IAC-02-R.P.12
SPACE SOLAR POWER CONCEPTS: DEMONSTRATIONS TO PILOT PLANTS
Connie Carrington, Ph.D.
Flight Projects Directorate
NASA Marshall Space Flight Center
Huntsville. AL USA
Connie.Carrington@msfc.nasa.gov
Harvey Feingold, Ph.D.
Science Applications International Corporation
Schaumburg, IL USA
Harvey.Feingold@saic.com
ABSTRACTThe availability of abundant, affordable power where
needed is a key to the future of exploration anddevelopment of space, as well as to the future
economic growth of communities on Earth. One
untapped energy resource uses wireless powertransmission (WPT) from space-based facilities thatcollect the sun’s energy and beam it where it isneeded. Both microwaves and lasers can be used totransmit power. Our studies have concentrated on theuse of microwaves, but laser WPT has recently
generated increasing interest, as a promisingalternative that could overcome some of the major
challenges of microwave WPT. This paper discusses
the variety of space-based concepts for collecting the
sun’s energy and transmitting it to either the Earth or
to orbiting spacecraft, organized in categories ranging
from near-term demonstrations to a full-size pilot
power plant in space.
INTRODUCTION
Abundant, affordable power, at locations where it isneeded, is a key ingredient in the future of the
exploration and development of space. The same canbe said for terrestrial futures that support economic
development and prosperity on Earth. Both Earth andspace development will also benefit if their powersources are clean and renewable. Affordable spacesystems need self-sustaining power sources that do
not depend on re-supply launches, and countries thatare deficient in coal and oil supplies need to be
independent of imports from other countries. Space
Copyright @ 2002 by the American Institute of
Aeronautics and Astronautics, Inc. No copyright isasserted in the United States under Title 17, U.S.
Code. The U.S. Government has a royalty-freelicense to exercise all rights under the copyrightsolar power can address the energy needs of both
space and terrestrial development.
SSP SYSTEMS AND CONCEPTORGANIZATION
NASA has been taking a new look at collecting solar
energy in space and transmitting it to Earth, toplanetary surfaces, and to spacecraft. A variety ofinnovative concepts have been studied for the spacesegment component of solar power beaming. These
concepts have been organized into five “model
system categories” (MSCs), primarily on the basis ofincreasing power levels and projected dates fortechnology readiness. Planning schedules include
down-selection of concepts and their technologies forflight demonstration at roughly five-to-seven-yearincrements, contingent on funding.
The first free-flying demonstration is MSC-1, at
approximately 100kW, with a technology readinesslevel of 7 in 2007-2008; it will demonstrate WPT,
advanced solar power generation (SPG), andadvanced power management and distribution(PMAD). MSC-2 is also a 100kW system, but will
operate on a planetary or lunar surface, todemonstrate advanced technologies that cannot beadequately tested in terrestrial or Earth-orbiting
environments, and to further advance science andexploration goals. MSC-3 will provide a significant
jump in power to approximately 10MW, withtechnologies ready in 2015-2017, and will
demonstrate solar electric propulsion (SEP) in
addition to advanced WPT, SPG, and PMAD. MSC-4
represents the first SSP pilot plant at approximately
1-2GW delivered to the ground, in a time period after
claimed herein for Governmental purposes. All otherrights are reserved by the copyright owner.

MSC 1

TRL 7 in 2007-8~ 100 kW

Free-flyer; WPT, SPG and SEPS; Demo-scale; Commercial space option

2. RF Transmission
to Ground Demo:
Pilot beam fromground provides phasing
info, Multiple ground
receivers for beam
pattern calibration

1. STS releases Spartan
and SP satellite,
RF In-space WPT demo
to Spartan

5. Optional release of
6a. Optional low
Target Vehicle,
microgravity
In-spaceapplication: Laser/Optical Transmission
6b. Optional
Rendezvous & Crystal-growth
to Target Vehicle
free-flying
Capture for SEP
7. Comet Rendezvous:
furnacereturn to LEO,
Laser to surface,recovery of crystals
Spectrometry determines
Constituents
4. Optional release of
Payload3b. Comet rendezvous SEP
3a. Orbit raising SEP

Fig. 1 MSC-1 Candidate Mission Scenarios
2025. A fifth category, at the 10-100GW power levelsuitable for an interstellar power station, will not be
addressed in this paper. Within each MSC several
concepts emerge, each having salient benefits andweaknesses. An attempt was made to develop withineach concept class an evolutionary series of designs
that grow from the MSC-1 demonstration stage to the
MSC-4 pilot plant.
NASA organized the SSP work breakdown structure
into 14 teams, including research and technology
groups defined by technical discipline, applications,market, and environmental groups, and a systems
integration working group (SIWG). Faced withdefining and analyzing a matrix of concepts, the
SIWG first focused on MSC-4 concepts and thedelivery of power to the grid on Earth. Hence thepilot plant MSC-4 concepts and their technologies
will be discussed first. Concepts for lower-power,near-term MSCs will then be presented.
At several SSP workshops, both the systems andapplications working groups developed options formission scenarios that could be enabled by the power
platforms in the MSCs. These options are depicted in
Figs. 1 and 2. For the near-term demo in MSC-1, one
possible missions scenario is STS-deployment of thesmall retrievable freeflyer Spartan into the same orbit
as the 100kW solar power satellite. Spartan would beequipped with a small microwave rectenna or a PV
array for laser power beaming, and the solar powersatellite would beam power to Spartan to charge its
batteries. The power on Spartan could be used for

microgravity experiments, such as crystal growth,and then STS could rendezvous and recover Spartan
and its payload for return to Earth. Another potential
mission for MSC-1 is RF transmission to a suitableground site. Due to the relatively modest power
levels provided by MSC-1, received power levels on
the ground would be low, but would provide an
opportunity to perform pattern calibration
experiments and to demonstrate ground-based pilot
beam functionality. The 100kW power levels of
MSC-1 would also be suitable for SEP
demonstrations, such as direct drive of Hall thrusters.
A science mission scenario could entail SEP transfer
of MSC-1 to a comet rendezvous, where the powerrich platform could power remote laser spectrometry
to determine comet constituents.
MSC-2’s objective is to utilize SSP-technologies forlunar exploration. Boeing has completed a study oflunar polar applications to explore totally unlit polar
craters for ancient ice, which may provide clues toorigins of the universe, and which could be used for
LOX/LH2 propellants. A lunar lander would touch
down near a crater of interest on a tall mountain at
the pole, which is almost continuously illuminated,
where a large solar array would be deployed. The
lander then would deploy a rover, whose batteries arecharged by remote power beaming from the lander.
The rover would explore up to 150km inside the darkinteriors of the crater, receiving power as neededfrom the illuminated lander. In addition to the lunar
polar applications study, Boeing and several
universities have built and tested several small roversthat have been remotely powered both by microwave
and laser beaming.

MSC 3> 2015

2-10 MWFree-flyer;
6. Cargo delivery
to Ma rs, power
beaming to surface
or orbi ting S/C

WPT, SPG and PMAD,
SEPS; Large demo
1. Deployment& Assemblyin LEO

2. Rendezvous& Captureof Cargo Payloads
in LEO

5. Power bea ming
to a steroid surface

analyses, and sensitivity studies were performed forthese concepts, with various technologies that werecompatible with their configurations. Size and mass
data presented here represent the minimum massconcepts, for technology configurations that
presented the best case. Technology performance has
been extrapolated to expected performance in the2020 time frame.

Gravity-Gradient Concepts(Sun-Tower Derived)
Options include rotating andnon-rotating arrays
to track Sun

3. Spiral outward
using SEP

4. Optional deployment of payloads,
power bea ming to lunar surface
or orbi ting facility

Fig. 2. MSC-3 Potential Mission Scenarios•

Rotating arrays have shadowing
losses at noon & midnight•Non-rotating arrays on SPS can be
combined with terrestrial PV to follow
profile of daily electrical demand

The 2-10MW of power for MSC-3 would enable
exploration scenarios such as power-beaming to lunar
or Mars surfaces, or to operations on an asteroid
surface, as depicted in Fig. 2. Due to the size of thesolar power generation surfaces, MSC-3 would need
to be assembled in LEO. With these power levels, itwould be capable of providing enough power totransport cargo from LEO out to exploration
destinations.
MSC-4 Microwave Concepts
Three primary concept classes have been developed
and analyzed for microwave power beaming to Earth:sun-tower
configurations,abacusreflector
configurations, and the integrated symmetricalconcentrator (ISC). All of these concepts are ingeosynchronous orbit, so that they dwell for 24 hours
over a given receiving site on Earth. Earlier studies
performed trades with LEO and MEO satellites, anddetermined that GEO orbits provided the necessary
delivery times to sites, and simplified satellite
maneuvers. All of the microwave concepts must
accommodate a single, monolithic, massive phasedarray transmitter that faces Earth. Spatiallydistributed microwave transmitters would produce
grating lobes at unacceptable levels of radiation on
the ground that are located hundreds of kilometers
away from the desired receiving site.
Mass and size estimates for all three primary MSC-4
concepts were based on 1.2GW of power delivered to
the ground. Configuration comparisons, economic

Fig. 3. Sun-Tower and Gravity-Gradient
Configurations
Sun-towers and other gravity-gradient configurations
as shown in Fig. 3 are characterized by solar arraypanels aligned along a vertical backbone that collectand conduct power down to the microwave
transmitter facing Earth. All power must be
conducted by a massive cabling system through the
backbone, through a single rotary joint between thesolar array assembly and the transmitter, and theninto an array of power convertors on the back of thetransmitting array. Hence the masses of Sun-towers
and their derivative configurations are dominated by
the masses of the PMAD and power transmission
subsystems.

Another characteristic of sun-towers and gravitygradient configurations is self-shadowing of thearrays at solar noon and midnight. Trades included
rotating and non-rotating arrays. The spacing
required to minimize shadowing for rotating arrays
increases the solar array structure length and mass,
but this needs to be compared to the larger number ofclosely-packed non-rotating arrays, which have tocompensate for longer shadow times. An additional
consideration is that power provided by the nonrotating arrays could be augmented on the ground bypower from terrestrial PV arrays to match a given
location’s profile of daily electrical demand.
One of the lowest mass sun-towers is a 20 km
tethered-backbone configuration using magnetrons in
a 500m diameter transmitter, stretched-lens arrays,
and AC power cabling. Approximately 460 pairs ofrotating rectangular 40m x 200m arrays are arranged
up the tethered backbone, and have been sized toprovide the 3.4GW of required power to thetransmitter. Array voltage is modeled at 1000V, witha bus voltage of 100kV and transmitter magnetron
voltage at 6kV, so converters are required at both thearrays and the transmitter. The initial mass in LEO of
this sun-tower is approximately 22,300 MT.
Other MSC-4 concepts were developed to eliminate
the self-shadowing problems of the gravity gradientconfigurations. These new concepts reverted to theperpendicular-to-orbit-plane orientation of theReference SPS Concept used in the 1970’s study ofspace solar power, but sought to address concerns
about the single-point-of-failure of the powerconducting rotary joint between the solar arrayassembly and the transmitter that was inherent to thatearlier concept. One newly proposed configuration
replaces the single joint by multiple joints between
solar arrays mounted on spars that rotate about a nonrotating backbone mast mounted on the transmitter.
Another concept eliminates power-conducting joints
entirely by designing the solar arrays and transmitteras one unit and rotating a large but relativelylightweight microwave reflector in front of thetransmitter to direct the power to Earth. This concept
is depicted in Fig. 4 and has been named the abacusreflector.
The configuration shown on the left in Fig. 4 is a
“kite” arrangement of solar arrays in the abacus
frame that minimizes interference between the arrayedges and the transmitted energy beam. Eleven bays
contain sets of the 40m x 200m concentrator arrays,
supported by a 3D structure with masts and stays.
Alternative abacus configurations are rectangular, as
shown on the right, with a prismatic truss structure toprovide a reasonable aspect ratio. Fig. 4 also shows asmall diameter reflector bearing that could belaunched as a unit; earlier configurations had a large
diameter rollring around the outside of the transmitterto rotate the reflector. A yoke reflector mount isshown. A single centrally-located mast support for
the reflector, with a joint at the reflector center, is an
attractive mount option, but has not been pursued due
to beam interference and blockage concerns.

Fig. 4 Abacus Reflector Configurations
Advantages of this concept include continuous antisun viewing for radiators behind the array, and an
abacus structural frame for the solar wing that
accommodates PMAD cabling. The use of thestretched lens array concentrator with a shifting lensto provide seasonal beta-tracking eliminatesrotational joints between the cells and the abacus
frame, simplifying solar array installation andmaintenance. The abacus frame could also be utilized
to support other solar collection technologies, such as
solar dynamic units. Disadvantages of the reflector
approach are primarily challenges to technology,
including the assembly or deployment of a large
500mx750m reflector with a surface precision of
λ/20-λ/40, management of the reflector temperatureand thermal distortions, and a stable reflector mountthat can meet beam pointing requirements. In
addition, the abacus reflector concepts are stillplagued by massive power transmission and PMAD
systems, and tend to have the most massive structuresof all concepts studied.
Lowest mass configurations for abacus reflector
concepts include a Brayton solar dynamic system andstretched lens arrays. Both low mass configurations
use magnetrons in the transmitter and AC cabling.
For the stretched lens array configuration, 962 40m x
200m arrays are needed, and for the Brayton
configuration, 1380 units are required. Converters arenot needed between the 100kV power distribution
system and the 100kV Brayton units, but thisconfiguration still requires converters at the 6kV
magnetrons. Since the Brayton concentrators are
100m in diameter, the abacus structure isapproximately 4.6km x 3km. The stretched lensconfiguration, while requiring converters at both thearrays and the magnetrons, requires a much smaller
abacus structure, approximately 2.8km x 2.8 km. The
initial mass in LEO for the Brayton configuration isapproximately 27,000MT, and is 28,000MT for the
stretched lens cofiguration. Both are significantly
more massive than the sun-tower configurations.
The Integrated Symmetrical Concentrator (ISC)
concept in Fig. 5 was proposed to reduce PMAD andcabling masses, as well as structural mass. Incoming
sunlight is collected in two large “clamshells” locatedon the ends of a mast aligned along the orbit normal,
reflecting solar energy onto two centrally-locatedphotovoltaic arrays. The energy is converted into
electrical power and transmitted by relatively shortcables to the transmitting array. The clamshells face
the sun, and the mast, PV arrays, and transmitterrotate as a unit to point the transmitter to a receiving
site on Earth. The mast is aligned with the orbit
normal. A non-power-conducting joint between theclamshells and the mast provide rotational capability
for the once-a-day orbital tracking, and the seasonalbeta tilt.
Twoclamshell
configurations
weredeveloped: a 24-mirror
version, which provides
a 2-to-1 concentration
ratio on the solar arrays,
and a 36-mirror version,
which provides a 4-to-1
concentration
ratio.
Each mirror is planar,approximately 500m in
diameter,andismounted to the backplane structure at aslightly different angle
to form a segmented
clamshellprimary
Fig. 5 ISC
mirror. Since the ISC
is not an optical imaging assembly, the light reflectedfrom each mirror only needs to fall somewhere on the
PV array, with a goal of minimizing solar array hot
spots. The mast length is sized so that the focal lengthof the mirrors is greater than 10km, which provides areasonable spot size for the sun’s image on the PV
arrays. With this focal length and a local surface-flatness requirement of about 0.5 degrees on the
mirrors, hot spots and excessive light spillage aroundthe PV arrays are minimized. Mirrors on the outeredges of the clamshell, which could experience larger
deflections than those located interior to theclamshell, will reflect their energy on interior regionsof the PV array to reduce spillage. The mirror
reflectivity has been modeled at 0.9, and lightspillage has been estimated at 10%. Each mirror is
0.5 mill thick Kapton supported by a circumferential
inflatable toroidal ring and an inflatable back-planestructure.
An initial ISC concept placed the solar arrays on theback of the transmitter, to minimize power cabling
distances. However, the backs of both the solar arrayand the transmitter need to radiate heat, and thermalradiation estimates of a back-to-back configuration
are 90 kW/m2. Hence the ISC configurations
presented here have two separated solar arrays thatare each canted 10 degrees.
Since ISC is a concentrating optical system, solar
array temperatures have been a concern. The only
array technology that holds promise of efficient
operation at high temperatures is quantum dot, which
was used in all ISC solar conversion models. To
roughly size the solar arrays, a top-level arraytemperature estimate was made, using both 1km and
1.5km array diameters, and a worst-case seasonal sun
angle of 23.5o.
Several mass estimates for the mast have been made,ranging from 170 MT for conventional composite
trusses to 50 MT for inflatable trussed masts withstays. System models are using the inflatable mast
mass estimates.
The lowest mass ISC estimate is a high concentration
configuration using magnetrons, quantum dot solar
arrays, and AC power. The 36 mirrors are 470m in
diameter, mounted on clamshells that areapproximately 4km in diameter. The mast is 7.2km
long, and the solar arrays are 1070m in diameter. The
initial mass in LEO for this configuration is 18,000
MT.
Changing to a low-concentration ratio configuration,
but maintaining the magnetrons, quantum dots, andAC, produces a 24-mirror clamshell, with 573m
diameter mirrors.
The mast is 7.6km long.
Unfortunately the solar array is now 1770m in
diameter, and with the additional mass of these arrays
and their structure, the mass of the low-concentration
ISC is 31,500MT in LEO. Hence 2-to-1

concentration ratios are not competitive with theother MSC-4 configurations.
MSC-4 LASER CONCEPTS
Laser technology advancements in recent years, andfrequency allocation and power density issues of
microwave WPT systems, prompted consideration oflaser WPT options. Their ability to transmit powerfrom multiple apertures distributed over a satellite’s
structure or flown on separate small satellites is adistinct advantage over the monolithic single
transmitters required for microwave WPT. Lasers
enable modular design approaches, in which
localized SPG, thermal management, and PMAD
systems can be sized and built for individual lasers.
Systems can be tailored for particular markets, andthen expanded by adding modules as the power
demand increases. Beam pointing flexibility allows
redirection to other receiving sites as weather or
power demands change. In addition, the powerreceivers are PV arrays, which can be used not only
for laser power reception, but also for conventional
sunlight conversion.

1-3 suns. The individual lasers use optics in the 20-25
cm range, which reduces thermal management
challenges and reduces the mass and cost of the
optics. Innovative designs have been proposed, withlaser diodes directly mounted on the solar panels, sothat radiators can be shared. The modules also house
two laser transmitter panels, a portion of a deployable
backbone, and a mini-bus containing spacecraft
subsystems, including propulsion. The concept isshown in Fig. 6. For 1.2GW delivered to the ground,a constellation of 480 satellites is required, which
will fly in a 6000-km Halo orbit.
Boeing has also designed a modular laser powersatellite as a gravity-gradient stabilized ribbon-likearray assembled from modular units, aligning theirarray panels edge-to-edge. This non-rotating array
configuration will not receive solar energy at noonand midnight, but its PV receiver on the ground will
be powered by natural sunlight during noon,
providing power to supplement that lost from thepower satellite. The ISS-sized panels are 260m x
36m. to provide 1.2GW to the grid; perhaps as manyas 1530 panels will be needed, forming a 5.5km longsun-tower configuration, as shown in Fig. 7.

Solar Panel Segment Dimensions:
260 m x 36 m

Solar Array Panel /Laser Converter Panel

Decoupling
Ring

8 Ion Thrusters

Lasers and Optics PMAD

A vionics
Mini-bus

Laser Transmitter
Panels (2)Deployable
BackboneFig. 6. Aerospace’s Modular
5MW Laser WPT Satellite
Both Aerospace Corp. and Boeing have developed
MSC-4 concepts that take advantage of the benefits
provided by laser systems. Aerospace sized theirsatellite for a single 40MT launch with 20 laser
modules built on a common backbone and providing
5MW. An optional configuration is a 20MT satelliteof 10 laser modules, providing 2.5MW. Each module
has two laser emitters and two solar arrays.
Aerospace arranged their satellites in Halo orbits,
which are satellite constellations that appear to fly inconcentric rings when observed from the receiving
site on Earth. Separation of the satellites can ensurethat individual beams from satellites will meet laserskin and eye safety requirements, with intensities of

Deployable Radiator

Full Sun Tower Portion•1530 modules•55 km long
•Backbone can be eliminatedMultiple beams

Fig. 7 Boeing’s Modular Laser Unit.
MSC-3 CONCEPTS
MSC-3’s charter is dual use of SSP technologies for
exploration. MSC-3 is a 10MW class vehicle, andwill utilize SEP in addition to the advanced SPG and
PMAD technologies of power beaming applications.
Boeing has examined applications of SSP
transportation systems to missions beyond GEO,
including an SEP human Mars mission, an SEP lunar
transfer vehicle, and a crewless precursor “power
explorer.”A modest human Mars mission was developed thatutilizes conjunction class trajectories and cargo preemplacement strategies for use in multiple visits to
Mars. SEP performance was based on direct-drive
Hall thrusters using an Isp of 2000s and a specificpower of 3.4 kg/kW. Payload elements wereseparated into 3 components, each utilizing its own
SEP Mars transfer vehicle (MTV). The first element
is a cargo MTV with a 58MT cargo lander payload.
The second element is another unmanned cargo MTV
carrying a 58MT Mars ascent/descent stage. The
third element is a piloted transfer vehicle consistingof a 40MT crew-transfer habitat and 6MT Earthreturn capsule that can rendezvous with the second
MTV in Mars orbit and return to Earth. The SEP
Mars mission chosen for evaluation was a 2018 lowthrust Earth-Mars heliocentric trajectory with a 255day outbound time, a 600-day surface stay time, anda 156-day inbound trip time. Reusable andexpendable SEP vehicles were assessed, withpropellant for the reusable vehicle totaling about
twice that of the expendable, due to the end-ofmission Earth return propellant carried during theentire flight of thereusable vehicle. The
total initial mass in
LEO for the piloted
reusable SEP-MTV is
501MT, and 268MT
for the expendable
piloted MTV. A
sketch of the piloted
SEP MTV (witharray sizes not toscale) is shown in
Fig. 8. The cargoMTVs are 292MT forthe reusable vehicle
and 185MT for the
expendable. Powerrequirements
are1.70MW for thereusable vehicle and1.07MW for the
Fig. 8 Boeing
expendable.

vehicle staging, and the outbound cargo payload
included most of the support facilities and propellant.
A 4.7MW cargo vehicle with a 278MT payload
leaves approximately 2 years and 2 months before thecrewed vehicle leaves LEO. The crewed vehicle witha 35MT payload takes 185 days from LEO to Mars
aereosynchronous orbit, using 40MW from thestretched lens arrays to power 42 Hall thrusters. In
Mars aereosynchronous orbit, the crew vehicle willrendezvous with the crew ferry and its propellant
tanks and supplies, which has been put in placeearlier by the cargo vehicle. After a high thrusttransfer to a 500km Mars orbit, the crewed ferry
docks with the cargo vehicle. The crew transfers to
the ascent/descent vehicle on the cargo vehicle,
descends to the surface for a 3-week stay, and thenreturns to the cargo vehicle. The crew then transfersback to the crew vehicle for return to Earth. The
return trip to the Earth-Moon L1 point is about one
year, so that the total crew round-trip time is 1.6
years. The cargo vehicle launched mass is 656MT,
including 278MT of payload and 254MT of SEP
propellant. The crew vehicle launched mass is
255MT, of which 147MT is SEP propellant and 35
MT is crew, their habitat, and their return vehicle.
Abacus vehicle configurations are shown in Figs. 9
and 10.

Concept for
Piloted SEP Mars
Transfer Vehicle

SAIC
has
also
examined a variety of
SEP
exploration
scenarios using SSPderived technologies, for mission scenarios from
LEO to GEO, LEO to Earth-Moon L1, LEO to EarthSun L2, LEO to Mars, and round trip Mars missions.
One example of these missions is a fast round trip toMars using stretched lens arrays and 100 kW Hall
thrusters. The Earth-Moon L1 point was used forFig. 9 SAIC Abacus
Cargo Vehicle

Fig. 10 10MW
Abacus MTV

706kW of power is supplied by a pair of stretchedlens arrays mounted perpendicular to the orbital
plane, which rotate once per orbit to track the Sun.
The depot mass is approximately 69 MT withoutwater or LOX/LH2, and is shown in Fig. 12.

“Power Explorer”
Radar Mapper(2-4MW)
Direction• Transmission distance of flightof 10 - 1000 km forlunar & planetary
mapping• 19 MT dry mass
(45 MT initial)• 2 Shuttle launches,
assembly in LEO

2

Sun
Side
scanning
radar

3

26 m

38 m
5

1
6

7

7 6

5

4

Fig. 11 Boeing MSC-3 “Power Explorer”
Boeing also assessed a crewless SEP precursor
vehicle, depicted in Fig. 11. Named the “Power
Explorer,” it uses four 52m x 31m PV arrays to bothdrive Hall thrusters for propulsion and to power aradar beam for in-situ planetary prospecting or orbital
debris detection. Two identical units, which are
docked in orbit, are each deployed from a single
launch of a Shuttle or Delta IVH.
MSC-1.5 CONCEPTS
The significant increase in power capability from a
100kW MSC-1 to a 10MW-class MSC-3 prompted
the insertion of an intermediate model system
category named MSC-1.5, since MSC-2 was alreadyassigned to lunar/planetary applications. MSC-1.5
was categorized as a 200kW to 3MW classdemonstration of advanced SSP technologies, in the2012 time frame. Two concepts in this class are thecryogenic production and storage depot for LEO
applications, and Boeing’s “Skylight” megawatt-classsatellite.
The propellant depot is deployed in a 400 km circularequatorial orbit, where it will receive tanks of waterlaunched from Earth, will convert the water to liquidhydrogen and oxygen, and will store up to 500 MT of
cryogenic propellants. The propellant stored in the
depot can support transportation from low Earth orbit
to geostationary Earth orbit, the Moon, LaGrange
points, Mars, etc. The tanks are configured in an inline gravity-gradient configuration to minimize dragand settle the propellant. Temperatures can be
maintained by body-mounted radiators, which will
also provide some shielding against orbital debris.

4

1
2

1.
2.
3.
4.
5.
6.
7.

3

80 m

LOX/LH2 Storage Tank
Transfer Vehicle Docking Port
Radiators
Solar Arrays
Water Docking Port
Water Storage Tanks
Electrolysis System

Fig. 12. MSC-1.5 Propellant Product Depot
Another MSC-1.5 concept developed by Boeing is“Skylight.” It is a sun-tower configuration with ten
pairs of rotating solar arrays, similarly sized to the
ISS arrays but utilizing advanced higher performance
SPG technologies. The spacecraft overall length is
170m, and the 2.7MW of power from the arrays
could provide approximately 1kW of beamed powerusing a distributed laser WPT system. Skylight has an
estimated mass of 53MT without energy storage. Fig.
13 shows the Skylight concept. Boeing has developed
several concepts that utilize advanced ISS-sizedarrays, with the goal of eventually replacing the ISS
arrays with newer high-performance SPG
technologies developed for SSP.
MSC-1 Concepts
This model system category is a near-term 100kWclass space platform whose primary function is todemonstrate SSP-technologies in WPT, SPG, andFig. 13 Boeing’s
MSC-1/5“Skylight”
Powersat

PMAD,
and
whose
secondary functions are
SEP and non-missioncriticaladvanced
technologies. System design
mass goals are 5MT for thedry spacecraft, 2MT fortechnology
flightexperiments, and 2MT for
propellant.
Technology
performance goals include
100-150kW of on-board
power, SPG conversion
efficiencies of greater than
30% and at least 250 W/kg,
and
300-500V
power
systems. Demonstration ofadvanced
electric
propulsion is also a goal; an
example is direct-drive ofHall thrusters. A five-year
lifetime, minimal on-orbit
assembly, and a single
shuttlelaunchareguidelines.

power plug’s ISS-sized arrays, up to 18kW could betransmitted by the beam, over much longer distances.
Boeing has also investigated an orbiting lunar power
plug, to provide power or illumination at the Moon.
They have developed a charging/thrusting strategy
for the transfer orbit that takes about a year, and uses
SEP except for trans-lunar injection and lunar
capture, where chemical bi-prop is used. The initial
mass in LEO is 20MT, of which 2.8MT is SEP
propellant and 2.5MT of TLI&LOC propellant, and
260kg is payload. A Mars version of this power plug
has also been proposed, and even at the low
insolation levels near Mars, up to 17kW may beavailable through laser power beaming to a Mars
infrastructure. The power plug concept for Earth,
lunar, and Mars missions is shown in Fig. 14.

Boeinghasdeveloped
several variations of an
MSC-1 “power plug inspace,” for space-to-space
power beaming in Earth orbit, and lunar and Mars
power spacecraft (LAMP). The power plug was
designed to fit into the shuttle cargo bay, and to use
the existing ISS array structural design, but outfittedwith advanced solar cells. A parabolic WPT reflector
is sized at 4.5m in diameter, and the array span is
70.8m x 11.7m. Hall thruster and advanced battery
technologies would be demonstrated.
For Earth-based missions, the power plug would be
deployed in LEO and then spiral out to a 24-hour
period low eccentricity orbit at an inclination of 10o
to 14o, where it would remain in sunlight. From thisorbit it would be capable of beaming power to GEO
satellites in eclipse. For existing GEO satellites, thebeamed power would be optical, either at laser
frequencies tuned to the receiving satellite’s PV
arrays, or using high-intensity white incoherent lightfor multibandgap arrays. Microwave or millimeter
power beaming could be an option, but the rectenna
on a receiving satellite would most likely be no larger
than 10m, which limits beaming to only shortdistances (<1km for 5.8GHz, or <40km for 245GHz,for ~4% power incident on the rectenna). For laser
power transmission, and advanced solar cells on thePower PlugGEO Satellite in Eclipse

Fig. 14 Boeing’s MSC-1 “Power Plug”
Since the MSC-4 ISC is a promising long-term lowmass concept, an ISC-derived 100kW-class concept
was investigated. Several small-scale ISC
configurations are shown in Fig. 15.

made as the clamshells tilt to track the beta-angle ofthe sun.

1 d.o.f. rotation

Transmitter
Diameter 4.6 m

Reflectors
Solar array
Radiator

Boom
length49 mSolar array
Diameter 9.8 m

Instrument shelf
Laser

Mirror
Diameter
4.4 x 6.2 m

Fig. 15 MSC-1 ISC Configurations
Only one solar array is used, so that the back side ofthat array will have a better opportunity to radiate
heat. An instrument shelf is mounted in a ring aroundthe outside of the solar array and radiator unit, andeither a microwave or a laser WPT transmitter can beaccommodated.

Boomlengt h30 m

Mirror
Diameter4 x 5.6 m

Solar array
Diameter 13.8 m
Transmitter
Diameter
4. 6 mFig. 16 ISC Configuration of ellipticalmirrors with spherical curvature
Uneven and changing illumination on the solar array,from the multiple mirrors of the clamshell, presents
difficulties for current and near-term SPG, PMAD,
and thermal management technologies. A change
from flat to convex mirrors that produce an image
illuminating the entire solar array is a solution. Since
the clamshells are nominally tilted 45o from the solar
array, two options for correcting the elliptical image
could be applied: elliptical mirrors with a sphericalcurvature (Fig. 16), or circular astigmatic mirrors(Fig. 17). Seasonal corrections would also need to be

Fig. 17 ISC Configuration with circularastigmatic mirrors
An
offsettwoclamshell, two-solar
array configuration,
shown in Fig. 18,
has the best chanceof rejecting the solar
array heat. The
transmitter diameter
has
also
been
increased to help
reduce temperatures,
but will now have to
be assembled onorbit.

Fig. 18 Offset ISC

Since these MSC-1
Configuration
ISC configurations
have 2-to-1 or 4-to-1
concentration ratios,they have the same high temperature solar arrayissues that the MSC-4 ISC must manage.
Unfortunately, quantum dots is the only technologythat holds the promise of functioning at thetemperatures on the MSC-4 ISC solar arrays, and thistechnology is not at a maturity level to be used for thenearer-term MSC-1 applications.
Other issues for the MSC-1 ISC are the packaging
and deployment complexity of a concentrator optic
assembly versus conventional solar arrays, close
locations of the solar array and transmitter, which
increase cooling requirements for both, payload
accommodation (which can be done), and propulsion

system accommodation. Benefits include solar arrayviews of deep space to aid thermal management,
relatively reasonable reflector surface accuracies, thesmall solar array, short cable distances, no slip joints,
and lightweight structures.
A more traditional configuration for MSC-1 is shownin Fig. 19. A standard spacecraft bus, but with aredesigned, higher voltage PMAD system, supports
two 5.5m x 30m planar arrays populated withstretched lens arrays. SEP options include two 50kWHall thrusters, and a 70m2 radiator (not shown) to
manage up to 90 kW of heat. A preliminary massestimate of 7MT IMLEO includes 1.4MT of Krypton
for transfer to GEO, and 2MT for technology
experiments.

Fig. 19 Planar Array MSC-1
One possible mission scenario using this spacecraft is
an ISS laser power beaming experiment, in which asmall laser would beam power from ISS to the coorbiting MSC-1 platform, as shown in Fig. 20. Site 1of the ISS JEM-EF could provide 5kW of power andthermal management to a laser unit, which would beinstalled without EVA using the JEM robotic arm. A
12m diameter octagonal PV array on the MSC-1
satellite would receive the beamed power; at a
distance of 10-20km in front of ISS, modest laser
pointing accuracies of 1 arcmin would be needed. A
small amount of power running in a wire around theouter edge of the PV array structure would thermallymark the boundaries of the target for infrared sensors
at the JEM-EF laser site, to aid in target centroid
detection for beaming acquisition and pointing during
power transmission. After the ISS laser beaming
experiment is complete, the laser is removed from theEF-1 by the JEM arm and returned through the
airlock, where it can later be returned to Earth by theSTS. The MSC-1 satellite would then use its onboard arrays to direct drive the Hall thrusters andspiral out and away from the ISS orbit.

Fig. 20 ISS Laser Power Beaming
ExperimentTechnologies that could be demonstrated on a power
MSC-1 platform include energy storage, innovative
thermal management, distributed control of flexible
structures, microgravity manufacturing, propellant
production and storage, and alternative WPTtechnologies, such as small-scale ISC concentrator
systems. This platform would also provide
opportunities to demonstrate advanced materials,
deployment of inflatable structures, and spaceassembly and robotics.
After the WPT demo from ISS, MSC-1 could use its
array power to perform experiments in high power
communication, space-based radar, or space science.
It could also use the array power for WPTexperiments with other spacecraft, such as laser
annealing of solar arrays, power beaming for energy
and control of solar sails, or a lunar/planetary surface
demonstration, such as powering a rover. Possible
destinations include the Van Allen belts, HEO, GEO,
Earth-Moon L1, the moon, Mars, Earth-Sun L2,
asteroids, or comets.
Another scenario for a first technology flight demo
eliminates the MSC-1 satellite and just demonstrates
laser power beaming from ISS to the ground. This
experiment could use an existing beaming expanderthat would fit through the JEM airlock, and could put
a reasonably-sized spot on the ground, where a PV
array could receive the power. For 5kW at the JEM
EF-1 site, received power could be on the order of
0.5-1kW. This demonstration could also perform
retro-directive beam experiments.
Other suggestions include collaboration with the
DoD, such as using second copies of the Orbital
Express satellites to perform space-to-space power
beaming and receiving. Collaboration with othergovernments may also provide suitable satellites.
Using the ISS as a platform for space-to-space power
beaming experiments to MSC-1 means that the initialorbit for MSC-1 is 51.6o inclination. Hence the same
solar array orientation, attitude, energy storage,
propulsion system location, and configuration
compromises that ISS has contended with will also
effect MSC-1’s design. A simpler approach is todeploy MSC-1 in a sun-synchronous orbit,
eliminating the need for a second DOF in solar arraytracking, greatly simplifying energy storage andthrust vector orientation for propulsion, providing
more uniform illumination on the solar arrays, andreducing drag from the orientation of the solar arrays.
However, the sun-synchronous orbit option would
require a large launch vehicle to place MSC-1 in ahigh inclination orbit. In a sun-synchronous orbit,
WPT demos would need to be from MSC-1 to theground, or another satellite in this orbit would have tobe provided to receive power.

3.

The Boeing Co., Space Solar Power Systems
Studies and Analysis, Report No. NAS899140, Nov. 14, 2001.

4.

Penn, J.P. and Law, G.W., Systems Studies
and Analysis of the Space Solar Power(SSP) Exploratory Research andTechnologies (SERT) Concepts andApplications, Aerospace Corp. Report No.
F04701-00-C-0009, Mar. 9, 2001.

5.

Penn, J.P. and Law, G.W., Operational and
Demonstrator Laser Concept Model
Development, Aerospace Report No.
F04701-00-C-0009, Jun. 10, 2002.

6.

Feingold, H., and Carrington, C., Evaluation
and Comparison of Space Solar Power
Concepts, IAC-02-R.1.08, World Space
Congress, Oct. 2002.

7.

Perkinson, D., Solar Power Satellite
Integrated Symmetrical Concept, Sverdrup
Technology Final Report, Sept. 2001.

8.

Carrington, C. et. al., The Abacus/Reflector
and Integrated Symmetrical Concentrator:
Concepts for Space Solar Power Collection
and Transmission, AIAA-2000-3067, 35th
Intersociety Energy Conversion Engineering
Conference, Las Vegas, NV, July 24-28,
2000.

9.

The Boeing Co., Space Solar PowerTechnology Demonstrations for Lunar Polar
Applications, SERT Contract Report NAS899147, Nov. 1, 2000.

CONCLUSIONS
A great variety of SSP concepts have been developed
over the past few years. In each model system
category, development of concepts has helped ourteam identify design issues and the technology
advances that might be made to address them.
Systems modeling of the concepts has provided
insight into which design features and technologies
have the greatest impact in reaching the ultimate goalof usable space-based solar power.
ACKNOWLEDGEMENTS
The authors would like to acknowledge thetremendous amount of work done by the SERT/SSP
teams in developing concepts and missions. In
particular, Don Perkinson of Sverdrup, Jay Penn and
Glenn Law of Aerospace Corp., and Seth Potter andMark Henley of the Boeing Corp. have provided a
substantial amount of the work that has beensummarized in this paper.
REFERENCES
1.

An International Forum on Space Solar
Power, NASA HQ, Washington, DC, Jan.
12, 2001.

2.

Feingold, H. et. al., SERT Systems,
Integration, Analysis and Modeling, Report
No. SAIC-01/1045, NASA MSFC Purchase
Order No. H-32458D, June 2001.

10. The Boeing Co., Space Solar Power &Platform Technologies for In-Space
Propellant Depots, SERT Contract NAS899140, Mod. 2, Task 3 Final Report, Nov.
14, 2000.
11. Potter, Seth, et. al., A Cryogenic Propellant
Production Depot for Low Earth Orbit,
International Space Development
Conference, Albuquerque, NM, May 24-28,
2001