SPACE SOLAR POWER REVIEW Volume 1, Number 3, 1980 PERGAMON PRESS • New York/Oxford/Toronto/Paris/Frankfurt/Sydney
SPACE SOLAR POWER REVIEW Published under the auspices of the SUNSA T Energy Council Editor-in-Chief Dr. John Freeman Space Solar Power Research Program Rice University, P.O. Box 1892 Houston, TX 77001, USA Professor Arthur A. Few Rice University Mr. I.V. Franklin British Aerospace, Dynamics Group Dr. Owen K. Garriott National Aeronautics and Space Administration Professor Norman E. Gary University of California, Davis Dr. Peter E. Glaser Arthur D. Little Inc. Professor Chad Gordon Rice University Dean William E. Gordon Rice University Dr. Arthur Kantrowitz Dartmouth College Colonel Gerald P. Can Bovay Engineers, Inc. Dr. M. Claverie Centre National de La Recherche Scientifique Dr. David Criswell California Space Institute Mr. Leonard David PRC Energy Analysis Company Mr. Hubert P. Davis Eagle Engineering Professor Alex J. Dessler Rice University Mr. Gerald W. Driggers L-5 Society Mr. Arthur M. Dula Attorney; Houston, Texas Mr. Richard L. Kline Grumman Aerospace Corporation Dr. Harold Liemohn Boeing Aerospace Company Dr. James W. Moyer Southern California Edison Company Professor Gerard K. O’Neill Princeton University Dr. Eckehard F. Schmidt AEG— Telefunken Professor George L. Siscoe University of California, Los Angeles Professor Harlan J. Smith University of Texas Mr. Gordon R. Woodcock Boeing Aerospace Company Dr. John Zinn Los Alamos Scientific Laboratories Associate Editors Editorial Assistant: Elizabeth Wood-Gunn Editorial Office: John W. Freeman, Editor-in-Chief, Space Solar Power Research Program, Rice University, P.O. Box 1892, Houston, TX 77001, USA. Publishing, Subscription and Advertising Offices: Pergamon Press, Inc., Fairview Park, Elmsford, New York 10523, USA; and Pergamon Press Ltd., Headington Hill Hall, Oxford 0X3 OBW, England. Published Quarterly. (ISSN 0191-9067) Annual subscription rate (1980) $65.00; Two-year rate (1980/81) $123.50. Special reduced rate for individuals whose institution subscribes $30.00. (Members of the SUNSAT council receive the journal as part of their dues.) Prices include surface postage and insurance; air mail subscriptions extra. Microform Subscriptions: Simultaneous subscriptions on microfiche and microfilm supplied at the end of the volume with index are available from: Pergamon Press and/or its division, Microforms International Marketing Company, Fairview Park, Elmsford, New York 10523 USA; and Headington Hill Hall, Oxford OX3 OBW, England. Copyright © 1980 SUNSAT Energy Council
0191-9067/80/030167-01$02.00/0 Copyright ° 1980 SUNS AT Energy Council EDITORIAL: CALL FOR PAPERS It is the purpose of the Space Solar Power Review to serve as a forum for the exchange of information pertinent to the solar power satellite concept. The Editor welcomes papers of a basic research, engineering and development or general interest character pertaining to space solar power systems. Papers need not be limited to technical and engineering areas but may include subjects such as human factors, environmental and health considerations, legal, international and institutional issues, economics, public attitudes, siting availability, novel concepts, etc. Contributions may be submitted in three forms: regular full length papers (usually not more than twenty-five typewritten pages); short contributions (not more than five typewritten pages); and letters to the Editor. All full length papers will be referred by members of the board of Associate Editors or their associates. The short contributions may be brief technical notes, reports of research too short to form a full length paper, or suggestions for new concepts not yet fully evaluated. Short contributions and letters to the Editor may contain commentary material. These will be reviewed primarily by the Editor-in-Chief. A diversity of views is encouraged. The Review will also publish news and reports of meetings. Contributions for these categories are also solicited. Announcements of forthcoming meetings are also invited. Instructions for authors are found on the inside back cover. The quality of any journal depends on the efforts of its contributors and quality of its contributions. All readers are urged to submit papers and to invite their colleagues to do so. The first three issues represent an outstanding collection of papers. This high standard can only be maintained with the help of all concerned. Thank you. JOHN W. FREEMAN Editor-in-Chief
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0191-9067/80/030169-05$02.00/0 Copyright ° 1980 SUNSAT Energy Council SOME THOUGHTS ON SUNSATt T. O. PAINE President, Northrop Corporation Los Angeles, California You can understand why it is a particular pleasure for me to be with you today. After all, any proposal that combines the large-scale generation of electric power with bold new steps in space capabilities is obviously of extraordinary interest to someone like me with a background both at General Electric and NASA. I just wish I could spend the entire day participating in your discussions of photovoltaics, microwave transmissions, lasers, large antennas, and system tradeoffs and economics. Since I can’t do that, let me pass on some general thoughts to you on this stimulating subject. Looking backward for a minute, I recall the concept of an SPS published by Peter Glaser in Science magazine in 1968. I also remember in 1968 Eberhardt Rees giving me and Wemher von Braun a brief tutorial on the sun. This took place at a top level meeting at NASA to explore the agency’s future directions. There was much discussion of planetary exploration, of high energy astronomical observatories, of the “wet workshop” versus the “dry workshop” (which later became “Skylab”), of the Space Shuttle, of space stations, and of how many trips we should make to the moon — would we need Apollo 18, 19, and 20? When it was his turn to speak, Eberhardt thoughtfully but forcefully admonished us for not giving much higher priority to the sun — “the source of all of our energy,” as I recall his words. We were all impressed by Eberhardt Rees’s vision of the long range significance of the sun. At that same time, in 1968, it was popularly believed that the exploration of space was worthy of a major national commitment. Space exploration was seen as an alternative to war as a focus of national dedication, mobilization and unity. This “new frontier” would stimulate peaceful technological advances to the betterment of world society. More recently, President Carter described our national energy programs as “the moral equivalent of war.” So here we are considering new space systems to help solve the world’s energy problems. Which just goes to show that it’s hard to come up with a really new idea. Or, as the Romans said so appropriately, “There’s nothing new under the sun” (see also Ecclesiastes 1.9). Today serious consideration is being given Solar Power Satellites by professional societies, universities, industry, and the Federal Government. We’re all aware of the many possible ways to use solar energy — the power tower, residential house-top systems, terrestrial photovoltaic arrays and, in space, photovoltaic collectors, thermal engines, microwave and laser transmission, mirrors, low earth and geostationary orbits, etc. From these many possibilities, the Federal Government has selected the tLuncheon address given at the conference, “Solar Power Satellites — What, How, and When,” Los Angeles, February 28, 1980.
SPS as a specific concept to be evaluated thoroughly in a joint study by DOE and NASA. This study began in 1977 and a report is expected this year. For those of you not familiar with this baseline SPS concept, let me outline its main elements so we can explore some of the most important requirements and associated problems. As you will be discussing, the baseline SPS is a huge, box-shaped rectangular structure, with one of its largest sides covered with silicon or gallium arsenide photovoltaic cells. A large circular antenna at one end of the box collects electricity from the photocells and relays it to an earth based antenna via a microwave beam. Let’s recall just how big this space box we are talking about is: its principal surfaces, one of which is covered with the photocells, measures 10 km x 5 km and the box is V2 km deep. It’s hard to visualize a single structure more than 6 miles long and 3 miles wide and 150 stories high, but that’s what we’re considering. The circular antenna is a full km in diameter, and the entire structure weighs about 50,000 tons. Can we design a structure this size to withstand the thermal, solar wind, and other forces in space? If so, can we build it, recalling that we’ve never fabricated a structure in space? Can we control it, keeping the antenna pointed at a fixed spot 23,000 miles away on earth and the photovoltaic cells pointed toward the sun, while keeping the entire structure on its proper station in geo orbit? How long will it take to do this? What will it cost? Will we be able to mass produce Si or GaAs cells economically that will perform at acceptable efficiency, with affordable orbital maintenance, over a lifetime of 30 years? Will the 100,000 klystrons proposed to convert the photocell output from DC to microwave AC function efficiently and reliably? Will that huge 1 km antenna produce a controlled microwave beam focused continually on the receiving antenna on earth without scattering energy throughout the atmosphere? At this stage, an honest answer to those questions might be, “Yes, we probably can do these things, but it’s going to require a well managed organization plus a lot of time and money.” There are a host of other questions. The baseline SPS plan calls for a new heavy lift launch vehicle, which would make even the giant Saturn V of Apollo days look small. It may weigh 24 million pounds at lift-off to put nearly a million pounds of payload in low earth orbit. From low earth orbit most of this payload must be transported on up to geostationary earth orbit, since the SPS is intended to operate 22,000 miles above the equator over South America. Another huge box, about one square mile in area and 180 stories high, is proposed as the cargo truck from low earth orbit to geostationary orbit, with a new electric ion engine as its propulsion system. Additional launch vehicles and spacecraft will be required for personnel, and operations such as cargo-handling and on-orbit maneuvering between earth and LEO, between LEO and GEO, and around the LEO and GEO operations bases. We’ve built a lot of launch vehicles, spacecraft, and space propulsion systems, including small ion engines, but these SPS requirements are orders of magnitude larger and more complex than anything we’ve experienced to date. Can we do all this and make it work? Again, we probably can, but it will take sound management, time and money. These open questions illustrate the uncertainties surrounding only one aspect of SPS — the technical feasibility of lifting the components of an SPS into GEO, constructing the SPS in orbit, orienting it properly, operating it economically, and maintaining all parts of the system in working order. Naturally the ground receiving antenna, the ground control system, and the conversion and distribution of power are also critical to success, but once you find the place to locate the ground facilities,
they are at least readily accessible for testing, maintenance, and repair. In addition to the working parts, there are other areas to be examined, such as the environmental impact of the SPS. The proposed Heavy Lift Launch Vehicle would bum 20 million pounds of propellant enroute to LEO, and the electric ion engines intended for propulsion between LEO and GEO would release thousands of pounds of argon ions. The effects of these effluents on the atmosphere must be determined. Another major environmental issue concerns the impact of the powerful microwave beam on people, animals, and plant life, and on existing communications systems. Will radio astronomy have to move to the far side of the moon? Communication Satellite systems are the information and control lifelines for large segments of society — government, banking, commerce, transportation, industry, and defense. They must be maintained in operating condition. Recall that the electric power in the microwave beam of an SPS would exceed 5000 megawatts, almost 1% of the total electric generating capacity in the U.S. today. That amount of power must be handled very carefully or it could wreak unacceptable havoc on people, vital communications systems, and the environment. All of these uncertainties are in no sense a cause for panic or premature rejection of the feasibility of the SPS Program. They are, however, reasons for thorough analyses and thoughtful evaluation. Fortunately, the system and mechanisms for such analyses and evaluations are in place and at work. The DOE/NASA study is addressing not only every question that I’ve cited, but many others. In addition, the DOE has asked the National Academy of Science to convene a group of independent experts to examine the SPS question and specifically to scrutinize the DOE/NASA study to assure that no important issue has been overlooked. Neither the DOE/NASA study nor the NAS review is expected to reach a decision on whether the U.S. should proceed now to implement an SPS program. The objective is rather to determine whether the SPS concept offers sufficient promise to warrant a substantial five year program of exploratory development, tests, and additional in-depth analyses to enable a national decision to be made about 1986 on whether to proceed with implementation. This work should also identify the most sensitive assumptions and the development work required over a broad spectrum to move forward. Many issues go well beyond the technical questions of space operations and environmental impact, and it may be useful to consider briefly the overall scope and rationale for the SPS program. U.S. energy demand projections cover a wide range, depending on the assumptions used, but it seems clear that in the next 30 years total U.S. energy requirements will increase by at least 50%, even with effective conservation measures. It also seems clear that over the next thirty years our currently installed electric generating capacity of 600 Gigawatts will have to be more than doubled. In comparison to these numbers — 600 GW in 1980 and perhaps 1500 GW in 2010 — the capacity of a single 5 GW SPS is almost insignificant, and the 300 GW which could be provided by the proposed 60 unit SPS program would perhaps comprise only 20% of a 1500 GW requirement for 2010. My message here is that if the dimensions of the SPS program appear huge, so is the requirement it is intended to satisfy. U.S. energy requirements are so enormous that despite conservation there will be increasing demands for coal, for nuclear power, and for whatever hydropower, oil and natural gas can be acquired, even with the proposed 60 unit SPS program. There are other difficult issues, too. Some of them just as complex and even more perplexing. To begin at the beginning—who’s going to pay for, own, control, and use the
electricity from the SPS? If the answer is “we are!”, we’ve got a special set of matters to resolve. For instance, the SPS is intended to operate at 2450 megaherz, so the U.S. has requested the International Telecommunications Union to designate this frequency for use by the SPS. The ITU, however, has declined this request, calling for an extensive study of the entire subject and a future report to the United Nations. ITU acceptance is not a diplomatic necessity, but it is substantively important to the United States. Proceeding without ITU acceptance would require serious consideration by the U.S. Government. Closely related to use of the scarce frequency spectrum is the use of the valuable slots in geostationary orbit. A unilateral U.S. plan to put 60 SPS’s in the 90° equatorial zone over South America, would undoubtedly result in accusations that the U.S. is appropriating for its own use an invaluable natural resource in limited supply which belongs to all mankind. We’ll need a very persuasive U.S. Ambassador to the U.N. in the year 2000. You can imagine many other international disputes — for example, what if the secondary effects of the microwave beam of a U.S. SPS degrades the operation of another country’s satellites to an unacceptable level? What do we do if the other system effectively degrades the SPS microwave beam or the two-way radio communications necessary to control the SPS from earth? On the purely technical level, the volume of space traffic for an SPS program is tremendous—eight Heavy Lift Launch Vehicle flights to orbit per week. Articles associated with the LEO staging base are subject to orbital decay. These circumstances lead to a potential for accidents. If a U.S. HLLV or a crane from LEO drops on Calcutta, Foggy Bottoms will face new diplomatic challenges. Could these problems be solved by organizing at the outset as an international program, with international ownership, management, financing and control? INTELSAT works quite well, but could we get international agreement today for another INTELSAT? Widespread international participation may be called for, but if so it would probably involve initially the two strongest, wealthiest, and most technically advanced partners — the European Common Market and Japan. The technical and economic questions are challenging enough, though, without the added complications of international management — at least in the beginning. In general, about half the cost of delivering power to the bus bar in today’s fossil fuel generating plants are attributable to fuel costs, but environmental protection and added safety requirements are escalating costs rapidly. We can’t estimate the future costs of resolving the technical and environmental issues of nuclear, coal and SPS power, but our experience should allow us to bracket them. Similarly, we cannot confidently estimate the future price of nuclear and fossil fuels, or the environmental costs associated with their greatly increased use. At best, the economic input to a 1981 decision on whether to proceed with exploratory development of SPS can provide only comparative ranges of cost between SPS and its alternatives. So, where do we stand now on SPS and, given the litany of open questions and the absence of definitive answers, how do we decide what to do next? Perhaps the absence of definitive answers may be a fortunate circumstance. Lacking the refuge of specific technical or numerical certainty, we may be forced into broadening our evaluation horizons. Looking back, we can recall the controversy during the 1960s over the utility (or futility) of economical communication satellites for world and national communications systems. Today, no one questions the benefits of satellite communications or the great reduction in the cost of communications made possible by comsats. Now all analogies are slippery, so the comsat revolution doesn’t necessarily presage an SPS
breakthrough, but our rapidly advancing capabilities in economical space operations will make future space activities practical and economical, which are not competitive today. The security needs of our nation will demand continuing expansion of the scale and character of our ability to operate in space. The burgeoning use of space for communications, earth resources, navigation, and research calls for larger and more sophisticated space structures than we presently have. All of these trends, independent of an SPS program, will inexorably lead to new launch vehicles and propulsion systems for inter-orbit transportation, and to the construction of ever larger structures in space, and the establishment and operation of increasing numbers of semipermanent installations in both LEO and GEO. These are facts, not the dreams of a space buff. The privately owned and insured commercial satellites in space today already represent an investment of about one billion dollars. The SPS may be a natural evolution of our general social, economic and technical advance, with our increasing concern for environmental quality. If so, that adds another and more important dimension to the evaluation of the SPS. Let me conclude by admitting what you’ve probably surmised: I don’t know whether the country should proceed with an exploratory SPS development program over the next five years. Those of us on the NAS review committee have another year to conduct our review to better understand both the SPS and the alternative ways to meet the nation’s future needs for electric power. In our inquiries, we will draw on the best talent we can identify, wherever it may be, and we’ll take the broad national view. But don’t misinterpret my long list of issues and problems to be resolved as a sign of pessimism. The nation that carried out the herculean tasks of a global war in 44 months after Pearl Harbor, can accomplish vast and unprecedented tasks—and do them quickly. There were many more uncertainties in meeting Jack Kennedy’s goal of a safe lunar landing in the 1960s than we face today in an SPS program. Intelsat, Apollo-Soyuz, the Space-Shuttle/European Spacelab program and Land- sat all demonstrate our ability to organize effective international partnerships for global space systems. All we need is the vision, the resolve, and the leadership to solve our energy problems. I believe that America is getting tired of the defeatists, the alarmists, and the legalists whose indecision has led to our frustrating dependence on imported petroleum. Let’s lay out a sound technical program, reassert some national leadership, and once again give the world an abundance of safe, clean, reasonably-priced energy.
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APPARENT LUMINOSITY OF SOLAR POWER SATELLITES L. E. LIVINGSTON Lyndon B. Johnson Space Center Houston, Texas 77058 Abstract — The objective of this investigation was a quantitative characterization of solar power satellite luminosity as seen from the earth. The reflective characteristics of the reference photovoltaic satellite configurations are defined. Diffuse reflection from a single satellite will, at maximum, fall between magnitudes -4 and -5. A 60-satellite fleet will have a maximum total luminosity equivalent to a single object of magnitude -9. Specular reflections from the array and antenna will be of magnitude -12 to -15, but visible only occasionally for about 2 minutes at a given location. Methods of preventing specular reflections from striking the earth are presented. Thermal cycle power conversion systems are discussed qualitatively as a means of reducing diffuse luminosity. 1. INTRODUCTION Several large artificial satellites, such as Echo, have been easily visible to the naked eye. Since the proposed Solar Power Satellites (SPS) will be much larger than any satellite launched to date, they can be expected to appear much brighter than previous satellites. The objective of this paper is a quantitative characterization of SPS luminosity. Because optical astronomy will be directly affected, luminosity is expressed herein in terms of conventional astronomical magnitudes as seen from the ground. The two current reference configurations (1) are illustrated in Figure 1. The gallium arsenide configuration has a concentration ratio (CR) of two to reduce solar cell area; the silicon configuration is unconcentrated (CR1). In both cases, the satellite is in geostationary orbit, the solar array is oriented toward the sun with the long axis perpendicular to the orbit plane (POP) at all times, and the transmitting antenna is continuously pointed toward the rectenna on the ground. 2. REFLECTIVE CHARACTERISTICS For optimum thermal properties, the surface of the antenna is presently envisioned as a specular surface with a reflectance as high as 0.98. The radiating slots account for about one percent of the gross antenna area, and gaps between subarrays and other irregularities might be two percent, giving an effective reflectance of perhaps 0.95. Efficient microwave performance requires that the subarrays be parallel
Fig. 1. Reference SPS configurations. within about three minutes of arc (1). Since this is much less than the apparent diameter of the sun (about 32 arc minutes), there will be no appreciable additional spreading of the reflected beam. The antenna could, therefore, be regarded as an excellent mirror. The solar cells contribute most of the reflectance of the solar array. For the CR1 configuration, cell area is about 50.1 km2. Cell reflectance depends on the type of cell used. The silicon cell now planned (2) would incorporate a reflective underlayer for improved efficiency and heat rejection, and is expected to have a total reflectance of about 0.25 in the spectral range from 0.35 to 2.0 /xm. The relative magnitudes of the specular and diffuse components are not well established; in this analysis, they are assumed to be equal, i.e., 0.125 each. Other types of cells may have lower reflectance, e.g., 0.04 diffuse and 0.035 specular. Because the difference could be significant, both values are used. The Kapton substrate on which the individual solar cells are mounted has about 1.8 km2 of exposed area between the cells. With a diffuse reflectance (2) of 0.67, this can make an appreciable contribution to the total luminosity of the array. The exposed structure on the front surface of the array has an effective area of about 0.24 km2 (estimate by the author). A reflectance of 0.85 has been assumed. Both diffuse and specular reflections must be considered. Because the analyses are entirely different, they will be examined separately. 3. DIFFUSE REFLECTION If the reflected energy were distributed uniformly over the hemisphere of radius R visible from the array surface, the apparent luminosity would be that of the sun
multiplied by the factor pAHirR2, where p is the diffuse reflectance, A is the area of the reflector and R is the distance from the satellite to the observer (3). Using subscripts c, k, s and a for the solar cells, Kapton substrate, structure and antenna respectively, we have Because of the significant contribution of the Kapton substrate, the difference is less than a magnitude. If the cells were the only reflective component, the change of reflectance from 0.125 to 0.04 would result in a change of 1.26 magnitudes. Variation in other parameters has less effect. A reduction in substrate reflectance from 0.67 to 0.2 increases wmax (algebraically) by 0.1 to 0.3 magnitude, depending on the solar cell reflectance used. Eliminating structural reflectance completely gains at most 0.1 magnitude. If, on the other hand, the antenna were diffuse rather than specular, the magnitude would be decreased by 0.1 to 0.3. Thus, the maximum plausible luminosity of the SPS, including a diffuse antenna surface, is It should, of course, be kept in mind that the areas and reflectances quoted here are estimates and subject to change as the SPS evolves. However, a major variation in parameters is necessary to effect a significant change in luminosity. All luminosities quoted above assume the sun at the equinox and the observer at the equator. At the solstices, the illumination on the array is reduced by 8%, resulting in an algebraic increase of 0.09 in the apparent magnitude. For an observer not on the equator, the increased angle between the line of sight and the normal to the surface increases the apparent magnitude by a maximum of 0.013. The solar array is always oriented toward the sun. Thus, the illuminated face of the CRl configuration is visible from the earth from about 6 p.m. to 6 a.m., local apparent time at the longitude of the rectenna, with maximum luminosity at midnight. At other longitudes, the times will vary as much as half an hour due to the At 40° latitude, an observer is about 37,500 km from a satellite at the same longitude. Using this value for/? and taking the maximum diffuse reflectances given above, we obtain
difference in viewpoints (see Figure 2). The concentrators on the CR2 configuration partially block the arrays from view, so that the period of visibility is from about 8 p.m. to 4 a.m., with luminosity equivalent to CR\ only from 10 p.m. to 2 a.m. Assuming a cosine law luminosity distribution, then, These relationships are plotted in Figure 3. However, there will be not one SPS but many. For observers in the western hemisphere, all the satellites in the fleet will be visible simultaneously, although at different luminosities. If a satellite is t degrees east of the observer, the angle 0 between his line of sight and the normal to the surface is defined by (see Figure 4) Fig. 2. Variation of extinction time with longitude from rectenna.
Fig. 3. Luminosity vs. time (single satellite). Fig. 4. Angle of incidence of observer’s line of sight. if the previously noted small adjustments for solar declination and observer’s latitude are neglected. Since the observer is not, in general, at the same longitude as the
Fig. 5. Equivalent luminosity of 60-satellite fleet. satellite, the quantity 15t in equations (4a-c) must be replaced by 15t + r + 0. The total luminosity of the SPS fleet can now be estimated by summing the contributions of the individual satellites. The results are presented in Figure 5 as the magnitude of a single object of equivalent total luminosity. A 60-satellite fleet at longitude intervals of one degree is assumed. Magnitude is shown as a function of the observer’s local apparent time for CR\ and CR2 configurations with the observer at the same longitude as (1) the center of the fleet and (2) the east end of the fleet. Even in the best case from an astronomer’s viewpoint, CR2 at the end of the fleet, it would be difficult to make long observations without an appreciable contribution to sky brightness by the SPS fleet. As noted above, even if very low reflectances are assumed, the situation is improved by only about one magnitude. 4. SPECULAR REFLECTION FROM SOLAR ARRAY Specular reflections can be expected from both the solar array and the antenna. Since the two elements have different orientations, they must be analyzed separately. Any specular reflection striking the earth from the solar array will be within a few
Fig. 6. Specular reflection from solar array (ideal orientation). degrees of the normal to the surface of the array. Area reduction due to oblique angles of incidence may therefore be neglected. The apparent luminosity can be calculated by noting that if the array were large enough to reflect the entire solar disc, the luminosity would be that of the sun multiplied by the specular reflectance, p, of the solar cells. This would require an array area of n(Ra/2)2, where R is the distance from the observer to the satellite and a is the apparent diameter of the sun, 9.36 mrad. If Ac is the area of the array, then the luminosity is Using the same values as in the previous section and taking p as 0.125, m = -16.4, or about 30 times the luminosity of the full moon. Even with p = 0.04, we have m = -15.2. For the CR2 configuration, the same reflectances give a magnitude range of -15.6 to -14.4. It can be seen from Figure 6 that, because of the orientation of the array relative to
Fig. 7. Specular reflection from solar array with attitude error. the sun and the orbit plane, a reflected ray could strike the earth only when the earth is between the SPS and the sun. Hence, specular reflection from the array will not be a problem in the nominal orientation. However, as Ekstrom (4) has pointed out, the array will not in general be precisely in the nominal orientation, but somewhere within a small attitude deadband. This deadband has not yet been defined in detail, but can be assumed to be less than a degree. If the error is such as to tilt the normal to the array toward the sun when the sun is near the earth’s limb as seen from the SPS, then a reflection could strike the earth. This can occur only near midnight during six-week periods centered on the equinoxes. If the reflected beam is seen only near the horizon, atmospheric absorption will mitigate any adverse effects. The angle of incidence at the ground, 0, may be readily calculated from the attitude error, -y (see Figure 7): This function is plotted in Figure 8, from which it may be seen that a deadband of 0.1° will allow no reflection to be visible more than about 12° above the horizon. A significant decrease in this maximum elevation requires a major reduction in deadband which may not be practical. For y = 0.1°, the reflection will be visible no more than 26° from the terminator if atmospheric refraction is neglected. The maximum duration at a given location is about two minutes, and the times at which a reflection could occur can be calculated in advance. Consequently, the impact of such reflection on astronomical observations is not likely to be of major importance. If, because of possible retinal damage to observers (4) or for other reasons, all specular reflection from the array must be avoided, a non-POP attitude can be flown during the periods of possible reflection. If the commanded angle between the orbit plane and the normal to the array is and the maximum error is y, then the following
Fig. 8. Angle of incidence on earth of specular reflection from array. attitude profile will prevent any specular array reflection from striking the earth (see TTimir-o QV Here 8.7° is the semidiameter of the earth seen from the SPS and 8 is the declination of the sun. This procedure requires an attitude change of roughly nine degrees in about 22.5 hours at the equinox. Assuming a constant acceleration and deceleration, the peak velocity must be 3.64 x 10~6 rad/s and the acceleration 95.8 x 10~12 rad/s2. Estimating the moment of inertia from reference 1 at 6.1 x IO14 kg - m2, the necessary torque is about 73000 N-m. However, the torque due to solar radiation pressure on the asymmetrical reference configuration is about 960,000 Nm. Since this torque is countered continuously by the reaction control system, the equinox maneuver places no additional load on the system.
Fig. 9. Attitude profile of array to avoid specular reflection. The illumination loss is (1-cos i/i), or a maximum of about 0.003. This is insignificant compared to the 8% loss at the solstices resulting from the POP array orientation. 5. SPECULAR REFLECTION FROM ANTENNA The luminosity of the antenna due to specular reflection can be calculated in the same way as that of the solar array. In this case, if the reflective surface is a circle of diameter d - 1 km, the luminosity is As with the array, reflections can strike the earth only from nearly normal incidence on the antenna; area loss can therefore be neglected. Taking p = 0.95, we then have m = -14.1 if the other parameters are the same as in the preceding section. For efficient power transmission, the transmitting antenna must point toward the rectenna within about one minute of arc (1). Consequently, the antenna cannot be pointed so as to avoid reflections to the earth. Since the antenna will not, in general, point at the equator, the reflection can strike the earth at any point from which the SPS is visible. The relationship between the sun and the antenna is calculated as illustrated in Figure 10. If p is the longitude difference between satellite and rectenna and is the rectenna latitude,
Fig. 10. Geometry of specular reflection from antenna. Then, if t is the time in hours from local apparent midnight at the rectenna, the angle of incidence of sunlight on the antenna, 0, is given by the reflected beam does not strike the earth. The distance R from the satellite to the center of the spot is
Fig. 11. Geometry of specular reflection from antenna (continued). Fig. 12. Variation of latitude of antenna reflection with date and rectenna latitude. The latitude and longitude of the center of the spot relative to the rectenna are then given by The most significant characteristic of the motion of the spot is the variation in latitude with solar declination. This is plotted in Figure 12 for satellite and rectenna at the same longitude. The spot from each antenna passes a given latitude in the spring and in the summer. However, because the SPS is eclipsed by the earth until about April 12th in the spring and from about August 31st in the summer, the
Fig. 13. Number of sightings of antenna reflection. reflection can t>e seen only Between tnese two aates unless tnere is an appreciate longitude difference between the rectenna and satellite. If all rectennas are between 25° and 40° north latitude, a fixed observer will see all the reflections within two- week periods in the spring and summer; however, an observer north of about 24° north latitude would see no reflections at all under these circumstances. The apparent angular velocity of the sun is 2rr cos 8/86164 rad/s. Since the reflection must have the same angular velocity, the maximum contact time for a fixed observer is if the center of the spot passes his location. For an observer in the southern hemisphere who sees the reflection at the summer solstice, T could be as much as 140 seconds. The apparent diameter of the antenna from the ground is d/R rad, or from 0.024 to 0.028 mrad depending on elevation above the horizon. Since diffraction effects are negligible for a one-km aperture, the time required for the reflection to reach full luminosity will be 86164/2tt R cos 8, or between 0.33 and 0.42 seconds. By plotting the latitude extremes of the edge of the reflected spot, an approximate indication can be obtained of the number of sightings to be expected at a given location. For a typical rectenna latitude of 35°, Figure 13 shows that the reflection will be visible on about two successive nights at middle latitudes and about three nights near the equator. Comparison with Fig. 12 shows that the number of sightings varies only slightly with rectenna latitude except for rectennas above 50°, in which case observers around 60° south latitude would see the reflection for several days in succession. However, very few rectennas are likely to be located at such high latitudes. The spot moves across the earth from east to west. Figure 14 illustrates a few ground tracks for a rectenna latitude of 35°. The coverage of the spot is indicated at five-minute intervals. The track for May 1st assumes a satellite 15° east of the rectenna to illustrate the effect of a longitude differential.
Fig. 14. Typical ground tracks of antenna reflection. Although the specular reflection from the antenna is not as luminous as that from the solar array, it could prove to be nearly as objectionable. Since antenna pointing constraints do not permit redirection of the beam, the alternative is to use a diffuse surface, even at the cost of some thermal performance. As has been noted above, a diffuse antenna surface adds little to the diffuse luminosity. 6. SPECULAR REFLECTION FROM CONCENTRATORS With the CR2 configuration, reflections from the concentrators will be also visible. Since uniform illumination of the solar array is desired for optimum performance, the concentrators must extend past the ends of the cell blanket to provide uniform concentration at all solar declinations with POP orientation (Figure 15). Near the equinoxes the reflection from these extensions will strike the earth at about 8 a.m. and 4 p.m. satellite time. However, some observers could see the reflection as much as 3-1/2 hours before sunrise or after sunset. The total effective area reflecting at one time reaches a maximum of 0.46 km2, or 60% of the antenna area, because the angle of incidence is about 60° when the reflection strikes the earth. The maximum luminosity would, therefore, be somewhat less than that of the specular antenna. Because the reflections are visible in daylight for the most part, they are not expected to be as troublesome. If they are found to be unacceptable, on the other hand, a simple and relatively inexpensive fix would be to stretch a sheet of concentrator material across the gap at the end of the solar cell blanket (Figure 15).
Fig. 15. Specular reflection from solar concentrators. Fig. 16. Typical thermal cycle SPS. 7. THERMAL CONVERSION SYSTEMS Thermal cycle solar energy conversion systems, primarily Brayton and Rankine cycles, have been extensively studied as alternatives to the reference photovoltaic systems (5,6). The best thermal systems have been found to be somewhat heavier and more costly than photovoltaics and have consequently not been adopted as the reference SPS. For this reason, the visibility properties of thermal systems have not been examined in detail. However, a brief consideration of their basic characteristics can permit some general conclusions. In a typical configuration (Figure 16), the solar collector consists of a large con-
cave reflector, continuous or faceted, focussing the sun’s rays at the mouth of a cavity absorber about 70m in diameter. The turbines and generators are mounted outside the absorber, together with the radiator, which is arranged edge-on to the sun. The most important feature of this type of system from a visibility standpoint is that the bulk of the area (the reflector) has an extremely low diffuse reflectance and that the specular reflection, being directed to the absorber, cannot strike the earth. The next largest component, the radiator, is located so as to present minimum area to the sun, coincidentally minimizing reflected sunlight. The absorber and turbomachinery are relatively small in overall dimensions and will probably be sufficiently irregular in shape to reduce substantially any reflection. It has already been seen that the truss-type structure has very low total luminosity because of its small effective area. It may be concluded, then, that a thermal cycle SPS can be expected to have considerably less luminosity than the photovoltaic reference systems. This does not consider the transmitting antenna, which would be the same in either case. 8. CONCLUDING REMARKS Diffuse reflection of sunlight will give each SPS the visibility of a bright planet for most of the night. With the reference configurations, this appears to be unavoidable. With the exception of optical astronomy, the diffuse reflection does not seem to present any major problems. Specular reflection can also occur from two sources (three for the CR2 reference configuration) for brief periods. These may have a potential for retinal damage to telescopic observers. However, all of the specular reflections can be prevented from striking the earth by one means or another if further research should show it to be necessary. In short, although the visibility of the SPS may be an inconvenience, it will not be a hazard. Even the inconvenience could, if necessary, be mitigated by use of a thermal conversion system instead of photovoltaic. Acknowledgement — The editor wishes to thank Dr. Harold Liemohn and Professor Harlan J. Smith for their assistance in reviewing this paper. REFERENCES 1. Satellite Power System Reference System Report, U.S. Department of Energy and National Aeronautics and Space Administration, Washington, D.C., DOE/ER-0023, pp. 12-37, October 1978. 2. J.L. Cioni, Johnson Space Center, private communication, February 9, 1979. 3. H.H. Koelle, Ed., Handbook of Astronautical Engineering, McGraw-Hill, New York, pp. 8-11 and 8-12, 1961. 4. P.A. Ekstrom, Battelle Pacific Northwest Laboratories, private communication, April 26, 1979. 5. G.R. Woodcock, Solar Power Satellite System Definition Study — Final Report, Boeing Aerospace Co., Seattle, Washington, D180-22876, December 1977, contract NAS9-15196. 6. G. Hanley, Satellite Power Systems Concept Definition Study — Final Report, Rockwell International, SD78-AP-0023, April 1978, contract NAS8-32475.
LOGISTICS COSTS OF SOLAR POWER SATELLITES* R. H. MILLER and D. L. AKIN Space Systems Laboratory Department of Aeronautics and Astronautics Massachusetts Institute of Technology Cambridge, Massachusetts, USA 02139 Abstract — A brief discussion details the current level of understanding of the problems associated with the proposed deployment of solar power satellites (SPS). A costing rationale is presented, based on standard systems analysis techniques. The problem is broken down into major cost-driving parameters of transportation, productivity, and solar collection device costs, and further broken down into sets of parameters which specify a construction scenario for the SPS. An analysis of transportation to low earth orbit shows that earth launch costs may be reduced to below $40 per kg through the full utilization of launch system capabilities. A computer program is developed which performs an estimation of the overall program costs of the SPS, based on a set of 55 parameters which impact systems costs. Baseline cases are selected, and sensitivity studies performed to find the cost impact of the major variables. It is found that the use of nonterrestrial materials is economically attractive from a very early point in the program, and that a program of SPS construction can compete favorably (<$1000/kW installed) with systems in the 1990 time frame and beyond. 1. INTRODUCTION During the ten years since the concept of satellite solar power systems was first introduced (1), exploratory research has indicated that the state of the art in space transportation, zero-gee operations, and microwave transmission is sufficiently advanced to permit initiating development of the concept during the next decade. Granted that it is feasible to generate power in space and transmit it to earth for collection and distribution, questions still remain as to whether this potential solution to the world’s growing energy crisis is either economical or desirable. This paper will address the first question as a necessary prerequisite to the second, which can ultimately be decided only by the interested nations. Clearly the desirability of solar power satellites stem from their use of an in- exhaustable, nonpolluting source of energy which, if tapped in space, is capable of supplying base load continuously. This energy is available without the need for storage devices, and can be readily modulated as demand dictates. Concerns stem from the unknown environmental effects of the low density microwaves required for power transmission, both on the upper atmosphere and at the receiving antenna (rectenna). Additional research is needed to answer these concerns fully, nor should the SPS be developed until such research has been conducted. Present preliminary ♦Presented at The 29th Congress of The International Astronautical Federation in Dubrovnik.
findings indicate that the microwave transmission system will not interfere with the other users of the electromagnetic spectrum, and that there will be no serious biological hazards associated with the rectenna side lobes. The transmission problem may therefore reduce essentially to one of the economics of land usage. If it is accepted that the SPS is a desirable candidate energy system for relieving the world’s present almost total reliance on fossil fuels, then the question must be answered as to its competitive economic position relative to other candidate systems in the same time frame, such as nuclear. In considering this question, it is important to emphasize that the world’s energy needs in the next two decades will require a capital investment for new plants measured in the trillions of dollars. Consequently, the development and manufacturing costs for the SPS can only be rationally evaluated in the context of this major capital outlay, which is required if the world is to have any hope of maintaining its present standard of living in the face of a burgeoning population, or of satisfying the legitimate aspirations of the developing nations for an equal quality of life. Relative costs are therefore of more interest than absolute values: it is convenient to measure the economics of the SPS relative to the known costs of fission power plants, approximately $1000/kW for capital investment. However since the SPS does not use fuel the equivalent number for purposes of comparison is probably closer to $1500/kW to $2000/kW. Fusion remains a hope for the future, but too many scientific breakthroughs are needed to permit a rational evaluation of its economics at the present time. In evaluating the economics of the SPS many cost elements must be estimated, mostly on the basis of very little hard data. The approach which will be taken in this analysis will therefore consist of identifying the most sensitive cost parameters by considering the effect of their variations separately for a base case. It should then be possible to evaluate the economics of the SPS for any selected set of initial assumptions. 2. MAJOR COSTING PARAMETERS The driving costs for an SPS which will be addressed in this paper are: (a) transportation costs (b) industrial productivity in space (c) energy conversion element costs. Recent studies have clearly indicated that the costing algorithm is highly nonlinear, in that there is a strong interdependence between variables. Certainly, however, (a) has dominated the cost of space operations in the past. It is only with the advent of the reusable launch vehicle that it has become at all possible to consider the assembly of large systems in space. The dominating cost of space transportation has also led to two novel concepts which show great promise for reducing SPS costs: the use of lunar rather than terrestrial material for manufacturing (2), and the use of an electromagnetic reaction engine in place of chemical propulsion (3). A major purpose of this paper will be to project the potential costs of transportation to and in space, and to determine the conditions under which it appears that the use of more advanced concepts are desirable. The optimum scenario depends on the costs of transportation, including the costs of propellants at the various transportation origins and destinations. The moon is an
obvious source of certain basic materials, because of the much lower energy requirements for transport off the lunar surface. However, against this advantage must be weighed the logistics costs of maintaining an operational base on or near the moon and bringing up from earth those elements (hydrogen, for example) which are required for operations and are not available on the moon. Manufacturing costs may be conveniently broken down into recurring and nonrecurring costs. Nonrecurring costs are the initial costs associated with, for example, setting up lunar and other space bases, acquiring the necessary equipment, and prototype development. Recurring costs are those associated with the actual manufacture of the sub elements and their assembly. These latter costs are subject to an exponential learning curve in which the cost of production of a single unit decreases as the number of units built increases. Unit costs may therefore be expressed in the form where Cn is the cost of the nth unit in a production run of N. For a learning curve corresponding to a 20% reduction in cost for every doubling of the production run (an 80% learning curve), P has the value P = .32. It is therefore clear from (1) that the cost of producing an SPS will depend critically on N, and it may furthermore be expected that the optimum manufacturing scenario will also depend on N, which represents the total thruput of material through the system. Another primary factor determining industrial costs is the productivity of labor in space. This factor is important not only because of the wage scale, even when allowing for the much higher anticipated scale in space, but also because of the costs of life support and the need for crew rotation which, in turn, are critically dependent on transportation costs. These costs would be much reduced if it were possible to postulate the availability of space habitats with closed ecological systems: indeed, the future development of space power systems may well be the catalyst leading to the realization of the concept of space colonization. Recent experiments in a simulated weightless environment (4) have provided some information on the potential productivity of man in space; however, this dominant cost parameter has been kept a variable in the analysis because of the high degree of uncertainty involved in its initial definition, and in the rapidity of the learning process in space. Turning to (c), it is evident that the costs of the elements required for converting solar power to electricity is closely tied to (b), the manufacturing costs. For the purposes of this paper, only photoelectric solar cells will be considered in order to limit the scope, although the conclusions should be equally applicable to other systems such as closed-cycle thermal engines. The costs of producing solar cells on earth have been following a well developed learning curve, whose projection through the next two decades indicates a potential cost reduction by two orders of magnitude, even in the absence of the manufacturing
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