Design of a Gigawatt Space Solar Power Satellite Using Optical Concentrator System

Transcription

1 Design of a Gigawatt Space Solar Power Satellite Using Optical Concentrator System Brendan Dessanti, Narayanan Komerath, Shaan Shah Daniel Guggenheim School of Aerospace Engineering Georgia Institute of Technology Atlanta, GA [email protected] Abstract A 1-gigawatt space solar power satellite using a large array of individually pointable optical elements is identified as the key mass element of a large scale space solar power architecture using the Space Power Grid concept. The proposed satellite design enables a significant increase in specific power. Placed in sun-synchronous dynamic orbits near 2000km altitude, these satellites can maintain the constant solar view requirement of GEO-based architectures, while greatly reducing the beaming distance required, decreasing the required antenna size and in turn the overall system mass. The satellite uses an array of individually pointable optical elements (which we call a Mirasol Concentrator Array) to concentrate solar energy to an intensified feed target that feeds into the main heater of the spacecraft, similar conceptually to heliostat arrays. The spacecraft then utilizes Brayton cycle conversion to take advantage of non-linear power level scaling in order to generate high specific power values. Using phase array antennas, the power is then beamed at a millimeter wave frequency of 220GHz down to Earth. The design of the Mirasol concentrator system will be described and a detailed mass estimation of the system is developed. The technical challenges of pointing the elements and maintaining constant solar view is investigated. An end-to-end efficiency analysis is performed. Subsystem designs for the spacecraft are outlined. A detailed mass budget is refined to reflect reductions in uncertainty of the spacecraft mass, particularly in the Mirasol system. One of the key mass drivers of the spacecraft is the active thermal control system. The design of a lightweight thermal control system utilizing graphene sheets is also detailed. TABLE OF CONTENTS 1 INTRODUCTION OPTICAL CONCENTRATOR DESIGN OPTICAL DESIGN CONSIDERATIONS BRAYTON CYCLE TURBOMACHINE DESIGN END-TO-END EFFICIENCY ANALYSIS COMPARISON WITH SIMILAR PROPOSED CON- CEPTS SUMMARY AND CONCLUSIONS ACKNOWLEDGMENTS REFERENCES BIOGRAPHY INTRODUCTION A 1-gigawatt space solar power satellite architecture using Brayton cycle conversion was proposed by our team at last year s IEEE Aerospace conference [1]. The architecture had /12/$26.00 c 2013 IEEE. 1 IEEEAC Paper #2167, Version 5, Updated 12/11/2013 two seperate aspects that we call the Girasol and the Mirasol. One part consists of the conversion of solar energy using an efficient, high temperature Brayton cycle. This part of the spacecraft was detailed in last year s paper, which we refer to as the Girasol, while a detailed design of the Mirasol collector was left for future work. The Mirasol consists of a large array of individually pointable optical elements (with an area of roughly 1.5km 2 ) that serve as reflectors/concentrators of solar energy to the Brayton cycle heater. This paper focuses primarily on detailing the optical design of the Mirasol part of the spacecraft and secondarily on refining the cooling system of the Girasol part, as that is a key mass driver of the spacecraft. Space-Based Solar Power Background Space-based solar power (SSP) is the idea of collecting solar power in space and beaming the power to rectennas on the Earth s surface using RF waves. It was first developed in the 1960s by Peter Glaser [2] of the Arthur D. Little company. Glaser s architecture utilized geosynchronous earth orbit (GEO) for the location of the solar power satellite. His architecture utilized microwave power beaming for transmission and photovoltaic devices for energy conversion. Since then, a number of studies have been performed by NASA and the DOE to evaluate the technical viability and economic feasibility of space-based solar power concepts including the NASA Fresh Look study [3] and the Space Solar Power Exploratory Research and Technology Program (SERT) [4]. These studies found no technical showstoppers preventing space solar power from becoming a reality. However, the large initial investments required coupled with the significant technological risks posed by establishing a large space-based power infrastructure have prevented space solar power programs from getting started. Recently, there has been renewed interest in space-based solar power concepts around the globe, as several different countries have shown interest in exploring SSP concepts. JAXA (Japan Aerospace Exploration Agency) has perhaps shown the greatest interest of the space agencies in moving towards space solar power. JAXA has plans for a Low Earth Orbit demonstration of wireless power beaming by 2015 [5] and plans to establish a functioning solar power satellite by JAXA and the Japanese Ministry of Economy, Trade, and Industry (METI) have committed $21 billion towards the development of a 1GW solar power satellite. China also has plans to complete a space solar power demonstration by 2025 [6]. India has also expressed interest in developing space solar power as a potential strategic partnership with the United States with the NSS-Kalam initiative [7]. In the United States, the NASA Institute of Advanced Concepts has funded a study into a solar power satellite concept proposed by John Mankins [8]. 1

2 SSP architectures show a gap of orders of magnitude gap in economic viability. At the root of the problem is the sheer magnitude of the mass needed for a power plant in space. Launch cost dictates high system costs. The large beaming distance of geosynchronous orbit and low frequency of microwave based systems, coupled with the low efficiency of photovoltaic conversion have made it difficult to present a viable business plan for SSP. As a result of the orbit location and frequency selection of traditional approaches, the transmitter and receiver sizes are pushed to diameters on the order of kilometers, making it difficult to establish an evolutionary approach to jump-starting SSP. Large geosynchronous orbit based architectures show no evolutionary path towards space solar power: until the satellite and ground station are built to full size, no power can be transacted. Innovative solutions are needed to the space solar power problem, to offer significant improvements in economic viability and allow for an evolutionary approach, in order to generate interest among government and/or private entities. Space Power Grid The Space Power Grid (SPG) systems architecture is an evolutionary approach to large-scale space-based solar power. The Experimental Aerodynamics and Concepts Group at Georgia Tech has been working on the development of this architecture since The Space Power Grid architecture argues for three key technical departures from traditional space-based solar power approaches. The first is to move to millimeter wave beaming at a frequency of 220GHz, as opposed to traditional microwave frequency choices at 2.45GHz or 5.8GHz. By moving to millimeter wave frequencies, the transmitter and receiver diameter sizes required come down to much more manageable levels, scaling down the size of the spacecraft, and in turn the mass that must be placed into orbit, by a large amount. Traditional approaches have favored microwave transmission because microwave antenna and conversion devices have greater technological maturity and because frequencies over 10GHz experience poor transmission through rain and fog. Our architecture trades the technology risk of millimeter wave technology, which has been seeing widespread advances in recent years, in favor of the greatly reduced system sizes. Potential options for transmission through rain and fog using millimeter waves include burn-through techniques, beaming around areas of high precipitation, and to consider using tethered aerostats placed above the weather generating portion of the atmosphere to receive the beamed power and then transmit it to ground via waveguides built into the tethers. The second key departure is the use of Brayton cycle solar dynamic conversion of solar power, rather than photovoltaic arrays. The explanation for this trade is further detailed in the Brayton cycle turbomachine design section of this paper. In order to reach the high power and intensities required for the high efficiencies of the Brayton cycle, an optical concentrator system, the mirasol, is required. The design of the Mirasol is discussed in the paper. The third key architecture change is to transmit from dynamic orbits around 2000km rather than from satellites placed in geosynchronous orbit. This also greatly reduces the antenna sizes required, bringing the spacecraft to much more manageable sizes. Lastly, the Space Power Grid proposes a first phase that does not include space-based power generation. The initial phase would consist of a constellation of waveguide relay satellites that would serve as a power exchange with terrestrial power entities. Relaying power beamed from terrestrial power sites with excess power to high demand areas around the world. This key first step creates an evolutionary approach, with reasonably sized spacecraft (about 4000kg [9]) that can demonstrate and reduce many of the technical risks associated with a large scale space solar power architecture. This initial constellation would be replaced over time by the large gigawatt level spacecraft detailed in this paper. 2. OPTICAL CONCENTRATOR DESIGN Mirasol/Girasol Orbit Selection Options The Mirasols are required to concentrate solar power to the heater of the Brayton cycle heat engine. The Mirasol and Girasol aspects of each spacecraft pair are not physically connected and there are multiple options considered for the orbit placement of the two parts, outlined in this section. 1. Place the Mirasols as high altitude collector/concentrators: In this option, the Mirasols would be placed at an altitude around 20,000km or higher and reflect/concentrate power down to the LEO constellation of Girasols. This option has the advantage of being able to easily reach all parts of the globe and maintain near constant solar view. This option is limited however by the spot size issue of reflecting solar power over such large distances [10]. Thus, the area of the Girasol collector dish would have to be increased substantially for this option over the current design. 2. Place Mirasols and Girasols together in sun-synchronous dynamic orbits: This option has the advantage that the spot size issue is eliminated while constant solar view is maintained, and the collector dish of the Girasol is no longer a necessary component of the spacecraft. Spacecraft placed in sunsynchronous orbits around 2000km can maintain constant solar view by staying along the Earth s day-night terminator as identified by Potter and Davis [10]. The issue with this option is the ability to continuously beam to the dark side of the Earth. 3. Place Mirasols and Girasols in close orbits at 2000km: This solution allows for a compromise between the other two options. Although the Girasol still requires a collector dish, the Mirasols, placed predominantly in sun-synchronous orbits, beam to nearby Girasols in orbits around 2000km altitude. The result is that the spot size is small enough to be at manageable levels, while continuous beamed power delivery can be achieved to the dark side of the Earth. A conceptual drawing of the Girasol part of the spacecraft receiving concentrated solar power from the Mirasol collector/concentrator is shown in Figure??. Configuration The optical array design consists of a large array of around one thousand ultralight optical reflector elements. The configuration shape has not yet been optimized, but it can be conceptualized as some sort of concave shape with the target of the concentrated solar power being the heater of the Brayton cycle heat engine. This design eliminates the need for a large Power Management and Distribution System (PMAD). When utilizing photovoltaic arrays for conversion, the use of refractive optical elements can be advantageous, as fresnel lens optics can be used to increase the concentration of solar energy multiple times. O Neill has extensively researched the possible use of streched fresnel lens refractive concentrator arrays for use in space solar power [11]. O Neill argues that in the long term, the Stretched Lens Array (SLA) proposed can reach 1 kw/kg for MW level arrays. A 2.5kW Stretched 2

3 ADACS for each element, bringing the total ADACS system mass to 28,800kg. Figure 1. Conceptual Drawing of Girasol part of the system receiving concentrated solar power from the Mirasol part of the system Fresnel Lens array called SCARLET on the Deep Space One Probe demonstrated 45 W/kg specific power [11]. The SLA uses a 140 micron thick fresnel lens array made from a silicone rubber material. Certainly, the use of fresnel lens optics makes sense when using a photovoltaic array for a space-based solar power satellite. For a satellite using solar thermal conversion, such as the one described in this paper, the goal is to maximize the concentration of solar power at minimum mass while incurring the smallest losses. Ultralight reflective optical elements were chosen to achieve very high efficiencies at low mass and as a result were selected as the best conceptual design choice. Material The material used must be ultrathin and lightweight. The most likely optical materials to meet the design requirements would be some form of polyimide film. A good starting point for selecting a material is to look at materials used for solar sail applications, as solar sail materials are designed to be ultralightweight and highly reflective. Some solar sail demonstration spacecraft have already been deployed. JAXA has successfully deployed the IKAROS spacecraft which uses a polyimide sail. The United States has successfully deployed the Nanosail-D2 spacecraft. This solar sail is 10 square meters, and the spacecraft weighs only 4 kg. It uses an ultralight thin film material called LaRC-CP1 polyimide made by ManTech [12] and developed at NASA Langley, on the order of a few microns in thickness [13]. Solar sail material densities are on the order of 0.01 kg/m 2 [14]. The solar sail material mass was calculated by simply taking the specific mass of solar sail and multiplying by the area required determined from the power required from the endto-end efficiency analysis. ADACS Subsystem An attitude determination and control system (ADACS) is required with actuators and sensors to maintain the necessary sun tracking and pointing requirements. The ADACS includes an actuator and sensor system for each optical element. The mass estimate was determined by allotting 30kg for the Mirasol Thrusters and Solar Radiation Pressure The propellant mass requirement becomes the limiting factor on Mirasol specific mass (given that solar sail material can be so light), based on the deltav requirement from solar radiation pressure. Given the use of ultralight thin film mirrors, the propellant required for solar radiation pressure becomes a significant mass aspect of the Mirasol subsystem. The large area per unit mass value implies that solar radiation pressure force is substantially high to accelerate the spacecraft out of orbit, which needs to be balanced by some propulsive force. As the propellant mass is increased, the acceleration due to solar radiation pressure force itself drops as the area to mass ratio decreases. In other, more massive architectures using PV arrays, this is still a considerable issue, just not a large mass component, which explains why solar radiation pressure considerations are sometimes neglected in space solar power studies. The Aerospace Corporation SERT report [4] does not ignore it however, and it is considerable for their Halo concept evaluation. Another thing to keep in mind is that the solar radiation pressure acceleration will actually increase as the propellant mass is used up, but the structural and attitude control system mass estimates used in this analysis should be conservative enough to account for this. The Delta-V required due to the solar radiation pressure force is calculated by equation 1. V = S C r T M/A S=Solar Flux at 1 AU from Sun C r =surface reflection index T =time interval of thrusting M/A = mass per unit area, specific mass of the Mirasol From this equation it is clear that as the specific mass increases, less Delta-V will be required. The deltav requirements due to solar radiation pressure are plotted in Figure 2. Note that the plot starts at 0.01 kg/m 2, as there is no way to get a lower mass than this because the solar sail material itself has this specific mass value. The plot shows the exponential decrease in Delta-V required with increasing specific mass. The Mirasol array would use krypton thrusters which can achieve a specific impulse of 5300 seconds. Using the rocket equation, the propellant mass required for the mission was calculated over the 17 year spacecraft lifetime. The results are plotted in Figure 3. Mirasol Subsystem Mass Summary With the optical material system mass, propellant mass, and ADACS subsystem required mass estimates, the last remaining item is to account for the structural support system for the Mirasol concentrator. Roughly 15,000kg has been allotted for the structural support system, an equivalent amount of structural mass as the solar sail material mass was estimated for the support required to maintain tension of the solar sail material for each optical element. The optimized Mirasol system mass was calculated by plot- (1) 3

4 Figure 2. Delta-V requirement for different specific mass values for the Mirasol Array Figure 4. Mirasol System Mass for Varying Specific Power Values Table 2. Mirasol Concentrator Array Design Parameters Parameter Value Units Power from Sun Per Area in Space 1.36 kw/m 2 Power Required 2.09 GW Area Required km 2 Number of Optical Elements 960 Area Per Element 1600 m 2 Mirasol Specific Mass 0.06 kg/m 2 Mirasol System Mass 92,000 kg Figure 3. Propellant Mass Requirement for Varying Specific Mass Values ting the Mirasol mass required as the constraint, and then plotting the Mirasol mass available by the specific power value. Finding the intersection of these two lines gives the minimum specific mass value that meets the mass requirements for each subsystem. As seen in Figure??, the design point is at a specific mass value of 0.06 kg/m 2, with a total Mirasol system mass coming to about 92,000kg. Mirasol Design Summary A table summarizing parameters for the Mirasol concentrator array is shown in Table??. Using the power required to be concentrated value calculated from the end to end efficiency analysis shown later, the area required could be determined from the solar flux at 1 AU. The number of optical elements Table 1. Mirasol Subsystem Mass Summary Parameter Solar Sail Material Mass Required Mirasol Structural Mass Required ADACS Mass Required Mirasol Propellant Mass Mirasol System Mass Mass kg kg kg kg kg required was a parameter that was minimized in an effort to reduce the ADACS mass requirement to the extent possible. The total Mirasol concentrator system mass of 92,000 kg was determined from the analysis shown earlier in this section. 3. OPTICAL DESIGN CONSIDERATIONS Optical Tracking Since the required tracking is continuous and predictable, the solar tracking can be accomplished using a mechanical attitude determination and control system with sensors and actuators. Mirror Degradation Longterm degradation effects from Near Ultraviolet (NUV) and Vacuum Ultraviolet (VUV) radiation could result in a small decrease in performance over the lifetime of the array. Radiation from the space environment can cause darkening of the optical material, altering the solar absorptance characteristics of the material [15]. However, degradation is not anticipated to significantly decrease the performance of the satellite. Some level of robotic maintenance is required for all space-based solar power architectures and optical coatings could be replaced periodically through on-orbit servicing. This is easier to do at the low earth orbit constellation proposed by the Space Power Grid as opposed to GEO based architectures. Micrometeroid Impacts Micrometeroid impact damage will occasionally puncture holes in the array. The modular nature of the design means that overall performance of the concentrator will only be 4

5 slightly decreased. These impacts can also be repaired over time by on-orbit servicing of the spacecraft. Since the material to be used is ultralight, micrometeroid impacts should merely result in small holes that don t result in cracks or rips that propagate. In [14] testing on hypervelocity impacts for solar sails did not show rip propagation. Error Tolerance Imperfections in the optical elements lead to shape errors that result in an uneven flux reflected by the optical element. This uneven flux should not pose an issue for the gigawatt satellite proposed because the energy is being directed towards the collector of a heat engine in order to heat the working fluid rather than photovoltaics. 4. BRAYTON CYCLE TURBOMACHINE DESIGN The design of the Brayton cycle turbomachine used for the spacecraft, that we call the Girasol portion of the satellite, is detailed in [16]. In that paper, subsystem components and a cycle analysis for a spacecraft using solar thermal Brayton cycle conversion were shown. High Brayton cycle efficiencies can be achieved using ultra-high temperature ceramic materials such as hafnium carbide. This paper refines the Girasol design by refining the thermal control system design for the Girasol. Brayton cycle conversion was chosen over photovoltaic conversion because of the scaling advantages in specific power of Brayton cycle conversion over photovoltaic at high power levels. At high power levels, photovoltaic arrays do not exhibit favorable scaling. Since photovoltaic arrays are a surface-area based phenomenon at their root, an increase in area produces a proportional increase in power. Thus an improvement in specific power (power delivered per unit mass) is not seen in photovoltaic systems at higher power rating levels. Brayton cycle engines do exhibit favorable scaling. As was shown in [16], a study of historical jet engine data exhibited this trend of increasing specific power with increasing power levels. NASA Glenn has researched the use of Brayton cycle converters for space power applications since the Brayton Rotating Unit development program dating back to the 1960s. Currently, NASA Glenn has the 2kW Brayton Power Conversion Unit (BPCU) which has been tested and shown to work with nuclear fission heat sources [17]. The BPCU does not exhibit very high specific power because of its low power levels. In other words, it takes a lot more mass for a Brayton cycle heat engine to produce its first watt than to produce its ten millionth for example. Mason argues for increasing specific power that can be achieved at high power levels in [18], offering favorable predictions for Brayton cycle engine performance capabilities in the mid-far term. Thermal Control System Design In order to design a highly efficient lightweight cooling system several options were considered. Among those options, the exceptionally high thermal conductivity of Graphene and Carbon Nanotubes, along with their low mass densities, made these options ideal candidates for the radiator material of the cooling system. Graphene consists of a single plane of sp2 bonded carbon-carbon atoms and as a result can be considered to be a two-dimensional material. The fact that it is two-dimensional makes Graphene have somewhat unusual Table 3. Comparison of thermal conductivity and areal density of Graphene, Carbon Nanotubes and Graphite Thermal Material Conductivity Areal Density (W/mK) Graphene Sheet e-7 kg/m 2 Carbon Nanotubes e-5 kg/m 2 Pyrolitic Graphite kg/m 3 properties; two of them are that it has an extremely high thermal conductivity and is incredibly lightweight. Table 3 compares Graphene with other materials that have comparatively high thermal conductivity. Using a radiator design similar to the one described by Juhasz in [19], and with a high enough emissivity, a thermal management system of below 1 kg/m 2 can be easily achieved with the use of Graphene because of its incredibly low mass density. Apart from this, new and developing heat pipe technology similar to one described in [20] hold a strong potential for lightweight and highly efficient heat management systems in space. Although there is no definite value of thermal conductivity of Graphene that is agreed upon as of yet, the measured values fall between 3000 W/mK to 5300 W/mK for different methods and different conditions [21]. Even as we increase the number of layers, the conductivity decreases approaching that of graphite, but still remains significantly higher than most other materials. Even after considering various effects like losses due to wrinkling of the material, etc., the thermal conductivity of Graphene remains extremely high. Juhasz proposed a radiator that could achieve 1kg/m 2 in [22]. We conservatively use this value as a specific mass estimate for our thermal control system, noting that with the very low mass density of graphene sheets and the very large scale of the system, it is likely that we could reduce the specific mass considerably lower than this value. The radiator area required was determined from the radiative power equation (shown in equation 2). The system mass was then estimated using the specific mass value. P = ɛσat 4 (2) P = Power Required to be Radiated from Girasol ɛ = Emissivity σ = Stefan-Boltzmann Constant T = Equilibrium Radiator Operating Temperature A = Radiator Area Required A table summarizing the parameters of the Thermal Control System is shown in Table 4. Note that the emissivity value is higher than the emissivity of graphene. This is an estimate of the emissivity value of the graphene sheets after the application of coating and paint has been applied. The current design requires radiating a very large amount of power away from the spacecraft (400MW). As a result, a trade study is being performed to evaluate incorporating a recuperator into the Brayton cycle turbomachine design, to increase the efficiency of the Brayton cycle to see if the power required for radiation can be decreased, in order to decrease the size and mass of the thermal control system. 5

6 heliostat array elements produced by Practical Solar can achieve 99.5% reflectivity [23]. The 80% Brayton cycle efficiency value comes from the cycle analysis performed in [16], using hafnium carbide material for the turbomachine components. One area of uncertainty in the efficiency analysis is in the ability of the millimeter wave generator to efficiently convert the power generated from the Brayton cycle to 220GHz for beaming. It remains to be shown that high power millimeter wave conversion technology can achieve these efficiency values and this is one area of uncertainty for the spacecraft. These generators would have to be compact, lightweight, and efficient. Barnes has proposed compact, lightweight superconducting generators for use in military power packs [24] that were scaled to size the generator for the Girasol. Figure 5. Thermal Control System Mass Required for Varying Equilibrium Operating Temperatures Table 4. Thermal Control System Parameters Parameter Value Units Power Required To Be Radiated 400 MW Operating Temperature 662 K σ 5.67e-8 W/(m 2 K 4 ) ɛ 0.23 Area Required 160,000 m 2 TCS Specific Mass 1 kg/m 2 TCS System Mass 160,000 kg 5. END-TO-END EFFICIENCY ANALYSIS An end-to-end efficiency analysis was calculated to account for various losses present in the system. Sizing the spacecraft to deliver 1GWe to the terrestrial grid, the end-to-end efficiency allows us to determine the power required to be concentrated by the Mirasol array. A summary of efficiency values is presented in Table 5. Current commercially available ultralight optical reflector elements can achieve 92% efficiency using Mylar, Teonex, or LaRC-CP1 polyimide materials [14]. It is anticipated that with research and development this value would be able to be increased to around 99%. Current commercially available Table 5. Efficiency Breakdown Parameter Value Units Mirasol Power Required 2.09 GW Mirasol Array Efficiency 0.99 Brayton Cycle Efficiency 0.80 Millimeter Wave Generator Efficiency 0.90 Power After Brayton Conversion 1.49 GW Waveguide System Efficiency 0.97 Phased Array Antenna Efficiency 0.90 Power to Beam 1.30 GW Atmospheric Transmission Efficiency 0.90 Beam Capture 0.95 Rectenna Efficiency (RF to DC) 0.90 Power to Ground Required 1.00 GWe End-to End Efficiency 0.48 Efficiency values for power transmission components were investigated and detailed in [9]. A waveguide system is used to transmit power to the phased array antennas at very high efficiencies. As a result of the orders of magnitude decrease in the required antenna transmitter and receiver diameters by moving to LEO and using millimeter wave transmission, it was shown in [9] that several lobes of the Airy diffraction pattern could now be captured without reaching excessive antenna sizes. Thus, beam capture efficiency could be increased from 84% power captured by the first ring as in geosynchronous arhitectures to around 95% by capturing the first four to five lobes. For microwave frequencies, rectenna efficiency for RF to DC conversion exceeds 90%. According to Koert and Cha, as early as 1992, 70% efficient rectennas could be developed for millimeter wave frequencies with a hybrid rectenna circuit using high power schottky diodes [25]. The reason high power millimeter wave rectennas are more difficult to make than microwave rectennas is because the size of the diodes needs to be smaller in order to avoid impedance issues at higher frequencies. With research and development into high power millimeter wave conversion devices and current developments in Microelectrical Mechanical System (MEMS) manufacturing techniques, there does not appear to be anything preventing millimeter wave rectennas from surpassing over 90% efficiency. An end to end efficiency value of 0.48 was calculated for the Girasol/Mirasol 1 gigawatt spacecraft. The end-to-end efficiency value here is defined as the amount of power delivered to the ground per the amount of power incident on the Mirasol array. For the spacecraft to deliver a gigawatt to the electric grid, this allows us to size the Mirasol array by determining a Mirasol power required value. 6. COMPARISON WITH SIMILAR PROPOSED CONCEPTS Halo Concept As part of the NASA Space Solar Power Exploratory Research and Technology program (SERT), The Aerospace Corporation looked at the technology available for space solar power and evaluated a number of different space solar power architectures [4]. One of the proposed architectures, the Halo concept, was developed by Henry Harris at the Jet Propulsion Laboratory [4]. This concept uses a constellation of reflectors to reflect towards a photovoltaic array. This architecture is similar to the Mirasol design in the sense that, like a heliostat, a collection of optical reflectors is used to concentrate solar 6

7 Table 6. Comparison with Halo and SPS-Alpha Concepts Mirasol Halo SPS-Alpha Item Concept Concept Concept Optical 960 optical 100 orbiting 5000 Elements elements reflectors reflectors Energy Solar PV PV Conversion Thermal Array Array Power Mill Wave Solid State 200,000 Generation Generator Amplifiers Modules Beam 220GHz 5.8 GHz 2.45 GHz Frequency (Mill Wave) (Microwave) (Microwave) Power Level 1 GW 1.2GW 2 GW System Mass 677MT 13,044MT 9,350MT Table 7. Mass budget for 1 GWe Space Solar Power Satellite Mass % of % of Element (kg) Girasol Satellite Collector 30, Cooling System 160, Brayton Cycle 20, AC generator 50, Cryogenics 20, GHz Amp 17, Antennae 20, Propulsion 170, Misc. 30, Structure 58, Total Girasol 584, Total Mirasol 92, Total Mass 676, energy. Some key differences exist between the Halo concept and the Mirasol concept. A comparison of the two concepts is summarized in Table 6. Of note, of the four microwave based space solar power architectures studied by The Aerospace Corporation, the Halo concept performed considerably better in terms of the overall system level metrics of maximum specific power and minimum cost per kilowatt hour than the other architectures evaluated. The system mass for the Mirasol/Girasol satellite is significantly less than the estimate for the Halo concept given several factors. One key reason is the very large reduction in antenna size. The area of the antenna becomes orders of magnitude less when moving to 220GHz beaming and bringing the orbit down to LEO, reducing the beaming distance from 36,000km to 3,000km (2000km altitude, plus accounting for 45 degree coning angle). Assuming that antenna area scales with mass, the antenna mass of the system becomes orders of magnitude less. The power conversion system mass is the other key mass driver that comes down considerably, since Brayton cycle conversion exhibits scaling advantages at higher power levels over photovoltaic. Another important note on mass reduction is that given the much higher efficiency of Brayton cycle conversion over photovoltaic systems, the heat that must be radiated away from the spacecraft is reduced considerably. This greatly reduces the required thermal control system mass, another key mass driver of the system. SPS-Alpha The NASA Institute of Advanced Concepts is funding a study into the conceptual design of SPS-Alpha, a space solar power satellite concept proposed by John Mankins[8]. SPS-Alpha is similar to the Mirasol/Girasol satellite proposed here in that they both utilize a large array of individually pointable thin film mirrors to redirect sunlight. The key differences are that SPS-Alpha would be in geosynchronous orbit, would utilize photovoltaic conversion rather than Brayton cycle conversion, and microwave beaming (2.45 GHz) rather than millimeter wave beaming. The Mirasol/Girasol again shows a significant mass reduction, due primarily to similar reasons as outlined above for the Halo concept. 7. SUMMARY AND CONCLUSIONS Satellite Mass Summary A mass summary for the different subsystems is shown in Table 7. This system mass summary reflects refined estimates of the spacecraft cooling system and the Mirasol optical system mass calculations, as these systems have been investigated in further detail. Conclusions The design of an optical concentrator system for a gigawattlevel space solar power satellite was proposed. A large array of individually pointable ultralight, thin film optical elements can redirect and concentrate sunlight towards the heater of a Brayton cycle solar dynamic heat engine for conversion of solar power for space-based solar power applications. A cooling system using graphene sheets offers the potential for efficient, lightweight radiation of heat for the converter part of the spacecraft system. The design is shown to close using conservative mass and efficiency estimations. The spacecraft offers potential for large improvements in specific power, the amount of electric power that can be delivered to ground per unit mass required to be launched into orbit, over traditional space solar power architectures ( 1.5kW/kg vs. 0.2kW/kg). ACKNOWLEDGMENTS This work is partially supported as an advanced concept development study under the EXTROVERT cross-disciplinary innovation initiative by NASA. Mr. Tony Springer is the technical monitor. REFERENCES [1] N. Komerath, B. Dessanti, and S. Shah, A gigawattlevel solar power satellite using intensified efficient conversion architecture, in Proceedings of the IEEE Aerospace Conference,, ser. Paper 1548, no. DOI /AERO Big Sky, Montana: IEEE, March [2] P. Glaser, Power from the sun: It s future, Science, vol. 162, pp , [3] J. Mankins, A fresh look at space solar power: New architectures, concepts and technologies, Acta Astro- 7

8 nautica, vol. 41, no. 4, pp , [4] J. P. Penn and G. W. Law, The aerospace corporation systems studies and analysis of the space solar power (ssp) exploratory research and technologies (sert) concepts and applications, The Aerospace Corporation, Tech. Rep. ATR-01(7710)-1, [5] M. Mori, H. Kagawa, and Y. Saito, Summary of studies on space solar power systems of japan aerospace exploration agency (jaxa), Acta Astronautica, vol. 59, no. 1, pp , [6] D. Flournoy, How is sunsat development faring internationally? Solar Power Satellites, pp , [7] R. Gopalaswami, Kalam-national space society energy technology universal initiative: An international preliminary feasibility study on space based solar power stations, October [8] J. Mankins, N. Kaya, and M. Vasile, Sps-alpha: The first practical solar power satellite via arbitrarily large phased array (a nasa niac project), in 10th International Energy Conversion Engineering Conference, [9] B. Dessanti, R. Zappulla, N. Picon, and N. Komerath, Design of a millimeter waveguide satellite for space power grid, in Proceedings of the IEEE Aerospace Conference, ser. Paper 1549, no. DOI /AERO Big Sky, Montana: IEEE, March [10] S. Potter and D. Davis, Orbital reflectors for space solar power demonstration and use in the near term, in AIAA SPACE 2009 Conference and Exposition. California: AIAA, September [11] M. O Neill, M. McDanal, M. Piszczor, D. Edwards, M. Eskenazi, and H. Brandhorst, Recent technology advances for the stretched lens array (sla), a space solar array offering state of the art performance at low cost and ultra-light mass, in Photovoltaic Specialists Conference, Conference Record of the Thirty-first IEEE, jan. 2005, pp [12] Thin film... large payoff. [Online]. Available: images/documents/ doc.pdf [13] L. Johnson, R. Young, E. Montgomery, and D. Alhorn, Status of solar sail technology within nasa, Advances in Space Research, vol. 48, no. 11, pp , [14] D. L. Edwards, C. Semmel, M. Hovater, M. Nehls, P. Gray, W. Hubbs, and G. Wertz, Status of Solar Sail Material Characterization at NASA S Marshall Space Flight Center, Sep. 2006, p [15] J. B. Heaney, L. R. Kauder, S. E. Bradley, and D. E. Neuberger, Mirror degradation in orbit due to space radiation exposure, pp , [Online]. Available: + [16] N. Komerath and B. Dessanti, Brayton cycle conversion for space solar power, in Proceedings of the 48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, no. DOI: / Atlanta, Georgia: AIAA, August [17] L. S. Mason, R. K. Shaltens, J. L. Dolce, and R. L. Cataldo, Status of Brayton cycle power conversion development at NASA GRC, in Space Technology and Applications International Forum, ser. American Institute of Physics Conference Series, M. S. El-Genk, Ed., vol. 608, Jan. 2002, pp [18] L. S. Mason, A comparison of Brayton and Stirling space nuclear power systems for power levels from 1 kilowatt to 10 megawatts, in American Institute of Physics Conference Series, ser. American Institute of Physics Conference Series, M. S. El-Genk, Ed., vol. 552, Feb. 2001, pp [19] A. Juhasz, Design considerations for lightweight space radiators based on fabrication and test experience with a carboncarbon composite prototype heat pipe, NASA, Tech. Rep. NASA/TP /Rev1, September [20] Y. Qu et al., Superconducting heat transfer medium, Jul , us Patent 6,916,430. [21] W. Choi and J. Lee, Graphene: Synthesis and Applications. CRC Press, 2011, vol. 3. [22] A. Juhasz, High conductivity carbon-carbon heat pipes for light weight space power system radiators, in 6th International Energy Conversion Engineering Conference. AIAA, July [23] B. Rohr, The promise of small heliostats, Northeast Sun, vol. Spring, [24] P. Barnes, G. Rhoads, J. Tolliver, M. Sumption, and K. Schmaeman, Compact, lightweight, superconducting power generators, Magnetics, IEEE Transactions on, vol. 41, no. 1, pp , [25] P. Koert and J. Cha, Millimeter wave technology for space power beaming, Microwave Theory and Techniques, IEEE Transactions on, vol. 40, no. 6, pp , jun BIOGRAPHY[ Brendan Dessanti received his B.S. degree in Aerospace Engineering from the Georgia Institute of Technology in 2011 and is currently pursuing an M.S. Degree in Aerospace Engineering from Georgia Tech. Brendan is a graduate research assistant for the Experimental Aerodynamics and Concepts Group at Georgia Tech and leader of the Space Power Grid student team. His research focuses on space systems design as it applies to spacebased solar power. Brendan has obtained experience from internships at Sikorsky Aircraft, MIT Lincoln Laboratory and SpaceWorks Enterprises, Inc. 8

9 Narayanan Komerath is a professor in the Daniel Guggenheim school of aerospace engineering at Georgia Institute of Technology (G.I.T.), and director of the John J. Harper Wind Tunnel and the Experimental Aerodynamics and Concepts Group. He has served as a NIAC Fellow, and as a Boeing Welliver Summer Faculty Fellow. Shaan Shah is a senior at Georgia Tech, pursuing his B.S. in Aerospace Engineering. He has participated in research as a member of the Experimental Aerodynamics and Concepts Group, mainly involved with the development of a conceptual design for the Space Power Grid. His research focuses on studying breakthroughs in materials like Graphene and Ultra High Temperature Ceramics, and technological advances related to space-based solar systems in the fields of thermoelectrics, thermophotovoltaics and narrow-band photovoltaics. This experience has left him with an interest to study further the possibilities of Space-based Power Systems. He intends to continue graduate studies in the area of propulsion and combustion. 9