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

Glenn Research Center, Cleveland, Ohio

Adam M. Sajdak

Glenn Research Center, Cleveland, Ohio

Robert J. Scheidegger

Glenn Research Center, Cleveland, Ohio

Solar Electric Propulsion (SEP) Tug

Power System Considerations

NASA/TM—2011-217197

December 2011
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Thomas W. Kerslake, Kristen M. Bury, and Jeffrey S. Hojnicki

Glenn Research Center, Cleveland, Ohio

Adam M. Sajdak

Glenn Research Center, Cleveland, Ohio

Robert J. Scheidegger

Glenn Research Center, Cleveland, Ohio

Solar Electric Propulsion (SEP) Tug

Power System Considerations

NASA/TM—2011-217197

December 2011

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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sponsored by the Aerospace Corporation

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subject to revision as analysis proceeds.
Solar Electric Propulsion (SEP) Tug 

Power System Considerations 

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

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Adam M. Sajdak1

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Robert J. Scheidegger 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Solar electric propulsion (SEP) technology is truly at the “intersection of commercial 

and military space” as well as the intersection of NASA robotic and human space 

missions. Building on the use of SEP for geosynchronous spacecraft station keeping, 

there are numerous potential commercial and military mission applications for SEP 

stages operating in Earth orbit. At NASA, there is a resurgence of interest in robotic 

SEP missions for Earth orbit raising applications, 1-AU class heliocentric missions to 

near Earth objects (NEOs) and SEP spacecraft technology demonstrations. Beyond 

these nearer term robotic missions, potential future human space flight missions to 

NEOs with high-power SEP stages are being considered. To enhance or enable this 

broad class of commercial, military and NASA missions, advancements in the power 

level and performance of SEP technologies are needed. This presentation will focus on 

design considerations for the solar photovoltaic array (PVA) and electric power system 

(EPS) vital to the design and operation of an SEP stage. The engineering and 

programmatic pros and cons of various PVA and EPS technologies and architectures 

will be discussed in the context of operating voltage and power levels. The impacts of 

PVA and EPS design options on the remaining SEP stage subsystem designs, as well 

as spacecraft operations, will also be discussed. 

 

1

Michigan State University, USRP, summer intern. 

NASA/TM—2011-217197 1
Solar Electric Propulsion (SEP) 

Tug

Power System Considerations Power System Considerations

2011 Space Power Workshop

P S t A hit t Power Systems Architecture

April 20, 2011

Presentation by: Tom Kerslake

NASA GRC Co-authors

Tomas W. Kerslake

Kristen M. Bury

Jeffrey S. Hojnicki Jeffrey S. Hojnicki

Adam M. Sajdak

Robert J. Scheidegger

Also thanks to many unnamed, but highly 

appreciated contributors

NASA/TM—2011-217197 1
Outline

What is SEP and why use it?

SEP missions/spacecraft

SEP tug subsystem impacts on power system design

Solar array design considerations

PMAD design considerations

Cost challenge

Closing Comments

What is SEP and Why Use It?

SEP spacecraft have a solar electric power 

system (EPS) that provides power to electric 

thrusters

Save mission mass and/or costs

Achieved via high Isp electric propulsion

(~10X higher Isp than chemical)

Enhance/enable mission capabilities

Delta-V

Operating life

NASA/TM—2011-217197 2
1’s kW

station keeping

orbit topping

10’s kW

SEP Tech

Demo (SFD)

SEP Missions, Spacecraft, and Tugs

Space Science

Earth and 

Space Science

Earth Orbit 

Transfers, GEO 

Servicing

High delta V 100’s kW High delta-V

Maneuvering

100 s kW

Lunar/Mars

Human Missions

(HEFT)

Solar Array

EPS and SEP Tug Subsystem Designs

Are Highly Interdependent

Impact of

On

TCS EPS EP StrucMechACS

GNC

Comm Stowed

Config

Flt

Ops

Crew

TCS H

EPS M HH H H L H H M

EP H

Struct-Mech H

ACS-GNC H

Comm M

Stowed Config H

Flight Ops H

Crew n/a

H- High, M-Medium, L-Low

NASA/TM—2011-217197 3
SEP Flightmode and Pointing – EPS Impacts

From LEO to HEO, must point solar arrays and EP thrusters

With presence of large disturbance torques

1-, 2-DOF gimbal options, roll steering or solar inertial flight mode

Implications for solar array, ACS and EP thruster articulation Outboard Mast thruster 

Open truss 

canister

Deployable 

Mast

Root 

Joint

Outboard 

joints

Outboard Mast

Inboard 

Mast

Root 

Truss

Deployed

palette

Open truss 

canister

Deployable 

Mast

Root 

Joint

Outboard 

joints

Outboard Mast

Inboard 

Mast

Root 

Truss

Deployed

palette

Orbital Mechanics – EPS Impacts

SEP mission solar array / avionics radiation dose dominated 

(~98%) by trapped protons

Spiraling orbit inclination reduces dose by ~4x (0° to 51.6°)

GTO->GEO mission starting arg. of perigee can reduce rad dose by ~6.5X

Place perigee near nodes

Rad dose nearly independent of solar activity

EP steering Modes (minimize proton belt transit time) 

NASA/TM—2011-217197 4
Space Operations – EPS Impacts

Solar array and PMAD current/voltage sizing will be 

driven by EP subsystem design and ops

Conventional PPU-based EP subsystem

Direct drive DDU-based EP subsystem

Constant voltage (Isp) ops

Constant current (flow rate) ops

Variable voltage/current (such as Pmax, Max. Thrust, other)

SEP tug operations will drive solar array/gimbal design

Fast, uniform, robust, reliable solar array deployment

Tolerance of large docking/plume loads during RPOD

High deployed strength and stiffness

> ~0.1-g thrust-to-weight for chemical stage burns

Avoid SEP tug and chemical stage controls-structures interactions

Limit solar array deflections

Solar Array Design Considerations

Configuration: number of wings and articulation

Qual and recurring costs (modularity, optics), ground/flight 

testability 

Wing stowage, deployment, gimballing complexity, performance

EP plume avoidance (manage sputtering erosion) EP plume avoidance (manage sputtering erosion)

RASC-OASIS Human Mars DRM3.0

ISS

Low E

Main Beam

ScatteredScattered

NASA/TM—2011-217197 5
Solar Array Design Considerations

Ambient and EP induced plasma interactions

Parasitic electron collection (Dominated by EP plasma)

Plasma/Vacuum/Sustained Arcing Avoidance/Management

Radiation deg p radation optimization

Goal: Minimum SEP tug cost (or mass) by choice of:

Subsystem designs affecting EPS/EP performance, mission design 

and solar array design/sizing

Solar Array Design Considerations

Voltage selection

Large SEP tug missions optimize w/Hall Thruster Isp ~2000 sec (300-V)

300-V class solar array designs consistent with de-rated performance of 

SOA EEE parts, insulators and gimbal roll-rings/slip rings

300-V EPS saves 30-40% mass over 100-V to 160-V EPS SOA

Above considerations make 300-V class direct-drive option attractive

300-V solar array technical challenges include higher electron collection 

current and availability of PMAD electronic parts

Large solar array ground deployment testing

Key for risk mitigation/qualification without costly flight test

Designed for 1-g off-loaded, phased, thermal vacuum deployment

Designed to minimize qualification costs

NASA/TM—2011-217197 6
SEP Tug Primary PMAD

Distribution Architecture Options (Centralized or Channelized)

Centralized/Channelized Hybrid Option Is Attractive - Good cabling 

mass/efficiency, modularity, good fault tolerance with cross ties 

Voltage

LimitingLimiting

PMAD Design Considerations

Voltage level has large impact on PMAD design

Compared to SOA PMAD, 300+ V PMAD

Higher efficiency system Higher efficiency system

Significantly less thermal load (direct-drive)

Significant cable mass savings (>60%) due to relatively 

lower currents

Down conversion needed to feed housekeeping loads

Limited electronic parts selection, especially for high 

radiation mission: may require wide bandgap electronics radiation mission: may require wide bandgap electronics

NASA/TM—2011-217197 7
PMAD Design Considerations (con’t)

EP operating mode has large impact on PMAD design

Non-direct drive

Bus power to thruster PPU power conditioning and boost converter

Galvanic isolation decouples source (solar array) and thruster

Good bus power quality and prevents multi-thruster interactions

on the bus

PMAD must deliver predetermined I/V range to PPU

Direct drive offers

Bus power directly to thrusters via DDUs

Good: W/kg, power efficiency, reliability, recurring cost

Bus voltage control primarily tied to EP thruster operation

More work desired in the areas of:

Stability/ops during EP start-up/shut-down transitions

Cathode current sharing for multi-thruster ops

Effective grounding schemes 

PMAD Design Considerations (con’t)

Solar array regulator/limiter (Protects from bus high voltage excursions)

Fault tolerance a significant driver for human missions

Design for thruster-out capability (# failures tolerated?)

Cold-spare thrusters or nominally de-rated thrusters

Grounding (negative solar array grounding desired) Grounding (negative solar array grounding desired)

Positive solar array ground unacceptable due to arcing/sputtering introduced

600-V rated EEE parts w/derating just sufficient for 300-V class bus

Limited parts may lead to undesirable board and PMAD box designs

May need to increase to 1200-V rated parts (more limited selection)

EEE parts radiation tolerance (high flux, high energy protons)

Leads to high TID and enhanced SEE (MOSFET latch-up, gate rupture)

SEP mission unique combination, high voltage/power/rad dose 

further limits choice of available parts

May need custom parts development/screening (including SiC parts) and 

more rad testing (TID and SEE) – all increasing costs

May require more box-level radiation shielding adding significant mass

NASA/TM—2011-217197 8
SEP Cost Challenge

Cost estimates show high power SEP stage affordability 

challenge – need major cost reductions (~2X)

Major component cost challenge is the solar array

Recurring costs (particularly cell costs), qualification costs

Large-scale cell production availability

(excludes systems integration costs)

Closing Comments

High power SEP tug missions offer attractive benefits

Many {solvable} technical challenges remain and must be met

Yet programmatically, to progress beyond just SEP mission 

studies, major cost reductions are needed

NASA/TM—2011-217197 9
Thank You

Questions?

Appendix Charts

More detailed, back-up information

that is not part of the

main presentation

NASA/TM—2011-217197 10
Solar Electric Propulsion (SEP) Tug

Power System Considerations;

*** Back-up Material ***

2011 Space Power Workshop

Power Systems Architecture

April 20, 2011

Tom Kerslake, Kristen Bury, Jeff Hojnicki,

Adam Sajdak, Bob Scheidegger

List of Abbreviations a constant

Arg argument

BOL beginning of life

Comm communications

Conc concentration (optical)

DDT&E design, development, test & engineering

DDU direct drive unit

Delta-V change in velocity (of spacecraft)

DENI d i l ll i id

LEO low earth orbit

LVLH local vertical, local horizontal (flight mode)

Mech mechanisms

MOSFET metal oxide field effect transistor

OASIS orbital aggregation & space infrastructure systems

OPS operations

PDU power distribution unit

PMAD power management and distribution

Pmax maximum power DENI damage equivalent normally incident

DOF degree of freedom

DRM design reference mission

EEE electrical, electronic, and electromechanical

E-M L1 earth moon Lagrange point 1

EOL end of life

EP electric propulsion

EPS electrical power system

g acceleration due to gravity

GEO geosynchronous earth orbit

GNC guidance, navigation and control

GTO geosynchronous transfer orbit

Pmax maximum power

PPU power processing unit

rad radiation

RASC revolutionary aerospace systems concepts

Rev (orbital) revolution

RPC remote power controller

RPOD rendezvous proximity operations and docking

SEE single event effects

SEP solar electric propulsion

SFD SEP Flight Demonstration

SLA Stretch lens array (Entech Technology)

SOA state of the art GTO geosynchronous transfer orbit

HEFT human exploration framework team

HEO high earth orbit

Imp maximum power current

Isp specific impulse (of EP thruster)

ISS International Space Station

I/V current / voltage

Struct structures

TCS thermal control system

TID total integrated dose

V voltage or volts

Vmp maximum power voltage

NASA/TM—2011-217197 11
What is SEP and Why Use It?

SEP spacecraft have a solar electric power 

system (EPS) that provides power to electric 

thrusters

S d ave mass and/ t or costs

Achieved via high Isp electric propulsion

(~10X more than chemical)

Enhance/enable mission capabilities

Delta-V

Operating life

SEP Missions, Spacecraft, and Tugs

Commercial, Defense, NASA

1’s kW

station keeping

orbit topping

10’s kW

SEP Tech

Demo (SFD)

pp g

Space Science

Earth and

Space Science

Earth Orbit 

Transfers, GEO 

Servicing 100’s kW PMAD

High delta-V

Maneuvering

100 s kW

Lunar/Mars

Human Missions

(HEFT)

Solar Array

NASA/TM—2011-217197 12
SEP Spacecraft vs. a Stage or “Tug”

SEP Spacecraft

EP system is just one of the spacecraft loads

Spacecraft instruments/payloads are mission focus

Missions tend to start in higher energy orbits

Lower or moderate EP power levels Lower or moderate EP power levels

Evolutionary power system design challenges

SEP Stage or Tug (high Isp for multi-ton earth orbit transfers)

Spacecraft bus dedicated to SEP propulsion function 

No focus on instruments or small attached payloads

Prime purpose is to move mass (spacecraft) from point A to B in space

Missions tend to start in lower energy orbits 

Moderate to very high EP power levels

Many new power system design challenges

This presentation to cover high power SEP tug with focus on:

low-Earth orbit to high-Earth orbit spiraling missions

Solar array and PMAD elements of the EPS (no issues with energy 

storage)

EPS and SEP Tug Subsystem Designs

Are Highly Interdependent

Impact of

On

TCS EPS EP StrucMechACS

GNC

Comm Stowed

Config

Flt

Ops

Crew

TCS H

EPS M HH H H L H H M

EP H

Struct-Mech H

ACS-GNC H

Comm M

Stowed Config H

Flight Ops H

Crew n/a

H- High, M-Medium, L-Low

NASA/TM—2011-217197 13
SEP Flightmode and Pointing – EPS Impacts

From LEO to HEO, must fly tug to point solar arrays at the Sun and achieve desired 

EP thrust vector

Maintain attitude dead-band in presence of large disturbance torques

Options

LVLH w/1-DOF solar array gimbal (large solar array off-pointing)

LVLH w/2 LVLH w/2-DOF solar array gimbal (potential solar array shadowing) DOF solar array gimbal (potential solar array shadowing)

LVLH w/roll steering (ACS impacts, moderate solar array off-pointing)

Solar inertial (must move EP thrusters)

Open truss 

canister

Deployable 

Mast

Root 

Joint

Outboard 

joints

Outboard Mast

Inboard 

Mast

Root 

Truss

Deployed

thruster 

palette

Open truss 

canister

Deployable 

Mast

Root 

Joint

Outboard 

joints

Outboard Mast

Inboard 

Mast

Root 

Truss

Deployed

thruster 

palette

Orbital Mechanics – EPS Impacts

Mission design to reduce solar array/avionics radiation dose

1-year spiral LEO to E-M L1 Radiation Contributions:

Trapped electrons (~1%), Trapped protons (~98%), Solar flare protons (<1%)

Effective:

Increase inclination-reduce dose by factor of ~3.8x moving from 0° to 51.6°

inclination

EP steering Modes (minimize proton belt transit time) 

Ineffective: Launch date tied to min/max solar activity

NASA/TM—2011-217197 14
Orbital Mechanics – EPS Impacts (con’t)

Mission design to reduce solar array/avionics radiation dose

GTO to GEO Radiation Contributions:

Trapped electrons (~1%), Trapped protons (~98%), Solar flare protons (<1%)

Effective: selected Argument of Perigee

Can reduce dose by ~6.5X by placing perigee near nodes

Ineffective: starting perigee altitude, launch date tied to solar activity

EP Operating Mode – EPS Impacts

Conventional EP operating mode

EP PPU Buck/Boost Converter and solar array design EOL 

Vmp done in tandem (trade-off of costs, masses and 

efficiencies)

Individual EP thruster operation is isolated from the EPS Individual EP thruster operation is isolated from the EPS

Direct-drive EP operating mode

Individual EP thrusters not isolated from the EPS

Ops of each EP thruster affects ops of others and EPS I/V levels

Constant voltage (Isp)

Drives solar array string EOL voltage

Drives solar array/EPS channel current rating and shunt regulator size Drives solar array/EPS channel current rating and shunt regulator size

Constant current (flow rate)

Drives number of parallel solar strings to provide EOL current

May drive EPS channel voltage rating 

Variable voltage/current (such as Pmax, Max. Thrust, other)

Drives solar array EOL Vmp and Imp, EPS I/V design ratings 

NASA/TM—2011-217197 15
OPS/RPOD – EPS Impacts

Initial orbit post-insertion OPS drives solar array deployment

Rapid deployment and power gen (nominally <1 rev, avoid energy 

storage over sizing)

Uniform deployment (minimize attitude disturbance torques)

Robust/reliable deployment (avoid failures altogether or allow for Robust/reliable deployment (avoid failures altogether, or allow for

contingency mission ops)

Human SEP mission architectures include RPOD (in-space 

chemical stages)

Drives solar array deployed strength (plume and docking loads)

Drives solar array configuration and gimballing (docking vehicle ingress 

corridors, minimizing docking/plume loading, gimbal locking) 

In space operations (high In space operations (high-g chemical stage burns >0 1 g chemical stage burns, >0.1-g)

Drives solar array deployed strength (burn cut-off base g-loads)

Drives solar array deployed stiffness (displacements and frequencies 

for stack ACS during the burn)

Drives solar array configuration and gimballing (minimize bending 

moment, attain preferred orientation, gimbal locking) 

Solar Array Design Considerations

Configuration: dual large wings versus multiple small wings

Ability to stow wings, deployment ops, deployment reliability/robustness, 

gimballing complexity, EP thruster location/plume avoidance, ground 

and/or flight testability, qualification and recurring costs (modularity), 

performance (sun-pointing accuracy provided and self shadowing)

RASC-OASIS

ISS

Human Mars DRM3.0

* Cost and performance will 

factor into conc. optics 

configuration: planar, 2X 

CellSaver or 8X SLA

NASA/TM—2011-217197 16
Solar Array Design Considerations

Hall EP plume interactions (sputtering avoidance)

Optical/electrical coatings loss, structural material loss, contamination

Trade off complexity/mass of EP boom v notched/displaced solar array

Design configuration must avoid high energy main beam ions

45 cone rule cone rule-of-thumb from EP beam centerline thumb from EP beam centerline

Low energy charge exchange ions are non-sputtering

Moderately high energy, scattered ions demand special design attention

Plume ion uncertainties for high power, multi-thruster, far field, in situ

ScatteredScattered

Low E

Main Beam

Solar Array Design Considerations

Ambient and EP induced plasma interactions

Parasitic electron collection

Dominated by EP-induced plasma with high densities/energies

Solar array current loss mechanism, must oversize 

neutralizers/propellant

Design solar array strings with minimal exposed conductors

Plasma chamber coupon test data needed to verify collection levels

Arcing

Plasma/vacuum primary arcs not a concern during EP/neutralizer 

ops (ties spacecraft ground to plasma potential, minimizes voltage 

gradients)

Without EP ops, high orbit vacuum arcing must be managed 

conventionally (electrically bonded surfaces)

Sustained arcing avoided by proper design of solar array panel

Plasma chamber coupon test data needed to verify arc behavior

NASA/TM—2011-217197 17
Solar Array Design Considerations

Radiation degradation optimization

Goal: Minimum cost (or mass) of the SEP tug

Parameters affecting radiation dose

Avionics and tug subsystem designs, EP operation, mission design 

and solar array design/sizing and solar array design/sizing

Solar cell type (BOL performance, rad. tolerance, conc. ratio)

Solar cell shielding (coverglass, substrate thicknesses/densities)

Solar Array Design Considerations

Voltage selection

Using lower, state-of-the-art design voltages (100-V, 160-V) 

imposes a mass penalty on a high power SEP tug

30-40% for EPS subsystem alone (harnessing and power 

electronics)

Large SEP tug missions tend to optimize with EP Isp

~2000 sec (300-V, Hall Thruster)

Above items lead to 300-V class direct-drive design option

To save mass, solar array Vmp is typically matched to the 

desired operating voltage at demanding point of the mission

300-V class solar array designs consistent with de V class solar array designs consistent with de-rated

performance of state-of-the-art cabling/connectors, diodes, 

insulators and gimbal roll-rings/slip rings

Higher string design operating voltages will increase 

parasitic electron collection current ~ (1+aV)^0.7

300-V class solar array will drive PMAD parts selection

NASA/TM—2011-217197 18
Solar Array Design Considerations

Structural design (deployed strength/stiffness)

Very high deployed strength required (>0.1-g’s)

High thrust, chemical stage burns, RPOD plume impingement

Above events are planned, allowing for solar array preferred 

orientations

Very high deployed first fundamental frequency required

Consistent with SEP tug controller frequency and allowable

dead-band

Consistent with chemical stage/stack ACS control frequency

Minimize controls-structures interactions driven solar array loading

Achieve wing stiffness high enough to limit solar array deflections to 

acceptable levels acceptable levels

Solar Array Design Considerations

Ground deployment testing of large solar array

Key for risk mitigation/qualification without flight test

Flight testing will be an affordability challenge and offers limited 

range of qualification

Solar array must be designed for g off Solar array must be designed for g off-loaded thermal loaded, thermal

vacuum deployment

Structures off-loading, optics off-loading (as needed), in situ postdeployment thermal-electrical performance (as needed)

Solar array must be designed for an acceptable risk level of 

phased deployment of full size or limited size hardware

Deployment phases: tie-down release, yoke/phasing structure, solar 

array panels/wings array panels/wings

World’s largest vacuum facility (100-ft diameter and 122-ft height at 

NASA Plum Brook Space Power Facility) is not sufficient to test large 

SEP tug solar array designs envisioned

Solar array must be designed to minimize qualification costs

NASA/TM—2011-217197 19
SEP Tug Primary PMAD

Distribution Architecture Options

Centralized (All array power routed thru central distribution node, superior 

PMAD component mass/efficiency)

Combined Centralized/Channelized Option is Desired (Dual solar array 

wings feed dual centralized ARU inputs/outputs that feed channelized PDU segments for 

individual thrusters, reconfigurable with cross ties, superior cabling mass/efficiency, more 

mod lar s perior fa lt tolerance ith degraded performance and/or o er modular, superior fault tolerance with degraded performance and/or over-si ed solar arra sized solar array 

and power channels to route power between EP thrusters) 

Voltage

LimitingLimiting

PMAD Design Considerations

Voltage level has large impact on PMAD design

SOA voltage system (120 - 160 Vdc)

PPU required to raise voltage in order to get desired ISP from thruster –

less efficient, higher thermal load

Large cable mass required to handle high currents (mass inverse with Large cable mass required to handle high currents (mass inverse with

square of voltage)

Conventional parts available; still may need radiation screening

Housekeeping loads can be fed from bus power without conversion

High Voltage (300+ Vdc)

Direct drive option available – higher efficiency system and significantly 

less thermal load

Sig g nificant cable mass savings (>60%) y due to relatively lower currents

Down conversion needed to feed housekeeping loads at conventional 

voltages

Very limited solid state parts selection, especially for high radiation spiral 

– may require wide bandgap electronics

NASA/TM—2011-217197 20
PMAD Design Considerations (con’t)

Non-direct drive:

Delivers predetermined range of I/V to EP thruster PPUs

PPU power conditioning to the thruster

Decoupling, galvanic isolation between the source (array) and thruster

Simplifies ground testing of individual components (solar array, EPS, EP subsystem)

May simplifies design of solar array electrical simulator

Prevents interactions from multi-thrusters through the power bus

Improves overall power quality for the high voltage bus

Direct drive offers:

Highest kw/kg performance and superior power efficiency

Increased reliability, lower recurring cost

Requires no new hig gp h voltage/power electronics tech development

Bus voltage control is primarily tied to EP thruster operation

Past direct drive system ground tests show stable operation, but more 

work is needed

Stability/ops during EP start-up/shut-down transitions

Cathode current sharing for multi-thruster ops

Effective grounding schemes 

PMAD Design Considerations (con’t)

Solar array regulator/limiter

Protects from short-lived, high bus voltage post-eclipse with low load

Additional mass/efficiency hit, and adds to thermal load

Fault tolerance a significant driver for human missions

Reduced-power may be challenging at significant distance from Earth power may be challenging at significant distance from Earth

Design for ability to tolerate loss of thrusters and/or power feeds

Thruster out capability – How many failures tolerated?

Carry cold-spare thrusters, or

De-rate thrusters nominally and power up after failure(s)

Grounding

Negative solar array grounding to the tug chassis desired

Positive solar array ground is unacceptable Positive solar array ground is unacceptable

Introduces solar array plasma arcing

Introduces untenable solar array sputtering from EP plume ions

NASA/TM—2011-217197 21
PMAD Design Considerations (con’t)

EEE parts voltage level

600-V rated parts with derating, just sufficient for 300-V class bus

Beyond this, may need to jump to 1200-V rated parts (more limited selection)

Use of limited existing acceptable parts may lead to heavier, more 

voluminous, and less reliable / less efficient circuit board and PMAD box 

designs

EEE parts radiation tolerance (high flux, high energy protons)

Leads to high TID and enhanced SEE (MOSFET latch-up, gate rupture)

Dearth of manufacturer’s SEE data will necessitate dedicated testing

SiC parts could be a solution

SEP mission unique combination of high voltage/power and 

high radiation further limits choice of available parts

May need parts development, custom part builds, increased radiation 

testing, more stringent part screening, etc. which add costs

May require more box-level radiation shielding adding significant mass

SEP Cost Challenge

Cost estimates show high power SEP stage affordability 

challenge – need major cost reductions (~2X)

Major component cost challenge is the solar array

Recurring costs (particularly cell costs), qualification costs

Large-scale cell production availability 

(excludes systems integration costs)

NASA/TM—2011-217197 22
Closing Comments

High power SEP tug missions offer attractive benefits

Many {solvable} technical challenges remain and must 

eventually be met eventually be met

Yet programmatically, to progress beyond just SEP mission 

studies, major cost reductions are needed

NASA/TM—2011-217197 23
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01-12-2011 

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4. TITLE AND SUBTITLE

Solar Electric Propulsion (SEP) Tug Power System Considerations 

5a. CONTRACT NUMBER

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6. AUTHOR(S)

Kerslake, Thomas, W.; Bury, Kristen, M.; Hojnicki, Jeffrey, S.; Sajdak, Adam, M.; 

Scheidegger, Robert, J. 

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WBS 735785.01.02.01.03 

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John H. Glenn Research Center at Lewis Field 

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

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13. SUPPLEMENTARY NOTES

Adam M. Sajdak, Michigan State University, USRP, summer intern. 

14. ABSTRACT

Solar electric propulsion (SEP) technology is truly at the “intersection of commercial and military space” as well as the intersection of 

NASA robotic and human space missions. Building on the use of SEP for geosynchronous spacecraft station keeping, there are numerous 

potential commercial and military mission applications for SEP stages operating in Earth orbit. At NASA, there is a resurgence of interest in 

robotic SEP missions for Earth orbit raising applications, 1-AU class heliocentric missions to near Earth objects (NEOs) and SEP spacecraft 

technology demonstrations. Beyond these nearer term robotic missions, potential future human space flight missions to NEOs with highpower SEP stages are being considered. To enhance or enable this broad class of commercial, military and NASA missions, advancements 

in the power level and performance of SEP technologies are needed. This presentation will focus on design considerations for the solar 

photovoltaic array (PVA) and electric power system (EPS) vital to the design and operation of an SEP stage. The engineering and

programmatic pros and cons of various PVA and EPS technologies and architectures will be discussed in the context of operating voltage 

and power levels. The impacts of PVA and EPS design options on the remaining SEP stage subsystem designs, as well as spacecraft

operations, will also be discussed. 

15. SUBJECT TERMS

Electric power; Solar arrays; Electric propulsion 

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