Carolyn R. Mercer, Steven R. Oleson, Eric J. Pencil, Michael F. Piszczor, Lee S. Mason, 

Kristen M. Bury, David H. Manzella, Thomas W. Kerslake, and Jeffrey S. Hojnicki

Glenn Research Center, Cleveland, Ohio

John P. Brophy

Jet Propulsion Laboratory, Pasadena, California

Benefits of Power and Propulsion Technology for a 

Piloted Electric Vehicle to an Asteroid

NASA/TM—2012-217274

February 2012

AIAA–2011–7252
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Carolyn R. Mercer, Steven R. Oleson, Eric J. Pencil, Michael F. Piszczor, Lee S. Mason, 

Kristen M. Bury, David H. Manzella, Thomas W. Kerslake, and Jeffrey S. Hojnicki

Glenn Research Center, Cleveland, Ohio

John P. Brophy

Jet Propulsion Laboratory, Pasadena, California

Benefits of Power and Propulsion Technology for a 

Piloted Electric Vehicle to an Asteroid

NASA/TM—2012-217274

February 2012

AIAA–2011–7252

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Space 2011 Conference and Exposition

sponsored by the American Institute of Aeronautics and Astronautics

Long Beach, California, September 27–29, 2011
Acknowledgments

The authors gratefully acknowledge the efforts of the entire COMPASS team at the NASA Glenn Research Center: Les Balkanyi, 

Anthony Colozza, John Dankanich, Jon Drexler, Robert Falck, James Fincannon, James Fittje, John Gyekenyesi, Geoffrey Landis, 

Mike Martini, Tom Packard, Thomas Parkey, Paul Schmitz, Jeff Woytach, Joe Warner, Glenn L. Williams, Anita Tenteris, and 

especially Melissa McGuire and Carl Sandifer for their diligent efforts in keeping things straight. Thanks also go to Sal DiStefano at 

the Jet Propulsion Laboratory and to the Exploration Technology Development and Demonstration Program within the Exploration 

Systems Mission Directorate.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217274 1

Benefits of Power and Propulsion Technology for a

Piloted Electric Vehicle to an Asteroid

Carolyn R. Mercer, Steven R. Oleson, Eric J. Pencil, Michael F. Piszczor, Lee S. Mason, 

Kristen M. Bury, David H. Manzella, Thomas W. Kerslake, and Jeffrey S. Hojnicki

National Aeronautics and Space Administration 

Glenn Research Center

Cleveland, Ohio 44135

John P. Brophy

National Aeronautics and Space Administration

Jet Propulsion Laboratory 

Pasadena, California 91109

Abstract

NASA’s goal for human spaceflight is to expand permanent human presence beyond low Earth orbit

(LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission 

options include crewed destinations to LEO and the International Space Station; high Earth orbit and 

geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and 

the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference 

missions to drive out required functions and capabilities for these destinations, focusing first on a piloted 

mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion 

stage could reduce mission cost by reducing the required number of heavy lift launches and could 

increase mission reliability by providing a robust architecture for the long-duration crewed mission.

Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, 

landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power 

and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of 

the assessment was to determine technology drivers to advance the state of the art of electric propulsion 

systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion 

and power systems were done to determine potential system-level impacts of improved technology.

Starting with a “reasonable vehicle configuration” bounded by an assumed launch date, we introduced 

technology improvements to determine the system-level benefits (if any) that those technologies might 

provide. The results of this assessment are discussed and recommendations for future work are described.

1.0 Introduction

NASA’S goal for human spaceflight is to expand permanent human presence beyond low Earth orbit

(LEO). To achieve this goal, NASA is identifying potential missions and technologies needed to conduct 

those missions safely and cost effectively. Mission options include piloted destinations to LEO and the 

International Space Station (ISS); high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit,

and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of 

Mars. Through a process known as the Human Exploration Framework Team (HEFT) architecture 

planning, NASA generated a series of design reference missions to drive out required functions and 

capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One 

conclusion from this exercise was that a solar electric propulsion (SEP) stage could significantly reduce 

mission cost by reducing the required number of heavy lift launches, and could increase mission 

reliability by providing a robust architecture for the long-duration piloted mission. Similarly, solar electric 

vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, 

and visiting its moons (Ref. 1).
NASA/TM—2012-217274 2

This paper describes the parameterized assessment of power and propulsion technologies for a piloted 

solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine nominal 

impacts of technology improvements on vehicle-level mass and cost. One benefit of this vehicle-level 

approach is that it captures the interrelationships between multiple subsystems, rather than simply 

quantifying the mass savings from a technology on a single subsystem. 

Details of the design reference mission and the concept vehicle are described, including key design 

choices and all propulsion and power variants. The key system-level impacts for each propulsion and 

power technology are summarized, along with the relative costs of developing these technologies and 

using them for flight hardware. Although the results are not definitive because the full breadth of design 

space was not explored nor were the design impacts of contingency operations, we believe that they are 

representative and provide insight into the relative benefits of power and propulsion technologies for solar 

electric vehicles of this class. This work can help guide technology development investments to enable 

future missions to near-Earth asteroids and beyond.

2.0 Design

Asteroid 2008–EV5 was chosen as the representative mission because it approaches Earth relatively 

soon and is sufficiently large to be of scientific and engineering interest. The launch opportunity for this 

destination is 2024. Further details about this mission are described in Section 2.1, as well as a concept 

vehicle created by a previous study. Our figures of merit and guiding design principles are described in 

Section 2.2, and choices for the baseline vehicle are described in Section 2.3. Oleson et al. provide a

detailed description of the baseline vehicle and design methodology (Ref. 2).

2.1 Design Reference Mission

The scenario of interest requires a solar electric stage to transfer cargo from LEO to the Earth-Moon 

libration point L1 (E–M L1) and to transfer crew to and from the asteroid. The cargo consists of a deep 

space habitat and space exploration vehicle. A 100-mt vehicle launches the SEP stage, cargo, and a kick 

stage; the kick stage raises the SEP stage with its cargo to a 400-km circular orbit. The SEP stage brings 

the cargo to L1 and docks with a previously-positioned cryogenic propulsion stage and a crew transfer 

vehicle. The cryostage provides the high-thrust delta-V needed for Earth escape velocity and is then 

jettisoned. The SEP stage delivers the crew and cargo to the asteroid, and then propels the crew transport 

vehicle to a hyperbolic return to Earth. This operations concept is shown in Figure 1. Note that this 

mission has two operational modes: an unpiloted mission from LEO to an Earth staging orbit, followed by 

a piloted mission up to the asteroid and return to Earth.

The primary constraints are a piloted trip time of 400 days or less to prevent excessive exposure to 

ionizing radiation, and mass and volume sufficiently small to permit the launch of the SEP stage together 

with other required vehicles in a single 100-mt launch vehicle.

For the initial assessment of this design reference mission, the HEFT developed an SEP stage concept 

design. Their design employed two planar solar array panels producing 320 kW at end of life (EOL) using 

33 percent efficient cells; eight hall effect thrusters with 50 kW maximum power; a thruster boom to 

minimize plasma impingement on the arrays; xenon stored in eight cylindrical tanks; and a power 

processing unit cooled by heat pipe radiators (Ref. 3).
NASA/TM—2012-217274 3

Figure 1.—Concept of operations for a piloted mission to a near-Earth asteroid (Ref. 3). The solar electric stage used 

for the crew (SEP2) is circled in red.

2.2 Design Approach

To conduct our parametric assessment of propulsion and power technologies, the Collaborative 

Modeling for Parametric Assessment of Space Systems (COMPASS) (Ref. 4) team at the NASA Glenn 

Research Center started with a clean sheet design using the following figures of merit:

• Safely deliver crew to/from asteroid

• Single fault tolerance to loss of crew and loss of mission

○ Nominal mission time of 800 days with piloted trip time <400 days 

○ 2023 launch date

• Mass and volume 

○ 100-mt launch vehicle with 8.5- by 25-m shroud 

○ Nominally 47 mt allocated for SEP stage

• Performance

○ Target 300 kW of power to the propulsion system

○ Provide 15 kW of power to the payload (deep space habitat), except during eclipse

○ Provide 5 kW of power for housekeeping loads

○ Withstand cryogenic kick stage propulsive maneuvers up to 0.1-g steady-state longitudinally 

and docking event loading up to 0.02 g longitudinally and 0.01 g transverse

• Low cost, simple, and robust

○ Use commercial-off-the-shelf components when available
NASA/TM—2012-217274 4

The spacecraft is designed to be single fault tolerant in the design of the subsystems, where possible.

Exceptions to this include the electric power system, propellant tanks, and radiators that have zero fault 

tolerance, although they are designed to accommodate some performance degradation. Note that because 

contingency operations are not included in this analysis, conclusions about the relative merits of 

parameterized power and propulsion technologies must be treated as preliminary. 

Mass growth calculations were conducted according to AIAA S–120–2006, “Standard Mass 

Properties Control for Space Systems.” The percent growth factors specified in this standard were applied 

to each subsystem before an additional growth was carried at the system level to ensure an overall growth 

of at least 30 percent on the dry mass of the entire system. Growth in the propellant mass was carried in 

the propellant calculation. A 30 percent growth factor on the bottoms-up power requirements for the bus 

subsystems was used, with a 5 percent margin for the electric thruster power requirements.

The Spacecraft N-body Analysis Program (Ref. 5) was used to conduct a tangential steering spiral 

from LEO to 150,000-km semi-major axis (approximately the same energy state as E–M L1). The 

Mission Analysis Low-Thrust Optimization interplanetary low-thrust trajectory optimization tool (Ref. 6)

was used to determine the propellant mass needed to perform the heliocentric phase of the mission.

Vendor-supplied data was used where possible for mass and volume estimates to provide realism. This 

does not imply endorsement of any particular commercial system.

2.3 Baseline Vehicle

Following this design approach of using the simplest, most mature technology that can meet design 

objectives, we selected an array of nominally 30-kW Hall thrusters and planar solar arrays with 33

percent efficient inverted metamorphic cells for the baseline configuration. Boom-mounted arrays are 

notched to provide a 45° “keep away” zone from the thrusters’ plumes. This design eliminates the need 

for power and propellant transport to boom-mounted thrusters. Power from the arrays is delivered to the 

thrusters using a direct-drive configuration (Ref. 7) rather than a conventional power processing unit, with 

both the arrays and thrusters operating at nominally 300 V. Power processing and voltage levels were 

varied during the parameterization, as was solar cell efficiency.

This mission has two phases―an unpiloted rise to E–M L1 and a piloted transit to the asteroid and 

back. For this thrust class, a spiral trajectory is required for the initial phase, followed by a heliocentric 

trajectory for the second.

A total of 10 thrusters were used, each operating at nominally 30 kW. If a thruster string fails, the 

piloted mission can still be completed by reducing some of the nominally planned coast time. The specific 

impulse (Isp) was set to 2000 sec to match the Hall thrusters to the 300-VDC direct-drive solar array 

output, and as a conservative trade between trip time and propellant mass. To a certain extent, the Isp can 

be increased to reduce the propellant mass for variants that would otherwise exceed the mass allocation. 

This option is limited for direct-drive architectures with solar arrays, however, as the thrusters must match 

the array voltage. Even higher Isp is required for nuclear power systems because of their higher inert 

mass, but this is possible because of the corresponding high alternating current (AC) voltage output from

the power system. For either solar or nuclear systems, a higher Isp (at fixed power) will reduce 

interplanetary coast times and require longer Earth spiral mission times. The Earth spiral is unpiloted so a 

longer trip time for this portion is likely acceptable. Reduced coast times for the piloted heliocentric 

transfer phase might be undesirable once abort scenarios are assessed.

Xenon was chosen as the propellant because its low ionization energy enables high thruster efficiency 

and is more easily stored than other heavy noble gases. A single, spherical composite overwrap pressure 

vessel (COPV) xenon tank at 1200 psi was selected as the baseline; the number of tanks and storage 

conditions were traded as part of the parameterization.
NASA/TM—2012-217274 5

Exposure to the Van Allen belt and scattered ions from the Hall thruster plumes is expected to 

degrade the solar arrays; we assumed an 18.8 percent EOL degradation from these sources when sizing 

the arrays. No damage was assumed to occur in heliocentric space. As the array voltage degrades, the 

current is adjusted altering the thruster mass flow rate to maintain a fixed specific impulse. 

“No-roll” steering was chosen to eliminate the considerable mass of control moment gyroscopes and 

as a result, we accepted secondary axis Sun pointing errors and the attendant power loss. Since only one 

revolute axis is available for tracking, the arrays are revolved to minimize the Sun off-pointing angle 

while under thrust. Array tracking is controlled with ISS-derived Solar Alpha Rotary Joint (SARJ)

gimbals, with mass reduced by removing the ISS in-flight servicing requirement. Our assumed launch 

date was favorable for the use of no-roll steering because the maximum angle between the arrays and Sun 

occurs early in the spiral trajectory, when the beginning-of-life (BOL) power is still available. This beta 

angle may occur later in the trajectory for different launch dates, requiring either oversized arrays or 

longer trip times because of the reduced power. Four RCS thruster pods provide roll, pitch, and yaw 

control, augmented by four gimbaled Hall thrusters.

Body-mounted radiators located directly below the solar array gimbals provide the best view for 

thermal rejection since they point roughly perpendicularly to the Sun’s direction. All components of the 

thermal system were sized for the worst-case environmental conditions (LEO), with no redundancy.

Micrometeoroid and orbital debris shielding was used to protect critical systems such as the propellant 

tanks and exposed heat pipes. The thrusters are recessed into the cylindrical structure to provide 

protection. Shielding by the aluminum structure is expected to be sufficient to protect electronics from 

radiation. The composite thrust tube design was sized to carry the mass of the deep space habitat and 

space exploration vehicle during launch. All communications are assumed to be performed by the deep 

space habitat, including relaying housekeeping commands and data for the SEP module.

This baseline SEP stage is shown in Figure 2 and the key design features are listed in Table 1. Note 

that a rollout solar array based on mass and structure data provided by Deployable Space Systems, Inc. 

(DSS) was used as the baseline in part because other NASA studies have used structures based on data 

provided by other vendors (Ref. 8). NASA relies on vendor-provided data to add realism to these concept 

designs and does not endorse any particular approach.

The baseline configuration configured for launch is shown in Figure 3 and in its stowed position is 

shown in Figure 4. The cargo elements from bottom to top are kick stage, SEP stage, deep space habitat, 

and space exploration vehicle. The adapter to join the kick stage to the launch vehicle is not shown. 

Table 2 lists the nominal masses for each element as determined by the HEFT architecture study.

Figure 2.—Solar electric propulsion stage—baseline configuration.
NASA/TM—2012-217274 6

TABLE 1.—KEY FEATURES OF BASELINE SEP STAGE

S/C Element Description

Mission/Operations Phase 1: ~1 year unpiloted spiral to Earth-Moon L1, ~6 km/s

Phase 2: <400 days piloted heliocentric transit to NEO 2008–EV5, ~4 km/s

~10 km/s Total ∆V

Launch 100-mt heavy lift launch vehicle with 8.5- by 25-m shroud

Design Margins • 10 percent launch margin, 30 percent power growth except thrusters (5 percent power growth)

• 5 percent propellant margins, 95 percent SEP duty cycle

• Growth added to reach system level 30 percent

Guidance, Navigation, 

and Control (GN&C)

Star trackers, inertial measurement units (IMUs), monopropellant reaction control system (RCS), no-roll 

steering during spiral

Power 381 kW BOL, 300 V planar solar arrays with 33 percent efficient cells, arrays notched to avoid 45°

thruster plume

Propulsion Ten 30-kW Hall thrusters using direct drive at 2000 sec Isp, four thrusters gimbaled; one 3-m-diameter 

COPV xenon tank

Structures and 

Mechanisms

Composite thrust tube; ISS-derived solar array gimbals

Communications All communications assumed to reside on Deep Space Hab vehicle

Command and Data 

Handling (C&DH)

Shielded system to control SEP module, housekeeping relayed by habitat

Thermal Simple body-mounted radiators, ~350 kg of micrometeoroid orbital debris (MMOD) shielding

TABLE 2.—LAUNCH MASSES OF 

EACH ELEMENT IN MISSION

Element Mass 

(mT)

Space Exploration Vehicle ................................................ 6.7000

Deep Space Hab ................................................................ 24.363

Solar Electric Propulsion Stage ......................................... 46.985

Service Module Kick Stage................................................. 8.173

Launch Vehicle Adapter...................................................... 2.095

Cargo Element................................................................... 88.316

Margin................................................................................. 1.684

Cargo Element With Margin ............................................. 90.000

Manager’s Reserve............................................................ 10.000

Heavy Lift Launch Vehicle (HLLV) Lift Capability......... 100.00

Figure 3.—Launch configuration. Figure 4.—Baseline vehicle in stowed configuration.
NASA/TM—2012-217274 7

3.0 Parametric Assessments of Power and Propulsion

The baseline vehicle employed 300 V planar solar arrays with 33 percent efficient cells producing 

381 kW at BOL. Four power variants were studied to determine their effect on vehicle mass and cost:

planar arrays versus 2X concentrator arrays; 300 V versus 120 V arrays; solar cell efficiencies ranging 

from 29 percent to 37 percent; and a nuclear fission reactor versus a solar array.

The baseline vehicle employed ten 30-kW thrusters, direct-drive power management, and a single 

xenon tank. Four propulsion variants were studied to determine their effect on vehicle mass and cost: ten 

30-kW thrusters versus seven 50-kW thrusters; direct-drive versus power processor units (PPUs); a single 

xenon tank versus four tanks; and supercritical versus cooled supercritical xenon storage.

A summary of each variant is provided in Table 3, showing each case’s configuration, trip time, inert

mass, propellant mass, and wet mass with growth. The percentage dry mass savings for each variant is 

presented as compared to the case based on a planar array operating at the lowest voltage with a 

conventional power processing unit. Note that while the inert mass of the nuclear powered system (Case 

VIII) is larger than the solar power systems, the total launch masses are comparable. Note further that the 

wet mass for any configuration must not exceed the total amount allocated for the SEP stage plus margin 

(48.669 mt) listed in Table 2, and should properly allow ~4 mt of margin for a target SEP stage wet mass 

of 44.7 mt. A description of each variant is provided in the following two sections. Full details can be 

found in the COMPASS final reports by Oleson et al (Refs. 2 and 9).

3.1 Power Trades

The following four power parameters were studied to determine vehicle mass and cost impacts: solar 

concentrator versus planar arrays, voltage level, cell efficiency, and power source.

The planar array used in the baseline vehicle described in Section 2.3 assumed 33 percent efficient 

cells provided 381 kW at BOL to provide 300 kW to the thrusters and 20 kW for the deep space habitat 

and housekeeping loads at EOL. Array dimensions are shown in Figure 2. Array mass and stowed volume 

data were obtained from DSS, Inc., and the vehicle based on this array was found to fit within the launch 

vehicle and complete the mission on time. The mass of this baseline power subsystem is shown in Table 3 

as Case I.

Changing the bus voltage to 120 V has two primary impacts: the harness mass increases to 

accommodate higher currents and the use of conventional power processing units becomes required 

because of the thrusters’ narrow operating voltage range. As shown in Table 3, Case IIIa, these changes 

resulted in a vehicle that did not fit the mass requirement for launch. A 120-V SEP stage can be built for 

this mission, however, by increasing the specific impulse to 2100 sec thereby reducing the mass of xenon 

propellant. As shown in Table 3, Case IIIb, this increases the trip duration by 64 days over the baseline, 

although the piloted portion is only increased by 3 days and fits within the 400-day constraint.

The impact on vehicle mass from varying the cell efficiencies is shown in Table 4. As would be 

expected, higher efficiency cells reduce the overall mass of the vehicle by directly reducing the area of the 

array. For each kilogram of mass reduction in the dry vehicle, 0.5 kg of xenon propellant is saved, as well 

as 0.5 kg of chemical propellant for the cryopropulsion stage. The savings of going to 35 percent cells 

from 33 percent cells is 166 kg of inert mass (with growth), and therefore about 249 kg of total SEP wet 

mass. When considering the impact of cell efficiency, higher values are always better from a mass 

perspective. However, as shown in Figure 7, the single largest cost of a SEP stage is the solar arrays, and 

the largest cost component of the arrays is the cells. Therefore, we looked to see if our mission goals 

could be met with lower cost, lower efficiency cells. As shown in Table 4, cells with 29 percent efficiency 

should be good enough to meet the vehicle mass target and could substantially reduce vehicle cost.
NASA/TM—2012-217274 8

TABLE 3.—SUMMARY OF PROPULSION AND POWER VARIANTS.

PRIMARY VARIANTS ARE SHOWN IN RED FOR EACH CASE.

Piloted portion of trip is heliocentric phase.

*Radiators were configured as needed for each case.

**Isp was increased to meet mass target.

TABLE 4.—CELL EFFICIENCY TRADE
NASA/TM—2012-217274 9

Figure 5.—Nuclear electric concept vehicle.

A 2X concentrator array based on “pop-up” flexible reflectors was designed to reduce the total area of 

photovoltaic cells. The areal size of the array must increase slightly (~10 percent) to account for the 

higher operating temperature of the concentrator cells while collecting sufficient solar flux. The 

concentrators lower the mass of the power system by about 6 percent, and they can potentially lower the 

cost of arrays by replacing high-cost solar cells with lower cost reflective elements. It is difficult to assess 

this cost savings because the concentrators will add complexity that will have some associated costs. The 

pointing requirements needed to maintain full illumination were sufficiently lax to maintain the ability for 

no-roll steering, so no other changes to the baseline configuration were required. The mass breakdown for 

this variant is shown in Table 3 as Case IV.

Our final power variation assessed the impact of using a nuclear reactor instead of a solar array. This 

concept vehicle is shown in Figure 5. The 300 kW of electrical power is created by a uranium-fueled 

fission reactor coupled to two 100 percent redundant Brayton power converters. The Brayton alternators 

deliver three-phase AC power at 480 V that is rectified in a direct-drive unit to power ten 36-kW Hall 

thrusters generating 2640 sec of specific impulse with xenon propellant contained in a single, 3-mdiameter COPV tank. A simple body-mounted and single-hinge deployable radiator system dissipates the 

Brayton waste heat. A telescoping boom provides separation between the reactor and the payload, and a 

truncated cone radiation shield with a 22° half angle protects the vehicle and crew from nuclear radiation. 

As shown in Table 3, Case VIII, the nuclear power system is heavier than the solar power system, but this 

mass is compensated by a reduced xenon propellant mass enabled by the higher Isp. Even with the higher 

Isp, the nuclear vehicle meets the trip time requirement, primarily because it provides continuous thrust 

even during the occulted portion of the spiral out. Development (after technology readiness level (TRL)

6) and recurring costs are expected to be about 30 percent higher than a comparable solar electric vehicle, 

but development costs incurred for this mission could be directly applicable to higher power missions 

needed for cargo and crewed missions to Mars for instance (Refs. 10 and 11).

3.2 Propulsion Trades

The following four propulsion parameters were studied to determine vehicle mass and cost impacts: 

thruster power level, power processor design, propellant tank size, and propellant storage conditions.

The baseline concept vehicle described in Section 2.3 used ten 30-kW-rated Hall thrusters. No spares 

were provided since losing a thruster string will only reduce the coast time for the piloted portion of the 

trip. Direct-drive units were connected to each thruster. No cross-strapping was considered. Four of the 

thrusters were gimbaled and mounted at the perimeter of the stage’s bus structure to augment the attitude 

control provided by the RCS thrusters; the remaining six were statically mounted and clustered in the 

center as shown in Figure 6(a). Ten direct-drive units (one for each thruster) were mounted on the inside 

of the bottom deck. The first propulsion variant increased the size of each thruster to provide nominal 

operation at 50 kW. Six of these thrusters, plus one spare, and the associated seven direct-drive power 
NASA/TM—2012-217274 10

Figure 6.—Propulsion variants.

processors, could easily fit within the available space and mass constraints so no other changes to the 

baseline configuration were needed. A spare was included because an engine-out condition using only six 

thrusters would yield a trip time that exceeds allowable mission constraints. This variant is shown in 

Figure 6(b) and the resultant vehicle mass is shown in Table 3 as Case V. Note that the inert mass 

increased slightly as compared to the baseline, primarily because a spare thruster is required. Although 

there is no mass benefit, there may be a cost savings associated with using fewer units because of reduced 

qualification testing and parts count.

Direct-drive power processing units (DDUs) were used in the baseline configuration. The use of 

DDUs is only possible when the arrays produce a voltage within the operating range of the thrusters, 

nominally 300 to 600 V. Outside of that range, conventional power PPUs are required. The use of PPUs 

instead of DDUs increases the system mass as is shown in Table 3, Case II. The increased mass arises 

primarily from adding isolation transformers to regulate the voltage generated by the solar arrays to match 

that needed by the thrusters, and bigger radiators to reject the additional heat generated by the less 

efficient PPUs (~95 percent efficient PPU vs. 99 percent efficient DDUs). PPU mass was assumed to be 

100 kg each, although they may be as low as 88 kg. The mass difference of 12 kg each is small relative to 

the vehicle mass. Note that while the DDUs reduce system mass, they potentially increase operational risk 

because of their inability to operate over wide voltage swings.

To save mass, the baseline configuration used a single 3-m-diameter spherical COPV tank to store the 

xenon propellant as a supercritical gas at 1200 psia. A tank this large may be difficult to fabricate, so we 

studied the impact of replacing the single tank with four smaller COPV cylindrical tanks. To fit into the 

baseline stage structure, the maximum allowable diameter for each of the four tanks is 1.62 m. The 

overall length of the tanks is 2.83 m, easily fitting within the baseline structure as shown in Figure 6(c).

Additional structure was added to the inside of the stage structure to handle the polar boss tank mounting 

for the tanks. As shown in Table 3, Case VIa, this design did not meet mission mass requirements 

primarily because of the increased surface area (and hence mass) to store the same volume, increased wall 

thickness in the cylindrical portion of the tanks, and additional structure. Increasing the specific impulse 

of the thrusters to 2100 sec can compensate for this mass overrun at the expense of piloted trip time.
NASA/TM—2012-217274 11

To reduce the mass required by the four cylindrical supercritical tanks, active cooling was added to 

reduce the tank pressure and thus reduce the required wall thickness. The maximum allowable diameter 

for each of the four tanks was kept at 1.62 m, allowing for 10 cm of foam insulation and 1.5 cm of 

multilayer insulation. This increased the overall tank length to 3.79 m, and required the addition of two 

small buffer tanks and associated mounting structure (Fig. 6(d)). No other changes to the baseline 

configuration were needed. This configuration was more massive than the single tank configuration, but 

fit within the mass allocation as shown in Table 3, Case VII.

4.0 Results

A summary of the mass breakdown for the baseline SEP stage and each propulsion and power variant 

is shown in Table 3. Out of the eight variations studied, only two designs could not fit within the mass 

and volume constraints levied by the mission using thrusters operating at 2000 sec Isp: the 120 V bus 

voltage and the four supercritical xenon tanks. The mass associated with a 120-V spacecraft bus was high 

because the bus harness mass was eight times larger than that required for the high-voltage (300-V) 

system, and the use of conventional power processing units was required to boost the voltage to levels 

required by the electric thrusters. Multiple smaller tanks increased mass because the tanks were heavier. 

Bus voltages as low as 120 V and supercritical xenon storage in smaller tanks can, however, be used for 

this mission when the specific impulse is boosted by 5 percent to 2100 sec. The crewed portion of the trip 

still meets the stated requirement of 400 days needed to keep human exposure to harmful radiation within 

allowable limits. As is usually found with electric propulsion, the Isp “knob” can be tuned to meet the 

launch mass at the expense of trip time. For this specific mission, the Isp tuning was fairly small 

(hundreds of seconds) and the required mission times could be met. However, once abort scenarios are 

considered, this ability to increase specific impulse may not be possible.

Each of the other variants can accomplish the mission within mass and volume limits, subject to the 

caveat stated at the outset: contingency operations were not considered and thus these are only design 

concepts, not true designs. Given that, the relative benefits of the various trades are summarized in Table 

5. In this table, technologies in the first column (technology “A”) are compared to the technologies in the 

second column (technology “B”) for vehicle mass, technology development cost to TRL 6, flight 

hardware costs (nonrecurring and recurring, assuming TRL 6 is achieved), and technical risk. The relative 

indicators include “HH” representing a very large mass savings of technology “A” over technology “B”

and “H,” “M,” and “L” for large, medium, and small mass savings. Development costs are designated as 

low, medium, or high representing additional costs required to bring technology “A” to TRL 6 relative to 

technology “B.” Flight hardware costs are marked “+” to indicate a probable savings and “–” to indicate

more expensive hardware for technology “A” as compared to technology “B.” The technical risk of 

developing technology “A” relative to technology “B” to TRL 6 is denoted high, medium, or low for 

substantially more risk, more risk, and nominally the same risk.

From this summary, one can see that the largest potential mass savings may be realized by four 

technologies: a direct-drive architecture, high-voltage solar arrays, a large in diameter spherical propellant 

tank, and an active cooling for the xenon propellant. Note that high-voltage operation is a necessary 

condition for direct-drive architectures, and the combination of the two provides over a 4 mt mass 

savings.

Moderate mass savings can be realized by increasing photovoltaic cell efficiency to 37 percent. The 

inert mass difference is small for 2X concentrator arrays and 50-kW Hall thrusters, and the difference 

between the launch mass of the nuclear and solar vehicles is also small.
NASA/TM—2012-217274 12

TABLE 5.—MASS, COST, AND RISK COMPARISONS OF POWER AND PROPULSION TECHNOLOGIES

TABLE 6.—TECHNOLOGY READINESS LEVELS OF POWER AND PROPULSION 

TECHNOLOGIES NEEDED FOR 300-kW-CLASS ELECTRIC PROPULSION SPACECRAFT

The current TRLs of the power and propulsion technologies employed during this study is shown in 

Table 6, together with a listing of critical needs still left to develop for this application and the current 

state of the art.

A cost analysis showed that the largest single contributor to the cost of the SEP stage is the solar 

arrays, and the largest contributor to the cost of the arrays is the cell/interconnect/coverglass assemblies 

and their integration into the array blanket (Fig. 7). These elements constitute nominally 33 percent of the 

stage’s flight hardware cost. Reducing this cost is by far the largest means to reducing the cost of this 

vehicle. One way to do this is to use concentrator arrays, which should provide cost savings by replacing 

high-cost photovoltaic cells by lower cost reflective elements. The 2X concentrator did not impose 

pointing requirements sufficient to require additional control hardware, so the savings should be equal to 

the difference in cost between the installed reflective elements and half of the installed solar cells.

Increasing the efficiency of the cells will reduce costs by reducing the number of cells. However, lower 

efficiency cells may cost substantially less than higher efficiency cells, so there may be no net benefit or 

the lower cost cells may produce a lower vehicle cost.
NASA/TM—2012-217274 13

Further savings should be realizable by using direct-drive power processors because they eliminate 

the need for a number of high-voltage components. The cost of a single, large xenon tank will be higher 

than the smaller tanks, but the large mass savings may warrant the expense. The cost of developing a 

nuclear reactor from TRL 6 and deploying it is estimated to be roughly 30 percent higher than the 

corresponding solar arrays, but this cost could be justified because the foundation of such a system would 

provide for future missions to Mars.

Figure 7.—Mass and cost breakdown for baseline vehicle (Case I).
NASA/TM—2012-217274 14

5.0 Conclusion

A 300-kW-class SEP vehicle concept was created to assess the system-level mass and cost impacts of 

power and propulsion technologies. The following technologies were assessed: electric thruster power 

level, power processing architecture, propellant storage, bus voltage, solar array concentration, solar cell 

efficiency, and power generation (solar vs. nuclear).

We found that this specific design reference mission could not be accomplished without using a highvoltage bus (nominally 300 V) unless the Isp was increased and the trip time was subsequently increased. 

Increased Isp’s can be used to reduce mass instead of adopting the other technologies considered in this 

study but may not be feasible in abort scenarios and would not reduce recurring costs.

The technologies able to most significantly reduce mass are a high-voltage solar array and power bus, 

direct drive, and large xenon tanks. Each of these require technology development to bring to flight 

readiness.

Moderate mass savings could be realized by storing the xenon propellant as a liquid and by increased 

photovoltaic cell efficiency. The 2X concentrator arrays and 50-kW Hall thrusters provide at most a small 

mass improvement, and the difference between the launch mass of the nuclear and solar vehicles is also 

small.

By far the largest contributor to the cost of the arrays is the cell/interconnect/coverglass assemblies 

and their integration into the array blanket, which constitute nominally 33 percent of the stage’s flight 

hardware costs (from TRL 6 to flight hardware). Direct-drive processors can also reduce costs by 

eliminating a number of high-voltage components. A nuclear reactor would cost nominally 30 percent

more than solar arrays, but the cost could be justified because of the foundation such a system would 

provide for future missions to Mars.

There are limits to the results of these studies. If a different mission is selected, the results will change.

If contingency operations are factored in, the results will change. A different suite of technologies could 

also be considered, again possibly changing the results. However, we believe that these results are 

representative and provide insight into the relative benefits of key power and propulsion technologies for 

solar electric vehicles of this class.

In addition to understanding the system-level benefits of particular technology advancements as 

summarized in Table 3, the feasibility, cost, and time to develop must also be considered when 

determining investment priorities. We believe that all of the technologies listed in Table 6 have a feasible 

path for development, and the time needed to bring them to TRL 6 scales roughly inversely with their 

current TRL level, excluding the time needed for life testing. Relative costs for both technology 

development and flight hardware are estimated in Table 5. A robust technology development program 

that attacks any subset of these technologies will substantially improve the state of the art of electric 

propulsion stages, laying the groundwork for a new capability for commercial and exploration 

opportunities in space.

References

1. NASA Exploration Systems Mission Directorate presentation, “Human Space Exploration 

Summary,” http://www.nasa.gov/pdf/525162main_HEFT_Final_Brief_508_20110309.pdf, January 

2011. [cited 19 Aug 2011].

2. S. Oleson, et al., “COMPASS Final Report: Exploration Technology Development and 

Demonstration (ETDD) Program Design Reference Mission (DRM) #1: Piloted Solar Electric 

Propulsion (SEP) to a Near Earth Orbiting (NEO) Asteroid,” NASA CD–2011–56 (to be published 

as a NASA TM).

3. J.R. Brophy, R. Gershman, N. Strange, D. Landau, R.G. Merrill, T. Kerslake, “300-kW Solar 

Electric Propulsion System Configuration for Human Exploration of Near-Earth Asteroids,” AIAA–

2011–5514, Presented at the 47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, August 1, 

2011.
NASA/TM—2012-217274 15

4. Collaborative Modeling for Parametric Assessment of Space Systems 

http://www.grc.nasa.gov/WWW/compass/ [cited 19 Aug 2011]. 

5. “Rapid Calculation of Spacecraft Trajectories Using Efficient Taylor Series Integration,” NASA 

Tech Briefs, Jan. 1, 2011 http://www.techbriefs.com/component/content/article/9031 [cited 19 Aug 

2011].

6. The In-Space Propulsion Technology Project Low-Thrust Trajectory Tool Suite: MALTO 

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080047350_2008047178.pdf [cited 19 Aug 

2011].

7. D. Hoffman, et al., “Concept Design of High Power Solar Electric Propulsion Vehicles for Human 

Exploration,” Presented at the 62nd International Astronautical Congress, Cape Town, South Africa, 

Oct 3-7, 2011, IAC-11-D2.3.5.

8. L.A. Capadona, et al., “Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: Issues 

and Solutions,” AIAA, Presented at the AIAA Space 2011 Conference, Long Beach, CA, September 

27–29, 2011.

9. S. Oleson, et al., “COMPASS Final Report: Exploration Technology Development and 

Demonstration (ETDD) Program Design Reference Mission (DRM) #2: Piloted Solar Electric 

Propulsion (SEP) to a Near Earth Orbiting (NEO) Asteroid Photovoltaic and Propulsion Trades,” 

NASA CD–2011–60 (to be published as a NASA TM).

10. L. Mason, “Evolutionary Development Path for Viable Nuclear Electric Propulsion Systems for 

Exploration,” presented at the 47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, August 1, 

2011.

11. J.H. Gilland et al., “MW-Class Electric Propulsion System Designs for Mars Cargo Transport,” 

presented at the AIAA Space 2011 Conference, Long Beach, CA, September 27-29, 2011.
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Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to an Asteroid 

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6. AUTHOR(S)

Mercer, Carolyn, R.; Oleson, Steven, R.; Pencil, Eric, J.; Piszczor, Michael, F.; Mason, Lee, 

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14. ABSTRACT

NASA’s goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to 

achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar 

orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to 

drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a 

solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust 

architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its 

surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth 

asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity 

analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with 

a “reasonable vehicle configuration” bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those 

technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

15. SUBJECT TERMS

Spacecraft propulsion; Spacecraft power supplies; Technology assessment; Solar electric propulsion; Nuclear electric propulsion

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