Timothy R. Sarver-Verhey, Thomas W. Kerslake, Vincent K. Rawlin, Robert D. Falck,

Leonard J. Dudzinski, and Steven R. Oleson

Glenn Research Center, Cleveland, Ohio

Solar Electric Propulsion Vehicle Design Study

for Cargo Transfer to Earth-Moon L1

NASA/TM—2002-211970

October 2002

AIAA–2002–3971
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Timothy R. Sarver-Verhey, Thomas W. Kerslake, Vincent K. Rawlin, Robert D. Falck,

Leonard J. Dudzinski, and Steven R. Oleson

Glenn Research Center, Cleveland, Ohio

Solar Electric Propulsion Vehicle Design Study

for Cargo Transfer to Earth-Moon L1

NASA/TM—2002-211970

October 2002

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

38th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Indianapolis, Indiana, July 7–10, 2002

AIAA–2002–3971
Acknowledgments

The development of the SEP transfer vehicle system model was a joint effort by the Power and Propulsion Office,

the Power and On-board Propulsion Technology Division, and the Systems Engineering Division, through the

Systems Assessment Team. The authors would like to acknowledge the assistance of these organizations in the

system modeling. Also, the authors acknowledge the service of NASA LaRC personnel on the RASC program who

are responsible for creation of the graphics of the SEP vehicle shown in this report, as well as many others.

Available from

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7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

This report contains preliminary

findings, subject to revision as

analysis proceeds.

Available electronically at http://gltrs.grc.nasa.gov
NASA/TM—2002-211970 1

SOLAR ELECTRIC PROPULSION VEHICLE DESIGN STUDY FOR

CARGO TRANSFER TO EARTH-MOON L1

Timothy R. Sarver-Verhey,* Thomas W. Kerslake, Vincent K. Rawlin,‡

Robert Falck,† Leonard J. Dudzinski,* and Steven R. Oleson*

NationalAeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

ABSTRACT

A design study for a cargo transfer vehicle using solar electric propulsion was performed for NASA’s Revolutionary Aerospace Systems Concepts program. Targeted for 2016, the solar electric propulsion (SEP) transfer

vehicle is required to deliver a propellant supply module with a mass ofapproximately 36 metric tons from

Low Earth Orbit to the first Earth-Moon libration point (LL1) within 270 days. Following an examination of

propulsion and power technology options, a SEP transfer vehicle design was selected that incorporated largearea (~2700 m2

) thin film solar arrays and a clustered engine configuration of eight 50 kW gridded ion

thrusters mounted on an articulated boom. Refinement of the SEP vehicle design was performed iteratively to

properly estimate the required xenon propellant load for the out-bound orbit transfer. The SEP vehicle performance, including the xenon propellant estimation, was verified via the SNAP trajectory code. Further efforts are underway to extend this system model to other orbit transfer missions.

INTRODUCTION

In 2001, one element of the Revolutionary Aerospace

Systems Concepts (RASC) program was to develop

space related concepts in an Orbital Aggregation and

Space Infrastructure Systems (OASIS) framework.1

OASIS is a set of concepts that provide common infrastructure for enabling a large class of space missions.

The concepts include communication, navigation and

power systems, propellant modules, tank farms, habitats, and transfer systems using severalpropulsion technologies. OASIS features the in-space aggregation of

systems and resources to support the mission

objectives.

A reusable Hybrid Propellant Module (HPM) that combines both chemical and electric propulsion (EP) propellant in conjunction with modular orbital transfer/

engine stages has been targeted as the first OASIS concept. The HPM would provide chemical propellant for

crew transfers and EP propellant for slower HPM delivery and return. The HPM will be driven by a set of mission and concept requirements that encompass the

“sweet spot” for given set of Exploration Design

Reference Missions and future Low Earth Orbit (LEO)

commercialization scenarios. The concept will feature a

high level of reusability and will be supported by inexpensive launch of propellant and logistics payloads

from the Earth. The anticipated benefits of modular

transfer vehicles and propellant supplies are reduced

future mission costs and increased mission robustness

and flexibility for future space exploration and commercialization initiatives.

The initialHPM “baseline” resulted from a focused Sun

– Earth L2 mission concept evaluated in fall of 2000.1

The Earth-Moon L1 (located between the Earth and

Moon at a distance of approximately 320,000–

325,000 km from the Earth and referred to as LL1)

mission for the HPM begins with the assumption that

humans will return to the Lunar surface for scientific

operations and that a Gateway station at LL1 facilitates

this return and also provides a facility for science missions and missions beyond the moon. In order to provide this capability, HPMs are used along with

propulsion and crew transfer vehicles to ferry a crew

out to and return from the gateway. A SEP transfer

vehicle was selected for the cargo phases, which

–––––––––––––––––––––––––

* TechnicalStaff, Member AIAA

‡ Senior Member AIAA

† Student Member AIAA
NASA/TM—2002-211970 2

included delivery of loaded HPMs and other non-time

critical (i.e., not crewed) elements, as well as return of

empty HPMs to LEO. After delivery of the HPM module, the SEP transfer vehicle is placed in a safe parking

orbit untilit is needed.

During the preliminary phase of the HPM program, a

SEP transfer vehicle was selected to deliver the HPM to

LL1 because it is a non-time criticalcargo. The HPM

vehicle consists primarily of cryogenically stored hydrogen and oxygen propellants required to supply the

rocket engines on the crew transfer vehicle for a return

trip to LEO. It was also decided that the xenon propellant for the EP system would be cryogenically stored

within the HPM. Xenon gas is supplied to the thrusters

on the SEP transfer vehicle during operation via a fluids

interface that serves as the main attachment point between the two vehicle elements. The remainder of the

HPM mass, approximately 11%, is support structure

and operating systems. A summary of the mission requirements for the SEP transfer vehicle is provided in

Table 1. These requirements were developed by the

HPM program leads at NASA Langley Research

Center.

Table 1. RASC Hybrid Propellant Module

cargo mission specifications

Requirement Value

Delivered payload (not SEP,

xenon), metric tons

36.3

Initial Altitude/ Inclination,

km/deg.

400 km/

51.6 °

Final Altitude/ Inclination,

km/deg. (Earth-Moon L1)

320,000/

19.5 °

Outbound Trip Time Limit, days < 270

Mission Timeframe 2016

STUDY OBJECTIVE

NASA Glenn Research Center (GRC) was tasked with

supporting the SEP transfer vehicle development. This

effort evolved into the team at GRC developing a complete spacecraft system model for the SEP transfer vehicle with sufficient versatility to be readily applied to

other high-power orbit transfer missions. This system

modelwas the result of the investigation into the requirements of the power and propulsion subsystems

that would support the HPM cargo mission. The goals

of this investigation were threefold: first, determine the

mass of the SEP transfer vehicle with the HPM that is

able to deliver the required HPM fuel load within the

proscribed trip time; second, determine propulsion and

power generation characteristics that enable successful

completion of this mission; and third, determine the

characteristics of the remaining subsystems of the SEP

transfer vehicle so that a vehicle mass could be determined to a relatively high degree of fidelity. The SEP

vehicle system model was also desired to execute trade

studies for different power and propulsion technology

options. This report will describe the spacecraft configuration developed for the system model, technology

options examined for primary subsystems, and the

methodology employed to determine orbit transfer performance. It concludes with discussion of other applications of this system modeland on-going activities in

2002.

SPACECRAFT DESIGN

Early in the examination of the SEP vehicle performance it became apparent that the elements of the vehicle

needed to be distributed on a relatively large structure

so that the propulsion system did not destructively interfere with the solar arrays. Employing a boom arm

similar to a previous system developed at NASA GRC

for a Mars mission,2 the propulsion subsystem was

mounted at one end of the boom while the solar arrays

were mounted on the main vehicle structure to which

the other end of the boom is attached. This separation

insures that the solar arrays are protected from effluent

from the thrusters.

Subsystem Technology Selection

Options for the primary technologies used in the SEP

transfer vehicle were investigated to determine the best

fit for the HPM delivery mission. These technologies

will be discussed in the following sections. At the request of the NASA LaRC program leads, near-term and

15-year projections for the various technologies employed were developed. The far-term technology projections developed by the relevant technologists at

NASA GRC and used in this design study are summarized in Table 2 (located at the end of this report).

Propulsion Of the three propulsion technologies listed

in Table 2, Gridded Ion and Hall Effect thrusters provided the best performance for a low thrust spiral trajectory out to LL1. Gridded ion thrusters were selected

because they were found to have a total lower propellant requirement for this mission and have a substantially longer lifetime than the Hall thrusters, which

reduced hardware replacement costs. While these

‘benefits’ were achieved at the expense of longer trip

times, the ion thrusters were able to meet the mission

requirements. The specifications for the thruster incorporated into the system model are compiled in Table 3.

The number and configuration of the ion thrusters are

discussed in the next section.
NASA/TM—2002-211970 3

Table 3. Gridded Ion Thruster Specifications

Characteristic Value

Operating Power, kW 49.5

Thruster Efficiency, % 70.0

Specific Impulse, sec. 3300

Thruster Size, cm dia. 60.0

Thruster Lifetime, hours 15,000

Specific Mass, kg/kW 1.25

PPU/ThermalControlsystem

Specific Mass, kg/kW

4.40

PPU Efficiency, % 94.0

Power Generation For solar power generation, thin

film and multi-band gap photovoltaic cells were investigated as the most qualified options. Since it was estimated that approximately 450 kWe would need to be

delivered to the thrusters, it was expected that any array

meeting these requirements would be quite large.

Hence, size and weight were important drivers in the

technology selection. Preliminary results showed that

thin-film based arrays, while significantly larger than

the multi-band gap (MBG) arrays, were also typically

lighter than those arrays. The mass benefit of the thinfilm arrays increased further when an advanced array

support architecture was used. This support architecture, which is under development by AEC-Able, is referred to as the ‘SquareRigger.3

’ Consequently, thinfilm arrays were selected for the SEP transfer vehicle;

the parameters of the arrays are listed in Table 4. In

addition to having a light-weight structure, the arrays

are also operated at 500 V nominally. While operation

at this high voltage is aggressive, it is necessary to keep

the mass of the power cabling at an acceptable level.

While radiation degradation effects were neglected for

this mission, it is expected that arrays will be designed

to limit the impact. Additionally, the usable lifetime of

the arrays was l imited to two years, or two trips between LEO and LL1. Energy storage capability was

minimalon the SEP vehicle because thrusting was not

performed during shadow periods and therefore engine

power did not have to be supplied at these times.

Table 4. Photovoltaic Array Specifications

Characteristic Value

Cell Type Cu(In,Ga)

(Se2 or S2)

Cell Efficiency 19%

Support Architecture Square Rigger

Array ArealDensity, kg/m2 0.33

Array Specific Power, W/kg 390

TotalSpecific Power, W/kg

(PMAD, thermalincluded)

315

Power Processing For the purposes of this study, a

generic architecture for the power management and

distribution system for the photovoltaic arrays was used

with the following critical parameters. The system efficiency was 95% and the specific power was 3,000

W/kg. These values were provided for the technology

projections compiled in Table 2. The power processor

system for the gridded ion thrusters had slightly different values because it was developed in conjunction with

the engine rather than a generalpower management

architecture. For the thrusters used in the study, the

efficiency of the power processor was 94% and the specific mass was 4.0 kg/kW.

Thermal Processing Like power processing, a generic architecture was employed for the thermal control

system used to handle heat rejection from the solar arrays power processing system. With an arealdensity of

11 kg/m2 and standard assumptions about surface characteristics, the thermalcontrolsystem mass was estimated by simple radiative heat transfer to a background

environment assumed to be at 50 °C. The sizes of the

radiators were subsequently determined. The radiators

on the base pallet were two rectangular wings with dimensions of 2.5 m wide by 7.0 m long. As noted in the

previous section, the thermalprocessing system, sized

for the thruster system power processing units, was

developed with the thruster system and is projected to

have a specific mass of 0.4 kg/kW. The radiators on the

thruster pallet were also comprised of two wings with

the dimensions of 2.5 m wide by 6 m long.

Severalother subsystems were included in the SEP

transfer vehicle. However, since they are interest only

as mass contributions, they will not be discussed further

in this report. The masses for each of these subsystems

are shown in Tabl e 5.

Table 5. Miscellaneous Spacecraft Elements

Component Mass, kg

Thruster Pallet Structure 75

Base Pallet Structure 175

Micrometeriod & OrbitalDebris

Protection

260

Maneuvering Propellant (Xe) +

tank

2150

HPM Interface 235

Fluid Transfer Interface 400

Attitude ControlSystem (Mom.

Bias wheels, Docking system)

300

Auxiliary Batteries 25

Guidance, Nav., & Controlsystem 30

Control, Data, & Handling system 150
NASA/TM—2002-211970 4

Vehicle Configuration

The SEP transfer vehicle consists of three elements, as

illustrated in Figure 1. The first is a flat, disc-shaped

metallic pallet upon which nine high power electric

thrusters are mounted along with associated power

processing and thermalcontrolsystems. This thruster

pallet is attached to the second element, a semicollapsible articulated boom.

Figure 1. SEP Transfer Vehicle shown without PV

arrays and HPM.

Thruster Pallet

Coilable Mast;

deployed

Base pallet

Mast

Canister

Rigid End

Element

Solar Array

Stand-Off Boom

Umbilical

(in yellow)

Radiators

The boom consists of two rigid end elements and a

coilable mast, which allows the thrusters to be repositioned to maintain a constant thrusting vector while the

remainder of the spacecraft, including the solar arrays,

is flown in a solar inertial attitude to maximize power

generation. Additionally, the boom enables the thruster

pallet to be positioned so that the thruster exhaust

plume does not impinge upon and subsequently degrade

the solar arrays. To obtain a plausible estimate for mass

contributions, two pieces of existing spacecraft hardware were used to size the boom arm elements. For the

rigid end booms, the characteristics of the Shuttle Remote Manipulator System (SRMS)4 and the InternationalSpace Station Remote Manipulator System

(SSRMS)5 were used to estimate the masses for each

segment. The length and mass of the rigid segments are

listed in Table 6. For the middle segment, a collapsible

structure was required for vehicle launch packaging.

Using the coilable mast design developed by AEC-Able

for the Shuttle Radar Topography Mission,6 a coil abl e

boom was sized by scaling the length to the SEP transfer vehicle requirements. The length and mass of the

boom, along with the storage canister mass, are listed in

Table 6. Also listed in Table 6 are mass contributions

for two-axis wrist joints that were used to join the three

segments of the boom together as well as attach the

boom at each end. Simple mass contributions were assumed because no development work for appropriately

sized wrist joints has yet been undertaken. The final

elements of the articulated boom are materials required

for operation of the thruster pallet. Electrical power,

propellant, and command/control lines are required to

be run out to the thruster pallet from the base pallet.

Because of the high power delivered to the thruster, the

power cabling was comprised of 12 pairs of 0 gauge

copper wire. There were two propellant feed lines run

out to the end pallet, with one for redundancy. The

command/control lines were assumed to be collected

into a generic ribbon cable with the necessary redundancy. All of these lines were bundled together within a

flexible sheath that served as protection against micrometer damage. This ‘umbilical’ was routed through

the core of the boom elements. The mass contributions

of the elements of the umbilical are included in Table 6.

Table 6. Boom Arm Specifications

Specification Value

Rigid Boom Segment Length

(thruster side), m

3.0

Rigid Boom Segment Mass

(thruster side), kg

31.5

Rigid Boom Segment Length

(base side), m

4.0

Rigid Boom Segment Mass

(base side), kg

42

Coilable Mast Length, m 20

Coilable Mast Mass. kg 96

Coilable Mast Canister Mass, kg 405

Wrist Joint Mass, kg 50

Electrical Cabling Mass, kg 576

Propellant Line Mass, kg 80

Command/ControlLine Mass,

kg

20

Protective Cover Mass for electrical, propellant, and command

lines, kg

20

The design of the boom arm will be refined as part of

the 2002 activities, which is discussed later in this

report.

The third element of the SEP transfer vehicle is a base

container to which the solar arrays are mounted and

which contains associated power processing and thermalcontrolsystems, along with command and control,

guidance and navigation, and attitude controlsystems.

In addition, a supply of xenon propellant for pre-HPM

maneuvering is located in the base structure. Finally,

the propellant interface and mating fixture for attachment to the HPM are located on this pallet. The mass

contributions from each of these systems are compiled

in Table 5.
NASA/TM—2002-211970 5

ORBIT TRANSFER PERFORMANCE

Orbit transfer vehicles operating between LEO and

higher orbits, particularly GEO, under low-thrust continuous propulsion have been studied extensively in the

past.7,8,9 Traditionally, the spacecraft is spiraled out

from the starting orbit to its destination. This approach

minimizes the propellant required for the transfer, at the

cost of increased transfer time. Additionally, the

slowly-changing nature of the spiral transfer enables the

orbit-averaging technique to accurately assess the trajectory for transfer.10

Vehicle configuration can contribute to the orbit transfer capability if it imposes limits on power generation

or thrusting. In order to maximize power generation, the

SEP transfer vehicle with the HPM maintains a solar

inertial attitude with the solar array plane perpendicular

to the Sun vector throughout the orbit transfer. When

the thrusters are operating, the articulated boom points

the resultant thrust vector as close as possible through

the vehicle center of mass. This has the twofold benefit

of maintaining the vehicle’s attitude and maximizing

the net thrust. Additionally, the thrust can be maintained and varied optimally while any periodic attitude

changes in the SEP/HPM vehicle (i.e. array pointing)

can be minimized.

Preliminary Assessment An iterative process for

vehicle development was taken because of the following considerations:

1. Evolving vehicle definition

2. Increasing fidelity of vehicle subsystems designs during the course of this study

The orbit transfer performance via a continuous, l owthrust spiral(tangentialthrusting primarily with inclination change distributed over the transfer) was investigated using a relatively simple orbit averaging

technique derived from Stuhlinger’s optimal specific

impulse algorithm.11 Using the systems parameters described in Tables 2 through 6, the technique determines

transfer characteristics based on initialmasses, locations, and propulsion and power characteristics. Because the starting and finallocations and trip time were

fixed, the initialmass was the primary input and the

propellant mass, propulsion performance and power

requirements were the primary outputs for the iterated

analysis. After each iteration, the initial vehicle mass in

LEO was adjusted, which changed the propellant load.

With increasing vehicle mass, the propulsion system

was adjusted by changing the number of thrusters. This

then resulted in revising the solar arrays, power processing, and cabling characteristics to support the

required power demand, as well as changed the required propellant load. Once the vehicle configuration

was updated to reflect these changes, the transfer performance was tested again. The analysis was repeated

until the target payload mass was delivered to LL1

within the required time period. After convergence to

the target payload was achieved, the vehicle performance and finalmass breakdown was then verified with

the high fidelity orbit trajectory code.

High Fidelity Assessment In the high fidelity analysis, the orbit transfer was executed in severalphases.

The first phase is a geocentric spiralwhere the thrust

vector direction is defined by the weighted blend of

three steering laws: tangential steering, eccentricity

change steering, and inclination change steering. Each

steering law defines the varying thrust direction

throughout each orbit passage that, for a given amount

of thrust, provides a maximum change in semi-major

axis, eccentricity, and inclination, respectively. The unit

thrust vectors instantaneously defined by each law are

weighted and combined to obtain the appropriate thrust

vector direction at a given point in the orbit. The instantaneous weight of each steering law is optimized so that

the transfer is accomplished for a minimal propellant

mass. Transit time is also minimized at the expense of

increased power level. The thruster output always defines the thrust vector magnitude, which is maximized

except when the Sun’s illumination of the transfer vehicle is blocked by the Earth. During these shadow periods the thruster are turned off. Typically, the

dominance of one law over the others changes during

the transit. Tangentialsteering dominates the trajectory

from LEO out to a semi-major axis of approximately

13,000 km to minimize the time spent in the radiation

belts, after which eccentricity change steering is phased

in with increasing weight as orbit eccentricity increases.

Once the orbit becomes eccentric, thrusting to change

inclination is efficient near apogee. Here inclination

change steering is blended each orbit when the spacecraft true anomaly is such that maximum eccentricity

steering is inefficient. The finalphase of the transfer

occurs after the SEP/HPM vehicle orbit reaches an apogee of about 300,000 km to 325,000 km where, when

apogee passage is phased with the Moon, the Moons

gravity pulls the spacecraft into LL1. When this happens. the SEP/HPM vehicle inserts into a halo orbit

about LL1. Again, thrust vector steering is defined to

accomplish this optimally, for a minimal propellant

mass.

The SEP transfer vehicle performs the inbound transit

from LL1 to LEO in the same way as the outbound

transit, albeit with the thrust phases in reverse. The

SEP/HPM vehicle thrusts to move the spacecraft out
NASA/TM—2002-211970 6

of LL1 halo orbit and the spacecraft then falls into a

geocentric orbit. Again, the thrust vector follows an

optimized blend of tangential steering, eccentricity

change steering, and inclination change steering laws to

move the spacecraft down to circular low Earth orbit at

400 km at 51.6° inclination. These results were used to

further revise the system modelfor the finalvehicle

configuration.

Typically in this approach, the orbit averaging technique tended to over-estimate the propellant load by

approximately 5%. However, the initial mass in LEO

was over-estimated and the trip time was underestimated by less than 1%. The relatively small variances for these parameters support the validity of this

approach.

SEP TRANSFER VEHICLE DESIGN RESULTS

The finalvehicle configuration that met the RASC requirements was determined to have an IMLEO of

59,800 kg. Of this mass, the SEP transfer vehicle dry

mass was 11.3 metric tons (with 20% contingency). The

spacecraft propulsion system consisted of eight gridded

ion thrusters that produced a net thruster of 20 N. The

solar arrays needed to generate 448 kWe to support

these engines as well as provide 5 kW for on-board

systems operation. These arrays consisted of two wings

with a totalarea of 2,685 m2

. The xenon propellant load

for the outbound orbit transfer was found to be 12.2

metric tons. A mass breakdown of the major elements is

shown in Figure 2. Additionally, the SEP transfer vehicle is shown graphically in Figure 3. This vehicle performance satisfied the HPM program requirements and

was incorporated into the 2001 HPM mission report.

Figure 2. Mass Breakdown ofthe SEP & HPM

spacecraft. Vehicle IMLEO = 59,800 kg.

Figure 3. Mass Breakdown ofthe SEP Transfer

Vehicle ‘Dry’ Mass. This mass includes 2000 kg of

xenon stored on base pallet for parking maneuvers

only. ‘Dry’ Mass = 11,300 kg.

Figure 4. SEP Transfer Vehicle. PV arrays shown

to scale. Thruster pallet at top offigure. HPM not

included in figure.

OTHER MISSIONS

In addition to the HPM mission for the RASC program,

the SEP transfer vehicle system model has been successfully used in other recent studies. First, a study was

undertaken by NASA Johnson Space Flight Center to HPM Location

SEP Transfer Vehicle with HPM Mass

Distribution

* SEP Dry mass includes 2,000 kgof xenon stored on-board for

parkingmanuevering

Xenon

Propellant

Mass

20%

SEP Vehicle

Contingency

Mass

3%

SEP

Transfer

Vehicle Dry

Mass*

Payload 16%

Mass (HPM)

61%

SEP Transfer Vehicle Mass Distribution

0%

5%

10%

15%

20%

Vehicle Element

Parking Xe Supply (fixed)

Power Processing

Contingency

Articulated Boom

PV Arrays

TCS

OtherThrusters

Structures

Momentum Bias System

C&DH/GN&C

ACS
NASA/TM—2002-211970 7

look at the Lunar L1 Gateway mission architecture,

with primary focus on the Gateway station and lunar

missions. NASA GRC supported the definition of an

SEP transfer vehicle to deliver 30 metric tons of propellant to LL1 within 180 days. Hall effect thrusters were

employed in this mission because the shorter trip times

mandated a higher thrust device. The important parameters of the SEP transfer vehicle developed for this mission are summarized in Table 7. The propellant storage

differed significantly between the Gateway Habitat

mission and the HPM mission because in the former, all

the xenon propellant resides at the base container, including return propellant for the SEP vehicle. A replaceable propellant container was added to the SEP

transfer vehicle to facilitate propellant replenishment,

which accommodated the change. For the RASC/ HPM

mission, xenon propellant was stored on the HPM and

piped into the SEP transfer vehicle during operation.

The finalvehicle design for the Gateway mission is

illustrated in Figure 5.

Table 7. JSC Lunar Gateway Mission

Requirements & Results

Requirement Value

InitialMass in LEO 66,800 kg

Delivered payload (not SEP,

xenon), 30,000 kg

Initial Altitude/ Inclination 400 km/

28.6 °

Final Altitude/ Inclination 320,000 km/

19.5 °

Thruster type, # of thrusters

12 - 50 kW

Hall Thrusters

(2,500 sec.)

Power Generated 584 kW

Array area 7,300 m2

Xenon Propellant Load, Total

(Inbound)

21,500 kg

(4950 kg)

Outbound Trip Time, days 168

Mission Timeframe 2011

As part of the 2002 activities, the RASC program initiated a study into the possibility of utilizing power

beaming from a large power station located at LL1.

This central power station, which is also being analyzed

at NASA GRC, is currently being sized to provide electrical power via laser beams for up to four vehicles simultaneously. This estimate was based on the traffic

modelfor HPM fuelvehicles that supports activities

Figure 5. Lunar Gateway SEP Transfer Vehicle.

PV arrays shown to scale; Hall thruster pallet

shown at the bottom ofthe illustration, and the

payload mass is the approximately square element

shown attached to the base pallet near the center

ofthe image.

based at LL1. At this time, power generation technologies are being investigated—Nuclear fission versus

photovoltaic, along with orbit analysis of the HPM

vehicles and the central power station. In addition, the

electric propulsion transfer vehicle is being investigated

to determine the performance of a vehicle configured to

use beamed power.

Finally, a study has been initiated to examine and assess

the structuralrequirements and dynamics of the articulated boom implemented in all of the studies described

within this report. This study will determine the limits

of feasible application of the conformable boom and

any subsequent limitations this may impose on the SEP

stage application.

CONCLUDING REMARKS

A conceptual SEP vehicle design has been developed

for a 2016 mission to transfer of a fuelvehicle to an

Earth-Moon L1 libration point. This vehicle is able to

deliver the target payload mass of approximately 36

metric tons to the destination within 270 days. The

transfer vehicle with the fuel module has an IMLEO of

approximately 59.8 metric tons, of that approximately

12 metric tons is xenon propellant. The vehicle uses 8

gridded ion thrusters operating simultaneously powered

by large (~2,700 m2

) thin-film arrays. The spacecraft

system model has been extended successfully to similar

applications and continues to be evolved for larger mission sets. Further studies to refine the SEP transfer

vehicle system model are underway in 2002.
NASA/TM—2002-211970 8

REFERENCES

1 Mankins, J.C., and Mazanek, D.D., “The Hybrid Propellant Module (HPM): A new concept for space transfer in the Earth’s neighborhood and beyond,” presented

at the 52nd InternationalAstronauticalCongress meeting, Paper IAF–01–V3.03, Toulouse, France, October

2001.

2 Dudzinski, L.J., “Design of a Solar Electric Propulsion Transfer Vehicle for a Non-Nuclear Human Mars

Exploration Architecture,” Presented at the 26th International Electric Propulsion Conference, Kitakyushu, Japan, October 1999.

3 Murphy, D., and Spence, B., “An Ultra-lightweight

Thin-Film Solar Array System for PowerSail,” presentation at Space Power Workshop, April5, 2001.

4 MacDonald Dettwiler Space and Advanced Robotics,

Itd. (MD Robotics), Canadarm description,

http://www.mdrobotics.ca/canaframe.htm

5 MacDonald Dettwiler Space and Advanced Robotics,

Itd. (MD Robotics), Canadarm 2 description,

http://www.mdrobotics.ca/aboframe.htm

6 AEC-Able Corp. Datasheet, Feb. 2000.

http://www.aec-able.com/srtm/srtm_spex.htm

7 Edelbaum, T.N., Sackett, L.L., and Malchow, H.L.,

“OptimalLow Thrust Geocentric Transfer,” AIAA

Paper 73–1074, Lake Tahoe, NV, October 1973.

8 Kaufman, H.R., and Robinson, R.S., “Electric

Thruster Performance for Orbit Raising and Maneuvering,” J. Spacecraft and Rockets, 1984, 21, No. 2.

9 Multiple reports, Progress in Astronautics and Aeronautics: Orbit-Raising and Maneuvering Propulsion:

Research Status and Needs, Caveny, L.H., Ed.; AIAA:

New York, NY, 1984, 89.

10 Edelbaum, T.N., “Propulsion Requirements for Controllable Satellites,” American Rocket Society Journal,

August 1961, pp. 1079–1089.

11 Stuhlinger, E., Ion Propulsion for Space Flight,

(C 1964: McGraw-Hill, New York), pp. 77.
NASA/TM—2002-211970 9

Table 2. Required technology advancements projected to 2016 for SEP transfer

vehicle development

Technology Figures ofMerit Comments

Power Generation Array

Specific

Power,

W/kg

Efficiency,

%

Thin Film Photovoltaic 400 20

Advanced Array Designs 300 >40

Quantum Dots >500 60 High risk/ high potentialtechnology

Energy Storage Specific

Energy,

Wh/kg

Lifetime Depth of

Discharge,

%

Full polymer batteries 300 20 yrs

(GEO)

70 Does not include

power electronics

mass

Future Flywheel 150 > 95 kCyc. 90 Includes power electronics mass

Passive Unitized Regenerative Fuel

Cell

> 1000 80 %

efficiency

– Specific energy is

function of discharge

time

Power Processing Specific

Power,

kW/kg

Efficiency,

%

Operating

Temperature, °C

High Temperature PMAD system 3.0 95 350

1200 V Power Distribution 1.4 – –

Electric Propulsion Thruster

Specific

Mass,

kg/kWe

Lifetime,

hours

50 kW Gridded Ion Thruster 1.25 15,000 Xenon propellant

50 kW Hall Effect Thruster 1.2 10,000 Xenon propellant

1.0 MW MPD Thruster 0.3 8,000 Lithium propellant
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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AIAA–2002–3971

E–13613

WU–706–87–23–00

15

Solar Electric Propulsion Vehicle Design Study for Cargo Transfer to

Earth-Moon L1

Timothy R. Sarver-Verhey, Thomas W. Kerslake, Vincent K. Rawlin, Robert D.

Falck, Leonard J. Dudzinski, and Steven R. Oleson

Electric propulsion; Ion thruster; Orbit transfer; Low thrust; L1; Hybrid propulsion

module; Solar electric propulsion; SEP; LEO

Unclassified -Unlimited

Subject Category: 16 Distribution: Nonstandard

Prepared for the 38th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE,

Indianapolis, Indiana, July 7–10, 2002. Responsible person, Timothy R. Sarver-Verhey, organization code 6920,

216–977–7458.

A design study for a cargo transfer vehicle using solar electric propulsion was performed for NASA’s Revolutionary

Aerospace Systems Concepts program. Targeted for 2016, the solar electric propulsion (SEP) transfer vehicle is required to

deliver a propellant supply module with a mass of approximately 36 metric tons from Low Earth Orbit to the first EarthMoon libration point (LL1) within 270 days. Following an examination of propulsion and power technology options, a SEP

transfer vehicle design was selected that incorporated large-area (~2700 m2) thin film solar arrays and a clustered engine

configuration of eight 50 kW gridded ion thrusters mounted on an articulated boom. Refinement of the SEP vehicle design

was performed iteratively to properly estimate the required xenon propellant load for the out-bound orbit transfer. The SEP

vehicle performance, including the xenon propellant estimation, was verified via the SNAP trajectory code. Further efforts

are underway to extend this system model to other orbit transfer missions.