Thomas W. Kerslake

Glenn Research Center, Cleveland, Ohio

Effect of Voltage Level on Power System Design

for Solar Electric Propulsion Missions

NASA/TM—2003-212304

April 2003

https://ntrs.nasa.gov/search.jsp?R=20030054462 2018-05-15T16:31:27+00:00Z
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Thomas W. Kerslake

Glenn Research Center, Cleveland, Ohio

Effect of Voltage Level on Power System Design

for Solar Electric Propulsion Missions

NASA/TM—2003-212304

April 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

International Solar Energy Conference

cosponsored by the American Society of Mechanical Engineers’ Solar Energy Division,

Japan Society of Mechanical Heat Transfer Division, Japan Solar Energy Society,

Korean Solar Energy Society, and CSIRO Australia

Kohala Coast, Hawaii, March 15–18, 2003
Available from

NASA Center for Aerospace Information

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National Technical Information Service

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This report is a formal draft or working

paper, intended to solicit comments and

ideas from a technical peer group.

This report contains preliminary

findings, subject to revision as

analysis proceeds.

Available electronically at http://gltrs.grc.nasa.gov

Acknowledgments

The author acknowledges Dr. Ioannis “Yiangos” Mikellides and Dr. Gary Jongeward of SAIC for developing

the expression for plasma electron collection current density, Mr. Andy Hoskins of General Dynamics

(Aerojet Corp.) for many productive technical discussions on high-voltage SEP systems

and Mr. Ken Metcalf of Boeing Canoga Park for developing

the detailed PMAD sizing spreadsheets.
NASA/TM—2003-212304 1 

EFFECT OF VOLTAGE LEVEL ON POWER SYSTEM DESIGN 

FOR SOLAR ELECTRIC PROPULSION MISSIONS

Thomas W. Kerslake 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Phone: 216-433-5373 

Fax: 216-433-2995 

Email: Thomas.w.kerslake@nasa.gov

ABSTRACT

This paper presents study results quantifying the 

benefits of higher voltage, electric power system 

designs for a typical solar electric propulsion spacecraft 

Earth orbiting mission. A conceptual power system 

architecture was defined and design points were 

generated for system voltages of 28-V, 50-V, 120-V 

and 300-V using state-of-the-art or advanced 

technologies. A 300-V “direct-drive” architecture was 

also analyzed to assess the benefits of directly powering 

the electric thruster from the photovoltaic array without 

up-conversion. Fortran and spreadsheet computational 

models were exercised to predict the performance and 

size power system components to meet spacecraft 

mission requirements. Pertinent space environments, 

such as electron and proton radiation, were calculated 

along the spiral trajectory. In addition, a simplified 

electron current collection model was developed to 

estimate photovoltaic array losses for the orbital plasma 

environment and that created by the thruster plume. The 

secondary benefits of power system mass savings for 

spacecraft propulsion and attitude control systems were 

also quantified. Results indicate that considerable 

spacecraft wet mass savings were achieved by the 

300-V and 300-V direct-drive architectures. 

INTRODUCTION 

Solar Electric Propulsion (SEP) systems offer 

attractive spacecraft mass benefits for a range of Earth 

orbiting and interplanetary missions (Mason and 

Oleson, 2000, Sarver-Verhey et al., 2002 and Kerslake 

et al., 2000). In recent years, the trend has been toward 

increasing SEP system power levels (Reinhardt et al., 

1998). As power levels increase, the spacecraft designer 

has greater motivation to select higher power system 

operating voltages to reap mass savings from smaller 

power cables and components. At the 300-V level, it 

becomes possible to implement a “direct-drive” 

architecture whereby the main discharge chamber of a 

Hall Effect Thruster (HET) is powered without a stepup converter (Jongeward et al., 2001, 2002, and 

Hoskins et al., 2002). This Direct-Drive Hall Effect 

Thruster (D2HET) configuration greatly simplifies the 

thruster power processor design and offers additional 

mass savings. However, high voltage photovoltaic 

arrays must be designed with arc tolerance and to limit 

parasitic electron collection currents. 

To quantify the benefits of higher voltage, power 

system designs were developed for a state-of-the-art, 

10-kW spacecraft envisioned for a 270-day Earth orbit 

transfer mission. This conceptual spacecraft is 

illustrated in Figure 1. The power system architecture 

was defined and design points were generated for 

system voltages of 28-V, 50-V, 120-V and 300-V. For 

each power system voltage design point, component 

technologies were chosen consistent with: (a) flight 

proven technologies and (b) advanced technologies 

available within the next 5 years. This paper discusses 

the power system design, modeling, and mission 

performance and the SEP spacecraft mass savings 

achievable with higher power system voltage levels. 

 

POWER SYSTEM DESIGN 

Power System Requirements 

The Electric Power System (EPS) design must 

provide 10-kW to the HET system, during orbital Sun 

periods only, over the course of the spiraling mission 

phase. This phase typically requires 3- to 9-months. In 

addition, the EPS must provide 0.4-kW continuously 

throughout the life of the mission to power spacecraft 

house keeping functions and payloads. If required, the 

spacecraft EPS can then support a multi-year GEO 

mission at a nominal power level of 10-kW. 
NASA/TM—2003-212304 2 

PMAD Architecture 

The Power Management and Distribution (PMAD) 

architecture is shown in Figure 2. In this direct-energy 

transfer architecture, there are two Photovoltaic Array 

(PVA) wings to generate electric power. Each wing is 

connected to an Array Regulator Unit (ARU), to control 

EPS voltage, and a Solar Array Drive Assembly 

(SADA), to articulate the wing for Sun tracking and 

transfer power via slip rings. SADA power is fed into 

the Main Distribution Panel (MDP) that provides 

switching for propulsion system and spacecraft loads. 

In addition, the MDP provides fault detection/isolation 

and grounding of EPS components to the spacecraft 

chassis. Spacecraft power is fed to the Power System 

Enclosure (PSE) that contains an isolation diode, a 

battery charge unit, a DC-to-DC Converter Unit 

(DDCU) and a Remote Power Controller Module 

(RPCM) for spacecraft load switching and fault 

protection. The SADA controller unit is also housed in 

the PSE. 

While spacecraft loads are assumed to operate at 

28-V, the spacecraft primary power bus voltage is 

treated as a parameter with values of 28-V, 50-V, 

120-V or 300-V. The rationale for choosing these 

voltage levels is as follows. 28-V was selected based on 

long standing flight heritage. 50-V was chosen to match 

the operating voltage of VAriable Specific Impulse 

Magnetoplasma Rocket (VASIMR) thruster radio 

frequency plasma generators (Baity, 1999). The 

VASIMR electric propulsion technology was included 

as part of the NASA Radiation and Technology 

Demonstration (RTD) Project flight experiment study. 

120-V was selected to match operating high-voltage 

power systems on NASA spacecraft, i.e. the 

International Space Station (ISS) EPS secondary bus 

and the Terra spacecraft primary bus (Kurland et al., 

2000). 300-V was selected since this is the lowest 

practical voltage to directly operate a HET without upconversion. The impacts of selecting a 300-V voltage 

level on component designs and PMAD architecture are 

shown in Table 1. The main impacts include the need 

for an ARU, the PPU design and location of the battery, 

i.e. tied to the EPS primary (high) or secondary (low) 

voltage level. 

Photovoltaic Array 

The State-of-the-Art (SOA) PVA wings consist of 

deployable rigid-panels populated with three-junction 

GaInP2/GaAs/Ge solar cells (Spectrolab, 2002) used on 

today’s high power satellites, such as the Lockheed 

Martin A2100 (Ehsani and Salim, 2000). Consistent 

with past work (Kerslake et al., 2000), 25% efficient 

solar cells are utilized and covered with massoptimized, 10-mil thick ceria-doped coverslides. Panels 

are constructed of composite face sheets bonded to a 

1-inch thick aluminum honeycomb core. The power 

harness consists of multiple flat copper conductors 

sized to allow a 3% voltage drop and sandwiched in 

layers of polyimide. The PVA mass has a calculated 

mass density of 3-kg/m2

 including a 10% factor for 

panel hinges, fittings and tie-down/release hardware. 

The SADA specific mass, 0.6-kg/kW, is based on the 

Moog Type 2 SADA (Moog, 2002). 

The SOA+5-year PVA wings consist of multiple, 

open-weave gore panels that deploy like an umbrella 

and are populated with 29% efficient multi-junction 

cells. This design is based on that reported by Jones 

(1993) and built for the Mars 2001 Lander spacecraft. 

The PVA has a total calculated mass of 1.8-kg/m2

. 

Energy Storage 

The energy storage subsystem is based on 

electrochemical battery cells. Both nickel-hydrogen 

(NiH2) individual pressure vessel and lithium ion 

battery cell technologies are considered. Cell properties, 

such as charge/discharge efficiencies, charge/discharge 

voltage limits, dimensions, mass, and operating 

temperature range, are obtained from commercial 

battery product data sheets (Eagle Picher, 2002, and 

AEA Technology, 2002). For this mission, a modest 

1000-cycle cell life is required. Cells are seriesconnected and housed in an aluminum enclosure to 

afford environmental protection and enhance thermal 

control. The battery mass is calculated at 1.1 times the 

cell mass to account for cell interconnects, by-pass 

diodes, cell voltage control (Li-ion only), cell heaters, 

the enclosure and connectors. The SOA+5-year battery 

assumes cells with 10% improved W-hr/kg 

performance and a carbon-carbon material enclosure. 

Power Electronics 

 ARUs contain multiple shunt elements that are 

switched on or off for coarse voltage regulation. The 

last shunt element is pulse-width modulated for fine 

voltage control (<1% ripple). Shunt elements are 

controlled by generating a voltage error signal used by 

switch gate control logic. The ARU contains 

input/output filtering and a PVA string blocking diode 

board. The MDP contains two input and three output 

hybrid switches and control electronics. Each switch 

consists of an electro-mechanical contactor in parallel 

with a solid state switch (MOSFET) and a current 

sensor. The PSE contains a notional solid-state charge 

limiting BCU, a 92% efficient DDCU and a 10-channel 

RPCM based on MOSFET switches. All components 

are mounted within aluminum enclosures with the 

associated control/monitoring electronics, power 

harnessing and thermal control (cold plates and 

heaters). The SOA+5-year power electronics are 

assumed to have a 10% improved kW/kg power density 

and carbon-carbon enclosures. Power cabling consists 
NASA/TM—2003-212304 3 

of twisted-pair copper conductors with ethylene 

tetrafluoroethylene insulation. 

Power electronic component and power cable 

masses were estimated using two separate approaches. 

The first approach, was to use a simplified mass-power 

scaling of similar ISS components (Kerslake et al., 

2000) while cable conductor gauge sizes were selected 

based on number of parallel conductors (two hot, two 

ground), operating temperature (100 °C) and bundled, 

vacuum current derating. In the second approach, 

component masses were estimated using a more 

sophisticated spread-sheet model. This model, derived 

from ISS power electronics designs, breaks down 

components into major subassemblies and utilizes mass 

curve fits based on design requirements and technology 

options (Metcalf, 1992 and Metcalf, 2002). In this 

approach, power cables were sized based on specified 

materials, efficiency (voltage drop) and the redundancy 

specified. 

ANALYSIS 

The spacecraft and orbit trajectory were designed 

to meet the RTD mission requirements (Kerslake et al., 

2000). However, this class of SEP mission has 

generalized applicability to a range of Earth science 

missions, interplanetary missions, micro-satellite and 

propulsion technology demonstrations and 

geosynchronous commercial and military satellites. A 

brief description of spacecraft system analyses and 

mission orbit mechanics analysis is given in the 

following sections. 

 

Mission Design 

Trajectory The mission entails launching a 

~1500-kg (wet mass) spacecraft into a 400-km, 51.6o 

inclination, circular low Earth orbit (LEO). This is 

followed by a ~270-day spiral transfer orbit to 35,000-

km equatorial orbit (nearly GEO). This maneuver is 

accomplished using a 10-kW electric propulsion system 

operating only during orbital Sun periods and thrusting 

tangential to the orbit plane. Phased, out-of-plane 

thrusting is used to achieve the inclination change. 

During the planar spiral out, micro-satellites are 

deployed to provide simultaneous radiation 

measurements of various Van Allen Trapped Radiation 

Belt locations. The spacecraft mission altitude history is 

shown in Figure 3. 

To estimate the impact of shadowing and launch 

date on the performance, the computer program 

SEPSPOT (Sackett, 1975) was used. This program used 

orbital averaging and solved for the minimum time 

trajectory between two closed conic orbits. Low-thrust 

trajectories were generated using the Glenn Research 

Center's high fidelity integrator “SNAP” which 

incorporated an eighth order Runge-Kutta integrator 

and the JPL DE403 Ephemeris file of the solar system. 

Earth shadow crossings were precisely modeled and an 

eighth order earth gravity model was used along with 

solar and lunar gravitational effects. EPS sizing (affects 

spacecraft initial mass) and mission trajectory (affects 

PVA degradation) analyses were performed iteratively 

until a converged solution was obtained. 

Electric Propulsion System The spacecraft is 

equipped with a 10 kW HET system that operates at a 

2,100-second specific impulse and a 54% combined 

thruster/PPU efficiency with xenon propellant. The 

electric propulsion system includes a propellant storage 

and flow control system and a Power Processing Unit 

(PPU). The PPU consists of an input EMI filter, the 

high power main anode discharge supply (with a boost 

converter), the lower power heater/keeper/magnet 

supply, auxiliary supplies and control electronics for the 

PPU and xenon flow control system. The PPU is 

actively cooled using an ammonia loop heat pipe 

system (Dynatherm, 2002) and a deployable rigid panel 

radiator. For 1000-W of heat rejection, the estimated 

loop mass was 14-kg while the estimated radiator mass 

was 30-kg for a 44-kg total mass. 

If the spacecraft EPS bus voltage is 300-V, a 

“direct-drive” PPU configuration is possible. In this 

case, filtering requirements are reduced and the main 

discharge converter is eliminated with an attendant 

electrical efficiency gain and reduced thermal 

dissipation requirement. This results in a 50% savings 

in PPU mass and heat rejection as well as reduced unit 

cost and complexity. Hoskins et al. (2002) provides an 

excellent review of “direct-drive” EPS and SEP system 

benefits. 

Spacecraft Flight Mode and Attitude 

Control The nominal spacecraft flight mode has the 

thruster axis in the orbit plane and tangential to the orbit 

ellipse. The solar array wings are perpendicular to the 

orbit plane and rotate about a single axis. This nominal 

flight mode is used for absolute solar beta angles, |β|, 

less than 25°, where β is the angle between the orbit 

plane and Earth-Sun line. During this flight mode, solar 

array cosine pointing losses are limited to 9%. 

For |β| greater than 25°, the spacecraft switches to a 

so-called "yaw-steering" flight mode (Chetty, 1991 and 

Hosken & Wertz, 2002). For this mission, the flight 

mode is more properly called "roll-steering" as the 

spacecraft rolls about the thruster axis (maintained in 

plane and tangent to the orbit ellipse). This type of 

spacecraft steering was studied for a solar electric 

transfer vehicle mission (Jenkin, 1992). The spacecraft 

is rolled through an angle ≤2*(90°-|β|) in one-half orbit 

period such that the combined spacecraft roll angle and 

single axis SADA rotation enables Sun-tracking PVA 

pointing. 
NASA/TM—2003-212304 4 

The spacecraft Attitude Control System (ACS) 

consists of 50 N-m-sec momentum wheels, a hydrazine 

monopropellant propulsion system (220-sec specific 

impulse) for momentum wheel desaturation, attitude 

sensors and control electronics. ACS jet thrust level and 

total propellant requirement were estimated based on 

engineering calculations for spacecraft maximum 

disturbance torques (Wertz and Larson, 1999) and roll 

steering torques. These torques would generally act 

about all three spacecraft axes, but for simplicity, were 

conservatively assumed to act additively about a single 

axis. Contributing disturbances included gravity 

gradients, aerodynamic drag, thruster misalignment, 

solar pressure and magnetic torques, in order of 

magnitude. Orbital momentum accumulation was also 

calculated. 

Orbital Environment Modeling 

Earth orbit environments affect PVA design and 

performance markedly. The largest contributor to PVA 

cell degradation is radiation damage from energetic 

protons and electrons trapped in the Earth’s magnetic 

field. The fluences and energy spectra of these particles 

was determined using the AP-8/AE-8 models (Jordan, 

1989). A secondary contributor to PVA degradation is 

impact damage for meteoroids and debris. Meteoroid 

fluences were calculated using the Cour-Palais model 

(Cour-Palais, 1969) while debris population fluences 

were determined using the Kessler model (Kessler et al. 

1989, 1996). The resulting impact damage and solar 

cell electrical performance loss were determined using 

empirical models (Eichelberger and Kineke, 1967, and 

Myre, 1991). 

Electron Collection Model Conventional PVA 

designs that are negatively grounded to the spacecraft 

chassis and that operate at voltages above ~120-V will 

collect electrons from the ambient and induced (HET 

plume) plasma environments. This results in a parasitic 

current loss that reduces solar cell string current. This 

parasitic current can become a major PVA design driver 

for high voltage arrays operating in the vicinity of high 

density plasmas created by electric propulsion thrusters. 

As such, plasma electron current collection must be 

modeled. However, current collection physics are 

complex and are still the focus of research (Jongeward, 

2002). Thus, for the purpose of these EPS parametric 

studies, a conservative engineering approach to 

modeling current collection was adopted. 

A preliminary, empirical model for solar cell 

plasma current collection, Ic (amps), for voltages, V, up 

to 300-V, was formulated as:

 Ic = F * Ac * Jc, (1) 

where F is an empirical “integral factor,” Ac is the 

exposed solar cell edge collection area (m2

) and Jc is the 

plasma electron collection current density (amp/m2

). 

For the two-dimensional, orbit-limited case ignoring 

potential barriers, Jc can be approximated as:

 Jc = Jth * [ 4*χ / π ]**0.5, (2) 

where χ = e*V/(k*Te) >>1 (Laframboise & Parker, 

1973, and Mikellides, 2002). Jth is the plasma electron 

thermal current density (amp/m2

) given by:

 Jth = (Ne*e/4) * [ 8*k*Te / ( π*m) ]**0.5 . (3)

In the above equations, e is the electron unit charge, m 

is the electron mass, k is the Boltzmann constant and Ne

and Te are the orbit altitude-dependent electron number 

density (#/m3) and temperature (K), respectively. For 

the ambient plasma, the approximate average orbit 

altitude-dependence of Ne and Te are given by (Smith 

and West, 1983 and Anon., 1991). The induced plasma 

environment at a PVA surface from an operating HET 

depends on the thruster design and the spacecraft 

configuration. For the current study, reasonable 

estimates can be derived from Jongeward (2001) that 

Ne ~ 2.5E13 #/m3 and from Gallimore et al. (2001) and 

Myers and Manzella (1994) that Te ~2.5 eV (or 29,000 

K). 

The electron collection area is given by: 

 Ac = 2 * (L + W) * δ * #cells, (4) 

where L is the solar cell edge length (m), W is the solar 

cell edge width (m), δ is the solar cell thickness and 

#cells is the number of series-connected cells in the 

solar cell string. 

The empirical integral factor, F, represents the 

average current collected from voltages 0 to Vmax 

(=300-V), divided by the maximum current collected at 

Vmax. This accounts for the distribution of cell voltages 

in a solar cell string and the different charging 

phenomena that take place as voltage increases, i.e. 

barrier formation and snap-over behavior. Current 

collection data versus voltage were obtained with a 

representative solar cell string coupon during ground 

testing with a laboratory-generated plasma (Jongeward 

2001). These data nicely fit a power-law equation. The 

average current collection was thus determined by 

integrating the power-law equation and dividing by the 

voltage range = Vmax. The maximum current collected 

was determined by evaluating the power-law equation 

at Vmax. The resulting value of F was found to be 0.33. 

Power System Computational Methods 

A dedicated Fortran code was developed to predict 

power system performance and estimate system mass. 

EPS component design and mission information were 
NASA/TM—2003-212304 5 

read in via data input files. Mission data, provided at 

15-minute intervals throughout the 270-day mission, 

included spacecraft position/velocity, Sun/Earth angles, 

orbit sun/eclipse indicator, insolation strength and local 

particulate radiation fluences. This information was 

used to calculate environmental heating rates, solar cell 

equivalent radiation dose (Anspaugh, 1996 and Marvin, 

2000), solar cell micrometeoroid/orbital debris damage 

area and ambient plasma electron collection current. 

EPS performance analysis was performed in a 

time-stepping, load-driven fashion. Based on load 

demand and set point voltages, PMAD system currents 

and voltages were calculated for the current time step. 

Component and cable losses were calculated based on 

input resistances, diode voltage drops and converter 

efficiency (if present). PVA string current was 

iteratively determined to satisfy solar cell and ARU 

voltage and current constraints. The number of string 

series-connected solar cells was determined iteratively 

such that cell operating voltage was no larger than 

maximum power voltage throughout the mission. The 

number of strings per wing was also determined 

iteratively such that minimum number of shunted 

strings was < 3. This minimized array area while 

ensuring that the ARU can maintain sun time voltage 

regulation throughout the mission. 

Solar cell electrical performance was modeled 

using a single exponential current-voltage (IV) function 

that was adjusted for operating temperature, 

illumination intensity, PVA sun pointing error and 

flatness, coverslide transmittance, environmental 

degradation and cell mismatch. Solar cell IV operating 

point and temperature were iteratively determined. Cell 

operating temperature was calculated using a lumpmass, transient thermal model accounting for 

environmental heating/cooling, electrical power 

extraction and interconnect wiring ohmic heating. 

Array area was calculated as the total solar cell area 

divided by a packing factor, 0.85. A wing length-towidth ratio of 6 was selected to obtain a width of 

approximately 2-m that corresponds to the approximate 

length of the spacecraft microsat storage section (see 

Figure 1). 

The battery was sized based on the input design 

characteristics of Table 2. These inputs were used to 

calculate the number of cells, cell capacity, design 

Depth-Of-Discharge (DOD), design charge/discharge 

rates and trickle charge rate. For mission analysis, the 

battery charge and discharge rates were determined 

based on the required load, battery charge and 

discharge efficiencies and the orbit sun and eclipse 

time. For orbits with little or no eclipse period, the 

battery charge current was set to a trickle charge value 

(C/50). 

The bulk temperature of EPS components was also 

calculated using a simplified, lumped-mass transient 

model. Calculated values of component power 

dissipation and environmental heating were used to 

determine operating temperatures and the required 

thermal control heating or cooling to satisfy operating 

temperature limits. 

RESULTS AND DISCUSSION 

EPS Performance 

The EPS component performance most affected 

through the mission was that of the PVA. A profile of 

PVA power is shown in Figure 4. The results shown are 

for the 120-V PMAD case but are indicative of the 

behavior of the array at other voltage levels as well. At 

the start of the mission, the full array power capability 

exceeds the delivered power requirement and thus, 

excess power is shunted in the ARU. Between mission 

days 25 and 75, a large drop in array power is seen 

followed by a more gradual decay throughout the 

remainder of the mission. By the end of the mission, the 

array full power capability is just slightly more than the 

delivered power (by design). The primary cause of 

array power degradation is solar cell radiation damage. 

This is evident in Figure 5 that plots various PVA 

current degradation components (voltage degradation 

components were similar). During mission days 

25 through 75, the spacecraft is crossing the most 

energetic region of the trapped proton belts at high 

magnetic latitude (orbital altitude of approximately 

2000-km to 7500-km). High energy protons are 

particularly damaging to the solar cell electrical 

performance. 

As the solar cells incur radiation damage, the array 

voltage capability is degraded and individual cells must 

operate at a relatively higher operating voltage to meet 

the PMAD voltage set point. This effect is seen in 

Figure 6 that plots the voltage ratio of solar cell 

operating voltage to peak power point voltage (Vmp). 

This voltage ratio starts at ~0.9 at the beginning of the 

mission with a fresh array and then approaches a value 

of ~1.0 (by design) on mission day 75 after sustaining 

radiation damage. Throughout the remainder of the 

mission, the voltage ratio recovers (decreases) slightly 

in response to cooler array operating temperatures that 

offer higher Vmp. Array temperatures and several 

PMAD box temperatures are shown in Figure 7. After 

mission day 100, the spacecraft no longer encounters 

Earth eclipses and extreme cold temperatures. For the 

remainder of the mission as altitude gained, thermal 

interactions with the Earth are diminished and 

component temperatures gradually cool off. 

Using multilayer insulation and thermal control 

coatings, calculated component electrical heating and 

active cooling thermal loads to keep hardware within 

acceptable operating temperature limits were negligible. 

For each power electronics box, the calculated values 
NASA/TM—2003-212304 6 

for dissipated heat and parasitic heat (i.e., power 

supply), respectively, were small: 123-W / 63-W for the 

ARU, 126-W / 80-W for the MDP and 88-W / 70-W for 

the PSE. With proper heat sinking, passive cooling is a 

reasonable assumption for the ARU and PSE. For the 

MDP, heat dissipation is localized at the switches and 

may require an active cooling cold plate interface to 

stay below semiconductor maximum operating 

temperature limits. 

Impact of PMAD Voltage 

Power system mass and sizing results for this SEP 

mission are shown in Table 3 as a function of EPS 

primary PMAD voltage. EPS technology assumptions 

and mass estimates were kept consistent with those of 

Kerslake et al. (2000). These results show that EPS 

mass decreases with increasing voltage. With higher 

voltages, current levels are reduced as are voltage 

losses throughout the system. With reduced currents, 

conductors can be sized smaller and mass saved. Lower 

system power losses allow for a smaller PVA active 

area and smaller power harness mass. These reductions, 

in turn, allow for a reduction in PVA ancillary structure 

mass and perhaps a reduction in PVA panel mass 

(though not assessed in this study). EPS mass 

reductions are approximately proportional to voltage 

level. Referenced to the 120-V case, the 50-V EPS is 

19% more massive while the 300-V EPS is 17% less 

massive. 

The battery mass did not vary monotonically with 

PMAD voltage level for three reasons. First, the battery 

design approach assumed the use of existing aerospacequalified cell designs (Eagle Picher, 2002, and AEA 

Technology, 2002) with discrete amp-hr capacities. The 

cell capacity selected had to be equal to, or greater than, 

the required cell capacity calculated. Thus, the battery 

mass depends on how closely the required capacity is 

matched to an existing cell product. For example, the 

50-V NiH2 battery required capacity was 17.6-amp-hr 

and matched well with a 20 amp-hr cell. But the 28-V 

battery required a 31 amp-hr capacity which could only 

be matched with a 50 amp-hr cell. Thus, the 28-V 

battery mass was considerably more than the 50-V 

battery. The impact of non-optimum battery size on 

EPS mass was also discussed by Gaddy, 1995. 

The second reason for non-monotonic battery mass 

was that cell performance varied with cell capacity. For 

example, a 20 amp-hr NiH2 cell has a specific energy of 

42-W-hr/kg while that of a 76 amp-hr cell is 

52-W-hr/kg. The third reason applies only to the 300-V 

battery options whereby the battery design is 

dramatically affected by the decision to locate the 

battery on the input side or output side of the DDCU. 

At 300-V, the NiH2 battery requires 201 seriesconnected cells which introduces reliability concerns. 

To address this issue, the NiH2 battery could be 

relocated to the DDCU 28-V output side and use the 

DDCU for charge control. This has the advantage of 

decoupling the battery design from the primary power 

bus voltage and allows the designer to tailor the 

secondary power bus voltage to the available battery 

cell capacity (Button, 2003). But without discharge 

control electronics, the spacecraft payloads will see a 

variable eclipse time voltage. However, these payloads 

could employ commercially available power converters 

already designed to handle the 25% voltage variation 

during battery discharge. Alternatively, for a modest 

mass penalty, battery discharge voltage could be 

regulated with a Series Connected Boost Regulator 

(Button, 1998). 

Another approach for a 300-V battery would be to 

select lithium ion cell technology with three times the 

average cell voltage capability compared to NiH2 cells. 

With lithium ion cells, a reasonable 74 series-connected 

cells are required which is comparable to the 76 seriesconnected NiH2 cells flying on the ISS US-segment 

EPS (Dalton and Cohen, 2002). 

Power cable masses also did not scale truly 

proportional to PMAD voltage for two reasons. First, a 

ground rule was imposed to limit minimum conductor 

size to AWG 20 (for robustness). Thus smaller, lower 

mass conductors could not be used even if current 

levels permitted. This limit was reached in the MDP to 

PSE power cable set at a PMAD voltage of 50-V. 

Second, as the conductor size decreased and voltage 

level increased, the insulation mass fraction of cable 

increased. This tended to off-set the mass savings in 

copper conductor mass. 

The 300-V direct-drive system offered an 11% 

decrease in EPS mass compared to the nominal 300-V 

system. There are two primary savings afforded to the 

EPS by implementing a direct-drive system. First, the 

PPU converter unit was removed increasing the box 

efficiency by 5 percentage points. This reduced the EPS 

power requirement by 500-W allowing for a smaller 

PVA area and mass. Second, the HET xenon flow 

control assembly was used to regulate primary PMAD 

voltage. This allowed the ARUs to be eliminated with 

an attendant 11-kg mass savings. In addition to EPS 

mass savings, the direct-drive HET system also realized 

mass savings of 17-kg in the PPU (without converter) 

and 15-kg in radiator mass from the PPU active thermal 

control system. Note that it would be possible to 

employ HET xenon flow control for primary bus 

voltage regulation for the 28-V, 50-V and 120-V 

voltage options. This would allow elimination of the 

ARUs, saving mass, but the PPU converter magnetics 

would have grow in mass to handle the resulting input 

voltage variations (Button, 2003). 

To improve EPS mass estimates and reconfirm the 

PMAD voltage trends discussed above, more detailed 
NASA/TM—2003-212304 7 

models for sizing power electronics and power cabling 

were utilized (Metcalf, 2002). These models were run 

separately from the dedicated Fortran code, but used 

component current/voltage results from the Fortran 

code as inputs. Results are presented in Table 4 for 

SOA technologies and SOA+5-year technologies. 

The overall trends of reduced EPS mass with 

increasing voltage and mass benefits from the directdrive architecture are still evident. However, the PMAD 

box and cables masses calculated were greater, while 

the mass savings realized with higher voltages were 

more modest. For the example comparing SOA 120-V 

NiH2 system to the 300-V lithium ion system, a 28% 

mass savings was calculated using simple scaling rules 

whereas only a 15% mass savings was calculated using 

the detailed PMAD component sizing models. The 

primary reasons for the more modest PMAD box mass 

savings are: (1) the enclosure & cold plate mass scale 

only weakly with component voltage level and (2) box 

ancillary components (power supply, control and 

monitoring card, local data interface, harnesses and 

sensors) are essentially independent of PMAD 

operating voltage. The mass savings of direct-drive for 

a 300-V PMAD was 15%. Assuming SOA+5-year 

technology (Table 4b), the PMAD voltage versus EPS 

mass trends and direct-drive mass savings remain the 

same. The largest difference afforded by advanced 

technologies was mass savings in the PVA. The openweave, multi-gore design with its highly efficient 

structure and higher efficiency / cooler operating cells, 

had a calculated mass nearly half that of the rigid panel 

SOA PVA designs. This is in spite of higher end-ofmission radiation dose (damage equivalent 1-MeV 

electron fluence of 6.0E14/cm2

 compared to a fluence 

of 4.0E14/cm2

 for the rigid panel). 

The improved power cable mass estimates, 

although slightly higher, compare reasonably well with 

the original mass estimates, except for the 300-V case. 

In the original mass estimate, insulation mass was 

scaled with conductor mass, whereas the improved 

model sized insulation thickness (mass) based on 

limiting electric field strength to half that of the 

insulator dielectric strength. With smaller conductor 

gauge size at 300-V, the improved model calculates a 

more representative insulation mass than the previous 

scaling approach. 

Plasma Electron Collection Currents 

PVA electron collection currents were negligible 

for array operating voltages of 120-V or less. For the 

300-V array designs, electron collection current was 

dominated by the HET induced plasma environment. 

The maximum calculated collection current was 4.6% 

of the PVA operating current. The maximum ambient 

current collection, only 0.4% of the PVA operating 

current, occurred at the start of the mission in LEO 

when electron number density was highest. Thus, for a 

~40-amp PVA peak operating current, approximately 

2-amps of parasitic electron current was collected. 

Therefore, to maintain the spacecraft chassis near 

plasma potential, the HET cathode must operate at a 2-

amp higher current level. 

If the HET is not operating, the ARU will shunt the 

majority of PVA strings and the electron collection 

current will decrease proportionally. In a direct-drive 

HET configuration, it is possible to operate the cathode 

only in a “plasma contractor” mode to maintain 

spacecraft chassis potential close to plasma neutral. 

Such an approach has minor design penalties since 

slightly more xenon propellant is required as is a 

slightly more complex, dual propellant flow control 

assembly to separately control xenon flow to the HET 

anode and cathode. 

 

Spacecraft Attitude Control System 

The calculated mass and PVA area savings from 

higher PMAD voltages and a direct-drive architecture 

result in a modest improvement in the spacecraft ACS 

mass. A summary of spacecraft mass properties, 

torques and momentum values are shown in Table 5 for 

the SOA and SOA+5-year EPS technologies. Torque 

and momentum values shown are for the most 

demanding part of the mission at the lowest orbit 

altitude of 400-km. 

Spacecraft moment-of-inertia and environmental 

disturbance torques are dominated by the PVA mass 

and dimensions. Thus any array mass or size savings 

afforded by PMAD voltage or direct-drive 

configuration results in an ACS mass savings. These 

mass savings are driven by reduced hydrazine 

propellant and tankage mass from fewer autonomous 

desaturation burns for momentum dumping. For the 

120-V SOA case, approximately 68-kg of propellant 

was required for momentum dumping over the mission 

duration. More frequent momentum dumps were 

required for the larger, lower voltage PVA cases, but as 

the spacecraft spirals out to higher altitude, the 

disturbance torques acting on the arrays quickly 

diminished. Therefore, the overall propellant savings 

from the 120-V case over the 28-V is only about 5% 

(about 4-kg of hydrazine and tankage). Increasing 

voltage level above 120-V offered very little propellant 

mass savings. Similar trends were observed for the 

SOA+5-year EPS technologies although spacecraft 

momentum management requirements were greatly 

reduced by the low-mass, circular-geometry, advanced 

technology PVAs. For this mission, the HET thrust 

offset torque dominated the spacecraft accumulated 

momentum term and thus, the added mass and 

complexity of a thruster gimballing is likely justified. 
NASA/TM—2003-212304 8 

Overall Spacecraft Mass Savings 

For the mission studied, the spacecraft wet-mass to 

dry-mass ratio is about 1.4. Thus, for every 1-kg saved 

in spacecraft dry-mass, 1.4-kg is saved in spacecraft 

wet-mass. For the SOA EPS, a 46-kg mass savings was 

obtained when increasing PMAD voltage from 120-V 

to 300-V. This translates into a 22-kg savings in xenon 

propellant and tankage and a spacecraft wet-mass 

reduction of 68-kg. Introducing direct-drive saves 

39-kg for 300-V EPS and 32-kg in the PPU and PPU 

heat rejection radiator. This translates to a 33-kg 

savings in xenon propellant and tankage and a 

spacecraft wet mass savings of 104-kg from directdrive alone. Introducing SOA+5-year EPS technology 

to a 300-V direct-drive system saves 84-kg of mass 

which translates into a 123-kg spacecraft wet mass 

savings. A fraction of these mass savings could be put 

into greater payload mass. Alternatively, the mass 

savings could enable a shorter LEO to GEO transfer 

time or perhaps a launch to orbit on a smaller, less 

costly launch vehicle. 

SUMMARY AND CONCLUDING REMARKS 

A dedicated Fortran EPS performance code was 

used in concert with spreadsheet models to quantify the 

potential mass benefits of high voltage spacecraft 

power systems and a direct-drive architecture for a 

representative, Earth-orbiting solar electric propulsion 

mission. Results showed that spacecraft EPS mass 

decreased as PMAD voltage was increased over the 

range of 28-V to 300-V. The mass savings from higher 

voltage, lower current were achieved by the use of 

lighter power cabling and component down-sizing due 

to lower system losses. Results from a simplified model 

showed that 300-V PVA plasma electron current 

collection was less than 5% of the array operating 

current. Further mass benefits were afforded by the 

direct-drive architecture by eliminating the ARU, 

removing the PPU converter, reducing the PPU heat 

rejection radiator and down-sizing EPS components due 

to lower system losses. These mass benefits were 

achieved assuming either SOA or advanced EPS 

technologies. The impact of EPS and propulsion 

subsystem mass savings on the spacecraft ACS system 

was also calculated. Results showed that mass savings 

of a 120-V system over that of a 28-V system can save 

only a few percent in ACS propellant mass over the 

mission duration. Further increases in voltage and 

direct-drive implementation offer very little added ACS 

propellant mass savings. However, increasing system 

voltage from 120-V to 300-V reduced spacecraft dry 

mass to save 68-kg in spacecraft wet mass (out of an 

initial 1500-kg total wet mass) and implementing 

direct-drive saved an additional 104-kg of wet mass. 

These added mass savings were achieved by reduced 

xenon propellant and tankage mass in the spacecraft 

primary electric propulsion system. 

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NASA/TM—2003-212304 9 

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NASA/TM—2003-212304 10

Table 1.—SEP Mission EPS Design Impacts 

Design Element ≤120-V 300-V 300-V D2HET 

Battery Technology NiH2 NiH2 NiH2 Li Ion Li Ion 

Battery Location: DDCU 

Input (high voltage) or 

DDCU Output (low voltage) 

Input Output Input Input Input 

Payload Bus Regulation 

(Sun/Eclipse) 

Reg/Reg Reg/Unreg Reg/Reg Reg/Reg Reg/Reg 

Thruster PPU Input Voltage 

Regulation 

Reg Reg Reg via propellant 

flow controller 

Power cable conductor gauge Baseline Smaller / 20 gauge lower limit Smaller / 20 gauge 

lower limit 

Solar Array Power Harness Baseline Reduced # traces, reduced ∆V/V Reduced # traces, 

reduced ∆V/V 

ARU Present? Yes Yes No 

PPU Efficiency (%) 90 90 95 

PPU Input Power (kW) 10.0 10.0 9.5 

Table 2.—Battery Design Inputs 

Design Feature NiH2 Li Ion 

Maximum Cell Voltage (V) 1.5 4.1 

Minimum Eclipse Time Bus Voltage 

(% of Sun Time Voltage Set point) 

75 75 

# of Cycles 1000 1000 

Recharge Ratio 1.1 1.0 

# of Failed (Open-Circuited) Cells 1 1 

By-pass Diode Voltage Drop (V) 0.7 0.7 

Battery Round Trip Energy Efficiency 

(%) 

79 85 

Design Temperature Limits (°C) 

Minimum 

Maximum 

 0 

20 

 -5 

30 

Table 3.—Design Impacts of System Voltage With Power Electronics Mass Scaling 

Design 

Characteristic

28-V 

NiH2

50-V 

NiH2

120-V 

NiH2

300-V 

Low-V/NiH2 NiH2 Li Ion 

300-V 

D2HET 

Li Ion 

EPS Masses (kg) 

 Solar Array 

 Battery 

 PMAD Boxes 

 PMAD Cabling

 Total 

253.2 

 33.5 

 116.4 

 20.7

423.8 

198.6 

 26.2 

 72.4 

 10.6

 307.8 

 167.2 

 35.6 

 51.5 

 4.1

 258.4 

 159.1 153.9 153.9 

 33.5 88.4 10.3 

 20.3 20.0 20.6 

 0.8 0.8 0.8

 213.7 263.1 185.5 

 148.3 

 10.3 

 9.3 

 0.8

168.7 

Array Wing Area (m2) 31.7 26.6 24.9 25.2 24.4 24.4 23.5 

# Solar Cells / String 18 28 64 154 154 154 154 

# Array Strings / Wing 624 336 138 58 56 56 54 

Battery Cell Capacity 

(Amp-hrs) 

50 20 10 50 10 4.5 4.5 

Battery # Series Cells 20 34 81 20 201 74 74 

 HET PPU Mass (kg) 35 35 35 35 35 35 18 

PPU Current (Amp) 382 205 84 33 33 33 32 

PPU Heat Rejection (W) 1000 1000 1000 1000 1000 1000 500 

PPU Heat Rejection 

Mass (kg) 

44 44 44 44 44 44 29 
NASA/TM—2003-212304 11

Table 4.—Design Impacts of System Voltage With Improved Power Electronics Modeling 

Design 

Characteristic

28-V 

NiH2

50-V 

NiH2

120-V 

NiH2

300-V 

Li Ion 

300-V 

D2HET Li Ion 

EPS Masses (kg) 

 Solar Array 

 Battery 

 PMAD Boxes 

 PMAD Cabling

 Total 

253.2 

 33.5 

132.4 

 24.5

443.6 

 198.6 

 26.2 

 110.7 

 12.1

 347.6 

 167.2 

 35.6 

 96.6 

 6.8

 306.2 

 153.9 

 10.3 

 90.9 

 5.6

 260.7 

 148.3 

 10.3 

 57.3 

 5.6

221.5 

(a) SOA Technologies 

Design 

Characteristic

28-V 

Li Ion 

50-V 

Li Ion 

120-V 

Li Ion

300-V 

Li Ion 

300-V 

D2HET Li Ion 

EPS Masses (kg) 

 Solar Array 

 Battery 

 PMAD Boxes 

 PMAD Cabling

 Total 

126.6 

17.6 

102.6 

 24.5

271.3 

 104.0 

 15.7 

 84.6 

 12.1

 216.4 

 92.9 

 11.2 

 72.9 

 6.8

 183.8 

 85.6 

 9.3 

 68.4 

 5.6

 168.9 

 80.4 

 9.3 

 42.7 

 5.6

138.0 

 (b) SOA + 5 years Technologies 

Table 5.—EPS Design Impacts to ACS 

Design 

Characteristic

28-V 

NiH2

50-V 

NiH2

120-V 

NiH2

300-V 

Li Ion 

300-V D2HET 

Li Ion 

Max. Spacecraft Mass Moment of Inertia 

(kg/m2

) 

16479 16479 13436 12217 11474 

Max. Sum of Environmental Torques 

(N-m) 

0.0524 0.0373 0.0312 0.0291 0.0277 

Max. Roll Steering Torque 

(N-m) 

0.3264 0.2245 0.1831 0.1665 0.1564 

Max. Sum of Environmental Momentum 

(N-m-sec / orbit) 

109.48 89.30 81.20 78.95 77.22 

Max. Roll Steering Momentum 

(N-m-sec / orbit) 

39.17 26.95 21.97 19.98 18.76 

Number of Momentum Dumps / Day 

BOL at 400 km altitude 

EOL at 35000 km altitude 

46 

9 

36 

9 

32 

9 

31 

9 

30 

9 

(a) SOA EPS Technologies 

Design 

Characteristic

28-V 

NiH2

50-V 

NiH2

120-V 

NiH2

300-V 

Li Ion 

300-V D2HET 

Li Ion 

Max. Spacecraft Mass Moment of Inertia 

(kg/m2

) 

6321 4975 4375 4209 3828 

Max. Sum of Environmental Torques 

(N-m) 

0.0181 0.0149 0.0135 0.0135 0.0125 

Max. Roll Steering Torque 

(N-m) 

0.0861 0.0678 0.0596 0.0574 0.0522 

Max. Sum of Environmental Momentum 

(N-m-sec / orbit) 

71.12 64.39 61.30 62.08 59.45 

Max. Roll Steering Momentum 

(N-m-sec / orbit) 

10.34 8.13 7.15 6.88 6.26 

Number of Momentum Dumps / Day 

BOL at 400 km altitude 

EOL at 35000 km altitude 

26 

9 

23 

9 

21 

9 

21 

9 

20 

9 

(b) SOA + 5 years EPS Technologies 
NASA/TM—2003-212304 12

Figure 1.—Conceptual SEP Mission Spacecraft 

Figure 2.—PMAD Architecture 
NASA/TM—2003-212304 13

Figure 3.—Mission Orbital Altitude History 

Figure 4.—Photovoltaic Wing Power 

Figure 5.—Photovoltaic Wing Degradation 
NASA/TM—2003-212304 14

Figure 6.—Photovoltaic Wing Operating Voltage 

Figure 7.—EPS Component Temperatures 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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E–13876

WBS–22–755–12–16

20

Effect of Voltage Level on Power System Design for Solar Electric

Propulsion Missions

Thomas W. Kerslake

Electric power; Solar arrays; Electric batteries; Solar electric propulsion; Electronic

equipment; Plasma (physics); Orbit transfer vehicles; Performance prediction

Unclassified -Unlimited

Subject Categories: 18 and 20 Distribution: Nonstandard

Prepared for the International Solar Energy Conference cosponsored by the American Society of Mechanical

Engineers’ Solar Energy Division, Japan Society of Mechanical Heat Transfer Division, Japan Solar Energy Society,

Korean Solar Energy Society, and CSIRO Australia, Kohala Coast, Hawaii, March 15–18, 2003. Responsible person,

Thomas W. Kerslake, organization code 6920, 216–433–5373.

This paper presents study results quantifying the benefits of higher voltage, electric power system designs for a typical

solar electric propulsion spacecraft Earth orbiting mission. A conceptual power system architecture was defined and

design points were generated for system voltages of 28-V, 50-V, 120-V, and 300-V using state-of-the-art or advanced

technologies. A 300-V “direct-drive” architecture was also analyzed to assess the benefits of directly powering the

electric thruster from the photovoltaic array without up-conversion. Fortran and spreadsheet computational models

were exercised to predict the performance and size power system components to meet spacecraft mission requirements.

Pertinent space environments, such as electron and proton radiation, were calculated along the spiral trajectory. In

addition, a simplified electron current collection model was developed to estimate photovoltaic array losses for the

orbital plasma environment and that created by the thruster plume. The secondary benefits of power system mass

savings for spacecraft propulsion and attitude control systems were also quantified. Results indicate that considerable

spacecraft wet mass savings were achieved by the 300-V and 300-V direct-drive architectures.