design of a millimeter waveguide satellite for space...
TRANSCRIPT
Design of a Millimeter Waveguide Satellite for SpacePower Grid
Brendan Dessanti, Richard Zappulla, Nicholas Picon, Narayanan KomerathDaniel Guggenheim School of Aerospace Engineering
Georgia Institute of TechnologyAtlanta, GA 30332
Abstract—A central element in the use of Space as a powergrid for near-real-time power exchange, is a constellation ofsatellites, capable of receiving and relaying beamed power atmulti-megawatt level with extremely high efficiency and lowthermal losses. The conceptual design of such a satellite isconsidered in this paper. Preliminary calculations indicate thatpower delivery and transmission will use millimeter waveguidesinside the satellite to guide the power between antennae, andas feeds for the antennae. The paper seeks solutions for themilimeter waveguides, antennae and thermal control systemsin order to refine the mass and efficiency estimates. While thestandardized design is for 2000-kilometer orbits, an option touse fewer satellites between as few as two participating nationsat 5500 km is also considered at the startup of the system. TheNASA system design process is followed wherever practical.However, some innovations are required beyond to move intothe regime of 60MW power beams with millimeter waveguides.Corrugated waveguides are identified as a solution for 220 GHzbeam redirection, directly coupled to antenna elements.
TABLE OF CONTENTS
1 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 DESIGN PROCESS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13 DEFINITION OF NEED AND DESIGN REQUIRE-
MENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24 PRELIMINARY PARAMETERS AND CONFIGU-
RATION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 PROPELLANT AND MASS BUDGETS . . . . . . . . . . . . . 26 WAVEGUIDE SYSTEM . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27 ANTENNA SIZING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38 ACTIVE THERMAL CONTROL SYSTEM . . . . . . . . . 49 OTHER SUBSYSTEMS DESIGN . . . . . . . . . . . . . . . . . . . 510 SATELLITE DESIGN SUMMARY . . . . . . . . . . . . . . . . . 511 DEMONSTRATION SATELLITE SIZING . . . . . . . . . . 612 CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
ACKNOWLEDGMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . 7BIOGRAPHY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
1. INTRODUCTIONThe purpose of this paper is to show that a certain type ofsatellite can be designed, to address a special need. This is amillimeter waveguide satellite capable of relaying significantamounts of power that meets the requirements for a PhaseI Space Power Grid (SPG) satellite as outlined in [1]. Theconceptual design in this paper will show that the satellite can
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meet the given requirements needed to keep spacecraft massand throughput efficiency within the constraints necessary forthe Space Power Grid architecture to become economicallyviable before any power is generated in Space. This is notclaimed to be an optimal design, but only to prove basicfeasibility. Given that one solution exists, a better optimumcan be sought in future work.
The Space Power Grid (SPG) architecture seeks to developa comprehensive space and terrestrial renewable power so-lution for the planet, by first enabling remotely-located ter-restrial solar plants to sell and receive power worldwide,and use this market to develop the infrastructure into whichspace-generated solar electricity will be added. Thus inthe first stage, the architecture calls for a constellation ofsmall satellites, each of which essentially functions as anintelligent waveguide junction. This satellite is to receivebeamed power from terrestrial stations and other satellitesthat track it using phase array transmitters, and distributesthe power immediately to other satellites and to ground-basedor aerial receiver platforms. The subject of this paper is thedesign of such a satellite. Traditional cost estimation and theSPG conceptual model calculations yield a mass of roughly4000 kg for each such satellite, with a nominal choice ofsun-synchronous and equatorial orbits at 2000 to 6000 km,transacting 60MW of 220GHz beamed power with a 17-yearlifetime.
2. DESIGN PROCESSThe iterative process used to design the spacecraft was similarto the approach in Table 10.1 of [2]. The Space MissionAnalysis and Design (SMAD) process shown below has beenmodified slightly to be used for a waveguide satellite designas part of the Space Power Grid Architecture. The processused in the design outlined in this paper follows the stepsbelow:
1. Define the need and design requirements from establishedSpace Power Grid Architecture
2. Determine preliminary spacecraft parameters and overallconfiguration selection
3. Calculate budgets for power and mass
4. Develop waveguide system and subsystem designs
5. Develop spacecraft configuration
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Parameter ValueOrbit Altitude (km above Earth) 2000Frequency (GHz) 220Design Power (MW) 60Satellite Lifetime (years) 17Total antennas Per Satellite 3Space-Space antennas 2Ground-Space Antenna 1
Table 1. Design Requirements Outlined by SPGArchitecture
3. DEFINITION OF NEED AND DESIGNREQUIREMENTS
The Space Power Grid architecture is a three stage evolu-tionary approach to bring about terrawatt level space-basedsolar power generation. In the first phase of the architecture,a constellation of relay satellites at a mid-low earth orbit areused to beam power from terrestrial renewable energy plantswith surplus energy to locations around the globe with highdemand. Systems architecture level trade studies have beenperformed that have established the requirements shown inTable 1.
The two specific design requirements that drive the need forestablishing a millimeter waveguide spacecraft capable ofrelaying significant quantities of power are the design fre-quency selection (220 GHz) and the design power (60 MW).The millimeter wave frequency design choice is a tradeoffbetween a number of factors. The key driver in selectinga millimeter wave frequency was the relationship betweenantenna and transmitter diameter. The high frequency choicebrings down the antenna diameter, decreasing mass of thespacecraft. For more detail, on this system design choice orother system design choices such as orbit altitude and satellitelifetime, reference other papers from our group[3][4][5].
4. PRELIMINARY PARAMETERS ANDCONFIGURATION
The general spacecraft configuration chosen is a 3-axis stabi-lized spacecraft with an equipment compartment in the shapeof a hexagonal cylinder. The general arrangement is similarto the TDRS satellite. The space to ground antenna pointstoward nadir, while the two space to space antenna are locatedon opposite sides of the spacecraft. The radiators are locatedon the backside of the space to space antennas. This takesadvantage of the large amount of area available here. Asystem of two waveguides is used to transmit power from areceiving antenna to a transmitting antenna of the spacecraft.An illustration showing the general arrangement and relativesizes of the space to space antennae, space to ground antenna,and spacecraft bus is shown in Figure 1.
Table 2 shows the initial parameters necessary in order todevelop the power and mass budgets. The space to spaceantenna and space to ground antenna diameters were deter-mined using the information from section 7. The antennamass per unit area is a key metric in the feasibility of a solarpower satellite. As a result of the large distances associatedwith beaming from space, even with a high frequency choiceand Low-Earth Orbit selection, the antenna area will still belarge. Since the payload mass is a key driver in the totalspacecraft mass, which in turn drives the launch costs, the
Parameter ValueSpace-Space Antenna Diameter (m) 90Space-Ground Antenna Diameter (m) 50Specific Impulse (sec) 5300Antenna Mass/Unit Area (kg/m2) 0.05
Table 2. Preliminary Spacecraft Parameters
Figure 1. Illustration showing general arrangement ofantennas and spacecraft body
ability to achieve a low value for this metric is essential to theviability of any solar power satellite. The value of 0.05kg/m2
has been chosen for this analysis. Using a mesh flat platearray antenna made of lightweight material, this seems like areasonable estimate of antenna mass per unit area.
5. PROPELLANT AND MASS BUDGETS
With the antenna sizes and a mass per unit area value deter-mined, a payload mass was estimated. From here, a massbudget was created for the different subsystems as shownin Table 3. Subsytem mass budget values were determinedbased on estimated percentage of dry mass using historicalspacecraft data[2]. The overall spacecraft results and di-mensions for the spacecraft bus are shown in Table 4. Thepropellant budget was calculated by estimating the delta-vrequirements for a satellite in a 2000km altitude orbit.
A spacecraft of the size and weight determined by the massbudget could be launched to LEO using a Delta II rocket.The Delta II rocket can carry satellites with a loaded mass ofroughly 6000kg up to Low-Earth Orbit at an estimated launchcost of about $40 million[6].
6. WAVEGUIDE SYSTEMThe waveguide system is designed to transmit power froma receiving antenna of the satellite to a transmitting antennawith minimum losses. Any power losses from the waveguidesystem adds heat to the spacecraft that must be radiatedfrom the spacecraft, driving the mass of the thermal controlsystem. Corrugated waveguide structures can be designed
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Subsystem Mass (kg)Payload (3 antennas) 734.3Propulsion 75.0Attitude Control 179.6C & DH 64.3Thermal 217.1Electrical Power 774.5Structure and Mechanisms 570.9Waveguide 100.0Communications 64.3
Table 3. Initial Mass Budget for the Different Subsystems
Parameter Mass (kg)Total Dry Mass (kg) 2680Dry Mass w/ Contingencies (kg) 3015Propellant Mass (kg) 510Loaded Mass (kg) 3526Spacecraft Volume (m3) 17.73Spacecraft Diameter (m) 2.5
Table 4. Overall Spacecraft Sizing, Mass, and PropellantBudget
to transmit power in the millimeter wave frequency rangearound 220GHz in a near lossless manner[7].
General Atomics produces straight corrugated waveguideswith very low losses in the HE11 mode[8]. General Atomicsproduces corrugated waveguides at a number of frequencieswith a diameter of 31.75mm and a diameter of 63.5mm.As frequency increases, attenuation decreases exponentiallyuntil the frequency reaches the region where Bragg scatteringcan occur[8]. For a frequency of 220GHz, the 63.5mmdiameter waveguide has a near zero attenuation loss. Anexample of what the corrugated waveguide structure lookslike can be found on the General Atomics website[8].
The waveguide system consists of two waveguides. Oneconnecting the space to space antennas through the space-craft, and the other connecting the space to space antennasto the space to ground antenna. The antennas were placedsufficiently far away from the spacecraft body such that theradiated heat from the thermal control system would not heatthe spacecraft body. This drives the values for length shownin the table. The material chosen for this design is copper dueto its conductive properties and because it is inexpensive.
The waveguide system is capable of transmitting large quan-tities of power by leveraging the fact that space is a vacuum.The amount of power a waveguide can transmit is limitedby the Electric Field Breakdown Limit of the medium[9].The electric field breakdown limit in a vacuum is an orderof magnitude higher than that of air[10]. The values forwidth, depth, period, and diameter that define the corrugatedstructure of the proposed waveguide system shown in Table5 match the 63.5mm corrugated waveguide produced byGeneral Atomics.
Efficiency through the waveguide was assumed to be 99%because of the near zero attenuation losses of the corrugatedwaveguide. Estimating a 1% efficiency loss at the transmit-ting antenna junction and receiving antenna junction, a total
Parameter ValueLength Waveguide 1 (m) 18.5Length Waveguide 2 (m) 20.3Total Length (m) 38.8Material CopperMedium VacuumMode HE11
Corrugation Period (mm) 0.66Corrugation Width (mm) 0.46Corrugation Depth (mm) 0.41Diameter (mm) 63.5Frequency (GHz) 220Max Power Transmitted (MW) 60Attenuation (dB/10m) 0.001Efficiency Through Waveguide 0.99Efficiency Waveguide-Antenna Junction 0.99Total System Efficiency 0.97Power Loss (MW) 1.8Density Material (g/cm3) 8.94Wall Thickness (mm) 2Mass/Unit Length (kg/m) 1.81Mass (kg) 70.3
Table 5. Waveguide System Properties
waveguide system efficiency was estimated to be 97%. Usingthe design power of 60MW, this results in a loss of 1.8MW ofpower that must be radiated away from the spacecraft as heat.Using a nominal wall thickness value of 2mm, a system masswas estimated to be roughly 70kg.
7. ANTENNA SIZINGThe SPG concept attempts to minimize system costs bygreatly reducing antenna transmitter and receiver sizes com-pared to GEO based microwave concepts. The antennatransmitter/receiver diameter relationship comes from thefundamental limit to resolution, diffraction. The derivation ofthis relationship is given in standard textbooks on diffractiontheory. Fraunhofer Diffraction at a circular aperture can berepresented using a Bessel function of the first kind [11]The amount of power received is calculated from this in theRayleigh Limit[11]. The results are shown in Figure 2. Thethree data points correspond to the first three rings of the Airydisc. Using the transmitter and receiver diameter relationshipfor different fractions of power received gives Figure 3. Thisplot shows the results of capturing the first three rings onpower received. An additional data point shows that inorder to receive 98.5% of the transmitted power, an order ofmagnitude increase in this constant value is required. Theseplots show the design challenge traditionally faced by an-tenna designers. For radar and communications applications,designers optimize by choosing to receive only the centrallobe (84%). This practice, which is immensely wasteful forpower beaming applications, has been assumed in most SpaceSolar Power architectures that use GEO and low microwavefrequencies, because the antenna size is already close to beingprohibitive there. With the choice of 2000 km altitudes and220 GHz frequency, the designer can have the luxury ofspecifying at least 95 % capture.
Using this information, the spacecraft antennas were sized.The space to ground antenna design diameter is 50 meters, asthis is sufficient to get about 95% efficiency with reasonably
3
Airy Ring % Power J1 Zeros kR kD
1st Ring 83.8 3.8317 1.220 2.442nd Ring 91.0 7.0156 2.233 4.473rd Ring 93.8 10.1735 3.238 6.484th Ring 95.2 13.3237 4.241 8.485th Ring 96.1 16.4706 5.243 10.49
Table 6. Values for Diffraction of Beam
Figure 2. Amount of Power Received in terms of x fromBessel Function of the first kind
Figure 3. Amount of Power Received in terms of theconstant relating transmitter and receiver diameter
Figure 4. TCS Mass vs. Equilibrium Temperature
sized ground receivers (100-200m diameter). The two spaceto space antennas need to be larger to account for the fact thatthe space to space receiving antenna is limited to the space-space transmitting antenna diameter. The design choice of90 meter diameter was chosen to provide the same level ofcapture efficiency. From these antenna sizes and using the0.05kg/m2 mass per unit area value shown in Table 2, a massestimate for the three antennas was determined.
8. ACTIVE THERMAL CONTROL SYSTEMAn active thermal control system was chosen to radiate excessheat from the spacecraft. Figure 4 shows the relationshipbetween the thermal control system mass and equilibriumtemperature for varying waveguide efficiencies. Using awaveguide efficiency of 97%, the plot shows that the equi-librium temperature of the spacecraft body must be rela-tively high in order to bring the mass within feasible levels.Choosing a design equilibrium temperature of 720K bringsthe Thermal Control System mass to roughly 1,000kg. Inorder to achieve an equilibrium temperature of 720K, thedecision was made to create a dual box spacecraft body. Allthe temperature sensitive Flight System Components, such asthe Attitude Determination and Control system, Commandand Data Handling system, Propulsions system, and Commu-nication system will be housed in the cold body in order tokeep these components within operating limits; the waveg-uide system is housed in the hot body. Figure 5 shows therelationship between equilibrium temperature and requiredradiator area. Using the chosen design point yields a requiredradiator area of 250 m2. This area is far too large to be locatedon spacecraft bus; however, the backside of the antennasprovides adequate area for the Thermal Control System. Theplots show that if a greater efficiency of the waveguide systemwas achieved, the system mass and equilibrium temperaturecan be lowered significantly.
In a space environment, heat is transferred via radiationor conduction. As a result, the Flight Systems bus canbe thermally isolated from the waveguide payload by (1)placing the radiators in areas where the Flight Systems boxis out of the radiation path; (2) minimizing the contact areabetween the payload and Flight Systems box; and (3) utilizing
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Figure 5. Required Radiator Area vs. Equilibrium Temper-ature
Figure 6. Illustration of Two Box Main Spacecraft Body
low thermally conductive materials at the contact interfaces.Figure 6 illustrates an example of the thermally isolated two-body configuration. Lastly, the number of supports thatseparate the two boxes is governed by the ability of thesupports to carry the loads imposed upon it during launch.
9. OTHER SUBSYSTEMS DESIGNPropulsion System
The propulsion system uses a Krypton ion thruster. Themost ideal ion thruster propellant in terms of performaceis Xenon. However, due to the relative abundance of thetwo materials, Krypton is a much cheaper option (about an
order of magnitude cheaper). Using a krypton thruster, aspecific impulse of about 5300 seconds can be achieved. Theion thrusters perform orbit stationkeeping for the spacecraft.Since the spacecraft uses a 3 axis control system, 6 thrustersare required to perform the necessary thrust maneuvers tomaintain the desired orbit. Two thrusters are placed on eachside of the spacecraft. This will allow for fore and aft thrust-ing along with rotation about an axis. The last two thrusterswill be placed on the antenna booms. This configuration willallow for thrusting about the primary pointing axes.
Command and Data Handling System
The Command and Data Handling (C&DH) system is re-sponsible for processing telemetry and basic satellite house-keeping functions, such as stationkeeping. Due to the largeamounts of power being transmitted in the close vicinity ofthe spacecraft, the satellite data bus will utilize EIA-422 (RS-422) differential signaling. This not only allows for datatransfer rates up to 10 Megabits per second at distances up to12 meters but also decreases the susceptibility to any Electro-magnetic Interference (EMI). Moreover, asynchronous datalink protocols can be used to check for and protect againsterroneous transmission on the data bus. Additionally, both theC&DH working and storage memory will utilize non-volatileflash memory that is passively shielded to guard against theradiation environment of space. Lastly, the C&DH subsystemwill utilize Class S reliability components due their higherreliability and single string redundancy.
Attitude Determination and Control System
In order to maintain the pointing accuracies to minimize thepointing losses of the antennas, the satellite will be a three-axis stabilized satellite utilizing zero momentum reactionwheels and six (6) ion thrusters. The reaction wheels willbe used to slew the satellite as well as counter any externallyapplied torques imposed on it by the environment. The ionthrusters will be used for basic station keeping as well asreaction wheel desaturation.
Electrical Power System
Given the unique mission and life of the spacecraft, thissatellite will utilize the waste heat produced by the ineffi-ciencies of the waveguide system to power the Flight Systemcomponents. The target metric for the heat-recovery engineis approximately one (1) kilogram of mass for every kilowattof power generated.
Structure and Mechanisms
The primary structure of the satellite must be able to copewith the temperature flux and exhibit good corrosion resis-tance characteristics. Given these requirements, Aluminium6061-T6 was chosen as the primary structural material as itprovides the optimum combination of durability, strength-to-weight ratio, ease of manufacturing, and cost. The internalstructure of both the payload and Flight System structuralbuses will feature an internal frame with cross-bracing be-tween the structural members that provides structural rigiditywhile reducing the overall mass of the satellite.
10. SATELLITE DESIGN SUMMARYTable 7 gives refined mass estimates for the spacecraft afteriterating through the design process. The biggest change isin the estimate for the thermal control system. The totalspacecraft mass reflects the changes in subsystem mass.
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Subsystem Mass (kg)Payload (3 antennas) 734.3Propulsion 75.0Attitude Control 179.6C & DH 64.3Thermal 989.0Electrical Power 774.5Structure and Mechanisms 570.9Waveguide 70.3Communications 64.3Total Spacecraft Dry Mass 3422Dry Mass w/ Contingencies (kg) 3757Propellant Mass (kg) 510Loaded Mass (kg) 4267Spacecraft Volume (m3) 17.73Spacecraft Diameter (m) 2.5
Table 7. Mass Summary
Atmospheric and Ionospheric Propagation Issues
In prior work [12] we have dealt at length with the atmo-spheric propagation issue. Data from observatories at 4000meters altitude shows over 90 percent reception of 220GHzsignals through the atmosphere, versus well over 95 percentfor frequencies below 10 GHz. However, as mentionedbefore, the use of 84 percent power capture using undersizedantennae below 10 GHz, versus 95 percent capture at 220GHz, reverses this comparison, and shows that 220 GHzis effectively much more efficient. The difficulty comes inthe transit of the atmosphere below 4000m, where densityand moisture make losses at 220 GHz quite high, especiallywhen moisture is present. Three arguments can be advancedas avenues to break through. The first is that many ideallocations for solar and wind plants, are indeed in dry lo-cations, and many are also at high altitude, such as theNew Mexico desert and the Greenland, Antarctic and Tibetanplateaux. Secondly, data [13] show that the actual numberof hours with significant precipitation over a wide swath ofthe USA from southern California to the north-central plains,is small enough to not pose an issue. Thirdly, with theSPG, finding alternative sites with low or no precipitationwithin a couple of hundred miles is usually easy, given theconvective thunderstorm nature of precipitation at least inSpring through Fall. Other options are to find ways of burningclear conductive paths in the cone above a ground station.The interaction of multi-megawatt-level power with moist airis a subject of interest. Recently [14] a more elegant solutionhas been proposed. This is to use Lighter Than Air (LTA)platforms as the actual antenna for ground stations. LTAs canbe tethered at 4000 to 5000 meters altitude, and the tethersthemselves can be used as highly effective waveguides (asshown in the waveguide section here) to convey millimeterwave beams through the moist, dense lower atmosphere.Evacuating the waveguides inside the tethers is not a difficultor expensive concept. Alternatively, very large antennae canbe located on statospheric platforms (Straforms) that hoverat altitudes such as 21000 meters, capturing beams fromSpace, and distributing them on to much smaller retail groundantennae.
Ionospheric effects on millimeter wave radiation have beenconsidered in the literature. This becomes an issue of interestbecause the SPG satellites will use polar orbits, and hence
Parameter ValueEfficiency through Atmosphere 0.90Ground Receiver Capture Efficiency 0.95Satellite Receiver Capture Efficiency 0.95Space Receiver Antenna Efficiency 0.90Space Transmitter Antenna Efficiency 0.90Efficiency of Waveguide System 0.97Total Spacecraft Efficiency 0.79End-to-End Efficiency 0.43
Table 8. Summary of Efficiency Values
transit the strongly ionized regions near the poles. Twomain effects have been noted [15]: propagation delay, andrefraction. These cause error when the signal is used for rangecalculation. Both are significant only below 10 GHz, withthe slopes of the curves suggesting that these are not issuesabove 100 GHz. No doubt, these will cause pointing errors,but these are easily compensated. Some speckle is expectedin the beam; however the intent here is power delivery and notsignal clarity. Some effect on efficiency is expected, howeverthere is no reason to fear that this will be a significant issue at220 GHz. If losses are found to be significant, then the long-term effects of such a constellation on the ionosphere itselfmust also be considered, but at present there is no evidence tosuggest any such issue.
End-to-End Efficiency
Table 8 gives a breakdown of the predicted efficiency lossesexperienced in beaming power from one ground station toanother. Using high altitude aerostats as outlined in the mostrecent Space Power Grid architecture, atmospheric losses areestimated to be about 10%. Assuming the receiving andtransmitting efficiency of the waveguide spacecraft to be 90%each, assuming 95% capture at the satellite receiver, andassuming the waveguide system has a 97% efficiency, eachspacecraft is estimated to be 75% efficient. In this calculationthe power is beamed from one satellite to another before be-ing beamed back through the atmosphere. Finally, using theantenna transmitter and diameter relationship derived earlier,assume that the ground receiver has a 95% capture efficiency.Total end-to-end efficiency is calculated as follows.
ηend−to−end = ηatm1∗ηsat1 ∗ηsat2 ∗ηatm2
∗ηground capture(1)
ηend−to−end = 0.90 ∗ 0.75 ∗ 0.75 ∗ 0.90 ∗ 0.95 = 0.43 (2)
Thus, it is predicted that almost half of the power transmittedfrom the transmitting Earth location could be captured by thereceiving Earth location.
11. DEMONSTRATION SATELLITE SIZINGThe need for a near term demonstration of beamed powerfor demonstrating technical feasibility of spaced-based solarpower concepts has been identified[16][17]. The SpacePower Grid architecture proposes a US-India Power Ex-change as a demonstration that would serve as a first step
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Parameter Phase I Sat Demo SatSat Receiver Capture Efficiency 0.95 0.84Total Spacecraft Efficiency 0.79 0.75End-to-End Efficiency 0.43 0.34Space-Space Antenna Dia. (m) 90 236Payload Mass (kg) 734.3 4472.5Total Spacecraft Dry Mass (kg) 3422 20500
Table 9. Comparison between SPG Phase I Satellite Designand Demonstration Satellite Design
towards a full space-based solar power architecture[5]. Usingthe 4 satellite model at an altitude of 5500km, the satellitesare not close enough to each other to be able to achieve 95%capture efficiency and maintain the same antenna diametersize. In the SPG Phase I architecture, there are enoughsatellites such that the 2000km altitude satellites are closeenough to each other to get 95% capture efficiency at the 90mspace to space antenna diameter. With four satellites equallyspaced at this altitude, the distance between any two of thesatellites can be found using the law of sines.
(Re + h)
sin45=
x
sin90(3)
with Earth radius Re = 6370km and altitude h = 5500km,the distance x = 16787km.
At this distance, the antennas must be resized using equation7. Since receiving diameter and transmitting diameter ofthe space to space antennas must have the same diameter,set Dr = Dt. Solving the equation yields a diameter of236m at a capture efficiency of 84%. Plugging this satellitecapture efficiency value into equations 9 and 10, the end-to-end efficiency for the demonstration model becomes 34%.
Estimating total spacecraft mass to be proportional to themass of the antenna payload. The demonstration satelliteswill be roughly 6 times the mass of the Phase I satellite.Table 9 shows a comparision of values calculated for thedemonstration model satellite and the SPG Phase I satellite.
The demonstration satellite is significantly heavier and lessefficient. The proposed demonstration would not be prof-itable in and of itself. However, that would not be the goal ofthe demonstration. By demonstrating the technical feasibilityof the concept, and the mechanisms of international collabo-ration needed for such an exchange across the world, such ademonstration could significantly reduce the technology andpolicy risks associated with developing a large space solarpower architecture. Viewed in this context, the cost of such ademonstration is miniscule.
12. CONCLUSIONSThe millimeter waveguide spacecraft design sizing estimatefits within the bounds necessary to meet the Space Power Grideconomic model laid out at the 2011 IEEE Aerospace Con-ference. The mass is low enough so that the launcher classfor launch to LEO is that of a Delta II rocket. The satelliteefficiency values are sufficient for providing enough powerto intended markets at reasonable prices such that the SpacePower Grid Phase I satellites can reach breakeven withintheir 17 year design lifetime. There does not appear to beany technical show stoppers that would prevent a high power
millimeter waveguide spacecraft from being developed.
ACKNOWLEDGMENTSThe authors thank NASA for supporting the presentation ofthis paper under the EXTROVERT cross-disciplinary learn-ing resource project at Georgia Institute of Technology. Mr.Tony Springer is the technical monitor.
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-basedsolar power. Brendan has obtained invaluable experiencefrom internships at Sikorsky Aircraft, MIT Lincoln Labora-tory and SpaceWorks Enterprises, Inc. At Sikorsky, analyzedflight loads for the Black Hawk Helicopter as a member of theLoads and Survivability group; at MIT Lincoln Laboratory,analyzed gravitational harmonics for the purpose of missileflight simulations in the Missile Defense Systems Integrationgroup; and at SpaceWorks Enterprises performed variousanalysis and presentation preparation tasks in support ofNASA Lunar Surface Systems cost estimating and systemintegration.
Nicholas Picon is in his second year atGeorgia Tech pursuing his B.S. degreein Aerospace Engineering. He is alsoworking on his minor in Computer Sci-ence with a concentration on ArtificialIntelligence. Nicholas has spent threesemesters working in the Experimen-tal Aerodynamics and Concepts Groupat Georgia Tech, where his main rolewas designing constellations to optimize
coverage time for various orbital networks. Nicholas hasco-authored several papers on the subject of wireless powertransfer through space and advised Boeing on trade studiesrequiring optimization of orbital networks. He plans to internwith Rolls-Royce in the Summer.
Richard Zappulla 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. Richard is a GraduateResearcher with the Space System De-sign Laboratory and the ExperimentalAerodynamics and Concepts Group. Hisresearch focuses on the development of
novel launch systems for CubeSats as well as SpacecraftSystems Development, in particular, the Command & DataHandling and Flight Software systems. Additionally, Richardis the Project Manager for Georgia Techs NASA UniversityStudent Launch Initiative (USLI) team. Richard has gained
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invaluable experience from internships and CO-OPS at the JetPropulsion Laboratory (JPL) and HEICO Aerospace. Whileat JPL, he was responsible for the development and testingof the Wind Guard System for the Mars Science Labora-tory (MSL) sample drop-off system; at HEICO aerospace,he gained aided in optimizing the manufacturing processesfor critical-to-flight aerospace parts as well as the reverseengineering of various aircraft parts.
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.
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