Lynn A. Capadona, Jeffrey M. Woytach, Thomas W. Kerslake, David H. Manzella, Robert J. Christie, 

Tyler A. Hickman, Robert J. Scheidegger, David J. Hoffman, and Mark D. Klem

Glenn Research Center, Cleveland, Ohio

Feasibility of Large High-Powered Solar Electric 

Propulsion Vehicles: Issues and Solutions

NASA/TM—2012-217275

February 2012

AIAA–2011–7251
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Lynn A. Capadona, Jeffrey M. Woytach, Thomas W. Kerslake, David H. Manzella, Robert J. Christie, 

Tyler A. Hickman, Robert J. Scheidegger, David J. Hoffman, and Mark D. Klem

Glenn Research Center, Cleveland, Ohio

Feasibility of Large High-Powered Solar Electric 

Propulsion Vehicles: Issues and Solutions

NASA/TM—2012-217275

February 2012

AIAA–2011–7251

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 wish to recognize the entire SEP team at the Glenn Research Center for their dedication to this study. This paper would 

not have been possible without: Carl Blaser (CAD modeling), Bob Buehrle (Propulsion), Jose Cruz (S&MA), Tom Doehne (SE&I), 

Charles Doxley (Communications), Daryl Edwards (Propulsion), Rob Falck (Mission Design and Analysis), Kurt Hack (HEFT 

liason), Tyler Hickman (Reaction Control), Jeff Hojnicki (Power), Richard Kelsch (Structures), Tom Kerslake (Photovoltaics), 

Tom Kraft (Structures), Steve Mainger (Avionics), Dave Manzella (Propulsion) , Mike Martini (Attitude Control), Kelly McEntire 

(Structures), Henry Nahra (MMOD), Mike Piszczor (Photovoltaics), John Ramsey (Structures), Carl Sandifer (Communications), 

Bob Scheidegger (PMAD), Paul Solano (Structures), Anita Tenteris (Risks), Lisa Van der Aar (Software), and Joanne Walton (SE&I).

Available from

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either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 

This report contains preliminary findings, 

subject to revision as analysis proceeds.
NASA/TM—2012-217275 1

Feasibility of Large High-Powered Solar Electric Propulsion 

Vehicles: Issues and Solutions

Lynn A. Capadona, Jeffrey M. Woytach, Thomas W. Kerslake, David H. Manzella, Robert J. Christie, 

Tyler A. Hickman, Robert J. Scheidegger, David J. Hoffman, and Mark D. Klem

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Human exploration beyond low Earth orbit will require the use of enabling technologies that are 

efficient, affordable, and reliable. Solar electric propulsion (SEP) has been proposed by NASA’s Human 

Exploration Framework Team as an option to achieve human exploration missions to near Earth objects 

(NEOs) because of its favorable mass efficiency as compared to traditional chemical systems. This paper 

describes the unique challenges and technology hurdles associated with developing a large high-power 

SEP vehicle. A subsystem level breakdown of factors contributing to the feasibility of SEP as a platform 

for future exploration missions to NEOs is presented including overall mission feasibility, trip time

variables, propellant management issues, solar array power generation, array structure issues, and other

areas that warrant investment in additional technology or engineering development.

1.0 Introduction

In the wake of the February 2010 cancellation of NASA’s Constellation program, and in response to 

the FY2011 budget request, NASA was challenged to embark on a new human space exploration 

program. The Human Exploration FrameworkTeam (HEFT) was formed and chartered to create an 

evolvable framework for efficient, sustainable human exploration of multiple destinations in the Solar 

System. Within the framework, the knowledge, capabilities and infrastructure that NASA would require 

to support a new exploration portfolio were examined, with a focus on demonstrating critical enabling 

technologies. Through a series of Design Reference Missions (DRMs), the HEFT team evaluated multiple 

scenarios to achieve a sustainable architecture, all of which stem from the basic premise to be flexible and 

evolvable over time and contain certain core elements. Solar electric propulsion (SEP)was quickly 

recognized as an element that may be able to capitalize on existing technology to accelerate vehicle 

development while reducing upfront costs and mitigating risk because of its many advantages: “gear 

ratio” (payload mass fraction) significantly higher than traditional chemical stages, greater mission 

flexibility of departure and return windows, less catastrophic failure modes over traditional chemical 

propulsion, potential for a reusable architecture , and finally, the substantial availability of power at the 

destination and during coast periods required for more ambitious exploration missions. SEP has the 

potential to cost effectively move payloads from low Earth orbit (LEO) to higher energy Earth orbits 

(geostationary Earth orbit (GEO), Earth-Moon L-1, Lunar, etc.). In order for SEP technology to be 

infused into NASA mission (human or cargo), the trip time has to be acceptable (generally in the range of 

1 year or less) for the crew, and the payload must also occupy a significant mass fraction of the vehicle (at 

least equal to the wet mass of the vehicle itself). To realize the potential of SEP for large missions, several 

challenges unique to SEP technology must be understood and overcome to make this a sustainable 

exploration option: large scale solar array deployment, array pointing during spiral out of LEO, array 

stability during load events like large attitude control thruster firings and during docking operations, 

radiation belt array degradation, and plume impingent, to list a few.

NASA’s Glenn Research Center (GRC) is uniquely positioned to examine these technological hurdles 

and understand how to overcome them with respect to mission feasibility. GRC has over 30 years 

experience in power and propulsion system development, and has demonstrated success on a number of 
NASA/TM—2012-217275 2

other SEP related vehicle activities over the years. As far back as 1997, a GRC team won an engineering 

award for the design of a Human Mars SEP tug driven by a total of 800 kW Hall-effect thruster (HET)

power from LEO to a highly elliptical orbit (HEO) (Ref. 1). More recently, the Center has had 

involvement in multiple concept design demonstration proposals and studies. These include a Radiation 

and Technology Demo (RTD)—10 kW SEP tug powering a HET and Variable Specific Impulse

Magnetoplasma Rocket (VASIMR) proposal in 2000 (Ref. 2), various 10 kW class interplanetary concept 

design studies (Refs. 3 and 4), and a lunar gateway Tug design (400 kW HET, LEO to E-M L1) in 2001

(Refs. 5 and 6). Last February, a SEP Technology Demonstration Mission (TDM) was proposed in the 

Office of Chief Technologist with a FY12/13 start. The objective of this project is to provide the 

maximum advancement of SEP technologies for the funds available as specified in the FY2013 Space 

Technology Program Resource Guidance (PRG). NASA Project Managers at GRC were directed to

develop and implement the appropriate balance of technology maturation activities, including ground 

and/or flight elements, such that a SEP technology infusion point is possible in 2017 that supports the 

high powered (300 kW) capability needed for deep-space human exploration missions. By harnessing the 

resident expertise at the Center, a study team was formed that was able to generate a SEP vehicle concept 

to a level of detail sufficient to examine the factors that contribute to the development of a tug of this type 

and class, and the cadre of technology issues that should be addressed before development of a large, 

high-power SEP stage.

2.0 Overview

The concept developed by GRC was not intended to generate an optimized vehicle configuration. It’s 

purpose was to examine the engineering and technology development challenges contributing to the 

development of a 300 kW class SEP freighter to support human exploration missions to NEOs, a DRM 

proposed by the HEFT. Development of such a class of vehicle would represent a new orbit transfer 

capability that does not presently exist. The concept was grounded in four figures of merit: mass, stowed 

payload volume, minimization of complexity (keep it simple), and use of single fault tolerance across the 

concept. The use of these FOMs ensured maintenance of concept validity for the study duration. Each 

vehicle system concept always began by employing existing technology to minimize the introduction of 

risk into the system, and help reduce programmatic cost to develop a vehicle of this class. This was 

particularly effective in refining most subsystems, however, the subsystems with the highest masses

(power system solar arrays, as one example), did require use of low TRL materials to achieve acceptable 

solar array mass. Use of this “keep it simple” philosophy resulted in a system integration approach that

focused on technology streamlining or mass reduction efforts in most areas. Figure 1 shows the relative 

breakout (by mass) of every subsystem of the concept vehicle generated by the Study Team. Systems with 

largest impact are obvious key areas to focus streamlining/reduction efforts. Notable mass trades in each 

subsystem area will be discussed in each respective section of this paper as well as any technology or 

engineering development areas of need that would help increase the technology readiness level (TRL) of 

that subsystem.

A broad systems engineering approach is imperative to the successful end-to-end design of such a 

concept vehicle. Treatment of an SEP tug as a collection of individually managed subsystem disciplines 

integrated at the end of the design process is not sufficient to understand the vehicle interdependencies as

a whole, especially in the design phase. An integrated, iterative approach to vehicle development is 

critical. This paper will detail, subsystem by subsystem, the choices that were made and the technology 

challenges that were overcome. In almost all cases, design modifications within a subsystem were made 

in reaction to a change in another subsystem that produced a systematic impact. The following areas will 

be addressed: Mission Design and Analysis, Power/Propulsion, Environments, Structures and 

Mechanisms, Thermal, and Attitude Control.
NASA/TM—2012-217275 3

Nominal Spacecraft Dry Subsystem Masses (kg)

Figure 1.—Vehicle dry mass breakout by subsystem. 

3.0 Subsystems Issues and Solutions

Certain ground rules and choices were made at the onset to establish a starting point for our Study 

Team. For example, the end-of life power goal was 300 kW at the input to the thrusters. A payload 

envelope was also assumed with a 7.5 m height and an 8.5 m diameter (based on a 10 m fairing). Our 

concept study limited trip time from LEO to EM-L1 to approximately 1 year. We also limited crewed 

mission duration from EM-L1 to the NEO and back to Earth at no greater than 400 days. Single fault 

tolerance was imposed across the vehicle in accordance with NASA’s requirements for human rating

(Ref. 7) of space systems except in documented areas of exception such as pressure vessels and primary 

structure (as implemented in the Orion Crew Exploration Vehicle project). Certain other design choices 

were made because of existing knowledge with commercial off-the-shelf (COTS) products, or due to 

previous experience gained on International Space Station (ISS) or Constellation. The vehicle concept 

described in this paper is pictorially represented in Figure 2.

3.1 Mission Design and Analysis

The design of an SEP vehicle is a lesson in trades between the spacecraft’s power and electric 

propulsion (EP) systems and the manner in which the mission is to be flown. These three are intimately 

connected in an SEP vehicle; much more so than with a spacecraft utilizing a chemical propulsion system. 

Figure 3 illustrates the interplay involved in these system trades. The top-level mission requirements and 

ground rules mentioned above determined the type of EP system best suited for the mission. The long 

duration of the mission and, by extension, the higher throughput required of the thrusters were the key 

drivers. A Hall thruster system represents the most mature electric propulsion technology (Ref. 8) with 

characteristics desired for this mission, and the required propellant throughput and mission lifetime

required for this EP system are achievable based on testing of smaller (~5 kW class ) thrusters (Refs. 9 

and 10). Since the spacecraft must spiral from LEO to the EM L1 point, Hall Effect thrusters were
NASA/TM—2012-217275 4

Figure 2.—GRC SEP vehicle concept showing main spacecraft body 

and notched solar arrays. 

Figure 3.—Graphic showing the relationships involved in SEP 

vehicle mission design. 

selected so that the higher thrust they provide would help minimize the time the spacecraft would spend 

in transit though the Van Allen belts. This minimizes degradation of the solar arrays that occurs when 

exposed to these charged particle environments, as well as the total radiation dose on the vehicle’s 

electronics. A 30 kW per thruster power level was chosen for our study, with an operating current of 

100 A because these power and current levels are the largest that have been previously demonstrated in 

laboratory thruster testing at GRC (Ref. 11). The operating current is determined by thruster mass flow 

rate. High flow rates require very high performance space simulation chambers for ground testing. A 

100 A electric propulsion current relates to a propellant mass flow rate that is compatible with the space 

environment test facilities extant at GRC. This includes the capability to test a system of multiple 

thrusters in a vacuum environment while still maintaining vacuum quality (Ref. 12).
NASA/TM—2012-217275 5

In a traditional EP system, a power processing unit (PPU) takes power from the spacecraft’s Power 

Management and Distribution (PMAD) system and converts it to the voltages and currents that the 

thruster requires. All prior EP missions have utilized state-of-the-art (SOA) bus voltages; usually ≤ 100 V 

regulated (±2 V typically) or 80 to 160 V unregulated. For any high power SEP vehicle, optimizing the 

solar array characteristics, PMAD, and EP characteristics as an integrated system will result in best 

vehicle performance. Since we were extremely time limited on our study, we proceeded without 

performing this optimization. For a flight mission, time and resources upfront would be invested to fully 

assess this. For our concept study it was decided to design a 300 kW SEP vehicle for 300 V—the voltage 

a SOA Hall thruster requires to provide 2000 sec of specific impulse.

PMAD and thrusters were chosen to operate directly from the solary array voltage (absence of PPU). 

This is referred to as direct drive (Refs. 13, 14, and 15). Direct drive minimizes the need for voltage 

conversion inside the PPU, thus decreasing mass and waste heat generation. Section 2.0 details the 

concept PMAD system. We refer to direct drive PPUs in our study as direct drive units (DDUs). Our 

analyses showed that DDU’s are less than half the mass of conventional PPU’s, and operate at 99 percent

efficiency versus a PPU’s 95 percent efficiency.

In our trajectory analysis, solar array degradation was modeled as a function of time during passage 

through the Van Allen belts. We investigated three methods of flying our concept vehicle from LEO to 

the EM-L1 libration point:

1. Constant voltage, where propellant flow rate is varied such that the array voltage remains constant 

during its degradation

2. Constant current, where propellant flow rate remains constant to maintain 100 A of current draw per 

thruster, until the array can no longer support the total current draw requirement due to degradation.

After this point, a constant voltage is maintained.

3. Peak power, where the propellant flow rate is varied such that the current/voltage remains at the peakpower point of the array I/V curve throughout the spiral out.

The power and propulsion system and trajectory analysis interdependencies become even more 

intimate when the direct drive mode of operation of the EP system is baselined. For example, every time a 

trajectory was run with a different operating mode of the thrusters, the amount of main propellant

changed. Sometimes this required resizing of the xenon tanks, which occasionally drove tank size 

growth to the point that the spacecraft bus structure also had to grow in order to accommodate the larger

tanks. Since these changes increased the spacecraft dry mass, the trajectory analysis was repeated to reoptimize the trajectory, which would then impact the amount of main propellant, which caused us to 

revisit tank size, etc. Again, the time frame available to complete the study did not allow vehicle 

performance optimization for this mission. In the end, a constant current mode of operation for the 

thrusters which minimized initial mass in low Earth orbit (IMLEO) was chosen for the very limited trade 

space we were able to cover. This was not intended to be a true end-to-end mission optimization.

The mission was modeled using a combination of trajectory analysis tools and techniques. The Earth 

spiral to L1 was modeled as a tangential thrust spiral trajectory to a semi-major axis of 150,000 km. The 

vehicle attitude control was modeled to be capable of rolling to an attitude such that the arrays directly 

face the Sun (when not in Earth’s shadow). The GRC SNAP trajectory propagation tool (Ref. 16) was 

used for the Earth spiral to EM L1, and JPL’s MALTO software (Ref. 17) was used for the heliocentric 

portion of the mission. For pre-Phase A and Phase A analyses, using these tools in this manner is 

reasonable, but performing trajectory analysis for a flight mission will require a much higher fidelity level 

of simulation. For mission analyses utilizing large SEP vehicles (particularly for human missions), a 

single, high-fidelity 6-DOF tool is required using a combination of analytic steering laws for the spiral out 

combined with polynomial fits for the heliocentric phase. The Mission Design and Analysis Branch and 

Power Systems Engineering Branch in the Systems Engineering and Analysis Division at GRC are

engaged in developing a high-fidelity mission analysis suite for large SEP vehicles, combining the 

required trajectory analysis fidelity with a detailed solar array performance model along withoptimization 
NASA/TM—2012-217275 6

of other mission parameters (i.e., time spent in transit of the Van Allen belts, crewed and non-crewed trip 

times, thruster performance limitations and degradation, solar array performance limitations and 

degradation, PMAD operating schemes, etc.

3.2 Power and Propulsion

3.2.1 Solar Array Sizing and Loading

The SEP vehicle solar array shown in Figure 4 was sized to provide 300 kW and 300 Vdc at the 

electric propulsion interface at the end-of-life (EOL). For comparison, the eight solar array wings on the 

International Space Station could produce a combined power output of approximately 250 kW at 160 Vdc 

at beginning-of-life (BOL) but have a mass approximately five times higher than that of the SEP vehicle 

solar array. Accounting for SEP vehicle solar array integration loss factors (such as harnessing voltage 

drop), operational loss factors (such as sun pointing error and solar cell operating temperature) and natural 

and induced environmental degradation factors (discussed in Section 3.0), the resulting solar array was 

size was 594 strings of 134 series connected inverted metamorphic, 34 percent efficient solar cells.

These solar cells populated flexible blankets with a 1300 m2 total area divided amongst 56, 2.5- by 

5.0-m bays comprising each of the two solar array wings approximately 25 m on a side, modeled after 

ATK Aerospace System’s “SquareRigger (Ref. 18)” concept. Each solar array wing has a mass of 

approximately 1300 kg including at 30 percent mass growth allowance. 

The corner bays along the inner edge of the two solar array wings were removed to avoid excessive 

electric propulsion thruster plume impingement effects (sputtering erosion and contamination). One of the 

chief solar array production challenges is the required surge in high efficiency space solar cell production 

to allow for fabrication of these very large solar array wings on each of two SEP vehicles to be launched 

within the period of a year. This surge in space solar cell production would require a significant increase 

in the annual output of U.S. manufacturers over the planned period of flight solar cell production.

Figure 4.—Relative sizes of SEP concept notched solar arrays relative to Space Shuttle dimensions. 
NASA/TM—2012-217275 7

The SquareRigger bay composite structure and yoke were sized to handle the required deployed 

g-loading and also to maximize deployed frequency. Deployed g-loads were assessed for external, high 

load events such as a high thrust propulsion stage main engine cut-off. This event in particular was found 

to drive deployed g-load based on preliminary assessments. Other loading events evaluated were docking 

loads and docked vehicle attitude control thruster plume impingement loads during proximity operations.

The driving solar array wing deployed g-load was 0.2-g based on the in-space stack thrust-to-weight ratio 

of 0.1-g times a dynamic amplification factor of 2.0 based on the assumption of a step fall-off in high 

thrust main engine cut-off. Stowed wing launch restraint and release hardware and secondary structures 

were sized to meeting ascent loading conditions (see Section 4.0).

3.2.2 Power Management and Distribution

Power management and distribution supporting a spacecraft with a propulsion system that consumes 

90 percent of the total power generated created unique issues for power system stability and 

controllability. Furthermore the use of direct drive introduces a direct connection between the solar array 

and the propulsion system that highlights potential power system stability and power quality issues. 

The power distribution system was broken up into a high voltage main bus and a low voltage 

secondary bus as seen in Figure 5. The main power bus provides primary power to the propulsion system 

and primary power to the secondary dc-dc down converters. The main power distribution architecture was 

optimized by evaluating multiple distribution configurations. The low voltage secondary bus is supported 

by batteries that provide power to avionics and internal housekeeping loads during eclipse. The power 

source for the secondary bus was evaluated between main bus derived or segmented array derived power.

Main bus power was selected, which down converts a secondary voltage optimizing the use of all solar 

array power for any subsystem requiring the unused power. The secondary bus is isolated from the main

bus by a dc-to-dc power converter. The secondary bus batteries provide power to the spacecraft loads and 

propulsion heaters during eclipse. The SEP also carries requirements to provide payload power which can 

be provided directly from the main bus or by the secondary bus.

Controllability of the bus voltage and overall power quality are impacted by the source-load 

characteristics for the array and the primary load (propulsion system). The manner in how the system is 

started, how much noise is generated by the system, and dealing with electrical faults must be accounted 

for in the power system design. Bus voltage control was evaluated via two approaches; a) control the

ehicle bus voltage by the propulsion system, or b) introduce a dedicated bus regulator. Option a) was 

abandoned due to the long coast periods when the propulsion system is not active. During coast, the solar

Figure 5.—Bus structure details.
NASA/TM—2012-217275 8

array is very lightly loaded and bus voltages will approach array open circuit voltages, in addition the 

main bus voltage will experience very high voltages from the solar arrays when transitioning from 

eclipse. For these reasons a regulated bus approach was adopted. Two types of regulation approaches 

were evaluated, one using a peak power tracking (PPT) series regulator and a shunt regulator or

sequential shunt unit (SSU). The PPT counteracts the advantages of direct drive. The SSU provided 

regulation without the disadvantages of the PPT. In addition the SSU can take advantage of full array 

power at BOL to the propulsion system by allowing for adjustments in main bus voltage for elevated 

BOL levels. An additional benefit of the SSU can be to provide short circuit protection of array strings 

that may experience high voltage plasma arcing.

Early experimental hall thruster data indicates the thruster produces substantial current oscillations 

during operation. To control the current fluctuations produced by the thrusters, an input filter is required 

to limit inrush currents and decouple the ripple current that each thruster contributes. The input filter 

would be comprised of passive devices to isolate the ripple currents from the main bus. The input filter is 

housed with the other housekeeping power supplies required to operate the thruster in a DDU.

Stability of this power system is achieved both through design and operation. Design of the SSU will 

provide power system stability and avoid voltage swings as power levels change throughout the mission.

In addition, providing proper input power filtering will isolate the main power bus from voltage variations 

produced by thruster interactions with the main bus. Thruster operating procedures will be required so 

that current surges during start-up do not interfere with the operation of the power system, and so that 

current inrushes to the thrusters do not result in collapsing the arrays. Generally, stability can be achieved 

by operating the propulsion system below the knee of the solar array I-V curve. On the positive side the 

thruster is very tolerant to moderate voltage swings with small effects to the propulsion system 

performance, therefore perturbations of the propulsion system do not directly couple back into the power 

system. There are still other unknown issues that include thruster plasma interactions due to the multithruster configuration and current sharing with respect to the thruster cathodes for multi-thruster 

configurations. In addition, for the case of direct drive, due to the direct connection of the propulsion 

system to the solar array, momentary high current events can introduce instability in the power system 

that must be further assessed. 

System grounding methods were also briefly investigated; both positive and negative grounding was 

briefly evaluated. System grounding refers to how the spacecraft is connected to the solar array and how it

is impacted by the charged plasma surrounding the spacecraft. Both alternatives have advantages and 

disadvantages that this concept vehicle did not fully address

3.2.3 Addressing Insolation Periods

Spiraling out from LEO to the EM-L1 libration point means that the vehicle is flying in and out of 

eclipse. To maintain a sensible battery mass, thruster operation is precluded during eclipse. However, the 

thruster should be kept in a “ready state” to begin firing as soon as possible once the vehicle moves into 

the sunlight portion of the orbit. To accomplish this, our concept study adopted an operational scenario 

where we utilized the vehicle’s sun sensors and array power to register when the vehicle entered the 

penumbra of Earth’s shadow. At this time, the vehicle’s flight computer would begin to ramp thruster 

power down from nominal to zero over short period of time. Shut down is achieved by removing 

propellant flow as opposed to removing power from the thrusters. The ramp down was accomplished by 

shutting off xenon flow (time constant determined by volume between xenon flow control valves and the 

thrusters). Cathode flow was maintained (5 to 10 percent of total nominal propellant flow) during eclipse 

to allow continued cathode operation. This flow rates provides essentially zero thrust. The keeper power 

supply was maintained to sustain cathode operation when current dropped below 30 percent of nominal.

When discharge current reached zero, magnet power supplies were deactivated. During eclipse the 

thruster cathodes run at 200 W (10 A and 20 V), anode flow rate zero, magnet power zero.

When flying out of eclipse, magnet power and anode flow are reinitiated. Discharge will reignite at 

open circuit voltage once flow rate rises above zero. The initial current level should be low as flow rate 
NASA/TM—2012-217275 9

ramps up. Current and power will ramp up to nominal as flow rate reaches nominal value. Keeper supply 

is turned off after current reaches more than 30 percent of nominal value.

3.3 Environments

When considering aggressive missions beyond LEO, the natural and induced environments the 

vehicle needs to withstand must be well understood. Three main environmental issues were examined: 

induced impacts of the EP plume on the vehicle solar arrays, environments during spiral out of Earth orbit 

(Van Allen belt trapped proton and electron radiation), and the micrometeoroid and orbital debris 

(MMOD) environment and shielding.

The EP plume can completely erode the solar cell optical coatings on the arrays resulting in power 

losses and could completely erode thin flexible solar array blankets leading to structural failure. With 

careful, deliberate design choices, the solar array design performance loss in this case was limited to

~5 percent by positioning solar array surfaces away from the EP plume at the expense of greater solar 

array structure mass.

One other option considered to mitigate this environmental issue was to house the thrusters and 

gimbals on a platform at the end of a deployable boom to physically distance the plume from the solar 

array to mitigate the degradation risk. This adds complexity and mass to the system, requiring propellant 

lines, high power cables and command and data lines to be flexible in nature and deployed along with the 

boom. The study team’s power system and main propulsion leads worked together to perform a study of 

array degradation due to thruster plume effects and concluded we could eliminate the boom and “notch” 

the arrays to avoid direct plume impingement. Their study also showed that plume impingement would 

occur to some degree whether or not a boom was used. Though the greatest degree of plume 

contamination was bounded within a 45° half-cone angle centered on the thruster, the plume wraps back 

around itself, 360°. By notching the arrays we eliminate the greatest effects and then design the arrays to 

accommodate secondary and tertiary plume effects (see Figure 2).

With respect to radiation environments, travel through the Van Allen belts was a significant concern 

on spiral out. The most obvious impact is degradation to the solar array solar cells from trapped electrons 

and protons during the slow transit through these belts. Of particular concern, are high energy (10 MeV 

and above) trapped protons encountered at orbital altitudes of approximately 2,000 to 15,000 km. These 

high energy trapped protons produce the bulk (>95 percent) of solar cell damage compared to trapped 

electrons and solar flare protons. The traditional approach to handling solar cell radiation damage 

performance loss is to shield the cell with cover glass to limit the equivalent 1-MeV electron dose to 

approximately 1×1015/cm2 and then over-size the solar array current and voltage capability. This approach 

was adopted, utilizing the minimum amount of glass shielding of 5-mils thickness on the front and back 

sides (to restrict solar array areal mass increase) and resulting in approximately 20 percent solar array 

power loss at EOL for the envisioned mission. 

Vehicle MMOD protection was also a challenging problem to solve since the traditional MMOD risk 

assessment code BUMPERII (Ref. 19) is designed for LEO (~2000 km). The MMOD shielding on the 

vehicle was sized based on the “no damage allowed to the propellant tank” criterion, which has 

Constellation Program heritage for the LEO portion. Orbital debris was factored in over a very brief 

period of time because the vehicle spends limited time in such an orbit. The micrometeoroid portion was 

assumed to be the major contributor to vehicle damage over the life of the mission. Micrometeoroid flux 

penetration was calculated with respect to SEP distance from Earth by extrapolation of the Grun model

(Ref. 20) used in the ISS version of the BUMPER code models. To account for the flux behavior with 

altitude, the Earth gravitational focusing and shielding factors were calculated as well. The Earth 

gravitational focusing factor accounts for the increased flux due to the Earth gravitational field. This 

factor decreases with altitude. The Earth shadowing or shielding factor accounts for the fact that Earth 

shields from impacting micrometeoroids. This shielding factor varies with altitude as the SEP spirals 

away from Earth LEO. The flux of penetrating particles is calculated based on a critical particle diameter 

that penetrates the designed shielding which is determined from a ballistic limit equation (Ref. 21).
NASA/TM—2012-217275 10

Particles of diameter greater than the critical diameter present a risk to the SEP. The probability of no 

penetration was then calculated from the penetrating flux using Poisson statistics (Ref. 19). The model 

results were applied to the outer closeout panel structural composite skin on both the solar array gimbal 

sides of the vehicle as well as the top closeout panel. No additional shielding was needed on the radiator 

sides of the vehicle.

3.4 Structures and Mechanisms

The primary drivers in designing the spacecraft structure for the concept vehicle were launch loads.

The effect of launch loads was magnified since we assumed a launch configuration that not only included 

the SEP vehicle, but also included two other elements mounted on top of the SEP vehicle, the crew 

habitat module and the multi-mission space exploration vehicle. Buckling loads were the key driver of the 

strength of the spacecraft structure, and hence the structure’s mass. The concept structure began as 

aluminum thrust tube and morphed through several iterations into an aluminum-lithium structure with 

aluminum honeycomb outer panels with composite face sheets.

The solar array single axis gimbals were based on the only analogous gimbal mechanism with flight 

heritage—the ISS Solar Alpha Rotary Joint (SARJ). Utilizing two SARJ mechanisms along with a growth 

contingency resulted in a total mass for both gimbals of nearly 2,700 kg. Again, time limitations on the 

study did not permit optimization of this mass. The SARJ is designed with a high level of redundancy for 

its 15 year service life at 100 percent duty cycle of 360° rotations. An engineering development program 

based on SEP vehicle mission requirements and incorporating a sizeable mass reduction effort should be 

initiated for development of solar array gimbals for an SEP tug.

3.5 Thermal

The thermal subsystem may be one of the cleanest examples of interdependence with a SEP vehicle

using Direct Drive. Thermal control went through several re-designs as work progressed. One of the 

positive impacts of going to the direct drive architecture was that waste heat was dramatically reduced.

Direct drive forgoes the use of PPUs for the thrusters. As a result, as the power system architecture 

matured we were able to go from a deployed set of thermal radiators to a relatively simple configuration 

of body mounted radiators. The design challenges are described here.

For the initial configuration of the SEP, it was assumed that the thermal conditioning system would 

be similar to the system used for the ISS Photovoltaic Module. GRC managed the development and 

testing of this system, which consists of a pumped ammonia system with deployable radiators. A similar 

system is also used for the ISS Environmental Control and Life Support System (ECLSS). Due to the 

numerous distributed heat sources with large high power densities and high dissipation rates and the 

requirement that the system operate in LEO and NEO environments, it was more practical to use a 

pumped fluid for cooling rather than other heat removal methods, such as heat pipes.

The initial design required a constant heat rejection capability of 17.4 kW. The main source of this heat 

was the Array Regulator which was assumed to have an efficiency of 95 percent because it incorporated a 

voltage converter. In order to continuously reject 17.4 kW the radiators needed to have a total radiative area 

of 232 m2

. Therefore two two-sided radiators with dimensions of 5- by 12-m were selected. Because of this 

large size the radiators had to be deployable, similar to the ISS Photovoltaic module.

As the vehicle design matured, additional heat sources were identified but the Array Regulator 

efficiency was increased to 99 percent. This increase in efficiency was enabled by incorporating 'direct 

drive', i.e. the thrusters would be operated at the same voltage as the solar array output and no voltage 

conversion was required. This reduced the total heat rejection to 10.8 kW and the required area to 55 m2

. 

Since the radiator size was now 3.5- by 8-m, it was deemed feasible to mount the radiators on the boom 

that supported the thrusters. Although the radiators were still deployable, they were now partially 

supported by the thruster boom structure and their structural mass could be reduced.
NASA/TM—2012-217275 11

In a final design refinement, the heat decreased slightly to 9.56 kW and the required radiative area 

decreased to 44 m2

. The larger reduction in radiator area was achieved by increasing the coolant temperature 

by 30 °C and changing the fluid type to HFE 7000.Because the radiator now radiated at a higher 

temperature, the radiative area could be significantly reduced. The dimensions of the vehicle body also 

increased and it was found that one-sided body-mounted radiators could be used on two sides of the vehicle, 

and deployable radiators were not necessary. The radiator panel design was changed to be similar to the 

radiator panels being developed for the Orion Service Module which GRC is providing technical insight and 

oversight for. At the end of the final design iteration, the total mass of the thermal control system was 

reduced to 28 percent of the initial estimate. This reduction was mainly accomplished by incorporating a 

direct drive power management system and increasing the coolant fluid temperature. Table I provides a 

summary of the progression of the thermal system design values through the study iterations.

TABLE I.—CHRONOLOGICAL DEVELOPMENT OF THERMAL SUBSYSTEM DESIGN

VALUES DEMONSTRATING THE 28 PERCENT REDUCTION IN SYSTEM MASS

Iteration Required heat 

rejection rate,

kW

Radiating area, 

m2

Thermal system 

mass,

kg

0 17.4 232 2,050

1 10.8 110 1,240

2 9.56 44 586

3.6 Attitude Determination and Control

The AD&C subsystem for the concept vehicle consists of star trackers, horizon sensors, sun sensors, 

control moment gyros (CMGs), an inertial measurement unit (IMU), and a bi-propellant propulsion 

system. The subsystem is single fault tolerant and all components are off-the shelf with a high technology 

readiness level. A block diagram is shown in Figure 6. During the spiral out of LEO a roll steering is used

twice per orbit to maintain pointing of the solar arrays. This maneuver requires three CMGs to provide 

momentum storage capability about all three axes (including redundancy) and a propulsion system for 

wheel de-saturation. It was assumed that 20° of array off-pointing is acceptable and two of the SEP 

thrusters can be gimbaled up to 15° to offset some of the disturbance torques. The CMG specifications are 

based on what is currently flying on the International Space Station. Each one has a mass of 272 kg, 

stores 4760 Nms of momentum, and has a maximum output torque of 258 Nm. More analysis is required, 

but it is possible that the use of roll steering could be eliminated. This would result in a significant mass 

savings because the CMGs could be eliminated but the trip time would increase due to the cosine loss on 

the arrays. Clearly, if CMGs are to be utilized on an SEP tug, an engineering development program 

should be initiated to create a significantly lower mass CMG than that developed for the ISS while still 

maintaining a comparable momentum storage capacity.

The worst case total propellant load required to desaturate the CMGs in every axis was determined to 

be 367 kg. Monomethylhydrazine (MMH) and nitrogen tetroxide (NTO) were selected for the bi-prop 

reaction control system because of the higher specific impulse they offer over a blowdown hydrazine 

system and because of recent experience and developments on the Orion Service Module Propulsion 

System. Four pods of six R1E engines (25 lbf thrust each) were set as the baseline. The thruster size and 

locations are notional and require further analysis. If the CMGs get eliminated or the total propellant load 

decreases, the propellants and engine selection would likely be re-evaluated to find the most mass 

efficient design solution.
NASA/TM—2012-217275 12

Figure 6.—AD&C block diagram.

4.0 Feasibility

The overall challenge to implement an SEP vehicle into an exploration architecture for NASA 

without question hinges on affordability. Investment in technology and engineering development of key 

components is imperative before a large SEP vehicle can be developed. The risk and cost reduction 

provided by these developments will enable development of the SEP vehicle. Without them, mass and 

cost will continue to be a challenge, especially given launch vehicle payload envelope and up-mass 

restrictions.

As illustrated by this paper, development of light weight solar array support structures will be critical.

Though these structures must be low in mass, they must be of high strength and stiffness to support the 

large area arrays during docking and other loading events. Keeping the solar arrays to a manageable size 

will also keep impacts to the vehicle’s attitude determination and control system to the minimum 

necessary for mission success. In addition to this, solar cell and blanket technology must mature 

sufficiently to provide the high output efficiency required to keep the large arrays at a manageable size for 

launch vehicle stowage. To design an SEP vehicle today, solar cell production capability would need to 

increase as currently available off-the-shelf photovoltaic cell providers have an estimated annual

production rate on the order of ~ 300 kW and a SOA 29 percent efficiency. Although it would be possible 

to create a vehicle using existing technology, development activities especially in the photovoltaics area 

could certainly enhance efficiency over the current SOA, and mass in turn.

In addition to low mass, high strength solar arrays, the gimbal mechanisms that will drive these arrays 

must also be low mass yet robust enough to meet mission requirements. The SOA gimbal for large solar 

arrays (the ISS SARJ) must be significantly reduced in mass, yet be able to pass substantially higher 

voltage and current than the ISS SARJ.

Externally, several factors also impact the ability to cost-effectively produce and operate a vehicle of 

this type. A large SEP vehicle will require a xenon propellant mass nearing 30,000 kg. Xenon is a trace 

element found in the Earth’s crust, sea water, and most prevalently the atmosphere at a density on the 

order of 100 parts per billion. It is estimated that there is more than 1012 kg of naturally occurring xenon 
NASA/TM—2012-217275 13

worldwide (Ref. 22). As a result, the 30,000 kg required for this application is a very, very small amount 

with respect to the total amount available. However, this amount of xenon is in excess of 50 percent of 

current total annual worldwide production (Ref. 23), resulting in a significant acquisition challenge.

Xenon is produced as a by-product from air separation process, most commonly associated with the steel 

industry and is used in a number of applications other than electric propulsion including the lighting 

industry and various medical applications. While production capability beyond that needed to meet 

current worldwide annual usage currently exists, an acquisition strategy would need to be developed to 

purchase the required amount of xenon without having a dramatic impact on the world xenon market

(Ref. 24). Such strategies include stockpiling over the course of several years and entering into long term 

contractor with suppliers.

Propellant management on the vehicle presents some engineering development challenges, as well.

Large composite overwrapped pressure vessels (COPV) will be required to store the xenon for the SEP 

mission. These tanks are larger COPV tanks than have ever been manufactured before. Though they must 

be kept relatively light weight, these tanks must also demonstrate their pressure integrity in addition to 

their ability to provide protection against M/MOD impacts that could occur during the spiral out from 

LEO. Manufacturing large COPV tanks of this size represents and engineering development challenge 

that must be answered.

To maintain solar array pointing during the spiral out from LEO, the SEP vehicle’s attitude control 

system will need momentum wheels of low mass but high momentum storage capacity. Again, the SOA 

momentum wheels for a large vehicle (the ISS CMGs) must undergo an engineering development 

program to decrease their mass while maintaining or increasing their momentum storage capacity.

Mission design and trajectory optimization of a SEP mission will require a suite of tools to manage 

the interplay between trajectory and vehicle impacts. With variables like IMLEO, trip time, solar array 

degradation from passage through the Van Allen belts, all vying for importance, an integrated suite of 

analysis tools is needed for this SEP mission to insure that all variables are addressed so that the mission 

can meet all of its requirements.

Lastly, cluster Hall thruster operation, particularly in the direct drive mode of operation, need to be 

demontstrated in the space environment of the outward LEO spiral in order to increase the the TRL level 

of the thrusters and the DDUs for future application to the 300 kW class vehicle. This technology 

maturation will also validate the integrated software suite for mission design.

5.0 Conclusion

A point concept SEP vehicle was created by GRC with significant analysis completed. Through 

concurrent systems engineering, we have demonstrated a workable solution for an SEP vehicle 

independent of many outside constraints. More importantly, we have a sound understanding of the issues 

and challenges presented to us as the design matured, and intimate knowledge of the interdependent 

nature of SEP vehicle development. The resident expertise at GRC makes the Center extremely well 

positioned to understand and overcome the challenges involved with SEP system design. Appropriate 

investment in technology development is critical toward development of a reusable SEP system that can 

cost effectively transport NASA assets to and from more aggressive mission destinations.

References

1. Dudzinski, Leonard A., Hack, Kurt J., Gefert, Leon P. and Kerslake, Thomas W., “Design of a Solar 

Electric Propulsion Transfer Vehicle for a Non-Nuclear Human Mars Exploration Architecture,” 

26th International Electric Propulsion Conference, paper IEPC-99-181, Kitakyushu, Japan, Oct. 17-

21, 1999.

2. Kerslake, Thomas W., Haraburda, Francis M. and Riehl, John P., “Solar Power System Options for 

the Radiation and Technology Demonstration Spacecraft,” 35th Intersociety Energy Conversion 
NASA/TM—2012-217275 14

Engineering Conference, AIAA-2000-2807, Las Vegas, Nevada, July 24-28, 2000. (See also 

NASA/TM—2000-210243.)

3. Kerslake, Thomas W., and Gefert, Leon P., “Solar Power System Analyses for Electric Propulsion 

Missions,” 34th Intersociety Energy Conversion Engineering Conference, SAE 99-01-2449, 

Vancouver, British Columbia, Canada, August 1-5, 1999. (See also NASA/TM—1999-209289.)

4. Kerslake, Thomas W., “Solar Electric Propulsion Mission Architectures,” Space Power Workshop, 

The Aerospace Corporation, Redondo Beach, CA, April 21-24, 2003. (See also NASA/TM—2003-

212456.)

5. Sarver-Verhey, Timothy R., Dudzinski, Leonard J., Kerslake, Thomas W., Oleson, Steven R., and 

Hoffman, David J., “Solar Electric Propulsion Vehicle Design Study for Cargo Transfer to EarthMoon L1 ,” 5th SAE Power Systems Conference, Society of Automotive Engineers, Coral Springs, 

FL, October 29-31, 2002. (See also NASA/TM—2002-211970.)

6. Sarver-Verhey, T. R., Kerslake, T. W., Rawlin, V. K., Falck, R. D., Dudzinski, L. J. and Oleson, 

S.R., “Solar Electric Propulsion Vehicle Design Study for Cargo Transfer to Earth-Moon L1,” paper 

AIAA-2002-3971, 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 

Indianapolis, IN, July 7-10, 2002.

7. Human Rating Requirements for Space Systems, NASA NPR 8705.2B. Rev 1,12/ 2009.

8. Goebel, D. M. and Katz, I., Fundamentals of Electric Propulsion: Ion and Hall Thrusters. John 

Wiley & Sons, New York, 2008.

9. De Grys, K. H., et al., “4.5 kW Hall Thruster System Qualification Status,” AIAA Paper 2005-3682, 

July 2005. 

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Presented at the 53rd JANNAF Propulsion Meeting, Monterey, CA, Dec. 5-8, 2005

11. Manzella, D. H., Jankovsky, R. S., and Hofer, R. R., “Laboratory Model 50 kW Hall Thruster,”

AIAA Paper 2002-3676, July 2002.

12. Jankovsky, R., Tverdokhlebov, S., and Manzella, D., “High Power Hall Thrusters,” AIAA-99-2949, 

June 1999. 

13. Hamley, J. A., et al. “Hall Thruster Direct Drive Demonstration,” AIAA Paper 1997-2787, July

1997. 

14. Hoskins, W.A., et.al. “Direct Drive Hall Thruster System Development,” AIAA-2003-4726, July 

2003.

15. Hamley, J.A., “Direct Drive Options for Electric Propulsion Systems,” NASA TM 106576, 1995.

16. Martini, M., “S.N.A.P. 2.3 User’s Manual, Spacecraft N-Body Analysis Program,” 2005.

17. Sims, J.A., Finlayson, P.A., Rinderle, E.A., Vavrina, M.A., Kowalkowski , T.D. “Implementation of 

a low-thrust trajectory optimization algorithm for a preliminary design,” AIAA Paper 2006-2513,

2006. 

18. Adler, A.P, et. al. “Powersail: The Challenges of Large, Planar, Surface Structures for Space 

Applications,” AIAA Paper 2003-1823, 2003. 

19. Prior, T.G., “BUMPER II Training Guide,” version 1.79d. JSC-28725, 1999. 

20. Grun, E. Zook, H.A., Fechtig, H., and Giese, R.H., “Collisional balance of the meteoric complex,” 

Icarus, Vol. 62, 1985, pp. 244-274.

21. Christiansen, E. “Meteoroids / Debris Shielding,” NASA/TP—2003-210788, 2003.

22. Crofton, M.W. and Hain, T.D., “Environmental Considerations for Xenon Electric Propulsion,” the 

30th International Electric Propulsion Conference, IEPC-2007-257, Sept. 2007. 

23. Karwacki, G., “Recovering Xenon – and Costs: A Special Report to Cryogas International,” 

CryoGas International, Vol. 47, No. 7, p. 39, July 2009. 

24. Welle, R.P., “Xenon and Krypton Availability for Electric Propulsion: An Updated Assessment,” 

AIAA-1993-2401, June 1993.

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Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: Issues and Solutions 

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Human exploration beyond low Earth orbit will require the use of enabling technologies that are efficient, affordable, and reliable. Solar 

electric propulsion (SEP) has been proposed by NASA’s Human Exploration Framework Team as an option to achieve human exploration 

missions to near Earth objects (NEOs) because of its favorable mass efficiency as compared to traditional chemical systems. This paper 

describes the unique challenges and technology hurdles associated with developing a large high-power SEP vehicle. A subsystem level 

breakdown of factors contributing to the feasibility of SEP as a platform for future exploration missions to NEOs is presented including 

overall mission feasibility, trip time variables, propellant management issues, solar array power generation, array structure issues, and other 

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Solar electric propulsion 

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