclear Electric transportation technologies required to put humans on 

Mars and return them safely to Earth. The study’s objectives were to: 1) Establish the feasibility of transportation 

systems for a ~2-year human Mars mission in the 2030s with acceptable risk and minimal development, and 2) 

Provide the agency with necessary knowledge and insights to inform future architectures, concept of operations 

(ConOps), campaigns, and development of the required capabilities. MTAS provides a detailed technical 

analysis of these two propulsion technologies using a common set of ground rules that were developed based 

upon the anticipated human Mars mission architecture at that time.

This analysis is intended to be illustrative and will contribute to the broader discussion on how to send humans 

to Mars. Thus, NASA is releasing these study results to inform the ongoing assessment of Mars transportation 

propulsion options in the evolving Mars architecture development. Specifically, MTAS directly supports the 

Transportation and Habitation Goal identified in NASA’s Moon to Mars Objectives1 and is relevant to the following 

objectives:

• TH-5: Develop a transportation system that crew can routinely operate from the Earth-Moon vicinity to Mars orbit 

and the Martian surface.

• TH-6: Develop a transportation system that can deliver large surface elements from Earth to the Martian surface.

• TH-12: Develop systems capable of returning large cargo mass from the Martian surface to the Earth, including 

the capabilities necessary to meet scientific sample return objectives.

As NASA’s ground rules for human Mars mission transportation continue to evolve with the mission architecture,

this study’s findings add valuable depth to the body of knowledge necessary for the transportation propulsion 

trades. The MTAS team performed analyses of each propulsion option without constraints on the concept of 

operations, beyond the common ground rules of trip time, crew operations, payloads, launch vehicles, 

rendezvous orbits, and infrastructure. This allowed each propulsion technology to take an optimal approach for 

the defined mission with technology choices made to balance required performance with development risk. As 

a result, the mission performance comparisons can be taken as relevant to any Mars transportation mission 

where each technology is similarly constrained. Furthermore, the technology comparisons and findings are 

independent of mission architecture, and valuable to the trade space evaluation. NASA encourages further 

discussions within the external technical community on the potentials of these two nuclear propulsion systems.

MTAS was supported by NASA’s experts in propulsion, power, mission design, trajectory analysis, cost analysis, 

systems engineering, and programmatic development. The team also relied on the Department of Energy’s 

nuclear engineering expertise.

NASA Headquarters

Washington, D.C.

1 https://www.nasa.gov/sites/default/files/atoms/files/m2m-objectives-exec-summary.pdf
NASA 6

Table of Contents

Table of Contents.............................................................................................................................6

1.0 Summary................................................................................................................................................8

2.0 Introduction.........................................................................................................................................11

3.0 Ground Rules and Assumptions........................................................................................................16

4.0 Mars Transportation Assessment......................................................................................................19

4.1 Nuclear Thermal Propulsion Reference Case ..........................................................................19

4.1.1 NTP Technology...................................................................................................................23

4.1.2 NTP Campaign......................................................................................................................25

4.2 Nuclear Electric Propulsion Reference Case ...........................................................................28

4.2.1 NEP Technology...................................................................................................................33

4.2.1.1 Chemical Propulsion Technology.........................................................................35

4.2.2 NEP Campaign......................................................................................................................35

4.3 Comparison between Propulsion Alternatives.........................................................................38

4.3.1 Performance ..........................................................................................................................38

4.3.2 Development.........................................................................................................................39

4.3.2.1 Technology-NTP...................................................................................................39

4.3.2.2 Technology-NEP: .................................................................................................40

4.3.2.3 Reactor – NTP: .....................................................................................................42

4.3.2.4 Reactor-NEP:........................................................................................................44

4.3.2.5 Propellant-NTP:....................................................................................................44

4.3.2.6 Propellant-NEP:....................................................................................................45

4.3.2.7 Testing and Development Facilities-NTP:............................................................46

4.3.2.8 Testing and Development Facilities-NEP.............................................................46

4.3.3 Operational Distinctions .......................................................................................................47

4.3.3.1 Operational Distinctions: Thrust...........................................................................47

4.3.3.2 Operational Distinctions: Extensibility.................................................................48

4.3.3.3 Operational Distinctions: Reusability ...................................................................49

4.3.3.4 Operational Distinctions: Robustness...................................................................50

4.3.4 Campaign ..............................................................................................................................52

4.3.5 Cost .......................................................................................................................................55

4.3.6 Other Considerations ............................................................................................................56

5.0 Conclusions..........................................................................................................................................57
EXECUTIVE S U M M A RY

S T U D Y G O A L

The MTAS Team was tasked by NASA leadership to establish the feasibility of transportation system options for a ~2 

year opposition class human Mars mission in the 2030s to reduce overall risks to astronauts. The initial focus of the 

team was on nuclear systems given the known challenges of chemical systems for opposition class trajectories. The 

team was to assess acceptable risk and minimal development for a viable mission, and provide the Agency with 

guidance for potential architectures, concept of operations (ConOps), campaign, and development of the required 

capabilities. Further assessment of the non-nuclear alternatives is recommended as follow-on work to ensure 

completeness of results; but is not currently anticipated to change the overall results. 

KE Y T A K E AWAY S

( C O R R E S P O N D I N G S E C T I O N S I N P A R E N T H E S E S )

1. Both Nuclear Thermal Propulsion (NTP) and Nuclear Electric Propulsion with Chemical Combustion 

Propulsion (NEP/CCP) system concepts have the potential to fulfill the mission requirements for fast 

Mars transits in the 2030s, but with significantly different challenges (4.3) 

2. A balanced portfolio of both NTP and NEP investment in technology development is required to retire 

risk and provide confidence for a down select decision (5.0) 

a. The current states of knowledge and technology are insufficient to commit to one without high risk 

(4.1.1, 4.2.1) 

3. If the Mars Transportation Assessment Study (MTAS) technology projections are validated through 

technology development, NEP augmented with chemical propulsion would be the more 

advantageous option for human Mars transportation and has stronger synergistic benefits (5.0) 

a. One-third of the launches, with overall lower mission mass required vs NTP (4.3.4) 

b. Significantly lower development and overall/subsequent use cost vs NTP (4.3.5) 

c. Mission robustness offered by the agility of having both high and low thrust systems (4.3.3.4) 

d. Option to use the more affordable CCP stage alone for predeployed cargo deliveries (4.2) 

e. Greater performance potential by reducing system specific mass & increasing Isp (4.3.2.2) 

f. Mitigates variance in mission performance across the Earth-Mars synodic cycle (4.3.3.4) 

g. Development of NEP has greater benefits to the range of NASA’s exploration applications

1. Fission power development for surface applications is synergistic with NEP (4.3.3.2) 

2. Small NEP enables high ∆V robotic exploration precursor and follow-on applications (4.3.3.2) 

h. NEP is an overall lower risk technology development with stronger synergy to terrestrial nuclear 

1. The use of Chemical Propulsion with NEP reduces the technology required (4.2.1) 

2. NEP can be tested in modular subsystems, with simulated interfaces, reducing facility 

requirements (4.2.1) 

4. A DOD-NASA partnership is possible for either NTP or NEP. DOD/DARPA is interested in NTP due 

to its high thrust acceleration and rapid response capability. A concept for ~10 kWe NEP based on 

the Kilopower reactor is also under consideration for development by the U.S. Space Force. (4.3.3.2) 

5. The use of High-Enriched Uranium (HEU) leverages past nuclear fuel and reactor development for 

both NTP and NEP and results in the simplest, highest performance systems (4.3.2.3, 4.3.2.4) 

a. The use of High-Assay, Low-Enriched Uranium (HALEU) introduces significant system 

complexity and development risk.

6. For the faster human missions that nuclear propulsion enables, committed and significant investment 

in nuclear propulsion technology must begin now, for Mars missions in the 2030s. (4.1.1, 4.2.1) 

NASA 7
NASA 8

1.0 S U M M A R Y 

NASA is studying a strategic objective to put humans on Mars and return them safely to Earth by the mid to late 

2030s. To investigate the feasibility of these objectives, NASA chartered the Mars Transportation Assessment 

Study (MTAS) to establish a human transportation solution based on technologies that would be achievable in 

the time frame of interest. Due to the current understanding of mission and crew risks associated with long 

duration interplanetary missions, the study was challenged with a two-year roundtrip, and 30-sol (Mars day) Mars 

surface stay as ground rules. Furthermore, a minimally viable overall architecture was given as a constraint to 

limit required developments and enhance programmatic affordability and achievability. Subsequently, MTAS was 

extended to refine what was initially established as a “proof of concept” design to address identified challenges, 

and balance overall risk between mission operations and technology development. This is the documentation of

the findings of Phase 1 of this study, including two reference design points for two propulsion system alternatives 

that can achieve the Agency objectives.

The MTAS team conducted analyses to determine the feasibility of performing a two-year roundtrip human 

mission to Mars, and the cargo missions to support it. A two-year round trip Mars mission requires a higher 

energy “opposition-class” trajectory, with an associated short stay time at Mars. By comparison, these 

trajectories require more propulsive energy than “conjunction class” trajectories, which require approximately a 

3-year roundtrip with longer stay times in the Martian system. Conjunction class trajectories take advantage of 

more favorable planetary alignments to minimize energy requirements. Chemical, solar-electric, nuclearthermal, nuclear-electric, and hybrid combinations of these propulsion systems were initially considered for the 

assessment. The study team was tasked to evaluate technical development requirements to determine if the 

required technologies could be ready for a 2035 mission. Since the developed vehicle design would be expected 

to be useful for additional Mars-mission opportunities, requirements were assessed against very difficult 

planetary alignments in 2039 and 2042 to establish bounding performance requirements. Any vehicle capable 

of such a mission would be too large to be launched on a single rocket so the study team assumed elements 

would be assembled in space across multiple launches. Launch vehicle options considered include the Space 

Launch System (SLS), Heavy Commercial Launch Vehicles (H-CLVs), and Super Heavy Commercial Launch 

Vehicles (SH-CLVs). Because the Lunar Artemis Program will also need some of the limited SLS vehicles, 

constraints were placed on how many could be used for the Mars mission. However, no constraints were placed 

on the availability of the H-CLVs or SH-CLVs.

To support Agency strategic planning, and due to limited resources, the MTAS focused on the two most 

promising propulsion technologies: Nuclear Thermal Propulsion (NTP), and Nuclear Electric Propulsion (NEP).

Because of propellant efficiency, as measured in specific impulse (Isp), each of these propulsion technologies 

has significant advantages over Chemical Combustion Propulsion (CCP or “Chemical Propulsion”) and could 

potentially enable more robust and faster transits to Mars. The study innovated a hybrid NEP/CCP approach, 

similar to the hybrid SEP/CCP approach identified as a non-nuclear alternative for Mars conjunction class 

missions in the 1990s, whereby a balance of high thrust chemical propulsion and low thrust electric propulsion 

can perform fast missions with less propellant using less advanced technology. Additional studies are

recommended to objectively assess both SEP/CCP and all-Chemical Propulsion against the MTAS NTP and 

NEP/CCP reference cases to provide the justification for the yet-to-be-selected Mars transportation propulsion 

option. Table 5.0-1 contains a compilation of recommended follow-on work. 

MTAS Phase 1 concluded that both NTP and NEP/CCP propulsion system concepts have the potential to fulfill 

the mission requirements for fast Mars transits in the late 2030s, including enabling two-year roundtrip mission

durations. But these options use different approaches with unique challenges. By technology comparison, NTP 

has technology challenges related to the high-temperature reactor operation and long duration liquid hydrogen 

storage and is more costly to develop and implement (according to the MTAS cost assessment). NEP requires 

a complex integration of five major subsystems: reactor, power conversion, heat rejection, power management 
NASA 9

& distribution (PMAD), and EP thrusters. The NEP development benefits from previous efforts and synergies 

with current NASA electric propulsion (both space- and aircraft-based) and terrestrial and space nuclear energy 

developments. MTAS found that the augmentation of NEP with high thrust chemical propulsion reduces the 

required NEP power level by a factor of 3 to 4 to perform the objective mission, which should simplify ground 

testing and in-space deployment. NTP has a significant technology challenge to provide long-term, low-loss 

storage of liquid hydrogen (LH2) as a cryogenic propellant, which is critical to maintain performance. NEP as 

proposed uses supercritical xenon propellant, which is much denser than hydrogen and easier to store but 

requires increased terrestrial production to meet the quantities needed for human Mars missions. Furthermore, 

the hybrid approach selected for NEP uses liquid oxygen (LOX) and liquid methane (LCH4) as propellants, which 

simplifies cryogenic storage compared to LH2 due to the higher storage temperature. By operational comparison: 

NTP has key utility as the highest Isp high thrust capability of currently achievable technology, which benefits 

human Mars missions with rapid propulsive response to contingencies, and comparatively short capture into and 

departure from planetary gravity wells; NEP is a low thrust system with three-to-ten times higher Isp resulting in 

greater performance potential for advanced missions, and with chemical propulsion adds the high thrust benefits 

of NTP. Previous studies suggest NEP is more versatile and extensible (Dudzinski, 1995; Dudzinski, 1992) than 

NTP when considering a range of applications for subscale precursor development, supporting science missions 

at 10-100 kWe, and more advanced follow-on human missions (Moomaw, 2005; NASA, 2011; National 

Academies, 2022) however, additional studies are needed, and this has been captured on the list of follow-on 

work. NTP requires approximately three times more unique launches to support the campaign than NEP/CCP

for the MTAS piloted mission objective, relying on Super Heavy Commercial Launch Vehicles largely because 

of the low density of hydrogen propellant.

Both NTP and NEP require an aggressive development program if we are to meet a 2030s launch date.

Furthermore, the federal policy guidance in SPD-6 to emphasize High-Assay, Low-Enriched Uranium (HALEU) 

presents an additional schedule challenge because it is a departure from the High-Enriched Uranium (HEU) 

heritage fuel and reactor development, which established heritage feasibility for both systems. HEU has been 

the focus for space nuclear power and propulsion fuel since the dawn of the space age because it results in the 

simplest, highest performance systems. HALEU will require advanced technology for fuels and moderators, 

adding cost, schedule, complexity, and development risk. The only reasons to consider HALEU for space 

applications is to address proliferation concerns with the use of HEU, and security and licensing for the handling 

of HEU. However, robust public-private partnerships for space nuclear power and propulsion are possible with 

HEU systems utilizing existing licensed facilities at U.S. government labs and in industry. 

NASA is currently investing in NTP components, with near-term plans for fuel characterization at the Department 

of Energy (DOE) Idaho National Laboratory (INL). NASA is also investing in the critical technologies for cryogenic 

fluid management of liquid hydrogen (as well as liquid oxygen and methane) to solve the propellant transfer and 

long-term storage challenges crucial for both NTP and cryogenic chemical propulsion systems. NTP is also 

receiving initial investment from DOD/DARPA due to its high thrust acceleration and rapid response capability 

with potential to enable more agile cislunar operations. If developed appropriately, the DOD systems could be 

adapted for NASA use, but coordinated development will be essential. Progress-to-date represents a small 

fraction of development needed for human Mars missions, and current HALEU NTP designs are a departure 

from the HEU design heritage established in Project Rover/NERVA in the 1960s.

NASA currently does not have a coordinated development activity for an NEP system but is investing in 

technologies for LOX-LCH4 chemical propulsion, and is independently investing in several relevant subsystem 

technologies, such as electric propulsion and high voltage PMAD, among others. NEP may benefit from current 

synergistic investment in relevant technologies for terrestrial applications that can be leveraged by NASA, 

including megawatt-scale microreactors and power conversion, etc. NEP also has technology heritage from the 

Snapshot Program (Systems for Nuclear Auxiliary Power), the Space reactor Prototype (SP-100) program, and 

Project Prometheus, which was developing NEP for the Jupiter Icy Moons Orbiter mission. The heritage 
NASA 10

technology from these investments offers a promising starting point for Mars-relevant NEP development. 

Currently, a complete NEP concept based on a ~10kWe evolution of the Kilopower reactor is also under 

consideration for development by the U.S. Space Force (Helios-N). This development could be leveraged by 

NASA for science and Mars NEP precursors.

The MTAS Programmatic Assessment found moderate (~37%) cost advantage for NEP/CCP over NTP when 

considering both development and first Mars campaign use. The MTAS assessment found a ~16% lower Design, 

Development, Test, and Evaluation (DDT&E) cost for NEP. However, the costs for both have large uncertainties 

due to technology challenges and obtaining accurate DDT&E cost estimates will require further technology

investments to benchmark performance projections. Since the end of Project Rover/NERVA, open-air testing at 

the Nevada Test Site is no longer acceptable. NTP development and qualification requires a full scale, integrated 

ground test, which requires a costly new facility to capture and process the, likely radioactive, propellant exhaust. 

This facility is a major development challenge with significant cost and schedule uncertainly. The development 

of NEP will also require some facility development, notably large, high-pumping capacity vacuum chambers for 

testing multiple large electric thrusters. That said, several nuclear facilities are in development for terrestrial 

microreactors that could be leveraged or adapted for the development of space reactors supporting NEP.

Furthermore, the characteristics of NEP system development allows for separate subsystem ground testing of 

modular elements and simulated interfaces, which could be accomplished in separate facilities, easing facility 

development requirements.

Additional distinctions between NTP and NEP/CCP emerge when considering an ongoing Mars exploration 

campaign. A benefit of higher specific impulse electric propulsion is the mitigation of the energetic challenges 

that result from more distant planetary alignments between Earth and Mars reducing the variance in mission 

performance across the synodic cycle. An NEP reactor has greater lifetime potential than NTR for vehicle 

reusability. Furthermore, this study’s implementation of NTP utilizes significant drop-staging to minimize roundtrip propellant mass, which requires more hardware production and Earth launch to support a subsequent 

mission. A majority of NEP/CCP concept hardware is returned to Earth. Finally, NEP/CCP offers cost 

advantages over NTP for development and first Mars campaign use, and with reuse cost advantages, the 

NEP/CCP cost advantage could be multiplied for an ongoing Mars exploration campaign. 

In summation, NTP offers high-thrust capability for rapid response to contingencies and gravity-well departures 

to support a human Mars mission with higher efficiency than chemical propulsion. NTP advantages result from 

the highest Isp for currently achievable high-thrust propulsion. However, low-thrust NEP, when augmented with 

high-thrust LOX-LCH4 chemical propulsion looks very promising as the more advantageous alternative. The 

NEP/CCP approach results in comparatively: fewer launches with fewer number and types of elements; lower 

launch costs, unit costs & campaign costs; and easier launch, storage, and management of propellants.

Furthermore, NEP, when augmented with Chemical Propulsion offers both mission agility and architecture 

flexibility, whereby separate high and low thrust propulsion elements can be used where advantageous. In 

comparing the technology challenges, the study found that NTP has at least three high risk potential 

showstoppers: a reactor fuel achieving 900 sec Isp while solving the hydrogen material and reactivity interactions; 

development of a full-scale engine test facility; and long-term, zero-leak hydrogen cryogenic fluid management

(CFM). Whereas NEP, though requiring the integration of 6 subsystems, there are no subsystem technologies 

that are high risk to be show-stoppers2. Additionally, there are important NEP synergies and feed forward from 

terrestrial energy and lunar power and leveraging of existing electric propulsion (Gateway SEP Power and 

Propulsion Element), and PMAD (electrified aircraft). There are additional synergies with ongoing LOX-LCH4

2 Measured by lower average and peak Advancement Degree of Difficulty (AD2) provided by independent NESC study.
NASA 11

propulsion development. Many of the NEP challenges will be addressed by fission power for the lunar surface, 

and electric propulsion for Gateway. Other studies (Noca, 2001; Jones, 1984; Kokan, 2021; Pawlik, 1977; 

Cassady, 2007) have identified NEP in general has potential applications to robotic science, and missions 

beyond Mars, which represent opportunities to leverage NEP development for precursor and follow-on 

applications (Oleson, 2019). Similar benefits from NTP technology to other applications are more limited (Houts, 

2012; Dudzinski, 1995). In consideration of the broad NASA needs for propulsion, and the broader national 

needs, both NTP and NEP possess an imperative for investment, as they each have noteworthy advantages in 

specific applications. However, if the Mars NEP/CCP performance metrics can be verified in technology 

development, NEP/CCP would be the more advantageous option for human and cargo transportation to Mars.

To support a 2030s Human Mars Mission, committed and significant investment in nuclear propulsion must begin 

now. NASA is currently investing in NTP to support synergy with other agencies and Congressional direction; 

however, the current state of understanding is insufficient to make a down-selection. The results of MTAS

suggest a balanced portfolio of investment in both NTP and NEP is necessary in order to answer key feasibility 

and development questions where the risks are sufficiently understood and retired to focus investments for future

Mars transportation needs. A coordinated technology research plan is suggested, with interim go/no-go decision 

points, and key figures of merit to inform a down-select decision. However, current NASA funding is insufficient 

to support a dual-track approach. NASA needs flexibility to rebalance investments between NTP and NEP to

support an informed down-select decision as soon as practical.

The decision on propulsion for human transportation beyond the Moon could set the course of human exploration 

for decades to come. It will be a decision that defines a decade-long development program costing more than 

$10B. This decision must be made based on sound, objective analysis, and strong technical confidence, for 

which the current state of knowledge is insufficient. Thus, MTAS Phase 1 recommends more detailed analyses 

to guide investments along with active funding of strategically focused investments in both NTP and NEP to 

inform the propulsion decision for humanity’s first interplanetary spaceship.

2 . 0 I N T R O D U C T I O N 

The Mars Transportation Assessment Study was supported by most of NASA’s best experts in propulsion, 

power, mission design, trajectory analysis, cost analysis and programmatic development to assess the 

challenges with human transportation to and from Mars and added the expertise of the Department of Energy in 

nuclear engineering to assess the nuclear aspects of the focus propulsion technologies. Together the team 

provided the Agency and the nation with an objective assessment of potential propulsion technologies to achieve 

the ground rule objectives and provided guidance for an appropriate transportation architecture, concept of 

operations, integrated Mars transportation campaign, and development approach for the required capabilities.

Concurrent to MTAS, the National Academies of Science (NAS) performed a similar study of nuclear propulsion 

options and received input from the MTAS team. Though the NAS study was very thorough, it concluded before 

the MTAS study completed iteration 2. Thus, the NAS did not receive a final report on the two propulsion 

systems, particularly the revised NEP/CCP concept. As a result, the NAS study is complimentary to MTAS, but 

does not reflect the most recent assessment of the technologies. 

The objective of Phase 1 of the Mars Transportation Assessment Study was to provide sufficient guidance to 

establish a technically grounded investment plan for propulsion technologies necessary to accomplish the 

reference Agency concept of operations for a two-year class human Mars Mission. This ConOps is defined in 

the ground rules in Section 3.0. A two-year round trip Mars mission requires an opposition-class trajectory, which 

requires the propulsion systems to perform more ∆V than conjunction-class trajectories, which are optimized to 
NASA 12

minimize ∆V (or change in vehicle velocity – a measure of propulsive effort3). Mars opposition missions can 

achieve total trip times of two years or less with short stays at Mars, whereas conjunction missions are generally

three years or more in total duration with required Mars stays of almost two years while waiting for optimum

planetary alignment. However, higher ∆V missions, such as opposition class missions, put more emphasis on 

the performance of propulsion technologies (and highlights the distinctions between options).

The ∆V required for a Mars mission varies across mission opportunities defined around Earth-Mars closest 

approach, which occurs every ~26 months. Some opportunities are relatively easy such as 2033 and 2035, 

while other opportunities can be very challenging such as 2039 and 2042. Furthermore, the required ∆V varies 

much more significantly for opposition class trajectories, since more distant planetary alignment affects both 

outbound and inbound legs of challenging opportunities. Figure 2.0-1 shows the ∆V variance for both conjunction 

and opposition class missions from 2032 to 2050. The ∆V for the most difficult opposition class opportunities 

can be as much as twice the ∆V of the average conjunction class opportunity. The alignment of Venus with Earth 

and Mars is an important consideration for opposition class missions; most opposition opportunities offer a 

beneficial Venus Gravity Assist (VGA) on either the outbound or inbound leg, which can reduce the overall ∆V. 

Figure 2.0-1 depicts where a Venus gravity assist is opportune or not by color variance. While the study 

considered all the Mars mission opportunities in the mid to late 2030s and early 2040s, the 2039 opportunity was 

used as the sizing case for both the NTP and NEP options since it provides a reasonable bounding case for 

performance that can support all but the most difficult opportunities. This approach recognizes that any longterm endeavor, such as a human Mars exploration campaign, is likely to miss targeting the most advantageous 

opportunities, and will ultimately need to support most opportunities in the synodic cycle.

Figure 2.0-1. ∆V variance across Mars opportunities 2032-2050.

3 Refer to appendix X for full definition.
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The study was conducted in two iterations: Iteration 1 efforts are referred to as “1.1”, and Iteration 2 efforts are 

referred to as “1.2”.

Iteration 1: Proof of Concept

The objective of the first iteration was to establish a proof of concept that the required propulsion technology can 

realistically meet the agency human Mars transportation ConOps in the timeframe of interest, which was 2035.

To accomplish this, the study team assessed the propulsion technology, elements, operations, infrastructure, 

and programmatics necessary to meet the ConOps ground rules, documenting the necessary assumptions. For 

each propulsion option, technologies for fuels, reactor designs, coolant schemes, power conversion, electric 

propulsion, and related subsystems were chosen based on expert assessments of the most mature, lowest risk 

options to meet performance goals in the mission timeframe. Thus, the MTAS reference cases provide a proof 

of concept based on a risk mitigation approach, and do not represent optimal performance from an exhaustive 

trade study. In many cases, alternative trades exist that should be evaluated as proposed in follow-on efforts 

from this study.

Subsequently, the study identified the areas where the technology, mission, campaign, and programmatics are 

high risk and/or insufficient in order to address these issues in a second iteration.

At the completion of Iteration 1, a defined process point in the design of several alternatives, an internal nonadvocate critical review was performed to assess the designs and critique the process. This “Red Team” 

representing a balance of relevant technical experts and nonadvocates invited from NASA (not part of the MTAS 

team), DOE, and the private sector, reviewed and provided a valuable independent assessment to ensure the 

credibility of the study results. Iteration 2 proceeded to address the challenges identified by the Red team, which 

included significant development risk for the technologies required in 2035, and significant cost, schedule, and 

operational risk for the number of launches required to support each option.

Iteration 2: Risk reduction and concept refinement

The objective of the second iteration was to identify alternative technologies, mission approaches, campaign 

approaches, and architectures to address the risks and insufficiencies from the proof of concept. After revising 

the 1.1 mission, design, and performance, the study team reassessed the technology necessary to support the 

revised ConOps ground rules, again documenting the necessary assumptions. In some cases, where risk was 

assessed as high, or performance was insufficient, technology subsystem trades performed resulting in 

alternative technologies for the 1.2 reference cases. DOE representatives were added to the team to support an 

expert assessment of the nuclear subsystems. Finally, the team reassessed the programmatics to ensure that 

an executable investment plan with acceptable risk was feasible.

The goal of Iteration 2 was to establish one or more reference design points with a balance of overall risk between 

the mission operations and technology development. These reference design points can be used to trade 

alternatives for technology, operations concepts, and mission architectures as a standard of measure to guide 

investment and implementation decisions. Phase 2 of the study is recommended to formally perform these trades 

as the Agency moves forward with formulating a human Mars mission and investing in the necessary 

transportation technology, including defining strategic investments looking toward advanced capabilities.

Due to constraints on study resources, and timing of the Agency need for a proof of concept, MTAS focused on 

two promising propulsion technologies: Nuclear Thermal Propulsion (NTP), which uses a fission reaction to heat 

a propellant, which is then accelerated using a rocket nozzle to produce thrust in a Nuclear Thermal Rocket 

(NTR) engine; and Nuclear Electric Propulsion (NEP), which converts the heat generated by a fission reaction 

to electricity, which is then used to power a propellant acceleration device, or “thruster”, to produce thrust.

Example system schematics of an NTP and NEP system are shown in Figures 2.0-2 and 2.0-3, respectively. 

MTAS selected these options because of the propellant efficiency advantage, measured in higher specific 

impulse, that NTP (~900 sec) and NEP (~2,000-10,000 sec) would have over Conventional Chemical Propulsion 
NASA 14

(~330 - 460 sec). Greater propellant efficiency reduces the propellant required to perform changes in vehicle 

velocity necessary to transit from Earth orbit to Mars orbit and back. Without an efficient propulsion system, the 

propellant mass for a Mars mission can become prohibitive, especially for fast Mars transits. Higher propellant 

efficiency also enables more robust missions, as the propulsion system can afford propellant margin to respond 

to contingencies.

Figure 2.0-2. Example NTP system schematic from the NESC assessment of NEP/NTP.

One of the challenges for the use of any electric propulsion technology is the low thrust acceleration, limited by 

power, which results in extended time to propel into and out of planetary gravity wells (regions where gravitational 

forces are much higher than deep space). To provide faster transits with NEP, the study identified a hybrid 

NEP/CCP approach, an innovation based on the hybrid SEP/CCP approach identified as a non-nuclear 

alternative for Mars transportation in the 1990s (Uppal, 2020; Davis 2011; Mani, 2020) whereby high thrust

acceleration CCP is used for gravity-well maneuvers. 

MTAS was initially chartered to complete a comprehensive assessment of credible propulsion options for human 

Mars transportation. It is recommended that a future phase of study assess SEP/CCP against the NTP and 

NEP/CCP reference cases. Furthermore, innovative implementations of Chemical Propulsion in alternative 

architectures should be assessed. These additional assessments should be completed objectively to establish 

the justification for the selected Mars transportation propulsion option, especially if it is nuclear.
NASA 15

Figure 2.0-3. Example NEP system schematic from the NESC assessment of NEP/NTP.

Two engineering teams were simultaneously employed to perform mission and systems analysis for MTAS: the 

Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center (MSFC), and the Compass concurrent 

engineering team at NASA Glenn Research Center (GRC). Because of both the diversity of technologies and 

engineering specialties involved, and the comprehensive nature of the assessment, a concurrent engineering 

approach was advantageous. Thus, a diversity of perspectives was gathered to provide a thorough and objective 

assessment, as well as promote greater team understanding of the issues and challenges to produce more 

creative solutions. The two teams studied both propulsion concepts in Iteration 1 in order to provide independent 

validation of results. Because the resulting mission analyses and concept vehicles built around each propulsion 

system were similar, the two teams specialized for Iteration 2 - the ACO team performed two iterations of the 

NTP vehicle design, and the Compass team performed two iterations of the NEP/CCP vehicle design. After the 

first iteration, both teams received feedback from the independent Red Team review. Nuclear reactor subject 

matter experts from the DOE joined the teams in Iteration 2 to provide expertise in the design and testing of the 

reactors and reactor components as the teams refined their technology assessments. Furthermore, a Campaign 

Assessment Team assessed logistics and operational integration of Mars transportation into an assumed 

ongoing Artemis campaign of lunar exploration, and an internal Programmatic Assessment Team reviewed both 

iterations to assess issues of cost, schedule, risk, infrastructure, and implementation. The Programmatic Team 

defined figures of merit for comparison broadly as: Affordability, Safety and Mission Assurance, and Program 

Robustness. A DOE expert also participated on the Programmatic Assessment team to provide estimates of fuel 

fabrication cost and schedule, testing requirements and their cost, and the need for new test facilities, what they 

would cost, and how long they would require to become available. Each sub team presented their progress at 

weekly meetings for review by the entire team to facilitate transparency and coordination across the entire MTAS 

team. MTAS was coordinated with Agency activities through the NASA Federated Board. The study process is 

depicted in the flowchart in Figure 2.0-4.
NASA 16

SEP NEP

Steering Team

NTP EP Nuclear 

Power

Solar 

Power

Chemical 

Prop CFM

M. Elements 

& Concepts

Federated 

Board

Red Team

COMPASS Mission Analysis

Campaign Analysis

Programmatic Analysis

Findings

ACO Mission Analysis

Technology 

Assessments

Assessments of 

Options 

vs 

the ConOps

Review & 

Integration

Ground Rules & Assumptions

Report Writing

EDL

DOD DOE Study

Leadership

Figure 2.0-4. MTAS Process Flow.

This document contains the summary details of the 18-month Mars Transportation Assessment Study. A more 

detailed documentation of the study and results for the reference cases, including proposed technology plans, 

will be published as a NASA Technical Memorandum.

3 . 0 G R O U N D R U L E S A N D A S S U M P T I O N S 

NASA is studying a strategic objective to put humans on Mars and return them safely to Earth in the mid-to-late 

2030s. Furthermore, NASA is considering a minimally viable overall architecture to limit required developments 

and enhance programmatic affordability and achievability. This is, in part, achieved by limiting the stay time on 

the surface of Mars, and thereby reducing the surface infrastructure required to support initial Mars landings. 

Short stays do not require robust habitation, ISRU, or associated large surface power supplies. It is further 

achieved with smaller crew-size and shorter mission time, reducing logistic masses and costs. 

It is envisioned that initial short-stay missions would grow humanity’s understanding of living and working on 

Mars, build infrastructure, and prove it out in order to support later, long-stay missions.

The ground rule architecture for an initial Mars mission campaign consists of a crew mission and the predeployment of assets to support the crew mission and crew operations. An “all-up” mission architecture, whereby 

all required mission assets travel with the crew, was not considered for Phase 1 of this study due to the increased 

requirements on the propulsion systems. The crew missions require a piloted Deep Space Transport (DST) 

consisting of the propulsion system, Deep-Space Habitation (DSH) for the crew, and associated spacecraft 

systems, propellant, and logistics supplies. In support of the crew mission, additional resources must be predeployed to Mars’ orbit and surface. The predeployed systems include the descent system to take the crew to 

the surface of Mars, the ascent system to return the crew from the surface back to Mars orbit, and the surface 

assets needed for the surface mission. The study reference predeploy campaign involves delivering three 25,000 

kg payloads to the surface of Mars. For the transportation system, this translates to delivering three 65,000 kg 

Mars descent systems to Mars orbit. The lander delivery cadence as specified by the ground rule is that two of 

the three landers must arrive at least 18 months prior to crew arrival, and one lander must be delivered ~4 years 
NASA 17

prior to crew arrival. The first lander is intended to be the first, full-scale EDL demonstration, and must be fully 

successful before committing additional assets and the exploration crew to Mars. This ground rule for the second 

and third landers is set to allow the payload enough time to complete autonomous operations in preparation for 

crew arrival4, which retires risk for the mission before the crew departs Earth. 

The Deep-Space Habitat is predeployed to a Near-Rectilinear Halo Orbit (NRHO) about the Moon, where it will 

be co-orbital with the Lunar Gateway. There it will be outfitted, checked out, and perform a shakedown cruise 

as an independent, free-flying vehicle. The DSH is returned to NRHO at the return to Earth after Mars, where it 

will be refurbished and potentially reused. Because of the ground rule, a propulsive rerendezvous with Earth is 

the reference Concept of Operations (ConOps). Also, due to perceived added risk to the crew, a Mars 

aerocapture or direct Mars entry was not considered. However, more propulsive advantageous mission 

approaches are considered for follow-on work.

The key ground rules for the vehicles and mission architectures levied on MTAS Phase 1 are as follows.

• 2-year roundtrip, opposition class mission in 2030s

• 30 day stay for 2 crew on surface, 2 remain in orbit

• Leverage lunar exploration infrastructure where appropriate and advantageous

• Optimize aggregation point for the Deep-Space Transports (DST) for crew and cargo

• The same Deep-Space Habitation (DSH) is assumed for both concepts and will be predeployed in 

NRHO about the Moon and returned to NRHO at the end of mission.

• Mars supporting SLS launch timing must consider ongoing Artemis SLS support

• 8.6 m diameter, long SLS fairing (7.5 m by 19.8 m dynamic envelope)

• Three 65,000 kg Landers predeployed to Mars prior to crew departure

Why two years?

The MTAS ground rules are based on the strategic objective to land humans on Mars and return them safely 

using lowest risk approach for mission success. This includes considering the risk of in-space vs surface time 

for the exploration crew. There remains debate among human mission planners whether there is greater risk to 

the crew in space, where humanity has decades of experience living and working, or on the Mars surface, where 

there is gravity, atmosphere, and some radiation protection, but also is increased infrastructure dependency and 

unknown hazards. The complexity of human operations on Mars, given the number of additional technologies 

required to keep astronauts alive and productive, with no rescue or resupply opportunities, may make the surface 

riskier than space for initial missions – until experience is gained and infrastructure is proven. MTAS studied the

first mission for humans on Mars or its vicinity. It was considered that an initial use of shorter opposition class 

round-trip trajectories, with associated short stays on Mars, may help mitigate these unforeseen risks. Lessons 

learned and experience gained from this initial mission builds toward long-stay, conjunction class missions. 

Furthermore, NASA has studied long-stay missions as the baseline approach in the Mars Design Reference 

Missions (DRMs) since the Space Exploration Initiative in the early 1990s and has numerous studies of 

Conjunction Class missions to understand the challenges. MTAS supported the need to understand the risk and 

benefits of Opposition Class missions.

Secondly, the global data set to date for humans in zero-g is limited to 6-12 months with most individual 

experiences only approaching ~6 months. Interplanetary space environment risks increase with time, such as 

equipment malfunctions, weightlessness, vacuum, radiation from Galactic Cosmic Rays (GCR) and solar particle 

4 The first lander was also discussed as an Entry Descent and Landing (EDL) demonstration, but this needs to be critically 

examined.
NASA 18

events (SPE), and space debris. Many of these risks continue on Mars, where additional crew risks include

physical injury, weather, and failure of surface systems, etc. Furthermore, mission challenges increase with 

time, such as logistics of consumables, and the psychological effects of isolation, constrained living space, and

being far away.

Thirdly, crew radiation from both natural sources, such as GCR and SPE, as well as the propulsion system are 

experienced throughout the mission, though at a somewhat lesser rate on Mars. Radiation is identified as a “Red 

Risk” by NASA as the highest-priority astronaut health challenge. Radiation damages DNA and can lead to 

mutations that can trigger cancers. It causes cardiovascular health problems such as heart damage, the 

narrowing of arteries and blood vessels (Belzile-Dugas, 2021). Radiation also causes neurological problems that 

can lead to cognitive impairment during the mission (Michaud, 2019; Turnquist, 2020; Smart, 2017).

By reducing trip time with mission success and crew survival as the focus, the risks to the crew and mission are

proportionally reduced. For these reasons, two years was established by the Agency as a target round-trip 

mission duration, and MTAS was challenged to assess the viability of the propulsion technology to achieve this 

goal. Significant reductions in round trip time below two years may be possible with advanced NTP and NEP 

technology, but these technologies are considered beyond viability for the 2030s.

The key assumptions and trades are as follows:

1. 2035 is the primary human mission opportunity, and the development target for programmatic 

assessment. 2039 will be analyzed to envelope the performance required for a ~2-year Mars opposition 

class mission across the range of synodic opportunities.

a. Roundtrip mission time is counted as time from Earth orbit departure (TMI) to Earth orbit capture 

(EOC) and is afforded 760 days or less.

b. A Venus gravity assist will be included in the mission when advantageous to add mission value.

c. Earth aggregation orbit for cargo and crew transportation systems will be optimized separately for 

NEP and NTP.

2. The reference case will be a design point with minimum technology to accomplish the Agency objectives 

as defined in the GR&A. The nuclear propulsion unit for the first Mars mission will not be reused, but the 

potential for future units to be reused will be considered.

a. The study will trade off the reference case to identify technology to improve performance, mitigate 

risk, and/or reduce cost.

3. The reference case will use SLS, requiring no more than two cargo launches in a single year. If two are 

required, the following year is limited to one SLS cargo.

a. Commercial Launch Vehicles, including Super Heavy CLVs, will be used to fill in where the SLS 

cannot meet the required launch cadence.

4. Primary reactor trade study variables include fuel type, uranium enrichment, reactor type (fast or

moderated), and moderator for both NTP and NEP. For NEP, additional trades include coolant type, and 

reactor outlet temperature.

5. Crew dose from the onboard nuclear systems will be limited to no greater than the mission GCR dose.

6. First lander supporting the crew mission will successfully land on Mars before the second lander is 

launched to provide a proof of the Entry, Descent, and Landing (EDL) system for this size-class of 

payloads.

7. Remaining landers supporting the crew mission must be in Mars orbit 18 months prior to crew departure.
NASA 19

8. Orion is used for crew rendezvous in LDHEO prior to TMI, and to return the crew at EOM.

4 . 0 M A R S T R A N S P O R TAT I O N A S S E S S M E N T 

Phase 1 of MTAS established two reference design points with a balance of overall risk between mission 

operations and technology development. The following sections summarize these reference cases in terms of 

mission design, vehicle concept, performance, and technology requirements. Also summarized is the 

assessment of each reference Mars mission in the context of an ongoing Artemis lunar campaign. Finally, a 

high-level technology development plan is provided. Greater detail on the MTAS assessments, the reference 

cases, and the technology plans is provided in focused MTAS reports, published as NASA TMs.

4 . 1 N U C L E A R T H E R M A L P R O P U L S I O N R E F E R E N C E C ASE

The trajectory for the NTP reference case is shown in Figure 4.1-1. All velocity changes are performed by Nuclear 

Thermal Rocket engines. This mission is 690 days for the 2039 opportunity as a bounding case for performance. 

This trajectory is a conservative case, assuming an NTR engine-out prior to Mars departure; a 670–680-day

mission is possible with no engine out. The planetary alignments in 2039 provide a Venus gravity assist on the 

in-bound leg which enables a round trip ~30 days shorter than two years for this reference case. A longer mission 

in 2039 would eliminate the VGA, and save propellant mass.

Figure 4.1-1. NTP Schedule and Trajectories for the 2039 Reference Mission.

The landers supporting the crew mission would be delivered to Mars by two, separate NTP systems. This 

approach would provide direct and early experience with the NTP system in deep space prior to the piloted 

mission in 2039, but it would also require earlier development of NTP. Alternately, to offset the development 

schedule risk associated with this earlier need date is the fact that the first cargo mission can easily be performed 

with a subscale nuclear thermal rocket in the 10-15 klbf thrust class, perfectly in line with an envisioned subscale 

demonstrator. For the 2039 reference mission, the single lander would be launched in 2035. In 2037, the 

remaining landers would be launched also using an NTP propulsion system.

Figure 4.1-2 shows the Concept of Operations (ConOps) for the NTP cargo predeployment campaign. The first 

cargo transfer vehicle would be launched to a Medium Elliptical Earth Orbit (MEEO) for aggregation. The first 

cargo transfer vehicle consists of a single NTR “Stub Core” Stage and three Super Heavy Class Commercial 
NASA 20

Launch Hydrogen Inline Stages (CL-HIS). The MEEO of 1,200 km x 7,000 km altitude is determined to maximize 

the mass of LH2 propellant delivered in each drop tank that fills the launch vehicle shroud volume of an SH-CLV. 

The first lander would be launched to MEEO where it is integrated with the NTP cargo transfer vehicle, then 

raised to Lunar-Distant High Earth Orbit (LDHEO) of 400 km x 400,000 km as the departure orbit for Mars. 

Once the first lander is successfully on Mars, the second lander would be launched in 2036 to MEEO and mated 

with a chemical propulsion stage, which will place the second lander in NRHO to wait for the second NTP Mars 

cargo transfer vehicle. In 2037, the final lander and the elements of the NTP cargo transfer vehicle would be 

launched to MEEO. The second NTP Mars cargo transfer vehicle consists of a second, single NTR “Stub Core”

Stage, four CL-HIS, and a Multi-Dock Truss Element (MDTE) to dock two of the CL-HIS and the two landers.

After assembly, the NTP stage would transfer the stack to NRHO to add the other lander, then transfer to the

LDHEO departure orbit for Mars. After arrival, the MAV would descend to the Mars surface and be verified 

operational before the piloted vehicle departs Earth orbit in 2039.

AR&D 

(Multiple)

AR&D 

(Multiple)

x2

NRHO 

5-sol 

Mars Orbit 

202-day transfer

TMI – Depart on June 27th, 2035 SLS 1 Lander CLV × 4

x1 x3

Launches in 

2035

MOI – Arrive on January 15th, 2036 SLS 1 Lander CLV × 6

Launches in 

2037 CLV × 2

Wait in NRHO for NTP 

vehicle SLS 1 Lander

Launches in 

2036

Pick up lander

x2x1

351-day transfer

MOI – Arrive on August 4th, 2038

x1

2-sol 

Mars Orbit 

Element Icons NOT to scale

LDHEO

400 x 400,000 km

MEEO

1,200 x 7,000 km

Figure 4.1-2. Mars 2039 NTP Cargo Lander ConOps.

Figure 4.1-3 shows the ConOps for the piloted DST. Its components would be launched to the MEEO for 

aggregation. The NTP DST consists of: 

• an SLS-launched core stage with two-NTR engines; 

• one SLS-launched Hydrogen Inline Stage (SLS-HIS); 

• two side-mounted “Stub Core” stages, each with two side-mounted CL-HIS; 

• two MDTE, docking fourteen CL-HIS, and a Power and Propulsion Module (PPM). 

Once assembled, the DST would propel itself to LDHEO where it would jettison the two Stub Core Stages and 

their associated four CL-HIS, along with four of the CL-HIS docked to the MDTE. These jettisoned stages could 

later be reused or disposed of in a safe orbit. The Deep-Space Habitation requires a chemical propulsion transfer 

stage to descend from NRHO to LDHEO to rendezvous with the DST. An Orion capsule brings the crew to 

LDHEO and rendezvous with the DST. After adding the DSH and crew, the DST departs for Mars from LDHEO. 

By the time it arrives at Mars, it will discard the remaining ten CL-HIS along the way. It will enter a two-sol orbit 
NASA 21

and two crew members will transfer to the crew lander that will take them to the surface of Mars. After 30 sols of 

surface operations, the surface crew will use the MAV to return to the DST. The DST will then return to Earth

with a Venus Gravity Assist (VGA). NTP would be used to capture the DST back at Earth into LDHEO, where 

an Orion capsule would rendezvous to bring the crew back to the Earth’s surface. The DST would transfer to 

NRHO via NTP for disposal or potential future use.

NRHO 

2-sol 

Mars Orbit 

~297-day transfer

TMI - Depart on April 26th, 2039

Surface Ops 

(30-sol)

Lander Disposal 

via RCS SLS

AR&D 

(Multiple)

Deep Space 

Maneuver (DSM)

Lander 

Descent SLS 

Element Icons NOT to scale NTP Core

Deep 

Space 

Habitation CLV × 4 SLS Crew Crewed Orion

Orion 

Return

Lander Orbital Ops 

from previous cargo 

launches × 3 

Ascent

TEI

EOI

~354-day transfer SLS Crew Orion

Orion 

Return 

(With crew)

Hab Cis-Lunar 

Ops for next 

mission

LDHEO

400 x 400,000 km

Venus 

Venus Flyby

Mars Stay

(~39 days / 38-sol)

MEEO

1,200 x 7,000 km

x2

x4

x4

x8

MOI - Arrive on February 16th, 2040

x2

Shakedown stage 

& logisticsCLV × 1

DSH transfer 

stage

Figure 4.1-3. Mars 2039 NTP Piloted Mission ConOps.

The NTP configuration that would depart Earth orbit for Mars with a crew in the Transit Habitat is shown in center 

isometric view in Figure 4.1-4. The initial MEEO configuration, as well as the Earth departure and Earth return 

configuration are shown in the lower left of the figure. As discussed above, two side-mounted Stub Core Stages, 

and eight of the SH-CLV launched hydrogen inline stages will be jettisoned after the vehicle transfers from MEEO 

to LDHEO. The NTP option requires a Power and Propulsion Module (PPM) to provide the primary power, 

communications, command, and avionics for the full integrated DST vehicle. The DSH concept on the right-side

of the image will rendezvous with the vehicle in LDHEO (see campaign discussion below). It will be connected 

to the PPM and then to the Multi-Dock Truss Element (MDTE) to which the additional CL-HIS (silver in the center 

of the figure) are mounted. Each Commercial-Launch Hydrogen Inline Stage is launched with a spacecraft BUS

for power and orbital maneuvering to reach the aggregation point and has a radiator to reject heat from the 

cryocoolers that maintain the propellant at cryogenic conditions. The MDTE mounts to the SLS-launched 

hydrogen inline stage and the SLS-launched, two-NTR core stage element. Solar panels on the PPM and NTR 

core stage provide 60 kWe of electrical power for the habitat, avionics, communications, reaction control system 

(RCS), and cryogenic coolers. The individual components are described in greater detail in a NASA TM.
NASA 22

DST DST

Cargo Cargo

Mission 1 Mission 2

Piloted DST Piloted DST Piloted DST

at Aggregation at Earth Departure at Earth Return

Figure 4.1-4. Mars 2039 NTP Piloted DST (Earth Departure Configuration).

Figure 4.1-4 also depicts how the NTP Stub-Core Stages and Commercial-Launched Hydrogen Inline Stages 

can be configured to propel the two cargo predeployment missions. The first cargo mission delivers one lander 

using a single subcore stage and three CL-HIS. In the second cargo mission delivering two landers requires an 

additional CL-HIS and a Multi-Dock Truss Element.
NASA 23

Based on MTAS vehicle sizing studies for the 2039 reference case, four NTR engines are needed, each

producing 25,000 lbf of thrust at a specific impulse of at least 870 sec for the mission to be completed with a 

realistic amount of propellant. Reducing Isp below 870 sec would rapidly increase the amount of propellant 

required and increase vehicle mass. The only propellant able to achieve this Isp is hydrogen, which is discussed 

later in this document. For margin, the study selected a design goal of 900 sec average specific impulse, which 

includes engine start-up and cool-down inefficiencies. Each engine also needs to be able to operate for as long 

as 4 total cumulative hours to be able to continue the mission if one engine were to fail before Mars departure. 

The maximum number of engine start-ups (thermal cycles the reactor fuel must withstand) for the mission is 

about eight. 

To achieve 900 sec Isp, the hydrogen must be heated to an exit temperature of 2700K (higher with losses). Peak 

temperatures in the fuel can be a few hundred degrees above the propellant exit temperature from the reactor. 

To achieve these temperatures, the reactor fuel elements must be made of materials that can hold together at 

the high temperatures and withstand corrosive exposure to hot hydrogen. The reactor must have the endurance 

to provide the Isp and thrust over the thrust lifetime to accomplish the mission requirements with reliability margin. 

Based on four hours of operation, CerMet (ceramic and metal) types of fuel elements are projected to reach Isp

close to 870 seconds and CerCer (all ceramic) can reach an Isp of 900 seconds with margin. However, CerCer 

fuels have never been subject to NTP conditions, and their performance is planned to be proven under NASA’s 

NTP development efforts. Two promising advanced carbide fuels are also predicted to be robust under these 

conditions and offer the potential for even higher Isp. Fuel maturity and development challenges, as well as 

associated reactor design challenges, are being considered in the selection of a fuel that can meet these 

performance requirements by the late 2030s.

4.1.1 NTP Technology

The NTP development activities can be defined in three phases: (1) technology maturation, subscale component 

development, and non-nuclear testing; (2) full-scale development, integration, and testing; and (3) fabrication 

and assembly of the flight Nuclear Thermal Rocket (NTR) engines and associated stage systems. The MTAS 

NTPS development roadmap is depicted in Figure 4.1-5. Phase 1 includes design, fabrication, and testing of 

fuel elements. Phase 2 includes design, fabrication, and testing of the reactor and engine. Once a fuel and fuel 

element design are selected, fabricating the fuel will take an estimated ten years for the six NTR flight engines 

(one for each of two cargo vehicles, four for the piloted DST), plus a development qualification engine. Reactor 

development and qualification will also drive the overall schedule. Including the design, construction, training, 

Nuclear Quality Assurance (NQA-1) certification, and testing, the process is estimated to take nine years. For 

the development testing, extensive facility modifications will be necessary and new facilities will be necessary

for integrated full-scale testing. The development roadmap culminates with the fabrication, assembly, and 

operational testing of the mission vehicles. Based upon this schedule for NTP development, it is unlikely NTP 

can support a Mars campaign until the late 2030s, assuming a fuel selection by 2025.

The first focus of an NTP development program will be the development of the fuel elements, other reactor 

components, non-nuclear engine components, instrumentation and control, and system engineering and 

integration. Fuel and fuel-element development will start subscale in Phase 1, proceeding to full scale design 

and testing in Phase 2. The best fuel and fuel design will be determined through subscale nuclear and nonnuclear testing. Full-scale design and development of the tanks, avionics, stage power, guidance navigation and

control, and stage structures are allocated more than ten years and will be worked in parallel during Phase 1. 

The most capital-intensive and time-consuming element of either vehicle concept is the development and 

qualification of the new reactor. This includes manufacturing large quantities of the selected fuel, designing and 

building specialized test facilities, and building and qualifying a unique reactor design, which are all long leadtime activities for Phase 2.
NASA 24

Since the end of Nuclear Thermal Rocket testing in the Nevada desert during Project Rover/NERVA, open air 

testing is no longer acceptable, and the development of an NTP test facility has been a major challenge to cost 

and schedule. The fuel, reactor core components, reflectors, Instrumentation & Control (I&C), and shielding, as 

well as the non-nuclear rocket engine components will all need to be developed and tested as an integrated 

system for flight qualification. This will require a new facility to capture and process the rocket exhaust, which 

may contain radioactive elements. Demonstration of viable NTP ground test exhaust capture and filtering to 

support full-scale facility development is a second focus of NTP technology maturation in Phase 1 to retire 

development risk for Phase 2.

A third focus for NTP Phase 1 technology maturation is the demonstration of Cryogenic Fluid Management 

(CFM) technologies necessary for long-duration hydrogen propellant storage and transfer. Storing large 

quantities of cryogenic propellants with negligible losses has always been a challenge. Storing liquid hydrogen 

for years compounds the challenge. Hydrogen is a small molecule and can readily leak through very small gaps 

in valve seats and seals at other interfaces. Further, to prevent boil-off losses, liquid hydrogen must be 

maintained at extremely cold temperature (20 K or -423 °F). Zero boil-off (ZBO) of liquid hydrogen and very low 

hydrogen leakage during the two-year mission has been assumed. Achieving this will require the development 

of no/low leakage valves and couplers, advanced insulation, and high-capacity, high-efficiency cryocoolers. 

Finally, the NTP concept leverages integration of multiple liquid hydrogen tanks to satisfy the full propellant 

requirement. Managing a cryogenic propellant, like liquid hydrogen, during microgravity coasts greater than a 

few hours and transferring that propellant from one tank to another in space has never been demonstrated; these 

represent additional technology maturation challenges. The Artemis missions, and other commercial space 

transportation developments share the need to solve the challenge of long-term cryogenic fluid management 

and transfer for hydrogen as well as other cryogens. NASA and industry are currently investing in technologies 

to solve these challenges, but many of the technologies remain low maturity.

Because of NASA contract obligations related to NTP development at the time of this writing, a more detailed 

technology plan is not provided here.
NASA 25

FY20 FY21 FY22 FY23 FY24 FY25 FY26 FY27 FY28 FY29 FY30 FY31 FY32 FY33 FY34 FY35 FY36 FY37

SNP Project

Level II Schedule - NTP

Non-Nuclear Engine Development 

Turbo Machinery

Chamber/Nozzle Thrust Chamber 

Assembly

Engine Valves

Lines & Ducts

Instrumentation & Controls

Materials Radiation Testing

25K Engine Testing Facility

Eng Sys

Integr. Begins

Subscale

Eng Del Dev

Eng

Dev

Eng

Cert

Eng

Cert

Eng

Flt

Eng

Flt

Eng

Flt

Eng

Flt

Eng

Flt

Eng

Full Scale Engine

Assemblies

PDR TRR

Subscale Component Test

Full Scale Component Dev

Full Scale Component Test

SRR PDR TRR

Subscale Component Test

Full Scale Component Dev

Full Scale Component Test

Acq., Test and Integration

Subscale Component Test

Full Scale Component Dev

Full Scale Component Test

Design, Analysis & Testing

Subscale Component Test

Full Scale Component Dev

Full Scale Component Test

Full Scale Component Dev

Full Scale Component Test

Design of Experiments, Acq. and Test of Components

Annual Facilities Reviews

Assembly & GSE Facility Facility Readiness Review

25K Engine

Test Facility ATP Facility Readiness Review

Environmental Impact Statement

NEPA

E.I.S. Complete

NEPA Complete

RECS Construction

Exhaust Capture System Test

Test Stand Mods

Real-Time Exhaust Filtering

Integrate Subscale Engine into Test Stand Test Readiness Review

Subscale Engine Testing Subscale Engine Test Data Report

Adaptation for 25K Engine Test Stand Readiness Review

Integrate Full Scale Engine into Test Stand Test Readiness Review

Full Scale Engine Dev & Cert Testing

Dev

Eng

Test

Dev

Eng

Test

Cert

Eng

Test

Cert

Eng

Test

Certification Series

Test Report

FY20 FY21 FY22 FY23 FY24 FY25 FY26 FY27 FY28 FY29 FY30 FY31 FY32 FY33 FY34 FY35 FY36 FY37

SNP Project

Level II Schedule - NTP

Subscale Engine(SE) Milestones

Full Scale Stage(FS) Milestones

Full Scale System Milestones

Tanks

Avionics

Main Prop System (MPS)

Stage Structures

Fuel Development

CerMet

CerCer

Fuel Testing

Coated UN Production

Fuel Rod Production

Reactor Design & Development

ATP

SRR/CoDR-SE PDR-SE CDR-SE SNP SE Delivery Boots on Mars

FS ATP FS SRR/CoDR FS PDR FS CDR Stage to Vehicle Integration

SLS Launch

SLS Launch

Aggregate for Mars

FS Systems

SRR/CoDR

FS Systems

PDR

FS Systems

CDR

Core Stage

Del. for Integ.

DCR

DCR

Drop & In-Line Tank Deliveries Complete

Fuel Qual

Plan

Coated UN 

Particle Study

Fuel Dev Plan

Decision Gate Fuel Downselect for Demo

SIRIUS 2 Test

SIRIUS 3 Test

PRIME 1 PRIME 2

NTREES 3 Test NTREES 5 Test

NTREES 4 Test NTREES 6 Test

Cermet Transient Thermal Mechanics Test Ceremt Vol Max Test

Moderator Fab. TREAT Hydrogen Loop

Facilities Prep Production

Facility Prep Production

RFP/BAA Prep

ATP

 Base Period

30% DR

Option 1

90% DR

Option 2

Full Scale Fac. Prep.

Subscale NTR Del.

Full Scale Reactor Build

Full Scale Reactor Del.

CLV Launches

Figure 4.1-5. NTP Development Plan executing to a 2034 launch date.

4.1.2 NTP Campaign

For the Nuclear Thermal Propulsion option, the mission architecture has six unique transportation system 

elements (in bold, below), and requires a total of 32 launch elements to support the piloted mission, including 26 

that are unique to the NTP campaign. These elements are: 

• Space Launch System launched Nuclear Thermal Propulsion Core Stage (1)
NASA 26

• Space Launch System launched Hydrogen Inline Stage (1)

• Heavy CLV launched Nuclear Thermal Propulsion Stub Core Stage (2)

• Heavy CLV launched Multi-Dock Truss Element (2)

• Heavy CLV launched Power and Propulsion Module (1)

• Super Heavy CLV launched Hydrogen Inline Stage (18)

• Heavy CLV launched Chemical Propulsion Stage for DSH transfer (1)

Common piloted launch elements:

• Space Launch System Orion launches for crew delivery and return (2)

• Heavy CLV launched Deep Space Habitation (2)

• Heavy CLV launched Chemical Propulsion Stage for DSH shakedown cruise (1)

• Heavy CLV launched logistics for Deep Space Habitation (1)

As an example of the variation between different opportunities in the Earth-Mars synodic cycle, a 2035 mission 

would require fewer than half the commercial launched hydrogen inline stages needed for the 2039 mission.

The Deep Space Habitation (DSH) is launched early in the campaign to support Artemis and Gateway activities. 

The Mars mission campaign includes two H-CLV launches to deliver and outfit the DSH in NRHO where it will 

be checked-out at Gateway. Then a third H-CLV launch provides a chemical propulsion stage to perform a DSH 

shakedown cruise. Finally, an H-CLV delivers logistics to the DSH prior to Mars departure.

The Nuclear Thermal Propulsion Core stage is the primary propulsion system for the integrated vehicle. It is 

launched on a Block 2 variant of the SLS. The Nuclear Thermal Propulsion Stub Cores Stages are smaller 

elements that can be launched on heavy class commercial launch vehicles. The architecture requires two of 

these Stub Core Stages for the crew mission. There are two types of inline stages to hold the hydrogen propellant 

needed for the mission: one large inline stage launched on a Block 2 variant of the SLS and eighteen smaller 

inline propellant stages that are sized to launch on super heavy class commercial launch vehicles. Two, multidock truss elements are required to support the propulsion system and the hydrogen inline stages that make up 

the integrated stack. These are launched on heavy class commercial launch vehicles. The PPM element is also 

launched on a heavy class commercial launch vehicle. Finally, as with the NEP/Chemical hybrid option

(discussed in the next section), the DSH is prepositioned in cislunar NRHO for checkouts. However, because 

the NTP vehicle does not travel to cislunar space for aggregation, an additional Heavy-CLV launched chemical 

propulsion stage is needed to retrieve the DSH for rendezvous in LDHEO with the transportation system to 

support the piloted mission to Mars. 

The launch campaign for the NTP reference case is depicted in Figures 4.1-2 and 4.1-3. For the campaign build 

up, the Nuclear Thermal Propulsion option optimized its aggregation orbit to a single medium elliptical Earth orbit 

of 1,200 km x 7,000 km altitude, which is considered safe for nuclear operation. All transportation system 

elements are aggregated in this orbit except for the Deep Space Habitation, and unlike the NEP/Chemical Hybrid 

option, all elements in the NTP option are launched fully fueled to the aggregation orbit. This orbit was chosen 

to maximize the orbital energy provided by the SH-CLV while at the same time maximizing the mass of LH2

propellant delivered in each hydrogen inline stage that fills the launch vehicle shroud volume. The optimization 

of the aggregation orbit for NTP crew and cargo vehicles was the major campaign improvement between Iteration 

1 and Iteration 2. Iteration 1 considered NRHO as the aggregation point for both options, significantly reducing 

the payload delivery capability of all launch vehicle options, while placing the Mars propulsion systems in a much 

higher energy orbit, reducing their necessary ∆V. By optimizing the aggregation orbit vs NRHO, the launch 

campaign unique to NTP for the piloted mission was reduced from 72 launches to 26
NASA 27

The NTP predeploy campaign is supported by initial uses of transportation system elements for the piloted 

mission, and requires a total of 15 launch elements, including 12 that are unique to NTP. These elements are: 

• Heavy CLV launched Nuclear Thermal Propulsion Stub Core Stage (2)

• Heavy CLV launched Multi-Dock Truss Element (1)

• Super Heavy CLV launched Hydrogen Inline Stage (7)

• Heavy CLV launched Chemical Propulsion Stage (2)

Common cargo launch elements:

• Space Launch System launched Mars Landers (3)

For the NTP launch campaign, the landers will be delivered to Mars by using the smaller Stub Core and series 

of hydrogen inline stages, all launching on commercial launch vehicles. The first lander, delivered on the 2035 

Mars conjunction mission opportunity, utilizes a single Stub Core Stage launched on a Heavy-CLV and three 

Super-Heavy-CLV-Launched hydrogen inline stages (see Figure 4.1-6). For this mission, the transportation 

system will rendezvous with the lander at the MEEO drop off point of the lander by the SLS before departing for 

Mars. The second and third landers, delivered on the 2037 Mars conjunction mission opportunity, utilizes a single 

Stub Core Stage, four CL-HIS, and a single MDTE launched by an H-CLV. To help reduce the launch cadence 

requirement of the SLS, the second lander is boosted to cislunar space for aggregation using commercial launch 

vehicles and the third lander is picked up by the NTP cargo transfer vehicle in MEEO. After the NTP cargo 

transfer vehicle rendezvous with the third lander, the system transits to cislunar space to retrieve the awaiting 

second lander before departing for Mars.

Figure 4.1-6. NTP Cargo Vehicles and Piloted Vehicle Launch Campaign.

Overall, to support the complete crew mission to the Mars surface, the Nuclear Thermal Propulsion option 

requires a total of forty-seven launch vehicles for the full campaign, including 38 that are unique to NTP. These 

launches are mostly clustered between 2035 and 2041 (including the crew return), with significant concentration 

in 2037 and 2038. A total of seven Space Launch Systems are required: two crew variant and five Block 2 cargo 

variant. Fifteen heavy class commercial launch vehicles deliver chemical propulsion stages and supporting
NASA 28

elements where propellant volume is not a consideration, like the Stub Core Stages, the Power and Propulsion 

Module, and the MDTE for the cargo and crew transfer vehicles. Finally, twenty-five super heavy class 

commercial launch vehicles are needed to deliver twenty-five inline stages full of hydrogen.

4 . 2 N U C L E A R E L E C T R I C P R O P U L S I O N R E F E R E N C E C ASE

The 2039 design reference mission interplanetary trajectory for the hybrid NEP – Chemical Propulsion concept

is shown in Figure 4.2-1. The trajectory is optimized to employ the high thrust and low thrust systems where they 

are most advantageous. High thrust chemical propulsion is employed to propel the vehicle away from 

Earth/Moon at Trans Mars Injection (TMI), then again at Mars Orbit Insertion (MOI) and Mars Departure/Trans 

Earth Injection (TEI) after which point it is jettisoned. The low-thrust NEP system would continuously accelerate 

the vehicle more than halfway to Mars, and then decelerate it for the remainder of the trip. The return trip would 

be accomplished after firing the chemical stage and using the NEP thrusters to escape Mars followed by an 

extended coast for a large fraction of the return trip. After a Venus Gravity Assist (VGA), the NEP system will 

decelerate the vehicle so it can be captured into a high elliptical orbit around Earth. The total time away from 

Earth would be 760 days (25 months).

Figure 4.2-1. NEP/Chem Schedule and Trajectories for the 2039 Design Mission.

Figure 4.2-2 shows the ConOps for the cargo predeployment campaign utilizing the NEP/CCP elements. The

first cargo transfer vehicle is a unit of the Chemical Propulsion Module (CPM) also used for the piloted DST. It 

is launched partially filled to NRHO, based on the capability of a Block 2 SLS, and LOX and LCH4 propellants 
NASA 29

are brought to NRHO to fill the LOX and LCH4 propellent tanks. The first lander would be launched at MEEO, 

with a maximum energy determined by the capability of an SLS and assembled and fueled there. The lander is 

raised to NRHO by chemical propulsion upper stages for integration to the CCP transfer vehicle and from there 

departs for Mars. 

After the first lander is successfully on Mars, the second and third landers are similarly shown being assembled 

and fueled in MEEO before being raised to NRHO. This happens simultaneously with the assembly of the first 

unit of the NEP Propulsion Module (NEPM) with a Xenon Interstage Module (XIM) in 500 km circular LEO, to be 

used for the second cargo transfer vehicle. The NEP cargo transfer vehicle is raised to a 1100 km circular 

nuclear safe orbit by a chemical propulsion upper stage, where the NEP system is commissioned and used to 

raise the vehicle to NRHO. Once the NEP cargo transfer vehicle and the two landers are mated, they depart for 

Mars from NRHO. The Mars Ascent Vehicle (MAV) descends to the Mars surface and is verified operational 

before the piloted vehicle departs Earth orbit.

Mars

Earth

NEP to Mars 

(C3 ~ 2)

NEP Thrusting

Chemical Burn

LGA

Landers

Chemical 

raise 

Lander launch and delivery to 

NRHO occurs simultaneously 

with NEP deployment and spiral 

to NRHO

Dock 

Landers to 

NEP

~6 month NEP 

spiral to NRHO

MAV descend to 

surface to prepare for 

crew

Reuse NEP stage for further Mars 

missions: Comms, safe haven, taxi, 

fuel depot, alternative return 

stage...

NEP maintains crew lander for 

until crew arrives

~500 day NEP 

spiral to Mars 5 

sol

NEP 

Launch/mating: 

~2 months Super Heavy Xe Interstage SLS: NEP Element SLS : Chem Stage Propellants SLS : Lander #2

Dispose 

chem 

stage

Reactor 

Commission: 

~2 monthsSLS : Lander #3

Chemical To Mars Lander 

#1

Cargo lander descends 

to surface to prepare 

for crew

Chem 

Stage 

Disposed

Chem/NEP Cargo Option: 5 SLS, 1 super Heavy for Xe Interstage and 11 CLV stages/tankers SLS : Lander #1

Figure 4.2-2. Mars 2039 NEP/Chem Cargo Lander ConOps.

Figure 4.2-3 shows the ConOps for the NEP/CCP piloted Deep Space Transport mission. Its components are a 

second unit of the NEP Propulsion Module, a second unit of the Chemical Propulsion Module (CPM), and three 

Xenon Interstage Modules containing additional propellant. These components are delivered to a 500 km 

circular LEO to utilize the maximum payload capability of Block 2 SLS and Super-Heavy CLVs. Once assembled, 

the Chemical Propulsion Module propels the DST more rapidly through the low Earth orbits where it will need to 

avoid a relatively high concentration of space debris to an 1100 km nuclear safe orbit. Once the NEP system is 

commissioned, and the Chemical Propulsion Module propellants are refilled, the NEP Propulsion Module propels 

the DST through a spiral trajectory to NRHO. The DST will rendezvous with the Deep Space Habitation and 

jettison two depleted XIMs in NRHO. Using NEP, the Deep-Space Transport transfers to Lunar-Distant High 

Earth Orbit of 400 km x 400,000 km altitude, where Orion will rendezvous with the DST to bring the crew. The 

DST departs LDHEO for Mars via a chemical propulsion burn. NEP is used in interplanetary space to accelerate 

the trajectory, and then chemical propulsion is used to capture the DST at Mars before a series of NEP burns 
NASA 30

brings the DST to a two-sol staging orbit. In a two-sol orbit, two crew members transfer to the lander that takes 

them to the surface of Mars. After 30 sols of surface operations, the surface crew uses the MAV to return to the 

DST. Following the final chemical stage burn, the chemical stage is jettisoned, and the DST returns to Earth 

using NEP and via a Venus gravity assist. Finally, a Lunar Gravity Assist (LGA) maneuver captures the DST 

with the crew and DSH back to LDHEO. A second Orion capsule rendezvous with the DST in LDHEO to bring 

the crew home. The DST is parked in NRHO using NEP for disposal or potential future use.

Mars

Earth

Crew to LDHEO SLS-Crew

NEP Thrusting

Chemical Burn

LGA

Rendezvous with 

Pre-Deployed Descent Lander

Crew travel to 

destination

CPS/NEP 

performs MOI

Crew return from 

destination

Crew in Ascent Vehicle 

rendezvous with 

DST

30 crew 

days on 

surface

Crew return to Earth SLS-Crew

LGA to C3 = 2

NEP/Chem Stack 

Launch/mating: ~6 month

NEP-Chem vehicle Launch Reqt’s: 2 SLS, 5 super Heavy for Xe Interstages and chem stage top-off

Dispose of 

chem 

stage SLS : Chem Stage Super Heavy Xe SLS: NEP Element Interstages

Reactor 

Commission/Top off 

Chem Stage: ~2 months Starship Chem Tanker

~ 3 month

transfer to 

LDHEO Integrate Hab into 

NEP/Chem Stack

Dump 2 Xe 

Interstages Starship Chem Tanker

Chemical 

raise 

~14 month NEP 

spiral to NRHO

NEP to Mars 

(C3 ~ 2) NEP aligns 

for departure

DST 38 days 

in Mars Orbit

NEP to Earth 

(to C3 = 2)

NEP performs EOI

Venus Flyby

Reuse for further Mars missions?

LGA to 

LDHEO

C3 = -2 4 CLV-H

Deep Space Habitation

Shakedown stage & logistics

Figure 4.2-3. Mars 2039 NEP/Chem Piloted Mission ConOps.

The NEP/CCP reference concept is shown in isometric view in Figure 4.2-4 in the Earth departure configuration.

This reference design was the result of a COMPASS conceptual level design study, which was not exhaustive, 

and should not be considered optimal. More advanced configurations were identified and merit further study. 

Planform views of the aggregation orbit configuration, with additional Xenon Interstage Modules, and the earth 

return configuration are shown below it. The NEP Propulsion Module is on the left. A 50 m telescoping boom 

holds the nuclear reactor far from the radiator, electronics, and Deep Space Habitation. A radiation shield at the 

base of the reactor creates a radiation-free cone that envelopes all those vehicle components. The electric 

thrusters are positioned to propel the vehicle in the direction of the Deep-Space Habitat and Chemical Propulsion 

Module. The radiators are depicted blue, and the electronics and Power Management and Distribution (PMAD)

are within the box at the base of the radiators. The XIM is between the radiators and the DSH, containing several 

xenon propellant tanks. The companion LOX-CH4 Chemical Propulsion Module (CPM) is at the right end of the 

figure. The NEP system was sized to maximize the power system that could be launched in a single SLS 8.6 m 

diameter, long SLS fairing. The CPM was also sized to volumetrically fit in a second, single SLS 8.6 m diameter, 

long SLS fairing.
NASA 31

DST

Cargo

Mission 1

DST

Cargo

Mission 2

Piloted DST

at Aggregation

Piloted DST

at Earth Departure

Piloted DST

at Earth Return

Figure 4.2-4. NEP/CCP Reference Vehicle Concept.

Figure 4.2-4 also depicts how the NEP module and the Chemical Propulsion Module are used separately to 

propel the two cargo predeployment missions. The NEP module can propel all three landers to Mars on 

subsequent missions, when the first lander is not required to predeploy before the remaining two as a capability 

verification.
NASA 32

MTAS assumed Hall-effect thrusters for the NEP system, based on the development of high-efficiency, 12.5 kWe

Advanced Electric Propulsion System (AEPS) thrusters for the Gateway SEP Power and Propulsion Element. 

Additionally, higher power Hall thruster technology has been demonstrated by the NASA 457M, and a separate, 

nested channel design, which have been tested up to 100 kWe. Based on these technologies, a specific impulse

(Isp) of ~2600 seconds was assumed for estimating the propellant required for the NEP mission. This Isp is nearoptimum for the reference Mars trajectory. Other thruster technologies are possible in this size and Isp class and

should be considered in subsystem trades as identified in follow-on work at the end of this document. The Isp of 

the LOX-CH4 chemical propulsion system was assumed to be ~350 sec.

A major consideration in the design of a nuclear electric propulsion system is the size of the power conversion 

radiators, which define packaging and deployment requirements as well as the need for in-space assembly of 

large structures. Radiator size is determined by the size and efficiency of the thermal-to-electric power 

conversion, and the designed heat rejection temperature for the operating environments encountered. The 

radiators are a critical element of the power conversion cycle, and they are required for mission success. To 

minimize the risk of coolant leaks, the current best practice is to assemble, fill and check the radiators for leaks 

before launch, and avoid disassembly/reassembly in space. Thus, it is advantageous to package the radiators 

in a single launch vehicle with the reactor and power conversion, with a reliable and structurally sound means to 

deploy them on orbit. A concept was developed to fold a fully assembled radiator as large as 2500 m2to fit

within the volume of an SLS 8.6 m diameter, long cargo fairing, based on technology developed under Project 

Prometheus. A more advanced radiator folding concept was identified for about 4000 m2in the same volume. 

However, to minimize the packaging complexity, number of folds, and flexible fluid joints, 2500 m2 was assumed

as a constraint to size the MTAS NEP power system. As in-space assembly capabilities improve, more advanced 

systems requiring larger radiators can be considered.

The MTAS NEP reference concept assumed the use of a liquid-metal cooled reactor with Brayton power 

conversion, utilizing Supercritical-CO2 as a working fluid. The reactor concept reference is a HALEU TRISO fuel 

in a moderated core, based on the DOE Transformational Challenge Reactor (TCR) in development, or 

alternately an HEU UN pin fuel in a fast-neutron spectrum core based on the heritage SP-100 design. Other 

coolant, power conversion, and working fluid options can be considered in this mission time frame and are 

identified as subsystem trades for follow-on work. Assuming these fuel and power conversion technologies and 

the desire to minimize the use of refractory metals due to manufacturability and testing challenges, the reactor 

outlet temperature was designed to be 1200K permitting a power conversion hot-side temperature of 1150K with 

nickel-based superalloy construction. These design choices and constraints define the nuclear electric power 

system size to be about 2 MWe. With conservatism, the mission was designed with 1.8 MWe provided to the 

propulsion system, with an additional 0.1 MWe needed for the habitat and vehicle systems. With a chemical 

propulsion system providing the high thrust needed near high-gravity bodies such as the Earth, Moon, and Mars,

1.8 MWe propulsion power was found to be near-optimum for maximizing overall mission performance.

In comparison, an all-NEP Deep-Space Transport for the same 2-year piloted mission and ConOps would require 

5+ MWe for propulsion power. While such power levels and propulsion systems are feasible, such a system 

introduces additional challenges on technology development and system packaging. Most significantly, the 

power conversion radiators would be 2.5 times larger, requiring multiple launches due to volume packaging 

constraints, and have associated fluid connections that would have to be mated, sealed and verified as reliable 

on orbit. Such in-space assembly and check-out operations are considered feasible in the 2030s, but the study 

considered it lower risk to avoid them. The higher propulsion power is beyond the limits of current Hall thruster 

technology as well as existing test facilities, driving the need for more advanced thruster concepts and larger 

test facilities unlikely to support a 2030s mission. Thus, the hybrid NEP/CCP propulsion concept innovated 

under MTAS enables the consideration of a 2030s NEP mission with achievable technology.
NASA 33

4.2.1 NEP Technology

A preliminary MTAS NEP development plan is presented in Figure 4.2-5. It consists of three major phases: (1) 

technology maturation, (2) engineering development (DDT&E), and (3) flight system development. The first step 

is the development of detailed government reference designs of the NEP reactor power system and electric 

propulsion system to inform hardware procurement efforts with industry. While the MTAS concept studies provide 

a good starting point, they are not of sufficient fidelity to establish design requirements for industry RFPs. A 

Failure Modes and Effects Analysis (FMEA) would be a key element of the system engineering process that 

starts under the government reference design and is continued by the industry subsystem suppliers. The results 

would inform subsystem testing requirements and higher-level vehicle decisions being made by the system 

integrator. In lieu of the FMEA, the team referenced prior projects and used engineering judgement to determine

testing requirements for this study.

Figure 4.2-5. Preliminary NEP Power Development Plan.

The technology maturation phase would include three main elements: power and EP Technology Demonstration 

Units (TDUs), reactor fuel characterization, and PMAD/DDU parts assessment. The power and EP TDUs will 

focus on representative-scale breadboard subsystems that could be combined for integrated system tests to 

advance the TRL. A non-nuclear power system TDU, shown in Figure 4.2-6, would perform a 1000-hour 

performance test at GRC Vacuum Facility #6 (VF6), integrating:

• a 500 kWt reactor simulator, 

• 1200 K lithium primary loop, 

• 100 kWe Brayton unit, 

• 500K NaK heat rejection loop, 

• and two 40 m2radiator panels 

The Power TDU would include an external 1000 Vac PMAD assembly coupled to simulated 650 Vdc Hall thruster 

loads. Similarly, full-scale 100 kWe Hall thruster, 650 Vdc DDU and XFC breadboard test articles will be procured 

and tested at GRC Vacuum Facility #5 (VF5). In parallel, DOE would lead efforts to characterize and qualify the 

NEP reactor fuel and moderator through several test iterations. The DOE fuel development effort is envisioned 
NASA 34

to be like the Advanced Gas Reactor TRISO fuel qualification performed by INL, Oak Ridge National Laboratory 

(ORNL), and BWX Technologies with tasks for production, irradiation, post-irradiation examination, and 

modeling. The final element of the technology maturation phase is the electronic parts assessment for the PMAD 

and DDU. This activity will focus on a combination of radiation and thermal testing to evaluate high voltage 

electronics with an emphasis on wide band gap devices such as SiC or GaN. The technology maturation phase 

will conclude with the Mission PDR in year 4 at which time most of the power and EP technologies will have 

reached TRL 5. ROM cost estimates by the MTAS NEP team estimate the technology maturation effort to be 

about $350M over four years.

Figure 4.2-6. Non-nuclear Power TDU.

The engineering development phase of the NEP power system development plan starts with the mission SRR 

in year 3. The design of full-scale engineering model subsystems will occur in parallel and be informed by the 

TDU testing. The plan is to build multiple full-scale Engineering Models (EMs) that would be used to support 

various test objectives. A complete full-scale power string will be assembled with a high-fidelity reactor simulator, 

single Brayton unit, heat rejection loop, and PMAD power channel for an integrated 1000-hour non-nuclear 

system test, possibly at the Armstrong Test Facility (ATF), formerly known as Plum Brook Space Power Facility

(SPF). A second set of EM units will be subjected to launch vibration environments and placed on long-duration 

tests to evaluate performance relative to the mission design life, currently estimated at 27,100 hours for the 

power system and 22,860 hours for the EP system. Life verification will be accomplished through a combination 

of testing and analysis, similar to the approach being used for the AEPS thrusters planned for Gateway. A third 

Brayton unit will be supplied to the DOE reactor team for use in a full-scale 100-hour reactor prototype test to 

demonstrate nuclear performance. Candidate nuclear facilities to accommodate the reactor prototype test 

include the INL Experimental Breeder Reactor II (EBR-II) facility or the Nevada National Security Site, U1A 

facility. Similar testing will be performed on full-scale EM thrusters, DDUs, and XFCs culminating in a thruster 

performance verification test at GRC VF6. The three separate integrated tests of full-scale EM test articles (nonnuclear power string, reactor prototype, and EP string) combined with the subsystem launch vibration tests would 

provide critical data to support Mission CDR and establish TRL6 for the key NEP elements by year 6.

The presumed use of GRC and ATF vacuum facilities for the non-nuclear power and EP testing is based on the 

need to provide a large volume and high pumping speed. There are other vacuum facilities at NASA (e.g., JSC

Chamber A), Department of Defense (e.g., Arnold Engineering Development Complex), industry, and academia, 

but they are smaller and lack the pumping speed needed for high power EP thrusters. Current facilities prohibit
NASA 35

simultaneous testing of multiple 100 kW thrusters, but there is no precedent to do so. Tests performed of multiple 

thrusters at lower power levels indicate negligible thruster interactions when appropriate separation distance is 

maintained. An upgraded VF6 could be utilized to test full-scale EP EMs. Analysis shows that pumping speeds 

of 1600-1800 kL/s xenon could be achieved corresponding to a background pressure of 1.5E-5 torr for a 2600s 

Isp, 60% efficiency 100 kW thruster. Alternatively, ATF could be used, but vacuum facilities there would require 

upgrades that are more extensive than VF6.

One of the primary MTAS ground rules was to deliver crew to Mars by 2039, which drove the technology 

decisions (and thus the facilities) selected for the mission concept. If NASA postpones the Mars missions to the 

late 2040s or 2050s, lower TRL technology and construction of new facilities could be pursued.

The flight system development phase will include design, fabrication, qualification, and acceptance testing of 

flight components. A full assembly of subsystem qualification hardware elements will be built and tested. The

first complete set of flight components would be used for an NEP demonstration mission, possibly an all-NEP 

Mars cargo delivery mission that would precede the crew mission. That flight mission would validate the overall 

NEP propulsion stage performance in space before a duplicate version would be used on the crew vehicle. The 

cargo mission is fundamental to demonstrating in-space performance, life and reliability of the nuclear power 

and propulsion element before it is used to transport crew. The NEP demo vehicle would be assembled, and 

checkout testing of the power and EP elements would be performed using a reactor simulator. After checkout, 

the flight reactor would be integrated, and a zero-power critical (ZPC) test would be performed to validate reactor 

neutronic performance before the demo vehicle launch in year 11. The second set of flight components will be 

fabricated and tested in parallel with the demo mission and readied for launch with the crew vehicle in year 13.

Assuming a 2039 launch of the crew vehicle and working backward using the 12-year development life-cycle 

results in a required project start in 2027. While the proposed NEP development schedule in Figure 4.2-5 is 

achievable, it is very aggressive and success-oriented with little margin for setbacks. If 2035 remains the target 

for the first human Mars mission, the required project start would be in 2023.

4.2.1.1 Chemical Propulsion Technology

Technology maturation is also required for the CCP, demonstrating of Cryogenic Fluid Management (CFM) 

technologies necessary for long-duration liquid oxygen and liquid methane cryogenic propellant storage and 

transfer. Storing large quantities of cryogenic propellants with negligible losses has always been a challenge. 

While storing LOX and LCH4 is easier than LH2, this capability and the CFM technologies necessary for multiyear missions have never been demonstrated in space. Oxygen and methane can readily leak through small 

gaps in valve seats and seals at other interfaces. Further, to prevent boil-off losses, the propellants must be 

maintained at extremely cold temperature (90 to 100 K or -297 to -280 °F).” Zero boil-off (ZBO) of LOX and 

LCH4 and very low leakage during the two-year mission has been assumed. Achieving this will require the 

development of no/low leakage valves and couplers, advanced insulation, and high-capacity, high-efficiency 

cryocoolers. Finally, the NEP/CCP concept leverages refueling of the cryogenic propellant tanks in orbit. 

Managing the cryogenic propellant during microgravity coasts greater than a few hours and transferring 

propellant from one tank to another in space has never been demonstrated and represents additional technology 

maturation challenges. As with NTP, these technology challenges are shared with the Artemis missions and 

other commercial space transportation developments, but the risks for developing this low maturity technology

remain.

4.2.2 NEP Campaign

For the NEP/CCP Hybrid propulsion option, the mission architecture has predominantly four unique 

transportation system elements (in bold, below) to support the piloted mission, and requires 13 total launch 

elements to support the roundtrip piloted mission, including 7 that are unique to the NEP/CCP campaign. These 

elements are:
NASA 36

• Space Launch System launched Nuclear Electric Propulsion Module (1)

• Space Launch System launched Chemical Propulsion Module (1)

• Super-Heavy CLV launched Xenon Interstage Modules (3)

• Super-Heavy CLV launched LOX-LCH4 propellant and transfer capability (2)

Common piloted launch elements:

• Space Launch System Orion launches for crew delivery and return (2)

• Heavy CLV launched Deep Space Habitation (2)

• Heavy CLV launched Chemical Propulsion Stage for DSH shakedown cruise (1)

• Heavy CLV launched logistics for Deep Space Habitation (1)

The Deep Space Habitation is potentially part of the initial Artemis & Gateway program and is launched in the 

early 2030s to support various lunar activities and Mars analogue missions. The Mars mission campaign includes 

two H-CLV launches to deliver and outfit the DSH in NRHO where it will be checked-out at Gateway. Then a 

third H-CLV launch provides a chemical propulsion stage to perform a DSH shakedown cruise. Finally, an HCLV delivers logistics to the DSH prior to Mars departure.

The Nuclear Electric Propulsion Module and the Chemical Propulsion Module are launched on a Block 2 variant 

of the Space Launch Systems, while the Xenon Interstage Modules and the LOX-CH4 propellant refueling 

vehicles are launched on super heavy classed commercial launch vehicles. The remaining elements are 

launched starting in the mid-2030s to support the piloted mission to Mars in 2039. 

The launch campaigns for NEP/CCP predeployment and piloted missions are depicted in Figures 4.2-2 and 

4.2-3. Both figures show the launch of elements to a 500 km circular Low Earth Orbit (LEO) aggregation orbit. 

This orbit was chosen for the NEP/CCP option to maximize the launch vehicle payload delivery, where shroud 

volume for propellant is not a consideration. The five hardware elements and two refueling flights must be 

launched and aggregated in a relatively short period of time, as the thermal and orbital debris environment in 

low Earth orbit can pose challenges to the systems. After the elements have been aggregated, the integrated 

stack performs orbit raising maneuvers using the Chemical Propulsion Module to a higher Earth orbit of 1,100 

km altitude, which is considered nuclear safe, with consideration for orbital debris and the Van Allen radiation 

belts. From here, the nuclear reactor is deployed, and the stack performs orbit raising maneuvers using the 

electric propulsion system over an 18-month period to reach the final aggregation orbit at NRHO in cislunar 

space, where the propulsion system rendezvous with the Deep Space Habitation in preparation for the piloted 

mission. Finally, approximately six months prior to crew departure for Mars, the integrated deep space transport 

performs a low energy transfer from its aggregation orbit in cislunar space to a lunar-distance high Earth orbit of 

400 km x 400,000 km altitude to rendezvous with the crew and prepare for the piloted mission to Mars. 

The optimization of the aggregation orbit for NEP crew and cargo vehicles was the major campaign improvement 

between Iteration 1 and Iteration 2. Iteration 1 considered NRHO as the aggregation point for both options, 

significantly reducing the payload delivery capability of all launch vehicle options, while placing the Mars 

propulsion systems in a much higher energy orbit, reducing their necessary ∆V. By optimizing the aggregation 

orbit vs NRHO, the overall launch campaign for NEP/CCP was reduced from 41 launches to 30.

To delay the first need date of the nuclear system, the predeployment campaign concept of operations for the 

NEP/CCP option utilizes a combination of the two primary propulsion elements to deliver the three lander

systems. The first lander is delivered using the Chemical Propulsion Module identical to the piloted mission 

during the 2035 Mars conjunction mission opportunity. Landers 2 and 3 are delivered to Mars using a Nuclear 

Electric Propulsion Module and a single Xenon Interstage Module during the 2037 Mars conjunction mission 

opportunity. To support this cargo campaign, a total of 17 elements are required, including 14 that are unique to 

NEP/CCP:
NASA 37

• Space Launch System launched Nuclear Electric Propulsion Module (1)

• Super-Heavy CLV launched Xenon Interstage Modules (1)

• Space Launch System launched Chemical Propulsion Module (1)

• Heavy CLV launched Chemical Propulsion Boost Stage (7)

• Heavy CLV launched LOX-LCH4 propellant (4)

Common cargo launch elements:

• Space Launch System launched Mars Landers (3)

The ~65,000 kg Mars Lander systems are launched on Block 2 variants of the Space Launch System. They are 

meant to aggregate in cislunar space, at NRHO, but because the SLS launch capability is limited to ~45,000 kg

to translunar injection condition, they are delivered to an elliptical, Medium Earth Orbit. Additional chemical 

“boost stages” are required to deliver the landers to the NRHO aggregation orbit. For the first lander Mars 

transfer, the Chemical Propulsion Module, launching on a Block 2 variant of the SLS to NRHO, is only launched

partially fueled due to the payload limitation of the SLS. Thus, additional H-CLV LOX-LCH4 propellant flights are 

required to supply the CPM enough propellant to deliver the lander to Mars.

The NEP Module and XIM are delivered to a 500 km circular LEO by a Block 2 SLS and SH-CLV, respectively.

If placed in a 500km orbit, astronauts could help with assembly and checkout, which has proven very useful for 

space station assembly and checkout. To reduce the time needed to deploy the NEP module, a chemical 

propulsion boost stage is delivered by an H-CLV to rendezvous with the NEP cargo transfer vehicle and is

employed to take the vehicle to the 1100 km circular, nuclear safe orbit. From there the NEP Module is employed 

to spiral to a rendezvous with Landers 2 and 3. In all, seven H-CLV launched, chemical propulsion boost stages 

are required to deliver all three lander systems to cislunar space for aggregation, two per lander and one for the 

NEP module. 

Overall, to support the complete crew mission to the Mars surface, the Nuclear Electric/Chemical Hybrid 

Propulsion system option requires a total of 30 launch vehicles for the campaign, 21 of which are unique to 

NEP/CCP. As an example of how the high Isp of NEP reduces the variance between launch opportunities, the 

same number of launches would be required for a 2035 piloted mission. For 2039, these launches are spread 

out between 2030 and 2041 (including the crew return) and are shown in Figure 4.2-7. A total of nine Space 

Launch Systems are required; two will be the crew variant and seven will be the block 2 cargo variant. Fifteen 

heavy class commercial launch vehicles serve mainly as support vehicles, delivering propellant and boosting 

stages to their final aggregation orbits. Finally, six super-heavy class commercial launch vehicles are required 

to provide the propellant requirements in the Xenon Interstage Modules, and the LOX-LCH4 propellant refueling 

operation for the piloted vehicle during its initial low Earth orbit aggregation. 
NASA 38

Figure. 4.2-7. NEP/Chem Cargo Vehicles and Piloted Vehicle Launch Campaign.

4 . 3 C O M P A R I S O N B E T W E E N P R O P U L S I O N A L T E R N A T I V E S

MTAS Phase 1 concluded that both NTP and NEP/CCP propulsion system concepts have the potential to fulfill 

the mission requirements for fast Mars transits in the late 2030s, including enabling two-year roundtrip mission

durations. But these options use different approaches with unique challenges. While a 2035 mission is possible, 

it is programmatically high risk, and would require substantial investment within two years of the date of this 

study. The following section provides a discussion of the distinctions between the propulsion approaches.

4.3.1 Performance

Figure 4.3-1 provides a summary of the second iteration (1.2) NTP and NEP/CCP performance assessment and 

vehicles designed for the objective two-year piloted mission to Mars in the late 2030s. These are considered 

reference cases resulting from trades for technology, systems size, mission performance, and operational 

approach and define what performance is possible for the mission ground rules and ConOps, but do not define 

optimality and are open to further trades. In both cases, the cargo predeployment campaign was performed with 

the propulsion elements defined by the piloted mission architecture. The NTP reference mission design is for a 

round-trip time away from Earth orbit of 690 days, assuming an engine-out before Mars departure. The trip time 

below two years is enabled by a Venus gravity assist on the return trajectory in 2039. A 760-day round-trip 

mission comparable to the NEP/CCP reference case would save propellant mass, but the VGA would no longer 

be beneficial in 2039. The mission requires four, 25,000 lbf NTR engines with ~900 sec Isp, each with ~4-hour 

full-thrust life. The four NTR engines are needed to provide sufficient thrust to leave the aggregation orbit with 

acceptable gravity losses. By comparison, the NEP reference mission design is for a round-trip time away from 

Earth orbit of 760 days to minimize the total number of launches and accommodates partial power loss on the 

Earth return leg. A shorter NEP/CCP mission would require a larger LOX-CH4 stage. The NEP/CCP mission 

requires: one, 1.9 MWe reactor with twenty, 100 kWe Hall thrusters operating at 2600 sec Isp; and three, 25,000 

lbf LOX-CH4 engines for high thrust. The difference in round-trip mission time for the two reference cases is an 

example of how each propulsion option optimizes differently around the two-year mission objective; in this case 

the timing of a Venus gravity assist for the 2039 impulsive trajectory provides an advantage. Both the NTP and 

NEP systems are capable of mission times significantly less than two years at the expense of increased 

propellant mass, and/or larger propulsion systems, and/or more advanced technology. However, the two-year 
NASA 39

round-trip objective is a reasonable target, resulting in performance requirements for Nuclear Thermal Propulsion

and Nuclear Electric Propulsion (when augmented with Chemical Propulsion) that are achievable by the late 

2030s.

NTP for 2039 NEP/Chem for 2039

Vehicles in Low 

Earth Aggregation 

Orbit

(Habitat added later)

Primary 

Technologies

Ø Nuclear Thermal Rocket (NTR): 900s Isp / 25klbf Ø Zero Boil-off LH2 storage & transfer

Ø High capacity 90K & 20K cryocoolers

Ø Cryogenic Feed System: 

Ø Low-leak 8” and 3” cryogenic valves

Ø Interstage 7” cryo couplers

Ø Nuclear Power: 1.9 MWe HEU or HALEU Reactor, SCO2 Brayton 

Convertors, Deployable modular radiators

Ø Electric Propulsion:100 kWe Xenon Hall thrusters @ 2600s, 13t throughput

Ø Large High pressure Xe storage Tanks

Ø Zero Boil-off LOX/LCH4 storage & transfer: high capacity 100K cryocooler

Ø Chemical Propulsion: LOX/LCH4 derived from Mars lander ascent engine

2039 Crew Mission Ø 690 round trip, 30 Sols on surface Ø 760 round trip, 30 Sols on surface

Number of reactors &

engines

Ø Four 25 klbf NTR: 4 hour life, 2700K, 530 MWth,

HALEU reactor/engines

Ø One reactor: 2 year life, 1200K, 10 MWth/1.9 MWe HEU or LEU

Ø Twenty Xe Hall thrusters: 100 kWe Ø Three 25 klbf LOX/LCH4 engines

Unique Elements

Ø NTR Core Stage (1) SLS

Ø Hydrogen Inline Stage (1) SLS

Ø NTR Stub Core Stage (2) H-CLV

Ø Multi-Dock Truss Element (2) H-CLV

Ø Power and Propulsion Module (1) H-CLV

Ø Hydrogen Drop Stage (18) SH-CLV

Ø Additional launch for DSH transfer stage (1) H-CLV

Ø Nuclear Electric Propulsion Module (1) SLS

Ø Chemical Propulsion Module (1) SLS

Ø Xenon Interstage Modules (3) SH-CLV

Ø LOX-LCH4 propellant & transfer capability (2) SH-CLV

Required Launch 

Fleet

Ø 2 SLS, 18 SH-CLV, 5 H-CLV

Ø 934 MT in 1200 x 7000 km Aggregation Orbit

Ø ~850 MT of LH2 propellant

Ø 2 SLS, 5 SH-CLV (alternately, H-CLV could be employed)

Ø 678 MT in 500 km Circular Aggregation Orbit

Ø ~256 MT Xe EP propellant, ~ 252 MT LOX-LCH4 propellant

Figure 4.3-1. Top Level Summary of NTP and NEP/Chem Approaches to Mars 2039.

MTAS performed a comparative analysis of these two reference cases between alternative propulsion 

technologies, considering distinguishing characteristics, development, operations, programmatics, and 

performance metrics. Details on these assessments are summarized in the following subsections.

4.3.2 Development

4.3.2.1 Technology-NTP: 

Nuclear Thermal Rocket (NTR) engines had achieved a high degree of maturity from significant past investment 

in both the U.S. and Russia. From 1955-1973 Project Rover focused on NTR technology, while in 1961 the 

companion Nuclear Engine for Rocket Vehicle Applications (NERVA) contract (awarded to the AerojetWestinghouse team) focused on the development of a flight engine. The combined Rover/NERVA efforts

developed several engine designs using HEU in UO2, UC graphite composite (UC-ZrC)C, and carbide (U, Zr)C 

fuel forms, in channeled prismatic fuel elements, to high technical maturity; with a thorough test program solving 

many of the technical issues. The NERVA cores operated in the epithermal neutron spectrum, with graphite as 

a moderator. Project Rover tested smaller cores (25 klbf vs 75 klbf) using a Zirconium Hydride (ZrH) moderator 

in the cooled structural tie tubes to achieve criticality. By the end of Rover/NERVA, the XE’ engine was a 

successful test of a flight configuration and had been certified by the Space Nuclear Propulsion Office to meet 

the requirements for a human mission to Mars as defined at the time. The XE’ achieved greater than 2400K 

average propellant exhaust temperature, sufficient for about 840 sec Isp, and demonstrated successful start-up 
NASA 40

and shut-down transients. However, NASA’s Mars mission plans never materialized, and in 1973 Project Rover 

and the NERVA contract were cancelled. From 1987 to 1994, the U.S. Air Force was developing an ultra-high 

performance NTR concept under the Space Nuclear Thermal Propulsion (SNTP) Program. The SNTP NTR 

design was to utilize HEU UC2 fuel particles in a cooled block moderated core to achieve a very high thrust-toweight ratio (T/W) of ~30. Comparatively, in the 1990s, Russia developed HEU (U,Zr,Ta)C fuel in a bundled 

twisted ribbon form to be operated in a cooled moderator block. These later programs achieved some design, 

modeling, and component testing, but did not complete reactors for testing. Much of the expertise developed 

during these programs has retired and is at risk of being lost.

For Nuclear Thermal Propulsion, a Nuclear Thermal Rocket engine must be integrated to a stage, which provides 

propellant to the system, along with spacecraft systems for guidance, navigation, and control. Briefly during 

Rover/NERVA, from 1961-1963, the Reactor In-Flight Test (RIFT) project focused on developing the additional 

technologies for systems necessary for the operation of an NTP stage. This project was cancelled prior to much 

development in these areas.

For NTP the main factor of performance is specific impulse, which is primarily determined by the achievable fuel 

temperature, which drives the temperature of the propellant. Low molecular weight propellant is key. A second 

factor of performance is thrust-to-weight ratio of the propulsion system – which affects vehicle acceleration and 

achievable single mission ∆V. Additionally, reactor full-power lifetime also affects the achievable ∆V over the 

mission life of the system. For the NTP reference case, the required Isp is 900 sec, while the assessed thrust-toweight ratio for the engines is 3.6, and the assumed full-thrust life of the reactor is 4 hours. The NTP technology 

plan is to meet these performance numbers, while performance significantly beyond this has yet to be 

demonstrated.

NASA is currently investing in Nuclear Thermal Propulsion system component technologies. Nuclear Thermal 

Rocket (NTR) engine subsystems, such as: engine gimbals; regeneratively cooled exhaust nozzles; and 

cryogenic hydrogen turbopumps, valves, and flexible fluid lines, benefit from strong synergies with the SSME 

and ongoing liquid chemical combustion engine development. Chemical rocket components will need to be 

developed for operation in a high radiation field for application to an NTR engine. NASA has near-term plans for 

characterization of novel fuel concepts at DOE Idaho National Lab (INL). NTP is also receiving some initial

investment from DOD/DARPA due to the potential for high thrust acceleration and rapid response capability, and 

this provides agility in the Earth and Moon gravitational environments. Some of these technologies could be 

adapted to NASA use for Mars transportation. Both current NTP efforts are focused on HALEU reactor designs,

which are a departure from the HEU design heritage established in the 1960s – 1990s. Progress-to-date 

represents a small fraction of development needed for flight; thus, significant development risk remains.

4.3.2.2 Technology-NEP: 

NEP was demonstrated on-orbit in 1965 when the SNAP-10A reactor powered an early cesium ion thruster, 

though there was no follow-on mission application. SNAP-10A used U-Zr-H fuel and NaK coolant with 

thermoelectric power conversion to produce ~500 We. Russia has flown over 30 nuclear reactors in space, the 

most recent design being the Topaz I reactor, which first flew in 1987 with an experimental Hall thruster. They 

also developed the Topaz II reactor, which was tested and exhibited but never flown. The Topaz reactors use 

HEU UO2 fuel with ZrH moderation, and in-core thermionic power conversion. They were short-lived reactors. 

The strongest U.S. design heritage for fission power in space was developed from the 1983-1994 SP-100 

program, which was a DOD/DOE/NASA tri-agency program developing a 100 kWe space nuclear reactor power 

system. NEP was one of several NASA mission applications envisioned. SP-100 developed HEU UN pin fuel 

and liquid lithium cooling technologies with strong feed forward to current NEP requirements. SP-100 achieved 

significant design development and component testing but did not complete a reactor system for testing.

More recently, the 2002-2005 Project Prometheus advanced NEP development for planetary science missions.

Project Prometheus was developing 100-200 kWe NEP for the Jupiter Icy Moons Orbiter (JIMO) mission in 2002-
NASA 41

2005 for flight in 2016. Most of the early JIMO government studies considered SP-100 derivative concepts with 

HEU UN fuel pins, lithium cooling, and HeXe Brayton conversion. At the conclusion of the project after the Naval 

Reactor Prime Contracting Team conducted a reactor down-select study, the Prometheus Baseline-1 design 

assumed a gas-cooled HEU UO2 pin fueled core and direct HeXe Brayton cycle power conversion system. JIMO 

was cancelled before any significant component testing was accomplished. There have also been several

relevant efforts to study and/or develop NEP technologies including the 1990s Space Exploration Initiative, the 

2014 Mars Design Reference Architecture 5.0 (Addendum 2), and the Constellation-era Fission Surface Power 

project, which developed and tested a 12 kWe Stirling TDU test in GRC VF6. 

In 2018, the STMD Kilopower project successfully demonstrated the inherent control of an HEU U-Mo solid fuel 

core using negative temperature reactivity feedback. The Kilopower design transferred heat to Stirling convertors 

using heat pipes in contact with the core perimeter and was designed to produce ~1 kWe. The ground test 

demonstrated load-following and stable thermal control over a number of power system operations and 

contingencies. NASA has explored derivatives of the Kilopower concept for a 10 kWe surface power reactor, 

and combined with low-power electric thrusters, for NEP to support outer planet robotic science missions.

For Nuclear Electric Propulsion, a space nuclear reactor is combined with thermal-to-electric power conversion, 

thermal heat rejection, and power management & distribution to deliver regulated electric power to the electric 

propulsion subsystem. The EP subsystem includes power processing equipment to convert and distribute the 

input power among the EP loads and a propellant feed system for the storage, management, and supply of 

propellant. Thus, NEP requires complex integration of all these distinct elements. Past development of the 

subsystem technologies for NEP has been at a smaller scale than required for the 1.9 MWe MTAS NEP reference 

case. For instance, NASA GRC developed a 25 kWe Brayton converter with HeXe working fluid, heat 

exchangers, and heat rejection in the 1980s for the Space Station Freedom solar dynamic power module and 

demonstrated that technology during the 1990s 2 kWe Solar Dynamic Ground Test Demonstration. Also, electric 

propulsion, both Hall-Effect Thrusters and Gridded-Ion Thrusters, have been used in space since the 1990s at 

the ~1 to ~10 kWe scale. During Project Prometheus, the 2 kWe Brayton unit from SD GTD was connected to a 

2 kWe NSTAR Ion thruster to demonstrate power delivery and stable electric control under thruster transient 

operations.

A major system challenge in developing an NEP power system is the size, mass, packaging, and deployment of 

large space radiators to reject the waste heat generated by a thermal-to-electric power conversion. Important 

advancement of large, high-temperature composite radiator technology occurred during the Fission Surface 

Power project that supported NASA’s Constellation Program after Project Prometheus. Also as important is the 

radiator system developed and deployed on the International Space Station. Although the ISS External Active 

Thermal Control System is operated at a lower heat rejection temperature and uses different materials and fluids 

than required for NEP, the total size (approximately 1200 m2) is comparable to the size required for the MTAS 

NEP reference design – demonstrating that deploying and operating radiators of this size is feasible. Important 

engineering remains in the development of flexible fluid lines and radiator coolant manifolds that are suited to 

the NEP heat rejection temperatures.

The main factor of performance for Nuclear Electric Propulsion is the mass-to-power ratio of the reactor power 

system and electric propulsion. This is often measured in specific mass (kg/kWe) or specific power (Watts/kg) for 

smaller systems. This performance factor translates in effect to the thrust acceleration potential of the NEP 

system, where lower specific mass results in higher acceleration potential, and thereby faster missions or larger 

payloads. For the NEP/CCP reference case the specific mass of the reactor power system alone was about 12 

kg/kWe, and the EP subsystem (thrusters, DDUs, and XFCs) adds about 2.5 kg/kWe. The overall mass of the 

dry propulsion system (including the power system, structure, stage subsystems, thrusters, and propellant 

tankage mass) resulted in a specific mass of ~40 kg/kWe.

NASA currently does not have a coordinated development activity for an NEP system but is independently 

investing in several relevant subsystem technologies. To begin with, the Artemis Gateway Power and Propulsion 
NASA 42

Element project is planning a 50 kWe Solar Electric Propulsion system with multiple 12 and 6 kWe Hall thrusters, 

and this technology can be leveraged for larger scale devices to suit Nuclear Electric Propulsion. A 50 kWe class 

Hall thruster, the NASA 457M, was fabricated and tested up to 100 kWe at GRC as a building block for 300 kWe 

SEP missions. Under the NextSTEP High Power EP project, HEOMD awarded contracts to EP developers to 

design 100 kWe thrusters and test them for 100 continuous hours. The necessary NEP power management and 

distribution technology could significantly benefit from current ARMD PMAD development for MWe electric 

aircraft.

Importantly, there is current synergistic investment in space fission power relevant technologies for terrestrial 

applications, including microreactors and power conversion at the scale required for Mars NEP. Relevant 

terrestrial microreactor technology is being funded by DOE under the Advanced Reactor Demonstration Program 

and by DOD under the Pele project. DOE, DOD, NASA, and industry have also joined forces to establish

production capabilities for TRISO fuel that could be used for NEP reactors. Further, DOE has been investing in 

Brayton power conversion in the MWe class using supercritical CO2 as a working fluid under the Solar Energy 

Technologies Office. Further study and development are required to adapt this technology for space. The NEP 

concept’s compact nuclear power generation capability has strong interest from terrestrial energy and can 

strongly benefit from synergistic component development.

Finally, in consideration of the companion conventional chemical propulsion technologies required in the hybrid 

NEP/CCP approach, many programs, both government and commercial, are developing LOX/LCH4 propulsion 

systems for both launch vehicles and in-space propulsion. Leveraging opportunities exist for the development 

of engines and stage technologies, as well as the important cryogenic fluid management of LOX and liquid 

methane.

4.3.2.3 Reactor – NTP: 

NTP reactors operate on the extremes of material capabilities and engineering design. This is necessary to 

achieve the highest reactor temperature to maximally add energy to, and accelerate propellant to high exhaust 

velocity, which translates to high specific impulse (propellant efficiency). Past NTR reactors designed for high Isp

operate close to the melting point of the fuel (within ~100-200 K). Since hydrogen is not only the propellant, but 

the reactor coolant, any significant imbalance in H2 flow could lead to hot spots and melted fuel. Furthermore, 

NTP reactors have been the highest power density nuclear reactor cores ever designed in order to provide thrust 

at minimal mass, which translates to high thrust-to-weight ratio. However, high power density translates to 

quicker reactor response to transients, and higher risk to melting the core.

The SNTP NTR design used HEU UC2 fuel particles, with high surface-to-volume for high heat transfer, in a 

moderated core to minimize size and mass, to achieve a very high thrust-to-weight of ~30. However, most past 

Nuclear Thermal Rocket (NTR) engine designs achieved thrust-to-weight of ~3; over 10-times lower thrust-toweight than conventional chemical rocket engines, which negates some of the advantages of higher Isp.

Additionally, the use of hydrogen as a propellant introduces engineering complexities to the design and operation 

of a fission reactor and requires rapid and precise reactor control. This is due to the neutronic interaction of 

hydrogen atoms in the balance of a critical nuclear reaction (hydrogen slows neutrons when they collide, 

thermalizing them and increasing their fission potential with uranium-235 – more hydrogen leads to more heat, 

but also more cooling). This interaction requires precise control to balance hydrogen flow and criticality control 

during reactor startup, when the core is cold and dense, hydrogen flow is introduced; and during transition to full 

thrust, as hot hydrogen becomes less dense, while the hot core requires more hydrogen cooling to keep from 

melting. Similar control complexity occurs during the shut-down transient when hydrogen flow is still required 

for cooling the reactor while not sustaining the nuclear reaction.

Finally, high temperature hydrogen corrosion of fuel elements and coatings in the core is a significant 

consideration for materials development and corrosion rate increases with fuel temperature. This issue typically 

limits engine lifetime to a few hours of thrust operation. Because of this, higher Isp NTR designs are driven to 
NASA 43

shorter-life, and longer-lived NTR designs are driven to lower Isp. Longer operation is also possible with fuel 

choice, and with hydrogen resistant material coatings, but the technology for full-thrust life greater than a few 

hours at high Isp has yet to be demonstrated. Thus, Isp, thrust-to-weight ratio, and full thrust lifetime are important 

considerations driving NTR reactor technology to extremes.

The consideration of high-assay, low-enriched uranium (HALEU - uranium enriched to 5-20% 235U, and typically 

19.75% for space applications) for space nuclear power and propulsion is very recent, resulting from the policy 

push to use such fuel where possible to promote nonproliferation of weapons-grade uranium. Uranium-235 is 

the only naturally occurring isotope in any practical abundance that is fissile with thermal neutrons. However, 

the neutrons generated by each 235U fission are born fast and must be slowed down (thermalized) to efficiently 

sustain the chain reaction with HALEU. If the neutrons are not thermalized, a HALEU core (even at 19.75% 

enrichment) must be large and heavy to contain enough 235U for a critical mass. Typical terrestrial nuclear 

reactors used for power are thermal neutron spectrum cores using LEU (some like the Canadian CANDU reactor 

even use natural uranium at 0.7% 235U), but size and mass are not key design limitations. For space applications, 

size, mass, complexity, lifetime, and reliability are all major design considerations.

The reference NASA NTP fuel development plan focuses on achieving the necessary performance from HALEU.

Current concepts are evolved from heritage 1960s HEU NERVA and 1990s HEU SNTP fuel designs. The use 

of HALEU requires a new fuel, and a thermal spectrum core design to keep the fuel mass and core size 

comparable to HEU designs. A thermal spectrum core requires moderator material (usually a high temperature 

hydrogen compound, or less effectively, graphite to slow fast fission neutrons to the thermal spectrum) and 

additionally requires cooling to maintain the moderator integrity over the operational life of the core. Hydrogen 

compound moderator materials, such as ZrH and Yttrium Hydride (YH), must be maintained at temperatures 

much cooler (~1000 K or less) than the high temperature fuel (~2700 K) for the hours of thrust operation. This 

introduces design complexity & challenges with integration and cooling of moderator materials. Furthermore, 

high power density requires this cooling to be precise and responsive. These materials are also low TRL and

require significant material development. The NERVA cores used graphite in the fuel elements to moderate the 

neutrons to an epithermal spectrum, and Project Rover experimented with ZrH material in cooled structural tietubes. The MTAS reference NTR design, assuming an HALEU core, requires full moderation to a thermal neutron

spectrum, and thus uses a block moderator design with coolant channels in high temperature areas similar to 

the 1990s Russian design. ZrH moderator is under development at Los Alamos National Lab and remains low 

TRL.

The heritage knowledge base for space nuclear power and propulsion, both NTP and NEP, is fast-spectrum 

reactors using HEU (uranium enriched to greater than 20% 235U and typically around 93%), which has been 

developed since the late 1950s. This includes for NTP the fuel and reactor development under Rover/NERVA 

and SNTP, and for NEP includes significant development of UN fuel under SP-100. While much of this legacy 

technology will require efforts to redevelop, the successes of the past represent feasibility proofs, while the 

promises of new technology have yet to be proven.

HEU reactors typically operate in the fast neutron spectrum, rather than the thermal neutron spectrum, to 

maintain simplicity and reliability, and facilitate long life by avoiding the use of moderator materials. Fast spectrum 

neutrons can sustain a fission reaction, albeit with less efficiency than thermal neutrons, and this inefficiency can 

be afforded if 235U is abundant, as in HEU. A thermal spectrum, HEU reactor has the potential to be the lowest 

mass solution for space nuclear power and propulsion, but it comes at the expense of design complexity and 

likely lifetime. The engineering of lightweight, compact, long-life, reliable reactors utilizing moderators, especially 

hydrogenated moderator materials, will require significantly more schedule and cost for technology development. 

Until the challenges inherent in this engineering and development are solved, the implementation of moderated, 

HALEU reactors represents significant risk for a 2030s nuclear propulsion system.
NASA 44

4.3.2.4 Reactor-NEP:

NEP reactors operate at lower temperatures than NTP reactors, but for longer duration. Overall, the NEP reactor 

is projected to be less difficult to develop given the lower operating temperatures and steady-state operation,

which is more analogous to terrestrial reactors. Higher temperature reactors are desirable to increase the 

efficiency of power conversion systems, as well as increase the heat rejection temperature to minimize radiator 

size. The operating temperature of space power reactors are typically limited by the material limits of the fuel, 

pressure vessel, and power conversion hot-side (e.g., the turbine for Brayton or Rankine systems). These limits 

are projected to be ~1200K for the use of super-alloy power conversion, ~1500K for refractory alloys, and 

~1800K for ceramic materials.

NEP cores typically are designed to operate behind a conic frustrum “shadow” shield to mitigate gamma and 

neutron emissions from the operating reactor and protect spacecraft systems and crew. Shields are dense and 

heavy, typically larger and heavier than the reactor itself, and are sized by the shape of the reactor. Because of 

this, compact cores are desirable, which has implications on reactor heat transport (which drives the volume of 

coolant channels in the core), as well as the choice between HEU and HALEU (HALEU requires more volume 

for both fuel and moderators). The use of HEU in a fast neutron spectrum core design results in the simplest, 

most compact, lightest space power reactor, especially for smaller systems up to a few MWe.

MTAS found that designing for much lower fuel density HALEU system requires employing hydride moderators

for a thermal spectrum core. This holds similar design and technology development challenges with HALEU NTP 

and introduces design complexity and development risk. A moderated space power reactor core must maintain 

hydrogen compound moderator materials, such as ZrH or YH, at less than ~1000K for years of operation while 

the fuel in the core is typically hotter (~1200K for the NEP reference design). Time at high temperature 

determines how long a hydrogen compound moderator material will maintain effectiveness, as hydrogen 

dissociates and escapes over time (an effect that increases with temperature); thus, the use of hydrogen 

compound moderators will limit reactor lifetime. Reactor temperature, core sizing factors, and full power lifetime

are important considerations driving NEP reactor technology and favoring the use of HEU. The only reasons to 

consider HALEU for space power applications are to alleviate policy concerns with the use of HEU and to address 

issues with security and licensing for the handling of HEU. However, robust public-private partnerships for the 

development of HEU space nuclear power and propulsion are possible, leveraging existing facilities, 

organizations, and entities licensed for handling HEU with adequate security. Furthermore, new facilities, 

organizations, and entities will have to be licensed to handle HALEU (as discussed in 4.3.2.3) and could be 

developed to the higher standards of HEU if the business case warrants it. Until technology challenges with 

HALEU fuel, and moderated space reactor cores are solved, such a space fission power approach remains a

high cost, schedule, and technical risk, while the performance required for human Mars propulsion may drive a 

return to HEU solutions.

The reference NEP plan builds on the SP-100 heritage for liquid metal reactor cooling, and on anticipated

development of HALEU nuclear fuels and moderators under the NASA Fission Surface Power project. The 

reference NEP approach departs from the SP-100 fuel heritage but considers the leveraging of advanced 

terrestrial nuclear development of HALEU fuels and moderators. NEP has higher heritage with HEU, and 

alternately could employ HEU UN fuel from SP-100 to reduce development risk, with a performance advantage

due to a smaller, lighter core and shield. In all MTAS NEP cases, a 1200K reactor operating temperature was 

assumed, based on the material limitations of super-alloy metals for the Brayton converter, but still requiring 

refractory metals in the reactor and primary heat transport.

4.3.2.5 Propellant-NTP: 

Nuclear Thermal Rockets may employ a few potential propellants, but the main benefit of NTP comes from high 

specific impulse. High Isp is derived from accelerating propellant molecules to high velocity, and the kinetic energy 

equation, Ek=

1

/2mv2, tells us that lower mass molecules achieve higher velocity for a given energy extracted 

from the reactor. H2 is the lowest mass molecule, and results in the Isp potential of 900 sec or higher. The next 
NASA 45

lowest mass molecules that could be considered as a practical propellants, methane or ammonia, would have 

similar or slightly higher Isp to LOX-LH2 CCP. With the lower thrust-to-weight of NTP there would be no propulsive 

advantage.

A critical technology for NTP performance is the storage, transfer, management, and containment of hydrogen 

as a propellant. To minimize volume of propellants for launch, to minimize mass for containment during in-space 

operation, and to minimize leakage over the course of a mission, hydrogen must be maintained in liquid form at 

20K cryogenic temperatures. Storing large quantities of cryogenic propellants with negligible losses has always 

been a challenge. Storing liquid hydrogen for years required for a human Mars mission compounds the 

challenge. Molecular or atomic hydrogen can leak through very small valve gaps, and even diffuse between the 

atoms in the metal lattice meant to contain the hydrogen. Any loss of hydrogen over the course of a mission is, 

in effect, a loss of propellant efficiency. Zero boil-off (ZBO) of liquid hydrogen, and very low hydrogen leakage 

during the 5+ year campaign is required. Achieving this will require the development of no/low leakage valves, 

disconnects, advanced insulation, and high-capacity, high-efficiency 20K cryocoolers. These technologies 

represent significant development risk. Though CFM has been used throughout the history of human spaceflight, 

the requirements have never been as stringent, and the criticality has never been as high as on an NTP mission 

of this duration. 

NASA is currently investing in the critical technologies for cryogenic fluid management in order to solve the longterm hydrogen, oxygen, methane, and other fluid storage challenges for propellants, ISRU products, and other 

exploration systems. Cryogenic hydrogen requires 20K cryocoolers, and advanced thermal insulation systems, 

which is a significantly greater challenge than higher temperature cryogens (the only more difficult cryogen is 

liquid helium). This technology is being pursued for multiple applications and has benefits for the terrestrial 

utilization of hydrogen.

The use of liquid hydrogen as a propellant has significant implications on vehicle design, launch of propellant, 

and orbital assembly due to its low volumetric density, and requirements for cryogenic fluid management from 

the moment liquid hydrogen is tanked on the launch pad. The low volumetric density (~70 kg/m3 compared to 

water at 1,000 kg/m3) typically limits the amount of hydrogen propellant that can be launched on any rocket by 

payload volume, rather than launch mass capability. Thus, NTP systems are driven to utilize large payload 

shrouds, and require significantly more launches to provide the necessary propellant for a human Mars mission.

Each propellant launch element requires a large propellant tank maximized to the payload shroud that includes 

either cryogenic fluid management, with its associated power system, or massive tank insulation to mitigate 

hydrogen boiloff until the propellant is integrated with orbital assets. Each propellant element in the MTAS NTP 

reference case includes a fully capable transfer vehicle to deliver the propellant element to the aggregation orbit 

where the NTP Mars Transfer Vehicle is being assembled. This transfer vehicle also provides power for active 

CFM. Both the number of launches and the number of vehicle elements adds considerably to the development 

and operational cost of an NTP human Mars mission.

4.3.2.6 Propellant-NEP: 

Electric propulsion using Hall Effect Thrusters (HET) or Gridded Ion Thrusters typically favor propellants that 

have a low first ionization energy, to minimize the energy needed to create an ion for acceleration, and a high 

atomic weight, to maximize the momentum generated by the acceleration (with low thrust acceleration 

propulsion, thrust is more important than specific impulse for fast missions). Noble gases have these advantages, 

as do cesium and iodine. Xenon, krypton, and argon have been considered for Mars class electric propulsion, 

and among these xenon is the best for relatively low Isp needed for fast missions. Xenon has additional 

advantages since it can be stored under pressure (~1,000 to 2,000 psi or ~7,000 to 14,000 kPa) in a supercritical 

state without cryogenic fluid management. Under these conditions, supercritical Xe is about as dense as a brick 

(comparatively, cryogenic hydrogen is about a dense as Styrofoam, cryogenic methane is about as dense as 

oak wood, and cryogenic oxygen is slightly denser than water). Pressure vessels for these conditions are typically 

flown in space for He pressurants, but the size needed for Xe propellant will require a much larger scale tank.
NASA 46

The main challenge with Xe is acquiring the quantities needed to propel each human Mars mission. Xe is overall 

abundant, but it is a small fraction of the constituents of Earth’s atmosphere. Air is liquified in many locations 

throughout the earth, where purified oxygen is particularly needed for steelmaking. Air is liquified in a series of 

increasingly cold condensation steps, and the noble gasses argon, krypton, and xenon are the last to condense.

Many liquification plants do not have the equipment to separate these three gasses at the end of the 

condensation line. A human Mars mission will require several years of the world’s supply of xenon at current 

production rates. Solutions to this challenge include stockpiling xenon over several years and increasing the 

infrastructure to purify xenon at existing air liquification plants throughout the world. This represents an 

opportunity for all nations to contribute to space exploration.

The hybrid NEP approach uses liquid oxygen (LOX) and methane (LCH4) as propellants for the companion 

chemical combustion propulsion. LOX and LCH4 both have cryogenic storage temperatures near ~90 to 100K, 

which is comparatively easier for cryogenic fluid management relative to hydrogen. NASA is also investing in the 

critical technologies for cryogenic fluid management of liquid oxygen and methane to solve the long-term 

propellant storage challenges. These include no/low leakage valves, disconnects, advanced insulation, and highcapacity, high-efficiency 90K cryocoolers.

4.3.2.7 Testing and Development Facilities-NTP: 

Because of the complex operation of a Nuclear Thermal Rocket engine, where core reactivity is tightly coupled

with hydrogen propellant flow, NTP development requires a full scale, integrated ground test to prove the system 

and qualify it for spaceflight. This includes start-up and shut-down transients, and because of the potential for 

reactivity and material changes over the thrust life of the reactor, should include the full mission profile of startups, shutdowns, and propellant throughput. Several integrated full system nuclear qualification tests, including 

post-irradiation examination of suspect fuel elements, would be necessary to provide confidence that the final 

design would achieve 900 sec Isp. Subsequently, a full system demonstration would be necessary to certify the

overall system for human spaceflight.

During Rover/NERVA such testing was performed in open air at the Nevada Test Site (now the Nevada National 

Security Site), where the full-scale engine exhaust was neither captured nor filtered. While the fission product 

inventory for an NTR test is very small due to the short run times of the engines, such testing cannot be performed

today without capturing and/or filtering all rocket exhaust to eliminate any risk of releasing radioactive fission 

products into the environment. This requires a new test facility. The Rover/NERVA program successfully tested 

the “Nuclear Furnace” NF-1, which demonstrated one concept for an effluent treatment exhaust system with a 

44 MWt reactor. Such a facility will be costly, and require significant time for construction, including the necessary 

approvals under the National Environmental Policy Act. The development of this facility adds significant cost and 

schedule risk for NTP. Since the end of Project Rover/NERVA, the design and development of this facility has 

been, and remains, a major hurdle for the development of NTP.

An alternative approach to avoid the development of this facility is an in-space development test. Such an 

approach is success oriented, and still has major risk. The challenge is troubleshooting a test failure, anomaly, 

or nonconformance without physical post-test inspection and the opportunity for disassembly. Without adequate 

instrumentation, a test failure, anomaly, or nonconformance may only be discovered through physical post-test 

inspection. As development of NTP moves forward, NASA will need to work closely with DOE to determine what 

fraction of NTP testing can adequately be accomplished in space.

4.3.2.8 Testing and Development Facilities-NEP: 

The NEP system development allows for separate subsystem ground testing of modular elements and simulated 

interfaces, which could be accomplished in separate facilities, easing both facility requirements and development 

and testing processes. Several nuclear facilities are in development for terrestrial small modular reactors that 

could be leveraged or adapted for testing space reactors supporting NEP. The MTAS study did consider the 

characteristics and need for new test facilities for NEP. The development of NEP will require some facility 
NASA 47

development, notably large, high-pumping capacity vacuum chambers for testing multiple, large electric 

thrusters.

4.3.3 Operational Distinctions

4.3.3.1 Operational Distinctions: Thrust

The amount of thrust acceleration any propulsion system provides in relation to the acceleration of the local 

gravity field determines the efficiency of providing change in velocity, as well as the time required. High thrust 

acceleration allows for quick changes in orbital velocity, in turn resulting in agility and quick response, as well as 

rapid departures from, and arrivals into, planetary orbits from interplanetary space. Lower thrust acceleration

results in slow orbital changes, which introduces inefficiencies from thrusting against gravity, termed “gravity 

losses.” However, with time, high propellant efficiency (high Isp) propulsion overcomes gravity losses and can 

result in an overall more efficient performance. The best of all propulsion has both high thrust acceleration and 

high Isp.

Thrust-NTP: Nuclear Thermal Propulsion represents the highest Isp, high thrust propulsion technology currently 

achievable. By comparison, NTP offers ~900 seconds while Chemical Combustion Propulsion typically offers 

~330 – 460 seconds. The effect of propellant efficiency is to significantly reduce the propellant necessary to 

perform the large changes in vehicle velocity necessary to perform the transit from Earth orbit to Mars orbit and 

back. The faster the Mars mission, and the more distant the Earth-Mars planetary alignment during the year of 

opportunity, the larger the required changes in velocity.

Thrust-NEP: The advantage of electric propulsion is that much higher propellant exhaust velocity can be 

achieved through either electrostatic, electromagnetic, or electrothermal acceleration of ionized propellant. 

Specific impulse of ~1,000 – 10,000 seconds is achievable and is a matter of how the thruster is designed, the 

voltages applied to acceleration, and the propellant choice. 

One of the challenges for the use of any electric propulsion technology is the low thrust acceleration, limited by 

power, which results in significant time to propel into and out of planetary gravity wells (regions where 

gravitational forces are much higher than deep space). Low thrust applied in a strong gravity field results in long, 

spiral shaped trajectories as thrust acceleration is applied over time. However, low thrust acceleration applied 

during interplanetary transit can be very efficient as the gravity field is much weaker, and there is significant time 

to apply acceleration. This results in potentially very fast interplanetary transit, but significant additional time to 

transit from low Earth orbits and into low orbits around planetary destinations. NEP mission times can be reduced 

with higher thrust acceleration, achieved with either advanced, lower specific mass technology, or higher power 

systems with lower specific mass by economy of scale. 

Hybrid mission approaches, utilizing both high thrust and low thrust where they are most advantageous, can 

improve overall mission performance if the mass of the combined propulsion systems is not prohibitive. Typically, 

high thrust is used in gravity well maneuvers, such as at Earth and Mars, to reduce or eliminate the long spiral 

trajectories, while low thrust is used for acceleration in interplanetary space to reduce the size of the gravity well 

maneuvers. The optimization of the propulsive maneuvers depends on the Isp of each propulsion type and the 

demands of the mission objective. The study found that the augmentation of NEP with high thrust LOX-LCH4

chemical propulsion significantly reduces the size of the electric propulsion system required, and with it the 

technologies required for the objective 2-year Mars mission.

An additional benefit of high specific impulse, combined with the efficient application of low thrust acceleration 

in interplanetary space, is the mitigation of the energetic challenges that result from more distant planetary 

alignments between Earth and Mars during more difficult mission opportunities (such as 2039 and 2042 when 

Mars is more distant from Earth at closest approach due to its elliptic orbit). In effect, higher Isp electric propulsion 

reduces the variance in mission performance between best and worst Earth-Mars planetary alignments. This 

effect was observed in the MTAS analyses between the mission performance for 2035 and 2039.
NASA 48

4.3.3.2 Operational Distinctions: Extensibility 

Both NTP and NEP require significant technology development. The development roadmap for either should 

consider appropriate and valuable steppingstones that may involve precursor systems that benefit other missions 

or applications. Furthermore, both NTP and NEP can evolve to advanced capabilities that support advanced 

human missions to Mars and human and robotic exploration beyond. The extensibility of either propulsion to 

precursor and follow-on applications is a distinction on the value of the technology investment. Future conjunction 

class missions, that might follow initial opposition class missions, may benefit from these technologies that could 

provide either: lower launch mass, faster transits, higher payload delivery, or some combination of the three. 

While extensibility was considered under MTAS, no specific mission studies were conducted to evaluate the 

merits of the various technology options for alternative applications. The following conclusions are drawn from 

past studies with appropriate references.

Extensibility-NTP: Nuclear Thermal Propulsion has been studied over the years for several applications where 

both higher Isp and high thrust acceleration is required. Project Rover initially conceived of an NTR as an upper 

stage engine for ICBMs during the Cold War, and the Air Force SNTP program considered the same. NASA’s 

NERVA effort was initially focused on an NTP upper stage for the Saturn or Nova rocket to support the Apollo 

human missions to the Moon. Only after NASA settled on the Apollo Lunar Orbit Rendezvous architecture and 

the sufficiency of the Saturn V rocket did NERVA focus on human Mars mission application. The upper-stage 

application of NTP is no longer considered necessary or appropriate. 

DOD/DARPA is considering NTP due to its high thrust acceleration and rapid response capability, and this 

provides agility in the Earth and Moon gravitational environments. These systems could be adapted to NASA for

more efficient Earth-Moon transportation supporting a robust, ongoing lunar exploration campaign.

Lower thrust-class NTP has been studied for smaller robotic science missions (Dudzinski, 1995; Ballard, 2019) 

where a higher ∆V Earth orbit departure is valued, but the benefits of such an application are highly dependent 

on the Isp and stage dry mass (Gabrielli, 2015; Rawlins, 2022). Because of NTR lower T/W and the use of large 

tanks and CFM for hydrogen storage, an NTP stage is often not advantageous over Conventional Chemical 

Propulsion stages at smaller scales. The utility of small NTP for distant destinations could be enhanced by 

bimodal operation of the reactor for both thrust and power – whereby the NTR engine could supply power to 

maintain CFM for hydrogen during the long transit, to the spacecraft, and potentially even to electric propulsion.

Such a bimodal system would enable NTP to provide ∆V at the destination of the mission as well. This is an 

advanced capability that was not considered for a mid-2030s human Mars mission development.

The extensibility of NTP to advanced applications, requiring higher ∆V, will be defined by the potential for 

advancements in specific impulse, stage dry mass (largely determined by NTR thrust-to-weight and hydrogen 

CFM system mass), NTR full-thrust lifetime, and the long-term storage of cryogenic hydrogen. The material 

limitations of solid core nuclear fuel will limit NTR Isp to about 1050 seconds. More advanced concepts for liquid 

cores and extreme concepts for gas cores may reach 1,300 seconds and 3,000 seconds, respectively, but these 

systems are merely conceptual at the time of this writing.

Extensibility-NEP: The use of nuclear power for electric propulsion was conceived of in the 1950s and depicted 

as a propulsion system for a human Mars mission in the 1957 Walt Disney film “Mars and Beyond”, where 

Wernher von Braun and Ernst Stuhlinger were consultants. NASA has considered NEP for the Jupiter Icy Moons 

Orbiter, and NEP has been studied for missions such as a Neptune-Triton orbiter, and an Interstellar Probe, 

where extreme ∆V is needed at distances from the Sun where high photovoltaic power is impractical (Jones, 

1984; Yam, 2012; Noca, 2003; Nock, 2012; Satter, 2005; Oleson, 2004; Oleson 2019). The practical mission ∆V 

for NEP is limited by the mass-to-power ratio (specific mass), and specific impulse. The Isp potential of electric 

propulsion (10,000+ sec) is enabling for even more advanced applications as the specific mass of space nuclear 

power is lowered by advances in technology.
NASA 49

The development of NEP technology for Mars has potential to both leverage other development and be 

leveraged to benefit other missions as a precursor application, such as subscale systems for robotic science. 

Mars NEP systems could evolve from systems already being developed for near-term NASA missions 

including the 10 kWe Fission Surface Power Technology Demonstration Mission and the 50 kWe Gateway 

Power and Propulsion Element. A potential intermediate development step between FSP and Mars NEP is to 

adapt the 10 kWe FSP system for use with off-the-shelf EP thrusters to perform a robotic orbiter mission to 

Uranus, Neptune, or Pluto. A complete NEP concept based on a ~10kWe evolution of the Kilopower reactor is 

also under consideration for development by the U.S. Space Force (Helios-N) and a partnership could possibly 

be leveraged by NASA for these missions. NEP technology has potential feed forward benefits to future NASA 

science and exploration applications at various scales, such as a heat source (e.g., ice world melt probes at 

~50 kWt) a large-scale surface power system (e.g., commercial ISRU propellant production at ~1 MWe) as well 

as high ∆V, large payload propulsion (e.g., Europa subsurface mission at ~200 kWe) (Oleson 2019; Oleson 

2022; Oleson 2022). MTAS did not study these applications but identified them from past studies (Gowen, 

2011).

4.3.3.3 Operational Distinctions: Reusability 

The development of NTP or NEP to support a human Mars mission will require significant investment in time, 

effort, and funding, and may also require the development of unique facilities. Once developed, each propulsion 

system will be more valuable if it is reusable. However, the additional requirements for reusability and extended 

lifetime will themselves increase the investment necessary for the first system. This value proposition must be 

explored and fully assessed in consideration of a reusability requirement. MTAS Phase 1 considered but did not 

fully assess the implications of reusability. Such an important assessment is recommended forward work.

Reusability-NTP: There are several factors limiting NTP lifetime that will affect reusability for Mars missions.

High temperature hydrogen corrosion of fuel elements and coatings in NTR cores is a significant consideration 

for materials development and fuel temperature (which drives Isp) and sets a limit on engine operational lifetime 

to a few hours of thrust operation. Extending the lifetime of NTR cores, given a specific material type, requires 

operating at lower hydrogen temperatures. Thus, longer life, reusable NTRs are possible with reduced specific 

impulse5. An additional approach to enabling NTR reusability is to reduce the total engine operation time in a 

mission by increasing the number of engines on the vehicle at the expense of higher engine mass. Increasing 

from two engines to three would reduce burn times by a third, but increase the propellant required by the 

additional engine mass.

Use of hydrogen compounds as moderator materials also place limitations on lifetime based on hydrogen 

dissociation over time, which is accelerated at higher temperatures. Moderated NTR cores can address this 

issue by designing to maintain hydrogen compound moderator materials at lower operating temperatures to 

extend lifetime at the expense of performance.

Finally, the ConOps for the reference NTP case utilizes drop-staging of empty propellant tanks to minimize roundtrip propellant mass (eighteen propellant tanks and two stub core stages are dropped for the 2039 mission). This 

significantly improves performance but requires the resupply of that hardware to reuse the core DST vehicle. 

The NTP reference case also requires 70% more propellant mass than NEP/CCP for the same mission. Thus, 

a reusable NTP system would require additional hardware production as well as greater propellant mass

resupply to support subsequent missions.

5 The current con-ops presented here does not reuse NTRs
NASA 50

Reusability-NEP: The Mars Nuclear Electric Propulsion system has significant potential for reusability, since

space power reactors can be designed for long operating life. The factors for NEP reusability include reactor 

lifetime, power conversion wear and failure mechanisms, and electric propulsion thruster throughput capability.

Space nuclear power reactors are typically limited by fuel burnup and issues with fission product build-up in the 

core over time. Such reactors can be designed for long life by adding U235 fuel to the core – resulting in larger, 

more massive cores and the corresponding shielding. This has some effect on performance, but at MWe scales 

the effect is small. NEP reactors may also experience operational life challenges from lithium corrosion and 

refractory alloy embrittlement, which must be addressed by materials development. Helium gas generation over 

time must be captured and controlled in the system design. More significantly, the incorporation of hydride

moderator materials will limit lifetime based on hydrogen dissociation over time. As with NTP, fast neutronspectrum HEU reactors are inherently simpler with greater potential for long-life.

Brayton power conversion for space applications is engineered to eliminate sources of wear by employing gas 

or magnetic bearings. Other potential areas of concern include material creep in the high temperature 

turbomachinery and the hot-side plenums. Radiators and other fluid systems will also be subjected to micrometeoroids and orbital debris during the mission and may develop leaks and holes from impacts. The fault 

tolerance of these systems will need to be addressed through both shielding and redundancy. Utilizing heat 

pipes with high-conductivity fins in the NEP radiators helps to reduce the vulnerable area and minimize the 

impacts of unit failures. The reliability of electric circuits and components operating at high voltage and high 

temperature in a radiation environment will also need to be addressed through design and testing, which will 

impact schedule and cost. Finally, most electric thrusters have a limitation on propellant throughput due to 

erosion of thruster components by interaction with the accelerated propellant. The thruster plume also may erode 

surfaces of the power and propulsion system, and mitigation of this concern may drive the vehicle configuration.

A straightforward approach to address EP thruster life is to replace them between missions, and such 

requirements can be designed into the system. Alternatively, an NEP system can carry spare thruster sets, as 

the thruster masses are a comparatively small percentage of the overall propulsion system.

4.3.3.4 Operational Distinctions: Robustness

Robustness involves the propulsion systems agility to respond to contingencies and prevent loss of mission, and 

more importantly loss of crew. While robustness can be designed into a propulsion system, some elements of 

robustness are natural characteristics of each propulsion system and may present distinctions between them. 

The issue of mission robustness involves complex mission modelling and analysis, and was not assessed in 

MTAS, but is recommended for additional analysis in a future study. This assessment will be required to support

actual mission design for a human mission to Mars. For this study, both NTP and NEP system designs and 

mission analyses accounted for engine-out and partial thrust and power contingencies.

Robustness-NTP: One benefit of increased propellant efficiency is that contingency propellant can be carried 

throughout the mission for less penalty, thereby affording mission robustness. The higher Isp of NTP over CCP 

allows for this. Other mission operations also can provide for mission robustness, such as predeploying

contingency propellant or replacement propulsion stages at Mars using the cargo mission capacity afforded by 

NTP. High thrust acceleration, such as with NTP, also provides for rapid response to contingencies.

Additionally, The multiple NTR engines within a Nuclear Thermal Propulsion system provides for engine-out 

capability. However, the resulting assessment of mission reliability with multiple reactor engines must consider 

all failure modes, including possible catastrophic reactor failure, which could affect multiple engines and even 

destroy the vehicle. The NTP reference case includes four NTR engines to provide the necessary thrust at Earth 

departure, when the vehicle is most massive, while reducing both burn time and gravity losses. Two of these 

engines are jettisoned on parallel stack stages after a deep space maneuver. The reference performance 

assumes one-engine out operation on the two-engine core stage returning from Mars.
NASA 51

Robustness-NEP: NEP offers even higher propellant efficiency than NTP, and therefore, better affords the 

mission lien of carrying contingency propellant. Responding to contingencies in interplanetary space with low 

thrust acceleration will require more time, but that time is available during interplanetary transit. However, while 

the time for low thrust acceleration response to contingencies in planetary orbit may be an issue, the hybrid 

NEP/CCP approach provides a high thrust capability for rapid response if necessary. The availability of both high 

Isp low thrust capability and high thrust capability allows for increased mission flexibility in response to 

contingencies. The MTAS NEP/CCP reference case assumes the LOX-CH4 stage is present and available until 

Mars departure. The MTAS NEP/CCP reference case also assumes contingency propellant is carried throughout 

the mission, though half of the contingency propellant is left at Mars to increase performance on the return leg 

(this ConOps should be reconsidered for robustness in the next phase of analysis).

The reference NEP system was designed with one reactor and four, independent Brayton power conversion 

strings, and can operate at partial power with fewer thrusters if one or more power conversion strings fail. Each 

Brayton string includes a dedicated heat rejection segment (radiator) and PMAD channel that supplies power to 

five Hall thrusters. Consideration was given to including spare power strings for the reference design, but that 

approach was rejected because the benefits were negated by the added complexity. Again, the resulting mission 

reliability with multiple strings must consider all failure modes, including possible catastrophic reactor failure,

which could affect all strings or eliminate the prime power conversion and even destroy the vehicle.

MTAS performed a cursory analysis demonstrating that the mission could be completed on partial power thrust, 

albeit with a longer mission time. A more thorough analysis is required. Past studies (Barnett, 1991; Jansen, 

2016; Casani, 2020) have documented analyses that show NEP has a robust partial power abort potential 

(Dudzinski, 1992) throughout the mission due to high Isp. These studies also identified that NEP offers wide 

departure windows from Earth and Mars, adding to the mission robustness.

Finally, electric propulsion provides an additional benefit to the robustness of the overall human Mars campaign 

by reducing the performance variance across Mars transit opportunities. This effect was observed in the MTAS 

analyses between the mission performance for 2035 and 2042, where the same size power system, and same 

specific impulse EP system completes both opportunities with similar performance. This allows for more 

confidence in mission planning and mitigates schedule risk for development.
NASA 52

4.3.4 Campaign

CREW MISSION LAUNCH MASS

SUMMARY

~3yrs

1

st Element to 

Landing

678 t

Total Launch Mass

NEP/Chem

~2.5yrs

1

st Element to 

Landing

934t

Total Launch Mass

NTP

NEP/Chem NTP

2029 2030 2031 2032 2033 2034 2035 2036 2037 2038 2039 2040 2041 2042 2043 2044

CREWED MISSION MILESTONES

NEP/Chem

NTP Deploy Habitat Deploy Habitat Crew Habitat Launch Shakedown Crew on Mars Habitat Shakedown Crew on Crew Mars Launch Crew Return Crew Return

MARS CREWED MISSION TIMELINE

01/2036

01/2037

01/2038

01/2039

01/2040

01/2041

07/2036

07/2037

07/2038

07/2039

07/2040

07/2041

NEP/Chem

NTP

0

100

200

300

400

500

600

700

800

900

1000

Final Earth Departure Mass ()

Propellant

Tankage Dry

Propulsion Dry

Nuclear Sys Dry

Logistics

Habitat NEP Deploy NTP Deploy

Crew NEP Module

Xenon Module

Xenon Module

Xenon Module

Chemical Module

SLS NTP Core

SLS NTP Inline

Truss

NTP Stub Core

PPM

CLV Inline

SLS Launch

Loiter 

EP Spiral

HEO-NRHO Transit 

H-CLV Launch Interplanetary Transit

Mars Orbit / Surface 

Crew Launch Window

SH-CLV Launch

Figure 4.3-2. Comparison of NTP and NEP/CCP Crew Mission Campaigns.

The MTAS campaign assessment found the launch, assembly, and mission characteristics of the two propulsion 

systems are similar in many aspects. Where they differ is summarized in Figure 4.3-2 for the first crew mission, 

and in Figure 4.3-3 for the supporting cargo missions. The mass of the vehicles as they would leave Earth orbit 

differ by 37%. The propellant density of the two propulsion systems drives this difference, as the dry mass of the 

propulsion systems is about the same. As discussed earlier, an NTP system optimizes with a propellant with low 

atomic mass to increase Isp, whereas an electric propulsion system optimizes with a propellant with high atomic 

mass to maximize thrust. As a result, the mass of the propellant required only differs by ~10% (~60,000 kg) but 

the mass of the tanks required to store those propellants differ by a factor of 7 (more than 200,000 kg). The mass 

of hydrogen propellant and tanks, as well as the volume, drive NTP’s need for 18 SH-CLVs to support the crew 

mission alone. Because the CFM of hydrogen is so challenging, the desire is that they be delivered over a short 

period. By comparison, the denser NEP and LOX-LCH4 propellants are delivered in 5 SH-CLVs for the same 

mission and could be launched on a greater number of smaller CLVs without concern for launch volume.

Comparing the cargo delivery approaches; because the plan with the NTP campaign is to send the first lander 

using NTP, it can deliver the first lander, a campaign pacing item, to Mars on a 2.5-year mission. The NEP/Chem 

plans to use the chemical propulsion system to deliver the first lander on a slower, 3-year transit time. This 

means the lander development plan for the NTP option can be 6-12 months longer. However, under this 

scenario, the first NTR engine and NTP system must be deployed about a year earlier, in 2035 to support a 2039 

piloted mission, than the NEP option, which deploys the first NEP module in 2036. Thus, NTP would require a 

shorter development time than NEP/CCP. 
NASA 53

MARS CARGO STACK MASS

SUMMARY

~3.9yrs

1

st Lander Lands

To Crew Landing

479t t

Total Cargo

Stack Mass

NEP/Chem

~4.1yrs

1st Lander Lands

To Crew Landing

487t

Total Cargo

Stack Mass

NTP

NEP/Chem NTP

2029 2030 2031 2032 2033 2034 2035 2036 2037 2038 2039 2040 2041 2042 2043 2044

CARGO MISSION MILESTONE

NEP/Chem

NTP Lander 2/3 Arrives Arrive Crew Lander 1 Arrives Lander 1 Arrives Arrive Crew Lander 2/3 Arrives

MARS CARGO MISSION TIMELINE Departs Lander 1 Departs Lander 1 Lander 2/3 Departs Lander 2/3 Departs

0

100

200

300

400

500

600

Total Reuqired Mass ()

Propellant

Propulsion Dry

Nuclear Sys Dry

Lander 3

Lander 2

Lander 1 NEP Deploy NTP Deploy

01/2034

01/2035

01/2036

01/2037

01/2038

07/203

07/2034

07/2035

07/2036

07/2037

07/2038

NEP/Chem

NTP

Cargo NEP Module

Xenon Module

Lander #3

Lander #2

Lander #1

Cargo Chemical 

Module

Refueling Modules

Lander #2

Truss

Cargo Stub Core

CLV Inline

Lander #3

Cargo Stub Core

Lander #1

CLV Inline

01/203

SLS Launch

Loiter 

EP Spiral

HEO-NRHO Transit 

Interplanetary Transit

H-CLV Launch

Mars Orbit / Surface 

SH-CLV Launch

Figure 4.3-3. Comparison of NTP and NEP/CCP Cargo Mission Campaigns.

Both NTP and NEP/CCP crew mission approaches require two SLS launches for the propulsion systems. 

Additionally, the NTP system requires five Heavy-CLV launches for truss structures, power, and spacecraft 

systems, and additional side mounted Stub Core Stages. Overall, the NTP crew mission requires 26 unique

launches, while an NEP/CCP crew mission requires 7 unique launches. The NTP cargo delivery approach 

requires 12 unique launches, while the NEP/CCP approach requires 14 unique launches to support a single 

Mars mission. However, again because of the larger launch volume of hydrogen propellant, the NTP option 

requires Super-Heavy CLVs, with larger payload volume. The number and type of launch vehicles required to 

support the NTP and NEP/CCP Mars campaign is shown in Figure 4.3-4. The overall higher number of launches 

required to support the NTP campaign, as well as the higher number of larger, more costly launch vehicles is 

reflected in the higher NTP mission cost in the comparison, below.
NASA 54

Figure 4.3-4. Comparison of Launch Count for NTP and NEP/CCP Mars Campaigns.

From an integrated campaign launch cadence perspective, the aggregation of the elements in lower Earth orbit 

significantly reduces the number of launches required to support both the crew and the cargo missions. However, 

the elements cannot loiter in the lower Earth orbit for long duration due to thermal and Micrometeoroid and Orbital 

Debris (MMOD) considerations (like ISS). The aggregation orbit must account for the thermal heat from the Sun 

and Earth reflection, and possible MMOD impact. To minimize the MMOD risk, both NTP and NEP/CCP 

reference cases selected an altitude with minimum MMOD, and the reference tank designs account for the worstcase total heat loads. With this in mind, the aggregation strategy increases the burden on the overall launch 

cadence of the elements during the aggregation phase with a significant number of launches required in a 

relatively short period of time. Significant improvements to the launch rate of commercial launch vehicles are

currently being implemented by several commercial vendors, but this risk to the overall mission should be 

monitored.

Note that, due simply to the number of launches required, the probability of successful completion of either 

launch campaign is quite low (0-20%), with NTP probability perforce being lower. However, if as few as four 

spare launch vehicles and associated lower value elements are budgeted, this probability rises above 50% for 

both campaigns.
NASA 55

4.3.5 Cost

As a cost comparison between these alternatives has never been performed to this level of detail, the MTAS 

Programmatic Assessment Team performed a detail cost assessment of the NTP and NEP/CCP 1.2 reference 

cases based upon the best available historical data, bottoms-up and heuristic approaches, and collective expertjudgments. These assessments contain significant uncertainties in absolute cost figures but are valuable for 

relative comparisons between the alternatives.

The MTAS Programmatic Assessment found a significant (~37%) cost advantage for NEP/CCP over NTP when 

considering both development and first Mars campaign use. The major reason for the cost difference results

from the larger number of unique NTP elements and launches. Figure 4.3-5 shows an overall cost comparison

of the summation of costed elements and activities supporting a first Mars mission, and Figure 4.3-6 compares 

a breakdown of these cost considerations. The development (DDT&E) cost estimates for NEP and NTP each 

have large uncertainties due to technology challenges, and the differences identified to date are not significant 

compared with those uncertainties. Further technology maturation is required to answer key technology 

questions that will better bound the DDT&E costs, and significant differences between the options may be 

revealed once the technologies are more mature. DDT&E cost estimates for NTP were also affected by choices 

in the NTP vehicle design made to minimize the number of launches at the expense of requiring the development 

of additional, unique propellant elements. This approach should be revisited in future analyses to assess the 

potential to reduce NTP cost. Forward work to assess subsequent mission costs is also recommended, 

particularly to assess new build/launch of elements that are jettisoned during first mission use, and to assess 

retained element life limit extension costs for subsequent missions.

Figure 4.3-5. Cost Comparison of NTP and NEP/Chem Approaches to Mars 2039 in Millions.

Another area of cost uncertainty in DDT&E is the need for, and cost of, new facility development. It has long 

been considered, since the end of the Rover/NERVA program, that any new NTR engine development requires 
NASA 56

a new test facility to capture and/or filter the propellant exhaust. An estimate of these costs is reflected in the 

NTP DDT&E. A new facility for an integrated NEP system test is also included in the NEP DDT&E for comparison; 

however, several nuclear facilities are in development for terrestrial small modular reactors that could be 

leveraged or adapted for the development of space reactors supporting NEP with significant cost saving 

potential. Furthermore, the modular characteristics of NEP systems allows for separate subsystem ground 

testing of modular elements with simulated interfaces for qualification, which could be accomplished in separate 

facilities, easing the requirements for new facility development. Thus, the NEP DDT&E estimate is considered 

potentially conservative. That stated, the development of NEP will also require some facility development, 

notably large, high-pumping capacity vacuum chambers for testing multiple, large electric thrusters, which is 

included in the NEP DDT&E.

Figure 4.3-6. Cost Breakdown of NTP and NEP/Chem Approaches to Mars 2039 in Millions.

A tangible advantage of the high Isp and high propellant density of the NEP/CCP approach is the significant 

reduction in launches and launch costs due to the comparatively low propellant mass required, as well as 

significantly smaller launch volume for the propellants. This advantage results in not only ~20% lower launch 

costs, but over two times lower hardware cost for the NEP/CCP vehicle, which has significantly fewer elements.

This advantage is compounded when considering the comparison of cargo delivery vehicle costs for the 

campaign.

4.3.6 Other Considerations

The effect of system generated radiation on the crew and overall system must also be considered and may offer 

a distinction between NTP and NEP. MTAS assessed the shielding that would provide equivalent system doses 
NASA 57

to the crew to provide comparable performance. However, to be complete, this assessment must consider the 

crew habitation and its configuration and integration with the propulsion system, which is not fully know at the

time of this study. An integrated assessment with the Transit Habitat research team and with the Office of the 

Chief Health and Medical Officer within NASA is still necessary. Furthermore, NASA’s Office of the Chief Health 

and Medical Officer must develop crew standards for radiation exposure that consider both onboard radiation 

and natural sources in space, such as Galactic Cosmic Radiation (GCR) and Solar Particle Events (SPE).

Therein lies a significant mission – technology trade between the propulsion to drive even faster missions to 

reduce crew exposure to natural sources, and the radiation those systems generate. This more detailed and 

comprehensive assessment has been identified as work for future study.

A further distinction between NTP and NEP may also be revealed upon completion of a Loss-of-crew/Loss-ofmission (LOC/LOM) assessment for each of the options. For example, while the presence of multiple fission 

cores in the NTP concept (vs. the single core in the NEP option) may appear to provide protection against a 

failure mode leading to a SCRAM of a reactor core (i.e., “engine out” capability), this factor also presents a 

notably increased exposure to a failure mode resulting in a reactor explosion. An assessment of which option 

would offer lower LOC/LOM risk is not trivial and is beyond the scope of this study.

5 . 0 C O N C L U S I O N S

This study is an initial step toward a complete understanding of the Mars transportation trade space and options. 

The objective of the short-stay approach is to minimize the overall mission risk, including risk to the exploration 

crew, as well as the programmatic risk of developing the numerous technologies needed to support a long 

surface stay for the exploration crew on Mars. The study found that the high ∆V required for fast Mars missions 

with short stay times drives the need for nuclear propulsion technology and highlights the performance 

differences between promising alternatives. Numerous past studies have shown that non-nuclear options require 

extremely aggressive technologies and concepts of operations to close a fast Mars mission. An apples-to-apples 

all-chemical ConOps is not likely to be viable. However, an objective assessment of non-nuclear options must 

be completed, and is recommended in the table of follow-on work, below.

In summation, both NTP and NEP/CCP propulsion system concepts have the potential to fulfill the mission 

requirements for fast Mars transits in the late 2030s and meet the Agency’s objectives, including enabling ~2-

year roundtrip mission durations. But these options use different approaches with unique challenges. NTP offers 

more efficient high thrust capability for rapid response to contingencies and gravity well departures to support a 

human Mars mission as well as DOD applications. NTP advantages result from the highest Isp for currently 

achievable high acceleration propulsion. However, low thrust NEP, when augmented with high thrust LOX-LCH4

conventional chemical propulsion looks very promising as the more advantageous alternative. The NEP/CCP 

approach results in comparatively: fewer launches with fewer number and types of elements; lower launch costs, 

unit costs & campaign costs; and easier launch, storage, and management of propellants. An additional benefit 

of higher specific impulse with NEP is the mitigation of the energetic challenges that result from more distant 

planetary alignments between Earth and Mars reducing the variance in mission performance across the synodic 

cycle. Furthermore, NEP/CCP has the potential for easier development due to the possibility of separate 

subsystem ground testing possible with modular elements and interfaces that can be accurately simulated.

Additionally, there are stronger NEP synergies and feed forward from terrestrial energy and lunar power, and 

strong leveraging of existing development for electric propulsion (Gateway SEP Power and Propulsion Element), 

PMAD (MWe all-electric aircraft). There are additional synergies with ongoing LOX-LCH4 propulsion 

development. NEP technology challenges for most of the subsystems will be addressed by fission power 

development for the lunar surface, and by SEP for Gateway. None of these challenges are at high risk to become 
NASA 58

showstoppers6. In contrast, NTP has at least three technology development challenges that are at high risk to 

become show stoppers: a fuel and core design to achieve lifetime at 900 sec Isp, test facilities for full scale 

integrated testing, and long term hydrogen CFM. Finally, NEP offers stronger benefits for precursor and followon potential applications, such as robotic science and missions beyond-Mars. In consideration of the broad 

NASA needs for propulsion, and even broader national needs, both NTP and NEP possess an imperative for 

investment, as they each have significant advantages in specific applications. However, if the apparent Mars 

NEP/CCP performance metrics can be verified in technology development, NEP/CCP would be the more 

advantageous option for human and cargo transportation to Mars. Furthermore, NEP/CCP offers significant cost 

advantages over NTP for development and first Mars campaign use, and these could be multiplied by greater 

potential for reuse.

MTAS programmatic assessment found both concepts could complete Operational Readiness Review (ORR) 

by ~2035 assuming no schedule delay, but only if significant investment in nuclear propulsion begins 

immediately. Furthermore, the federal policy guidance in SPD-6 that HEU should be limited to applications for 

which the mission would not be viable with other nuclear fuels. The consideration of HALEU presents an 

additional schedule challenge because it is a departure from the HEU heritage fuel and reactor development for 

both systems. HALEU will require advanced technology for fuels and moderators, adding cost, schedule, 

complexity, and development risk. The only reason to consider HALEU for space applications is to address 

policy concerns with the use of HEU.

Both NTP and NEP require an aggressive development program to meet a 2030s launch date. A strategically 

focused investment plan in both NTP and NEP is recommended to address key feasibility and development 

questions and retire risk before selecting a propulsion system for human Mars transportation. The technology 

development plan recommended in this document matures both NTP and NEP to TRL 5 with an integrated

component test in a relevant environment. For NTP, this requires three focus areas: first, developing and testing

full scale fuel elements to validate ability to produce an Isp of 900 seconds or higher for greater than 4 hours;

second, demonstrating viable NTP ground test exhaust capture and filtering to support full-scale facility 

development necessary for flight validation; and third, demonstration of CFM technologies necessary for longduration hydrogen propellant storage and transfer. For NEP this is defined as a 1,000-hour thermal vacuum test 

campaign with a non-nuclear representative scale power subsystem Technology Demonstration Unit (TDU), 

parallel fuel and moderator irradiation testing at the required temperature, fluence and burnup, and a separate 

vacuum chamber test of an EP subsystem TDU. LOX-CH4 chemical propulsion will require demonstration of 

CFM technologies necessary for long-duration liquid oxygen and liquid methane propellant storage and transfer. 

The down selection of a propulsion system would happen after both NTP and NEP subsequently complete a 

preliminary design review. NASA is currently investing in NTP to support synergy with other agencies and

Congressional direction; however, as a result of the favorable NEP/CCP conclusions provided by the MTAS 

assessments, a balanced portfolio of investment in both NTP and NEP is necessary until the PDR milestones 

are met, and sufficient risk is retired to confidently support a decision.

The propulsion choice for human transportation beyond the Moon could set the course of human exploration for 

the U.S. and humanity for decades to come. It will be a decision that defines a decade-long development worth 

more than $10B. This decision must be made based on sound, objective analysis, and strong technical 

confidence, for which the current state of knowledge is insufficient. Thus, the MTAS team recommends further, 

6 Measured by lower average and peak Advancement Degree of Difficulty (AD2) as assessed by the NESC independent 

study.
NASA 59

more detailed studies to guide investment, and funding for a strategically focused technology maturation in both 

NTP and NEP to support the best propulsion decision for humanity’s first interplanetary spaceship. 

RECOMMENDED FOLLOW -O N WORK

• Complete initial study scope . 

• Complete a non-nuclear reference case and conjunction reference case for comparison 

• Thorough assessment of issues touched on in Phase 1: mission robustness, human rating, 

reusability, extensibility 

• Complete remaining HEU vs HALEU independent assessment in order to provide an objective, 

comprehensive assessment of the relative merits and challenges. Include an assessment of fast/epithermal NTP 

• Assess key technology trades, including subsystem technologies, to inform design down-selection 

toward focused technology investments, and quantify performance enhancing potential for advanced 

technology investments 

• Provide an analytical foundation for investment and support Agency partnerships for long-term success 

• Engage DOD to participate in synergistic requirements development 

• Continue engagement with DOE and DOE Labs to build understanding of NASA missions and 

requirements 

• Look at other demand side missions for NASA, other Agencies, emerging commercial interests, and 

international partnerships 

• Answer key questions emerging at the end of MTAS Phase 1 

• Assess mission sensitivity to Key Performance Parameters. 

• Assess the value to the Agency of precursor missions using subscale systems on the development 

path. Consider a precursor mission to demonstrate both propulsion system and full-scale Entry, 

Descent, and Landing capability 

• Assess potentially advantageous mission architecture trades, such as predeployment of propellant to 

Mars, crew direct entry at Mars, and crew direct entry at Earth return with subsequent rerendezvous 

of the Deep-Space Transport habitat and propulsion elements 

• Deeper evaluation of the facilities needed for technology, development, I&T, and flight 

• Address 1.2 shortcomings - Cargo delivery scenarios need refinement 
NASA 60

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