Stud y o n European Aspects of Solar Power Satellites VOLUME I: Final Report The work described In this report was carried out under ESA contract. ; Responsibility for the contents resides in the author or organisation that prepared it. EUROPEAN SPACE AGENCY CONTRACT REPORT'.
STUDY ON EUROPEAN ASPECTS OF SOLAR POWER SATELLITES Volume I : Final Report PREPARED BY J. RUTH AND W. WESTPHAL Technische UniversitAt Berlin Institut fOr Luft- und Raumfahrt SALZUFER 17 - 19 D-1000 Berlin 10 PREPARED FOR EUROPEAN SPACE AGENCY ( ESA ) ESA STUDY MANAGEMENT : G. SEIBERT , D. KASSING UNDER CONTRACT N° 3705 / 78 / F / DK ( SC ) JUNE 1979 european space agency agence spatiale europdenne estec (european space research and technology centre) Domeinweg. 2200 AG NOORDWIJK, The Netherlands — telex 31698 tel (nat ) 01719-8 , (mt.) +31 - 1719 - 8 12 9
ABSTRACT This report presents the results of a six-months study on implications of the potential electricity supply for Western Europe by Solar Power Satellites (SPS) and of their implementation problems around the turn of the century. Three objectives in this scope have been pursued to provide a decision platform for on-going study and research activities: (a) to establish an information base including relevant data collection and basic SPS-systems understanding, (b) to identify and preliminarily assess specific European problems of SPS utilisation, and (c) to accomplish recommendations for early study activities of particular European concern. The last item is directly derived from the evaluations of the second study objective and must be seen in the context with SPS activities of the USA. The authors conclude that the special European environment calls for special European utilisation feasibility analyses, while the principle technical feasibility of a SPS-system should be taken over from U.S. studies. Moreover, if the USA stop their activities, an implementation of a SPS-sytem for Western Europe would be principally impractical in the considered time frame (2000 - 2030). Keywords: ELECTRIC POWER PLANTS / ENERGY POLICY / ENERGY TECHNOLOGY / ENVIRONMENT EFFECTS / MICROWAVE TRANSMISSION / SATELLITE POWER TRANSMISSION (TO EARTH) / SATELLITE SOLAR POWER STATIONS / SPACE PROGRAMS / SPACE TRANSPORTATION / TECHNOLOGY ASSESSMENT
ABSTRACT ii ABBREVIATIONS iv EXECUTIVE SUMMARY v INTRODUCTION 1 PART I : LITERATURE REVIEW 2 1.1 SPS SYSTEM-CONCEPTS 3 1.2 SUMMARIES ON THE STATE OF THE ART 12 1.3 BIBLIOGRAPHY 59 PART II : PARTICULAR CONCERNS OF EUROPEAN SPS UTILISATION 72 II.1 GEOGRAPHICAL CONSIDERATIONS FOR WESTERN EUROPEAN SOLAR POWER SATELLITES 73 II.2 STATUS OF WESTERN EUROPEAN ELECTRICITY SUPPLY AND DEMAND FORECASTS 92 II.3 ORBITAL LIMITATIONS FOR EUROPEAN SPS 96 II.4 ORGANISATIONAL PROBLEMS OF EUROPEAN SPS DEVELOPMENT PARTICIPATION AND SPS COMMERCIALISATION 105 II.5 SPS AND WESTERN EUROPEAN INDUSTRIES 108 PART III : RECOMMENDATIONS FOR EUROPEAN STUDY PROGRAMMES 112 III.1 GENERAL CONDITIONS 113 III.2 RECOMMENDATIONS FOR PRELIMINARY EUROPEAN SPS ASSESSMENTS 1 1 4 III.3 RECOMMENDATION FOR A MEDIUM-TERM SPS STUDY PROGRAMME IN EUROPE 120 III.4 SPECIFICATION OF NEAR-TERM STUDY TASKS I 24
ABBREVIATIONS AC - Alternate Current AMO - Air-Mass-Zero AOCS - Attitude & Orbit Control System AU - Accounting Unit BOL - Begin-of-Life COTV - Cargo Orbital Transfer Vehicle CR - Concentration Ratio DC - Direct Current DDT&E - Design, Development, Test & Evaluation DOE - (U.S.) Department of Energy EC - European Communities EMI - Electromagnetic Interference EOL - End-of-Life ERDA - (U.S.) Energy Research & Development Administration FEEP - Field Emission Electric Propulsion GEO - Geosynchronous Earth Orbit GSSPS - Gravitational Stabilized SPS GW - Gigawatt HLLV - Heavy Lift Launch Vehicle LEO - Low Earth Orbit LH2/LOX - Liquid Hydrogen / Liquid Oxygen LPTS - Laser Power Transmission System MHD - Magneto-Hydro-Dynamic MOSES - Modular Solar Energy Satellite MPTS - Microwave Power Transmission System MW - Megawatt OECD - Organisation for Economic Co-operation & Development OLL - Orbital Longitude Limitation OTV - Orbital Transfer Vehicle PLV - Personnel & Priority Cargo Launch Vehicle POTV - Personnel & Priority Cargo OTV PRS - Power Relay Satellite R&D - Research & Development RF - Radio Frequency RFI - Radio Frequency Interference SPS - Solar Power Satellite / Satellite Power System / Satellite Power Station SSME - Space Shuttle Main Engine SSPS - Satellite Solar Power Station VTOHL - Vertical Take-of/Horizontal Landing
EXECUTIVE SUMMARY The scope of this study is to establish an information base, on which the major problems of SPS utilisation in Western Europe can be identified and preliminarily assessed, and finally to accomplish recommendations for early study programmes to evaluate principle solutions of these problems. The study starts with a broad analysis of the SPS-literature of the past years to provide a basic data collection and systems understanding. Dealing with concepts which are referred to as serious competitors for a potential realisation only, several concepts with potential realisation too far in the future or with technology breakthroughs required have not been taken into consideration. Also some non-photovoltaic power conversion devices and non-microwave power transmission systems, such as Power Relay Satellites (PRS), Solettas, Lunettas etc. have not been taken into account because of their presently just global status of investigation. Of the 372 publications which have been selected initially, 84 were rejected after a first evaluation. They proved to deal with the subject only marginally. Of the remaining publications 120 were selected for the literature review in Part I of the study. The other 168 were taken as additional information only, because of their very special contents mainly dealing with investigations in the field of high efficiency solar cells and concentration devices. The selected publications have been evaluated thoroughly due to special categories which have been identified prior by the authors: Power Conversion, Power Distribution, Power Transmission (DC to DC) Structure and Attitude and Orbit Control System, Space Transportation, Space Construction and Maintenance, Programme Cost and Risk, Energy Demand, Cost Comparison, Funding, Legal Aspects, Public Acceptance and Safety, Microwave Impacts, Rectenna Impacts, Launcher Emission and Space Environment, Resource Problems, Programme Planning and Development. These categories as well as abstracts and listings of the most important data, figures and illustrations are available in the Appendix (Volume II of this report). The analysed reports were our information base for the summaries on the state of the art. These summaries are held in a concise form according to the special categories to enable the reader to inform himself roughly. Since it appeared impractical to present all detailed information in the summaries, the reader can find the numbers of the publications we referred to at the beginning of each category in Part I and, if he wants to learn more about a special category, seek them out in the Appendix. (The reader who is mainly interested in restrictions or problems typical for Europe may start with Part II). In Part II of this study a lot of differences of a potential SPS application in Europe compared to the USA were identified and evaluated. Five major issues (by the authors investigation) are to be seen and checked prior to any increasing European attempt towards SPS: Geographical Siting Requirements of the Rectenna, Energetical Situation and Demand Forecasts, Orbital Restrictions, Organisational Problems, and Influences by European Industries. A first rough evaluation of these concerns was performed by actual data and partially by authors' own estimates and by applying two different working scenarios.
Part III presents the concluded recommendations for a near-term and a medium-term study programme concerning a potential SPS-im- plementation for Western Europe. These study programmes have to find out whether .there are problems which would definitely prevent a SPS-system from being introduced to Western Europe or not, and to improve the understanding of special European problems. To accomplish early results, which should provide a decision platform for on-going activities (late in 1980), the authors identified 22 near- term study tasks amounting roughly between 9.75 and 13.5 man-years. The medium-term study tasks which mainly should be in-depth investigations of the European concerns, are pointed out more globally. They at first should be defined by the results gained from the near-term efforts. STUD/ ON EUROPEAN ASPECTS OF SOLAR POWER SAT ELUTES The authors acknowledge Dieter KASSING, Visiting Scientist at ESTEC, for his substantial contribution to this study, in particular for the weighty literature procurement and useful discussions during consultations and study reviews, where he represented ESA as a technical study advisor.
INTRODUCTION When Dr. Peter Glaser in 1968 for the first time proposed Satellite Solar Power Stations (SSPS) to generate electricity for earth, he made a big step towards realisation of an old idea of Hermann Oberth. Glaser was stimulated by the 1960 experiments of William Brown of Raytheon Co. to transmit larger amounts of energy by microwaves, and by the fact, that in certain orbits the sun is shining over 99 % of the year with full intensity without being attenuated or scattered by weather or by the day and night cycle. Since that time several conceptual designs of Solar Power Satellites (SPS) have been investigated and a very large number of studies in that field has been conducted, nearly all of them in the USA and only a very few in Europe. The support of NASA started in 1972, when the feasibility study on SPS was conducted under the leadership of Arthur D. Little, Inc. In 1977 the U.S. Department of Energy (DOE) became responsible for all SPS activities and a three-year "Concept Evaluation" programme was started with several studies in the technical areas being managed by NASA, and DOE itself directs studies in the field of social, economical, ecological and environmental impacts of SPS. This programme will end in mid 1980. Presently, there can be seen SPS to be a real chance for an ecological and economical future energy source with several advantages compared to conventional ones and others of to-day (potential less pollution and other environment burdening, unexhaus- tible energy source) .
PART I LITERATURE REVIEW - SPS - SYSTEM CONCEPTS - SUMMARIES ON THE STATE OF THE ART - BIBLIOGRAPHY
1.1 SPS - SYSTEM CONCEPTS (References: 74/1 ,74/3,77/2,77/7,77/9,77/11 ,77/15,77/16, 77/18,78/2,78/3,78/16,78/19,78/27,78/37,78/55,78/63.) 1.1.1 SPS from 1968 to 1979 The first approach in 1968 to imagine a SPS was to attach a microwave antenna to a huge solar array by a cable. The evolution of a photovoltaic SPS configuration is shown in fig. 1.1Fig. 1.1: SPS CONFIGURATION EVOLUTION The philosophy of using concentrating mirrors was to reduce the cost of the solar collector. At the end of this evolution chain (not shown in this figure) was the SSPS-concept of A. D. Little, Grumman, Raytheon and Spectrolab of 1973. This design was the first concept which was investigated deeper (refer to I. 1.2). In 1974 other engineering teams than that around Dr. Glaser started to publish their SPS-imaginations. So Woodcock, Patha, Gregory et. al. from Boeing Co. proposed solar thermal options (refer to I. 1.3). J. Ruth from the Technische Universitat Berlin made recommendations for a high-degree-modularity of the photovoltaic solar collector (refer to 1. 1.4.) . The efforts of the A. D. Little team were continued with NASA’s financial support. A lot of improvements compared to the SSPS- concept were made. At the Boeing and Rockwell companies work continued also and in 1976 NASA found it worthwhile to conduct two inhouse studies. One at MSFC and the other at JSC. The results of these evaluations led to two in depth studies from 1976 to 1978. The one was conducted by Boeing and was managed by JSC (refer to I. 1.5) and the other was conducted by Rockwell and was directed by MSFC (refer to I. 1.6). In 1976/77 a novel concept was published in the USA. It is a gravitationally stabilised photovoltaic power satellite (refer to I. 1.7). In 1977 the responsibility of all SPS-activities was transferred to the ERDA (now DOE) by the U.S.-Congress. A task group was established which defined a "Concept Evoluation" programme. This led to a three year plan of investigations in the field of technical, social, environmental, ecological, and economical aspects of SPS. One of the first milestones of this programme was a reference concent selection in late 1978 (refer to I. 1.8). Note: Technical and cost data on the following 8 pages are received from the literature. The validity of the cost figures was not evaluated in this study.
1.1.2 The SSPS Concept ( A.D.Little et al.) 1973 Ground Power Output Total Mass Solar Cell Type Efficiency with CR = 2 Microwave Generator Type Generator Efficiency System Efficiency Total Cost (Without DDT&E) Electricity Cost DDT&E Cost Characteristics: Prefabricated parts of the SSPS will be lifted to LEO by a HLLV(200-400 ton class). There the satellite will be partially mounted and these parts will be delivered to GEO by ion-thrusters. In GEO these parts will be finally integrated to a SSPS.
1.1.3 Solar Thermal SPS ( Boeing) Ground Power Output Total Mass Conversion Concentration Ratio Cycle Thermal Efficiency SPS System Cost (Without Transport Construction,Rectenna and DDT&E) Transportation Cost Characteristics:This concept consists of thermoelectric modules (dishes) of 2.5 GW each. 10000 Heliostats per dish concentrate the incoming sunlight on a cavity absorber. 14 turbines are assembled around each cavity to generate electricity. The power is transmitted to Earth by microwaves.
I. 1.4 The MOSES- Concept (TU-Berlin). (MOSES - Modular Solar Energy Satellite) Ground Power Output Total Mass Conversion Efficiency with CR = 2 Total Cost (Without DDT&E) Electricity Cost (1 973 estimate) Module Size Module Power Characteristics:In a low technology advancement scenario the complete modules will be stacked into each other and transported into GEO by a high volume HLLV. In a medium technology advancement scenario only the blankets will be fabricated on Earth. Structure and concentration mirrors will be fabricated in a space factory and then mounted to the blankets. The high technology advancement scenario proposes total space-manufacturing in an automatic facility in space. The modules will be integrated to large solar collectors of variable sizes.
1.1.5 The GSSPS-Concept ( The Aerospace Corporation ) Characteristics:This 73 km long SPS uses the gravity gradient to stabilise the attitude. 48 solar panels are linked with waveguide rotary joints to the central waveguide, where the microwave power, which is generated by solid state devices within the panels, is delivered to a lens type solid state phased array antenna.
1.1.6 The Boeing SPS-Concept 1977/78 Ground Power Output Total Mass Conversion Efficiency without CR Microwave Generator Type Generator Efficiency System Efficiency Total Cost (Without DDT&E) Electricity Cost Characteristics: Eight major pieces would be assembled in a huge LEO construction base and then shipped by electric self-propulsion to GEO. In Geo the eight modules would be integrated to the SPS with the help of a final construction base. In the operational phase each of the two microwave antennas would deliver 5 GW to a ground based rectenna. The construction philosophy is based on a general simplicity of all satellite parts.
1-1-7 The Rockwell SPS-Concept 1977/78 Characteristics: After delivery to LEO by HLLV's, dedicated electrical orbital transfer vehicles woult take the raw materials to GEO where the SPS is constructed by a crew of 500 or more. The GaAlAs cells ( silicon might be another option ) would operate at 125 °C. At this temperatures the cells will regenerate from the particle radiation damages ( self-annealing ). The construction philosophy of this SPS-concept is more sophisticated and requires some technology advancements.
1.1.8 The DOE/NASA Reference Concept 1978 ( Silicon Option ) Characteristics: Two reference concepts have been selected as a preliminary design for comparisons within the scope of the DOE - SPS activities. This option is a derivative of the Boeing concept ( compare to 1.1.6 ). Total Power has been reduced to 5 GW and so only one transmitting antenna is required. Construction location is in GEO with raw-materials transport by a dedicated electrical OTV.
The DOE/NASA Reference Concept 1978 ( GaAlAs - Option ) Characteristics are very similar to the silicon Option. The particular difference lies in the concentration of the sunlight by mirrors. Since the solar cells are operated at 125 °C , they will be self annealed ( compare to 1.2.7 ).
1.2 SUMMARIES ON THE STATE OF THE ART 1.2.1 Power Conversion (References: 76/1,76/2,76/3,77/2,78/12,78/19,78/55.) I.2.1.1 General This category implies the conversion of solar radiation power into electrical DC power within the SPS. This includes optical solar-thermo-electric and photovoltaic devices. Not implied are nuclear-thermo electric and directly solar powered laser or maser devices. 1.2.1.2 Solar-Thermo-Electric Conversion In this sub-category six methods appear worth to be mentioned : - Thermionic conversion - Combined thermionic/Brayton cycle conversion - Brayton cycle turbomachine - Rankine cycle turbomachine - Cascaded dielectric device - Magnetohydrodynamic conversion The thermionic conversion appears not to be feasible for a long reliable working period mainly because of the required high temperatures. The Brayton cycle has been considered as the most promising method within the thermo-electric conversion sub-category because the single phase working fluid precludes internal system corrosion. The cascaded dielectric conversion device is only mentioned in reports of the inventors. So it is not possible to be classified within this range. The Rankine cycle is no longer seriously considered as a competitor to Brayton conversion because of the corrosion problems due to the liquid metal working fluid. The development state of the MHD device is not sufficiently advanced, so it is not possible to criticise it. 1.2.1.3 Photovoltaic Conversion In this sub-category the following cell types must be mentioned as the most promising ones: - Single crystal silicon cells - Polycrystalline silicon cells - Amorphous silicon cells - Gallium Arsenide Cells In addition to thesezCadmium-Sulfide cells appear to have a certain potential for SPS-application.
The single crystal silicon cells are the photovoltaic devices with which the most experience has been made during the spacemissions of the recent two decades. The development state is rather high and 19.7 % efficiency has been achieved with special small-area experimental cells under laboratory conditions. Since the solar collector of a SPS is the heaviest part of the satellite, the cell weight is a very important factor. In the case of silicon, the minimum weight has been reached with the minimum thickness of 0.05 mm for optimum efficiency. Gallium Arsenide has the potential of weight savings, but this is not as much as one might expect because of its relatively high specific gravity of 5.32 compared with silicon which is 2.34. But another advantage of GaAs over Silicon makes it the most serious contender to displace Silicon as the prim photovoltaic material in the future: it has the potential of higher efficiencies - more than 20 % have been achieved under laboratory conditions - it has a lower temperature sensitivity, and its radiation resistance is much better. Disadvantages of GaAs are the facts, that this semiconductor material is binary compound and therefore more difficult and expensive to produce, and that the availability of Gallium is very scarce. Recent approaches in replacing the single crystal Silicon cells by polycrystalline or amorphous ones have shown encouraging results. Efficiencies of more than 10 % have been achieved. Two additional novel devices for photovoltaic power converters are the photoemissive cells and the beam splitter generators. It is expected to achieve efficiencies of 30 % and higher by these more complex devices with considerably higher masses. 1.2.1.4 Summary Fig-2.1.1: SILICON SOLAR CELL CROSS SECTION [ Source: Boeing ]
The two most deeply investigated SPS concepts, the Boeing and the Rockwell designs - and consequently both reference concepts - use photovoltaic devices for power conversion, although a closed cycle thermodynamic conversion is considered as a potential competitor. Boeing proposes 50 urn Silicon cells <fig. 2.1.1) without concentration. The cells are expected to perform 16.5 % efficiency at 36.5 C, AMO. This should be partially achieved by the v-shaped grooves in the c^ver glass. The blanket2unit mass is expected to be 0.427 kg/m with a cost of £ 35/n/. These cells are planned to be angealed by a laser device at 500° C (a more recent figure says 350 C, or even lower, because of severe problems expected with the higher temperature). Fig. 2.1.2: GaALAs SOLAR CELL BLANKET CROSS SECTION [Source: Rockwell ] The Rockwell team chose GaAlAs cells with a concentration ratio (CR) of 2 (fig. 2.1.2). These cells are expected to perform an efficiency of 18.2 % at 125 C (operating temperature) or 20.0 % at 28°C, AMO. The GaAlAs cells under investigation show very good annealing characteristics at operating temperature, this is called self-annealing because no extra device will be required. The specific weight of the blanket is 0.252 kg/m^ at a cost of $ 71/m . As a material of the reflectors 12.5 urn aluminised Kapton was selected with reflectivities of 90 % BOL and 83 % EOL respectively. The specific weight is 0.018 kg/m2.
I.2.2 Power Distribution (References: 77/15,77/16,78/1,78/5,78/22,78/55.) I.2.2.1 General This category implies the delivery of the electrical DC power from the solar cell modules to the microwave generators, including connections, rotary joints, conductor materials, and any regulation and control equipment. Most SPS-concepts use high-voltage DC-power with voltage levels between 20 and 44 kV. The GSSPS-concept distributes the power by microwave in waveguides. The cascaded-dielectric-de- vice concept by Drummond uses 60 Hz or DC alternatively. The high voltage levels in the DC-distribution concepts were selected considering the requirements of the microwave generators and the conductor mass. In the GSSPS-concept a 2 kV DC-subsystem was selected because of the voltage limits for the microwave transistors, which are used for RF-generation. The weight of the power distribution system should be in a range of 8 % - 14 % of total SPS mass. I.2.2.3 Conductors In this sub-category the literature shows very different approaches. It is not quite clear if the conductors shall be actively cooled as superconductors. Only the choice of aluminum as material was corresponding throughout the literature because of its low mass to conductivity ratio and good radiation cooling properties. The geometric shape of the conductors depends on the amount of waste heat to be rejected. The GSSPS concept uses cryogenically cooled aluminum conductors for the 2 kV DC-distribution and overmoded cylindrical waveguides for the microwave distribution. The waveguide material is a graphite/aluminum composite. 1.2.2.4 Interfaces There are little data available in literature concerning this sub-category. Some items are: - B£tary_joints are used as slip-rings for the high voltage transmission to the antenna. Waveguide rotary joints are used in the GSSPS together with tapered flares. - Switching_gear will perhaps be designed as superconducting devices, but these are currently limited to a few hundred amperes and a few thousand volts. Other approaches are liquid metal plasma valves or solid state gates. - DC-DC converters or regulators are identified as a severe problem for the high voltage distribution.
1.2.2.5 Efficiency The efficiency range of the power distribution systems is expected to be 88 - 96 %. I.2.2.6 Problems The most severe problems are expected with: - Interactions of the power distribution system and the space plasma, - Development of the required high-power switchgear with high reliability, - Development of conductor insulation, - Development of DC-DC regulators with high efficiency and high reliability, and - Reliability of the rotary joints. 1.2.3 Power Transmission (References: 74/1,76/2,77/4,77/15,78/3,78/4,78/5,78/18,78/37, 78/41,78/42,78/55.) 1.2.3.1 General This category implies generation, transmission, reception, and rectification of power. • I.2.3.2 Microwave Generation 1.2.3.2.1 Ampl.itrons_(Crossed Field Amplifiers! In earlier studies amplitrons were considered as the best suited candidates because of their high efficiency (estimated 88 %), their relatively low mass (2.3 kg/5 kW tube), and their potential for passive cooling. Disadvantages are: - Due to the individual 5 kW power level, a huge amount of amplitrons are needed for the antenna, - Amplitrons have to be baked out prior to full operation, so the facility in space has to be complex, - Amplitrons need a brushless motor to control the pole gap, and - Amplitrons have a relatively high RF-noise level. I.2.3.2.2 Klystrons ( Linear Beam Amplifiers ) Klystrons are currently considered as the best suited RF-gene- rators. The klvstrons have a fixed tuninn a high RF-gain with relatively low noise (50 kHz from carrier - 125 dB/kHz down) and can be easily baked out. Disadvantages are:
- Klystrons have a relatively high specific weight, - Klystrons have larger dimensions than amplitrons, - Klystrons operate at a higher temperature (400°C), - The efficiency is slightly lower than with amplitrons (86 %), and - Klystrons have to be actively cooled (heat pipes). I .2.3.2.3 Other RF-Generators Since the amplitron and the klystron are the devices which have been developed much more than all other candidate RF- generators, these will be discussed here in one sub-category. The most promising type of alternative microwave generator is a solid-state-device. Weight might be greatly reduced (0.1 kg7kW~compared-to~O.7 kg/kW for klystrons) and reliability could be significantly improved (lifetime is expected to be 30 years compared to 10 years with klystrons or amplitrons). Estimated figures for a field effect transistor generator are: voltage 10 V output 1-100 W overall efficiency 80-85 % gain 10-20 dB spec, weight 0,1 kg/kW passive cooling Similar devices are proposed for the GSSPS. Here 1,744,500 transistors are directly implanted in the surface of the 2100 m long, 4 m diameter waveguide mounted on the back of each solar panel, delivering 159 MW of RF-power. Three other types of potential RF-generators were only mentioned in literature: - the travelling wave tube, - the magnetron, and - the gyrocon. 1.2.3.3 Transmitting Antenna and Beam Control In this sub-category two substantial different approaches were found in the literature. Most SPS-concepts use antenna types consisting of slotted waveguide elements with directly mounted RF-generators. These concepts differ only in the design of sub-modules, power density distribution, etc. Typical figures for this kind of antenna are:
Fig. 2.3.1: RETRODIRECTIVE BEAM STEERING Fig. 2.3.1 explains roughly the function of an active retro- directive array. The left figure represents an undeflected beam, while in the middle the beam is missing the rectifying element of the rectenna it is attached to. On the right figure, the displacement is measured on the ground and reported to the SPS. There the phase control system is able to redirect the deflected beam. The other antenna type is the horn fed 797 m diameter lenstype solid-state phased array design used with the GSSPS. The antenna consists of 2^6 crossed dipole elements with 6-bit digital diode phase shifters and about 100 beam steering computers. Each element operates at a power level of 98 W. Another potential antenna type is the electronically steerable multibeam antenna. 1.2.3.4 The Microwave Beam The geometry of a microwave beam depends on the size of the phased array transmitting antenna. Fig. 2.3.2 shows the transmission efficiency as a function of the areas of the transmitting and receiving antennas, the separation distance, and the wavelength. For the example of a 1 km diameter antenna and a 10 km diameter receiving antenna the reception efficiency
Fig.2.3.2: TRANSMISSION EFFICIENCY [Source:77/41 Fig. 2.3.3: GROUND POWER OUTPUT LIMIT [Source: 78/42]
of the main lobe will be better than 95 % (2.45 GHz-frequency and distance to geosynchronous orbit). Several antenna illumination functions have been investigated for the best performance in sidelobe and grating lobe suppression. Summarising the results, the 10 dB Gaussian taper has relatively high overall performance considering all constraints. The beam reception efficieny with the above mentioned example is more than 87 %. Fig. 2.3.3 represents the combination of two diagraphs to express the selection of the 5 GW output power level. Using klystrons £or RF-power generation, a maximum power density of 22 kW/m should not be exceeded in the antenna because of thermal problems. The maximum power-density in the ionosphere should not exceed 23 mW/cm^, higher levels might cause nonlinear heating. Fig. 2.3.3 shows, that 5 GW output power meets the restriction. 6 GW would exceed either the antenna power density limit or the ionospheric power density limit. 4 GW transmission is possible only with higher specific cost. Beam attenuation by the atmosphere is expected to be in the range of 2 % (good weather conditions) and 6 % (heavy thunderstorm with wet hail). However 10 - 13 % attenuation basing on another theoretical model were also reported. Depolarisation and deflection effects appear to be of minor influence on the beam. I.2.3.5 Receiving Antenna (Rectenna) The rectenna will be constructed by hundreds of millions of halfwave dipole rectifying elements (Fig. 2.3.4) mounted on reflecting screens (groundplanes). The rectenna elements must be placed perpendicular to the incoming beam. A feature of the reported rectenna design is that it can be continuously fabricated at high speed in different lengths. The area required for a 5 GW rectenna is 117 km^ at 36$ latitude to intercept 88 % of the beam power. Fig.2.3.4: RECEIVING ELEMENT [Source: 78/41 ]
1.2.3.6 Efficiency The reported range of expected DC-DC efficiency for microwave power transmission is 57-68 %. I.2.3.7 Problems The reported main problems with microwave power transmission are: - Reliability and lifetime of microwave amplifiers (up to 30 years) - Beam steering and control, phase-control - Acceptable thermal deflection of transmitting antenna structure I.2.3.8 Laser Power Transmission System ( LPTS ) Recent approaches show, that although the microwave power transmission system MPTS) is much more advanced in understanding than a laser system, the laser transmission has quite a potential for application in a SPS. Advantages are: - Considerably reduced land requirements for receiving sites, - Absence of radio-frequency-interferences (RFI), - Not dependent on 5 GW - size Disadvantages are: - Very low demonstrated receiving and conversion efficiencies, - Extremely high power densities in the beam. - Transmission dependent on weather conditions 1.2.4 Structure and AOCS (Attitude and Orbit Control System) (References: 76/2,78/3,78/23,78/29,78/34,78/35,78/55.) I.2.4.1 General This category implies the support functions of the structure, the investigation of mechanical and thermal loads to the structure, materials for the structure, and considerations of an attitude and orbit control system.
1.2.4.2 Structural Fundamentals The very large dimensions of a SPS will require an open structure for low weight and has to be 3-dimensional if high stiffness is required, (compare to I. 2.6) The most significant dynamic loading frequencies to the lar ge array are the twice daily (2.3 x 10~b Hz) and daily gravity gradient loading cycles. To keep the dynamic motions of the overall station to a reasonable level, a minimum natural frequency criterion of 10 times of the loading frequencies (2.3 x 10 Hz) should be selected. It is estimated that the control system will provide loads at frequencies which will not exceed 0.2 Hz, but there will exist potential interactions between the control system and individual structural elements. 1.2.4.3 Loads Since the mission loads were found to be very low, the transportation - and mounting - loads gain significance for the structural design. The mission loads can be divided into the following categories: A.1 Earth_grayity_gradient_forces These cyclic forces Induce a moment of about 1 million Nm to a 10 GW satellite. This moment requires a maximum control force of 100 N. Earth gravity gradient forces produce also a cyclic tension of 50 N (insignificant). In LEO gravity gradient forces are about 230 times greater compared to GEO. A.2 Solar Radiation Pressure This pressure induces a force of 600 N to a 10 GW satellite which effects mainly an orbit perturbation to a slight eccentricity. An additional shear force of 90 N will appear because of a difference in the area/mass ratios of the solar array and the antenna. A. 3 Solar_and_Lunar_Gravity_and_Earth2.s_Ellipticity These forces cause no significant structural loads but substantial orbit perturbations. A. 4 Atmospheric Drag Atmospheric forces appear only at LEO-construction sites. At an orbit of 500 km an atmospheric force of 400 N has been calculated for a 10 GW satellite. A. 5 Eccentrical Orbit Eccentricity of 0.04 of geosynchronous orbit will produce a torque of about 240 Nm. B.) Induced Loads B.1 Control System Thrusters A thrust level of 100 - 300 N is expected.
B.2 Current Loop/Magnetic_Field_Interaction These~Impacts~can- be made zero by particular design, but could be advantageously used by integrating them into the control system! A maximum of 10 million Nm could be obtained in GEO. In LEO the magnetic field is about 200 times as high as in GEO. B.3 Interaction_Between_Conductors This impact”Is~configuration dependent and can be held to reasonable levels. B.4 Antenna_Recoil This~reactlon force of the microwave beam is 22 N for each 5 GW antenna. I • 2.4.4 Structural Materials Fig.2.4.1: PHYSICAL PROPERTIES OF ALUMINUM AND GRAPHITE COMPOSITE [Source: 78/ 34] As the prime structural material carbon/epoxy composite is considered rather than aluminum because of the benefits due to weight, axial stiffness, thermal and mechanical behaviour. A cost equalization between aluminum and composite is estimated in mid to late 80’s at about £ 20/kg.
1.2.4.5 Attitude and Orbit Control System For the AOCS high, performance electric thrusters (ion-bombardment) were selected with a specific impulse between 7000 s and 13000 s, diameters between 100 cm and 120 cm, and a weight of 120 kg to 150 kg each. As propellant argon was selected (see fig. 2.4.2). The dominant stationkeeping propellant requirement is the correction of the solar pressure perturbation. This requirement, however, could be reduced by higher than 0.5° longitude drifts allowed. Fig.2.4.2: ION PROPELLANT SELECTION CRITERIA [Source: 78/35 ] 1.2.5 Space Transportation (References: 76/2,76/7,77/5,77/8,77/11,77/13,77/24,78/35, 78/36,78/55,78/63.) 1.2.5.1 General This category implies all space-transportation activities. The objectives of a total SPS transportation fleet are:
A. Earth surface to LEO transport of personnel and high priority cargo, B. Earth surface- to LEO transport of large amounts of cargo, C. LEO to GEO transport of personnel and high priority cargo, and D. LEO to GEO transport of large amounts of cargo. 1.2.5.2 The PLV ( Personnel Launch Vehicle ) Fig. 25.1: PERSONNEL LAUNCH VEHICLE [Source 78/55] The DOE/NASA selected a PLV for their SPS-reference system with the following characteristics: A consequent utilisation of the Space Shuttle Orbiter with a reduced-size External Tank and a winged liquid-propellant fly-back booster. The orbiter has a transportation capabiltiy for 75 passengers or an equivalent of high priority cargo. The resized external tank carries 546 tons of propellant instead of 715 tons currently. 1.2.5.3 The HLLV ( Heavy Lift Launch Vehicle ) The selected HLLV concept is a two stage winged vehicle with a LEO-payload capability of 424 tons and a gross lift off weight of 11040 tons. This launcher uses 16 CH4/O2 engines on the booster stage with a total thrust of 156.65 x iOb N and 14 standard Space Shuttle Main Engines (SSME’s) on the or biter with a total thrust of 29.26 x 106 N. The booster has, in addition, 4 turbojets providing flyback capability.
Fig.2.5.2: 2-STAGE WINGED SPS LAUNCH VEHICLE [Source: 78/55] The landing weights are 934 tons for the booster and 453 tons for the orbiter including returned pay load of 63.5 tons. The operating characteristics are vertical take off and horizontal landing (VTOHL) for both stages. 1.2.5.4 The POTV (Personnel Orbit Transfer Vehicle ) Fig.2.5.3: PERSONNEL ORBIT TRANSFER VEHICLE (POTV) [Source: 78/551 The concept which has been selected for the reference system is a two (common) stage propulsion system with three attached modules, the supply module with supplies for 480 man months, the GEO passenger module for 160 passengers, and a flight con-
trol module with a crew of two. This vehicle has payload capabilities of 151 tons up and 55 tons down. The separate weights are the followings I.2.5.5 The COTV ( Cargo Orbit Transfer Vehicle ) Fig.254: CARGO ORBITAL TRANSFER VEHICLE (COTV) OPTIONS [Source: 78/551 Two options have been selected, one for each reference SPS. The one uses silicon cells without concentration as the reference SPS does. It has a payload capability of 4000 tons, weights 1206 tons and requires a round-trip time of 160 days. The total mass in LEO is 6191 tons. The ion bombardement thrusters are identical with those used at the AOCS of the SPS. The other option uses GaAlAs solar cells with a CR of two. It has a payload capability of 3469 tons, weights 715 to, and requires a round-trip time of approx. 180 days. The total mass in LEO is 4396 tons.
I.2.5 •6 Flights per Year for the Reference Systems 1.2.6 Construction and Maintenance (References: 74/1,77/5,77/12,77/13,78/2,78/13,78/15,78/26, 78/27,78/29,78/32,78/33,78/34,78/55,79/1.) 1.2.6.1 General This category implies space manufacturing of structural elements, the mounting of these elements to large structures and the maintenance of the SPS. Fig. 2.6.1 : ON-ORBIT AND EARTH CONSTRUCTION POTENTIALS [Source: 77/5]
The construction philosophy is to fabricate nearly all structural parts in space, to fabricate all components (solar cells, klystrons, etc.) on earth, and to assemble the SPS in GEO. Fig. 2.6.2 : BENEFITS OF SP/CE FABRICATION [Source : 78/291 The SPS structures require high tolerances, especially the antenna ( 1 mm for 10 GHz!). For these tolerances and for good mountability the joint designs will be essential.
I.2.6.2 Construction Elements Five potential large space structure construction elements were identified: - Triangular beams - Geodetic beams - Conical beams - Unfolding beams - Unfolding tetrahedrons From these options the reference SPS-system uses the triangular beams (automatically fabricated continuous chords). The Boeing team compared it with earth fabricated conical tubes: The tube structure weights less than the continuous chord structure, has a higher manufacturing rate, easier structural- integrity verification and more structural design flexibility. The continuous chord structure has the advantage of higher packaging density if manufactured in space, less joint slop, and lower machine complexity. This comparison found no significant difference. Fig.2.6.3: BASELINE BEAM CONFIGURATION For the automatical fabrication of continuous triangular beams in space two optional machines have proved their functionality in 1978, a beam Building machine for aluminum and one for carbon/epoxy beams. A typical application to large space structures is shown in the next figure.
Fig.2.6.4: CONCEPT OF MICROWAVE ANTENNA STRUCTURE [Source:79/1] 1•2.6.3 Construction Site and Construction Bases Fig.2.6.5: LEO CONSTRUCTION SITE VERSUS GEO CONSTRUCTION SITE [Source: 77/13]
The final selection of a construction site (LEO, GEO or combined) depends on many factors. A summary of advantages and disadvantages is shown in fig. 2.6.5. The Boeing team chose a concept with module construction in LEO (8 Modules»1 SPS), electric self propulsion into GEO, and there the final assembly of the power satellite. For this SPS concept this method is considered to offer the greatest potential mainly because of the expected significant savings in transportation cost ( 30 %). This concept requires very large construction bases: LEO: 2.9 x 1.8 x 1 km, 5.6 million kg weight GEO: 1.6 x 1.4 x 0,1 km, 0.9 million kg weight. The Rockwell concept and the DOE/NASA baseline concept chose total construction in GEO mainly because of lower structural loads and therefore lighter structure. A Grumman study says, that the size of a construction base depends mainly on the number of satellites to be mounted per year. 1.2.7 Programme Cost and Risk (References: 76/1,76/2,76/3,76/4,76/6,76/7,76/8,76/9,76/10, 77/3,77/9,77/10,77/11,77/14,77/16,77/17,77/18,77/19,77/24, 77/25,78/2,78/7,78/11,78/13,78/16,78/24,78/34,78/35,78/44, 78/48,78/54,78/57,79/2,79/3.) I.2.7.1 General This category implies critical cost/risk areas evaluation, cost ranging of the overall SPS system and sub-systems (including transportation systems) and cost estimates of full duration programme phases. 1.2.7.2 Critical Cost-Risk Areas Several areas of SPS design, development and implementation concerns have been yet identified, which will remain uncertain due to cost for many years along the DDT&E phase. The most risky ones - i.e. the most sensitive to SPS overall cost - are meant to be the following: - Space transportation cost and its sensitivity concerning construction site location of SPS - Solar cell cost and efficiencies - Rectenna site and construction cost (land use, local environmental and structural impacts, or offshore base construction) - Costs by environmental impacts (acquisition of atmospheric/ ionospheric environmental standards, - of RFI standards on earth and in space, - of surface radiation density standards, etc.)
1.2.7.3 Less Critical Cost Areas Areas of similar .uncertainty of cost estimation but with minor contribution to SPS overall cost, and therefore less critical are: - Fabrication and assembly facilities and support cost (ground and orbital) - Power distribution and antenna cost and efficiencies - Personnel cost (ground and space) I.2.7.4 Main SPS Cost-Risk Sensitivity Parametrics Fig.2.7.1: MAIN SPS COST-RSK SENSITIVITY FARAMETRICS (Sources: 76/2,76/6,78/57]
Besides the SPS sub-systems cost uncertainties the principle risk factor due to cost accrues from total GEO-mass uncertainty of the overall system (satellite and several facilities and supports.) Figure 2.7.1 gives an idea of the GEO-mass ranging of cost correlated to critical cost-risk areas and a total range of DDT&E and commercial busbar electricity cost uncertainty. I.2.7.5 Preliminary Programme Cost Estimates Fig.2.7.2: SPS PRELIMINARY RESEARCH AND DEVELOPMENT COST ESTIMATE [Source: NASA,79/57] Recent cost estimates for a full duration SPS development and commercialisation programme are performed by NASA. The following cost items are analysed and several cost ranges (in 1977 £) are identified: - Technology Development & Verification (Range: $4-8 Billion)
- Design, Development, Test & Evaluation (DDT&E) (Range: £ 36 - 48 Billion) - First Unit (5 GW) (Range: £32-38 Billion) - Total Development (Range: £ 72 - 94 Billion) - Capital Investment Requirements per Unit (Range: £ 12 - 24 Billion) Best estimates for research and development costs out of these items evaluated by NASA are given in figure 2.7.2. Figure 2.7.3 presents a cost break-down of capital investment requirements for the commercialisation phase of SPS, which are very preliminarily analysed by NASA: Fig.2.7.3:CAPITAL INVESTMENT REQUIRED PER SPS (PRELIMINARY ESTIMATE BY NASA)
1.2.8 Energy Demand {References: 76/6,78/9,78/11,78/44,78/45.) I.2.8.1 General This category implies the present and the projected future figures of the total energy requirements and the demand of electricity in the USA and in Western Europe (EC). The potential of a SPS-system is added to some figures. 1.2.8.2 Proven Reserves of World Primary Fuels Fig. 2.8.1: WORLD PRIMARY ENERGY RESERVES [Source: 76/4] 1975 Fig. 2.8.1 contains the total reserves in two categories, reasonable assured and total estimated reserves. The figures for potential fusion fuels have only been listed for reasons of completeness.
I.2.8.3 World Energy Consumption Fig. 2.8.2 represents a projection of total world energy consumption up to the year 2000. The yearly growth is totally 3.5 % and for the. USA and Western Europe 2.9 %. Fig.2B.2: PROJECTION OF FUTURE ENERGY CONSUMPTION [SOURCES: EC, OECD 1977, 78/44,78/45] 1.2.8.4 U.S. Requirements for Electricity Fig.2.8.3: PROJECTIONS OF FUTURE US. ELECTRICITY
In fig. 2.8.3 projections are given for the future U.S. requirement for electricity. It is important to notify, that a more recent forecast (by ERDA) estimates a much lower growth rate (4 %/year). Two SPS-scenarios are implied in this figure. The one provides for 112 SPS's in 2025 starting in the year 1995 (earliest date for operational SPS). The other scenario starts in 2005 with one SPS yearly. I.2.8.5 Requirements for Electricity in the EC Fig.2.84: PROJECTIONS OF FUTURE ELECTRICITY CONSUMPTION IN THE EC.[SOURCES:OECD 1977,EC 1978] In fig. 2.8.4 two projections for electricity demand in the EC up to the year 2000 are presented. It is remarkable, that the growth rates are considerably higher than in the U.S. forecasts. Noticei These projections are just examples of the growing numbers of energy forecasts that could be found in the literature. However, the projections are presented by governmental organisations and therefore selected here. 1.2.9 Cost Comparison 1.2.9.1 General This category only represents the comparison of busbar cost of alternative power generation options because no other cost figures were available from the literature for a reasonable comparison.
Fig.2.9.1: BUSBAR ELECTRICITY COST COMPARISON OF ALTERNATIVE POWER GENERATION SYSTEMS [Sources: 76/2, 76/3, 77/3,77/17, 78/3,78/11] Natural Gas Oi1 Coal Hydropower Oil Shale Nuclear LWR Nuclear LMFBR Fusi on Geothermal power Wind Ground Solar-Thermal Ground Solar-Photov. Oc ean therma1 power Thermionic SPS Brayton-Type SPS Photovoltaic SPS
Fig. 2.9.1 compares actual and predicted busbar cost of 16 power generating options referred to two different time frames, the one currently and the other in the year 2000. The very large range for photovoltaic SPS busbar cost results from the fact, that naturally every publication being considered for this category contains own" predictions. 1.2.10 Funding (References: 76/1 , 77/3,7 8/16,7 8/24,78/25,7 8/37,78/4 5,7 9/1 .) 1.2.10.1 General This category implies the funding figures which have been found in the literature. This includes actual fundings in the past and present and recommended funding figures for proposed research programmes. 1.2.10.2 Actual Fundings Several studies were performed in the U.S. until 1977 dealing with system concepts and subsystems representing a budget of £ 12 million . Fig. 2.10.1: FUNDING OF THE DOE/NASA SPSPROGRAMME DEFINITION PLAN(in $ [Source: 78 / 37] The FY 80 funding of 3.427 million has been recently increased to 8 million by the U.S. Congress.
1.2.10.3 Recommended U.S. Fundings In a hearing befpre a committee of the U.S. Congress in early 1976 a NASA-official requested $ 230 million for a five year research programme. A legislation effort by 11 U.S. Representatives in 1978 calls for $ 25 million in the first year (1978/79) of a SPS research and development programme. This effort was supported by the Congress but the Senate postponed it. After a recent combination of this approach with other future space legislation efforts, the approval of both U.S. houses appears to be certain . In early 1979, Jerry Grey from the AIAA recommended, that the present SPS-programme definition phase should be increased to a 5 year - $ 150 million effort! 1.2.11 Legal Aspects (References: 76/4,76/6,78/24,78/43,7 8/64,78/65,78/66,78/67.) 1.2.11.1 General This category implies aspects of military implications, of political, legal and organisational problems, and of regulation for the erection of rectennas. 1.2.11.2 Military Implications The special literature says, that the SPS-system could be utilised as a power supply for future beam weapons or could be utilised as a psychological weapon itself. The system could support military preparedness by providing energy for a strong and stable economy. A SPS with military capabilities may have a strong negative impact on international relations if it is not internationalised. Vulnerability of the satellite itsplf could be high only by military action of an enemy with high developed space capability. The receiving antenna or the ground control equipment is as vulnerable as many large industrial plants. 1.2.11.3 Polltical,Legal and Organisational Aspects Problems in this sub-category may occur from nearly every interaction of the technical SPS-system with the socio-economic structure of society, with the environment, and with particular political situations. The following critical areas have been identified in the literature: - the use of the GEO for SPS,
- the requirement for international frequency agreements, - international efforts to adjust to an acceptable RFI-level, - the accumulation in space of debris from SPS activities, - the potentially reduced oil or natural gas imports for the SPS-owning countries and other users of SPS’s, and - the organisation for the development and the operation of SPS’ s. 1.2.11.4 Regional and Governmental Regulation for Rectenna Siting Only one publication tries to derive such regulations from existing ones for power generating stations (78/64) . 1.2.12 Public Acceptance (References: 76/1,76/2,76/3,77/17,77/18,78/44,78/69.) 1.2.12.1 General This category implies the potential social impacts from a SPS-programme. Areas of special concern are the safety considerations and the potential implications of the microwave power transmission. Since the public acceptance of large- scale programmes become increasingly important, most authors pointed out, that the defects in public discussion and information during the planning of nuclear power plants must be avoided in planning SPS's. 1.2.12.2 Environmental Impacts by SPS and the Public A. Land Requirements 2 Land primarily required for rectenna siting (^100 km without and1000 kmz with extensive buffer zones) and for launch- and recovery sites. This is one of the most frequently referred public concerns. B. Environmental Pollution by Launch Vehicle Operations This is another subject of serious concern In the public. The large number of required launches will mainly produce high altitude emissions. This could affect the ozone layer, which is vitally necessary for life on earth. Other potential pollutions are low atmospheric emissions by booster engines, ground and water pollution by exhaust fallouts and debris, and very high noise levels near the launch and recovery sites. C. Microwave Impacts on the Ionospheric Environment The potential heating of the ionosphere by microwaves is also seriously discussed as a public concern. These effects might cause a disturbance in communication services utilising the ionosphere. However, the understanding of ionospheric impacts by the beam is very immature.
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