Stanley K. Borowski

Glenn Research Center, Cleveland, Ohio

David R. McCurdy and Thomas W. Packard

Vantage Partners, LLC, Brook Park, Ohio

Nuclear Thermal Propulsion (NTP): A Proven, 

Growth Technology for “Fast Transit” Human

Missions to Mars

NASA/TM—2014-218104

October 2014

AIAA–2013–5354
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Stanley K. Borowski

Glenn Research Center, Cleveland, Ohio

David R. McCurdy and Thomas W. Packard

Vantage Partners, LLC, Brook Park, Ohio

Nuclear Thermal Propulsion (NTP): A Proven, 

Growth Technology for “Fast Transit” Human

Missions to Mars

NASA/TM—2014-218104

October 2014

AIAA–2013–5354

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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Acknowledgments

The authors wish to thank Chris Moore and John Warren from NASA Headquarters, Mike Houts from NASA Marshall Space 

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

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

artwork and images depicted in Figures 5, 6, and 7 of this paper.

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NASA/TM—2014-218104 1 

Nuclear Thermal Propulsion (NTP): A Proven, Growth Technology 

for “Fast Transit” Human Missions to Mars 

Stanley K. Borowski 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

David R. McCurdy and Thomas W. Packard 

Vantage Partners, LLC 

Brook Park, Ohio 44142 

Abstract 

The “fast conjunction” long surface stay mission option was selected for NASA’s recent Mars Design 

Reference Architecture (DRA) 5.0 study because it provided adequate time at Mars (~540 days) for the 

crew to explore the planet’s geological diversity while also reducing the “1-way” transit times to and from 

Mars to ~6 months. Short transit times are desirable in order to reduce the debilitating physiological 

effects on the human body that can result from prolonged exposure to the zero-gravity (0-gE) and 

radiation environments of space. Recent measurements from the RAD detector attached to the Curiosity 

rover indicate that astronauts would receive a radiation dose of ~0.66 Sv (~66 rem)—the limiting value 

established by NASA—during their 1-year journey in deep space. Proven nuclear thermal rocket (NTR) 

technology, with its high thrust and high specific impulse (Isp ~900 s), can cut 1-way transit times by as 

much as 50 percent by increasing the propellant capacity of the Mars transfer vehicle (MTV). No large 

technology scale-ups in engine size are required for these short transit missions either since the smallest 

engine tested during the Rover program—the 25 klbf “Pewee” engine is sufficient when used in a 

clustered arrangement of three to four engines. The “Copernicus” crewed MTV developed for DRA 5.0 is 

a 0-gE design consisting of three basic components: (1) the NTP stage (NTPS); (2) the crewed payload 

element; and (3) an integrated “saddle truss” and LH2 propellant drop tank assembly that connects the two 

elements. With a propellant capacity of ~190 t, Copernicus can support 1-way transit times ranging from 

~150 to 220 days over the 15-year synodic cycle. The paper examines the impact on vehicle design of 

decreasing transit times for the 2033 mission opportunity. With a fourth “upgraded” SLS/HLV launch, an 

“in-line” LH2 tank element can be added to Copernicus allowing 1-way transit times of 130 days. To 

achieve 100 to 120 day transit times, Copernicus’ saddle truss/drop tank assembly is replaced by a “star 

truss” assembly with paired modular drop tanks to further increase the vehicle’s propellant capacity. The 

HLV launch count increases (from ~5 to 7) and a fourth engine is needed to reduce total mission burn 

time and gravity losses. Using a “split mission” approach, the NTPS, in-line tank and the saddle truss/LH2

drop tank elements can be configured as a pre-deployed Earth Return Vehicle/propellant tanker 

supporting 90-day crewed mission transits. The split mission approach also eliminates the need for onorbit assembly. Mission scenario descriptions, key features and operational characteristics for five 

different vehicle configurations are presented. 

Nomenclature 

IMLEO initial mass in low Earth orbit 

K temperature (degrees Kelvin) 

klbf thrust (1000’s of pounds force) 

LEO low Earth orbit (= 407 km circular) 

LOX/LH2 Liquid oxygen/liquid hydrogen propellant 
NASA/TM—2014-218104 2 

RAD Radiation Assessment Detector carried on Mars Science Laboratory’s Curiosity rover 

SLS/HLV Space Launch System/Heavy Lift Vehicle 

Sv Sievert (1 Sv = 100 rem) 

t metric ton (1 t = 1000 kg) 

V velocity change increment (km/s) 

1.0 Introduction and Background 

In 2007 and 2008, NASA conducted a multi-center, agency-wide study that examined the mission, 

payload and transportation system options and requirements needed to land humans on Mars in the post2030 timeframe. The Mars Design Reference Architecture (DRA) 5.0 study (Ref. 1) provided an update 

to NASA’s earlier DRM 3.0 and 4.0 studies (Refs. 2 to 5) and was conducted in two phases. In Phase I, 

key architectural approaches, mission drivers and technology options were evaluated to support a downselection process. Selected architectural approaches and systems concepts were then further refined in 

Phase II while also evaluating different surface exploration strategies. A three mission set of Mars landing 

missions over an ~10 year period was assumed with each mission carrying a crew of six. Opposition and 

fast-conjunction trajectories supporting short and long surface stay landing missions were examined 

(shown in Fig. 1) during Phase I. Mission characteristics and features associated with each trajectory type 

were evaluated and compared against the in-space transportation system requirements and the amount of 

scientific investigation that could be achieved using each mission type. The scientific community 

overwhelmingly preferred going to different sites and staying as long as possible so the crew could 

sample the planet’s rich geological diversity and maximize the scientific return on each mission. While 

opposition trajectories typically have short round trip mission times, the stay times at Mars are also short. 

This trajectory class also usually has one short and one long transit leg with the latter involving a Venus 

swing-by (VSB) maneuver that can bring the spacecraft relatively close to the Sun. By contrast, fast 

conjunction trajectories have long round trip and surface stay times, but transit times to and from Mars 

can be relatively short and no VSB is required on either mission leg. The long surface stay mission also 

has a lower total V requirement than the short stay option, and therefore requires less propellant and less 

mass delivered to LEO. Because of its attractive features—lower total V, a long surface stay, short 

transit times, and no VSB—the fast conjunction/long stay mission option was selected as the baseline for 

DRA 5.0. 

Figure 1.—Features and characteristics associated with opposition and fast 

conjunction Mars missions. 
NASA/TM—2014-218104 3 

Recent measurements of the energetic particle radiation environment inside the Mars Science 

Laboratory (MSL) spacecraft during its 253-day journey to Mars were made by the RAD attached to the 

Curiosity rover located inside MSL (Ref. 6). The detector measured an average of ~1.8×10–3 Sv per day 

indicating that astronauts could receive a radiation dose of ~0.66 Sv (~66 rem)—the limiting value 

established by NASA—during a fast conjunction Mars mission involving a 1-year journey through deep 

space. This dose rate could also be higher depending on the time in solar cycle that the mission occurs. 

When coupled with the potential for additional crew dose received during the lengthy exploration phase 

on Mars, it is of interest to see what proven NTP can do to reduce transit times further. Can NTP cut 

transit times in half and if so what are the engine requirements and vehicle configuration impacts? 

The NTR was selected as the propulsion system of choice in DRA 5.0 because of its high thrust 

(10’s of klbf) and high specific impulse (Isp ~900 to 940 s), its increased tolerance to payload mass growth 

and architecture changes, and its lower IMLEO which is important for reducing the HLV launch count, 

overall mission cost and risk. With a 100 percent higher Isp than today’s best chemical rockets, the use of 

NTP reduced the required launch mass for the baseline long stay mission by over 400 t—the equivalent 

mass of the International Space Station. For the higher energy, short stay opposition missions examined 

during Phase I, NTP provided an even greater mass savings—over 530 t compared to chemical 

propulsion. This is important to note since short stay missions will likely be selected by NASA when it 

decides to send humans to orbit Mars in the early 2030s as outlined in the United States’ National Space 

Policy (Ref. 7). Lastly, it should not be overlooked that NTP is a proven technology successfully ground 

tested at the performance levels required for a human mission to Mars. The smallest and highest 

performing engine tested during the Rover/NERVA programs (Ref. 8)—the 25 klbf “Pewee” engine is 

sufficient for a wide variety of human mission to Mars when used in a clustered engine arrangement. This 

means there are no large scale-ups in size or performance needed for NTP as there are with other 

proposed non-chemical propulsion options. 

DRA 5.0 featured separate cargo and crewed Mars transfer vehicles each using a common “core” 

NTPS with three 25 klbf Pewee-class engines. Two cargo vehicles were used to pre-deploy surface and 

orbital assets to Mars ahead of the crew who arrived during the next mission opportunity (~26 months 

later). The “Copernicus” crewed MTV (Fig. 2) is a 0-gE vehicle (Refs. 9 and 10) consisting of three basic 

components: (1) the NTPS; (2) the crewed payload element; and (3) an integrated “saddle truss” and LH2

propellant drop tank assembly that connects the two elements. With a propellant capacity of ~190 t in its 

two tanks, Copernicus can achieve 1-way transit times to Mars and back ranging from ~150 to 220 days 

over a 15-year synodic cycle starting in 2033. 

 

Figure 2.—“Copernicus” crewed MTV in LEO prior to departure. 
NASA/TM—2014-218104 4 

This paper examines the possibilities for achieving short transits times by reconfiguring the 

Copernicus design to allow it to carry larger quantities of LH2 propellant needed for these more 

demanding missions. The paper covers the following topic areas. First, the operational principles and 

performance characteristics of the baseline 25 klbf Pewee-class engine used in this analysis are presented 

along with a summary of the Rover/NERVA programs’ technical accomplishments. Key NTP 

development activities currently funded under NASA’s Advanced Exploration Systems (AES) program 

are also mentioned. Mission and transportation system ground rules and assumptions used in the analysis 

are then presented along with trajectory data for a fast conjunction 2033 mission that covers a range of 

transit times. Copernicus’ key components and mission capabilities are reviewed next. To reduce 1-way 

transit times down to 130 days, then to 100 days, modifications are made to the Copernicus configuration 

incrementally—first with the addition of an in-line tank, then with the addition of paired sets of drop 

tanks—to increase the vehicle’s propellant loading. A “split crewed mission” option that does not require 

orbital assembly and uses Copernicus’ components is also described that can support a crewed mission 

with 90-day transit times. Mission scenario descriptions, key features and characteristics for five different 

vehicle configurations are presented along with the engines’ operational requirements. The paper ends 

with a summary of our findings and some concluding remarks. 

2.0 NTR System Description and Performance Characteristics 

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

damage. 

Figure 3.—Schematic of “expander cycle” NTR engine with dual LH2 turbopumps. 
NASA/TM—2014-218104 5 

The fuel elements (FE) tested in the Rover/NERVA program (Ref. 8) consisted of a “graphite matrix” 

material that contained the U-235 fuel in the form of either coated particles of uranium carbide (UC2) or 

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

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

were coated with niobium carbide (NbC) initially, then with zirconium carbide (ZrC) using a chemical 

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

helped reduce hydrogen erosion of the graphite. Individual elements were 1.32 m (52 in.) in length and 

produced ~1 MWt. 

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

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

of engines (Ref. 8). The Rover program’s Pewee engine (Ref. 8) was designed and built to evaluate higher 

temperature, longer life fuel elements and improved coatings. It set several performance records including 

the highest average fuel element exit gas temperature of ~2550 K, and the highest peak fuel temperature 

~2750 K. Other performance records included average and peak power densities in the reactor core of 

~2340 MWt/m3

 and ~5200 MWt/m3, respectively. A new CVD coating of ZrC was also introduced in 

Pewee that showed performance superior to the NbC coating used in previous reactor tests. 

In addition to FEs, the engine reactor cores also included tie tube (TT) elements of the same 

hexagonal shape that provided structural support for the FEs. A coaxial coolant tube of Inconel inside 

each TT supplied a source of heated hydrogen for turbine drive power and a sleeve of zirconium hydride 

(ZrH) moderator material could also be incorporated in the TTs to help raise neutron reactivity (shown in 

Fig. 4). In the larger size engines tested during the Rover/NERVA programs, a “sparse” FE—TT 

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

surrounding elements (Ref. 11). In this sparse pattern, the FE to TT ratio is ~3 to 1. 

In the Small Nuclear Rocket Engine (SNRE) design developed by Los Alamos National Laboratory 

near the end of Rover/NERVA (Ref. 12), shorter (0.89 m/35 in.) FEs were used so additional TTs were 

included in the reactor to increase core reactivity. With the “SNRE” FE—TT pattern each FE has three 

adjacent TTs and three adjacent FEs surrounding it (shown in Fig. 4) and the FE to TT ratio is ~2 to 1. An 

important feature common to both the sparse and SNRE FE—TT patterns is that each tie tube provides 

mechanical support for six adjacent fuel elements. 

 

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

arrangement. 

 
NASA/TM—2014-218104 6 

Recent Monte Carlo N-Particle (MCNP) transport modeling of engine reactor cores by Schnitzler et 

al., (Refs. 11, 13, and 14) has shown that the SNRE design can be scaled down to even lower thrust levels 

(~5.3 to 7.4 klbf) or up to the 25 klbf Pewee-class engine used in Mars DRA 5.0. For low thrust engines, 

short length elements (~0.89 m) and a “dense” FE—TT pattern are used consisting of parallel rows of FEs 

and TTs. Each FE has four adjacent TTs and two adjacent FEs surrounding it and the FE to TT ratio is ~1 

to 1. By using the SNRE FE—TT pattern and increasing the FE length from 0.89 to 1.32 m (the same 

length used in the Rover program’s Pewee engine), the U-235 fuel loading in the core can be lowered 

from ~0.60 to 0.25 grams/cm3

 allowing the FEs to operate at higher peak fuel temperatures (Tpeak) while 

still staying safely below the melt temperature for composite fuel of ~3050 K. It also allows higher 

hydrogen exhaust/higher chamber inlet temperatures thereby increasing the engine’s Isp capability. Higher 

Isp operation can help stretch the available LH2 propellant loading to meet mission requirements, or in the 

case of an emergency to allow a safe return of the crew. Engine and reactor performance characteristics 

for small and Pewee-class engines using composite fuel are summarized in Table 1. 

The reference NTR engine used in DRA 5.0 and this paper is a 25 klbf Pewee-class engine that is 

based on an “axial-growth” version of the SNRE. It utilizes an expander cycle and has the following 

nominal performance parameters: Tex ~2790 K, chamber pressure ~1000 psia, nozzle area ratio (NAR) 

~300:1, and Isp~906 s. The LH2 flow rate is ~12.5 kg/s and the engine thrust-to-weight ratio is ~3.50. The 

overall engine length is ~7.01 m, which includes an ~2.16 m long, retractable radiation-cooled nozzle 

skirt extension. The corresponding nozzle exit diameter is ~1.87 m. The higher chamber pressure helps to 

maintain reasonable nozzle dimensions at the assumed NAR. The engine’s reactor core contains 564 FEs 

and 241 TTs that are 1.32 m long like those used in Pewee. The core power level and fuel matrix power 

density are ~560 and ~3.44 MWt/liter, respectively. The U-235 fuel loading used in the reactor FEs is 

~0.25 grams/cm3 and the inventory of 93 percent enriched U-235 in the core is just under 37 kg. At this 

fuel loading the engine is able to operate at a Tpeak of ~2860 K providing a temperature margin to fuel 

melt of ~190 K. During enhanced operation Tpeak is raised to ~3010 K (a margin-to-melt temperature of 

~40 K) and Tex to ~2940 K resulting in an ~35 s increase in Isp to ~940 s if needed. Table 1 also shows a 

“radial-growth” version of a Pewee-class engine using shorter FEs and a sparse FE—TT pattern for 

comparison. 

TABLE 1.—PERFORMANCE CHARACTERISTICS FOR SMALL AND FULL SIZE “COMPOSITE” FUEL NTR ENGINES 
NASA/TM—2014-218104 7 

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

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

(Ref. 8) and withstood peak power densities of ~4500 to 5000 MWt/m3. They also demonstrated better 

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

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

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

helping to reduce coating cracking. Electrical-heated composite fuel elements were also tested by 

Westinghouse in hot hydrogen at 2700 K for ~600 min—equivalent to ten 1-hr cycles. At the end of 

Rover/NERVA program, composite fuel performance projections (Ref. 15) were estimated at ~2 to 6 hr at 

full power for hydrogen exhaust temperatures of ~2500 to 2800 K. 

In addition to these carbide-based fuels, a ceramic-metallic or “cermet” fuel consisting of uranium 

dioxide (UO2) in a tungsten (W) metal matrix was developed in the GE-710 and ANL nuclear rocket 

programs (Refs. 16 and 17) as a backup to the Rover/NERVA fuel. While no integrated reactor/engine 

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

and irradiation testing with promising results. Both fuel options are under development today in NASA’s 

Advanced Exploration Systems (AES) program and the Nuclear Cryogenic Propulsion Stage (NCPS) 

project (Ref. 18) which began in FY’12. This 3-year project is a collaboration between NASA and DOE 

involving five key task activities that include state-of-the-art engine modeling, mission design and 

requirements definition, fuel element fabrication and non-nuclear testing, and analysis of affordable 

options for nuclear ground testing and engine development. If successfully completed by the end of 

FY’14, a NCPS-Phase II effort would begin that could culminate in ground testing a small NTR engine at 

the Nevada Test Site (NTS) in the early 2020’s, followed by a flight demonstration mission several years 

later (Ref. 19). 

Finally, the motivation for selecting NTP as the propulsion system of choice for Mars is simple—it is 

a proven technology with a specific impulse that is twice that of today’s best chemical rockets. During the 

Rover/NERVA programs (1955 to 1972), a technology readiness level (TRL ~5 and 6) was achieved 

(Ref. 8). Twenty rocket reactors were designed, built and ground tested in integrated reactor/engine tests 

that demonstrated: (1) a wide range of thrust levels (~25, 50, 75, and 250 klbf); (2) high temperature 

carbide-based nuclear fuels that provided hydrogen exhaust temperatures up to 2550 K (achieved in 

Pewee); (3) sustained engine operation (over 62 min for a single burn achieved in the NRX-A6); as well as 

(4) accumulated lifetime at full-power; and (5) restart capability (>2 hr with 28 startup and shutdown 

cycles achieved in the NRX-XE experimental engine)—all the requirements needed for short transit time 

Mars missions. Just as important, NTP requires no large scale-ups in size or performance that are required 

with other advanced propulsion options. A cluster of three to four Pewee-class engines are more than 

adequate. 

3.0 V Budgets, Mission and Transportation System Ground Rules and 

Assumptions 

The fast conjunction 2033 opportunity was selected as the baseline mission for this analysis since 

missions performed in the 2033 to 2035 timeframe have V budgets near their minimum values over the 

15-year synodic cycle. As a result, it is expected that the shortest achievable transit times would be found 

in this opportunity. Table 2 shows the ideal V budgets as a function of decreasing transit time for the 

outbound Earth to Mars mission leg. It is assumed that the crewed MTV departs from a 407 km circular 

LEO, then captures into and departs from a 250 by 33,793 km elliptical Mars orbit (EMO) with a 24-hr 

orbital period. Gravity losses are also added to the ideal Vs and depend on the particular mission C3, 

vehicle thrust-to-weight (F/W), and propulsion system specific impulse. 

 
NASA/TM—2014-218104 8 

TABLE 2.—TRAJECTORY DATA FOR 2033 FAST CONJUNCTION MISSION—OUTBOUND EARTH-TO-MARS LEG 

TABLE 3.—TRAJECTORY DATA FOR 2033 FAST CONJUNCTION MISSION—MARS-TO-EARTH RETURN LEG 

Table 3 shows the ideal V budgets for the same range of transit times for the inbound Mars to Earth 

mission leg with Mars departure occurring in late 2035. The last column in the table shows the 

corresponding stay time at Mars assuming equal outbound and inbound transit times. As transit times 

decrease, it becomes necessary to perform a slowdown burn on the return leg to ensure that arrival Vinf at 

Earth is less than 6.81 km/s at an altitude of ~125 km. 

For 180-day transits to and from Mars and a 575-day stay time at Mars, the total round trip time and 

V budget (minus g-loss) for the mission are 935 days and 6.045 km/s, respectively. Cutting the transit 

times in half to 90-days increases the Mars stay time to ~673 days, but decreases the round trip time to 

~853 days. The mission V budget however, increases by ~250 percent to ~15.175 km/s! While 

challenging, transit times this short are possible using NTP. 

The mission and transportation system ground rules and assumptions used in this paper are 

summarized in Tables 4 and 5, respectively. Table 4 provides information about the mission profiles and 

operational scenarios, and the assumed parking orbits at Earth and Mars. In addition to the large V 

requirements for the primary propulsion maneuvers—trans-Mars injection (TMI), Mars orbit capture 

(MOC), and trans-Earth injection (TEI)—smaller V maneuvers are needed for rendezvous and docking 

(R&D) of vehicle components during the LEO assembly phase, for spacecraft attitude control during inspace transit, and for “split crewed mission” orbital operations at Mars. 
NASA/TM—2014-218104 9 

TABLE 4.—MISSION AND PAYLOAD GROUND RULES AND ASSUMPTIONS 

TABLE 5.—NTR TRANSPORTATION SYSTEM GROUND RULES AND ASSUMPTIONS 
NASA/TM—2014-218104 10 

For the crewed mission, the outbound payload mass varies with the crew size and mission duration. 

For long surface stay Mars missions, the crewed MTV carries contingency consumables equivalent to that 

found on the habitat lander. This allows the crew MTV to function as an orbital “safe haven” in the event 

of a major failure of a key surface system. For the nominal surface mission, the contingency consumables 

are jettisoned prior to the TEI maneuver. Fixed mass payload elements include the Orion Multi-Purpose 

Crew Vehicle (MPCV). For Mars DRA 5.0 and in this analysis it is assumed that ~250 to 500 kg of 

samples are returned. 

Table 5 lists the key transportation system ground rules and assumptions. The NTR engine and fuel 

type, operating characteristics, and thrust levels examined are summarized first. All engines use 

composite fuel with a U-235 fuel loading of 0.25 g/cm3

. With Tex ~2790 K and NAR ~300:1, the Isp is 

~906 s with higher Isp values achievable by increasing the fuel operating temperature. The total mission 

LH2 propellant loading consists of the usable propellant plus performance reserve, post-burn engine 

cooldown, and tank-trapped residuals. For the smaller auxiliary maneuvers, a storable bipropellant RCS 

system is used. All transfer vehicle configurations utilize a “split RCS” with 16 of 32 AMBR thrusters 

and approximately half of the bipropellant mass located on the rear propulsion stage and the forward most 

saddle truss adaptor ring just behind the TransHab module. 

The LH2 propellant carried by the various vehicles is stored in the same “state-of-the-art” Al/Li LH2

propellant tank being developed for the SLS/HLV that will support future human exploration missions. 

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

safety factor of 1.5. A 3 percent ullage factor is also assumed. All tanks use a combination foam/ 

multilayer insulation (MLI) system for passive thermal protection. A zero boil-off (ZBO) “reverse turboBrayton” cryocooler system is used on the core NTPS and “in-line” LH2 tanks (where required) to 

eliminate boil-off during LEO assembly and the remainder of the mission which can be as long as 

950 days in duration. The propellant tank heat load is largest in LEO and sizes the ZBO cryocooler 

system. Solar photovoltaic arrays are baselined for supplying the primary electrical power needed for all 

key transfer vehicle subsystems. Because of the decreased solar radiation at Mars (~486 W/m2

), array 

areas can become quite large (~10 m2

/kWe) necessitating multiple arrays for human Mars missions. 

Table 5 also includes the “dry weight contingency” (DWC) factors, along with the lift and payload 

envelope requirements for the “upgraded” SLS/HLV. A 30 percent DWC is used on the NTR system and 

advanced composite structures (e.g., stage adaptors, trusses) and 15 percent on heritage systems 

(e.g., Al/Li tanks, RCS, etc.). The NTPS and crewed payload element (PL) determine the required lift and 

PL envelope. The optimum tank diameter and mass for a Mars NTPS are 10 m and ~140 t (Ref. 9). The 

crewed PL element includes the “packaged” TransHab module with PVA power system, the short saddle 

truss, consumables container and transfer tunnel with secondary docking module (DM), and the Orion 

MPCV (shown Figs. 2 and 6). The PL envelope’s diameter is ~11 m (the saddle truss outer dimension) 

and its length can be up to ~33.8 m (the DRA 5.0 PL length including the Orion MPCV). 

4.0 Mars DRA 5.0: “7-Launch” NTR Mission Overview 

The 7-Launch NTR Mars mission strategy (Ref. 9) for a 2033 human landing mission is illustrated in 

Figure 5. It assumes a long surface stay, split cargo/piloted mission approach. Two cargo flights predeploy a cargo lander to the surface and a habitat lander into Mars orbit where it remains until the arrival 

of the crew on the next mission opportunity (~26 months later). The cargo flights utilize “1-way” 

minimum energy, long transit time trajectories. Four HLV flights carried out over 90 days (~30 days 

between launches), deliver the required components for the two cargo vehicles. The first two launches 

deliver the NTR propulsion stages each with three 25 klbf NTR engines. The next two launches deliver 

the cargo and habitat lander payload elements which are enclosed within a large triconic-shaped aeroshell 

that functions as a payload shroud during launch, then as an aerobrake and thermal protection system 

during Mars aerocapture (AC) and subsequent entry, descent and landing (EDL) on Mars. Vehicle 

assembly involves Earth orbit rendezvous and docking (R&D) between the propulsion stages and payload 

elements with the NTR stages functioning as the active element in the R&D maneuver. 
NASA/TM—2014-218104 11 

Figure 5.—DRA 5.0 long-stay Mars 2033 mission overview: “7-launch” NTR strategy. 

Once the operational functions of the orbiting habitat and surface cargo landers are verified, and the 

Mars Ascent Vehicle (MAV) is supplied with ISRU-produced ascent propellant, the crewed MTV is 

readied and departs on the next mission opportunity. The baseline Copernicus design is capable of 1-way 

transit times ranging from ~150 to 220 days depending on the particular opportunity. For the 2033 

mission, 150-day transit times are possible. Like the cargo MTVs, Copernicus is assembled in LEO using 

Earth orbit R&D. It uses the same “common” NTPS but includes additional external radiation shielding 

on each engine for crew protection during engine operation. Three HLV launches over 60 days are used to 

deliver the vehicle’s key elements which include: (1) the NTPS; (2) the integrated “saddle truss” and LH2

drop tank assembly; and (3) the crewed payload. The payload element includes the TransHab module 

with its six crew, the Orion MPCV for vehicle-to-vehicle transfer and “end of mission” re-entry, a 

secondary T-shaped DM, a contingency consumables container and connecting structure. Four 12.5 kWe/ 

125 m2

 rectangular PVAs provide ~50 kWe of electrical power at Mars for crew life-support (~30 kWe), a 

ZBO Brayton cryocooler system (~10 kWe), and high data- rate communications (~10 kWe) with Earth. 

When assembly is completed, the Mars crew is delivered to LEO on either the SLS or a commercial 

launch vehicle and docks with Copernicus on its underside using the secondary DM that connects the 

TransHab crew module and contingency consumables container (shown in Figs. 6 and 7). 

Following the TMI maneuver, the drained LH2 drop tank, attached to the saddle truss, is jettisoned 

and the crewed MTV coasts to Mars under 0-gE conditions with its four PVAs tracking the Sun. Attitude 

control and mid-course correction maneuvers are provided by Copernicus’ split RCS that uses 200 lbf

storable bipropellant Advanced Material Bipropellant Rocket (AMBR) thrusters located on the rear NTPS 

and the short saddle truss forward adaptor ring just behind the TransHab module. After the MOC burn, 

Copernicus rendezvous with the orbiting Hab lander using engine cool-down thrust and the vehicle’s 

RCS. The crew then transfers over to the lander using the MPCV. After crew transfer, the MPCV returns 

and docks to the TransHab autonomously. The crew then initiates EDL near the cargo lander and begins 

the surface exploration phase of the mission. After ~592 days on the surface, the crew lifts off using the 
NASA/TM—2014-218104 12 

Figure 6.—Copernicus’ secondary DM provides access to the MPCV and MAV during the mission. 

MAV and returns to Copernicus using its secondary DM (shown in Fig. 6). Following the transfer of the 

crew and samples, the MAV is jettisoned. The crew then begins a weeklong checkout and verification of 

all MTV systems, jettisons the DM and contingency consumables and performs the TEI burn to begin the 

journey back to Earth. After a 150-day return trip, the crew enters the MPCV, separates from the MTV 

and re-enters Earth’s atmosphere while Copernicus flies by Earth at a “sufficiently high altitude” and is 

disposed of into heliocentric space. Although Copernicus was operated in an “expendable mission mode” 

in DRA 5.0 to reduce total IMLEO and number of HLV launches, it can readily be modified to operate in 

a “reuse mode” by providing the vehicle with additional propellant capacity as discussed elsewhere 

(Ref. 10). 

The Copernicus crewed MTV has an overall length of ~93.7 m (Fig. 7) and an IMLEO of ~331.7 t for 

the 150-day transit 2033 mission. Included are (1) the NTPS (~138.8 t); (2) the saddle truss and LH2 drop 

tank (~125.6 t); and (3) the crew payload section (~67.3). The NTPS uses a three-engine cluster of 25 klbf

NTR engines and also carries additional external radiation shield mass (~6 t) for crew protection. The 

NTPS uses an Al/Li LH2 tank size which has a diameter (D) and length (L) of 10 m D by 19.7 m L. The 

LH2 tank has a propellant capacity of ~87.2 t but is slightly off-loaded for this particular mission and 

carries ~86.8 t. The NTPS also carries avionics, RCS, auxiliary battery and PVA power, docking and 

Brayton-cycle ZBO refrigeration systems located in the forward cylindrical adaptor section. To remove 

~78 W of heat penetrating the 60 layer MLI system in LEO (where the highest tank heat flux occurs), the 

2-stage cryocooler system requires ~8.9 kWe for its operation. Twin circular PVAs on the NTPS provide 

the electrical power for the ZBO system in LEO until the four primary PVAs on the crewed PL section 

are deployed prior to TMI. 

Copernicus’ second major component is its saddle truss and LH2 drop tank assembly. The saddle 

truss is a rigid, spine-like composite structure that wraps around the upper half of the LH2 drop tank and 

connects the NTR stage to the forward payload section. It is ~27.7 m long and has a mass of ~9 t. The 

saddle truss is open underneath allowing the drained LH2 drop tank to be jettisoned after the TMI burn is 

completed. The ~22.7 m long LH2 drop tank has a mass of ~21 t and a propellant capacity of ~102.4 t but 

is also off-loaded and only carries ~91.9 t of LH2 propellant. 
NASA/TM—2014-218104 13 

Figure 7.—Key features and component lengths of the Copernicus Mars transfer vehicle. 

Copernicus’ third and final component is its payload section. In DRA 5.0 it was designed for launch 

as a single integrated unit and thus determines the overall size of the PL envelope. The integrated payload 

element is ~33.8 m long and includes the short saddle truss, “T-shaped” DM and transfer tunnel, 

consumables container, TransHab and crew, and the Orion MPCV. The DM provides “secondary access” 

to Copernicus for the crew delivery MPCV and the MAV (see Figs. 6 and 7). Following the crew’s return 

from Mars and MAV separation, the DM and attached consumables container are both jettisoned to 

reduce vehicle mass prior to TEI (see Fig. 5). 

The total crewed payload mass at TMI is ~67.3 t distributed as follows: (1) short saddle truss (~5.1 t); 

(2) DM and transfer tunnel (~2.8 t); (3) contingency consumables and jettisonable container (~10.7 t); (4) 

TransHab with its primary PVAs (~27.5 t); (5) transit consumables (~4.6 t); (6) crew (~0.6 t); (7) MPCV 

(~10 t); and (8) forward RCS and propellant (~6 t). Copernicus’ total RCS propellant loading is ~11.9 t 

with the “post-TMI” RCS propellant load split between the NTPS (~7 t) and the short saddle truss 

forward cylindrical adaptor ring (~4.9 t). 

Lastly, for this particular mission, the performance requirements on operating time and restart for 

Copernicus’ three 25.1 klbf NTR engines are quite reasonable. For the round trip mission, there are four 

primary burns (three restarts) that use ~168.7 t of LH2 propellant. With ~75.3 klbf of total thrust and a Isp 

of ~906 s, the total engine burn time for the mission is ~74.6 min (~53.2 min for the “2-perigee burn” 

TMI maneuver, ~13.6 min for MOC, and ~7.8 min for TEI), well under the ~2 hr accumulated engine 

burn time and 27 restarts demonstrated by the NERVA eXperimental Engine—the NRX-XE (Ref. 8). 
NASA/TM—2014-218104 14 

5.0 Copernicus “Growth” Options Allowing Faster Transit Time Missions 

As discussed in Section 4.0, Copernicus was sized so it can perform all fast conjunction missions in 

the 2033 to 2048 timeframe. For the 2033 mission opportunity, with its propellant tanks near their 

maximum capacity of ~190 t, Copernicus can achieve 150-day transit times. Besides human Mars 

missions, the “3-element” Copernicus spacecraft (Fig. 8) has significant capability that can be exploited 

for a variety of other mission applications such as a reusable human mission to the large, high energy near 

Earth asteroid—Apophis in 2028. “Scaled down” versions of Copernicus can also be used for reusable 

lunar cargo delivery and crewed lunar landing missions. Mission details and vehicle characteristics for 

these applications are provided elsewhere (Ref. 10). 

Figure 9 shows the round trip mission V versus transit time for the 2033 opportunity using the 

trajectory data found in Tables 2 and 3. As is evident, the V requirements rise dramatically as the transit 

times are shortened. Decreasing the transit times by 60 days—from 180 to 120 days—increases the total 

mission V by ~46 percent. A further 60-day reduction in transit time—down to 90 days each way—

increases the total mission V by ~250 percent. 

To achieve transit times less than 150 days, it is necessary to modify the Copernicus spacecraft so it 

can carry additional propellant. This will be done in incremental steps—first with the addition of an inline tank, then with the addition of paired sets of drop tanks—to show what can be achieved, but also to 

show what is required in terms of extra hardware, engine operating time and additional HLV launches. 

Figure 8.—Growth paths and alternative missions for Copernicus spacecraft using modular components. 
NASA/TM—2014-218104 15 

Figure 9.—Total mission V versus transit time. 

Figure 10.—Copernicus-2 option: Supports 130-day transit times for the 2033 mission. 

5.1 Copernicus-2 Option: Addition of an In-Line Tank to Achieve 130-day Transit 

Times 

To reduce transit times by 20 days—from 150 to 130 days, the Copernicus-2 configuration positions 

an “in-line” LH2 tank between the NTPS and integrated saddle truss/drop tank assembly (shown in Figs. 8 

and 10). This configuration has an IMLEO of ~443.8 t consisting of the “wet” NTPS (~136.4 t), in-line 

tank (~118.5 t), saddle truss/drop tank assembly (~119 t) and the crew PL section (~69.9 t). The overall 

vehicle length is ~117.6 m including the Orion MPCV at ~8.9 m. The LH2 tank length in the NTPS is 

~19.7 m and it carries ~86.1 t of propellant (close to its maximum capacity of ~87.2 t). The in-line and 
NASA/TM—2014-218104 16 

drop tank lengths are the same at ~20.5 m and each tank carries ~91.5 t of LH2 propellant. With these 

equal tank lengths, the masses of in-line tank element and the saddle truss/drop tank assembly are 

balanced at ~118.8 t. 

For this 130-day transit mission, there are four primary burns (with three restarts) and ~253.8 t of LH2

propellant is used. A fourth engine is added to the NTPS to reduce the total engine operating time and 

minimize g-losses (~326 m/s) during the TMI maneuver but this increases NTPS inert mass at the 

expense of LH2 and RCS propellant loading. With ~100.4 klbf of total thrust and a Isp of 906 s, the total 

engine burn time for this mission is ~84.2 min. The first TMI perigee burn is the longest single burn at 

~35 min. After this burn, the drop tank is drained and then jettisoned to reduce vehicle mass and 

propellant consumption during the second perigee burn (~21.8 min). The in-line tank and NTPS supply 

the propellant for the remaining MOC (~17.5 min) and TEI burns (~9.9 min). 

Copernicus-2 also requires one additional HLV launch as well as the development of the in-line tank 

element with its own ZBO cryocooler system. For the longer in-line tank, the cryocooler system requires 

~9.13 kWe of electrical power supplied by PVAs on the NTPS and in-line tank. Even with a fourth engine 

added, the total burn time requirement on the engines is increased by 13 percent (from 74.6 to 84.2 min). 

An alternative to adding a fourth NTR engine would be to add oxygen “afterburner” nozzles to 

Copernicus’ three NTR engines. In the “LOX-Augmented” NTR (LANTR) option (Refs. 20 and 21), 

oxygen is injected into the divergent section of the nozzle downstream of the sonic throat. Here it mixes 

with reactor-heated H2 and undergoes supersonic combustion adding both mass and chemical energy to 

the engine’s exhaust. By operating the LANTR engines with a LOX-to-LH2 mixture ratio (MR) = 1 

during the first TMI perigee burn, the engine’s thrust level can be increased by over 62 percent—from 

25 to ~40.6 klbf. The addition of oxygen lowers the Isp to ~725 s but this is still ~270 s higher than that 

achieved by LOX/LH2 engines. 

5.2 Copernicus-3 Option: Addition of Twin Drop Tanks to Achieve 120-day Transit 

Times 

In the Copernicus-3 configuration (shown in Figs. 8 and 11), the saddle truss/drop tank assembly is 

replaced by a 4-sided “star truss” with twin LH2 drop tanks that supplement the propellant carried by the 

NTPS and in-line tank. Copernicus-3 has an IMLEO of ~551.3 t including the NTPS (~139.9 t), the inline tank (~133.8 t), the “star truss” with drop tanks (~210.3 t) and the crew PL section (~67.3 t). Two 

drop tanks (~190 t) carrying ~159.4 t of LH2 (~79.7 t per tank) are delivered to LEO on two HLV 

launches, then attached to the ~25.4 m long star truss (~20.3 t) during the LEO assembly phase. Each 

drop tank is ~18.2 m long. The overall vehicle length is ~118.2 m including the 8.9 m Orion MPCV. The 

LH2 loading in the off-loaded NTPS, in-line and twin drop tanks is ~83.1, ~98.6, and ~159.4 t, 

respectively. The lengths of NTPS and in-line LH2 tanks are ~19.7 and ~21.9 m. During the four primary 

mission burns ~321 t of LH2 propellant is used. With 100.4 klbf of total thrust and Isp ~906 s, the total 

engine burn time for this mission is ~106.4 min, close to the ~2 hr demonstrated on the NRX-XE engine. 

After completing a long first perigee burn (~52 min), both drop tanks are jettisoned. The remaining burn 

durations are as follows: second perigee burn (~20.7 min), MOC (~22.2 min) and TEI (~11.5 min). Four 

HLV launches are required to deliver the major transportation systems elements (the NTPS, in-line tank 

and two drop tanks) to LEO. An upgraded SLS is used to deliver the crewed payload element and the 

large but lightweight star truss. 

5.3 Copernicus-3a Option: Addition of Second Set of Drop Tanks to Achieve 100-day 

Transit Times 

In transitioning from 120 to 100-day transit times the total mission V increases by ~42 percent—

from ~8.83 to ~12.5 km/s. To satisfy the high-energy requirements of this mission, Copernicus-3a (shown 

in Figs. 8 and 12) adds a second set of drop tanks to the star truss. Adjustments are made to the PL as 

well. The crew size is reduced to four and the mass of the TransHab and consumables needed for the 
NASA/TM—2014-218104 17 

mission are reduced accordingly. Composite LH2 tanks are also used instead of Al/Li. Lastly, the 

operating temperature of the engines’ (UC-ZrC) graphite composite fuel is increased to achieve a specific 

impulse of ~940 s (see Table 1). The resulting vehicle has an IMLEO of ~720.7 t, which includes the 

NTPS (~137.6 t), the in-line tank (~123.7 t), the star truss with four common LH2 drop tanks (~397.5 t) 

and the crew PL section (~61.9 t). The PL element includes the Orion MPCV (~15.1 t) which carries 

additional propellant (~5.1 t) for the final Earth slowdown burn (V ~1.266 km/s) before re-entry while 

the MTV flies past Earth and is disposed of in heliocentric space. 

Figure 11.—Copernicus-3 Option: Supports 120-day transit times for the 2033 mission. 

Figure 12.—Copernicus-3a Option: Supports 100-day transit times for the 2033 mission. 
NASA/TM—2014-218104 18 

The LH2 tank lengths in the NTPS and in-line element are ~19.7 and ~21.9 m, and they carry ~87.2 

and ~95.5 t of propellant, respectively. Copernicus-3a’s four drop tanks carry ~327.8 t of propellant 

(~82 t per tank), and each tank is ~18.6 m long. The drop tanks are jettisoned in pairs after the first and 

second perigee burns to reduce vehicle mass and propellant consumption during the TMI maneuver. The 

overall vehicle length is ~117.6 m. 

The total usable propellant for the mission is ~480.7 t. With 100.4 klbf of thrust and Isp ~940 s, the 

total engine burn time is ~165.4 min. This value exceeds the ~2 hr demonstrated on the NRX-XE engine 

by ~45 min. The first perigee burn is the longest single burn at ~78.3 min and also exceeds the single burn 

duration record of ~62 min set by the NRX-A6 engine. The duration of the remaining burns are as 

follows: second perigee burn (~37.3 min), MOC (~33.3 min), and TEI (~16.5 min). Seven HLV launches 

are required to deliver the major transportation systems elements (the NTPS, in-line tank, four drop tanks, 

and PL element) to LEO. A single SLS with sufficient PL volume will be used to launch the large but 

lightweight star truss. 

5.4 Copernicus “Split Crewed Mission” Option for Achieving 100—90-day Transit 

Times 

Copernicus-3a is a large complex vehicle that requires eight launches to deliver the various elements 

to LEO. From the initial launch of the NTPS until the last drop tank is attached to the star truss (via 

orbital assembly) could take ~7 months assuming 30 days between SLS/HLV launches. Failure to deliver 

the elements on time and in the prescribed order could delay vehicle assembly and jeopardize mission 

departure within the specified launch window. 

Using the Copernicus-2 spacecraft plus its three key components—the NTPS, in-line tank and 

integrated saddle truss/drop tank assembly—now configured as an Earth Return Vehicle (ERV), a “split 

crewed mission” architecture can be utilized that can reduce vehicle size and complexity, eliminate the 

need for orbital assembly and an extra engine, and can achieve transit times as low as 90 days. Use of a 

pre-deployed ERV is not a new idea. In NASA’s DRM 1.0 study in 1993 (Ref. 22), the ERV was one of 

the key transportation system elements used in the mission architecture. 

The split crewed architecture for very fast 1-way transit missions is outlined in Figure 13. Using 

the Copernicus-2 spacecraft plus an ERV, a 2033 landing mission would require a round trip time of 

~857 days with ~667 days spent in Mars orbit and on the surface. The outbound transit time on 

Copernicus-2 is 100 days and the inbound transit time using the ERV is 90 days. In this split mission, the 

ERV is pre-deployed to Mars orbit in advance of the crew. It departs from LEO in December 2030 

(departure C3 ~10.794 km2

/s2, VTMI ~3.662 km/s), on a 283-day “minimum-energy” trajectory out to 

Mars. The ERV then arrives at Mars (arrival Vinf ~3.480 km/s) and propulsively captures into a 24-hr 

elliptical Mars orbit (~250 by 33,793 km, VMOC ~1.34 km/s) in October 2031 where it remains until the 

crewed Copernicus-2 MTV arrives during the next opportunity ~26 months later. 

Copernicus-2 departs LEO with its four crew on May 25, 2033 and arrives at Mars 100 days later. It 

then captures into the same parking orbit as the ERV on September 2, 2033. In the process, Copernicus-2

uses up ~97 percent of its available propellant. To return to Earth, Copernicus-2 rendezvous with the 

ERV and the forward crewed PL element is switched over to the ERV (shown in Fig. 13). No propellant 

transfer is required just a R&D maneuver. The reconfigured ERV then rendezvous with the orbiting Hab 

lander, followed by crew transfer, then EDL near the cargo lander at which point the surface exploration 

phase of the missions begins. After ~653 days on the surface, the crew lifts off using the MAV and 

returns to the ERV. Following the transfer of the crew and samples, the MAV is jettisoned. Before 

leaving Mars orbit, the crew finally jettisons the exterior consumables container and connecting tunnel 

from the PL element to reduce vehicle mass and propellant consumption. The ERV then performs the TEI 

burn on July 1, 2035 and begins the 90-day transfer back to Earth. On the final approach to Earth, the 

ERV does a final slowdown burn (V ~2.257 km/s) after which the crew separates and re-enters using the 

Orion MPCV capsule. The ERV flies past Earth at a sufficiently high altitude and is disposed of into 

heliocentric space. 
NASA/TM—2014-218104 19 

Figure 13.—Copernicus-2 MTV and components configured for a very fast split crewed mission. 

For the outbound mission leg, the Copernicus-2 spacecraft (Fig. 14) has an IMLEO of ~402 t 

consisting of the NTPS (~134.5 t), the in-line tank (~92.4 t), the integrated saddle truss/LH2 drop tank 

assembly (~122.8 t), and the crewed PL element (~52.3 t). The NTPS, in-line and drop tank elements 

(again using composite tanks) carry ~87.2, ~71.3, and ~95 t respectively, for a total LH2 propellant 

loading of ~253.5 t. By using two separate vehicles for the crewed mission, three engines are again 

sufficient for each. With ~75.3 klbf of total thrust, an Isp of 940 s, and ~239 t of usable propellant, the total 

engine burn time for this mission is ~109.6 min. The first TMI perigee burn is the longest at ~52.6 min. 

After this burn, the drop tank is drained and jettisoned to reduce vehicle mass and propellant consumption 

during the second perigee burn (~28.8 min). The NTPS supplies all the propellant needed for MOC 

(~28.2 min). 

The “round trip” ERV has an IMLEO of ~360.3 t consisting of the NTPS (~131.3 t), the in-line tank 

(~101.4 t) and the integrated saddle truss/LH2 drop tank assembly (~122.9 t). The composite tanks used in 

the ERV’s NTPS, in-line and drop tank elements carry ~87.2, ~77, and ~95 t respectively, for a total 

propellant loading of ~259.2 t. The ERV’s NTPS performs five primary burns (and four restarts) and its 

three engines provide ~75.3 klbf of total thrust. With ~244.3 t of usable propellant and a Isp of ~940 s, the 

total engine burn time for this mission is ~112.1 min. The first TMI perigee burn is the longest at 

~41.3 min. After this burn, the drop tank is drained and jettisoned. The durations of the remaining burns 

are as follows: second perigee burn (~13.7 min), MOC (~13 min), TEI (~28.6 min), and final slowdown 

burn (~15.5 min). The long TEI burn is attributed to the addition of Copernicus-2’s ~40.2 t crewed PL 

mass plus the high TEI V requirement (~3.075 km/s) for the 90-day transit back to Earth. The total and 

longest single engine burn times required for the ERV mission while substantial, are still below the 

capabilities demonstrated by the NRX-XE (~2 hr) and NRX-A6 engines (~62 min). 
NASA/TM—2014-218104 20 

Figure 14.—“Split Crewed Mission” Option: Supports 100 to 90-day transit times. 

Lastly, the total mission IMLEO for the two spacecraft is ~762.4 t. Seven HLV launches are required 

for the split mission option—four for the Copernicus-2 and three for the ERV—compared to eight total 

launches required for the “all up” Copernicus-3a configuration. No on-orbit assembly (e.g., attachment of 

multiple drop tanks) is required just autonomous R&D of several vehicle components. 

6.0 Summary and Conclusions 

In NASA’s recent DRA 5.0 study (NASA/SP—2009-566), the fast conjunction long surface stay 

mission was identified as the preferred approach for exploring Mars. Plenty of time for scientific 

investigation and study, coupled with short ~6-month transit times to reduce crew exposure to the 0-gE

and radiation environments of space were key reasons for this selection. Similarly, NTP was identified as 

the preferred propulsion option because of its high performance (100 percent higher Isp than LOX/LH2

chemical propulsion), its demonstrated technology record, plus the strong synergy that exists between it 

and chemical propulsion systems and stages. Recent measurements of the energetic particle radiation 

environment inside MSL spacecraft during it journey out to Mars indicate that astronauts could receive a 

radiation dose of ~0.66 Sv (~66 rem)—the limiting value established by NASA—during a 1-year journey 

out to Mars and back. With the potential for the crew to receive additional dose during the exploration 

phase of the mission, several questions immediately arise: Can NTP’s performance capability be used to 

reduce transit times further and by how much? Also, what are the engine operating requirements and 

vehicle design impacts? 

The answer to the first question is yes, and analysis presented here indicates transit time reductions as 

much as 50 percent are possible. The key requirements are more propellant and reduced U-235 loading in 

the engine’s reactor fuel elements to allow higher operating temperatures. For the 2033 mission 

opportunity baselined in this paper, Copernicus can achieve 1-way transit times of ~150 days when filled 

to its maximum propellant capacity of ~190 t. To shorten transit times further, the Copernicus spacecraft 

can be reconfigured in incremental steps that allow it to carry larger quantities of propellant needed for 
NASA/TM—2014-218104 21 

these more demanding missions. To reduce the 1-way transit times from 150 to 130 days, an in-line tank 

is inserted between Copernicus’ NTPS and saddle truss/drop tank assembly that increases the vehicle’ 

propellant capacity by ~42 percent to ~270 t. A fourth engine is also added to the NTPS to reduce the 

total engine operating time and minimize g-losses. 

To achieve transit times in the range of 120 to 100 days, Copernicus’ saddle truss/drop tank assembly 

is replaced by a 4-sided star truss to which paired sets of LH2 drop tanks are attached as needed. For 120-

day transits to Mars and back, twin drop tanks—each carrying ~80 t of propellant—are used. Further 

reductions in transit times from 120 to 100 days requires an ~42 percent increase in total mission V—

from ~8.83 to ~12.5 km/s. To compensate, the crew size is reduced to four, composite propellant tanks 

are utilized and the engines are operated at their maximum temperature capability to achieve a specific 

impulse of ~940 s. Four drop tanks—each carrying ~82 t of propellant—are also used. The “all up” 

vehicles for 120- and 100-day cases require 5 and 7 HLVs respectively, as well as on-orbit assembly. The 

engine operational requirements for the 100-day case also exceed those demonstrated in the NERVA 

program. 

A “split crewed mission architecture” is outlined that uses the Copernicus-2 spacecraft configuration 

plus an ERV. It can reduce vehicle size and complexity, eliminate the need for orbital assembly and an 

extra engine, and can achieve transit times as low as 90 days. Finally, it is important to note that both the 

total and longest single engine burn times required for the split mission, while substantial, are still below 

the capabilities demonstrated by the NRX-XE (~2 hr) and NRX-A6 engines (~62 min). 

References 

1. “Human Exploration of Mars Design Reference Architecture 5.0,” Drake, Bret G., ed., National 

Aeronautics and Space Administration, NASA-SP-2009-566, Washington, DC, July 2009. 

2. “Reference Mission Version 3.0 Addendum to the Human Exploration of Mars: The Reference 

Mission of the NASA Mars Exploration Study Team,” B.G. Drake, ed., Johnson Space Center, 

Exploration Office Document EX13-98-036, June 1998. 

3. Borowski, Stanley K., Dudzinski, Leonard A., McGuire, Melissa L., “Vehicle and Mission Design 

Options for the Human Exploration of Mars/Phobos Using Bimodal NTR and LANTR Propulsion,” 

AIAA-98-3883, July 1998. 

4. Borowski, Stanley K., Dudzinski, Leonard A., and McGuire, Melissa L., “Artificial Gravity Vehicle 

Design Option for NASA’s Human Mars Mission Using Bimodal NTR Propulsion,” AIAA-99-2545, 

June 1999. 

5. Borowski, Stanley K., Dudzinski, Leonard A., and McGuire, Melissa L., “Bimodal” Nuclear 

Thermal Rocket (NTR) Propulsion for Power-Rich, Artificial Gravity Human Exploration Missions 

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Nuclear Thermal Propulsion (NTP): A Proven, Growth Technology for “Fast Transit”

Human Missions to Mars

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Borowski, Stanley, K.; McCurdy, David, R.; Packard, Thomas, W.

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WBS 279585.04.01.22

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The “fast conjunction” long surface stay mission option was selected for NASA’s recent Mars Design Reference Architecture (DRA) 5.0 study because it provided adequate

time at Mars (~540 days) for the crew to explore the planet’s geological diversity while also reducing the “1-way” transit times to and from Mars to ~6 months. Short transit

times are desirable in order to reduce the debilitating physiological effects on the human body that can result from prolonged exposure to the zero-gravity (0-gE) and radiation

environments of space. Recent measurements from the RAD detector attached to the Curiosity rover indicate that astronauts would receive a radiation dose of ~0.66 Sv (~66

rem)-the limiting value established by NASA-during their 1-year journey in deep space. Proven nuclear thermal rocket (NTR) technology, with its high thrust and high specific

impulse (Isp ~900 s), can cut 1-way transit times by as much as 50 percent by increasing the propellant capacity of the Mars transfer vehicle (MTV). No large technology

scale-ups in engine size are required for these short transit missions either since the smallest engine tested during the Rover program-the 25 klbf

 “Pewee” engine is sufficient

when used in a clustered arrangement of three to four engines. The “Copernicus” crewed MTV developed for DRA 5.0 is a 0-gE design consisting of three basic components:

(1) the NTP stage (NTPS); (2) the crewed payload element; and (3) an integrated “saddle truss” and LH2

 propellant drop tank assembly that connects the two elements. With a

propellant capacity of ~190 t, Copernicus can support 1-way transit times ranging from ~150 to 220 days over the 15-year synodic cycle. The paper examines the impact on

vehicle design of decreasing transit times for the 2033 mission opportunity. With a fourth “upgraded” SLS/HLV launch, an “in-line” LH2

 tank element can be added to

Copernicus allowing 1-way transit times of 130 days. To achieve 100 to 120 day transit times, Copernicus’ saddle truss/drop tank assembly is replaced by a “star truss”

assembly with paired modular drop tanks to further increase the vehicle’s propellant capacity. The HLV launch count increases (from ~5 to 7) and a fourth engine is needed to

reduce total mission burn time and gravity losses. Using a “split mission” approach, the NTPS, in-line tank and the saddle truss/LH2

 drop tank elements can be configured as a

pre-deployed Earth Return Vehicle/propellant tanker supporting 90-day crewed mission transits. The split mission approach also eliminates the need for on-orbit assembly.

Mission scenario descriptions, key features and operational characteristics for five different vehicle configurations are presented.

15. SUBJECT TERMS

Nuclear thermal rocket (NTR); Mars transfer vehicle (MTV); Design reference architecture (DRA); Composite fuel

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