Mars Relay PowerStar RFI Concept

8/11/2019 Mars Relay PowerStar RFI Concept

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RFI RESPONSE, NAIC CODE

#54172

By: Dr. David Hyland, of: Texas A & M Universityemail: [email protected]

Shawn Paul Boike, of: American Industrial Co. (DUNS#962375700) email: [email protected] / [email protected]

Business Model for Commercializing of Mars Relay Services

BUSINESS APPROACH & RECOMMENDATION
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INTRODUCTION

This is our Industrial & Universities response to NASAs Request For Information (RFI) for how they mightsustain Mars relay infrastructure, consisting of orbiters capable of providing standardized telecommunicationservices for rovers and landers on the martian surface, in the martian atmosphere, or in Mars orbit. Usingmass production techniques for satellite structure, power & antenna relays. The abundant supply of fusion-

based energy produced by the sun still remains to be efficiently harvested. The collection of solar radiation inspace could potentially be an order-of-magnitude more effective than ground-based technology because in

space, solar insulation is continuous and un-attenuated by the atmosphere. These potential advantages havemotivated efforts to design space solar power systems since the early 1960s. Reference 2 gives a timely andthorough review of previously proposed designs. In the early 1960 NASA, ATT & Bell made a huge inflata-

ble balloon communication satellite called ECHO, we took this concept to new construction for absorbingsolar energy and integrated communication for delivery of both signals & energy, we call it the Power Star TM.

BENEFICIAL NEW BUSINESS APPRAOCH

Our new business approach to providing NASAs Relays to Mars, different than the Big 5 Aerospace Corpo-rations (Boeing, Honeywell, Northrop-Grumman, Lockheed-Martin, Raytheon) approach we integrate theBest of all Worlds for Fast, Economic advantages & delivery. This Power Star TM System is Patent Pendingto use the best of old and the best of new with mass produced satellite integrated systems for its own powerand rectenna system. The Power Star TM system shall deliver at and above all data acquisition volumes andspeeds listed in the RFI by communizing & modularization for expansion growth guaranteed. Meet and ex-ceed all frequencies & bandwidths with the addition of modular and multi-tiered, parallel tunable channelslisted in the RFI.Best of all, most (80%) financials will be commercially funded by private investment firms & Venture Capi-talists. We deliver all relay requirements and have added commercial telecomm services additional service isthat it is an Energy from Space delivery system. The PowerStar delivery system can transmit energy directlyto the space transport vehicle going to Mars. Understanding this allows NASA a new option on getting toMars or even further, just think of having >25K times more power than the sun being provided to the electric

propulsion can get you there faster & cheaper.Proposed System Overview:

1. Complete Full System Simulation in 3D with CAD, CAE, CFD, Thermals, E3, etc2. Common Satellite Design & Mass Production concept similar to NASAs proven ECHO 3. 5 Satellite Production; 4 in NEO & 1 Command & Control Sat in Outer GEO4. All 5 are common Multi-Purpose for Data, Communications Relay & Power Transfer Sat.

5. Full Launch Capability details are explained below

Benefits to NASA & taxpayers:

50% Faster & cheaper development with improved reliability in Satellite Production Additional Capabilities for added Relay Data & Power aids Future Missions & Growth

Integrating the Best Teams, Technologies & Investors to Lower NASAs Burden

Only 20% Commitment funding by NASA full ROI within 7 years & Profit Making then after

Team Consists of Program Team, Technical & Private Investors, new LLC prior to award.

The proposed financial approach is to benefit NASA & the US taxpayers with as much value and less burdenas feasible. We will be using Private Investors and Venture Capitalist for 80% program costs and NASA tofund 20% with a 7 year complete ROI, providing profit from onward for decades. The Team will have a newLLC prior to program award includes our Program Leadership, investors have an on board Audit & PayoutMgr, we will hire a NASA Mgr which gained their PHD under Dr. David Hyland 4 decades as Professor. Thefull manufacturing will be from one of the Big 5 based on a bid package from them or maybe from foreign


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allies. We discourage the big 5 leading us because over 3 decades experience working at them has showntheir Corporate Mgt. can drain a program of Time & Money with little value to show for it (examples; FCS,JSF, A12, NASP, SDI, etc).

Team Mates on Program:

Boeings Robert Friend & their Space Electric Propu lsion System which can gain power from us.

Northrops new business Director Alberto Conti and their open minded ap proach.

Dr. Paul Werbos (NSF), Dr. John Mankins (NASA), Dr. Neha Satak (Uof FL), Dr. Gary Barnhard(UN) & other which provide High Value Leverage Asset(s) or Best bang for the buck.

Guy Webster 818-354-6278, Jet Propulsion Laboratory, [email protected]

Nancy Neal Jones 301-286-0039, Goddard Space Flight Center, [email protected]

The ROM estimated costs are listed below:Financial ROM Proposed Approach:

1. Full Common Satellite System $100M x 5 Satellites Avg.: $500M

2. Launch/Delivery & Integration Systems $100M each: $500M3. Complete MARS Power Star TM System w/developments: $1 Billion

4. Private Investors & Venture Capitalists Funds: $800M (80%)

5. NASAs Commitment with a 7 Year ROI & then Profit: $200M (20%)

Program Schedule/Timeline: 42 Months + 10 Month Launches

a) Phase 1 Prelim Design & Development Completionwith 20% Funds @ startup 12 Months (20% Funds)

b) Phase 2 Eng. Tooling & Mfg Development with 30% Funds: 18 Months (30% Funds)

c) Phase 3 Full Scale Production, Test & Eval. & Cert.: 12 Months (25% Funds)

d) Phase 4 Launch/Delivery Schedule every 2 months: 10 Months (25% Funds)

Full Delivery System Acquisition in Service by: 52 Months (less than 4 Years)

Standard Chart for Government ROI, Ref: NSF (shown for Ref Only)
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Commercial Relay Services, for Odyssey & MRO Relay plus growth capability:

Our Relay Services are suggested a 20% deposit from NASA with full ROI within 7 years & profitthere-after. Reason for deposit/partial investment is because, political change & budget uncertaintymakes NASAs interests to be served properly and proving to investors tangibility (skin in the game).

Mars Landers and rovers are highly constrained in mass, volume, and power an existing significantlimitation is in the data rates and data volumes that can be communicated on the direct link betweenthe Mars surface spacecraft and Earth.

At large Earth- Mars distances, the Curiosity rovers X -band direct-to-Earth (DTE) link operates at da-ta rates of less than 500 bps when communicating to a Deep Space Network 34m antenna.

Mars Exploration Program (MEP) has employed a strategy of including a proximity-link telecommu-nication relay payload on each of its Mars science orbiters. Currently, operating in the UHF band(390 450 MHz).

Curiosity rover operating at rates of up to 2 Mb/s, and data volumes average over 500 Mb/sol (or mar-tian day). Similar relay support has been provided to the prior Spirit rover and Phoenix Lander mis-sions.

Lunar Atmosphere and Dust Environment Explorer (LADEE) spacecraft at the Moon to Earth, withdownload rates of 622 megabits per second (Mbps).

We shall demonstrated an error-free data upload rate of 20 Mbps transmitted from the primary groundstation in New Mexico to the spacecraft orbiting the Moon.

Although No NASA Mars science orbiters are currently manifested beyond MAVEN, we shall bemodular & flexible enough to allow growth into these upgrades.

We shall accommodate the Mars Atmosphere and Volatile Evolution (MAVEN) spacecraft, RelayRadio on Mars.

This radio hardware, the Electra UHF Transceiver on NASA's MAVEN mission to Mars, is designed to provide communication relay supportfor robots on the surface of Mars.

Synergies with other commercial Mars interests for cost-sharing :As described, in the following PowerStar system description the 80/20 cost sharing and profit sharing will

be further understood. Please read the following:


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Power Star TM: APPROACH TO: MARS Advanced relay system

David C. Hyland *, Shawn Boike & AIC Team

This new business model is response to NASAs Request For Information (RFI) for how they might sustainMars relay infrastructure, consisting of orbiters capable of providing standardized telecommunication ser-

vices for rovers and landers on the martian surface, in the martian atmosphere, or in Mars orbit. Usingmass production techniques for satellite structure, power & antenna relays. Space Solar Power also knownas: Energy From Space refers to the concept of a space system that collects solar power via photovoltaicsand transmits it to ground collection stations using visible or microwave radiation. Previous system designsdeveloped over the past several decades entail gigantic structures with many moving parts and require on-orbit infrastructure and in-space construction. Here we combine very new and very old technologies toform a design that has no moving parts, requires no in-space construction and can be packaged in manyexisting launch vehicle payload fairings.

INTRODUCTION

This is our Industria l & Universities response to NASAs Request For Information (RFI) for how they mightsustain Mars relay infrastructure, consisting of orbiters capable of providing standardized telecommunication

services for rovers and landers on the martian surface, in the martian atmosphere, or in Mars orbit. Usingmass production techniques for satellite structure, power & antenna relays. The abundant supply of fusion- based energy produced by the sun remains to be efficiently harvested. The collection of solar radiation inspace could potentially be an order-of-magnitude more effective than ground-based technology because inspace, solar insolation is continuous and un-attenuated by the atmosphere. These potential advantages havemotivated efforts to design space solar power systems since the early 1960s. Reference 2 gives a timely andthorough review of previously proposed designs.

A solar power system consists of a space segment that collects solar energy, converts the energy into radiation(typically in a wavelength band to which the atmosphere is mostly transparent), then transmits the radiation toa ground facility that converts the radiation into electrical power. Since the ground-based power collectiontechnology is well developed, we concentrate here on the space segment, called the Solar Power Satellite(s)(SPS) . Moreover, the method of solar energy collection assumed here is photovoltaic, and the power transmis-sion to the ground is chosen to be microwave radiation with wavelengths near 10cm.Within the above restrictions, there are a wide variety of SPS design concepts. All previous approaches forSPS in this category involve very large, articulated structures, that must be assembled (in most cases robot-ically) in space and require many launches of the component parts into orbit (typically geostationary orbit) 1,2.These characteristics necessitate very large initial investments and technology developments to field an opera-tional system. An example for comparison that is fairly representative of previous concepts is the Naval Re-search Lab, 5MW SSP design 1. Figure 1 shows a summary of this concept. We choose this for later compari-son because it resulted from a quantitatively complete engineering design as well as a financial analysis. Ascan be seen from the Figure, this involves two 18,300 square meter solar arrays and a one kilometer diametermicrowave antenna. Rotating relay mirrors direct energy into the solar arrays, while the remainder of thestructure is nadir pointing. The study assumed an end-to-end efficiency of ten percent, and sought a FirstRevenue Unit design that could transmit 5 Megawatts of power. Typically, this type of design cannot belaunched by a single vehicle, but must be assembled on-orbit by either human or robotic agents.

* Professor of Aerospace Engineering, Aerospace Engineering Department, Texas A&M University, TAMU 3141, Collage StationTexas 77843.

NASAs Mars Relay RFI - Response


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Figure 1. Summary of the characteristics fo the NRL 5MW First Revenue Unit design

A significant improvement over previous efforts is the SPS-ALPHA (Solar Power Satellite via ArbitrarilyLarge Phased Array) 2. The main structure of SPS-ALPHA does not have to be slewed to follow the sun di-rection. The system is highly modular and good use is made of retro-directive phased array technology. Asandwich design combines the solar arrays with the microwave transmitters such that high voltage, centralized

power distribution is avoided. On the other hand, there are perhaps thousands of rotating mirrors used to redi-rect reflected sunlight onto the solar array, and the solar radiation redirection functions and photovoltaic ra-diation functions are segregated into different, very large structures. The very large structure cannot belaunched except through many launch vehicles, and the system must be assembled on-orbit via elaborate in-frastructure, including advanced robotic technology. Thus, although a significant step forward, the conceptinterposes the obstacle of a huge initial investment to achieve a first revenue system.The design concept discussed here carries modularity and multiple functionality several steps further. Theconcept combines a technology that is so new it is often overlooked with a technology that is so old it is al-most forgotten. The new technology is the printing (via photolithography, ink-jet processes, etc.) of solar cellsinterspersed with microwave patch antennas on thin, flexible sheets (Mylar, Kapton, paper, fabric, etc.). The

printed sheets are produced in mass quantities. The old technology is that of the Echo satellites. Large, thinsheets are assembled into a spherical balloon. For launch, the sphere is compactly packaged in a small con-tainer that fits into the launch vehicle payload faring. Once on orbit a volatile material is made to sublimate to

provide the gas pressure for initial inflation. Metallic layers within the printed sheets are forced into yield to provide rigidification and the Power Star TM sphere is then evacuated. Electromagnetic propagation theoryshows us that a completely decentralized control algorithm allows us to coordinate the numerous (printed)microwave antennas to transmit multiple beams to any desired ground-based power collection locations. Thesystem is a single, very simple structure and no slewing or mechanical motion is required. Further, the powerdistribution technique involves power transmission within the skin only over distances of a few centime-ters. Thus power transference is localized and requires neither complex and high voltage power distributionand management systems nor large power-conducting wires.


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The following sections describe these features in some detail and we substantiate that the system has no mov-ing parts, requires no slewing or rotating elements, can be deployed from a single launch vehicle, is extremelyrobust to component failures and is composed of material that can be manufactured in great quantity.

THE NEW: SOLAR-MICROWAVE FABRIC TM

The very new and rapidly advancing element of Power Star TM technology is illustrated in Figure 2. Largescale production of inexpensive solar arrays is well underway. Printed microwave antennas are also well

known and are being advanced at a rapid rate for numerous communication applications. Solar-MicrowaveFabric TM combines these two components on the surface of the same flexible substrate. The lower part of Fig-ure 2 illustrates a typical cross-section. The solar cells and patch antennas are interspersed (without overlap-

ping) with a randomized tessellation in order to eliminate grating lobes. This pattern is printed on what is to become the exterior surface of the substrate sheet or skin . In the full system, there may also be an arraycomposed solely of microwave transceivers (dual transmitters and receivers) printed on the opposite surface(due to become the interior surface of the sphere). Patch antennas on the exterior surface draw power fromhalf of the immediately adjacent solar cells (a few centimeters distance) or from the interior transceivers,through the thickness of the skin. Details of power transfer are described in the Intra-Satellite Power Distri-bution sub-section below. Besides the short power leads there is a grid of conducting wires for electricalground and for rigidizing the sphere prior to evacuation. In this section we discuss printed solar cells, printedmicrowave antennas and choice of the substrate material.

Printed Solar Cells

Presently, there is a range of solar cell printing technologies, where rapid manufacturability is traded offagainst cell efficiency. A notable example is that reported in Reference 3. The Victorian Organic Solar CellConsortium has demonstrated the capability to produce printed solar arrays at speeds of up to ten meters perminute, or one cell every two seconds. Up to 30cm wide, these cells produce 10-15 watts of power per squaremeter per square meter under maximum ground insolation. Subtrates include paper-thin flexible plastic orsteel. As illustrated in Figure 3, the cells combine various organic materials to capture power from different

parts of the solar spectrum.

Figure 2. Illustration of the basic concept of the Solar-Microwave Fabric TM .

Substrate layer

Transmitter

Solar cell Solar cell

Conductive coatingPowerconnectors

Printed Solar Ar-

Printed Patch An-

Solar-Microwave Fab-

The


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Figure 3. Composition of the Victorian Organic Solar Cell Consortium.

In comparison, MIT solar cells 4 use an ink-jet process to print cells on paper. Efficiency for most designs is presently 1% to 2%. However, 4% is a near-term goal and it is quite reasonable to anticipate 5 to 10% in thefuture. As a baseline we can say that 2% efficiency with rapid fabrication ability is the current capability.

Printed Microwave Patch Antennas

Antennas can be inkjet printed onto many flexible materials, even including cotton-polyester Multiple print-ing layers can be used to increase efficiency. As illustrated in Figure 4, a microwave patch antenna consists ofa metal patch mounted on a grounded, dielectric sub strate.

Figure 4. The basic configuration of a microwave patch antenna.

The dielectric provides a resonant cavity to amplify the transmitted signal. Since L is the resonant dimension,we must have:

2 L (1)

Where is the operating wavelength. W is usually chosen as 1.5 L to get higher bandwidth, but we shall as-sume 2W L here. The practical printing resolution is 15 microns and is quite sufficient to satisfy Equa-tion (1) to sufficient accuracy. Table 1 shows a survey of performance statistics for existing patch antennas 5.Efficiencies of up to 79% are presently attainable.

Table 1. Performance characteristics of various printed patch antennas.


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Substrate Heightin mm

BW = Bandwidth

Etched patch onFR45 substrate

Inkjet Patch(two layers ofink) glued on

FR45 substrate

Inkjet Patch(one layer ofink) on felt

Inkjet Patch(two layers of

ink) on felt

Patch size(mm) 37.4 x 28.1 37.4 x 28.1 47.7 x 36.9 47.7 x 36.9

Substrate height 1.6 1.6 1.9 1.9

Frequency(GHz)

2.378 2.480 2.405 2.505

SII (dB) -13.39 -14.89 -10.05 -9.95

10 dB BW(MHz)

22.5 24.5 17.5 N/A

Directivity (dBi) 7.39 7.55 8.38 8.72

Gain (dBi) 6.37 5.09 4.02 5.98

Efficiency (%) 79 57 37 53

Substrate Material

Although solar cells and patch antennas have been printed on a wide variety of materials, we have focused ontwo materials that have the closest connection to Echo satellite technology. The foremost, and the one withthe most heritage, is Mylar, a polyester film made from resin Polyethylene Terephthalate (PET). This materialretains its full mechanical capabilities at temperatures ranging from -70 C to 150 0C. Its melting point is 2540C. Its volumetric density is 1390 kg/m 3. An attractive alternative is Kapton, an organic polymeric materialthat, effectively does not melt or burn and functions well at temperatures ranging from -269 C to 400 0C. At1420 kg/m 3, its volumetric density is slightly larger that that of Mylar. Continuing studies will explore print-compatible materials with adequate tear resistance and minimum density.

THE OLD: ECHO SATELLITE TECHNOLOGY

Sheets of the multi-functional fabric described in the previous section are cut into gores (sectors of a sphere)and the several gores are assembled to form a spherical balloon (once inflated). Beyond this point, the PowerStar TM system makes full use of Echo satellite technology.

Project Echo 6 was the first passive communications satellite experiment. Each of the two satellites were de-signed as a passive reflector of microwave signal, and each was a metalized PET film balloon satellite. Soonafter the launch vehicle failure of Echo 1 in 1960, the 30.5m diameter Echo 1A was successfully placed inorbit by a Thor-Delta vehicle in the same year. It reentered Earth's atmosphere, burning up on May 24,1968. Following successful operation of Echo 1A, on January 25, 1964, the 41.1m diameter Echo 2 was suc-cessfully deployed on orbit. Echo 2 reentered Earth's atmosphere and burned up on June 7, 1969.


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Figure 5. Various aspects of Echo satellite technology: (a) Echo 1A stowage canister; (b) Canister closed; (c) Fold-

ed sub-scale prototype; (d) Inflated sub-scale prototype; (e) Echo 2 during inflation testing.

Figure 5 shows various aspects of the Echo technology. The satellites were made of 12.7 m thick biaxiallyoriented PET (Mylar) film, coated with a vapor deposited 0.2 m layer of aluminum (to provide RFreflectivity). Special folding techniques were devised to minimize the stowed volume (see Figure 5(c) and5(d)). This is an important feature since finite material strength sets a lower limit to radii of curvature in

bending so that any fold of a thin sheet introduces voids that reduce packing efficiency. The folded balloonsfor both spacecraft could be stowed for launch in small spherical canisters (See Figure 5(a) and 5(b)). In

particular, the 30.5m diameter inflated sphere of Echo 1A was stowed in a 0.71m seamed spherical canister.

Once on-orbit, the small canisters were opened and the balloons were inflated to form two of the largest andvisible artificial satellites ever created (Figure 5(e)). At launch, the Echo 1A balloon mass was 71.212 kg

which included 15.12 kg of sublimating powders of two types7: anthraquinone, and benzoic acid. These

coated the interior surface of Echo 1A, and sublimated once the balloon was exposed to the sun. On orbit,only several pounds of gas pressure were all that was required to inflate the sphere and maintain its shape.

Echo 2 used a refined inflation system to improve the balloon's smoothness and sphericity. In thiscase, a number of pillows containing sublimating powder were stor ed flattened against the interiorsurface of the balloon. See Figure 6 for the pillow inflation process. Once exposed to the heat fromthe sun, the pillows inflate, and vent gas through perforations in their surface, thereby inflating the rest of thesatellite. This deployment process prevents the gas from getting trapped in pockets and producing deleteri-ous stress concentrations. In the Power Star TM a copper grid (for electrical ground) is embedded in the skin.This is designed to yield at the inflation pressure. Like the aluminum coating in the Echo 2 satellite, the yield-ed grid provides rigidification of the structure, eliminating the need to sustain gas pressure. One of the pillowsis designed to rupture the outer surface of the balloon after deployment, allowing the Power Star TM to releaseexcess gas once the copper grid has just begun to yield. Once fully deployed the balloon is an evacuated shell.See Reference 8 for on-orbit video of Echo 2 inflation.

(b)(a)

(c) (d) (e)


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Figure 6. Echo 2 inf lation system pillows. Top: stowed configuration; Bottom: Pillow out -gassing from its per-forations.

SYSTEM COORDINATION AND OPERATIONIn this section we describe how the old and new elements of the Power Star TM system are combined and coor-dinated to work together once the system is launched and deployed. Figure 7 sketches the overall compositionand method of operation.

Figure 7. Overall Power Star TM operation once deployed

The exterior surface of the sphere is printed with solar cells and microwave transmitters (Figure 7, lowerright), where the placement of transmitters is somewhat randomized to prevent grating lobes (see below).There are power connectors between each transmitter and a subset of the immediately adjoining solar cells(Figure 7, top, center, red lines in the cross-section). Beneath the exterior coating is the substrate layer (gray

band in the Figure) with an embedded copper grid (orange lines in the Figure) for electrical ground andrigidification. The interior surface of the substrate is coated solely with transceivers (transmitter/receivers,

blue layer on the bottom of the cross-section). There are power connections through the thickness of the skinfrom the internal transceivers and the immediately proximate external transmitters. Power connections in the


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skin are very short (a few centimeters) and the power collection and transmission devices are on a microscop-ic scale, such that we anticipate an eventual halving of the Echo skin thickness to ~6 m.

Power is received at several locations on the ground by arrays of rectifying antennas ( rectennas ). At the loca-tion of each rectenna, a low-power microwave beacon is placed. At each patch antenna a local microprocessorrecords the beacon radiation that the patch receives; records the radiation wave form; amplifies the waveformand emits it back in reverse time (or, equivalently with conjugate phase). As will be elaborated below, thiscompletely decentralized transmitter control scheme produces transmitted radiation that, given the size and

shape of the Power Star, optimally matches desired power distribution on the ground. Note that the systemcan absorb power from the sun and transmit power in any other direction without the need for slewing or me-chanical motions. The system works with electrons and photons and has no moving parts.

In the next two subheadings, we discuss further details of the power transmission control, and the specific processes for transferring collected solar power within and across the sphere.

Power Beam Control

It is a rigorous result in electromagnetic propagation that the beacon-based control that coordinates the nu-merous transmitters as described above optimally approximates, in a mean-square sense, the desired powerdistribution on the ground. This power delivery scheme is a generalization of retro-directive beam technologyand has been applied to many areas. For example Reference 9 discussed its application to acoustics for medi-cal technology.

Indeed, the ground distribution actually produced is the spatial convolution of the desired distribution (as set by some pattern of beacons) with the Power Star aperture point-spread function (PSF), which is essentiallythe tightest, most concentrated beam that the total configuration of transmitters can produce. This PSF func-tion depends on the size, shape and distribution of the transmitters on the external surface of the power star.Thus, if the beacons can be approximated by point sources, then the ground distribution consists of severalPSF spots, each centered at one of t he beacon locations.

Recording the beacon signals, then amplifying them and playing them back in reverse time occur concurrent-ly. To simplify the explanation, we illustrate these steps separately. First, consider the beacon propagation,illustrated in Figure 8 by means of a simple two-dimensional wave propagation simulator. Here there are threeapproximately point sources (that is, a single pixel in extent) unevenly distributed along the vertical line to theleft, representing the ground plane. The circular region to the right represents the Power Star sphere. In part(a), radiation commences with a widening interference pattern. Then (part (b)), each pixel on the circumfer-ence of the circle records the time signal of the field amplitude measured at its location. Figure 9 shows whathappens when each pixel (representing a single patch transmitter) transmits the signal it recorded in reversetime. In part (a), note the converging wave fronts of the initial field amplitude. In part (b) Of the Figure, wesee three concentrated spots of intensity, centered at the beacon locations. These spots represent the PSF dis-tributions and are broader than the beacons. The broader width of the ground plane spots is mainly propor-tional to the overall size of the Power Star. The results also illustrate that, despite the usual assumption that

phased arrays are planar, the accuracy with which a desired ground distribution is duplicated is mostly de- pendent on size, not on shape. This spherical phased arrays work well.

Intra-Satellite Power Distribution

Since the directions of the sun and the beacons are not coincident, a mechanism for distributing power within

the satellite is needed. Figure 10 shows the geometry of irradiation from the sun and the beacons, where weassume that the angular separation of beacons is small so that a single, representative beacon direction may beconsidered. The quantity is the angle between the sun direction and the beacon direction. Recall that theinterior surface of the sphere is coated with transceivers operating at a higher frequency (to reduce diffractioneffects). These transceivers are to be oriented so that the resonant axes of each diametrically opposite pair are

parallel.

As illustrated in Figure 10, the surface of the sphere is divided into four sectors: The sector exposed to bothsunlight and beacon radiation (denoted by ,S B ); that receiving beacon radiation but no sunlight ( ,S B ); thatexposed to sunlight but not beacon ( ,S B ), and the region where neither sun nor beacon are visible ( ,S B ).


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Clearly, sectors ( ,S B), and ( ,S B ) are mirror images, such that each point on ( ,S B ) has a diametrically oppo-site point on ( ,S B ), and vice-versa . The same remark pertains to ( ,S B ), and ( ,S B ). The sector that a particu-lar transmitter and its adjacent solar cells are located is indicated by their output signals. Given this infor-mation, the power supply

(a)

(b)

Figure 8. Initial Propagation of beacon radiation. (a) Radiation commences, (b) Circular phased array recordsbeacon information.


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(a)

(b)

Figure 9. Phased array propagates amplified beacon in reverse time. (a) Transmission commences, (b) Three con-centrated spots, centered at the beacons appear on the ground plane.


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Figure 10. Geometry of the power distribution system. Angle denotes the angle between the directions to thesun and a beacon.

algorithm is indicated in Table 2. Note that no processing is needed for this algorithm. In essence, thetransmitters that need to be active because they receive a beacon signal are powered by either the proximatesolar cells or by the proximate internal transceivers, whichever is actually producing power. No beacon signalmeans the transmitter is blocked. Each transmitting antenna draws power from the solar cells in its immediatevicinity (within a few centimeters), or through the thickness of the skin. Each transmitter receives just a fewWatts, so there are no high voltages or large wires. This localized architecture means robustness against par-tial damage.

Table2. Power transfer algorithm

Sector Power Transfer

,S B External surface transmitter draws power from the adjacent solar cells

,S B Solar cells transfer power through the skin to their immediately proximateinternal surface transceivers. The internal transceivers emit power beams

through the center of the sphere to fall on the internal transceivers in sector

,S B .

,S B Internal transceivers transfer received power through the skin to their

immediately proximate external surface transmitters

,S B No action taken.

PERFORMANCE CHARACTERIZATION

Having described the basic design of the satellite, we next consider the analysis of its performance character-istics, viz. power transmitted to the ground, beam width, etc., under separate subheadings.


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Power Transmitted

To begin, a geometrically regular arrangement of the patch antennas on the exterior surface would produce anaperture PSF having, besides a main concentrated spot ( the central lobe ), several regularly spaced offsetspots (the grating lobes ). This tends to have a disastrous effect on the accuracy with which a desired powerdistribution may be approximated, since the actually produced distribution is the convolution of the desireddistribution and the PSF. However, a slight randomization of the transmitter antenna placements (that retainsthe same average number of antennas per unit area) suffices to disburse the grating lobes so that the central

lobe alone remains the only power concentration in the emitted radiation. In this case, the main lobe is propor-tional to the characteristic function (the Fourier transform) of the probability density function of patch antennalocations. For example, if the locations of all patch antennas are statistically independent Gaussian distribu-tions, then the angular distribution of radiated power, P produced by the entire phased array is:

22

2

2

1 2exp

2 2 A A D D s P s

(2)

where is the unit vector from the phased array to a point of observation, and where:

operating wavelength

balloon diameter average distance between the centers

of neighboring patch antennas

2

A D s

L W

(3.a-d)

The last equation repeats the assumption made in the remarks under Figure 4 that the patch antennas areroughly squares that are half a wavelength on a side. Note that the maximum value of 2 s has to be unity;in which case, patch antennas cover the entire exterior surface of the balloon, leaving no room forthe solar arrays. Thus the phased array must be sparse , and of necessity 2 s .

Equation (2) is a reasonable approximation for many different antenna position probability distributions. Thenaccordingly, the total power transmitted to the ground from the central lobe is:

2

2 2

Total power input from the solar arrays

and internal transceivers

t sa

sa

P d P P s

P

(4.a,b)

Note that the factor appropriately reflects the sparse aperture theorem.

For a given s, the fraction of the frontal area occupied by the solar arrays is 21 2 s , therefore:2

2

2

12 4

aggregate efficiency of solar arrays

and patch antennas

Solar insolation 1367W

sa A eff s

eff

S

P D Q s

Q m

(5.a-c)

The aggregate efficiency, eff is a function of both the solar array and transmitter efficiencies and the

beacon-sun angle, . Assuming roughly the same efficiencies for the exterior and interior transmit-ters, we have:


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2 212 1 1 cos

0,

Solar array efficiency

Transmitter efficiency

eff T S T T

S

T

(6.a-d)

Combining Equations (4)-(6), we have in summary:

2 22

2 212

12 2 4

1 1 cos

t A eff s

eff T S T T

P D Q s s

(7.a,b)

From this relation, it is clear that the optimal average spacing of the transmitters is 2optimal s .This means that the surface area of the balloon is equally divided between the solar cells and trans-mitters. Also, the total power to the ground becomes:

2max

2 212

14 4

1 1 cos

t A eff s

eff T S T T

P D Q

(8.a,b)

Note that the factor 14 arises from the sparseness of the array.

From Equation (8), we see that if the satellite is at geostationary altitude with the ground station beneath, the power transmitted rises to a maximum at midnight ( 0180 ) and declines to a minimum at noon ( 0 ).This conforms to the daily electrical power usage profile for street lighting of typical municipalities.

To see what Equation (8) predicts for power transmitted to the ground given current device capabilities, we let2%S and, consulting Table 1, set 79%T . Figure 11 shows the ranges of transmitted power

(over all sun-beacon angles) for S equal to 2% (current


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Figure 11. Power transmitted as a function of balloon diameter for various values of the solar cell efficiency.

capability), and for 5%, 10%, and 25% , representing different stages of development, all as func-tions of the balloon diameter. We see that even with the presently lowly capabilities of printed solarcells, a one kilometer balloon can deliver from 3 to 4 Megawatts comparable to the design of Fig-ure 1. Moreover, efficiency of ~4% is expected soon, in which case, a 1 km system gives ~6 to 10

MW. Printed cell technology is still in an early stage of development wherein cheap manufacturabil-ity is paramount over cell efficiency. But one can expect a progression toward the efficiency levelsof pres ently one -off laboratory devices, where 25% is typical. In this case the 1km balloon might

be capable of 30 to 50MW. Compare this with the system of Reference 1.

Minimum Beam Width (Rectenna Size)

Common to all SPS concepts is the minimum beam width on the ground expressed, by use of Ray-leighs angular resolution formula, as a function of wavelength, distance and transmitting aperturediameter. In the present case this is modified slightly because the aperture is sparse, not filled. In ac-cordance with the sparse aperture theorem (see also Equation (2)), the width of the central beam in

the system PSF is diminished by the factor of 2 . Therefore the minimum width of the power con-centration spot that can be put on the ground, x , is given by:

1.1 2

transmit distance

(35,786 km for GEO)

A

z x

D

z

(9.a,b)


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This sets the size of the rectenna. Assuming a geostationary orbit, Figure (12) shows the rectennadiameter as a function of balloon diameter for various values of the operating wavelength. It is seenthat we have inordinately large (~5.5km for a 1km balloon) rectenna sizes for the nominal wave-length of 10cm. This is not a problem peculiar to the power star. Indeed, the PowerStar beam widthis smaller than any filled aperture concept. There are three principal avenues. The first is to decreasethe operating wavelength to, maybe, 1cm thereby reducing the rectenna size to hundreds of metersinstead of kilometers. The second is to increase the aperture size to several kilometers. The third is to

reduce the transmit distance. This would entail a constellation of power collector Power Stars in sun-synchronous, lower orbits (~2000km) complemented by several relay satellites in lower inclination,MEO orbits that take turns beaming power continuously to the rectennas. These tradeoffs must beexamined for any design concept and this effort is underway for the Power Star.

Figure 12. Rectenna diameter (minimum spot size) as a function of balloon diameter for various values of the op-erating wavelength.

Packaging for Launch

The Power Star is to be folded compactly into a canister that can be accommodated in existing launch vehicle payload fairings. We assume here that the stowed configuration is a sphere of diameter S D . If w denotes thethickness of the skin, the total volume occupied by just the skin of the deployed balloon is 2 A D w . The small-est stowed diameter is obtained when this volume is equal to 3 6S D . However as remarked above, a thinmembrane folded many times has an external volume much in excess of just the volume of the material ofwhich it is composed. Thus we characterize the folding system by the packing efficiency, 1eff p , so that

3 26S A eff D D wp , or:


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1 326

Skin thickness

Packing efficiency 1

S eff A

eff

D p wD

w

p (10.a-c)

The packing efficiency is difficult to calculate and depends upon the precise geometry of the folds, the skin

thickness, and the material properties. However, we shall take the Echo satellite characteristics as our guide.Table 2 lists the Echo dimensions and the values of packing efficiency. These are quite close, so in the follow-ing we use simply 3.0eff p .

Table 2. Dimensions and packing efficiencies for the Echo satellites.

Satellite , mAD , mSD , m w eff p

Echo 1 30.5 0.71 20.3* 3.16

Echo 2 41.1 1.04 36.0** 3.08

*Includes Mylar, metallic coating and sublimating power coating.**Includes Mylar, metallic coating and average thickness due to pillows

Based on this value, Figure 13, shows the launch canister diameter as a function of the inflated balloon diame-ter. Evidently, a one kilometer Power Star, the same size as the FRS design microwave antenna of Figure 1,can be accommodated in several existing heavy-lift launch vehicles. In particular: the Delta Heavy (5.1 m di-ameter fairing), the Ariane 5 (5.4m), the Minataur VI (5.71 m) & the new SpaceX Falcon 9 (Custom Fairing).

Aerodynamic Drag and Orbit Lifetime

As is the case with the Echo satellites, Power Star would have a very low ballistic coefficient so that aerody-namic effects can set limits on orbit altitude such that orbit lifetime is more than a few decades. To analyzethis situation, we assume an initially circular orbit. For lifetimes greater than 10 years, the lifetime as a func-tion of the initial orbit radius of a circular orbit is nearly independent of the launch time relative to the solarmaxima or minima (see Reference 10). Thus the orbit lifetime can be estimated using the average atmosphericdensity as a function of altitude, as given by the U.S. Standard Atmosphere. Further we may assume smalldrag forces such that the decaying orbit takes the form of a tight spiral with a slow ly varying instantaneousorbit radius.


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Figure 13. Stowed diameter as a function of the inflated balloon diameter.

Then we have approximately that the orbit life time, f t , is:

Initial orbit radius

Radius of the Earth

Gravitational constant (GM) of the earth

Atmospheric density at orbit radius

1 E

f

i

E

Atm

Atm

ia

Rt

a

a

R

a a

daa

(11.a-e)

where is the ballistic coefficient:

2

2

4

Power star mass

Frontal area

Volumetric density of the skin

D

A skin

A

skin

M C A

M D w

A D

(12.a-c)

Thus, for the Power Star and assuming free molecular flow ( 2 DC ), we get:

2 skinw (13)


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Figure 14. Orbit lifetime as a function of initial orbit altitude.

which is just twice the areal density of the skin and independent of diameter. To get a conservative estimateof orbit life, we assume the smallest practicable thickness, 0.006w mm . Then we have:

2 2 1390 0.006 0.01668 skinw mm (14)With the above assumptions, Equations (11) give a conservative estimate of the orbit lifetime as a function ofinitial altitude as shown in Figure 14. The results indicate that a long lifetime is ensured by placing the PowerStar at roughly 2000km or above. Thus a MEO orbit or above is suitable for a long-term system. Note that the

de-orbit time function is independent of the diameter of the system and directly proportional to , which isapproximately twice the areal density of the skin. Hence results for larger skin thicknesses can be obtainedfrom the Figure by multiplying the ordinate by the ratio of new to old thicknesses.

CONCLUSION

In this paper we have proposed a novel design concept for a Space Solar Power Satellite the Power Star TM.With heritage dating back to Project Echo, this system is an inflatable balloon made of a thin, flexible skinwhereupon solar cells, and microwave patch antennas are printed via the most modern mass production tech-nology. Power Star TM operates with no moving parts and with no slewing or other mechanical motion. Atleast up to 1km diameter, it requires no on-orbit manufacturing or construction. Advanced adaptive phasedarray technique and insights from time-reversed acoustics, combined with low-amplitude beacons yield a

beam forming control algorithm that is entirely local to each patch antenna. The operation of the phased arrayis decentralized and adaptive so that even if severely damaged, the system can retain some level of useful per-formance. Power is regulated within the balloon such that transmission through the skin occurs within a fewcentimeters at most, obviating the need for a centralized, high voltage power distribution system. The powersystem permits solar power to be gathered from any angle and power to be beamed in any direction (s) with-out slewing or structural deformation.

Preliminary performance calculations show that even with the low efficiencies of presently available printedsolar cells, a 1 km Power Star can produce enough power for a First Revenue System. Using Echo technolo-gy, a 1 km Power Star can be packed for launch in several existing heavy-lift vehicles. Finally, despite its low


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ballistic coefficient, the orbit lifetime is of the order of a century if the initial (circular orbit) altitude is greaterthan approximately 2000km.

NOTATION

A Frontal area of the Power Star

ia Initial orbit radius

DC Aerodynamic drag coefficient

A D Balloon diameter

S D Diameter of launch canister

L, W Resonant length, and width, respectively of the microwave patch antennas

M Mass of the Power Star

eff p Packing efficiency

sa P Total power input from the solar arrays and internal transceivers

t P Total power transmitted from the central lobe

S Q Solar insolation at 1 AU

E R Radius of the earth

s Average distance between the centers of neighboring patch antennas

f t Orbit lifetime

w Skin thickness

z Transmit distance

Ballistic coefficient

x Minimum beam width; also approximate diameter of the rectennaeff Aggregate efficiency of solar arrays and patch antennas

,S T Efficiencies of the solar cells and microwave patch antennas, respectively

Operating wavelength

Gravitational constant of the Earth

m micron

Atm a Atmospheric density at orbit radius a

skin Volumetric density of the skin

Angle between the sun and beacon directions

REFERENCES1 A.C. Charania, J.R. Olds, and d. Depasquale, Operational Demonstration of Space Solar Power (SSP) ; EconomicAnalysis of a First Revenue Satellite (FRS).http://www.nss.org:8080/settlement/ssp/library/Economic_Analysis_of_a_First_Revenue_Satellite_for_SSP_(2011).pdf . 2 J.C. Mankins, The Case for Space SolarPower . Virginia Edition Publishing LLC, Houston Texas, December 2013.
http://www.nss.org:8080/settlement/ssp/library/Economic_Analysis_of_a_First_Revenue_Satellite_for_SSP_(2011).pdfhttp://www.nss.org:8080/settlement/ssp/library/Economic_Analysis_of_a_First_Revenue_Satellite_for_SSP_(2011).pdfhttp://www.nss.org:8080/settlement/ssp/library/Economic_Analysis_of_a_First_Revenue_Satellite_for_SSP_(2011).pdf


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3http://newsroom.melbourne.edu/news/ctrlp-printing-australia%E2%80%99s-largestsolarcells?_ga=1.64104170.436929901.13963592184D. L.Chandler, David L. (2012). "While youre up, print me a solar cell - MIT News Office" . web.mit.edu . Retrieved 20February 2012.5http://ieeexplore.ieee.org/stamp/stamp.jsp?tp=&arnumber=66937346 "Echo 1, 1A, 2 Quicklook" . Mission and Spacecraft Library . NASA . Retrieved February 6, 2010.

7H. M. Jones; I. I. Shapiro; P. E. Zadunaisky (1961). "Solar Radiation Pressure Effects, Gas Leakage Rates, and AirDensities Inferred From the Orbit Of Echo I". In H. C. Van De Hulst, C. De Jager and A. F. Moore. Space Research II,

Proceedings of the Second International Space Science Symposium, Florence, April 10-14, 1961 (North-HollandPublishing Company-Amsterdam).

8 Echo II Satelloon Inflation, 1964. https://www.youtube.com/watch?v=qz3-b7sB9CA&noredirect=1

9 Mathias Fink, Time -Reversed Acoustics Scientific American, November 1999.

10 Larson, and Wertz, Space Mission Analysis and Design , 3 rd Edition, Figure 8.4.
http://web.mit.edu/newsoffice/2011/printable-solar-cells-0711.htmlhttp://web.mit.edu/newsoffice/2011/printable-solar-cells-0711.htmlhttp://web.mit.edu/newsoffice/2011/printable-solar-cells-0711.htmlhttp://web.mit.edu/newsoffice/2011/printable-solar-cells-0711.htmlhttp://samadhi.jpl.nasa.gov/msl/QuickLooks/echoQL.htmlhttp://samadhi.jpl.nasa.gov/msl/QuickLooks/echoQL.htmlhttp://samadhi.jpl.nasa.gov/msl/QuickLooks/echoQL.htmlhttp://en.wikipedia.org/wiki/Mission_and_Spacecraft_Libraryhttp://en.wikipedia.org/wiki/Mission_and_Spacecraft_Libraryhttp://en.wikipedia.org/wiki/Mission_and_Spacecraft_Libraryhttp://en.wikipedia.org/wiki/NASAhttp://en.wikipedia.org/wiki/NASAhttp://en.wikipedia.org/wiki/NASAhttps://www.youtube.com/watch?v=qz3-b7sB9CA&noredirect=1https://www.youtube.com/watch?v=qz3-b7sB9CA&noredirect=1https://www.youtube.com/watch?v=qz3-b7sB9CA&noredirect=1https://www.youtube.com/watch?v=qz3-b7sB9CA&noredirect=1http://en.wikipedia.org/wiki/NASAhttp://en.wikipedia.org/wiki/Mission_and_Spacecraft_Libraryhttp://samadhi.jpl.nasa.gov/msl/QuickLooks/echoQL.htmlhttp://web.mit.edu/newsoffice/2011/printable-solar-cells-0711.html

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