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Acknowledgments

The reactor, shield, power conversion, power management and distribution, and heat rejection

systems were sized using SRPS_Opt, created by Lee Mason, NASA Glenn Research Center.

Robert Adams, NASA Marshall Space Flight Center, led the overall 2004–2005 RASC Mars

Orbiter study and is the point of contact for the NEP-Rankline and the chemical studies.

The Space Propulsion and Mission Analysis Office, led by Glen Horvat,

was instrumental in supporting this study.

Available from

NASA Center for Aerospace Information
7121 Standard Drive
Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

This report is a preprint of a paper intended for presentation at a conference. Because

of changes that may be made before formal publication, this preprint is made

available with the understanding that it will not be cited or reproduced without the

permission of the author.

This report is a formal draft or working

paper, intended to solicit comments and

ideas from a technical peer group.

This report contains preliminary

findings, subject to revision as

analysis proceeds.

Available electronically at

http://gltrs.grc.nasa.gov

NASA/TM—2006-214106

Use of High-Power Brayton Nuclear Electric Propulsion (NEP)

for a 2033 Mars Round-Trip Mission

Melissa L. McGuire

National Aeronautics and Space Administration

Glenn Research Center
Cleveland, Ohio 44135

Michael C. Martini and Thomas W. Packard

Analex Corporation

Brook Park, Ohio 44142

John E. Weglian and James H. Gilland

Ohio Aerospace Institute

Brook Park, Ohio 44142

Abstract

The Revolutionary Aerospace Systems Concepts (RASC) team, led by the NASA Langley Research

Center, is tasked with exploring revolutionary new approaches to enabling NASA to achieve its strategic
goals and objectives in future missions. This paper provides the details from the 2004-2005 RASC study
of a point-design that uses a high-power nuclear electric propulsion (NEP) based space transportation
architecture to support a manned mission to Mars. The study assumes a high-temperature liquid-metal
cooled fission reactor with a Brayton power conversion system to generate the electrical power required
by magnetoplasmadynamic (MPD) thrusters. The architecture includes a cargo vehicle with an NEP
system providing 5 MW of electrical power and a crewed vehicle with an NEP system with two reactors
providing a combined total of 10 MW of electrical power. Both vehicles use a low-thrust, high-efficiency
(5000 sec specific impulse) MPD system to conduct a spiral-out of the Earth gravity well, a low-thrust
heliocentric trajectory, and a spiral-in at Mars with arrival late in 2033. The cargo vehicle carries two
moon landers to Mars and arrives shortly before the crewed vehicle. The crewed vehicle and cargo
vehicle rendezvous in Mars orbit and, over the course of the 60-day stay, the crew conducts nine-day
excursions to Phobos and Deimos with the landers. The crewed vehicle then spirals out of Martian orbit
and

returns via a low-thrust trajectory to conduct an Earth flyby. The crew separates from the vehicle

prior to Earth flyby and aerobrakes for a direct-entry landing.

Introduction

This paper details a Revolutionary Aerospace Systems Concepts (RASC) study investigating a high-

power nuclear electric propulsion (NEP) space transportation architecture to support a manned mission to
Mars. The RASC project, led by the NASA Langley Research Center, is tasked with exploring
revolutionary new approaches to enabling NASA to achieve its strategic goals and objectives in future
missions. For this study, two vehicle concepts were designed, both using a high-power NEP system with
Brayton power conversion and magnetoplasmadynamic (MPD) thrusters. The first vehicle is the Mars
Transfer Vehicle (MTV) which carries the crew from the Earth to Mars and back again. The second
vehicle, the Cargo Transfer Vehicle (CTV), delivers additional cargo necessary for the mission from the
Earth to Mars.

This paper details one of the four space transportation architectures selected by the 2004-2005 RASC

Mars Obiter Study for analysis. The other three investigated were nuclear thermal propulsion (Borowski,
Packard, and McCurdy, 2006), NEP with Rankine power conversion, and chemical propulsion. In order
for the architectures to be compared across an even playing field, all four started with the same mission

NASA/TM—2006-214106

182 m

Two 5 MW

reactors, shield

Four Brayton Power

conversion units

Six 7.6 m x 19 m

LH2 tanks

TransHab,

ECRV

Four 2.5 MW

MPD

thrusters (two operational)

per arm

Main radiators:

2722 m

planform area

5444 m

effective area

182 m

Two 5 MW

reactors, shield

Four Brayton Power

conversion units

Six 7.6 m x 19 m

LH2 tanks

TransHab,

ECRV

Four 2.5 MW

MPD

thrusters (two operational)

per arm

Main radiators:

2722 m

planform area

5444 m

effective area

Figure 1.—Mars Transfer Vehicle.

and payload assumptions. The mission consisted of a split profile with the cargo elements sent out on one
vehicle and the crew sent out on a second vehicle. Each transportation architecture in the RASC study
assumed the same cargo and crew payloads. These study requirements led to a mission that was not
optimized specifically for an NEP system.

Vehicle Configurations

Mars Transfer Vehicle (MTV)

In order to provide the required artificial gravity for the crew during the Trans-Mars Injection (TMI)

outbound and Trans-Earth Injection (TEI) inbound trajectory legs, the Mars Transfer Vehicle was
configured to allow a rotation about the center of gravity. The crew is located in an inflated
Transportation Habitat (TransHab) at one end of the NEP vehicle while the Brayton power conversion
system and the nuclear reactors are located at the other end. To minimize the translation of the center of
gravity over the mission, the LH2 tanks are located at the center of the vehicle configuration. The MTV
uses two reactors, each providing 5 MW

, and a total of four Brayton power conversion units. There are

two thruster arms with four 2.5 MW

MPD thrusters (two operational, two spare) on each arm. Each

thruster arm has a radiator to reject heat from the power processing units (PPU). The total planform area
of the PPU radiators is 136.7 m

(273.4 m

effective radiating area). Six LH2 tanks that are 7.6 m in

diameter and 19 m long occupy the middle truss section of the vehicle and store the 279.4 MT of
propellant. The main radiator is comprised of two sections of double-sided flat panels attached to the
center truss structure on either side of the propellant tanks due to center of gravity requirements. The total
planform area of the main radiator is 2722 m

(5444 m

effective radiating area). The MTV is 182 m long

and must be assembled in orbit. The configuration of the MTV is shown in figure 1.

Cargo Transfer Vehicle (CTV)

The Cargo Transfer Vehicle is modeled after the NEP configuration used in the 2002 RASC Callisto

mission entitled HOPE (McGuire, et al., 2003, and Borowski et al., 2003). Since the cargo vehicle does
not require the artificial gravity spin, the propellant tanks are located at the far end from the reactor to
prevent splitting up the radiator into two sections. This avoids having the hot heat-rejection fluid routed
around the cryogenic tanks, as is required in the MTV configuration. Like the MTV, the CTV has double-
sided radiator panels attached to the central truss structure of the vehicle. The total planform area of the
main radiator is 1361 m

, which provides 2722 m

effective radiating area. The CTV uses one reactor and

NASA/TM—2006-214106

127

5 MW

reactor,

shield

Two Brayton power

conversion units

Main radiator:

1361 m

planform area

2722 m

effective area

Two 2.5 MW

MPD

thrusters (1 operational)

per arm

PPU radiator:

137 m

planform area

Two 7.6 m x 19 m

LH2 tanks

127

5 MW

reactor,

shield

Two Brayton power

conversion units

Main radiator:

1361 m

planform area

2722 m

effective area

Two 2.5 MW

MPD

thrusters (1 operational)

per arm

PPU radiator:

137 m

planform area

Two 7.6 m x 19 m

LH2 tanks

Figure 2.—Cargo Transfer Vehicle.

two Brayton power conversion units to provide 5 MW

power. The four 2.5 MW

MPD thrusters are

mounted on the outside of the truss section that contains the propellant tanks with two thrusters, one of
which is a spare, on each side. The total planform area of the PPU radiators is 273.4 m

(546.8 m

effective radiating area). The CTV only has two of the 7.6 m diameter, 19 m long LH2 tanks, storing the
63.9 MT of propellant. The CTV is 127 m long and, like the MTV, must be assembled in orbit. The
configuration of the CTV is shown in figure 2.

Assumptions

Mission Assumptions and Outline

The RASC Mars Orbiter mission was configured as an opposition class (short stay) Earth-to-Mars

round-trip mission. A crew of six is deployed to Mars, but does not perform any Mars surface operations.
Rather, they perform two nine-day excursions to Phobos and Demos before returning home to Earth. The
total stay-time in Mars orbit is 60 days. The components of the NEP stages of both vehicles are launched
on heavy-lift Magnum expendable launch vehicles (ELVs) and assembled in a circular Low Earth Orbit
(LEO) at 1000 km altitude and 28.5° inclination. The Magnum is assumed to be capable of delivering
80 MT into LEO in a payload shroud 7.5 m wide by 30 m long.

The CTV conducts a spiral escape from Earth and follows a low-thrust trajectory to Mars to pre-

deploy two moon landers (for landing on Phobos and Deimos) in Mars orbit prior to the crew’s arrival.
After assembly and checkout, a second NEP stage with the TransHab begins the spiral escape from LEO.
After the NEP stage has cleared the Van Allen belts and is ready to escape Earth, the crew is launched on
a smaller ELV (Delta IV Heavy class) in an Earth crew return vehicle (ECRV) and docks with the NEP
stage in a high orbit. At this point, the mated NEP stage with the inflated TransHab and ECRV is referred
to as the MTV. The MTV uses the reaction control system (RCS) thrusters to spin the MTV end-over-end
upon Earth escape to provide artificial gravity (38 percent of Earth gravity, equal to Mars gravity) to the
crew in the TransHab module. The MPD thrusters provide for “side thrusting” by thrusting along the axis
of rotation. Once the MTV has reached Mars space, the vehicle performs a spin down maneuver, and the
MTV spiral captures into the same Mars orbit as the CTV.

NASA/TM—2006-214106

After 60 days of Mars orbit operations, the MTV spiral escapes from Mars orbit and follows a low-

thrust trajectory back to Earth. During the heliocentric portion of the flight, the RCS thrusters induce
another end-over-end spin for artificial gravity (38 percent of Earth gravity) for the crew in the TransHab.
At Earth arrival, the ECRV separates from the MTV with the crew onboard to perform a direct-entry
aerobrake and parachute landing on Earth.

Payload Assumptions

With this mission architecture, the cargo (moon landers) is sent out on a separate vehicle than the

crew. The crew only carries enough supplies and cargo to last them through the TMI leg, the 60-day stay,
and the TEI leg of the trip. All cargo necessary to carry out moon-landing operations at the destination is
sent out on the CTV. The CTV payload consists of two moon landers designed by a team led by the
NASA Langley Research Center as part of this RASC study. These landers are designed to take three
crew on nine-day excursions to the surfaces of Deimos and Phobos, and then return to Mars orbit to
rendezvous with the MTV.

The MTV payload consists of an inflatable TransHab and an ECRV. The TransHab is similar to the

TransHab design from the Human Exploration and Development of Space (HEDS) design reference
mission 4.0 study (Joosten, 2002). The TransHab mass includes enough consumables for a 545-day
round-trip mission. Any missions with total trip times longer than 545 days must add an additional
2.45 kg/person/day to the dry-mass allocation. The crew are onboard the MTV for a total of 612 days, so
this adds 984.9 kg of consumables to the TransHab for this study. The mass of the TransHab also includes
approximately 1900 kg of water for radiation protection and 400 kg for the environmental control and
life-support system. The ECRV carries the crew during the final aerobrake for an Earth landing at the end
of the mission. Table 1 shows the masses for each of the piloted and cargo payload elements as set by the
RASC study. These masses already contain the appropriate contingency for each item, so no additional
contingency was added to the payloads in this study.

TABLE 1.—RASC 2004 PAYLOADS

Element Mass

(MT)

Vehicle

TransHab: includes food for 545 days, 6 crew

35.0

MTV

Earth crew return vehicle (ECRV)

7.0

MTV

ECRV docking structure

8.0

MTV

Two moon landers for 9-day missions

42.5

CTV

Power System

This study assumes a high-temperature, liquid-metal, fission reactor with a Brayton power conversion

system to generate the electrical power required to supply the MPD thrusters. The reactor was based on an
advanced version of the early reactor concept for the Jupiter Icy Moons Orbiter study. The fission reactors
use liquid-metal coolant loops, which operate at a temperature of about 1600 K, in order to represent
“mid-term” technology (Mason, 2001), consistent with the 2033 mission timeframe. Each reactor coolant
loop transfers heat to the Brayton system’s working fluid via a heat source heat exchanger (HSHX),
producing a Brayton turbine inlet temperature of 1500 K. The Brayton unit includes a recuperator to
improve system efficiency by pre-heating the working fluid from the compressor outlet with the turbine
exhaust before it reaches the gas cooler. The recuperator reduces the heat load of both the gas cooler and
the HSHX, which in turn reduces the size of the radiators and the reactor. The heat rejection system uses a
pumped NaK working fluid to remove heat from the Brayton working fluid via the gas cooler and
transfers that heat to the radiator panels via water heat pipes. A turbine inlet temperature of 1500 K
requires a very high-temperature turbine blade material (possibly ceramic) or active cooling of the blades,

NASA/TM—2006-214106

and a reactor temperature of 1600 K necessitates the use of refractory metals or other high temperature
material for the reactor. A schematic of the Brayton power conversion system is shown in figure 3.

The MTV uses two reactors sized to provide 5 MW

net electrical power, each. Neutron interactions

between the two reactors were not considered. The cargo vehicle only requires one reactor sized to
provide 5 MW

net electrical power. The component masses and radiator areas for both vehicles are

presented in table 2. The reactor system includes the radiation shield, which is composed of layers of
tungsten (gamma shield) and lithium hydride (neutron shield). The MTV’s shields are much heavier than
the CTV’s due to the crew’s more stringent radiation limits. The radiator is double-sided, so heat is
rejected from both sides of the radiator panels. Because of this, the effective area for rejecting heat is
double the physical area of the radiator panels. The radiator design is described by Siamidis and
Mason (2006).

TABLE 2.—POWER SYSTEM PARAMETERS.

MTV

CTV

Reactor system mass

18088

4973

Brayton power conversion system mass

8748

4374

Heat rejection system mass

33456

16728

PMAD system mass

20484

9648

Radiator area (effective)

5444

2722

Radiator area (physical)

2722

1361

HSHX

Reactor

Gas Cooler

Radiator

Alternator

Turbine

Comp.

Recuperator

HSHX

Reactor

Gas Cooler

Radiator

Alternator

Turbine

Comp.

Recuperator

Figure 3.—Schematic of the power conversion system.

0.1

0.2

0.3

0.4

0.5

0.6

1000

3000

5000

7000

9000

11000

Isp (sec)

lph

a (

/kW

MPD thruster: 1 MWe

MPD thruster: 2.5 MWe

MPD thruster: 5 MWe

30%

40%

50%

60%

70%

80%

90%

1000

3000

5000

7000

9000

11000

Isp (sec)

ici

ency (

)

MPD thruster: 1 MWe

MPD thruster: 2.5 MWe

MPD thruster: 5 MWe

(a) Near term MPD thruster alpha

(b) Near term MPD thruster system efficiency

0.1

0.2

0.3

0.4

0.5

0.6

1000

3000

5000

7000

9000

11000

Isp (sec)

lph

a (

/kW

MPD thruster: 1 MWe

MPD thruster: 2.5 MWe

MPD thruster: 5 MWe

30%

40%

50%

60%

70%

80%

90%

1000

3000

5000

7000

9000

11000

Isp (sec)

ici

ency (

)

MPD thruster: 1 MWe

MPD thruster: 2.5 MWe

MPD thruster: 5 MWe

(a) Near term MPD thruster alpha

(b) Near term MPD thruster system efficiency

Figure 4.—MPD system alpha and efficiency versus I

NASA/TM—2006-214106

Propulsion System

This study used magnetoplasmadynamic (MPD) thrusters using hydrogen propellant. Besides

operating at a high specific impulse (I

), the MPD thrusters also have the added advantages of a high-

power capability and a compact size.

This analysis used high power MPD thrusters operating at 2.5 MW

per thruster at a constant I

of 5,000 sec with a thruster lifetime of 7500 hr.

The MPD thrusters use cryogenically-stored liquid Hydrogen (LH2) propellant. This mission utilized

the 2.5 MW

thrusters assumed in the 2002 HOPE study (McGuire et al., 2003). The MTV vehicle used

four operating thrusters for a total power level of 10 MW

and had 4 non-operating spares for redundancy.

Likewise, the CTV used two operating thrusters at a total power level of 5 MW

and had two non-

operating spares for redundancy. The mass of the thrusters is I

dependant. Since a constant I

was used

in this analysis, the mass of the thrusters was calculated using the system alpha (mass/kW

) for an I

5000 sec. The thrusters were run at an I

of 5000 sec due to higher efficiencies at this specific impulse.

See figure 4 for the dependency of system alpha and MPD thruster efficiency versus operating I

One power processing unit (PPU) and one radiator are assumed per thruster. The system alpha of the

PPU and radiator is assumed to be 2.5 kg/kW

. This included the mass for the power conditioning at the

turbine (transformer to increase the voltage to 1 kV), the 1 kV transmission line, the PPU to convert
power at the other end, and the Parasitic Load radiator (PLR) to reject waste heat. The sink temperature
is assumed to be 250 K at Earth orbit for a worst-case sizing. The effective radiator areas for the two
vehicles were: CTV = 273.4 m

for a rejection of 5 MW

, and MTV = 546.8 m

for a rejection of

10 MW

Trajectory

Both the MTV and the CTV begin in LEO at 1000 km altitude, spiral out from the Earth, and

follow a low-thrust trajectory to Mars. The CTV captures into a 24.65-hr period Mars orbit (radius of
periapse = 3643 km, radius of apoapse = 37,186 km, orbital period = one Martian day). The MTV spiral-
captures into the same Mars orbit 12 days later. After a 60-day stay at Mars, the MTV returns the crew to
Earth on a flyby trajectory. The crew aerobrakes at Earth in the ECRV for a direct-entry that is
constrained to 13 km/sec at 125 km altitude.

The trajectories for the vehicles were optimized using the

VARITOP trajectory optimization program (Williams, 1994). The trajectory for the MTV and CTV are
shown in figure 5. The dates and times of each mission phase as well as the total mission time and total
crew time (for the MTV) are shown in table 3.

TABLE 3.—MISSION EVENT LIST

MTV CTV

Mission event

Date Time

segment

(days)

Mission

time

(days)

Crew

time

(days)

Date Time

segment

(days)

Mission

time

(days)

Begin Earth escape spiral

11/7/2032

12/23/2032

Escape Earth

3/24/2033

137

4/5/2033 103 103

Mars capture

10/22/2033 212 349 212

10/14/2033 192 295

Complete Mars capture spiral

11/3/2033 12 361

224

10/22/2033 9 304

Begin Mars escape spiral

1/2/2034 60

421

284

Escape Mars

1/14/2034 12 433

296

Arrive Earth

11/26/2034 316 749 612

NASA/TM—2006-214106

(a) MTV trajectory

(b) CTV trajectory

Depart Mars
January 14, 2034

Earth escape
March 24, 2033

Arrive Mars
October 22, 2033

Arrive Earth
November 26, 2034
Vinf=6.81 km/sec

60 Day Mars

Stay Time

12.1 days
spiral capture

Begin 11.7 day
spiral escape

Begin 137 days
Earth spiral escape
November 7, 2032

Thrust Arc
Coast Arc

Begin 24-hour
Mars orbit

Earth escape
April 5, 2033

Arrive Mars
October 14, 2033

Begin 102.9 days
Earth spiral escape
December 23, 2032

8.6 days
spiral capture

Thrust Arc
Coast Arc

Begin 24-hour
Mars orbit

(a) MTV trajectory

(b) CTV trajectory

Depart Mars
January 14, 2034

Earth escape
March 24, 2033

Arrive Mars
October 22, 2033

Arrive Earth
November 26, 2034
Vinf=6.81 km/sec

60 Day Mars

Stay Time

12.1 days
spiral capture

Begin 11.7 day
spiral escape

Begin 137 days
Earth spiral escape
November 7, 2032

Thrust Arc
Coast Arc

Begin 24-hour
Mars orbit

Earth escape
April 5, 2033

Arrive Mars
October 14, 2033

Begin 102.9 days
Earth spiral escape
December 23, 2032

8.6 days
spiral capture

Thrust Arc
Coast Arc

Begin 24-hour
Mars orbit

Figure 5.—CTV and MTV trajectories.

Figure 6.—Effect of trip time on the minimum distance to the Sun.

The study requirements on the total round-trip time and stay time at Mars were not optimal for an

NEP mission and led to a close approach to the sun (0.41 AU) in the return trajectory. Such a close
approach to the sun could require additional shielding to protect the crew, the power system, and the
electronics; however, these effects were not included in this study. VARITOP was used to determine how
the trip time affected the closest approach to the Sun for varying propulsion specific masses. While the
propulsion specific mass does not have much of an effect, figure 6 shows that increasing the trip time
actually decreases the closest distance to the Sun. Reducing the trip time could be accomplished by
increasing power and/or reducing I

, however, this would raise the power or propulsion system masses.

These trade-offs were not performed as part of this study. Since the MTV passes within the orbit of Venus
(0.7233 AU), a Venus flyby may be able to improve the trip time and/or the closest approach to the Sun.
A preliminary investigation using a Venus flyby, subject to the RASC mission constraints, did not show
any improvement to the trip time. Similarly, another mission profile using a stay time at Mars longer than
60 days might allow for a more favorable return trajectory, but this was beyond the scope of this study,
since the RASC mission requirements did not allow for changing the Mars stay-time.

0.35

0.4

0.45

0.5

0.55

500

550

600

650

700

Time From Earth Departure (Days)

Min

imu

m Distan

o Su

n (AU)

alfa=10

alfa=15

alfa=20

alfa=25

NASA/TM—2006-214106

Mission Mass Summary

The mass breakdown for the MTV and CTV is shown in table 4. A common contingency factor of

25 percent was applied to all dry masses other than the RASC payloads in this analysis in order to get a
final mass with contingency. A 1 percent contingency was applied to the LH2 propellant masses.

TABLE 4.—MASS AND POWER SUMMARY FOR THE MTV

MTV Mass (MT)

CTV Mass (MT)

Subsystem

Baseline Contingency Total Baseline Contingency Total

Power system

Reactor (liquid metal)—2 x 5 MW

18.1

4.5 22.6 5.0 1.2

6.2

Brayton conversion system

8.7 2.2

10.9

4.4 1.1 5.5

Radiator

33.5 8.4

41.8

16.7 4.2 20.9

Power management and distribution

20.5

5.1 25.6

9.7 2.4 12.1

Propulsion system

MPD thrusters—4 active, 4 spares 2.4 0.6 3.0

0.9 0.2 1.1

PPU, radiators, electrical lines,

etc.

23.3 5.8

29.1

11.7 2.9 14.6

Tank details

LH2 tanks—six 7.6 m x 19 m 41.5

10.4

51.9

15.6

3.9

19.5

LH2 feed lines

3.7

0.9

4.6

2.4

0.6

3.0

Tank attachments

1.2

0.3 1.5

1.2 0.3 1.5

Refrigeration System

6.0 1.5 7.5

LH2 propellant

276.6

2.8 279.4

63.3 0.6

63.9

Structure

5 m square box truss (total of

52) 13.9 3.5

17.4

11.2 2.8 14.0

Radiator connection to the truss 0.8

0.2

0.9

0.7

0.2 0.9

Artificial gravity balance

2.0 0.5 2.5

N/A N/A N/A

ECRV docking structure

8.0 2.0

10.0

N/A N/A N/A

Vehicle communications/avionics 0.2

0.1

0.3

0.2

0.1

0.3

RCS system, tanks, thrusters, prop. 7.0

1.8

8.8

N/A

RASC payload

Trans habitat

34.96

N/A

ECRV -

7.0

N/A

Additional Consumables for 67 days

1.0

N/A

Phobos/Deimos landers

N/A N/A

N/A

- 42.5

Total NEP stage mass

190.7 47.7

238.4

85.62

21.4

107.0

Total NEP stage mass with payload 233.7

47.7

281.4

127.12

21.4

149.5

Total NEP wet mass with payload

560.7

213.5

Conclusion

This study has developed a space transportation architecture based on high-power nuclear electric

propulsion using Brayton power conversion and magnetoplasmadynamic propulsion to support a manned
Mars mission. The architecture consists of a cargo transfer vehicle with one 5 MW

fission reactor and a

Mars transfer vehicle with two 5 MW

fission reactors. The RASC study fixed the mission parameters to

investigate the performance of 4 different technological approaches to accomplishing the mission.
Unfortunately, these mission parameters were not optimal for this architecture and led to several
difficulties. The Mars Transfer Vehicle makes a very close approach to the sun (0.41 AU) during the
return trajectory in order to rendezvous with Earth. Adjusting the stay-time at Mars and/or utilizing a
Venus flyby could be used to increase this distance. The trajectory sequence requires the MTV to begin
thrusting before the CTV; however, this occurs before the crew is on board. The MTV similarly departs
Earth and begins the heliocentric portion of the flight before the CTV, even though the CTV arrives at
Mars, first. A better approach would be to launch the CTV on an earlier opportunity than the MTV to pre-
deploy the cargo in the proper orbit before any resources associated with the MTV are launched.

NASA/TM—2006-214106

References

Borowski, S. K., McGuire, M. L., Mason, L. S., Gilland, J. H. and Packard, T. W., “Bimodal NTR

Propulsion for an Artificial Gravity HOPE Mission to Callisto,” in proceedings of

Space Technology

and Applications International Forum (STAIF-2003)

, edited by M. El-Genk, AIP, New York, 2003.

Borowski, S. K., Packard, T. W., McCurdy, D. R., “Crewed Orbiter/Moon Survey Mission to Mars using

‘Bimodal’ NTR Propulsion,” in these proceedings of

Space Technology and Applications

International Forum (STAIF-2006)

, AIP, New York, 2006.

Joosten, B.

K.,

“

Preliminary Assessment of Artificial Gravity Impacts to Deep-Space Vehicle Design”,

JSC Document no. EX-02-50, NASA Johnson Space Center, Houston, CA, 2002.

Mason, L. S., “A Comparison of Brayton and Stirling Space Nuclear Power Systems for Power Levels

from 1 Kilowatt to 10 Megawatts,” NASA/TM 2001-210593, NASA Glenn Research Center,
Cleveland, Ohio, 2001.

McGuire, M. L., Borowski, S. K., Mason, L. S., and Gilland, J. H., “High Power MPD Nuclear Electric

Propulsion (NEP) for Artificial Gravity HOPE Missions to Callisto,” in proceedings of

Space

Technology and Applications International Forum (STAIF-2003)

, AIP, New York, 2003.

Siamidis, J., Mason, L., “A Comparison of Coolant Options for Brayton Power Conversion Heat

Rejection Systems,” in these proceedings of

Space Technology and Applications International Forum

(STAIF-2006)

, AIP, New York, 2006.

Williams, S. N., “An Introduction to the Use of VARITOP, A General Purpose Low-Thrust Trajectory

Optimization Program,” Jet Propulsion Laboratory, JPL D-11475, 24 January 1994.

NASA/TM—2006-214106

Appendix—Nomenclature

CTV

Cargo transfer vehicle

ECRV

Earth crew return vehicle

ELV

Expendable launch vehicle

HSHX

Heat source heat exchanger

Specific

impulse

LEO

Low Earth orbit

LH2 Liquid

hydrogen

MPD Magnetoplasmadynamic
MT

Metric ton (1000 kg)

MTV

Mars transfer vehicle

NEP

Nuclear electric propulsion

PLR

Parasitic load radiator

PPU

Power processing unit

RASC

Revolutionary aerospace systems concepts

RCS

Reaction control system

TEI Trans-Earth

injection

TMI Trans-Mars

injection

TransHab Transportation

habitat

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

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E–15441

WBS 463169.03.03

Use of High-Power Brayton Nuclear Electric Propulsion (NEP)
for a 2033 Mars Round-Trip Mission

Melissa L. McGuire, Michael C. Martini, Thomas W. Packard, John E. Weglian,
and James H. Gilland

Cargo; Trajectories; Flyby missions; Aerospace systems; Nuclear electric propulsion;
Space transportation; Manned Mars missions; Mars missions; High temperature;
Brayton cycle

Unclassified - Unlimited
Subject Category: 13

Prepared for the Space Technology and Applications International Forum (STAIF–2006) sponsored by the University
of New Mexico’s Institute for Space and Nuclear Power Studies (UNM-ISNPS), Albuquerque, New Mexico, February
12–16, 2006. Melissa L. McGuire, NASA Glenn Research Center; Michael C. Martini and Thomas W. Packard, Analex
Corporation, 1100 Apollo Drive, Brook Park, Ohio 44142; and John E. Weglian and James H. Gilland, Ohio Aerospace
Institute, 22800 Cedar Point Road, Brook Park, Ohio 44142. Responsible person, Melissa L. McGuire, organization
code PBM, 216–977–7128.

The Revolutionary Aerospace Systems Concepts (RASC) team, led by the NASA Langley Research Center, is tasked with
exploring revolutionary new approaches to enabling NASA to achieve its strategic goals and objectives in future missions.
This paper provides the details from the 2004-2005 RASC study of a point-design that uses a high-power nuclear electric
propulsion (NEP) based space transportation architecture to support a manned mission to Mars. The study assumes a
high-temperature liquid-metal cooled fission reactor with a Brayton power conversion system to generate the electrical
power required by magnetoplasmadynamic (MPD) thrusters. The architecture includes a cargo vehicle with an NEP
system providing 5 MW of electrical power and a crewed vehicle with an NEP system with two reactors providing a
combined total of 10 MW of electrical power. Both vehicles use a low-thrust, high-efficiency (5000 sec specific impulse)
MPD system to conduct a spiral-out of the Earth gravity well, a low-thrust heliocentric trajectory, and a spiral-in at Mars
with arrival late in 2033. The cargo vehicle carries two moon landers to Mars and arrives shortly before the crewed
vehicle. The crewed vehicle and cargo vehicle rendezvous in Mars orbit and, over the course of the 60-day stay, the crew
conducts nine-day excursions to Phobos and Deimos with the landers. The crewed vehicle then spirals out of Martian
orbit and returns via a low-thrust trajectory to conduct an Earth flyby. The crew separates from the vehicle prior to Earth
flyby and aerobrakes for a direct-entry landing.