1.50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition AIAA 2012-062909 - 12 January 2012, Nashville, Tennessee Solar Power Satellite Demonstration System for Lunar and Planetary Exploration Jean N. Koster1University of Colorado, Department of Aerospace Engineering Sciences, Boulder, CO 80309-0429 The lunar outpost has an estimated need of 12 kW power to support fourastronauts, in situ energy production, and a rover. An Outpost Solar Satellite PowerPlant (S2P2) is proposed to provide energy from lunar orbit to the lunar outpost forlife support and in situ production needs. If such a satellite system is established firstbefore a moon landing by astronauts the development of an outpost on the surface ofthe moon will become easier as energy supply and some energy infrastructure isalready established. This concept applies to other planets, such as Mars, targeted foroutpost development as well. I.Introduction A few years ago the President of the United States had chartered NASA to develop a permanent lunarbase by 2025 [1, 2]. This effort requires an incremental buildup that begins around the year 2020. Theprimary goal was to prepare for human exploration of planet Mars. The endorsement may have shiftedmeanwhile, but a future return to the Moon remains likely. The major issue related to a lunar outpost isthe availability of energy at the base. Such a lunar outpost requires a supply of energy in the form of heatand electricity. A current concept to develop lunar based solar power systems includes concentrated solarpower for heat generation and photovoltaic systems for electricity production [3]. However, thatinfrastructure will be installed only after astronauts have landed. One concept not discussed at length is asolar power satellite system orbiting the moon which would beam the power to the surface of the moon.Such a system can be put in operation before astronauts land on the surface of the Moon thus providing aninfrastructure of energy supply soon after landing. This satellite solar power system (SSPS) concept wasfirst developed many years ago as a potential energy collector to supply Earths needs and has beenreconsidered only recently [4-6]. In 2001 the National Academies published an assessment of NASAsSpace Solar Power work [7]. Recommendations were made in favor of the research and development butthey were not sustainable.1 Professor, Dept. Aerospace Engineering Sciences, University of Colorado, Boulder, Colorado 80309-0429, U.S.A.,Associate Fellow.-1- American Institute of Aeronautics and AstronauticsCopyright 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2. Here we propose a demonstration satellite solar power system, the SOlar POwer SAtelliteDEmonstration System (SOPOSADES) that could evaluate feasibility for providing energy from orbitingsatellites. More specifically the following design proposes a Lunar Outpost Solar Satellite Power Plant(LOS2P2) to supply a moon base with power. A similar concept may apply for Mars and even othercelestial bodies. A lunar outpost shall have the capability to receive the energy from the solar powersatellite.The system may comprise one or more satellites orbiting the moon. The orbiters shall collect sunlightand beam the energy to the receiving station at the lunar outpost. It shall provide the outpost with energyfor life support, communication, exploration, research and regolith processing (in-situ production). Theparent requirements of the major sub-systems, mission design, satellite system, and receiving station aredetailed below. As an example some in-situ resource utilization processes are mentioned as energyconsumers. Figure 1: Parent Requirements of Major Sub SystemsThe SOPOSADES demo system may ascertain feasibility of a large scale solar power satellite forEarth applications. The proposed lunar satellite solar power system is an excellent test bed to develop-2- American Institute of Aeronautics and Astronautics 3. technologies that could be used in Earth orbiting satellites solar power systems, should that technologybecome important or necessary for energy supply to planet Earth.II.Mission Design requirementsThe SOPOSADES orbiter will surround the moon in a defined orbit. This orbit could be designed fora) the main mission in a piggyback launch scenario or b) it could be independently designed forSOPOSADES itself.In a piggyback scenario, it is expected that SOPOSADES will surround the moon with a lunarexploration or scientific mission. Such an orbit would be similar to the orbits used for the latest missionsto the moon. Most missions are nearly circular in altitudes from 50 km to 300 km. Low altitudes aredesired for high resolution images of the lunar surface.If SOPOSADES is designed as an independent mission, an optimal lunar orbit can designed. Themain challenge in lunar orbit design is the inhomogeneous gravitation field, which occurs due to massconcentrations below the lunar surface [8]. For orbit altitudes below ~ 750 km (Low Lunar Orbits), thoseinhomogeneities cause the main perturbations in a lunar orbit. Orbits above this altitude are moreinfluenced by the third body perturbation due to earth. Those perturbations require a higher amount ofpropellant for the attitude control system (ACS) and result in a higher launch mass.The mission design, when fully tested, calls for three solar power satellites to provide continuousenergy to the lunar outpost. Initially for development purposes a single orbiter would be sufficient toestablish confidence in the design. The orbit shall ensure extended energy beaming periods and shallmake the tracking process easier. The choice of the orbit will define fly-by periods, coverage and altitude,as well as the required number of satellites to achieve continuous energy beaming to the outpost receiver.Redundancy of energy supply for the operation (life support, production) at the lunar outpost isrequired for the safe operation of the outpost. That energy can be supplied by solar power systems to bedeployed at the lunar base itself.III.System ArchitectureThe proposed solar power satellite demonstration system design consists of a lunar orbiter and alander (Figure 2). The orbiter will collect solar radiation and convert it into a transmittable form ofenergy. The orbiter module includes communication and navigation and control for the satellite systemcontrol and command. It also includes a selection of data acquisition systems to ensure its functionality asa solar collecting and radiation system. The solar radiation collector is composed of a deployable largeenergy collection dish. The received energy has to be stored temporarily on the orbiter for periodicbeaming to the receiver at the lunar surface. The orbiter and the lander/receiver are integrated as onesystem for the launch phase of the mission. A mechanical connection device is required as well aselectrical umbilical connection between orbiter and lander. -3-American Institute of Aeronautics and Astronautics 4. The lander itself includes a production facility that will start operation at the lunar surface. As it isseparated from the orbiter the lander needs its own control and communication system. To ensure itsoperational success a select number of measurement systems need to be included. The landing device andpropulsion system may be of similar style as successfully used during the Apollo missions. The energyreceiving antenna may be a deployable dish structure that can track the orbiter. Energy storage is requiredfor the operation of the lander as well as the production facilities. To prevent failure excess energy mayneed to be dumped by radiation to the environment.Figure 2: Possible SOPOSADES architecture as designed for launch.The proposed lander is a demonstration package that lands on the surface of the moon andsimulates an energy consumer. The transmitted energy will be collected by the lander system onthe lunar surface, which converts the received energy form into electrical power. A conceivablelunar material production facility may include oxygen production for fuel and life support,hydrogen for fuel support or water generation with oxygen, for sintering lunar material intobuilding materials, and many other in-situ resource utilization processing. The productionfacilities require communication and control and must include measurement capabilities foroperational success. Three main steps are part of the processing: acquisition of regolith, -4-American Institute of Aeronautics and Astronautics 5. processing, storage and analysis of products. To ensure full success additional power resourcesmay have to be acquired at the lunar surface itself.Figure 3: Possible SOPOSADES Systems requirements. -5-American Institute of Aeronautics and Astronautics 6. A. System RequirementsA select number of system requirements for the proposed SOPOSADES are listed in Figure 3. Theorbit for each fly-by at the receiving shall be the same to make tracking easier. The period of the orbitdefines the amount of energy which needs to be beamed during every fly-by. During power beaming thesatellite has to track the target and continuously adjust the rectenna. Wobbling during power beamingmay cause the beam to miss the rectenna and hit structures next to the rectenna. To adjust the satellite forthe power beam and point the solar collector to the sun the attitude control system has to control themovement of the satellite. For the attitude control system different technologies are available; thrusters,spinning wheel, field coils etc. For the thrusters different kinds of fuels are available.The energy management system controls the energy of all other subsystems. It saves the energy in thestorage system and provides the beaming device with the energy for the beam. Storage systems can bebatteries, fuel cells, fly wheels, other. This means that the satellite can operate during possible orbitaleclipse. For the power beam are two possible scenarios possible: 1) Energy in energy beamed outwithout satellite storage capability; 2) Energy storage on satellite for intermittent transmission to thereceiving station. During an orbit around the moon the thermal conditions due to the sun and thebackground temperature are different, which necessitate thermal control. The thermal control system hasalso the function to manage excess energy in a useful way (heating for example). To ensure that thebatteries, electronic devices and other subsystems will not overheat or freeze the TCS has to handle thetemperature in a defined range. This thermal control can be done by an active or passive thermal system.Optimizing the efficiency of the solar radiation collector system is an important requirement. Only ahighly efficient system can satisfy all requirements for a manned lunar outpost as discussed in the parentrequirements. Pivoting mechanisms for solar collectors are well proven in many space missions. The solarradiation collector could be a photovoltaic or solar thermal power system. For a solar thermal powersystem a power conversion subsystem needs to be investigated. To collect a high amount of solar energythe SRC should be at all times pointed to the sun during the sun-phase. The device will convert electricityinto transmittable energy with high efficiency. During the fly-by the beaming device tracks the Lander onthe lunar surface. The choice of the physical device depends on the selected beaming technology(microwave or laser). To send the energy beam to the receiving station a tracking between the satelliteand the receiving station is required. The satellite has to send housekeeping and measurement data to thecontrol center on Earth to show feasibility of the mission. The control center has to send commands to getaccess to the satellite and the subsystems. There are major subsystem requirements for the receiving station system. The receiving station shallstore enough energy to operate all control systems, communication, data storage, rectenna pointing andthe processing subsystem for defined time. The antenna has to receive sufficient power to provide energyfor consumption by receiving station subsystems. The antenna has to unfold automatically upon landingand shall adjust the satellite automatically during each fly-by. The receiving station shall track the satelliteto communicate and receive the energy beam. The lander shall manage excess energy. The receivingstation shall communicate with the orbiter and the Earth.-6- American Institute of Aeronautics and Astronautics 7. B. Orbiter SystemThe orbiter satellite will collect the solar radiation and convert it into a transmittable energy form.This energy will be transmitted to the lander deployed on the lunar or Martian surface. The orbiterdesign (Figure 4) includes the orbiter vehicle design, the choice of attitude control system, the solarradiation collector, the power beaming technology, the energy storage device and the data acquisitionand communication system.Figure 4: Orbiter System Alternatives.1. Solar Radiation CollectorTwo different designs of the solar radiation collector system are investigated in this study: a) aphotovoltaic technology and b) a thermal power technology (Figure 5). The criteria for the quantitativedesign of the SOPOSADES Solar Radiation Collector are primarily reliability. Independent of the choice oftechnology, the solar radiation collector has to receive sufficient energy to provide a 100 Watt peak atthe receiving lunar lander. The power beaming device is an experimental technology. The other drivingassessment criteria are a lightweight and high power conversion. Lightweight is required because of themass requirement for the SOPOSADES orbiter system. The solar radiation collector also needs to providea high voltage power output which is required for some microwave generators.-7- American Institute of Aeronautics and Astronautics 8. Figure 5: Detailed View on the Solar Radiation Collector SubsystemsMost photovoltaic systems for satellites use a rigid solar array design. These panels are folded ontoeach other until deployment. In comparison to the rigid array the Ultra Flex (UF) solar array is a rigidpanel that reduces storage room and weight by 25%. The UF array is a fan-folded circular system thatdeploys around a central hub structure. Inflatable structures for solar arrays are also developed with afocus on inflatable-rigidizable structures. The solar arrays deploy as the central tube is inflated axially.LGarde Inc. has successfully developed an inflatable-rigidizable solar array for the NASA NewMillennium Program. Other types of panel design are solar concentrator arrays. This technology uses reflective surfaces oroptical concentrators that bundle the solar radiation and focus it at a solar cell. Available solarconcentrator arrays are the Cell Saver (Able Engineering), the Stretched Lens Array (SLA) and theStretched Lens Array Square Rigger (SLASR) (Entech Inc.). The SLA and SLASR are improved SCARLET(Solar Concentrator Array using Refractive Linear Element Technology) solar arrays which flewsuccessfully on the NASA Deep Space 1 mission in 1998-2001. SLA is a rigid concentrator panel that usesflexible Fresnel lenses to increase the output of the solar cells. This leads to a lower number of cells for -8-American Institute of Aeronautics and Astronautics 9. the equivalent power. The SLA on the Square Rigger Platform uses folded lever arms with rigid cells ona flexible polymer blanket. The flexible Fresnel lenses are also used for this SR design.Thermal power is another energy support system alternative for SOPOSADES. Such solar dynamicsystems bundle the solar radiation and specific spot and use the generated heat for a thermal cycle. Atypical configuration of a solar dynamic power system is shown in Figure 6. In general a solar dynamicsystem consists of a solar concentrator, the solar heat receiver, the thermodynamic power conversionunit, the waste heat radiator, and the power distribution system. NASA prioritized the Brayton cycle andthe Stirling engine for development and research in solar dynamic (SD) systems in 1999.Figure 6. General configuration of a solar dynamic power system [9]For a solar dynamic system, the incoming sun radiation must be concentrated on a defined spot at thesolar heat receiver. Solar concentrators may be designed to include one or two collector subsystems:the primary and secondary concentrator. The primary solar concentrator can be rigid panelconcentrator, deployable Spline Radial Panel (SRP) concentrator or inflatable thin film structures. Therigid design has a lowest mass per area of 6 kg/m[10]. The 1992 patented spline radial panelconcentrator from Harris Corp., should have a mass per area value of 2 kg/m [11]. SimultaneouslyLGarde Inc. developed a lightweight inflatable radio antenna, which deployed successfully in orbit in1996 [12].The advantage of using a secondary concentrator is, that concentrated solar radiation can beconcentrated a second time to provide higher radiation flux to the solar heat receiver [13]. Alternativedesigns for a secondary concentrator are refractive and reflective concentrators. If multipleconcentrators are used for a SD system, the solar flux of each primary concentrator must be"transported" to the shared solar heat receiver. For this purpose optical fiber cables were considered.NASA assessed both designs for solar thermal applications and came to the conclusion that refractivesecondary concentrators have a significant performance advantage. Refractive concentrators redirectthe focused light from a primary concentrator by refraction and internal reflection into the cavity of thesolar heat receiver.-9- American Institute of Aeronautics and Astronautics 10. The refractive secondary solar concentrator was originally invented for the NASA Shooting StarExperiment, but this project was canceled, with limited development efforts since year 2000 [13]. Asolar dynamic (SD) system for SOPOSADES would require further investigation of that technology.A solar dynamic (SD) system for SOPOSADES with a single primary concentrator could be designed incombination with a secondary concentrator. The secondary refractive concentrator offers better energythroughput efficiency than a secondary reflective concentrator. Beneficial is also that up to 80% of theIR radiant energy can be reflected by additional coating. Also no active cooling of the concentratormaterial is needed, which relaxes the thermal subsystem requirements. No application for a SD powersystem with secondary concentrator is currently under development, but the configuration of asecondary refractive concentrator in the solar thermal propulsion unit for the Shooting Star Experimentcould be adopted by SOPOSADES in a modified form.The main goal of a solar heat receiver is to collect the concentrated solar radiation with a high absorbingand thermal resistant material and heat up a working fluid for further energy transport. In general,two concepts of solar heat receivers exist: a) Direct Gain concept, University of Surrey, UK [14]; theconcentrated solar flux is heating the receiver walls, which are heating the working fluid. b) Thermalstorage concept; the working fluid is heated by a thermal storage material which was heated by thesolar flux. The direct gain receiver is lighter than the thermal storage receiver concept, because of theabsence of thermal storage material. The major disadvantage is that the direct gain concept works onlyduring sun exposure. Solar heat receivers are discussed for solar dynamic (SD) systems and alsoinvestigated for solar thermal propulsion system applications [14, 15].The Closed Brayton Cycle (CBC) was under development at NASA for dynamic power conversion systems[16]. As heat source, solar energy and nuclear energy are in discussion. Currently the CBC is consideredas the power conversion system for the Prometheus program, the NASA space reactor developmentproject. In the 1990s, NASA designed and tested a 2 kW SD-CBC as an energy support system of theSpace Station Freedom. This system was never integrated into the International Space Station (ISS) butseveral ground tests were made.Stirling convertor engines are currently under development for future use in radioisotope powersystems [17]. The so called Advanced Stirling Converter (ASC) is a product of Sunpower Inc. anddemonstrated a converter efficiency of more than 30%.Thermal storage [18] is desired for a solar dynamic (SD) system that shall work even through an eclipse.Such a system was ground-tested by NASA for a 2kWe at 120 VDC SD-Brayton process. The receiversubsystem combines three functional elements: the heat receiver, the heat source heat exchanger, andthe thermal storage device. The thermal storage device contains a phase change material (PCM). Thissystem configuration with an attached Brayton process was tested by NASA from 1994 - 1998, for 886hour of simulated solar heating and 783 hour of power generation in the solar simulator facility at Lewis(Glenn) Research Center.- 10 - American Institute of Aeronautics and Astronautics 11. 2. Power BeamingFor SOPOSADES it is essential to demonstrate the functionality of the power beaming technology.The power beaming to the lunar surface should be operational during sun exposure phases and duringeclipse phases. For the wireless transmission of energy from the demonstration satellite to the lunarsurface, two technologies exist: Microwave Power Transmission Laser Power TransmissionThe power transmission via microwaves was recommended by Glaser in 1973, for the transmission ofenergy from the Satellite Solar Power Station (SSPS) to Earth [19]. The most important parameter for asolar power satellite is the efficiency of the power beam.In general, a microwave power transmission system consists of three elements (Figure 2.44:Microwave Power Transmission 1.) the microwave transmitter, which converts the supplied direct current(DC) power into microwaves 2.) the free space microwave transmission, and 3.) the receiver, whichreconverts the received microwaves into DC power. For a continuous power transmission only microwave(MW) converters with a continuous wave (CW) output are considered.For the laser system beaming device, are several different kinds of lasers available. In thisassessment a selection of lasers are assessed to get an idea what kind of laser is the best forSOPOSADES. In general the energy beaming will consist of three different phases: a) electrical to opticalconversion, b) free space beaming and c) optical to electrical conversion. To operate with a laser systemthree devices are required. The first device is the power converter for the laser, second one is the lasersystem and the third one is a receiving station. These systems were mostly tested in communicationsystems such as the Mercury Laser Altimeter test on the Messenger mission in 2005.The antenna design has a big influence on the entire spacecraft design. Several antenna designs forSolar Power Satellites were proposed in the past. NASA recommended three designs for the antenna in1981, one for each power conversion technology: the Klystron, the Magnetron and the Solid-Statetransmitter [20]. One of the developed laser systems is the Airborne Laser Test Bed (ALTB) for the US Air Force [21].The ABL was designed for use against tactical ballistic missiles. For laser power beams in space, laserbeams of high optical quality and high average power are essential for effective power transmission.Diode laser arrays have high potential for high power lasers [22].2. Energy Storage SubsystemAs the satellite collects solar energy during most of the orbit, it is necessary to store energy fortransmission during fly-over times. The energy storage sub system may become the highest mass itemon the orbiter. Energy storage is also required for the lander system. There are several ways to storeenergy in spacecraft: chemical, mechanical, and physical.- 11 - American Institute of Aeronautics and Astronautics 12. Chemical energy storage subsystems are batteries and fuel cells. A mechanical energy storage alternative is the flywheel. Considered physical energy storages are latent and sensible heat storage materials.3. OrbitThe main goal for SOPOSADES is the successful demonstration of the power beaming technology.Therefore, the orbit of SOPOSADES shall offer sufficient time of sun exposure for energy conversion,storage, and the energy transmission to the lunar surface.For circular low lunar orbits (50 km altitude), the longest expected eclipse per orbit is 48 minutes with asun exposure time of ~75 minutes. This worst-case orbit scenario will be a baseline for the design of theSOPOSADES subsystems. For longer sun exposure periods, eccentric orbits would be beneficial.For the power transmission between the SOPOSADES Orbiter and the Lander, both need to track eachother periodically. Advantageous for the SOPOSADES pinpointing requirement and the wireless powertransmission efficiency would be a lower lunar orbit. Due to the negligible lunar atmosphere, it ispossible to perform orbits in a very low mean altitude. The tallest mountain on the lunar surface (heightof 12 km), is limiting the orbit altitude.The transmission period is also a design issue for SOPOSADES. The functionality of the system can bedemonstrated in short periods (a few minutes). Potential fast movements of the transmitting andreceiving apertures can cause moments and torques, which must be suppressed by the attitude controlsystem.Example in Figure 7: The defined maximum angle of sight limits the transmission to 51 km (26 km Orbit),96 km (50 km), 186 km (100 km), and 352 km (200 km) respectively. 400Transmitting Distance [km] 300 200 1000 -90-60-300 30 6090Angle of Sight [] 26 km Orbit 50 km Orbit 100 km Orbit Figure 7: Transmitting Distance as a Function of the Angle of Sight - 12 -American Institute of Aeronautics and Astronautics 13. The design drivers and critical parameters for SOPOSADES are: The Transmission Distance The Transmission Efficiency The Transmission PeriodOther critical parameters are the orbital eclipse and sun exposure period for the thermal household andlunar eclipses for the final mission design. For this study a circular lunar orbits will be assessed on thetransmission period and the aperture sizes. The aperture size is a function of the transmission distanceand efficiency (Table 1). From this limited analysis the orbit at 100 km altitude is the desired orbit forthis demonstration system. Table 1. Orbit Assessment Values 26 km 50 km100 km 200 km Orbit Orbit OrbitOrbit Transmission Period0.881.673.296.41 [min] Aperture Size [A]n/a n/a 1369 - 13852734 - 28094. Operational PhaseThis timeline describes the concepts of operations for the major architecture systems in a vertical flowcharts (Figure 8).The solar radiation collector mounted on the orbiter satellite will point automatically to the sun tocollect and store the solar radiation energy. During fly-over, the beaming device will track the receivingstation positioned at the lunar surface for example. If tracking is successful the collected solar energywill be beamed to the receiving station during line-of-sight period.The receiving antenna of the receiving station will track the orbiting satellite during fly-by and receivethe transmitted energy. This energy will be reconverted and stored in the receiving station energystorage subsystem. Acquired scientific data will be communicated to Earth. - 13 -American Institute of Aeronautics and Astronautics 14. Figure 8: SOPOSADES System Operational Timeline for the Operational Phase5. Lunar outpost and lander receiver A lunar outpost requires support shipments with needed supplies from Earth. To decrease thesupply shipment rate the outpost shall have the ability to use resources that are available at the lunarsurface. This live off the land also known as In-Situ Resource Utilization (ISRU) will increase theindependency of the lunar outpost from expensive and risky supply. This ability demonstration will havea significant benefit for further human space exploration missions to Mars and other celestial bodies.This design study considers a Lunar Outpost In-Situ Resource Utilization as a system that requires energysupply. Alternatively a lunar rover system could also be considered as an energy consumer for thedevelopment of SOPOSADES.- 14 - American Institute of Aeronautics and Astronautics 15. IV.ConclusionThis paper is about research and development of a satellite solar power energy supplyinfrastructure to a future moon or Mars outpost. This energy supply infrastructure is proposed tobe implemented before astronauts land on the surface of the Moon or Mars to set up an outpost.The moon has no atmosphere that absorbs radiated energy; the Martian atmosphere will havesome energy losses. With the development of such a lunar orbiting solar power satellite system,technology may be developed to a maturity where Earth applications of a SSPS would bewithout significant risk to Earthlings and would be available if an unexpected short term needrequirement would develop. AcknowledgementThe research and study is based on student independent studies by four students: J. Bergman, L.Haumann, M. Kluck, T. Stolze. References[1] A Renewed Spirit of Discovery, The Presidents Vision of Space Exploration, President George W.Bush, January 2004[2] NASA Authorization Act, PUBLIC LAW 109155, 109th Congress, December 30, 2005[3] Power Requirements for the First Lunar Outpost(FLO), NASA/TM-105925, R. L. Cataldo, J. M.Bozek, Lewis Research Center, Cleveland, OH, 1993[4] Glaser, Peter E. (1968). "Power from the Sun: Its Future" Science Magazine 162 (3856): 857861.[5] Glaser, Peter E. (December 25, 1973). "Method And Apparatus For Converting Solar Radiation ToElectrical Power". United States Patent 3,781,647.[6] Glaser, P. E., Maynard, O. E., Mackovciak, J., and Ralph, E. L, Arthur D. Little, Inc., "Feasibilitystudy of a satellite solar power station", NASA CR-2357, NTIS N74-17784, February 1974[7] Committee for the Assessment of NASAs Space Solar Power Investment Strategy, Aeronautics andSpace Engineering Board, National Research Council, Laying the Foundation for Space Solar Power: AnAssessment of NASAs Space Solar Power Investment Strategy, The National Academy Press, 2001.15American Institute of Aeronautics and Astronautics 16. [8] Lunar Frozen Orbits,AIAA-2006-6749-391, NASA Goddard Space Flight Center, D. Folta and D.Quinn, 2006[9] Technology Projections for Solar Dynamic Power, NASA/TM1999-208851, Lee S. Mason, LewisResearch Center, Cleveland, Ohio[10] ISUS solar concentrator array development, AIAA 96-3045, Borell, Campell, 1996[11] Splined Radial Solar Concentrator, US-Patent 5,104,211; Schumacher et al., Harris Corp., 14 Apr.1992[12] Inflatable Structures in Aerospace Engineering - An Overview, European Conference on SpacecraftStructures, European Space Agency, ESASP-468, 2001., p.93, Veldman, 2001[13] Refractive Secondary Concentrators for Solar Thermal Application, NASA/TM1999-209379,Wayne A. Wong, 1999[14] Solar Thermal Propulsion Augmented with Fiber Optics: - A System Design Proposal, AIAA 2005-3922, Paul R. Henshall, University of Surrey, UK, 2005[15] Early Results From Solar Dynamic Space Power System Testing, NASA/TM-1996-107252, RichardK. Shaltens and Lee S. Mason, July 1996[16] An Advanced TurboBrayton Converter for Radioisotope Power Systems, NASA/TM2007-214976, Space Technology and Applications International Forum--STAIF 2005. Vol. AIP ConferenceProceedings 746, pp. 632-640. 2005, Zagarola[17] Development of Advanced Stirling Radioisotope Generator for Space Exploration, NASA/TM-2007-214806, Jack Chan, 2007[18] Overview of energy storage technologies for space applications, http://trs-new.jpl.nasa.gov/dspace/handle/2014/38329 Pasadena, CA : Jet Propulsion Laboratory, NationalAeronautics and Space Administration, Surampudi et al., January 2006[19] The Satellite Solar Power Station, Microwave Symposium Digest, G-MTT International,Volume: 73,Issue: 1, pp. 186- 188, Glaser, P.E., Jun 1973[20] Space solar power programs and microwave wireless power transmission technology, MicrowaveMagazine, IEEE, Volume: 3, Issue: 4, pp. 46- 57, McSpadden,et al., December 2002[21] http://www.mda.mil/system/altb.html[22] US Government, Diode laser satellite systems for beamed power transmission, National Aeronauticsand Space Administration, Office of Management, Scientific and Technical Information Division, ISBN-10: 1234334755 16 American Institute of Aeronautics and Astronautics