Concept Development and Evaluation Program SATELLITE POWER SYSTEM DOE/ER-0023 US Department of Energy and the National Aeronautics and Space Administration Reference System Report October 1978
Available from: National Technical Information Service (NTIS) U.S. Department of Commerce 5285 Port Royal Road Springfield, Virginia 22161 Price: Printed copy: $11.75 Microfiche: $3.00
DOE/ER-0023 Dist. Category UC-11,41,60,63, 63a,b,c,e,64,66e,95f,97c Satellite Power System Concept Development and Evaluation Program Reference System Report October 1978 (Published January 1979) U.S. Department of Energy Office of Energy Research Washington, D.C. 20545 and the National Aeronautics and Space Administration
digitized by the Space Studies Institute Research Library ssi.org
FOREWORD The possibility of generating large quantities of electrical power in space and transmitting it to earth using satellites was first suggested in 1968. During the following years, several studies of the concept were conducted by the National Aeronautics and Space Administration (NASA) and industry. The energy shortages of 1973 spurred interest in the concept and in early 1976, the Department of Energy (DOE) (then the Energy Research and Development Administration) and NASA initiated an SPS Concept Development and Evaluation Program. This evaluation program is guided by a joint DOE-NASA plan which covers a period from mid-1977 to mid-1980. The key program milestones which guide all substudies and program activities are: Reference System Definition October 1978 Preliminary Program Recommendations May 1979 Updated Program Recommendations January 1980 Final Program Recommendations June 1980 The joint plan states that the Reference System selection milestone "will focus the evaluation effort in what is considered to be at that time the optimal direction." It will particularly emphasize technical and operational information required by DOE to conduct environmental, socioeconomic and comparative studies. This report is submitted in response to the first major program milestone. It defines a Reference System Concept based on the system definition effort to date. The concept presented is considered to be in the proper "direction," but is not optimum at this time. The system definition studies have, however, indicated technical feasibility of the reference concept and the concept will continue to be analyzed and changed as the results of preceding systems definition and other studies warrant.
TABLE OF CONTENTS I. SUMMARY 1 II. INTRODUCTION 4 III. REFERENCE SYSTEM DESCRIPTION 10 A. Guidelines and Assumptions 10 B. System Overview 10 C. Solar Cells and Blankets 14 D. Solar Array and Structure 17 E. Power Distribution 19 F. Rotary Joint 22 G. Attitude Control System 22 H. Microwave Power Transmission System 25 I. Mass Statement 47 J. Space Transportation 47 K. Natural Resources 58 L. Operations 64 1. Construction 2. Commercial Operation M. Costs 76 IV. TECHNOLOGY SUMMARY 79 V. DOCUMENTATION SUMMARY 83 APPENDICES A. Systems Analysis Results A-l References B. Contracted System Definition Studies B-l 1. Rockwell International B-2 2. Boeing Aerospace Corporation B-6(
LIST OF FIGURES Figure Page 1 SPS Reference System Concept 3 2 Simplified Diagram of SPS Concept Development 5 and Evaluation Methodology 3 Solar Power Satellite 7 4 SPS Reference System - Satellite Configuration 12 5 SPS Efficiency Chain (GaAlAs CR2 and Si CR1) 13 6 SPS Operations 15 7 Reference System Characteristics 16 8 Solar Cell Options for SPS 18 9 Laser Annealing Concept 18 10 Solar Cell Blanket Support 20 11 SPS Power Distribution 21 12 Antenna Yoke and Turntable 23 13 Attitude Control System Characteristics 24 14 Transmitting Antenna Functional Description 26 15 Microwave Power Transmission System Parameters 27 16 Microwave Power Transmission System 31 17 Microwave Array Power Distribution 32 18 Power Density at Rectenna as a Function of 33 Distance from Boresight 19 Peak Power Density Levels as a Function of Range 35 from Rectenna 20 Grating Lobe Characteristics 36 21 Grating Lobe Maxima 38 22 Peak Power Density for Sidelobes and Grating 39 Lobe as a Function of Range from Rectenna 23 Near-Field Antenna Pattern 40 24 Rectenna Patterns and Power Levels 42 25 Microwave Transmission Efficiency 44 26 Noise Power Density of Ground for a 1 km, 5 GW 46 SPS Antenna
2J Heavy Lift Launch Vehicle 50 28 Launcher/Erector Concept 51 29 SPS Heavy Lift Launch Vehicle Trajectory and 52 Exhaust Products Data 30 Personnel Launch Vehicle 54 31 LOX/LH2 Common Stage POTV 55 32 Cargo Orbit Transfer Vehicle Gallium Option 56 33 SPS Construction and Commercial Operations 65 34 Component Packaging Characteristics 67 35 Typical Component Mixing 68 36 Construction Base Buildup for Silicon System 70 37 Scenario for Buildup of Construction Bases 71 38 Construction Timeline for Two 5 GW Satellites/Year 72 39 Scenario for Construction of Two 5 GW Satellites/Year 74 40 SPS Operational Functions 77 41 SPS Operations Management Concept 78
LIST OF TABLES Table Page 1 SPS Mass Statement - Millions of KG's 48 2 GaAlAs Independent Electric COTV Mass Breakdown 57 3 Si Independent Electric COTV Mass Breakdown 58 4 Materials List for Reference System 59 5 Total Program Materials List 62 6 SPS Fleet Sizes 75
List of Abbreviations AC ALTERNATING CURRENT ACS ATTITUDE CONTROL SYSTEM ACSS ATTITUDE CONTROL AND STATIONKEEPING SUBSYSTEM AMO AIR MASS ZERO APU AUXILIARY POWER UNIT Ar ARGON BOL BEGINNING OF LIFE CMG CONTROL MOMENT GYRO CR CONCENTRATION RATIO CRM CREW AND RESUPPLY MODULE CW CONTINUOUS WAVE db DECIBEL DC DIRECT CURRENT DDT&E DESIGN, DEVELOPMENT, TEST AND EVALUATION DOE DEPARTMENT OF ENERGY EMI ELECTROMAGNETIC INTERFERENCE EOL END OF LIFE FET FIELD EFFECT TRANSISTOR GaAlAs GALLIUM ALUMINUM ARSENIDE GEO GEOSYNCHRONOUS EARTH ORBIT GFRTP GRAPHITE FIBER REINFORCED THERMOPLASTIC GHz GIGAHERTZ (10^9 CYCLES/SECOND)
GW GIGAWATT (109 WATTS) GWe GIGAWATT ELECTRIC HF HIGH FREQUENCY h2 HYDROGEN IMPATT IMPACT AVALANCHE TRANSIT TIME I sp SPECIFIC IMPULSE JPL JET PROPULSION LABORATORY JSC JOHNSON SPACE CENTER KSC KENNEDY SPACE CENTER kV KILOVOLT kW KILOWATT LaRC LANGLEY RESEARCH CENTER LBF POUNDS OF FORCE LEO LOW EARTH ORBIT LeRC LEWIS RESEARCH CENTER Mm MICROMETER MPD MAGNETOPLASMADYNAMICS MPTS MICROWAVE POWER TRANSMISSION SYSTEM MSFC MARSHALL SPACE FLIGHT CENTER MTBF MEAN TIME BETWEEN FAILURE MT METRIC TON MW MEGAWATT (106 WATTS) N NEWTON NASA NATIONAL AERONAUTICS AND SPACE ADMINISTRATION NM NAUTICAL MILES
NS NORTH-SOUTH O&M OPERATING AND MAINTENANCE 02 OXYGEN POP PERPENDICULAR TO THE ORBIT PLAN psi POUNDS PER SQUARE INCH RP ROCKET PROPELLANT RPC RECTENNA POWER COLLECTION RF RADIO FREQUENCY RFI RADIO FREQUENCY INTERFERENCE SCB SATELLITE CONSTRUCTION BASE SECS SOLAR ENERGY COLLECTION SYSTEM SE&I SYSTEMS ENGINEERING AND INTEGRATION SG SWITCHGEAR Si SILICON SPS SATELLITE POWER SYSTEM VAB VERTICAL ASSEMBLY BUILDING (AT KSC) VHF VERY HIGH FREQUENCY VLF VERY LOW FREQUENCY
Transportation Systems BLOW BOOSTER LIFT OFF WEIGHT ch4 METHANE COTV CARGO ORBITAL TRANSFER VEHICLE EOTV ELECTRIC ORBITAL TRANSFER VEHICLE ET EXTERNAL TANK GCR GAS CORE REACTOR GLOW GROSS LIFT OFF WEIGHT HLLV HEAVY LIFT LAUNCH VEHICLE HTO HORIZONTAL TAKEOFF IOTV INTRA-ORBIT TRANSFER VEHICLE lch4 LIQUID METHANE lh2 LIQUID HYDROGEN LOX LIQUID OXYGEN MOTV MANNED ORBITAL TRANSFER VEHICLE OTV ORBITAL TRANSFER VEHICLE PLV PERSONNEL LAUNCH VEHICLE POTV PERSONNEL ORBITAL TRANSFER VEHICLE SRB SOLID ROCKET BOOSTER SSBE SPACE SHUTTLE BOOSTER ENGINE SSME SPACE SHUTTLE MAIN ENGINE SSTO SINGLE STAGE TO ORBIT T/W THRUST TO WEIGHT ULOW UPPER STAGE LIFT OFF WEIGHT AV DELTA V - CHANGE IN VELOCITY
digitized by the Space Studies Institute Research Library ssi.og
I. SUMMARY The SPS Reference System concept as defined herein, is largely a product of system definition studies conducted by the Boeing Aerospace Company under contract to the Johnson Space Center (JSC) from December 1976 to December 1977, and Rockwell International under contract to the Marshall Space Flight Center (MSFC) from March 1977 to March 1978. The results of these two system definition studies combined with in-house efforts at both NASA Centers and several smaller contracted studies provided the data from which the Reference System was developed. The two parallel, but independent, system definition studies resulted in well-integrated system concepts in which some major elements were very similar, but others were markedly different. To meet the Reference System milestone with essentially a single concept, a third system was developed by the selection of system elements of the two individual concepts. Part III of this report describes the Reference System. Appendix A describes the various systems analyses that have been conducted. Appendix B describes the independent system definitions developed by Boeing and Rockwell. The Reference System description emphasizes technical and operational information required in support of environmental, socioeconomic, and comparative assessment studies. Supporting information has been developed according to a guideline of implementing two 5 GW SPS systems per year for 30 years beginning with an initial operational date of 2000 and with SPS1s being added at the rate of two per year (10 GW/year) until 2030. Figure 1 illustrates the Reference System concept, which features galliumal uminum-arsenide (GaAlAs) and silicon solar cell options. The concept utilizes a planar solar array (about 55 km ) built on a graphite fiber reinforced thermoplastic structure. The silicon array uses a concentration ratio of one (no concentration), whereas the GaAlAs array uses a concentration ratio of two. A one- kilometer diameter phased array microwave antenna is mounted on one end. The
antenna uses klystrons as power amplifiers with slotted waveguides as radiating elements. The satellite is constructed in geosynchronous orbit in a six-month period. The ground receiving stations (rectenna) are completed during the same time period. The other two major components of an SPS program are (1) the construction bases in space and launch and mission control bases on earth and (2) fleets of various transportation vehicles that support the construction and maintenance operations of the satellites. These transportation vehicles include Heavy Lift Launch Vehicles (HLLV), Personnel Launch Vehicles (PLV), Cargo Orbit Transfer Vehicles (COTV), and Personnel Orbit Transfer Vehicles (POTV). The earth launch site chosen is the Kennedy Space Center, pending further study.
Figure 1. SPS Reference System Concept
II. INTRODUCTION The DOE and NASA have initiated a joint program which is described in the document entitled SPS Concept Development, and Evaluation Program Plan (July 1977 - August 1980), D0E/ET-0034, dated February 1978. The objective of the program is to generate information from which a rational decision can be made regarding the direction of the SPS program after fiscal year 1980. Briefly, the plan states that NASA will conduct systems definition of the SPS and the DOE will evaluate health, safety, and environmental factors and will study SPS economic, international, and institutional issues. In addition, DOE will make comparative assessments of the concept relative to alternative power sources for the future. Figure 2 shows a simplified diagram of the development and evaluation methodology. As indicated, the major milestones of the plan are baseline concept(s) selection - October 1978; preliminary program recommendations - May 1979; updated program recommendations - January 1980; and final program recommendations - June 1980. In this report, the term Reference System is used instead of baseline concept(s) as being more appropriate for the current level of definition and understanding. Using the results of this evaluation program as a basis and considering other pertinent factors, it will be possible for the Administration to either recommend continuation of the program in accordance with a defined option or terminate the program. The purpose of this document is to present a description of the Reference System. It is submitted in response to the Baseline Concept(s) Definition program milestone (October 1978) established in the DOE/NASA plan as indicated above. Section III of the report presents a description of the Reference System with emphasis on technical and operational information required by DOE to conduct environmental, socioeconomic, and comparative assessment studies. It is recognized that the Reference System lacks maturity as reported herein. Definition work is continuing to develop further understanding of the system.
Figure 2. Simplified Diagram of SPS Concept Development and Evaluation Methodology
Section IV provides a discussion of technology advancement requirements which focuses on critical questions to be resolved that affect SPS feasibility. Also, alternate technologies which are delineated in the summary of study results section, are reviewed. A summary of SPS documentation is provided in Section V to assist the reader in locating reference documents and following the flow of study results that has occurred over the past several years. This documentation summary serves as a list of references for this report. Appendix A provides a summary of systems analysis results. The information presented is based on results of the key design and operational trade-off studies conducted to date in the major areas of investigation. The data base for the analysis summaries consists of study reports and other documents from current studies as well as those prepared in the late 1960's and early 1970's. Much of the early work was performed by A. D. Little, Raytheon Corporation, Spectro- lab, Inc., Grumman Aerospace Corporation and the NASA Lewis Research Center. The Jet Propulsion Laboratory also made significant contributions to the SPS data bank, particularly in the area of microwave power transmission. The primary sources of information for the Reference System description are the systems definition reports published by Rockwell International and Boeing Aerospace Company under contract to MSFC and JSC, respectively. Appendix B provides systems descriptions resulting from the studies conducted by the Boeing Aerospace Company and Rockwell International. Background This document deals with the solar power satellite concept as illustrated in figure 3. It is a primary electrical power source that involves generating electrical power (from solar energy) in geosynchronous orbit, transmitting the power to earth via focused microwave beams, and collecting and converting the microwave beams into useful electricity on the earth's surface. This concept was suggested in 1968 ("Power from the Sun: Its Future," Dr. Peter E. Glaser, Science, Vol. 162, November 22, 1968, pp 857-861).
Figure 3. Solar Power Satellites
A number of other potential energy system concept involving the use of space satellites have also been suggested: 1. Orbiting nuclear reactor power systems (in lieu of solar collector/ converters) with microwave transmission of power to earth. 2. The power relay satellite concept (Reference 12) in which power systems on the earth’s surface or in low orbit transmit power by microwave to geostationary satellites. These geosynchronous satellites then relay (reflect) the microwave energy to ground stations placed at desired locations on earth. 3. Solar reflecting satellites (mirrors) in earth orbit which reflect solar energy directly to earth to augment ground-based solar power plants, allowing night operation or increased output. 4. Laser power transmission (in lieu of microwave) from the satellite. The orbiting nuclear reactor concept has been evaluated to a limited extent. While it might offer the advantage of compactness relative to solar powered systems, its mass and complexity are significantly greater than solar powered systems. Significant safety and environmental questions remain to be addressed. The power relay satellite is a long distance power transmission concept rather than a primary electrical energy source; consequently, it is not viewed as a basic alternative to the SPS concept. The idea of placing large mirrors in earth orbit has been evaluated (Reference 14). Analysis indicates that with a mirror in geosynchronous orbit, the smallest illuminated "spot" on the earth would be about 330 km (205 miles) in diameter, governed by optical geometric considerations. If continuously illuminated at an average level of one sun, this large area would tend to rise in temperature to approximately 150°F, posing severe environmental problems. Placing large mirrors in lower altitude orbit reduces the size of the illuminated "spot." However, the mirrors would not be stationary with respect to a point on the earth. Thus, to achieve continuous illumination at a given location, numerous mirrors would have to be placed in low orbit. In addition, cloud cover and weather conditions would have an adverse effect on solar insolation precluding consideration
of this concept as a primary, baseload electrical power source. Laser power transmission has a significantly lower efficiency for longdistance power transmission than is estimated for microwave power transmission. Atmospheric attenuation is substantial compared to microwave frequency transmission. Therefore, this concept is presently less attractive than the microwave concept for transmitting power from geosynchronous orbit. Alternate system concepts such as solar collectors and laser transmitters in low earth orbit with relays in geosynchronous orbit have received preliminary consideration. Another SPS concept is being evaluated that makes use of materials derived from the moon to construct the SPS. The moon's lower gravitational force (onesixth of earth's) would allow much less propulsion energy to move payload to geosynchronous earth orbit. This idea appears to have merit in terms of conserving earth resources and possibly reducing the cost of space transportation; however, it would require development of moon-based mining, manufacturing and launch facilities. Consequently, the research and development requirements for such an approach would be greatly increased. While the above options offer interesting possibilities, the present DOE/ NASA program focuses evaluation on the SPS concept using terrestrial materials and deployed in geosynchronous orbit as illustrated in figure 3. This evaluation does not exclude the possibility of future consideration of the alternatives and options such as identified above.
III. REFERENCE SYSTEM DESCRIPTION The purpose of this section is to describe the SPS Reference System which has evolved primarily from system definition studies conducted by Boeing Aerospace Company and Rockwell International. The system concept presented herein is the result of numerous trade-off studies and engineering analyses, which are summarized in Appendix A of this report. It should be emphasized that the system described herein is preliminary and incomplete in detail in some areas. Evaluation by both DOEi and NASA will continue to progress with the Reference System evolving and maturing as further details are developed. A. Guidelines and Assumptions The guidelines and assumptions utilized in the Reference System definition are listed below. o SPS operational date is year 2000. o Rate of implementation is two 5 GW systems per year; 300 GW total capacity for costing purposes. o All ground rectennas sized for 5 GW. o SPS operation in geosynchronous orbit. o Systems operating frequency is 2.45 GHz. 2 o Microwave power density not to exceed 23 and 1 mW/cm at center and edge, respectively, of rectenna. o All materials derived from earth resources. o System life is 30 years with no salvage value or disposition costs, o Zero launch rate failure assumed. o Technology availability date is 1990. o No cost margins will be used, o Cost estimates in 1977 dollars, o System weight growth factor (25%) to be reflected in costs. B. System Overview The Reference System is sized for 5 GW DC power output into a conventional power grid. The satellite has one end-mounted antenna which transmits
to a rectenna on the ground. This concept is illustrated in Figure 4. The configuration of the satellite consists of a planar solar array structure built from a graphite composite material. Two conversion options are presented. One is the single-crystal gallium-aluminum-arsenide (GaAlAs) solar cells with a concentration ratio of 2 as illustrated in figure 4. The other energy conversion option is the use of single-crystal silicon (Si) solar cells with no concentration. The size of the solar array is dictated primarily by the efficiency chain of the various elements in the system. Figure 5 shows the end-to-end efficiency chain for the GaAlAs and silicon solar cell options. With the satellite designed to provide 5 GW of DC power to the utility busbar and an overall efficiency of approximately 7%, it is necessary to size the solar arrays to intercept approximately 70 GW of solar energy as indicated in figure 5. The quoted efficiency is the minimum efficiency, including the worst-case summer solstice factor (0.9675), the seasonal variation (.91), and the end-of-life (30 year) solar cell efficiency assuming annealing. For the GaAlAs case, the end- of-life (30 year) concentrator reflectivity is 0.83. Since only half of the intercepted solar energy is reflected by the concentrators, the equivalent overall efficiency is 0.915. The GaAlAs option is a five-trough configuration with a solar blanket area of 26.52 km^, a reflector area of 53.04 km^ and an overall planform area of 55.13 km^. The silicon option has the solar blanket with no concentration resulting in a blanket area of 52.34 km^ and a planform area of 54.08 km^. The satellite in either option is oriented so that the antenna main rotational axis remains perpendicular to the orbital plane. The end-mounted microwave antenna is a one kilometer diameter phased- array transmitter. The phase control system utilizes an active, retrodirective array with a pilot beam reference for phase conjugation. Klystrons are used as the baseline power amplifier with slotted waveguides as the radiating element. The ground rectenna has subarray panels with an active element area of 78.5 km . The satellite is constructed in geosynchronous orbit with construction time being six months. The initial estimates of construction crew size are 555
Figure 4. SPS Reference System - Satellite Configuration
Figure 5. SPS Efficiency Chain (GaAlAs CR2 and Si CR1)
for the silicon option (480 in GEO and 75 in LEO) and 715 for the GaAlAs option (680 in GEO and 35 in LEO}. The transportation system is made up of four major items. These include: (1) the Heavy Lift Launch Vehicle (HLLV), the Cargo Orbit Transfer Vehicle (COTV), the Personnel Launch Vehicle (PLV), and the Personnel Orbit Transfer Vehicle (POTV). The HLLV is a two-stage, vertical launch, winged, horizontal land-landing, reusable vehicle with 424 metric ton payload to low earth orbit. The earth launch site was chosen as Kennedy Space Center pending further study. The COTV is an independent, reusable electric engine-powered vehicle which transports cargo from the HLLV delivery site in low earth orbit (LEO) to the geosynchronous earth orbit (GEO). For the GaAlAs SPS option, the COTV is powered by GaAlAs solar cells, whereas a silicon solar cell power supply is used for the silicon SPS option. Personnel for the orbital construction and support functions are transported to LEO via the PLV which is a modified space Shuttle Orbiter with a passenger module. The POTV, a two-stage reusable, chemical fuel vehicle is used to transfer personnel from LEO to GEO and return to LEO. The satellite construction scenario for the Reference System is illustrated in figure 6. The HLLV is shown transporting cargo to LEO while the COTV and the POTV are illustrated transporting cargo and personnel, respectively, from LEO to GEO. A LEO operations base is used for temporary storage of supplies and propellant. One satellite is shown in GEO during the construction phase while another satellite is shown in the conventional operational phase transmitting energy to ground rectenna. Figure 7 summarizes the characteristics of the Reference System. C. Solar Cells and Blankets Both GaAlAs and single-crystal silicon solar cells are considered reference energy conversion devices. Figure 8 shows a cross-section of the GaAlAs and Si solar cells and blankets. The basic GaAlAs solar cell consists of a 5 ^m thick GaAs-P-N cell with a 0.03 to 0.05/pi thick GaAlAs front-side window. The solar cell efficiency is 20% at AMO, 28°C. The design operating
Figure 6. SPS Operations
Figure 7. Reference System Characteristics
temperature is 125°C, which produces an 18.2% cell efficiency. At 125°C, self-annealing of radiation-induced damage occurs in the GaAlAs cells. As indicated in figure 8, the solar cell blanket material is Kapton (25 ^m thick). The 20 yum synthetic sapphire (AI2O3) substrate, used in an inverse orientation, also acts as the cell cover. The blanket weight is 0.252 kg/nA The projected cost of the GaAs solar cell blanket is $71/m^. Solar reflectors for the GaAlAs CR2 concept consist of thin reflective membranes aluminized Kapton mounted in a 60° Vee trough configuration. The reflector membrane has a reflectivity of 0.9 BOL and 0.83 EOL. The effective end-of- 1ife efficiency is 0.915 as previously stated. The membrane mass is 0.018 kg/m^. The silicon solar blanket consists of 50 jqm thick single crystal silicon solar cells with borosilicate cover glass electrostatically bonded to the cells front and back. The cells are designed with both P and N terminals brought to the back of the cells. This feature makes it possible to use 12.5 jL|m silver- plated copper interconnections which are formed on the substrate glass. Complete panels are assembled electrically by welding together the module-to-module interconnections. The cell efficiency is 17.3% (AMO, 28°C) at beginning of life. At design operating temperature of 36.5°C, the efficiency is 16.5%. A key feature of the blanket design is the ability to perform in-situ annealing of the solar cells using a laser annealing concept. The laser annealing concept utilizes gimbal-mounted CO2 lasers. The gimbals would be mounted on an overhead gantry structure to permit transversing of the entire solar array by several lasers as illustrated in figure 9. The laser beam heats the solar cells to annealing temperature (500°C) without damaging cell interconnect and substrate materials. Annealing is required to recover radiation induced degradation of the cells. The o projected cost of the silicon solar blankets is $35/m . D. Solar Array and Structure The solar array consists of the deployed solar cell blankets attached to solar array structure. In the case of the CR2 GaAlAs option, the array includes the solar reflectors mounted on each side of the solar cell blankets. The solar cell blankets would be transported to orbit in modular packages in
Figure 8. Solar Cell Options for SPS Figure 9. Laser Annealing Concept
either a rolled or folded configuration. The blanket modules would be connected together in a parallel/series arrangement to obtain the desired voltage output of 40 to 45 kv. Figure 10 illustrates the method of attaching the silicon solar cell blankets to the support structure. The GaAlAs blankets would be attached in a similar fashion. The reflector panels for the GaAlAs option are pleated at intervals to produce an accordian fold and then rolled for storage and shipment to orbit. The reflector panels would be attached to the primary structure by cables (catenaries). The primary structure for solar array and microwave transmitting antenna is an open truss-type design. The structure material is a graphite-fiber reinforced thermoplastic composite. The basic elements (beams) are designed for automatic fabrication in space. The CR1 silicon cell option would utilize a rectangular configuration constructed with truss-type beams. The CR2 GaAlAs option utilizes similar construction elements for the solar array structure, but would include additional elements to form the structure for attaching solar reflectors. E. Power Distribution The prime function of the power distribution system is to accumulate and control prime power from the solar cell panels; control, condition, and regulate the quantity and quality of the electrical power generated for the microwave generators; provide for the required energy storage during solar energy occulation or system maintenance shutdown; and provide for monitoring fault detection, and fault isolation disconnects. Figure 11 shows a schematic diagram of the solar array power collection and distribution system. The solar array is divided into power sectors. Each power sector is switchable and can be isolated from the main power bus, facilitating solar cell annealing operations (for silicon cells) and/or other servicing. Solar array power is controlled by high voltage circuit breakers near the buses. Voltage is controlled by turning groups of strings on or off, depending on load requirements. Two sections of the array provide the required voltage at the sliprings using the sheet conductor voltage drop to achieve the required voltage at the sliprings.
Figure 10. Solar Cell Blanket Support
Figure 11. SPS Power Distribution
The microwave power transmitting antenna includes a power distribution system which transmits DC power from the slipring system to the DC-RF generators. Conductors from the slipring brushes are tied to DC/DC converters through switchgear tc allow isolation when performing maintenance. Conductors are then tied between voltage summing buses through other switchgear for transmitting the required power to the klystrons. An electrical energy storage power system is located on the antenna structure with a bus routed along the regular network for operation during powered-down periods such as may occur during solar eclipse periods. A battery energy storage system would be utilized for storage of about 12 MW- hours of electrical energy. F. Rotary Joint Power transfer from the solar array section to the microwave antenna is accomplished via a rotary joint (figure 12) using a siipring/brush assembly. Mechanical rotation and drive is provided by a mechnical turntable 350m in diameter. The antenna is suspended in the yoke by a compliant mechanical joint to isolate the antenna from turntable vibrations. The antenna is mechanically aimed by control moment gyroscopes (CMG’s) installed on its structure. A positive feedback with a low frequency band-pass allows the mechanical turntable to drive the yoke to follow the antenna and also provide sufficient torque through the joint to keep the CMG's desaturated. G. Attitude Control System (ACS) The attitude control system for the reference 5 GW system is described for a CR2 GaAlAs option. The ACS for the CR1 Si option will be similar, although the number of thrusters and propellant requirements will differ. Preliminary Baseline ACS Description - GaAlAs-CR2 - The baseline ACS features an argon ion bombartment thruster reaction control system (RCS) whose characteristics are given in figure 13. A total of 64 thrusters is included in the system to provide the required redundancy assuming: an annual maintenance interval, 5000 hour thruster grid lifetime and a 5-year thruster MTBF. Approxi-
Figure 12. Antenna Yoke and Turntables
Figure 13. Attitude Control System Characteristics
mately 36 operating thrusters are required. Sixteen thrusters are located on the lower portion of each corner of the collector. Each thruster is gimbal - led individually to facilitate thruster servicing using a servicing cab, to permit operation of adjacent thrusters during servicing, and to provide redundancy. The system is nominally designed for X-POP operation (long axis per- pendicular-to-orbit-plane). The pertinent features and locations of the Attitude Reference Determination System (ARDS) are also given in figure 13. The average power required for the system is 34 megawatts. H. Microwave Power Transmission System The reference microwave power transmission system was developed considering environmental effects, antenna size tradeoffs, antenna thermal heating limitations, and ionospheric heating effects. The present microwave system has DC-RF power converters feeding a 1 km diameter phased array antenna with a 10-decibel (dB) Gaussian taper illumination across the array surface. This antenna shown in figure 14, is composed of 7220 subarrays, approximately 10 meters X 10 meters on a side, having slotted waveguides as the radiating surface with DC-RF power tubes mounted upon the backside of the subarrays. The antenna structure is a graphite composite material while the slotted waveguides are aluminum. Each subarray has its own RF receiver and phasing electronics to process a pilot beam phasing signal from the ground. The subarrays are phased together to form a single coherent beam focused at the center of the ground antenna/rectifying system (rectenna). This power beam has approximately 88% of its energy within a 5 km radius perpendicular to boresight, with a resultant beam width of 1.2 arc-minutes. Microwave System Parameters and Sizing Considerations - Some of the key parameters of the SPS microwave system are presented in figure 15. The power capability of the SPS system was sized by: (1) thermal limitations of o 22 kW/m in the transmit array due to waste heat from the DC to RF power con- 2 verter tubes; (2) a peak power density limitation in the ionosphere of 23<mW/cm
Figure 14. Transmitting Antenna Functional Description
Figure 15. Microwave Power Transmission System Parameters
Figure 15. Microwave Power Transmission System Parameters (Continued)
to prevent non-linear heating interactions between the ionosphere and the microwave beam; and (3) microwave transmission efficiency considerations (particularly the RF levels into the rectenna). Studies into the microwave beam/ionosphere interactions indicate that non-linear thermal self-focusing instabilities in the F-region (200 to 300 kilometers altitude) and thermal runaway conditions in the D-region (100 kilometers) may limit the maximum power density at the center of the beam to approximately 23 mW/cm at the 2450 MHz operating frequency (references 26, 29, 30 and 31). Above this threshold power density level (which is a theoretical number, not yet verified by experiments), non-linear interactions between the power beam and the ionosphere begin to occur. These non-linear heating interactions are of concern because of possible degradations to existing HF and VHF communication and VLF navigation systems due to RFI effects and multipath degradations. The heated ionosphere may also introduce phase jitter and/or differential phase delays on the uplink pilot beam signal. These ionosphere/microwave beam interactions are now being studied, both theoretically and experimentally. Antenna Characteristics The microwave antenna has both a primary and a secondary structure composed of a graphite composite material. The primary structure is an open truss, 130 meters deep, with an octagonal shape over 1000 meters in width and length. The secondary structure is a deployable cubic truss, 9.93 meters in depth, which provides support for installation of the microwave subarrays. The aperture illumination function across the 1-kilometer transmit array was optimized to provide the maximum amount of RF power intercepted by the ground rectenna and to minimize the sidelobe levels. A number of different illumination functions, operating in the presence of phase and amplitude errors and element (subarray) failures, have been studied (see Appendix A). The 10- decibel Gaussian taper has the best overall performance of the optimized illumination functions after considering the maximum power density constraints in the transmit array and rectenna.
A 10 step, 10 dB Gaussian taper for the transmit array is given in figure 16. There are 36 DC-KF power converter tubes per subarray at the center of the antenna, decreasing in quantized steps down to four tubes per subarray at the edge to provide the 10 dB Gaussian taper. There will be a total of 101,552 tubes in the antenna, integrated into the subarrays as shown in figure 17. This particular configuration uses a 70 kw klystron tube; an alternative concept has a 50 kw klystron, which requires approximately 140,000 tubes. The number of tubes per subarray would then vary from 50 tubes at the center of the antenna to six tubes at the edge in order to provide the 10 dB illumination taper. The radiation pattern for the 10 dB taper, 1 km array (and cr= 10° RMS phase error, + 1 dB amplitude error, and 2% random failures) at the ground rectenna is shown in figure 18. The effect of the antenna errors is to produce a wider, lower intensity main beam with higher sidelobes. For the SPS sytem concept, only a portion of the main lobe will be collected; the sidelobe energy occupies a very large area at very low density levels and is not economically feasible to collect. The peak power densities are 23 milliwatts per square centimeter at the rectenna boresight, 1 milliwatt per square centimeter at the edge of the rectenna, and 0.08 milliwatt per square centimeter for the first sidelobe, o which is two orders of magnitude below the U.S. radiation standard of 10mW/cm . If there is a total failure within the phase control system (for example, the uplink pilot beam transmitter is shut off), the subarrays will no longer be phased together and the total beam will be defocused. As shown in figure 18, o the peak intensity of the beam drops to 0.003 mW/cm and the beam width greatly increases. This peak power density is significantly less than the USSR guideline indicated on figure 18. Consequently, this is a fail-safe feature of the phasing system. In addition, there are sensors near the rectenna to detect any large changes in incident power density; this information would immediately be transmitted to the antenna to cease operations. In addition to the sidelobe patterns near the rectenna, the far-sidelobe patterns have been calculated. There had been some concern about the radio
Figure 16. Microwave Power Transmission System 10 dB taper
Figure 17. Microwave Array Power Distribution
Figure 18. Power Density at Rectenna as a Function of Distance from Boresight
interference levels at large distances from any given rectenna because of frequency allocation problems. The SRS downlink power beam lies in the 2400-2500 MHz frequency band which has been reserved for Government and non-Government industrial, medical, and scientific (IMS) usage. By definition, anyone operating in an IMS band must accept interference from any other user within this band. However, this 100 MHz band is not recognized by some of the eastern European countries which reserve 2375 MHz + 50 MHz for the IMS band. The far sidelobe levels as shown in figure 19 indicate the peak levels for one 5 gigawatt SPS system are three to four orders of magnitude below 0.01 mW/cm . For simultaneous operation of 200, 5 gigawatt SPS systems, the average peak level is still one to two orders of magnitude lower than 0.01 mW/cm2. Grating lobes, which occur at 440 km intervals from the rectenna, are functions of subarray size and mechanical misalinement of the subarrays within the 1 km phase array. The grating lobes occur at spatial distances corresponding to angular directions off axis of the antenna array where the signals from each of the subarrays add in-phase. When the boresights of the subarrays are not alined with the uplink pilot beam transmitter at the rectenna, the unwanted contributions of the array factor of the antenna do not lie in the null-points of the subarray pattern as shown in figure 20. Even though the phase control system will still point the composite beam at the rectenna, some energy will be transformed from the main beam into the grating lobes. The amount of energy in the grating lobes depends upon the misalinement (or how far the array factor is displaced from the null points of the subarray pattern). These grating lobes are somewhat unique in that they do not spatially move with misalinement changes, rather they are stationary with an amplitude dependence upon the mechanical misaline- ments. This behavior is due to the operating characteristics of the retro- directive phasing system. Based on environmental considerations, the grating o lobes are constrained to be less than 0.01 mW/cm . The total mechanical aline- ment requirements for both the subarrays and the total array can be determined from this constraint. The 10.4 meter X 10.4 meter subarray which is considered
Figure 19. Peak Power Density Levels as a Function of Range From Rectenna
Figure 20. Grating Lobe Characteristics
to be the smallest entity for phase control, has the peaks of the grating lobe patterns at the ground as shown in figure 21. Since the distance between maxima for the grating lobes is inversely proportional to the spacings between subarrays, a 10.4 meter square subarray has peaks every 440 km. If the phase control system is extended down to the power module level, the grating lobes will be spatially smeared and the peaks greatly reduced in amplitude. This improvement in grating lobe pattern would be due to differences in spacings between the power tubes within the antenna. There are two types of mechanical misalinements: (1) a systematic tilt of the entire antenna structure, and (2) a random tilt of the individual subarrays. The systematic tilts have a greater impact than the random subarray tilts on the grating lobe peaks. An example of the first grating lobe peak for a total antenna/subarray tilt of 3.0 arcminutes is shown in figure 22. Other simulations have established mechanical alinement requirements of less than 1 arc^minute for the antenna tilt and less than 3-arc-minutes for the random subarray tilts. The near-field antenna pattern for distances close to the transmit array is shown in figure 23. A peak density of approximately 32 kk/m occurs at a distance of 1600 km. Rectenna Characteristics - The present ground receiving antenna (rectenna) configuration, which receives and rectifies the downlink power beam, has half-wave dipoles feeding Schottky barrier diodes. Two-stage low-pass filters between the dipoles and diodes suppress harmonic generation and provide impedance matching. For economic reasons, the rectenna is a series of serrated panels perpendicular to the incident beam rather than a continuous structure. Each panel has a steel mesh ground plane with 75-80% optical transparency. This mesh is mounted on a steel framing structure, supported by steel columns in concrete footings. Aluminum conductors are used for the electrical power collection system. The rectenna will produce RFI effects due to rescattered incident radiation and harmonic generation within the diodes. There will also be a small amount of RF energy leakage through the ground screen as well as knife eoge diffraction patterns at the top edge of each rectenna section. The general
Figure 21. Grating Lobe Maxima
Radius from Boresight (Km) Figure 22. Peak Power Density for Sidelobes and Grating Lobe as a Function of Range from Rectenna
Figure 23. Near-Field Antenna Patterns
rectenna parameters may be summarized as follows: Total Active Panel Area: 78.5 km^ Configuration; Panels (multiple antenna elements feeding a single diode. Panels should be openfaced to reduce wind loading, with a maximum of 1% leakage energy. Receive Elements: Half-wave rectifying diodes. To provide an estimate of the power levels around the rectenna, studies indicate that the dipoles have 98% collection efficiency under normal loading conditions (reference 9). The 2% reflected microwave power is directed upward and towards the southern horizon. This reflected energy at 2.45 GHz is only partially coherent since the regions of coherence for the incident beam are limited due to phase irregularities in the heated ionosphere and atmosphere. Harmonics of 2.45 GHz will be generated within the half-wave rectifying diodes and will be reradiated back through the low-pass filters and dipoles. Initial measurements of the harmonic levels relative to the fundamental indicate the second, third, and fourth harmonics are down by -25 dB, -40 dB, and ^70 dB, respectively, for the normal dipole/diode rectifier configuration (reference 9). There will also be leakage power through the rectenna. For a ground plane transparency of 80%, approximately 1% of the incident power appears as leakage through the wire mesh. Since the rectenna's receiving surface appears serrated with individual panels perpendicular to the incident radiation, there will be diffraction losses at the top edge of each section. An analysis of the knife edge diffraction pattern has been made to determine the variation in power density incident upon the adjacent rectenna section (reference 9). The power density will vary in the shadow region (area behind the rectenna section) as shown in figure 24. It may be necessary to extend the size of the rectenna section to intercept part of the shadow region. There will also be energy lost as heat in the rectenna due to
Figure 24. Rectenna Patterns and Power Levels
2 I R losses in the receiving elements and in the diodes. Approximately 7% of the incident energy is expected to be lost as heat. The expected power levels around the rectenna for reflected energy, leakage energy, harmonics, diffractions, and heat losses are summarized in figure 24. Microwave System Efficiency - The total microwave system efficiency from the DC power output at the rotary joint to the collected DC power output of the rectenna is 63%. A breakdown of the efficiencies of the microwave subsystems is given in figure 25. Phase Control System - The phase control system has an active, retro- directive array with a pilot beam reference for providing phase conjugation. Each subarray or possibly each power module (that portion of a subarray fed by one klystron tube), has its own RF receiver which processes the uplink pilot beam reference and inserts the proper phasing signal to form a single coherent beam at the ground rectenna. Tradeoff studies are now being conducted to determine if the phase conjugation should be at the subarray level or the power module level. Conjugation at the power module level improves main beam efficiency and microwave environmental effects, but increases costs and complexity. The Reference System includes: (1) phase lock loop around each power tube for phase stability and noise suppression. (2) double sideband, suppression carrier modulation which is symmetrical about the downlink power beam frequency f^, as shown below.
Figure 25. Microwave Transmission Efficiency
The two sidebands are demodulated in the RF receivers in the subarrays (or power modules) and the carrier is reconstructed. This prevents beam squint problems arising from different uplink and downlink frequencies, and it allows the proper phase conjugation to be made. The ionosphere constrains the frequency separation Af between the sidebands and the downlink carrier to be greater than 10 MHz, the maximum plasma resonance frequency. This limitation is to prevent intermodulation products between the uplink pilot beam and the downlink power beam from creating parametric instabilities associated with overdense ionospheric heating, (3) coding of the pilot beam for security and pilot discrimination. Since multiple SPS satellites will be illuminated by a single pilot beam transmitter, each satellite has to recognize which pilot beam signal it should respond to. In addition, coding will prevent power drain from any intentional interfering signals. (4) ground safety control system (ground sensors for interpreting beam shape) with a command link capability to the satellite. RFI Characteristics -■ The radio frequency interference comes primarily from the DC-RF power converter tubes. This interference can be divided into three main categories: (1) interference from the high power downlink beam due to sidelobe and grating lobe radiations, (2) spurious noise generated near the carrier frequency by the tubes, and (3) harmonic generation within the tubes. The sidelobe and grating lobe levels were previously examined. Within the phase control system, the phase lock loop around each power tube will reduce the spurious noise close to the carrier frequency. A representative loop might have a 5 MHz bandwidth with a second or third order filter. This loop will not affect the tube noise characteristics outside the 2450 + 50 MHz band. However, the klystron tubes will have a multiple cavity design which provides additional filtering (24 dB/octave) to reduce the out-of-band noise. The SPS noise density characteristics are summarized in figure 26 (reference 31), The CCIR (International Radio Consultative Committee) requirement for power flux density at the earth's surface is -180 dBW/m /Hz for S-band frequencies with an angle of arrival above 25°. As shown in figure 26, the RFI
effects due to spurious noise will be below the COIR requirements, provided the klystron tubes are phase-locked for noise reduction and a multiple cavity design is used. Figure 26. Noise Power Density at Ground for a 1 km, 5 GW SPS Antenna
I. Mass Statement A summary of the satellite mass properties is presented in Table 1. The masses are separated into three major segments: solar array, microwave antenna, and array antenna interfaces (the section of satellite between the array and antenna which includes the rotary joint, sliprings, and antenna yoke). The GaAlAs configuration utilizes a concentration ratio of 2 which reduces the required blanket area and in turn the blanket mass. The antenna section mass properties are the same for both options. The antenna mass is dominated by the transmitter subarray which includes the klystrons and waveguides. The other large items are the power distribution system and the thermal control system for the klystrons and DC/DC converters. The total mass for the two options, including a 25% contingency factor, is 34 and 51 million kilograms for the GaAlAs and silicon options, respectively. J. Space Transportation This section provides descriptions of the reference space transportation system vehicles. The alternative concepts from which the Reference System was selected are described in Appendix A. The vehicles are distinguished by their primary payload, either cargo or personnel, and their area of operations between earth and low earth orbit (LEO) or between LEO and geosynchronous earth orbit (GEO). Cargo is transported from the earth's surface to LEO by the HLLV and personnel (and priority cargo) are transported from earth to LEO and back by the PLV. fransportation between LEO and GEO is provided by the COTV and the POTV. The general groundrules followed in the development and evaluation of the transportation system are: • The SPS transportation system elements, with the possible exception of shuttle derived PLV's, are dedicated and optimized for the installation, operation, and maintenance of the SPS. • The SPS transportation system will be designed for minimum total program cost.
RkJQdWJsaXNoZXIy MTU5NjU0Mg==