Satellite Power System (SPS) Feasibility Study FINAL REPORT Rockwell International Space Division SD 76-SA-0239-2
SD 76-SA-0239-2 SATELLITE POWER SYSTEM (SPS) FEASIBILITY STUDY FINAL REPORT DECEMBER 1976 CONTRACT NAS8-32161 Prepared for National Aeronautics and Space Administration George C. Marshall Space Flight Center Rockwell International Space Division
FOREWORD This document presents the final report of the Satellite Power System (SPS) Feasibility Study, conducted by Rockwell International Corporation's Space Division for the National Aeronautics and Space Administration, George C. Marshall Space Flight Center, under Contract NAS8-32161. The results of the four-month technical effort are reported in this volume. A companion volume (SD 76-SA-0239-1) presents an Executive Summary of the study results. The results presented herein are due principally to the contributions of the following Rockwell personnel: Messrs. Will Cooper, Bob Jamieson, Dave Reed, Lee Smith, Charlie Tomita, and Duane Tonelli. Special acknowledgment should go to the NASA COR, Mr. Dale Wasserman, for his direction and coordination with other SPS activities. James F. Madewell Director, Advanced Systems William V. McRae, Jr. Study Manager
CONTENTS Section Page 1 INTRODUCTION ...... 1 1.1 Study Approach ................................ 1 1.2 Study Ground Rules and Assumptions .... 2 1.3 Study Results and Conclusions ............... 5 2 CONCEPT DEFINITION .................................... 11 2.1 Satellite Configuration ....................... 11 2.2 Power Conversion ............................... 32 2.3 Power Transmission ........................... 62 2.4 Attitude Control/Stationkeeping ............... 88 3 ORBITAL OPERATIONS .................................... 115 3.1 Structures Fabrication ....................... 115 3.2 SPS Construction Jig/Base........................ 118 3.3 Mirror and Solar Cell Blanket Fabrication Facilities...........................................122 3.4 Rotary Joint, Slip Ring Structure, and Electrical Track Buildup ...................... 124 3.5 Microwave Antenna Trunnion Structure Buildup . 126 3.6 Microwave Antenna Assembly .................... 126 3.7 Installation of Electronic Elements and Phased- Array Antenna.......................................132 3.8 Maintenance Facilities . . . . . . . 136 3.9 Crew Sizes and Assembly Sequence .... 138 4 TRANSPORTATION ........................................ 139 4.1 Earth Launch Vehicle ......................... 139 4.2 Orbital Transfer ............................... 144 4.3 Crew and Resupply Module . 149 4.4 Payload Integration ........................... 151 5 PROGAMMATICS ............................................ 155 5.1 Program Development ........................... 155 5.2 Economic Comparisons ........................ 157 5.3 Technology Advancement ........................ 174 6 SPECIAL STUDIES ........................................ 193 6.1 Nuclear Radiation at Geosynchronous Orbit . . 193 6.2 Electrical Spacecraft Charging at Geosynchronous Orbit...............................................200 6.3 SPS Microwave Radiations . . . . . . 205
ILLUSTRATIONS Figure Page 1.1- 1 Study Overview............................................ 2 1.2- 1 SPS Reference Configuration ............................ 4 1.3- 1 Assembly Schedule - Nth Satellite ........................ 7 1.3- 2 Air-Augmented HTO-SSTO Concept ........................ 7 1.3- 3 SPS Cargo Traffic Model Cumulative Cargo Masses to Orbit . 8 1.3- 4 Program Cost Relationships................................ 9 1.3- 5 System Cost by Fiscal Year................................ 10 1.3-6 Future Study Recommendations ............................ 10 2.1-1 SPS Reference Configuration ............................ 12 2.1-2 Configuration Alternatives ................................ 13 2.1-3 SPS Frame Girder Generation Sequence .................... 14 2-1-4 SPS Structural Geometry................................ 15 2.1-5 SPS Satellite Structure................................ 16 2.1-6 Girder Intersection Variations ....................... 16 2.1-7 Beam Interface Attachments................................ 17 2.1-8 Akron and Macon Frame Structure ....................... 18 2.1-9 Akron and Macon Partially Assembled .................... 19 2.1-10 Hindenberg Partially Assembled ....................... 20 2.1-11 Beam Element Section for SPS Array Structure .... 21 2.1-12 Built-up Beam Element Section for SPS Array Structure . . 21 2.1-13 Normalized Sections .................................... 22 2.1-14 Existing MPTS Concepts ... ....................... 23 2.1-15 Structural Concept for Compression Frame/ Tension Web MW Antenna ........... ........ 24 2.1-16 Tension Web/Compression Frame Interface ................ 24 2.1-17 Design Loads ............................................ 26 2.1-18 Compression Frame Analysis Technique .................... 27 2.1-19 Compression Frame Girder Configuration .................... 28 2.1-20 Tribeam Personnel Scooter ................................ 29 2.1-21 Catenary Rope Design .................................... 30 2.1-22 Tension Web Design .................................... 30 2.2-1 SPS Efficiency............................................ 32 2.2-2 Solar Cell Spectral Response Comparisons ................ 35 2.2-3 Typical 1 x 1 cm GaAlAs/GaAs Heteroface Solar Cell ... 36 2.2-4 I-V Characteristics of 1 x 1 cm^ Cell Measured on X-25 Solar Simulator............... 37 2.2-5 Solar Cell Efficiency Projections ........................ 37 2.2-6 Solar Cell Efficiency.................................... 38 2.2-7 Normalized Solar Cell Maximum Power Versus 1-Mev Electron Fluence .................... 39 2.2-8 Fractional Degradation in Short Circuit Current Versus 1-Mev Equivalent Fluence ............ 41 2.2-9 Experimental Radiation Effects Data .................... 41
2.2- 10 Temperature and Radiation Effects on Solar Array Performance.................... . . . 42 2.2- 11 Efficiency Versus Thickness (Optimum Conditions) ... 43 2.2- 12 Solar Cell Configurations ................................ 45 2.2- 13 Solar Cell Blanket Configuration ........................ 45 2.2- 14 Vee Trough Solar Cell Reflector System - Concentration Ratio and Dimensions as Function of Reflector Angle . . 47 2.2- 15 Vee Channel Solar Array Concentration Ratio as Function of Misorientation Angle .... . . 48 2.2- 16 Simplified Solar Array Configuration .................... 49 2.2- 17 Panel and Reflector Steady-State Temperatures .... 51 2.2- 18 Microwave Antenna Trunnion Structure Buildup .... 52 2.2- 19 Microwave Antenna Hexagonal Compression Frame .... 52 2.2- 20 Shoe and Slip Ring Drive Assembly........................ 53 2.2- 21 Pickup Shoe............................................ 54 2.2- 22 Solar Array Conductor Network (One Quadrant) .... 57 2.2- 23 Schematic of Solar Array Conductors .................... 57 2.2- 24 System Voltage Drops ............................ 58 2.2- 25 Conductor Cross-Section Area for Each of 4 SPS Modules . 58 2.2- 26 Total Array Conductor Weight of SPS .................... 59 2.2- 27 Conductor Weight ........................................ 59 2.2- 28 Power Distribution Lateral Power Flow .................... 60 2.3- 1 Conceptual Diagram Showing Array and Amplifier Cost Trends Versus Amplifier Size................... 63 2.3-2 Simplified Diagram of Klystron Mod-Anode and Beam Power Requirements........................ 64 2.3-3 PPM/PM High-Efficiency Klystron CW Amplifier .... 66 2.3- 4 Microwave Energy Conversion Efficiency Chain .... 68 2.3- 5 Efficiency Breakdown - Transmitting Antenna .... 68 3.2- 6 Raytheon MW Subarray Concept ............................ 69 2.3- 7 Typical TE^q SWR Array.................................... 70 2.3- 8 Amplitron Modified Heat Sink............................ 72 2.3- 9 Experimental RCR ........................................ 73 2.3- 10 Radiating Slot Plane (H-Plane) ........................ 74 2.3- 11 Feedline Plane (E-Plane) ................................ 74 2.3- 12 Far-Field Radiation Pattern (10-Square-Meter Subarray) . 75 2.3- 13 10-Square-Meter Element Factor ........................ 75 2.3- 14 RCR Element Maintenance ................................ 76 2.3- 15 Low-Density 10-Square-Meter Subarray .................... 77 2.3- 16 Low-Density 30-Square-Meter Subarray .................... 78 2.3- 17 High-Density 10-Square-Meter Subarray .................... 79 2.3- 18 High-Density 30-Square-Meter Subarray .................... 80 2.3- 19 MW Antenna Modular Sizes ......................................82 2.3- 20 Functional Block Diagram of Hytrodyne Phase Control Network ............................ 84 2.3-21 SPS MPTS Transmit Array Geometry Using Space Feed Reference System.................... 85 2.3- 22 Microwave Power Transmission System .................... 87 2.4- 1 100-cm Ion Thruster Single Cathode Concept .... 89 2.4- 2 Summary of Engine Requirements ........................ 89 2.4- 3 Inert Gas and Tankage Mass................................ 91
2.4- 4 Propulsion Requirement for Orbit Inclination Control . 95 2.4- 5 Satellite Motion ........................................ 97 2.4- 6 Canted Vehicle................................................ 103 2.4- 7 Cant Angle Comparisons.................................... 103 2.4- 8 Momentum Storage Requirements ............................ 105 2.4- 9 Capabilities of Existing Devices ........................ 105 2.4- 10 Momentum Wheel Design Parameters ........................ 106 2.4- 11 Large Momentum Wheel Conceptual Design .................... 106 2.4- 12 Analysis of Problem .................................... 107 2.4- 13 Periodic Motion ........................................ 109 2.4- 14 Required Initial Conditions ............................ 110 2.4- 15 Stable Oscillatory Motion ................................ 110 2.4- 16 Control Requirements .................................... 112 2.4- 17 Configuration B Periodic Motion ........................ 113 3.1- 1 SPS Equilateral Triangle Beam Verendeel Truss-Formed Sheet . 116 3.1- 2 SPS Structural Element Fabricator ........................ 117 3.1- 3 Tri-Beam Fabrication........................................ 118 3.2- 1 SPS Construction Jig/Base Complex ........................ 119 3.2- 2 Central Construction Base.................................... 119 3.2- 3 Intersection Facility - Perspective View ................ 120 3.2- 4 Intersection Facility - Orthogonal Views ................ 121 3.3.-1 SPS Construction Facilities Operational.....................122 3.3- 2 Solar Cell and Reflector Film Deployment Facility Center Vertex Crawler .................... 123 3.3- 3 Film Deployment Concept Cable System .................... 124 3.3- 4 Tribeam Personnel Scooter ................................ 125 3.3- 5 Reflector and Solar Cell Film Installation .... 125 3.4- 1 Rotary Joint, Slip-Ring Structure, and Electrical Track . 126 3.4- 2 Rotary Joint, Slip-Ring Structure, and Electrical Track Buildup................................127 3.5- 1 Microwave Antenna Trunnion Structure .................... 127 3.5- 2 Microwave Antenna Trunnion Structure Buildup .... 128 3.5- 3 Beam Interface Attachments.................................... 129 3.6- 1 Microwave Antenna Hexagonal Compression Frame .... 130 3.6- 2 Microwave Antenna Hexagonal Compression Frame Buildup . . 130 3.6- 3 Nth Satellite Assembly - 52nd Day..............................131 3.6- 4 Microwave Antenna RF Element Installation - Negative Lens Configuration .................. 131 3.6- 5 Assembly Schedule - Nth Satellite............................ 132 3.7- 1 Central 30-Meter by 30-Meter Subarray In-Line Assembly Facility........................................133 3.7- 2 Microwave Antenna 30-Meter by 30-Meter Subarray Modular Assembly and Installation ................ 134 3.7- 3 Microwave Antenna Modular Sizes ...................... 134 3.7- 4 Microwave Antenna Elements - 3-Membrane Negative Lens Concept . ........ ................... 135 3.7- 5 Microwave Antenna Modules and Catenary Structure Assembly Sequence ........................ 135 3.7- 6 Antenna Subarray Deployment Sequence .............. . 136 3.8- 1 Microwave Antenna Maintenance Concept .................. 137
4.1- 1 Activity Matrix Cost Characteristics .................... 140 4.1- 2 Candidate Earth Launch Vehicle Concepts ................. 140 4.1-3 Cost per Flight Comparisons.................................142 4.1- 4 Air-Augmented HTO-SSTO Concept .......................... 143 4.2- 1 Common Stage LO2/LH2 Concept .......................... 146 4.2- 2 OTV Flight Profile .........................................147 4.2- 3 OTV Performance Capabilities .......................... 148 4.3- 1 Passenger Module Mass Trend (No. of Passengers Versus Mass/Man) ........................ 150 4.3- 2 Crew and Resupply Module...............................150 4.4- 1 Payload Integration .................................... 151 4.4- 2 SPS Cargo Traffic Model - Cumulative Cargo Masses to Orbit . 152 4.4- 3 SPS Cargo Traffic Model Supply and Demand Rates . . . 152 5.1- 1 SPS Program Breakdown Structure ........................ 155 5.1- 2 SPS Program Development...............................156 5.2- 1 Program Cost Relationships . 161 5.2- 2 System Cost by Fiscal Year...............................165 5.2- 3 Economic Comparison - Example SPS Cash Flow Based on Estimated Investment Costs and Funds Generated by 4 Satellite Installations ........................ 173 6.1-1 Van Allen Belt Particle Dose...................195 6.1-2 Solar Flare Particle Dose.......................195 6.1-3 Galactic (Cosmic) Particle Radiation Dose ..... 196 6.1-4 Total Nuclear Radiation Dose ........................... 196 6.1-5 Low-Altitude Van Allen Orbit Doses ........................ 197 6.2-1 Local Time Dependence of Phenomena ....... 201 6.2-2 Secondary Electron Effects ................................. 203 6.3-1 Structure of Earth's Ionosphere ........................ 208 6.3-2 RF Frequency Interference . 209
TABLES Table Page 1.2- 1 SPS Concepts...................................... 3 1.3- 1 Satellite Weight Estimates ................................ 5 2.1- 1 Compression Frame Girder Design ................ . 29 2.1-2 Compression Frame/Tension Web MPTS Weight Statement ... 31 2.2-1 SPS Solar Cell and Blanket Preliminary Specification . . 34 2.2- 2 Solar Cell Description and Weight.......................43 2.2- 3 SPS Reflector Preliminary Specification .................... 46 2.2- 4 Optical Properties ........................................ 49 2.2- 5 Array Temperature as Function of Substrate Emissivity . . 50 2.2- 6 Slip Ring and Brush Block Preliminary Specification ... 55 2.2- 7 Power Distribution Weights ................................ 61 2.2- 8 Design Concepts for Conductor Weight Reduction .... 61 2.3- 1 High-Efficiency 50-kw Klystron CW Amplifier - Estimated Weights for Different Schemes ..... 66 2.3- 2 Theoretical Power Saving of RCR Over Conventional Standing Wave TE10 Slotted Arrays ............. 71 2.3- 3 Effects of Linear Phase Error ............................ 82 2.4- 1 Satellite Propellant Requirements ........................ 90 2.4- 2 Reference Attitude Control System Mass Summary .... 92 2.4- 3 Configuration A Advantages ................................ 112 2.4- 4 Configuration B Advantages ................................ 113 4.1- 1 SPS Program Applicability - Costs.............................143 4.1- 2 SPS Program Applicability - Cost Projections .... 144 4.2- 1 Dual Ascent/Return AV Budgets.................................148 4.3- 1 Crew Rotation/Resupply Logistics Profile ................ 149 5.2- 1 SPS Cost Summary.............................................158 5.2- 2 Preliminary Economic/Financial Findings Based on Capital Investment ......................... 169 5.2- 3 Economic Comparisons - SPS Cash Flow Performance Summary . 170 5.2- 4 SPS Cash Flow Summary Based on Estimated Investment Costs and Funds Generated by Profit and Depreciation and Amortization (Case II)................................ 171 5.2- 5 SPS Cash Flow Summary Based on Estimated Investment Costs and Funds Generated by Profit and Depreciation and Amortization (Case III) ..................... 172 5.3- 1 Technology Advancement Requirements ........................ 175 6.1- 1 Time Dependence of Nuclear Radiations .................... 193 6.1- 2 Recommended Astronaut Dose Limits ........................ 198 6.1- 3 Allowable Times Due to Radiation Doses .................... 198 6.2- 1 History of Charging-Induced Phenomena . . . . . . 201 6.3- 1 SPS Microwave Radiations.....................................205
ABBREVIATIONS ACS attitude control system Ag silver Al aluminum AMO air mass zero: No atmospheric filtering is solar cell performance measurements APS auxiliary propulsion system Au astrononical unit: Mean distance from earth to sun ( 92,900,000 miles) BOL beginning of life CER cost estimating relationship CRM crew and resupply module Cu copper CW continuous wave DBM decibels referenced to 1 milliwatt DPL diplexer ELV earth launch vehicle EM electromagnet: A type of Klystron EOL end of life EROI earth return orbit insertion EVA extra - vehicular activity FEP fluorinated ethylene propalene (teflon) FSTSAS Future Space Transportation Systems Analysis Study (Boeing) Ga gallium GaAs gallium arsenide GaAlAs gallium aluminum arsenide GEO geosynchronous equatorial orbit GFY government fiscal year GHz gigahertz GLOW gross lift off weight GSE ground support equipment GSTOI geosynchronous transfer orbit Insertion GW gigawatt HLLV heavy lift launch vehicle HTO-SSTO horizontal take off - single stage of orbit IOC initial operational capability I specific impulse (thrust/propellant weight flowrate) sp
KEV kilo-electron volt Kwh kilowatt hour LEM lunar module LEO low earth orbit LEOI low earth orbit insertion LSB lunar space base MeV million electron volts MIC microwave integrated circuit Mo molybdenum M0S2 molybdenum disulfide MPD magneto-plasma dynamics MPS main propulsion system MPTS microwave power transmission system MSS modular space station MW microwave NBF narrow band filter OLS orbiting lunar station O&M operations and maintenance OSC oscillator OTV orbit transfer vehicle PD Power divider PLL phase lock loop POP perpendicular to the orbit plane PPM/PM periodic permanent magnet/permanent magnet: type of klystron QF quality factors RBE relative biological effectiveness RCR resonant cavity radiator rem Roentgen equivalent man RF radio frequency ROI return on investment SB space base SbF5 antimony pentafluoride SE&I systems engineering and integration SEPS solar electric propulsion stage Si silicon SPS satellite power system SSME space shuttle main engine SWR standing wave radiator
Ta tantalum TBD to be determined TBMA time division multiple access (time share electronics) TEm,o transverse electric: mode m, 0 (1,0 is waveguide) TFU theoretical first unit Vdc voltage, direct current VSWR voltage standing wave ratio: a measure of reflected power CONTR phase controller
1. INTRODUCTION Studies conducted by government agencies and by industry have clearly delineated the urgency for developing new energy-producing sources to meet the ever-increasing future demands for electrical power. The controversies surrounding high oil prices and limited reserves, atmospheric pollution, and nuclear safety - whether justified or not - have served to bring the energy crisis to a national awareness. It is logical, therefore, that attention is turning to the potentiality of tapping the unlimited resources of the sun. Many ingenious approaches have been postulated and some are technogically feasible. Most, however, cannot meet reasonable economic criteria as evidenced by the current lack of large capital investments for operational systems. For the past few years, the aerospace industry, under the guidance and funding of joint NASA/ERDA teams, has been investigating the practicality of intercepting sunlight in space, converting it to electricity, and - via microwaves - transmitting it to earth for subsequent reconversion and distribution over typical earth-based transmission lines. Results from these preliminary studies are encouraging. They show that with proper emphasis, judicious planning, and adequate funding, the projected mid-1980's technology would suffice to develop such systems that are both technically feasible and economically viable. Historically, development of new space systems from a given technology base takes from five to fifteen years; therefore, it does not appear unrealistic to estimate an initial operational capability (IOC) date for the satellite power system (SPS) in the mid-1990's. The level of contribution SPS could make in meeting the projected national energy demands will ultimately depend on its degree of cost competitiveness and other key criteria. A reasonable goal has been established for SPS to supply about 25 percent of the projected electrical energy demands in the year 2025. In terms of energy production capacity, this converts to a requirement of about 600 gigawatts. In general, these technology levels, IOC dates, and power generation requirements constitute the broad NASA/ERDA guidelines for this and other recent studies. 1.1 STUDY APPROACH The Space Division of Rockwell International proposed to conduct an extensive four-month technical effort to study the technical feasibility and economic viability of SPS based on an end-to-end analysis of "reference" designs and operational concepts. A functional task overview of this study is depicted in Figure 1.1-1. The first task defined a satellite and each of its systems in sufficient detail to develop appropriate design, performance, and cost data. The second task required analyses of on-orbit operations in order to develop assembly schedules, define required crew complements, design construction base concepts, and to provide a basis for costing these operations and hardware. Having defined the on-orbit satellite system and described the processes and equipment needed for its assembly, the third task required definition of the transportation systems, traffic models, and costs associated with placing in orbit all of the SPS hardware elements and the equipment and
Figure 1.1-1. Study Overview crews needed for assembly. As indicated in the figure, constant iteration was required between these tasks throughout the study. Finally, in the fourth task, cost were compiled, a representative program schedule was developed, and the technical disciplines polled to define critical technology advancements required. 1.2 STUDY GROUND RULES AND ASSUMPTIONS NASA has analyzed potential SPS configurations and systems both in-house and utilizing contractor support. The reference configurations provided conceptual approaches to SPS that were judged to be within a cost-competitive regime. These configurations are called reference concepts as opposed to baseline concepts since the latter designation implies that a relatively firm set of design requirements has been established. Rather than duplicate analyses of existing concepts, Rockwell proposed to study alternatives which offered unique design and operations features not yet pursued in other studies. These alternative features and the rationale behind their choice were described in the proposal. In each case, the alternative feature proposed by Rockwell appeared to offer added cost-effectiveness over the present reference concepts. Further, they represented appropriate alternatives which would be analyzed to generate requirements, resolve design, operational, and developmental issues, and identify problems common to establishing the technical feasibility and economic viability of all concepts. The alternatives adopted for study are tabulated in Table 1.2-1.
Previous studies of photovoltaic satellite power systems had concentrated on a low aspect ratio configuration using silicon solar cells with assembly taking place in low earth orbit. The fully assembled satellite was transferred from LEO to GEO using advanced technology electric propulsion systems. Early results from investigations conducted at Rockwell's Science Center were indicating a great potential for gallium-aluminum arsenide (GaAlAs) solar cells in the areas of thin-film cell development possibilities, higher theoretical efficiencies, and strong resistance to radiation damage. The high aspect ratio configuration alternative was selected for study primarily for two reasons. First, the long, thin structure allowed the design of a microwave (MW) antenna rotary joint which would not require microwave transmission through structural elements. Second, preliminary calculations showed that significant reductions could be achieved in the annual propellant requirements for attitude control. The SPS configuration used as a reference concept is shown in Figure 1.2-1. The selection of equatorial geosynchronous orbit as the reference satellite construction location using conventional chemical systems (e.g., LO2/LH2) appeared to be a good compromise between the advanced GaAlAs solar-cell technology assumption and the existing propulsion technology. Furthermore, the uncertainties surrounding the impact of day-night cycles every hour and a half and drag effects would not be encountered at GEO. Also the concerns of possible collisions, revenue losses incurred with long-duration transfers, and all degradation due to Van Allen belt radiation would be minimized. Table 1.2-1. SPS Concepts
Figure 1.2-1. SPS Reference Configuration In-depth studies had been conducted by Raytheon for NASA of the microwave power transmission system (MPTS), and Rockwell proposed to use those concepts, weights, and cost data for this study. Design concepts for the MW antenna structure, however, were studied in two other NASA-funded efforts by Grumman Aerospace Corporation and the Martin Marietta Corporation. Both studies had developed a concept for a box-grid matrix-type structure to be assembled in orbit using relatively short-length beams that were fabricated either with on-orbit automated modules (Grumman) or pre-fabricated on the ground (Martin-Marietta). In both cases, on-orbit assembly times appeared as an area where significant reductions could be made by taking a different design approach. The reference structural concept proposed was a compression frame-tension web configuration which was under company-sponsored investigation. The compression frame structure would be roll-formed in orbit with a continuous-beam fabricator concept that was also being studied at Rockwell. Within the stated study guidelines, ground rules and assumptions, two variations occurred during the study; both resulting in added data. The first of these is an investigation conducted by the Autonetics Division of Rockwell to develop alternative approaches to the microwave power transmission system. This is discussed in Section 2.3. The second study addition was introduced when considering the transportation system requirements for SPS. An earth launch vehicle (ELV) concept new to SPS studies has been selected for its performance potential and operational versatility (Section 4.1). It should be noted that each of these alternatives represents major design challenges and needs considerable in-depth study. They are included in the study, however, because of the highly significant advantages to an SPS program should future efforts prove them feasible.
1.3 STUDY RESULTS AND CONCLUSIONS The most significant conclusion to be drawn from this study is that there appear to be no insurmountable technological, design, operational, or economic barriers which should deter or delay continued investigations into the viability of satellite power systems as a strong contender for future production of electrical energy. If the nominal environmental impacts of microwave radiation are determined to be negligible or well within acceptable limits, and if radio frequency pollution can be effectively circumvented (refer to Sections 6.2 and 6.3), then SPS might well prove to be the most beneficial means of power production. Since pursuit of the technology issues applicable to SPS are directly supportive of other space programs and national interests, it is recommended that a case for continued investigations at an accelerated level could be made. The results of the first study task, Concept Definition (Section 2), support the choice of GaAlAs as a preferred solar cell candidate for photovoltaic systems. Inasmuch as the reference satellite configuration with a concentration ratio of two may represent a more conservative design approach (as opposed to a higher concentration ratio configuration), the weight estimates shown in Table 1.3-1 should inherently contain some margin for weight growth. Other areas where significant weight reductions might be achieved as a result of continued study are the power distribution network and the microwave antenna. Reconfiguration of the satellite to achieve a lower aspect ratio and development of electrical subsystems to handle 40,000 volts de could reduce wiring weights by as much as 50 percent of those shown. Table 1.3-1. Satellite Weight Estimates
Should the Rockwell resonant cavity radiator (RCR) eventually prove successful, then by integrating amplitrons into the backside RCR panel for heat rejection, as much as 1.5 x 10$ kg additional weight savings may be effected. New concepts for attitude control and vehicle orientation were also introduced which hold promise for minimizing satellite weight. In summary, the reference configuration, systems, and satellite weight should represent realizable design goals. These results from the first task are not, however, without engineering challenges. The development of high-purity, very thin GaAlAs cells which can be produced as large blankets at low cost is a critical technology. Providing proof-of-concept for phase-control of the 1-km diameter microwave antenna is yet another and may represent the most difficult technological challenge to SPS feasibility. Assembly at geosynchronous oribt—the second major study task, Orbital Operations (Section 3)—poses no design or operational problems over LEO construction other than providing for increased radiation protection and minimizing extravehicular activities (refer also to Section 6.1). In the assembly concepts selected, it was assumed that EVA activities would be needed only for unforeseen failure modes which might require "hands-on" actions not designed into the equipment provided. The results from this study task show that the very large sizes and numbers of subsystem elements readily lend themselves to consideration of continuous-flow assembly processes. With these processes, relatively simple machinery operating continuously at appropriate periods and at very reduced rates (as compared to earth-based production rates) can result in greatly reduced assembly times and minimal crew size requirements. Concepts were developed and machine operating times were typically only 15 percent of the total assembly time requirements. Full 24-hour assembly operations were assumed and crew size estimates were made. Interestingly, major portions of the crew complements were devoted to performing on-orbit logistics functions, and it was determined that three construction bases were required for assembly of an SPS. Iteration of the logistics demands/ capabilities with construction processes showed that a staggered assembly schedule was needed wherein initiation of assembly of the satellite structure preceded solar blanket, reflector, and MW RF element installations by approximately a month. Figure 1.3-1 illustrates the resulting schedule. If the basic concepts and processes developed in this task are eventually proven by orbital tests, an assembly rate of four satellites per year—on a 90-day cycle—appears feasible. An earth launch vehicle (ELV) concept new to SPS studies was introduced in the third study task, Transportation (Section 4): a horizontal takeoff, single-stage-to-orbit (HTO-SSTO) vehicle using a combined airbreathing-rocket propulsion system. The ELV (Figure 1.3-2), is an uprated version of a design study initiated at Rockwell just prior to the decision that Space Shuttle studies would be limited to vertically launched concepts. Although insufficient hard data exist today as to its development feasibility (the earlier studies were dropped eight years ago), comparisons with other postulated vehicles indicate that this concept would be equally cost-effective for an SPS program.
Figure 1.3-1. Assembly Schedule - Nth Satellite Figure 1.3-2. Air-Augmented HTO SSTO Concept
As important, however, is the consideration of its applicability to programs of much lesser demand than SPS. Additionally, the versatility of the concept is particularly desirable for future missions. Using the airbreathirig HTO-SSTO vehicle as a reference, a fully reusable common-stage, LO2/LH2 orbital transfer vehicle (OTV) was selected from NASA-contracted studies and sized for compatibility with the ELV. Since the reference ELV can return with payloads, the empty OTV stages are returned to earth for subsequent refurbishment, thus extending their operational life cycles. The integration of cargo payloads was investigated and payload "mixes" were defined which satisfied the massdensity capabilities of the ELV and the construction sequence demands of the SPS. A traffic model which meets these requirements is shown in Figure 1.3-3. Figure 1.3-3. SPS Cargo Traffic Model Cumulative Cargo Masses to Orbit The first three tasks are directed toward establishing technical feasibility; however, they were conducted with a constant awareness of the significance of the fourth task, Programmatics (Section 5). Results of this task— the "bottom line" of economic viability—have yielded insights into the practicability of SPS as an economic venture. Cost estimates for the elements of the program are shown in Figure 1.3-4 based on the 1985 technology projections and construction of 120 5-Gw satellites over a 30-year period at a rate of four per year with a first-unit IOC date of 1995. The total program cost estimate, including $57.7 billion for DDT&E, is $850 billion. This equates to an average capital investment requirement of $1,400 per kilowatt. Under these investment conditions, user chargers could range from 30 to 50 mills/kwh.
Figure 1.3-4. Program Cost Relationships Much added study is needed to refine these cost estimates; nevertheless, there are areas indicated on the figure where, with continued efforts, cost reductions appear possible (e.g., launch support and logistics). Application of these cost estimates to a representative development schedule results in the funding curves shown in Figure 1.3-5. The peak funding level shown for DDT&E is, in actuality, within the same relative cost realm today as was funding for the Apollo Program in the mid-1960's. As indicated previously, results from each of these study tasks are quite encouraging. Recommended areas for future emphasis are shown in Figure 1.3-6. This study, like others preceding it, suffers primarily from lack of data that can come only from future verifying research, development, and test. Yet in each succeeding study of SPS to date, new ideas, concepts, and processes are developed which continually refine and reinforce the preceding results.
Figure 1.3-5. System Cost By Fiscal Year Figure 1.3-6. Future Study Recommendations
2. CONCEPT DEFINITION The study of SPS begins with a preliminary definition of the on-orbit satellite concept. This includes the overall configuration with its major structural systems and the power conversion, power distribution, power transmission, and attitude control/stationkeeping systems. The study of these systems and the evolution of design concepts was the objective of the concept definition task. Since orbital construction of the satellite will be required, the definition of equipment, facilities, crew sites, assembly schedules, and of the transportation systems required to support these operations must be considered in the design of a satellite concept. The iteration requirement between this and the other study tasks will become more apparent as each succeeding task is discussed. Finally, as the satellite configuration and its systems concepts progressed, an effort was made to reach realistic compromises between technology projections and costs. Wherever possible, the emphasis was on the conservative side. This approach is reflected in discussions of the satellite configuration, the major structural systems, and the wiring for power distribution. 2.1 SATELLITE CONFIGURATION The selection of a satellite configuration is tied inextricably to each of the satellite systems and to both assembly and operational processes. Alternatives to the reference concept selected for study were investigated. Although some configurations appeared promising for further study, it was determined that the reference concept should be analyzed—end-to-end—throughout the study. This decision was based on the rationale that the basic configuration can be analyzed for its apparent advantages yet it represented a more conservative design approach. The potential of other configurations is discussed in areas of the material to follow. 2.1.1 Configuration Concepts The reference concept for a 5-gigawatt satellite adopted for study is shown in Figure 2.1-1. This two-trough configuration was chosen for several potentially promising reasons. First-order analyses indicate that the long, deep structure reduces the energy required to control gravity gradient forces substantially over that of a wide, thin configuration. Due to its basic shape, even greater reductions may be achievable by redistributing some of the masses. At some point in the reduction - as yet undefined - the introduction of momentum-storage devices may eventually prove practical, thus alleviating the propellant resupply demands and significantly reducing the propulsion effluents emitted in the vicinity of reflectors, solar cells, high-voltage wiring, and amplitrons. Similarly, deep structures tend to achieve greater first-mode bending and torsional stiffness. Substantiating analyses of these potential advantages are discussed in Section 2.4. Also, to remove the problem of transmitting RF power through transparent structure, solar blankets, reflectors and their supporting structures are removed at the center of the satellite and
Figure 2.1-1. SPS Reference Configuration the microwave antenna is supported on a track-like ring, as indicated in the lower portion of the figure. The antenna is rotated about the large hexagonal structure which carries through the entire length of the satellite. Design concepts were generated for the antenna track ring and for power transfer across this rotary joint. However, a major disadvantage of the high-aspect- ratio concept was uncovered early in the study. Power distribution wiring weights were increased significantly over designs which have a more equal length-to-width ratio. Some configuration alternatives were briefly investigated based primarily on design variations for achieving higher concentration ratios and are shewn in Figure 2.1-2. A major concern with higher concentrations designs is in maintaining contour of the shaped reflectors. It was determined that to maintain a parabolic shape requires a more complex array structure with added rigidity. A contour may be formed by periodic attachments (a minimum of 5 points are required) of "tuning" wires, or similar devices, to the supporting substructure. In addition, the surface must be constantly supported in biaxial tension to assure shape retention. Other configuration alternatives are discussed in Section 2.4; however, for reasons stated above, the configuration of Figure 2.1-1 carried as the reference design. 2.1.2 Satellite Structure A library search of very lightweight structures was made to provide historical structural data from 1900 to 1976. Finite element theory using the NASTRAN computer program was used to analyze the steady-state thermal stresses
Figure 2.1-2. Configuration Alternatives (worst case) produced on a single frame of the SPS structure. Finally five structural shapes were compared (normalized to have equal cap area and equal perimeter) to determine which shape would be most effective structurally for the reference SPS configuration. The general properties sought were high allowable stresses per unit mass where column, local crippling, general instability and fatigue allowables are equal. The structural shape must also be capable of mass production such that all structural centroids would be concentric at all levels of assembly. Finally all interconnections of structural elements were to have end-fixidity coefficients from 1.4 to 1.6 using quick-connect assembly concepts. Physical Description Rockwell SPS Structure The reference configuration is a two-trough, solar photovoltaic SPS. The overall size is 2.15 km wide, 1.386 km deep, and 26.7 km long. Each trough contains two aluminized polished Mylar/Kapton mirrors set at 60 degrees to base of the trough which is the photovoltaic solar cell collector surface. The solar cell surface receives one solar sun of energy by direct exposure and one solar sun of energy from the inclined mirror surfaces. The resultant two solar sun energy level is uniformly incident to all areas of the photovoltaic surface when orientation of the SPS is normal to solar rays. The troughs are supported by a central hexagonal structure as shown in Figure 2.1-3. Figure 2.1-4 illustrates the locational arrangement of the basic roll-formed sheet metal structural element, the next larger tri-beam girder, and the longitudinal and lateral location of the girder structure. This structure is comprised of 11.longitudinal main structural columns and 14 interconnecting tribeams which form the frame support structure for the longitudinal tribeam columns.
Figure 2.1-3. SPS Frame Girder Generation Sequence A portion of the exposed framework, including tension-element crossbracing is illustrated in Figure 2.1-5. The tribeam girder is used as an SPS longitudinal column, and the basic sheetmetal triangular element is used as column elements and as support elements for the columns. The column and support element material gages are determined by environmental loading or assembly loads, whichever are most critical. Pretensioned cross-bracing by small diameter composite or metallic "wires" stabilizes the square bays formed by the tribeam column element and column support element for a tribeam section. Likewise, the square panels formed by the major column tribeams and the major frame tribeams are cross-braced by pretensioned wires. The amount of pretension in the cross-bracing is such that slack wire brace conditions are avoided. A typical frame station structure for SPS is illustrated in Figure 2.1-6. Typical major intersections of the tribeam structure are shown to indicate the orientation of individual elements. Cross-brace wires are not shown for reasons of clarity. Each type of major structural intersection was thoroughly analyzed for member orientation versus simplicity of fabrication and assembly. The results of the study are shown on Figure 2.1-7. The least complex joint (shown upper left) results when the apex orientation of the sheet metal equilateral triangle is oriented 180° away from the apex of the tribeam truss. A simple bent tube fastened to the sheet metal element as shown permits the ends of a second sheet metal element, equipped with quick-connect fasteners to be located and secured by automatic machine means.
Figure 2.1-4. SPS Structural Geometry
Figure 2.1-5. SPS Satellite Structure Figure 2.1-6. Girder Intersection Variations
Figure 2.1-7. Beam Interface Attachments The second joint (upper center of Figure 2.1-7) involving increased complexity occurs when two tribeams intersect at 90° to each other having a common base plane and apex centerlines lying normal to each other in a plane parallel to the base plane. The resulting joint is a compound 60® bevel, 90° miter-type joint. Fortunately, the four base plane corners of the tribeam intersection are identical and are oriented progressively at 0® (360°), 90°, 180°, and 270* (i.e., 90® apart in rotational orientation). The fitting is attached at four places by fasteners to the sheet metal element at the juncture of the sheet-metal flanged braces. Quick connect fittings located on the tubular structure and the ends of the two intersecting sheet metal elements permits automatic assembly of the intersection. A third joint (upper right Figure 2.1-7) is the most complex joint required to produce the entire SPS frame (except at the rotating joint and transmitting antenna). This intersection occurs when tribeam trusses in the same plane cross at an angle of 60° to each other. The vertex sheet metal elements of each truss form this particular joint. Due to symmetry of reflection or inversion of this joint, all parts of the frame structure intersecting at 60° or 120° may be connected. Quick-connect fitting halves are installed in the tubular members and the ends of the sheet-metal elements to be connected such that subsequent automatic assembly can be achieved. The results of the historical search for lightweight structures is illustrated by Figures 2.1-8 through 2.1-10. The utilization of the Vierendeel truss applied to square columns by the Akron and Macon dirigibles is shown by Figure 2.1-8. Framing and wire rigging is dramatically shown in Figure 2.1-9.
Figure 2.1-8. Akron and Macon Frame Structure The multi-spring concept of supporting long, continuous column elements in triangular array is shown in Figure 2.1-10. This particular type of structure emerged during World War I, where the Germans were constructing 600-foot battle zepplins once a month, for the bombing of London. This section was also incorporated in the Hindenberg and sister ship structures from 1927 to 1929. For the Rockwell SPS structure, two structural element candidates emerged. They are shown in more detail by Figure 2.1-11, an embodiment of the Vierendeel truss in an equilateral triangular cross-section, and Figure 2.1-12, an embodiment of the multi-spring stabilization of continuous column elements in an equilateral triangular cross-section. Figure 2.1-13 is a comparison of five cross-sections normalized to have the summation of area of column cap and polygon perimeters to be equal for all five sections. The comparison is valid for dispersed material, lightweight structure in which material thickness is approximately 0.010 inch or less. The comparison appears to be a means of rapidly obtaining a structural figure-of-merit as a function of geometrical shape. With local crippling, column, and general stability structural failure modes, all occurring at equally high allowables (a requirement of lightweight structures). Included as material property parameters, the trend to triangular cross-section and multi-convolution of cap element shapes appears to be most promising for SPS structure. Further, these triangular shapes permit coaxial alignment of centroids at all assembly sequence levels. Finally, such shapes permit all column elements to obtain end-fixidity via quick-connect attachment means. Cross-bracing of columns by preloaded tensile elements, provides the lightest possible structure available for the SPS framework.
Figure 2.1-9. Akron and Macon Partially Assembled
Figure 2.1-10. Hindenberg Partially Assembled Modularization (e.g., scalar sequencing of structural shapes) provides an essential parameter for permitting automatic machine fabrication of SPS structures. Modularization includes scalar size of the initial sheet metal element (cap centerline distance of triangular section), the longitudinal pitch distance for cap braces, and the use of odd or even number of sheet-metal bays to form the tribeam truss. In order that a properly supported cap element occurs at each of the sheet-metal column intersections, an even number of bays exist between the centroids of the tribeam truss. One-half bay per each of the column support element is involved in joining. This leaves an odd number of sheet-metal element bays between the longerons. The odd-number bay requirement locates one bay symetrically about the half-length centerline of the support element. Likewise, the tribeam cap sheet-metal element must have a similar even/odd sequence so that a bay centerline occurs at the longitudinal half-length centerline of the tribeam cap. Thus, when modularization is fully
Figure 2.1-11. Beam Element Section for SPS Array Structure Figure 2.1-12. Built-up Beam Element Section for SPS Array Structure
Figure 2.1-13. Normalized Sections invoked throughout the structure, all joint requirements reduce to the three joints illustrated by Figure 2.1-7. Most importantly, these joints can be assembled by automatic beam-making machinery. 2.1.3 Microwave Antenna Structure Prior to initiation of this study, Rockwell had conducted a review of microwave antenna concepts developed by Grumman (Contract NAS 3-17835) and by Martin Marietta (Contract NAS9-14319) (Figure 2.1-14). Both concepts provided a 1-km diameter rigid structure of significant depth. The results of these
Figure 2.1-14. Existing MPTS Concepts reviews indicated two major areas of concern which drove Rockwell to investigate and adopt an alternate antenna structure concept. The areas of concern are: 1. Construction Time—Both concepts required the fabrication (either on the ground or on-orbit) of thousands of individual structural beam-columns. The beam-columns would then be joined together to form a 1-km disc. This antenna structure assembly activity was estimated to take up to one year to complete. 2. Thermal Distortion—dc-to-RF conversion and transmission equipment (i.e., amplitrons/klystrons and waveguides) generates excess heat which is radiated to the antenna structure. As a result, a temperature delta exists between the two parallel surfaces of the structure. In addition, because of the Gaussian microwave distribution and resulting variation in heat flux, a temperature delta exists between the center of the structure and the circumferential perimeter. The result is structure surface distortion, and in turn, a decrease in beam efficiency. Counteracting this decrease in beam efficiency requires individual waveguide array adjustment devices which introduces additional system failures modes and system weight. With these two concerns as drivers (i.e., minimizing on-orbit construction time and structure thermal distortion), the tension web-compression frame concept shown in Figures 2.1-15 and 2.1-16 was developed. The concept as developed and analyzed is applicable to the Rockwell reference SPS configuration. Variations of this basic concept were made in the assembly operations tasks.
Figure 2.1-15. Structural Concept for Compression Frame/Tension Web MW Antenna Figure 2.1-16. Tension Web/Compression Frame Interface
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