Laura M. Burke and Stanley K. Borowski

Glenn Research Center, Cleveland, Ohio

David R. McCurdy and Thomas W. Packard

Vantage Partners, LLC, Brook Park, Ohio

A One-Year, Short-Stay Crewed Mars Mission Using 

Bimodal Nuclear Thermal Electric Propulsion 

(BNTEP)-A Preliminary Assessment 

NASA/TM—2013-216568

September 2013
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Laura M. Burke and Stanley K. Borowski

Glenn Research Center, Cleveland, Ohio

David R. McCurdy and Thomas W. Packard

Vantage Partners, LLC, Brook Park, Ohio

A One-Year, Short-Stay Crewed Mars Mission Using 

Bimodal Nuclear Thermal Electric Propulsion 

(BNTEP)-A Preliminary Assessment 

NASA/TM—2013-216568

September 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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Acknowledgments

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subject to revision as analysis proceeds.
A crewed mission to Mars poses a significant challenge in dealing with the physiological issues that arise with the crew being exposed to a near zero-gravity environment as

well as significant solar and galactic radiation for such a long duration. While long surface stay missions exceeding 500 days are the ultimate goal for human Mars exploration,

short round trip, short surface stay missions could be an important intermediate step that

would allow NASA to demonstrate technology as well as study the physiological effects

on the crew. However, for a 1-year round trip mission, the outbound and inbound hyperbolic velocity at Earth and Mars can be very large resulting in a significant propellant

requirement for a high thrust system like Nuclear Thermal Propulsion (NTP). Similarly,

a low thrust Nuclear Electric Propulsion (NEP) system requires high electrical power levels (10 megawatts electric (MWe) or more), plus advanced power conversion technology

to achieve the lower specific mass values needed for such a mission. A Bimodal Nuclear

Thermal Electric Propulsion (BNTEP) system is examined here that uses three high thrust

Bimodal Nuclear Thermal Rocket (BNTR) engines allowing short departure and capture

maneuvers. The engines also generate electrical power that drives a low thrust Electric

Propulsion (EP) system used for efficient interplanetary transit. This combined system

can help reduce the total launch mass, system and operational requirements that would

otherwise be required for equivalent NEP or Solar Electric Propulsion (SEP) mission. The

BNTEP system is a hybrid propulsion concept where the BNTR reactors operate in two

separate modes. During high-thrust mode operation, each BNTR provides 10’s of kiloNewtons of thrust at reasonably high specific impulse (Isp) of 900 seconds for impulsive

trans-planetary injection and orbital insertion maneuvers. When in power generation /

EP mode, the BNTR reactors are coupled to a Brayton power conversion system allowing

each reactor to generate 100’s of kWe of electrical power to a very high Isp (3000 s) EP

thruster system for sustained vehicle acceleration and deceleration in heliocentric space.

Nomenclature

Isp Specific impulse, s

g Earth standard gravity, 9.80665 m/s2

V∞ Hyperbolic Excess Velocity, km/s

C3 Characteristic energy, km2/s2

∆V (Delta-V) Imparted Velocity Change km/s

T /W Ratio of Thrust to Weight

BNT EP Bimodal Nuclear Thermal Electric Propulsion

BNT R Bimodal Nuclear Thermal Rocket

BRU Brayton Rotating Unit

A One-Year, Short-Stay Crewed Mars Mission Using 

Bimodal Nuclear Thermal Electric Propulsion 

(BNTEP)-A Preliminary Assessment 

Laura M. Burke and 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 

NASA/TM—2013-216568 1
DRA Design Reference Architecture

EOI Earth Orbit Insertion

EP Electric Propulsion

IMLEO Initial Mass in Low Earth Orbit

LEO Low Earth Orbit

MLI Multilayer Insulation

MOI Mars Orbit Insertion

MP CV Multi-Purpose Crew Vehicle

mt Metric Ton (1 mt = 1000 kg)

MT V Mars Transfer Vehicle

NAIF Navigation and Ancillary Information Facility

NEP Nuclear Electric Propulsion

NT R Nuclear Thermal Rocket

P CU Power Conversion Unit

PMAD Power Management and Distribution

RCS Reaction Control System

SEP Solar Electric Propulsion

SLS Space Launch System

T EI Trans-Earth Injection

TMI Trans-Mars Injection

ZBO Zero Boil Off

I. Introduction

A crew health and performance assessment was done for the Mars DRA 5.0 concluded that for both

conjunction and opposition class missions, exposure to solar proton events and galactic cosmic radiation

during interplanetary transit and solar proton event effects during close perihelion passage are well outside

current permissible exposure limits.1 Limiting total mission time to 1-year can reduce interplanetary time

by 75-300 days, significantly reducing exposure time.

The analysis presented in this paper covers a detailed trajectory design and mission architecture for a

2033 one-year round trip, 1 MWe-class BNTEP mission with an outbound transit time of 126 days, a 30

day stay in a 24 hour elliptical Mars orbit and an inbound transfer time of 209 days with direct re-entry at

Earth by the crewed MPCV. The dual-mode BNTEP vehicle enables a short round trip mission by reducing

the use of the high thrust BNTR engines. Having an initial mass in LEO of 389.9 mt, this mission requires

a total of 5 SLS Block-IB launches to deliver all the vehicle components for assembly in LEO. This is on-par

with the stay time and IMLEO for NT R opposition-class mission and accomplishes the mission in 365 days

instead of 400-600 days. Descriptions of the mission, propulsion and vehicle needed for this type of mission

are included in the following text. A conceptual design of the vehicle and major components is provided in

Figures 1 and 6.

Figure 1. Key Components of BNTEP Mars Transfer Vehicle

NASA/TM—2013-216568 2
II. 2033 1-Year Earth-Mars Roundtrip BNTEP Mission

II.A. Mission Definition

A 1-year mission during the 2033 opportunity requires a lower combined BNTR and low-thrust ∆V budget

than other opportunities in the 2031-2041 time frame (Appendix A). Although 2033 is not the preferred

mission opportunity for ballistic transfer NTR missions due to the high launch asymptote declination, the

compressed mission timeline causes the launch date to move later than the ballistic optimal and out of the

higher declination region. Plane change to adjust for declinations beyond 28.5

◦

is performed during the

low-thrust burns during the outbound transit.

This mission uses the BNTR engines in high-thrust mode to depart Earth on May 6, 2033. Following

injection into the hyperbolic Earth escape trajectory, the BNTR engines are then powered down to powergeneration mode to complete low-thrust EP burns during transit. The BNTR engines are powered up again

just prior to Mars arrival for a high-thrust burn to injection into a 24 hr Mars orbit on September 8, 2033.

The MTV stays in orbit around Mars for 30 days at which point the BNTR engines perform the last highthrust burn to escape Mars. During the return leg of the trajectory the BNTR engines are switched to

power-generation mode again to complete the return low-thrust EP burns. Upon arrival at Earth on May

6, 2034, the crew boards the MPCV, separates from the MTV and performs a direct re-entry into Earth’s

atmosphere. Over the course of the mission, the total time the BNTR engines are operating in high thrust

mode is 1.75 hours.

II.B. Mission Design Analytical Methods

The trajectory optimization program COPERNICUS2 was used to model and perform analysis for the 1-year

BNTEP missions. Impulsive burns were used to model the BNTR propulsive maneuvers and finite burns

were used to model the longer EP burns. Trajectories were optimized for the 2031-2041 opportunities. Earth

departure and arrival dates were optimized parameters but the total time of flight was constrained to be

no longer than 365 days. The MTV begins in an 407 km circular LEO orbit inclined at 28.5

◦

. The BNTR

engines perform TMI to put the MTV into a hyperbolic Earth-Mars transfer orbit. A minimum half-day

coast is constrained to occur before and after each BNTR burn to allow for targeting and checkout operations. Following TMI and a half-day checkout period, the optimization is allowed to initiate the first EP

burn. This burn usually occurs as early as possible in the transfer trajectory in order to reduce the burden

on the BNTR system to provide all the energy required for Earth departure. A coast period separates this

first EP burn from a second EP burn that occurs prior to Mars arrival. The second EP burn is terminated

no later than a half-day prior to MOI. The second EP burn is used to slow down the MTV to reduce the

MOI ∆V . MOI is performed using the BNTR engines to insert the MTV into a 300 km x 24 hr Mars orbit.

The stay time at Mars is constrained to be a minimum of 30 days. Due to the geometry of this mission,

stay time never extends past the 30 day lower bound. To depart Mars, the BNTR engines perform the

TEI burn to put the MTV into a hyperbolic Mars-Earth transfer orbit. After a half-day checkout coast

following TEI the optimization is allowed to initiate the third EP burn. A coast period separates this third

EP burn from a fourth EP burn. Typically the fourth EP burn is removed, as it is deemed unnecessary by

the optimizer. Unlike the other EP burns which occur as close to departure/arrival as possible, the fourth

EP burn, if performed, occurs near the mid-point of the inbound trajectory. Since the MTV employs a direct

entry at Earth rather than EOI to capture into an Earth orbit, the optimization uses the fourth EP burn to

shape/correct the trajectory to return to Earth within 365 days.

Similar to ballistic Earth-Mars trajectories, optimal BNTEP Earth-Mars trajectories can have outbound

legs that are either Type-I or Type-II trajectories. Type-I trajectories are characterized as having shorter

trip times while Type-II trajectories are characterized as having longer trip times. Type-I and Type-II

trajectories are identified as having heliocentric travel angles less than or greater than 180◦respectively.

While 2031, 2033 and 2035 are more optimal with Type-I outbound transfers, 2037, 2039 and 2041 are more

optimal having a Type-II outbound trajectory leg. For this analysis, only the optimal trajectory type was

analyzed for each opportunity.

NASA/TM—2013-216568 3
Figure 2. Ecliptic View of 2033 Reference 1-Year BNTEP Earth-Mars Roundtrip Mission

An ecliptic view of a 2033 BNTEP trajectory is showing in Figure 2. The green segments represent

coasting phases while the red segments represent EP burns as well as the thrust direction during those

burns. It should also be noted that although a minimum Spacecraft-Sun distance was not specified as a

constraint, for all trajectories analyzed for this study the MTV did not approach closer than 0.7 AU to the

Sun (Venus’ orbit) and that these trajectories were modeled without requiring a Venus flyby on the outbound

or return legs.

II.B.1. Analysis Tools

COPERNICUS was used to model the Earth-Mars BNTEP trajectories and perform the analysis presented in

this paper. COPERNICUS is a 3-Degrees of Freedom spacecraft trajectory design and optimization program

that was originally developed at the University of Texas with recent developments and maintenance being

done by Johnson Space Center.2 Trajectory models are built as a series of segments propagated in either

a point-mass or higher order gravitational field. The segments can be connected and constrained with the

addition of linear or non-linear constraints. COPERNICUS has the capability to model both low thrust

burns as well as high thrust burns as either impulsively or with finite maneuvers for propulsion systems

which can be based on BNTR, solar electric, or nuclear powered engines. SPICE files provided by Jet

Propulsion Laboratory’s NAIF is the basis of coordinate systems and ephemeris data.

II.C. Assumptions and Constraints

• TOF = 365 days

• Mars Stay Time = 30 days

• EP system efficiency = 65%

• Earth departure orbit: 407 km circular orbit inclined at 28.5

◦

• Mars arrival orbit: 300 km x 24 hr elliptical orbit

• Maximum Earth arrival V∞ = 11.5 km/s

• EP system Isp = 3000 s

• NTR Isp = 900 s

NASA/TM—2013-216568 4
• Propellant Constraints

– Maximum drop tank propellant capacity is 39.2 mt (8.4 m outer diameter x 13 m length drop

tank)

– The TMI-1 burn is constrained to use no more than the full propellant capacity of 2 drop tanks

– BNTR propellant required for TMI-2 beyond the capacity of the remaining 2 drop tanks will be

stored in the inline tank

• Dropped Elements

– Individual drop tank sets released after TMI-1 and TMI-2

II.D. Trajectory Details

Trajectory analysis for the 2033 Opportunity was completed to determine the variations in BNTR and EP

∆V with the MTV IMLEO. As expected, as the IMLEO of the MTV increases, the effectiveness of the EP

system to move the vehicle during the short transit time decreases. Hence, the amount of ∆V imparted by

the EP system decreases since the high-thrust BNTR engines must be utilized to a greater degree in order

to accomplish the mission in the constrained 1-year time period. Plots showing the relationship between

IMLEO and ∆V are shown in Figure 3 and Figure 4.

8.6$

8.7$

8.8$

8.9$

9$

9.1$

9.2$

9.3$

9.4$

9.5$

300000$ 325000$ 350000$ 375000$ 400000$ 425000$ 450000$

Total&BNTR&Delta,V&(km/s)&

IMLEO&(kg)&

Total&BNTR&Delta,V,&Varying&BNTR&Isp&&

2033&Opportunity&

EP$System$Isp$=$3000s,$LH2$Drop$Tanks$JeCsoned$following$impulsive$TMI

900s$

950s$

(a) Total

3.7$

3.72$

3.74$

3.76$

3.78$

3.8$

3.82$

3.84$

300000$ 325000$ 350000$ 375000$ 400000$ 425000$ 450000$

TMI$Delta*V$(km/s)$

IMLEO$(kg)$

TMI$Delta*V,$Varying$BNTR$Isp$$

2033$Opportunity$

EP$System$Isp$=$3000s,$LH2$Drop$Tanks$JeAsoned$following$impulsive$TMI

900s$

950s$

(b) TMI

2.2#

2.22#

2.24#

2.26#

2.28#

2.3#

2.32#

2.34#

2.36#

2.38#

2.4#

300000# 325000# 350000# 375000# 400000# 425000# 450000#

MOI$Delta*V$(km/s)$

IMLEO$(kg)$

MOI$Delta*V,$Varying$BNTR$Isp$$

2033$Opportunity$

EP#System#Isp#=#3000s,#LH2#Drop#Tanks#JeAsoned#following#impulsive#TMI

900s#

950s#

(c) MOI

2.7$

2.8$

2.9$

3$

3.1$

3.2$

3.3$

300000$ 325000$ 350000$ 375000$ 400000$ 425000$ 450000$

TEI$Delta*V$(km/s)$

IMLEO$(kg)$

TEI$Delta*V,$Varying$BNTR$Isp$$

2033$Opportunity$

EP$System$Isp$=$3000s,$LH2$Drop$Tanks$JeBsoned$following$impulsive$TMI

900s$

950s$

(d) TEI

Figure 3. BNTR ∆V s for 2033 Mission

NASA/TM—2013-216568 5
0"

1"

2"

3"

4"

5"

6"

300000" 325000" 350000" 375000" 400000" 425000" 450000"

Total&EP&Delta+V&(km/s)&

IMLEO&(kg)&

Total&EP&Delta+V,&Varying&BNTR&Isp&&

2033&Opportunity&

EP"System"Isp"="3000s,"LH2"Drop"Tanks"Je@soned"following"impulsive"TMI

900s"

950s"

(a) Total

0"

0.2"

0.4"

0.6"

0.8"

1"

1.2"

1.4"

1.6"

1.8"

2"

300000" 325000" 350000" 375000" 400000" 425000" 450000"

Outbound(EP(Delta/V((km/s)(

IMLEO((kg)(

Outbound(EP(Delta/V,(Varying(BNTR(Isp((

2033(Opportunity(

EP"System"Isp"="3000s,"LH2"Drop"Tanks"JeBsoned"following"impulsive"TMI

900s"

950s"

(b) Outbound, Earth-Mars

0"

0.5"

1"

1.5"

2"

2.5"

3"

3.5"

4"

300000" 325000" 350000" 375000" 400000" 425000" 450000"

Inbound'EP'Delta/V'(km/s)'

IMLEO'(kg)'

Inbound'EP'Delta/V,'Varying'BNTR'Isp''

2033'Opportunity'

EP"System"Isp"="3000s,"LH2"Drop"Tanks"Je@soned"following"impulsive"TMI

900s"

950s"

(c) Inbound, Mars-Earth

Figure 4. EP ∆V s for 2033 Mission

The data presented in these charts was interpolated for the IMLEO of the fully assembled MTV vehicle. Since these burns were modeled impulsively for the COPERNICUS analysis, a separate calculation of

gravity loss for each of the BNTR burns was completed and added to the impulsive ∆V . The BNTR and

EP ∆V values used in this analysis are presented in Table 1. The gravity loss calculations assume each

of the three BNTR propulsive maneuvers are performed as a 2-perigee burn departure/capture maneuvers.

For high-thrust burns of this magnitude, even with the high thrust BNTR engines there can be significant

gravity losses. By utilizing a 2-burn maneuver sequence instead of a single burn, the gravity loss of each

maneuver is minimized. In future assessments, to reduce complexity and risk the MOI and TEI maneuvers

could be reduced to a single perigee burn without a substantial increase in gravity loss ∆V . Since no mass

drops are performed during MOI and TEI, the ∆V in Table 1 is collective, representing the sum of the two

perigee burns for each of these maneuvers.

Phase Name EP DV

Impulsive DV G-Loss DV

(m/s) (m/s) (m/s)

TMI-1 Burn 1765

TMI-2 Burn 2029 317

EP Burn 1 639

EP Burn 2 781

MOI 2324 79

TEI 3060 100

EP Burn 3 2873

EP Burn 4 0

Mission Delta-V Summary

BNTR DV Table 1. BNTR and EP ∆V s for Reference 2033 1-Year BNTEP Earth-Mars Roundtrip Mission

Following assembly of the MTV in LEO, a 2-perigee burn TMI using the three 25 klbf BNTR engines

injects the MTV into a hyperbolic Earth-Mars transfer trajectory on May 6, 2033. TMI is separated into two

burns in order to reduce gravity losses as well as to take advantage of the drop tank design of the MTV. In

order to reduce the inert mass that needs to be accelerated throughout the mission, the MTV jettisons the

NASA/TM—2013-216568 6
depleted LH2 drop tanks following each of the TMI burns. The first pair of empty LH2 tanks are jettisoned

following TMI-1 such that TMI-2 does not have to push the empty drop tank mass. Likewise, the second

pair of LH2 drop tanks are jettisoned following TMI-2 in order to further reduce the inert mass carried

throughout the remainder of the mission. Figure 5 shows am image of how the four drop tanks separate

from the star truss.

Figure 5. Jettison of Drop Tanks from Star Truss following TMI Burn

At the start of the in-transit phase to Mars, the BNTR engines are powered down from 545 MWt (for

the NTR/thermal propulsion mode) to 1.76 MWt (for the EP/power generation mode). Energy generated

in the reactor fuel assemblies is removed using a secondary closed gas loop that carries a helium-xenon

(He-Xe) gas mixture. In a BNTR design derived from the Rover/NERVA technology program,3the existing regenerative-cooled core support tie tubes would carry recirculated He/Xe coolant gas. Power from

the NERVA fuel elements surrounding the tie tubes enters the tubes via conduction. In the fast reactor

BNTR design called ESCORT,4 a closed loop, coaxial energy transport duct, integrated into each UO2-W

fuel element, carries the He-Xe coolant gas. The heated gas is then routed to a high power (500 kWe) BRU

consisting of a Turbine-Alternator-Compressor (TAC) assembly that generates electricity at approximately

20% energy conversion efficiency to power the EP system and spacecraft systems. Waste heat is rejected to

space using a pumped-loop radiator.5 A 36.7 day low thrust EP burn follows the half-day forced coast after

TMI. This burn is used to add additional energy to the MTV such that the high thrust BNTR engines do

not need to provide the full departure energy that would be necessary for a true ballistic transfer. A 44 day

coast period separates this first EP burn from a second 43.8 day low thrust EP burn then ends a half-day

prior to MOI.

Following completion of the last EP burn just prior to Mars arrival, the power level and fuel temperature

of the BNTR engines are increased to full capability for the high thrust/high Isp 2-perigee burns required

to propulsively capture into a 24 hr elliptical Mars parking orbit on September 9, 2033. At this point the

crew will be captured in a 300 km x 24 hr Mars orbit for 30 days.

The BNTR engines are used again to perform the 2-perigee burn TEI maneuver which injects the MTV

into a Mars-Earth hyperbolic transfer trajectory on October 9, 2033. After a successful TEI, the BNTR

engines are again powered down to 1.76 MW for a 80.8 day EP burn occurring a half-day after TEI completion. A 127.7 day coast returns the MTV to Earth and aligns for a direct re-entry. V∞ at Earth was

constrained to be less than 11.5 km/s. The optimization eliminated the need for a final EP burn near Earth

on the return leg of the trajectory. One may be added if a case arises where Earth arrival V∞ exceeds 11.5

km/s or when the trajectory needs to be corrected part way through the return coast to align properly to

intercept Earth. Once the MTV has reached the Earth entry interface point, the crew boards the MPCV

and separates from the MTV. Re-entry occurs on May 6, 2034 while the MTV flies by Earth and is disposed

of in heliocentric space.

For the 1-year round trip BNTEP mission, the total operational time for the BNTR engines is 324

megawatt - days (MWD) which includes 1.75 hours of high thrust / high power mode operation totaling 40

MWD and 284 MWD of reduce power EP mode operation. Assuming 1.2 grams of U-235 consumed per

MWD, the total U-235 fuel loss in each engine is 389 grams. Since each cermet fuel BNTR has in excess of

NASA/TM—2013-216568 7
200 kg, the burn-up is less than 0.2% which is quite acceptable.

Table 2 shows the timeline for the 2033 reference mission, including pre-burn masses and propellant loads

for each of the BNTR and EP burns. Similarly to Table 1,the ∆V and burn time in the mission timeline

are collective for MOI and TEI, representing the sum of the two perigee burns.

Phase # Event Name BNTR DV EP DV

(m/s) (m/s)

1 Earth Departure

1.1 TMI-1 Burn 1765.0

1.2 Coast, drop 2 drop tanks

1.3 TMI-2 Burn 2345.3

1.4 Coast, drop 2 drop tanks

2 Earth-Mars Transit

2.1 Coast

2.2 EP Burn 1 638.8

2.3 Coast

2.4 EP Burn 2 780.7

2.5 Coast

2.6 MOI 2402.9

3 Mars Stay

4 Mars-Earth Transit

4.1 TEI 3160.4

4.2 Coast

4.3 EP Burn 3 2873.2

4.4 Coast

4.5 EP Burn 4 0.0

4.6 Coast

4.7 Direct Entry at Earth

Event Begin 

Date Duration

(hr)

5/6/33 0.00

5/6/33 0.52

5/6/33 8.00

5/6/33 0.52

5/6/33 0.00

5/6/33 0.00

5/6/33 12.00

5/6/33 881.93

6/12/33 1056.79

7/26/33 1052.23

9/8/33 12.00

9/9/33 0.36

9/9/33 720.00

10/9/33 0.00

10/9/33 0.35

10/9/33 12.00

10/9/33 1938.35

12/29/33 3064.94

5/6/34 0.00

5/6/34 0.00

5/6/34 0.00

Pre-Burn 

Mass

(kg)

389907.08

389907.08

303236.05

303236.05

216474.84

216474.84

216474.84

216474.84

211825.39

211825.39

206278.13

206278.13

157114.16

157114.16

157114.16

109825.47

109825.47

99606.68

99606.68

99606.68

99606.68

BNTR Prop 

Load

(kg)

0.00

70671.04

0.00

70761.20

0.00

0.00

0.00

0.00

0.00

0.00

0.00

49163.97

0.00

0.00

47288.69

0.00

0.00

0.00

0.00

0.00

0.00

EP Prop 

Load

(kg)

0.00

0.00

0.00

0.00

0.00

0.00

0.00

4649.45

0.00

5547.26

0.00

0.00

0.00

0.00

0.00

0.00

10218.79

0.00

0.00

0.00

0.00

Table 2. Mission Timeline for 2033 Reference 1-Year BNTEP Earth-Mars Roundtrip Mission

II.E. Launch Summary

The Space Launch System (SLS) Block-IB configuration, a Shuttle-Derived Heavy Lift Vehicle capable of

delivering 113.8 mt into orbit,6is used for delivering the MTV components to LEO for assembly. Vehicle

integration is performed via autonomous rendezvous and docking of major elements, along with crew assisted

orbital assembly to attach the four LH2 drop tanks. Five SLS launches over 120 days deliver the MTV’s

major elements. A graphic showing the fully assembled MTV with the main components identified is shown

in Figure 6.

Figure 6. Concept Design of Mars Transfer Vehicle (MTV) Following Assembly in Low Earth Orbit

A launch summary giving the launch order as well as the total wet mass of each launch is provided in

Table 3. The first launch carries up the core propulsion and power module. The second launch delivers the

large in-line LH2 propellant tank. The third launch delivers the 4-sided star truss and deployable, twin EP

thruster arms with Xenon Hall engines, and the attached TransHab module. Launches 4 and 5 deploy four

NASA/TM—2013-216568 8
modular LH2 propellant drop tanks manifested as twin tank sets. The MPCV and crew is launched on a

Block-I SLS or commercial launch vehicle to the MTV after the third Block-IB SLS launch and assists in

the attachment of the 4 LH2 drop tanks to the MTV’s star truss.

Launch Mass

Launch Elements (t)

1 Core Power and Propulsion Module 37.40

2 In-line tank 104.55

3 TransHab, Startruss + EP System 56.15

4 Crew, Consumables, Re-entry Capsule 18.38

5 Drop Tanks 1,2 88.67

6 Drop Tanks 3,4 88.76

Total 393.91

Launch Summary

Table 3. Launch Order and Wet Mass for 2033 Reference 1-Year BNTEP Earth-Mars Roundtrip Mission

Once fully assembled in LEO, the crewed MTV has an initial mass of 389.9 mt. The vehicles main

components and their masses are provided in Table 4.

Unit Inert 

Mass

Total Inert 

Mass

(mt) (mt)

Non-Dropped Elements 99.61

Core Propulsion and Power Module 1 37.40 37.40

In-line LH2 Tank 1 8.10 8.10

Star Truss+EP Subsystem 1 13.03 13.03

Crewed Payload Element 1 41.08 41.08

TransHab 1 22.70 22.70

Crew 1 0.80 0.80

Consumables 1 4.08 4.08

MPCV 1 13.50 13.50

Drop Tanks 32.00

Individual Tank 4 8.00 32.00

Element Masses

Table 4. Main Components of the Mars Transfer Vehicle

II.E.1. Launch 1

• Core Propulsion and Power Module

The core stage uses a cluster of three 25 klbf thrust BNTP engines each carrying an external radiation

shield for crew protection. These engines have a T /W of approximately 4-5. Analysis in this paper

assumed the use of cermet engines operating in the 1 MWt range utilizing UO2 in a tungsten metal

matrix fuel. However, analysis of this fuel option has to date only been completed for BNTR engines

operating at significantly lower electrical power levels, 10s of kWe.

The core propulsion and power module contains the following key components:

– Three BNTR engines, each with its own external radiation shield

– Three 500 kWe-class recuperated Brayton PCUs (containing the BRU, heat exchanger, heat rejection and PMAD)

– Combined thrust and cylindrical support structure (8.4 m outer diameter x 20 m length) which

houses the Brayton PCUs and provides support for the shared, 1-sided 970 m2cylindrical radiator

mounted to its exterior

– Avionics

– Auxiliary battery power

– Docking system

In order to provide the 1 MWe of power needed by the EP system, each 500 kWe BRU is operated

at 2/3 of the rated power, approximately 334 kWe. Operating in this way provides BRU redundancy.

Should a malfunction occur resulting in the loss of a BRU or loss of an engine during the mission, the

remaining two units can be ramped up and operated at the full 500 kWe power level.

NTR Engine Performance Parameters:

NASA/TM—2013-216568 9
– Fuel Type: UO2 in a tungsten metal matrix cermet

– Engine Thrust Level: 25 klbf

– Engine T/W: Approx. 4-5 for 25 klbf engines

– Number of Engines: 3 engine cluster

– Isp: 900 s

– Allowable Engine Thrust Mode Burn Time: 1.5 to 2.0 hrs

– Allowable Power Mode Operating Time: 1.0 to 1.5 yrs

Power Conversion System (PCS) Performance Parameters:

– Closed Brayton Cycle Total Electrical Power: 1 MWe

– Brayton Engine Unit Size: up to 500 kWe

– Turbine Inlet Temperature: 1300 K

Future launch vehicle assessments may determine that these components could be delivered using a

Block-I SLS with a larger faring rather than the Block-IB SLS. The additional lift capability of the Block-II

design is unnecessary for the 37.4 mt launch mass.

II.E.2. Launch 2

The in-line tank is sized to carry the BNTR propellant for the MOI and TEI burns as well as any excess

propellant required for TMI beyond the drop tank capacity. For the 2033 reference case, the total propellant

load carried in the in-line tank is 96.4 mt. The tanks are cylindrical with forward and rear cylindrical

adaptor sections with docking attachments. A 920 kg ZBO Brayton-cycle cryocooler system is located in the

forward adaptor section. The ZBO system removes heat penetrating the 60-layer MLI system throughout

the mission with peak use occurring during MTV assembly in LEO where the tank heat flux is the highest.

The shell of the tank is manufactured using an aluminum-lithium alloy, AlLi ”Weldalite” T6E4. Assuming

an exterior diameter of 8.4 m, the tank has a wall thickness of 0.3175 cm and a total inert mass of 8095.9

kg. Included in this mass is 600 kg to account for lines, valves, and fittings. Additional parameters and

assumptions used for sizing the in-line tank are provided in Table 5. Avionics, RCS and auxiliary battery

power are also carried on this launch.

Propellant Load 96452.66 kg

Additional Tank Masses 1520 kg

Pressure 30 psia

Tank Material

Tank Interior Diameter 8.39365 m

Wall Thick. 0.3175 cm

Volume Margin 5.00% %

Propellant Density 67.541 kg/m3

Tank Material Density 2700 kg/m3

Required Volume 1499.463 m3

End Cap Volume 218.946 m3

Cyl. Section Volume 1280.518 m3

Cyl. Section Length 23.142 m

Total Tank Length 29.083 m

Tank Outer Diameter 8.400 m

Tank Shell Mass 6575.870 kg

Total Tank Inert Mass 8095.870 kg

Total Tank Wet Mass 104548.531 kg

Tank Shape: Sqrt(2)/2 Cylinder with End Caps

In-line Tank Sizing Inputs

AlLi "Weldalite" T6E4

Table 5. Inline Tank Sizing using Sizer Tool Parametrics

NASA/TM—2013-216568 10
II.E.3. Launch 3

• Star Truss and EP Subsystem

The EP subsystem consists of ten 100 kWe Xenon Hall thrusters, having five thrusters located on each

EP thruster arm. The EP system is used to gradually accelerate or decelerate the spacecraft in interplanetary space., allowing the mission to maintain relatively low ∆V BNTR departure and injection

burns when compared to a 1-year all NTR mission. The electrical power generated by the BNTEP

reactor system will also provide power to the other spacecraft subsystems, eliminating the need for

photovoltaic arrays. Minimal attitude control will be required during interplanetary transit when the

BNTR engines are in low-thrust mode since the EP system can correct for trajectory inaccuracies

during the long burn periods. A RCS having thrusters and bipropellant located on the core bimodal

propulsion and power stage as well as the star truss forward adaptor ring, will provide additional attitude control which will be required for initial vehicle assembly, larger mid-course corrections, vehicle

orientation maneuvers, and attitude control while in high-thrust mode as well as during Mars stay.5

The EP subsystem has a total inert mass of 10.5 mt. This analysis leveraged results for sizing EP

subsystems for MWe-class Earth-Mars missions done for other studies.7 The EP mass for BNTEP

missions is similar except any additional power subsystem mass for solar arrays is unnecessary. For

the 2033 reference case, the total Xenon propellant load carried by the EP subsystem is 20.4 mt.

The star truss is a rigid, 4-sided composite structure, with integrated docking, electrical and drop tank

connections, and forward RCS and propellant. Mounted to the truss are two EP thruster arms each

with five 100 kWe ion thrusters plus the Xenon propellant and tankage. Truss sizing for the drop tank

was done by interpolating parametric data generated from the Sizer Excel Tool developed at NASA

GRC.8 The hang-mass and propellant load of the drop tanks were used to interpolate parametric curves

generated by the Sizer Tool. Star truss sizing parameters are provided in Table 6. The truss length is

set such that it is 5 m longer than the length of the in-line tank.

Truss Length 18.000 m

(Tank Wet Mass)*(T/WIMLEO) 15.127 mt

Truss Mass 2534.010 kg

Star Truss Sizing

Table 6. Truss Sizing using Sizer Tool Parametrics

• Crewed Payload Element

The crewed payload element consists of the transit habitat, 4 crew members with suits, consumables,

and the MPCV. The total initial mass of this payload element when delivered to LEO is 41.08 mt.

The 4 person crew will reside within the transit habitat (TransHab) during interplanetary transit as

well as throughout the stay in Mars orbit. Consumables are carried within the TransHab, amounting

to 2.45 kg/d per crew member. An additional 51 days were added to the total crewed time to budget

for consumables required for the period of time the crew will on-orbit assisting in final MTV assemble

and checkout in LEO. The MPCV will be used as a re-entry capsule which the crew will enter just

prior to arrival at Earth. At Earth entry interface, the crew and the MPCV will separate from the

MTV, enter the atmosphere, and perform a water landing.

II.E.4. Launch 4, 5

The final two SLS launches deliver twin sets of LH2 drop tanks that will attach to the star truss. Each

drop tank is 13 m long with an external diameter of 8.4 m and has an inert mass of 8 mt. Each drop tank

is capable of carrying up to 39.2 mt of LH2 propellant. For the 2033 reference case, the LH2 propellant

required for TMI does not fill the drop tanks so there is excess volume between the tanks to optimize the

∆V split between TMI-1 and TMI-2. For this analysis, TMI-1 and TMI-2 were sized in order to maintain

an approximately equal propellant load between the tank pairs.

In addition to the wet mass of the drop tanks, each drop tank pair requires a 2 mt intertank adapter

for launch. Once in orbit the intertank adapter is jettisoned prior to full vehicle assembly, hence, the total

launch mass in Table 3 is 4 mt higher than the IMLEO of the fully assembled MTV.

NASA/TM—2013-216568 11
II.F. Variations in Total Time of Flight

Allowing the mission time to extend beyond 365 days can provide significant savings in propellant mass.

Figure 7 and Figure 8 show the variation in BNTR and EP ∆V as total flight time is extended. The

extra time for interplanetary transit allows the BNTEP system to reduce the amount of energy added to

or removed from the MTV by the BNTR system. Specifically, the TMI maneuver can be reduced by approximately 110 m/s when 20 days are added to the mission time by allowing more of the required Earth

departure acceleration of the vehicle to be performed with the high efficiency EP system, increasing the

outbound EP ∆V by approximately 130 m/s. The reduction in LH2 and Xenon propellant by extending

the mission 20 days could reduce IMLEO of the MTV from 389.9 mt to 340.2 mt.

Although there may be significant mass savings, any mass savings made by extending the mission duration

must be carefully weighed against the increased risk to the crew by extending exposure time to harmful

radiation as well as the prolonging the period in a microgravity environment.

0"

1"

2"

3"

4"

5"

6"

7"

8"

9"

10"

365" 375" 385" 395" 405" 415" 425" 435" 445" 455" 465"

Total&BNTR&Delta,V&(km/s)&

Total&Mission&Time&of&Flight&(days)&

Total&BNTR&Delta,V&vs.&Total&Mission&Time&of&Flight&

2033&Opportunity,&Varying&IMLEO&

EP"System"Isp"="3000s,"BNTR"Isp"="900s,"LH2"Drop"Tanks"JeEsoned"Following"TMI"

350"mt"

400"mt"

(a) Total

3.1$

3.2$

3.3$

3.4$

3.5$

3.6$

3.7$

3.8$

3.9$

365$ 375$ 385$ 395$ 405$ 415$ 425$ 435$ 445$ 455$ 465$

TMI$Delta*V$(km/s)$

Total$Mission$Time$of$Flight$(days)$

TMI$Delta*V$vs.$Total$Mission$Time$of$Flight$

2033$Opportunity,$Varying$IMLEO$

EP$System$Isp$=$3000s,$BNTR$Isp$=$900s,$LH2$Drop$Tanks$JeFsoned$Following$TMI$

350$mt$

400$mt$

(b) TMI

0"

0.5"

1"

1.5"

2"

2.5"

365" 375" 385" 395" 405" 415" 425" 435" 445" 455" 465"

MOI$Delta*V$(km/s)$

Total$Mission$Time$of$Flight$(days)$

MOI$Delta*V$vs.$Total$Mission$Time$of$Flight$

2033$Opportunity,$Varying$IMLEO$

EP"System"Isp"="3000s,"BNTR"Isp"="900s,"LH2"Drop"Tanks"JeFsoned"Following"TMI"

350"mt"

400"mt"

(c) MOI

0"

0.5"

1"

1.5"

2"

2.5"

3"

3.5"

365" 375" 385" 395" 405" 415" 425" 435" 445" 455" 465"

TEI$Delta*V$(km/s)$

Total$Mission$Time$of$Flight$(days)$

TEI$Delta*V$vs.$Total$Mission$Time$of$Flight$

2033$Opportunity,$Varying$IMLEO$

EP"System"Isp"="3000s,"BNTR"Isp"="900s,"LH2"Drop"Tanks"JeFsoned"Following"TMI"

350"mt"

400"mt"

(d) TEI

Figure 7. BNTR ∆V s vs Total Mission Time of Flight for 2033 Mission Opportunity

NASA/TM—2013-216568 12
0"

1"

2"

3"

4"

5"

6"

365" 375" 385" 395" 405" 415" 425" 435" 445" 455" 465"

Total&EP&Delta+V&(km/s)&

Total&Mission&Time&of&Flight&(days)&

Total&EP&Delta+V&vs.&Total&Mission&Time&of&Flight&

2033&Opportunity,&Varying&IMLEO&

EP"System"Isp"="3000s,"BNTR"Isp"="900s,"LH2"Drop"Tanks"JeEsoned"Following"TMI"

350"mt"

400"mt"

(a) Total

0"

0.2"

0.4"

0.6"

0.8"

1"

1.2"

1.4"

1.6"

1.8"

2"

365" 375" 385" 395" 405" 415" 425" 435" 445" 455" 465"

Outbound(EP(Delta/V((km/s)(

Total(Mission(Time(of(Flight((days)(

Outbound(EP(Delta/V(vs.(Total(Mission(Time(of(Flight(

2033(Opportunity,(Varying(IMLEO(

EP"System"Isp"="3000s,"BNTR"Isp"="900s,"LH2"Drop"Tanks"JeFsoned"Following"TMI"

350"mt"

400"mt"

(b) Outbound, Earth-Mars

0"

0.5"

1"

1.5"

2"

2.5"

3"

3.5"

365" 375" 385" 395" 405" 415" 425" 435" 445" 455" 465"

Inbound'EP'Delta/V'(km/s)'

Total'Mission'Time'of'Flight'(days)'

Inbound'EP'Delta/V'vs.'Total'Mission'Time'of'Flight'

2033'Opportunity,'Varying'IMLEO'

EP"System"Isp"="3000s,"BNTR"Isp"="900s,"LH2"Drop"Tanks"JeFsoned"Following"TMI"

350"mt"

400"mt"

(c) Inbound, Mars-Earth

Figure 8. EP ∆V s vs Total Mission Time of Flight for 2033 Mission Opportunity

III. Conclusion

Ballistic opposition-class (short stay) Mars round trip missions are characterized by long in-space durations, shorter surface stays, and high ∆V requirements. Because of this, architectures for crewed Mars

missions, including Mars DRA 5.0, often select a conjunction-class (long stay) trajectory. However, the

overall shorter duration of an opposition-class mission is desirable for reducing risk associated with crew

behavioral health and performance as well as medical capabilities.1 Utilizing BNTEP technology for short

duration Mars missions can reduce the in-space duration as well as the IMLEO for comparable NTR missions.6 BNTEP technology enables a feasible 1-year Mars round trip mission by using the EP system during

transits to maintain relatively low escape and arrival energy requirements for the high thrust BNTR engines.

Reducing propulsive requirements for the BNTR engines allows the IMLEO of the MTV to remain relatively

low. The analysis presented in this paper shows that the MTV of a 2033 BNTEP 1-year mission would

have an IMLEO of 389.9 mt, requiring 5 SLS Block-IB launches to deliver all of the components. Future

investigation into the use of BNTEP technology for conjunction-class and split mission architectures could

potentially further reduce in-space time and potentially IMLEO for these missions.

Additional work in the future could also be done to assess vehicle reusability by capturing the MTV into

a highly elliptic Earth orbit or halo orbit around the Earth-Moon L2 libration point following MPCV and

crew separation. This would reduce cost and risk for future BNTEP missions by already having the main

structure of the MTV assembled and only requiring refueling.

NASA/TM—2013-216568 13
References

1Mars Architecture Steering Group., Human Exploration of Mars Design Reference Architecture 5.0, NASA-SP-2009-566,

2009.

2Ocampo, C., Senent, J., THE DESIGN AND DEVELOPMENT OF COPERNICUS: A COMPREHENSIVE TRAJECTORY DESIGN AND OPTIMIZATION SYSTEM, IAC-06-C1.4.04, 2006.

3Koeing, D., Experience Gained from the Space Nuclear Rocket Programs (Rover / NERVA) LA-10062-H, Los Alamos

National Laboratory, 1986.

4Joyner, C., The Synergistic Application of Chemical Rocket Component Technologies to the ESCORT Nuclear Bimodal

System AIAA-2000-3211, American Institute of Aeronautics and Astronautics, 2000.

5Borowski, S., Dudzinski, L., Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP), 2008.

6Oleson, S., Piloted Mars Combined SEP-Chem. Conjunction Piloted and Cargo Designs, Powerpoint Presentation, December 14 2012.

7Gilland, J., LaPointe, M., Oleson, S., Mercer, C., Pencil, E., Mason, L., MW-Class Electric Propulsion System Designs

for Mars Cargo Transport, AIAA, 2012.

8McCurdy, D., Results from Sizer Excel Tool, 2012.

9Adams, Constance, Kriss. ISS TransHab: A Space Inflatable Habitation Module, In proceedings of Space 2000: The

Seventh International Conference and Exposition on Engineering, Construction, Operations and Business in Space (2000);

American Society of Civil Engineers, Reston VA

10NASA, Reference Mission Version 3.0 Addendum to the Human Exploration of Mars: The Reference Mission of the

NASA Mars Exploration Study Team, NASA Special publication 6107-ADD, 1998

11Borowski, S., McCurdy, D., Packard, T., Modular Growth NTR Space Transportation System for Future NASA Human

Lunar, NEA and Mars Exploration Missions, AIAA, 2012.

12Burke, L., Falck, R., McGuire, M., Interplanetary Mission Design Hanbook: Earth-to-Mars Mission Opportunities 2026

to 2045, NASA/TM-2010-216764, 2010.

NASA/TM—2013-216568 14
Appendix A - ∆V s for 2031-2041 Mission Opportunities

A.1 Assumptions and Constraints for General Analysis

• TOF = 365 days

• Mars Stay Time = 30 days

• EP system efficiency = 65%

• NTR Isp = 900 s

• Earth departure orbit: 407 km circular orbit inclined at 28.5

◦

• Mars arrival orbit: 300 km x 24 hr elliptical orbit

• Maximum Earth arrival V∞ = 11.5 km/s

• Four 8 mt LH2 drop tanks are jettisoned following TMI

0"

2"

4"

6"

8"

10"

12"

14"

16"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

Total&BNTR&Delta,V&(km/s)&

Opportunity&Year&

Total&BNTR&Delta,V,&Varying&EP&System&Isp&&

OpportuniBes&2031,2041&

IMLEO"="350t,"BNTR"Isp"="950s"

2000s"

2500s"

3000s"

3500s"

4000s"

(a) Total

0"

1"

2"

3"

4"

5"

6"

7"

8"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

TMI$Delta*V$(km/s)$

Opportunity$Year$

TMI$Delta*V,$Varying$EP$System$Isp$$

Opportuni@es$2031*2041$

IMLEO"="350t,"BNTR"Isp"="950s"

2000s"

2500s"

3000s"

3500s"

4000s"

(b) TMI

0"

0.5"

1"

1.5"

2"

2.5"

3"

3.5"

4"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

MOI$Delta*V$(km/s)$

Opportunity$Year$

MOI$Delta*V,$Varying$EP$System$Isp$$

Opportuni?es$2031*2041$

IMLEO"="350t,"BNTR"Isp"="950s"

2000s"

2500s"

3000s"

3500s"

4000s"

(c) MOI

0"

0.5"

1"

1.5"

2"

2.5"

3"

3.5"

4"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

TEI$Delta*V$(km/s)$

Opportunity$Year$

TEI$Delta*V,$Varying$EP$System$Isp$$

Opportuni?es$2031*2041$

IMLEO"="350t,"BNTR"Isp"="950s"

2000s"

2500s"

3000s"

3500s"

4000s"

(d) TEI

Figure 9. BNTR ∆V s for 2031-2041 Mission Opportunities

NASA/TM—2013-216568 15
0"

1"

2"

3"

4"

5"

6"

7"

8"

9"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

Total&EP&Delta+V&(km/s)&

Opportunity&Year&

Total&EP&Delta+V,&Varying&EP&System&Isp&&

Opportuni?es&2031+2041&

IMLEO"="350t,"Chemical"Isp"="900s"

2000s"

2500s"

3000s"

3500s"

4000s"

(a) Total

0"

1"

2"

3"

4"

5"

6"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

Outbound(EP(Delta/V((km/s)(

Opportunity(Year(

Outbound(EP(Delta/V,(Varying(EP(System(Isp((

Opportuni@es(2031/2041(

IMLEO"="350t,"Chemical"Isp"="900s"

2000s"

2500s"

3000s"

3500s"

4000s"

(b) Outbound, Earth-Mars

0"

0.5"

1"

1.5"

2"

2.5"

3"

3.5"

2031" 2032" 2033" 2034" 2035" 2036" 2037" 2038" 2039" 2040" 2041"

Inbound'EP'Delta/V'(km/s)'

Opportunity'Year'

Inbound'EP'Delta/V,'Varying'EP'System'Isp''

Opportuni@es'2031/2041'

IMLEO"="350t,"Chemical"Isp"="900s"

2000s"

2500s"

3000s"

3500s"

4000s"

(c) Inbound, Mars-Earth

Figure 10. EP ∆V for 2031-2041 Mission Opportunities

NASA/TM—2013-216568 16