Design of a Gigawatt Space Solar Power Satellite UsingOptical Concentrator System

Brendan Dessanti, Narayanan Komerath, Shaan ShahDaniel Guggenheim School of Aerospace Engineering

Georgia Institute of TechnologyAtlanta, GA 30332

404-894-3017bdessanti@gatech.edu

Abstract—A 1-gigawatt space solar power satellite using a largearray of individually pointable optical elements is identified asthe key mass element of a large scale space solar power architec-ture using the Space Power Grid concept. The proposed satellitedesign enables a significant increase in specific power. Placedin sun-synchronous dynamic orbits near 2000km altitude, thesesatellites can maintain the constant solar view requirement ofGEO-based architectures, while greatly reducing the beamingdistance required, decreasing the required antenna size and inturn the overall system mass. The satellite uses an array ofindividually pointable optical elements (which we call a MirasolConcentrator Array) to concentrate solar energy to an intensi-fied feed target that feeds into the main heater of the spacecraft,similar conceptually to heliostat arrays. The spacecraft thenutilizes Brayton cycle conversion to take advantage of non-linearpower level scaling in order to generate high specific powervalues. Using phase array antennas, the power is then beamedat a millimeter wave frequency of 220GHz down to Earth. Thedesign of the Mirasol concentrator system will be described anda detailed mass estimation of the system is developed. Thetechnical challenges of pointing the elements and maintainingconstant solar view is investigated. An end-to-end efficiencyanalysis is performed. Subsystem designs for the spacecraft areoutlined. A detailed mass budget is refined to reflect reductionsin uncertainty of the spacecraft mass, particularly in the Mirasolsystem. One of the key mass drivers of the spacecraft is theactive thermal control system. The design of a lightweightthermal control system utilizing graphene sheets is also detailed.

TABLE OF CONTENTS

1 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 OPTICAL CONCENTRATOR DESIGN . . . . . . . . . . . . . 23 OPTICAL DESIGN CONSIDERATIONS . . . . . . . . . . . 44 BRAYTON CYCLE TURBOMACHINE DESIGN . . . . 55 END-TO-END EFFICIENCY ANALYSIS . . . . . . . . . . . 66 COMPARISON WITH SIMILAR PROPOSED CON-

CEPTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67 SUMMARY AND CONCLUSIONS . . . . . . . . . . . . . . . . . . 7

ACKNOWLEDGMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . 7REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7BIOGRAPHY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

1. INTRODUCTIONA 1-gigawatt space solar power satellite architecture usingBrayton cycle conversion was proposed by our team at lastyear’s IEEE Aerospace conference [1]. The architecture had

978-1-4577-0557-1/12/$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 anefficient, high temperature Brayton cycle. This part of thespacecraft was detailed in last year’s paper, which we referto as the ’Girasol’, while a detailed design of the Mirasolcollector was left for future work. The Mirasol consists of alarge array of individually pointable optical elements (with anarea of roughly 1.5km2) that serve as reflectors/concentratorsof solar energy to the Brayton cycle heater. This paper fo-cuses primarily on detailing the optical design of the Mirasolpart of the spacecraft and secondarily on refining the coolingsystem of the Girasol part, as that is a key mass driver of thespacecraft.

Space-Based Solar Power Background

Space-based solar power (SSP) is the idea of collecting solarpower in space and beaming the power to rectennas on theEarth’s surface using RF waves. It was first developedin the 1960s by Peter Glaser [2] of the Arthur D. Littlecompany. Glaser’s architecture utilized geosynchronous earthorbit (GEO) for the location of the solar power satellite. Hisarchitecture utilized microwave power beaming for transmis-sion and photovoltaic devices for energy conversion. Sincethen, a number of studies have been performed by NASAand the DOE to evaluate the technical viability and economicfeasibility of space-based solar power concepts including theNASA Fresh Look study [3] and the Space Solar PowerExploratory Research and Technology Program (SERT) [4].These studies found no technical showstoppers preventingspace solar power from becoming a reality. However, thelarge initial investments required coupled with the significanttechnological risks posed by establishing a large space-basedpower infrastructure have prevented space solar power pro-grams from getting started.

Recently, there has been renewed interest in space-basedsolar power concepts around the globe, as several differentcountries have shown interest in exploring SSP concepts.JAXA (Japan Aerospace Exploration Agency) has perhapsshown the greatest interest of the space agencies in movingtowards space solar power. JAXA has plans for a Low EarthOrbit demonstration of wireless power beaming by 2015 [5]and plans to establish a functioning solar power satellite by2040. JAXA and the Japanese Ministry of Economy, Trade,and Industry (METI) have committed $21 billion towardsthe development of a 1GW solar power satellite. China alsohas plans to complete a space solar power demonstration by2025 [6]. India has also expressed interest in developingspace solar power as a potential strategic partnership withthe United States with the NSS-Kalam initiative [7]. In theUnited States, the NASA Institute of Advanced Concepts hasfunded a study into a solar power satellite concept proposedby John Mankins [8].

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SSP architectures show a gap of orders of magnitude gapin economic viability. At the root of the problem is thesheer magnitude of the mass needed for a power plant inspace. Launch cost dictates high system costs. The largebeaming distance of geosynchronous orbit and low frequencyof microwave based systems, coupled with the low efficiencyof photovoltaic conversion have made it difficult to present aviable business plan for SSP. As a result of the orbit locationand frequency selection of traditional approaches, the trans-mitter and receiver sizes are pushed to diameters on the orderof kilometers, making it difficult to establish an evolutionaryapproach to jump-starting SSP. Large geosynchronous orbitbased architectures show no evolutionary path towards spacesolar power: until the satellite and ground station are built tofull size, no power can be transacted. Innovative solutionsare needed to the space solar power problem, to offer signif-icant improvements in economic viability and allow for anevolutionary approach, in order to generate interest amonggovernment and/or private entities.

Space Power Grid

The Space Power Grid (SPG) systems architecture is anevolutionary approach to large-scale space-based solar power.The Experimental Aerodynamics and Concepts Group atGeorgia Tech has been working on the development of thisarchitecture since 2006. The Space Power Grid architectureargues for three key technical departures from traditionalspace-based solar power approaches. The first is to move tomillimeter wave beaming at a frequency of 220GHz, as op-posed to traditional microwave frequency choices at 2.45GHzor 5.8GHz. By moving to millimeter wave frequencies, thetransmitter and receiver diameter sizes required come downto much more manageable levels, scaling down the size ofthe spacecraft, and in turn the mass that must be placedinto orbit, by a large amount. Traditional approaches havefavored microwave transmission because microwave antennaand conversion devices have greater technological maturityand because frequencies over 10GHz experience poor trans-mission through rain and fog. Our architecture trades thetechnology risk of millimeter wave technology, which hasbeen seeing widespread advances in recent years, in favorof the greatly reduced system sizes. Potential options fortransmission through rain and fog using millimeter waves in-clude burn-through techniques, beaming around areas of highprecipitation, and to consider using tethered aerostats placedabove the weather generating portion of the atmosphere toreceive the beamed power and then transmit it to ground viawaveguides built into the tethers.

The second key departure is the use of Brayton cycle solardynamic conversion of solar power, rather than photovoltaicarrays. The explanation for this trade is further detailed in theBrayton cycle turbomachine design section of this paper. Inorder to reach the high power and intensities required for thehigh efficiencies of the Brayton cycle, an optical concentratorsystem, the mirasol, is required. The design of the Mirasolis discussed in the paper. The third key architecture changeis to transmit from dynamic orbits around 2000km ratherthan from satellites placed in geosynchronous orbit. Thisalso greatly reduces the antenna sizes required, bringing thespacecraft to much more manageable sizes.

Lastly, the Space Power Grid proposes a first phase that doesnot include space-based power generation. The initial phasewould consist of a constellation of waveguide relay satellitesthat would serve as a power exchange with terrestrial powerentities. Relaying power beamed from terrestrial powersites 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 candemonstrate and reduce many of the technical risks associatedwith a large scale space solar power architecture. Thisinitial constellation would be replaced over time by the largegigawatt level spacecraft detailed in this paper.

2. OPTICAL CONCENTRATOR DESIGNMirasol/Girasol Orbit Selection Options

The Mirasols are required to concentrate solar power to theheater of the Brayton cycle heat engine. The Mirasol andGirasol aspects of each spacecraft pair are not physicallyconnected and there are multiple options considered for theorbit 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 altitudearound 20,000km or higher and reflect/concentrate powerdown to the LEO constellation of Girasols. This optionhas the advantage of being able to easily reach all partsof the globe and maintain near constant solar view. Thisoption is limited however by the spot size issue of reflectingsolar power over such large distances [10]. Thus, the areaof the Girasol collector dish would have to be increasedsubstantially for this option over the current design.2. Place Mirasols and Girasols together in sun-synchronousdynamic orbits: This option has the advantage that the spotsize issue is eliminated while constant solar view is main-tained, and the collector dish of the Girasol is no longer a nec-essary component of the spacecraft. Spacecraft placed in sun-synchronous orbits around 2000km can maintain constantsolar view by staying along the Earth’s day-night terminatoras identified by Potter and Davis [10]. The issue with thisoption is the ability to continuously beam to the dark side ofthe Earth.3. Place Mirasols and Girasols in close orbits at 2000km:This solution allows for a compromise between the othertwo options. Although the Girasol still requires a collectordish, the Mirasols, placed predominantly in sun-synchronousorbits, beam to nearby Girasols in orbits around 2000kmaltitude. The result is that the spot size is small enoughto be at manageable levels, while continuous beamed powerdelivery can be achieved to the dark side of the Earth.

A conceptual drawing of the Girasol part of the spacecraftreceiving concentrated solar power from the Mirasol collec-tor/concentrator is shown in Figure ??.

Configuration

The optical array design consists of a large array of aroundone thousand ultralight optical reflector elements. The con-figuration shape has not yet been optimized, but it can be con-ceptualized as some sort of concave shape with the target ofthe concentrated solar power being the heater of the Braytoncycle heat engine. This design eliminates the need for a largePower Management and Distribution System (PMAD).

When utilizing photovoltaic arrays for conversion, the use ofrefractive optical elements can be advantageous, as fresnellens optics can be used to increase the concentration of solarenergy multiple times. O’Neill has extensively researched thepossible use of streched fresnel lens refractive concentratorarrays for use in space solar power [11]. O’Neill argues thatin the long term, the Stretched Lens Array (SLA) proposedcan reach 1 kW/kg for MW level arrays. A 2.5kW Stretched

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Figure 1. Conceptual Drawing of Girasol part of the systemreceiving concentrated solar power from the Mirasol part ofthe system

Fresnel Lens array called SCARLET on the Deep SpaceOne Probe demonstrated 45 W/kg specific power [11]. TheSLA uses a 140 micron thick fresnel lens array made from asilicone rubber material. Certainly, the use of fresnel lensoptics makes sense when using a photovoltaic array for aspace-based solar power satellite. For a satellite using solarthermal conversion, such as the one described in this paper,the goal is to maximize the concentration of solar power atminimum mass while incurring the smallest losses. Ultralightreflective optical elements were chosen to achieve very highefficiencies at low mass and as a result were selected as thebest conceptual design choice.

Material

The material used must be ultrathin and lightweight. Themost likely optical materials to meet the design requirementswould be some form of polyimide film. A good startingpoint for selecting a material is to look at materials used forsolar sail applications, as solar sail materials are designedto be ultralightweight and highly reflective. Some solar saildemonstration spacecraft have already been deployed. JAXAhas successfully deployed the IKAROS spacecraft whichuses a polyimide sail. The United States has successfullydeployed the Nanosail-D2 spacecraft. This solar sail is 10square meters, and the spacecraft weighs only 4 kg. It usesan ultralight thin film material called LaRC-CP1 polyimidemade by ManTech [12] and developed at NASA Langley,on the order of a few microns in thickness [13]. Solar sailmaterial densities are on the order of 0.01 kg/m2 [14]. Thesolar sail material mass was calculated by simply takingthe specific mass of solar sail and multiplying by the arearequired determined from the power required from the end-to-end efficiency analysis.

ADACS Subsystem

An attitude determination and control system (ADACS) isrequired with actuators and sensors to maintain the necessarysun tracking and pointing requirements. The ADACS in-cludes an actuator and sensor system for each optical element.The mass estimate was determined by allotting 30kg for the

ADACS for each element, bringing the total ADACS systemmass to 28,800kg.

Mirasol Thrusters and Solar Radiation Pressure

The propellant mass requirement becomes the limiting factoron Mirasol specific mass (given that solar sail material can beso light), based on the deltaV requirement from solar radia-tion pressure. Given the use of ultralight thin film mirrors,the propellant required for solar radiation pressure becomes asignificant mass aspect of the Mirasol subsystem. The largearea per unit mass value implies that solar radiation pressureforce is substantially high to accelerate the spacecraft out oforbit, which needs to be balanced by some propulsive force.As the propellant mass is increased, the acceleration due tosolar radiation pressure force itself drops as the area to massratio decreases. In other, more massive architectures usingPV arrays, this is still a considerable issue, just not a largemass component, which explains why solar radiation pressureconsiderations are sometimes neglected in space solar powerstudies. The Aerospace Corporation SERT report [4] doesnot ignore it however, and it is considerable for their Haloconcept evaluation. Another thing to keep in mind is that thesolar radiation pressure acceleration will actually increase asthe propellant mass is used up, but the structural and attitudecontrol system mass estimates used in this analysis should beconservative enough to account for this.

The Delta-V required due to the solar radiation pressure forceis calculated by equation 1.

∆V =S ∗ Cr ∗ TM/A

(1)

S=Solar Flux at 1 AU from SunCr=surface reflection indexT=time interval of thrustingM/A = mass per unit area, specific mass of the Mirasol

From this equation it is clear that as the specific mass in-creases, less Delta-V will be required.

The deltaV requirements due to solar radiation pressure areplotted in Figure 2. Note that the plot starts at 0.01 kg/m2,as there is no way to get a lower mass than this because thesolar sail material itself has this specific mass value. Theplot shows the exponential decrease in Delta-V required withincreasing specific mass.

The Mirasol array would use krypton thrusters which canachieve a specific impulse of 5300 seconds. Using the rocketequation, the propellant mass required for the mission wascalculated over the 17 year spacecraft lifetime. The resultsare plotted in Figure 3.

Mirasol Subsystem Mass Summary

With the optical material system mass, propellant mass, andADACS subsystem required mass estimates, the last remain-ing item is to account for the structural support system for theMirasol concentrator. Roughly 15,000kg has been allottedfor the structural support system, an equivalent amount ofstructural mass as the solar sail material mass was estimatedfor the support required to maintain tension of the solar sailmaterial for each optical element.

The optimized Mirasol system mass was calculated by plot-

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Figure 2. Delta-V requirement for different specific massvalues for the Mirasol Array

Figure 3. Propellant Mass Requirement for Varying SpecificMass Values

ting the Mirasol mass required as the constraint, and thenplotting the Mirasol mass available by the specific powervalue. Finding the intersection of these two lines gives theminimum specific mass value that meets the mass require-ments for each subsystem. As seen in Figure ??, the designpoint is at a specific mass value of 0.06 kg/m2, with a totalMirasol system mass coming to about 92,000kg.

Mirasol Design Summary

A table summarizing parameters for the Mirasol concentratorarray is shown in Table ??. Using the power required to beconcentrated value calculated from the end to end efficiencyanalysis shown later, the area required could be determinedfrom the solar flux at 1 AU. The number of optical elements

Table 1. Mirasol Subsystem Mass Summary

Parameter MassSolar Sail Material Mass Required 15356 kgMirasol Structural Mass Required 15356 kgADACS Mass Required 28800 kgMirasol Propellant Mass 32753 kgMirasol System Mass 92000 kg

Figure 4. Mirasol System Mass for Varying Specific PowerValues

Table 2. Mirasol Concentrator Array Design Parameters

Parameter Value UnitsPower from Sun Per Area in Space 1.36 kW/m2

Power Required 2.09 GWArea Required 1.536 km2

Number of Optical Elements 960Area Per Element 1600 m2

Mirasol Specific Mass 0.06 kg/m2

Mirasol System Mass 92,000 kg

required was a parameter that was minimized in an effort toreduce the ADACS mass requirement to the extent possible.The total Mirasol concentrator system mass of 92,000 kg wasdetermined from the analysis shown earlier in this section.

3. OPTICAL DESIGN CONSIDERATIONSOptical Tracking

Since the required tracking is continuous and predictable,the solar tracking can be accomplished using a mechanicalattitude determination and control system with sensors andactuators.

Mirror Degradation

Longterm degradation effects from Near Ultraviolet (NUV)and Vacuum Ultraviolet (VUV) radiation could result in asmall decrease in performance over the lifetime of the array.Radiation from the space environment can cause darkeningof the optical material, altering the solar absorptance char-acteristics of the material [15]. However, degradation is notanticipated to significantly decrease the performance of thesatellite. Some level of robotic maintenance is required forall space-based solar power architectures and optical coatingscould be replaced periodically through on-orbit servicing.This is easier to do at the low earth orbit constellationproposed by the Space Power Grid as opposed to GEO basedarchitectures.

Micrometeroid Impacts

Micrometeroid impact damage will occasionally punctureholes in the array. The modular nature of the design meansthat overall performance of the concentrator will only be

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slightly decreased. These impacts can also be repaired overtime by on-orbit servicing of the spacecraft. Since the ma-terial to be used is ultralight, micrometeroid impacts shouldmerely result in small holes that don’t result in cracks or ripsthat propagate. In [14] testing on hypervelocity impacts forsolar sails did not show rip propagation.

Error Tolerance

Imperfections in the optical elements lead to shape errors thatresult in an uneven flux reflected by the optical element. Thisuneven flux should not pose an issue for the gigawatt satelliteproposed because the energy is being directed towards thecollector of a heat engine in order to heat the working fluidrather than photovoltaics.

4. BRAYTON CYCLE TURBOMACHINEDESIGN

The design of the Brayton cycle turbomachine used for thespacecraft, that we call the Girasol portion of the satellite, isdetailed in [16]. In that paper, subsystem components anda cycle analysis for a spacecraft using solar thermal Braytoncycle conversion were shown. High Brayton cycle efficien-cies can be achieved using ultra-high temperature ceramicmaterials such as hafnium carbide. This paper refines theGirasol design by refining the thermal control system designfor the Girasol.

Brayton cycle conversion was chosen over photovoltaic con-version because of the scaling advantages in specific powerof Brayton cycle conversion over photovoltaic at high powerlevels. At high power levels, photovoltaic arrays do notexhibit favorable scaling. Since photovoltaic arrays are asurface-area based phenomenon at their root, an increase inarea produces a proportional increase in power. Thus animprovement in specific power (power delivered per unitmass) is not seen in photovoltaic systems at higher powerrating levels. Brayton cycle engines do exhibit favorablescaling. As was shown in [16], a study of historical jet enginedata exhibited this trend of increasing specific power withincreasing power levels.

NASA Glenn has researched the use of Brayton cycle con-verters for space power applications since the Brayton Ro-tating Unit development program dating back to the 1960s.Currently, NASA Glenn has the 2kW Brayton Power Con-version Unit (BPCU) which has been tested and shown towork with nuclear fission heat sources [17]. The BPCU doesnot exhibit very high specific power because of its low powerlevels. In other words, it takes a lot more mass for a Braytoncycle heat engine to produce its first watt than to produceits ten millionth for example. Mason argues for increasingspecific power that can be achieved at high power levels in[18], offering favorable predictions for Brayton cycle engineperformance capabilities in the mid-far term.

Thermal Control System Design

In order to design a highly efficient lightweight cooling sys-tem several options were considered. Among those options,the exceptionally high thermal conductivity of Graphene andCarbon Nanotubes, along with their low mass densities, madethese options ideal candidates for the radiator material ofthe cooling system. Graphene consists of a single planeof sp2 bonded carbon-carbon atoms and as a result can beconsidered to be a two-dimensional material. The fact that itis two-dimensional makes Graphene have somewhat unusual

Table 3. Comparison of thermal conductivity and arealdensity of Graphene, Carbon Nanotubes and Graphite

ThermalMaterial Conductivity Areal Density

(W/mK)Graphene Sheet ∼3000-5300 6e-7 kg/m2

Carbon Nanotubes ∼3500 2e-5 kg/m2

Pyrolitic Graphite ∼1200-1600 ∼1900 kg/m3

properties; two of them are that it has an extremely highthermal conductivity and is incredibly lightweight. Table 3compares Graphene with other materials that have compar-atively high thermal conductivity. Using a radiator designsimilar to the one described by Juhasz in [19], and with a highenough emissivity, a thermal management system of below1 kg/m2 can be easily achieved with the use of Graphenebecause of its incredibly low mass density. Apart fromthis, new and developing heat pipe technology similar to onedescribed in [20] hold a strong potential for lightweight andhighly efficient heat management systems in space.

Although there is no definite value of thermal conductivityof Graphene that is agreed upon as of yet, the measuredvalues fall between 3000 W/mK to 5300 W/mK for differentmethods and different conditions [21]. Even as we increasethe number of layers, the conductivity decreases approachingthat of graphite, but still remains significantly higher thanmost other materials. Even after considering various effectslike losses due to wrinkling of the material, etc., the thermalconductivity of Graphene remains extremely high.

Juhasz proposed a radiator that could achieve 1kg/m2 in [22].We conservatively use this value as a specific mass estimatefor our thermal control system, noting that with the very lowmass density of graphene sheets and the very large scale ofthe system, it is likely that we could reduce the specific massconsiderably lower than this value. The radiator area requiredwas determined from the radiative power equation (shown inequation 2). The system mass was then estimated using thespecific mass value.

P = εσAT 4 (2)

P = Power Required to be Radiated from Girasolε = Emissivityσ = Stefan-Boltzmann ConstantT = Equilibrium Radiator Operating TemperatureA = Radiator Area Required

A table summarizing the parameters of the Thermal ControlSystem is shown in Table 4. Note that the emissivity value ishigher than the emissivity of graphene. This is an estimate ofthe emissivity value of the graphene sheets after the applica-tion of coating and paint has been applied. The current designrequires radiating a very large amount of power away fromthe spacecraft (400MW). As a result, a trade study is beingperformed to evaluate incorporating a recuperator into theBrayton cycle turbomachine design, to increase the efficiencyof the Brayton cycle to see if the power required for radiationcan be decreased, in order to decrease the size and mass ofthe thermal control system.

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Figure 5. Thermal Control System Mass Required forVarying Equilibrium Operating Temperatures

Table 4. Thermal Control System Parameters

Parameter Value UnitsPower Required To Be Radiated 400 MWOperating Temperature 662 Kσ 5.67e-8 W/(m2K4)ε 0.23Area Required 160,000 m2

TCS Specific Mass 1 kg/m2

TCS System Mass 160,000 kg

5. END-TO-END EFFICIENCY ANALYSISAn end-to-end efficiency analysis was calculated to accountfor various losses present in the system. Sizing the space-craft to deliver 1GWe to the terrestrial grid, the end-to-endefficiency allows us to determine the power required to beconcentrated by the Mirasol array. A summary of efficiencyvalues is presented in Table 5.

Current commercially available ultralight optical reflectorelements can achieve 92% efficiency using Mylar, Teonex,or LaRC-CP1 polyimide materials [14]. It is anticipated thatwith research and development this value would be able tobe increased to around 99%. Current commercially available

Table 5. Efficiency Breakdown

Parameter Value UnitsMirasol Power Required 2.09 GWMirasol Array Efficiency 0.99Brayton Cycle Efficiency 0.80Millimeter Wave Generator Efficiency 0.90Power After Brayton Conversion 1.49 GWWaveguide System Efficiency 0.97Phased Array Antenna Efficiency 0.90Power to Beam 1.30 GWAtmospheric Transmission Efficiency 0.90Beam Capture 0.95Rectenna Efficiency (RF to DC) 0.90Power to Ground Required 1.00 GWeEnd-to End Efficiency 0.48

heliostat array elements produced by Practical Solar canachieve 99.5% reflectivity [23].

The 80% Brayton cycle efficiency value comes from the cycleanalysis performed in [16], using hafnium carbide materialfor the turbomachine components. One area of uncertaintyin the efficiency analysis is in the ability of the millimeterwave generator to efficiently convert the power generatedfrom the Brayton cycle to 220GHz for beaming. It remainsto be shown that high power millimeter wave conversiontechnology can achieve these efficiency values and this is onearea of uncertainty for the spacecraft. These generators wouldhave to be compact, lightweight, and efficient. Barnes hasproposed compact, lightweight superconducting generatorsfor use in military power packs [24] that were scaled to sizethe generator for the Girasol.

Efficiency values for power transmission components wereinvestigated and detailed in [9]. A waveguide system isused to transmit power to the phased array antennas at veryhigh efficiencies. As a result of the orders of magnitudedecrease in the required antenna transmitter and receiverdiameters by moving to LEO and using millimeter wavetransmission, it was shown in [9] that several lobes of the Airydiffraction pattern could now be captured without reachingexcessive antenna sizes. Thus, beam capture efficiency couldbe increased from 84% power captured by the first ring as ingeosynchronous arhitectures to around 95% by capturing thefirst four to five lobes.

For microwave frequencies, rectenna efficiency for RF to DCconversion exceeds 90%. According to Koert and Cha, asearly as 1992, 70% efficient rectennas could be developed formillimeter wave frequencies with a hybrid rectenna circuitusing high power schottky diodes [25]. The reason highpower millimeter wave rectennas are more difficult to makethan microwave rectennas is because the size of the diodesneeds to be smaller in order to avoid impedance issues athigher frequencies. With research and development into highpower millimeter wave conversion devices and current de-velopments in Microelectrical Mechanical System (MEMS)manufacturing techniques, there does not appear to be any-thing preventing millimeter wave rectennas from surpassingover 90% efficiency.

An end to end efficiency value of 0.48 was calculated forthe Girasol/Mirasol 1 gigawatt spacecraft. The end-to-endefficiency value here is defined as the amount of powerdelivered to the ground per the amount of power incident onthe Mirasol array. For the spacecraft to deliver a gigawatt tothe electric grid, this allows us to size the Mirasol array bydetermining a Mirasol power required value.

6. COMPARISON WITH SIMILAR PROPOSEDCONCEPTS

Halo Concept

As part of the NASA Space Solar Power Exploratory Re-search and Technology program (SERT), The AerospaceCorporation looked at the technology available for space solarpower and evaluated a number of different space solar powerarchitectures [4]. One of the proposed architectures, the Haloconcept, was developed by Henry Harris at the Jet PropulsionLaboratory [4]. This concept uses a constellation of reflectorsto reflect towards a photovoltaic array. This architecture issimilar to the Mirasol design in the sense that, like a heliostat,a collection of optical reflectors is used to concentrate solar

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Table 6. Comparison with Halo and SPS-AlphaConcepts

Mirasol Halo SPS-AlphaItem Concept Concept ConceptOptical 960 optical 100 orbiting ∼5000Elements elements reflectors reflectorsEnergy Solar PV PVConversion Thermal Array ArrayPower Mill Wave Solid State ∼200,000Generation Generator Amplifiers ModulesBeam 220GHz 5.8 GHz 2.45 GHzFrequency (Mill Wave) (Microwave) (Microwave)Power Level 1 GW 1.2GW 2 GWSystem Mass 677MT 13,044MT ∼9,350MT

energy. Some key differences exist between the Halo conceptand the Mirasol concept. A comparison of the two conceptsis summarized in Table 6.

Of note, of the four microwave based space solar powerarchitectures studied by The Aerospace Corporation, theHalo concept performed considerably better in terms of theoverall system level metrics of maximum specific power andminimum cost per kilowatt hour than the other architecturesevaluated. The system mass for the Mirasol/Girasol satelliteis significantly less than the estimate for the Halo conceptgiven several factors. One key reason is the very largereduction in antenna size. The area of the antenna becomesorders of magnitude less when moving to 220GHz beamingand bringing the orbit down to LEO, reducing the beamingdistance from 36,000km to 3,000km (2000km altitude, plusaccounting for 45 degree coning angle). Assuming thatantenna area scales with mass, the antenna mass of the systembecomes orders of magnitude less. The power conversionsystem mass is the other key mass driver that comes downconsiderably, since Brayton cycle conversion exhibits scalingadvantages at higher power levels over photovoltaic. Anotherimportant note on mass reduction is that given the muchhigher efficiency of Brayton cycle conversion over photo-voltaic systems, the heat that must be radiated away fromthe spacecraft is reduced considerably. This greatly reducesthe required thermal control system mass, another key massdriver of the system.

SPS-Alpha

The NASA Institute of Advanced Concepts is funding a studyinto the conceptual design of SPS-Alpha, a space solar powersatellite concept proposed by John Mankins[8]. SPS-Alphais similar to the Mirasol/Girasol satellite proposed here inthat they both utilize a large array of individually pointablethin film mirrors to redirect sunlight. The key differencesare that SPS-Alpha would be in geosynchronous orbit, wouldutilize photovoltaic conversion rather than Brayton cycleconversion, and microwave beaming (2.45 GHz) rather thanmillimeter wave beaming. The Mirasol/Girasol again showsa significant mass reduction, due primarily to similar reasonsas outlined above for the Halo concept.

Table 7. Mass budget for 1 GWe Space Solar PowerSatellite

Mass % of % ofElement (kg) Girasol SatelliteCollector 30,000 5.21 4.49

Cooling System 160,000 27.79 23.96Brayton Cycle 20,000 3.47 2.99AC generator 50,000 8.68 7.49Cryogenics 20,000 3.47 2.99

220GHz Amp 17,000 2.95 2.55Antennae 20,000 3.47 2.99

Propulsion 170,300 29.58 25.50Misc. 30,930 5.37 4.63

Structure 58,470 10.00 8.62Total Girasol 584,700 100.00 86.22Total Mirasol 92,000 13.78

Total Mass 676,700 100.00

7. SUMMARY AND CONCLUSIONSSatellite Mass Summary

A mass summary for the different subsystems is shown in Ta-ble 7. This system mass summary reflects refined estimates ofthe spacecraft cooling system and the Mirasol optical systemmass calculations, as these systems have been investigated infurther detail.

Conclusions

The design of an optical concentrator system for a gigawatt-level space solar power satellite was proposed. A large arrayof individually pointable ultralight, thin film optical elementscan redirect and concentrate sunlight towards the heater ofa Brayton cycle solar dynamic heat engine for conversionof solar power for space-based solar power applications. Acooling system using graphene sheets offers the potentialfor efficient, lightweight radiation of heat for the converterpart of the spacecraft system. The design is shown to closeusing conservative mass and efficiency estimations. Thespacecraft offers potential for large improvements in specificpower, the amount of electric power that can be delivered toground per unit mass required to be launched into orbit, overtraditional space solar power architectures (∼1.5kW/kg vs.∼0.2kW/kg).

ACKNOWLEDGMENTSThis work is partially supported as an advanced concept de-velopment study under the EXTROVERT cross-disciplinaryinnovation initiative by NASA. Mr. Tony Springer is thetechnical monitor.

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BIOGRAPHY[

Brendan Dessanti received his B.S.degree in Aerospace Engineering fromthe Georgia Institute of Technology in2011 and is currently pursuing an M.S.Degree in Aerospace Engineering fromGeorgia Tech. Brendan is a graduateresearch assistant for the ExperimentalAerodynamics and Concepts Group atGeorgia Tech and leader of the SpacePower Grid student team. His research

focuses on space systems design as it applies to space-based solar power. Brendan has obtained experience frominternships at Sikorsky Aircraft, MIT Lincoln Laboratory andSpaceWorks Enterprises, Inc.

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Narayanan Komerath is a professorin the Daniel Guggenheim school ofaerospace engineering at Georgia Insti-tute of Technology (G.I.T.), and direc-tor of the John J. Harper Wind Tunneland the Experimental Aerodynamics andConcepts Group. He has served as aNIAC Fellow, and as a Boeing WelliverSummer Faculty Fellow.

Shaan Shah is a senior at GeorgiaTech, pursuing his B.S. in AerospaceEngineering. He has participated in re-search as a member of the Experimen-tal Aerodynamics and Concepts Group,mainly involved with the developmentof a conceptual design for the SpacePower Grid. His research focuses onstudying breakthroughs in materials likeGraphene and Ultra High Temperature

Ceramics, and technological advances related to space-basedsolar systems in the fields of thermoelectrics, thermophoto-voltaics and narrow-band photovoltaics. This experience hasleft him with an interest to study further the possibilities ofSpace-based Power Systems. He intends to continue graduatestudies in the area of propulsion and combustion.

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