Space Power Resources, Manufacturing and Development Volume 12 Numbers 1 & 2 1993
SPACE POWER Published under the auspices of the Council for Economic and Social Studies on behalf of the SUNSAT Energy Council. Editors: Dr. Gay E. Canough, ETM Solar Works and Dr. Andrew Hall Cutler, NASA Space Engineering Center, The University of Arizona Associate Editors: Fred KoomanofT, Dept, of Energy, USA Richard Boudreault, Consultant, Montreal, Canada Lars Broman, SERC, Sweden William C. Brown, Massachusetts, USA Lucien Deschamps, Paris, France Ben Finney, U of Hawaii, USA Peter Glaser, Aurther D. Little, Inc. USA Dieter Kassing, ESTEC, The Netherlands Mikhail Ya. Marov, U of North Carolina, USA Gregg Maryniak, International Space Power Program, USA Makoto Nagatomo, ISAS, Japan Mark Nelson, Institute of Ecotechnics, USA John R. Page, U of New South Wales, Australia Tanya Sienko, NASDA, Tsukuba, Japan Space Power is a quarterly, international journal for the presentation, discussion and analysis of advanced concepts, initial treatments and ground-breaking basic research on the technical, economic and societal aspects of: large-scale space-based solar power, space resources utilization, space manufacturing, space colonization, and other areas related to the development and use of space for the benefit of humanity. Recent subject coverage: • history' and status of national space power programs • technologies for large-scale space power e.g. solar power satellites • systems aspects of large-scale power, e.g. SPS and central space power utilities • potential extraterrestrial resources for use in space-based manufacturing • lunar and planetary science for understanding space resource location and availability • plasma and other space environment interactions with large space structures • medical, psychological, sociological and cultural aspects of human presence in space • forms of advanced space propulsion and power technologies and systems. Space Power is published four times per year. These four issues constitute one volume. An annual index and titlepage is bound in the December issue. 1993 is volume 12 ISSN = 0883-6272 Editorial Correspondence: Dr. Gay E. Canough, Space Power c/o ETM Solar Works, Inc., PO Box 67, Endicott, NY 13761, phone/fax = (607) 785-6499 e-mail (Internet): CANOUGH@BINGVAXA.CC.BINGHAMTON.EDU radio call sign: KB2OXA. Business Correspondence: including orders, subscriptions, advertisements and back issues should be addressed to the publisher: Council for Social and Economic Studies Inc.. 1133 13th St. N.W., Suite C-2, Washington D.C. 20005, U.S.A. Tel: (202) 371-2700; FAX (202) 371-1525. Subscriptions: libraries S288/year, individuals $144/yr: additional $25/yr for foreign airmail. Cover: Artist’s conception of a Solar Power Satellite in space with the inset showing how the rectennae or Earthbased receiving station would appear. The painting was done by Marc Martel of Dayton, Ohio, with scientific input from Dr. Seth Potter of New York City. The artwork is the property of Space Studies Institute of Princeton, New Jersey. Copyright Space Studies Institute. 1993.
Note from the Editors: Welcome back to Space Power! We have had some difficulties with the publication of the journal in the past year, however, things are now back on track and you can expect to receive the first 1994 double issue by the summer. Aerospace budgets are being cut everywhere and so work in the various space fields has grown less. However, there are some special bright spots for space power. It is becoming widely understood that the need for energy on Earth must be met with cleaner forms of energy, such as solar. Therefore, work in this area is starting to accelerate. Large renewable energy programs are beginning in industry and in some governments. Some of these programs include space solar power and wireless power transmission. We will also be including some news in issues of Space Power to keep you up to date on new developments. Dr. Gay Canough of ETM Solar Works will be taking over the editorship of the journal, so please contact her about submitting articles. Dr. Andrew Cutler has been the editor for the past 5 years and seen the journal through many great issues. Sincerely, Andrew Cutler Gay Canough
SPACE POWER Volume 12, Number 1 & 2, 1993 Table of Contents H. Matsumoto, N. Kaya, S. Kinai, T. Fujiwara, J. Kochiyama. A 1 Feasibility Study of Power Supplying Satellite (PSS) Keith Rogers. Design of a Low-Cost, Earth to Space Power Beaming 7 Demonstration Makoto Nagatomo, Kiyohiko Itoh. An Evolutionary Satellite Power 23 System for International Demonstration in Developing Nations Kiyohiko Itoh, Yasutaka Owaga. An Inland Rectenna Using Reflector 37 and Circular Microstrip Antennas Geoffrey A. Landis. Receivers for Laser Power Beaming: Summary of 51 the Workshop at SPRAT-XII Klaus Lotzerich. A Comparison of a Conventional Launch System vs. 55 Externally Supplied Vehicles for Installation and Maintenance of Solar Power Satellites Kitt C. Carlton-Wippern. An Introductory Analysis of Satellite 67 Collision Probabilities Benjamin Dessus' Francois Pharabod. Energy Development and 83 Environment: What about Solar Energy in a Long Term Perspective? Elgudja Medzmariashvili, Alexander lacobashvili, Guram Beducadze. 99 Creating and Testing Large Space Structures of High Precision Surface
A Feasibility Study of Power Supplying Satellite (PSS) H. MATSUMOTO* N. KAYA, S. KINAI.**, T. FUJIWARA,***, J. KOCHIYAMA, **** A feasibility study is given on a new type of an orbiting power station which supplies a power of the order of 100 kW to orbiting customers (satellites or space stations). We call the orbiting power station "PSS" which stands for Power Supplying Satellite. The PSS is composed of three main parts; a power generator, a power transmitter and a satellite bus system. The unique feature of the proposed PSS is the use of a module which has a solarcell array on one side and a microwave transmitting antenna array driven by FET- amplifiers on the other side. These autonomous transmitter modules are used to form a large disc-structured active phased array of 40 m diameter which transmits a 100 kW energy beam of 24 GHz microwave. The frequency of 24 GHz is chosen to reduce the size and volume of the transmitting antenna. DC electric power generated by the solar cells on the top plane of the module is fed directly to the semiconductor FET amplifiers located under the solarcells, converted to 24 GHz microwave, and then transmitted from the antenna arrays on the bottom plane of the module. These autonomous solar-cell-transmitter modules make it possible to eliminate the rotary joint and the DC power collecting network from the design of the PSS and SPS system. The PSS provides the following advantages; (1) availability of high power electricity to orbiting customer satellites or stations; (2) reduction of weight and volume of power system on customer space vehicles; and (3) technical go-forward for the future SPS. Introduction Man's prosperity has been possible by expansion of four quantities; the area of habitation, materials and food, population, and energy. It is obvious that all of the four quantities will eventually (but not in the far future) reach their limit of availability and thereby choke the future of mankind. From this point of view, space civilization is inevitable in the 21st century to guarantee ever-increasing activities of mankind. The Solar Power Station (SPS) proposed by P. Glaser in 1968 was originally supposed to become a rescue for the lack of electric power demand on our mother planet * Radio Atmospheric Science Center, Kyoto University, Kyoto, Japan **Dept, of Instrumentation, Kobe University, Kobe, Japan *** Space Development Section, Nissan Motor Co., Tokyo, Japan **** Rocket System Corp., Tokyo, Japan
without destroying the Earth's environment!!]. This idea is still valuable and should be realized before the world economy collapses due to the crisis of oil which will become shortly unavailable by the middle of the next century. However the SPS is only a milestone of the space civilization program. Not only the demand of power on the Earth but that in space will have to be satisfied by Space Power Facilities. The SPS, which is planned to be placed on the Geostationary Orbit, is not always suitable for power transmission to fast-orbiting space vehicles. A different design for such need must be studied other than those having been made for the SPS. The present paper presents a feasibility study of a Power Supplying Satellite (PSS) and proposes some new concept of design of the PSS. Concept of the PSS The PSS is an orbiting power satellite with a capability of feeding a maximum power of 100 kW to orbiting customer satellites or stations. A schematic illustration of the PSS is given in Figure 1. It uses a 24 GHz microwave as an energy carrier in contrast to 2.45 GHz proposed for the SPS. The reason for this is that the higher frequency reduces the size of the transmitting antennas and hence the volume and weight of the PSS. Regardless of this merit of using higher frequency of 24 GHz, the SPS is supposed to use 2.45 GHz microwave to avoid a strong absorption and damping by the Earth's atmosphere. However, in case of the PSS, the power transmission is limited only in space and thus is free from the problem of power absorption by the water molecules. The advantages of the establishment of the PSS as one of the space infrastructures are as follows. (1)A High Power Availability in Space Power in space is normally generated by deployed paddles of solar-cell array. However, a high power supply by the huge solar battery system requires a more complicated control system and enhances the weight and volume of the necessary solar paddles. Therefore, the attitude control of the heavy solar paddles produces the perturbation of an artificial gravity in the spacecraft, which destroys the ideal zerogravity environment for some material science and processing in space. Such bulky and complicated onboard power systems of the spacecraft could well be avoided by replacing the solar cell paddles with a much lighter and simpler microwave rectenna system. The PSS is capable of feeding power up to 100 kW to such rectenna-equipped customer satellites or stations. (2) Establishment of Key Technology of Microwave Energy Transmission Though the frequency of the microwave used for the SPS is an order of magnitude lower than the PSS microwave, most of the technologies developed for the PSS are transferable to the future SPS. One of the new technological aspects of the PSS
is the development of a unified module of power generator and power transmitter. The module, which we may call "autonomous transmitter module", consists of three layers; high efficiency solar cells on one side; transmitting antenna arrays on the other side; and F-class FET power amplifiers in between the solar cell and the antenna planes. The direct connection of the DC output of the solar cells to the FET amplifier of each active antenna element makes it possible to design a much simpler structure not only of the PSS but also of the future SPS. In the SPS Reference System studied by NASA/DOE in 1980 [2], DC electric power generated by the solar cells must be collected by a DC power collecting network in the huge SPS solar paddle. A critical technology of using super-conductor network for the current collection is being considered for the SPS. This DC power collection network is not necessary for the proposed PSS. If this technology is once established then the design of the future SPS would become much simpler and less expensive. Another critical technological problem raised in the SPS Reference System is a mechanical rotary joint which electrically connects the differentially rotating Solar Paddle and the microwave transmitting antenna in vacuum. This is one of the difficult technological point of the SPS. However, the present PSS does not need such a rotary joint because the DC electric power is directly fed to the FET amplifier situating below the Solar Battery Unit. Elements of PSS The PSS is composed of three main sub-system; a power generator system; a microwave amplifier and transmitter system; and a satellite bus system. A schematic illustration of the PSS structure is given in Figure 2. The PSS is a disc-shaped satellite with a large disc with a diameter of 40m. The disc is composed of the autonomous microwave transmitter modules. The PSS is folded at the time of the launch as indicated in Figure 2 and is deployed in space. One side of the disc is the solar-cell array and the other side is the microwave transmitter. The solar-cell side should be controlled to direct to the sun. This sun-oriented attitude control naturally limits the direction of the microwave beam, even though the beam can be steered in fairly wide angles by the active phase array. To increase the scanning range of the beam, we propose, as an option, to attach a reflector of sun light over the solar-cell array surface. The addition of the optional reflector of the sun light enlarges the service area to which the PSS can direct its microwave beam. Autonomous Microwave Transmitter Module A block diagram of the Transmitter Module is shown in Figure 3. The FET power amplifiers are power-supplied by the DC input from the solar-cells and generate 24 GHz microwave. A new scheme retro-directive phase control system is used for the beam control. A pilot signal of one third of the transmitting frequency is used. This simple conjugate phase generator can determine the phase of the transmitted microwave
with no phase ambiguity because only multipliers are used and no dividers are used in the circuit. Transmitting Antennas and Rectennas Microstrip antennas are used for the active phased array of the autonomous transmitter module. Inter-antenna distance is chosen to be a half wavelength (~ 0.63 cm) to suppress the grating lobes. The merit of using the microstrip antennas is their light weight and ease of mass production. A sphere-shaped rectenna is recommended to attain the omni-directional directivity as shown in Figure 4. Each antenna element imbedded on the surface of the sphere-shaped balloon is a crossed dipole to achieve a light-weight rectenna. Missions of the PSS Possible missions of the PSS are; 1. Power transmission to small satellites 2. Power transmission to space stations 3. Experimental tests of energy transmission via microwave in space. The first two missions have practical purposes providing a unique infrastructure of a large power (~ 100 kW) supply to customers through a light-weighted and simple rectenna system. The third mission is a verification test of the functions of the transmitter modules and beam control system developed and adopted for the PSS. The tests will cover 1. Pointing accuracy and time response of the microwave energy beam, 2. Spatial distribution of the transmitted power produced by the side and grating lobes of the transmitter, 3. Frequency stability of the FET microwave generator and its noise characteristics, 4. Heat control of the module, 5. Attitude stability and mechanical distortion of the large disc of the modules. References [1] GLASER, P.E., (1968) Power from the Sun: Its future, Science, vl62, p857-886 [2] HANLEY, G. M. (ed ), (1980) Satellite Power System (SPS) Concept Definition Study, NASA CR, 3317,.
Design of a Low-Cost, Earth to Space Power Beaming Demonstration KEITH ROGERSt Introduction The concept of collecting the sun's energy in space and transmitting the energy down to the Earth's surface has now been around for about a quarter of a century. In that time, many studies have been done, but little actual experimentation has been completed. Power beaming experiments and demonstrations have so far covered distances of at most a few hundred meters, and only one such demonstration has been conducted in space. [1] One of the reasons for this lack of hardware experimentation is the formidable size of the proposed solar power stations. Even the more modest examples, such as Japan's SPS 2000 program, have price tags measured in billions of dollars. [2] At various times a terraced approach to developing the necessary technologies has been suggested, but even these plans soon progress to the point where expensive space-based demonstrations are necessary. [3] This paper will examine a simpler, cheaper way of demonstrating some of the important technologies for the eventual goal of providing beamed power for space and Earth. What really needs to be examined before taking the leap to megawatt level power satellites is the effects of beaming power through the atmosphere. Doing this with space to Earth power beaming experiments is expensive; one needs to put into orbit not only a fairly hefty power source but also a transmitting antenna with an aperture large enough to focus the beam so that it can be received in a reasonably sized area on Earth. Combine these factors with the control needs for pointing such a ponderous structure and fuel requirements for maintaining its orbit, and the costs really begin to add up. Yet if we reverse the equation, and beam power from Earth to space, things look quite different. Now our hefty power source and large aperture antenna do not need to be sent into orbit; in fact they have already been built. Our target can be a small, cheap microsatellite with minimal control systems and almost no fuel. Associated with this vastly reduced cost is a corresponding decrease in the time and organizational work necessary to carry out the project. Clearly, an Earth to space power beaming experiment would appear to be cheaper than a corresponding experiment in the other direction. But would it accomplish the same scientific objectives? A space to Earth power beaming experiment would test a number of different things. It would measure atmospheric absorption under different weather conditions. Atmospheric scattering effects would be examined, as would sidelobe strengths and distributions. All of these things could be assessed equally or nearly as well going from Earth to space. In addition, an Earth to space power beaming experiment could accumulate data on rectenna efficiencies in space and their degradation over time, information that would be useful to future space t MIT, 790 Main St., #10, Cambridge, MA 02139
to space power beaming projects. Sacrificed would be data on rectenna performance on Earth, but this is easy enough to obtain elsewhere. Also missing from this experiment would be the experience with mid-sized deployable structures that might be gained from a mid-range space to Earth demonstration. The structures used in a mid-sized satellite would probably bear little resemblance to those used for a large-scale power satellite, however, so experience with the smaller satellites would likely be little help with the large scale construction needed for larger satellites. This paper examines the possibilities for conducting an inexpensive Earth to space power beaming demonstration using existing facilities. Design trade-offs will be discussed for both ends of the link: on the ground, potential facilities will be examined for power, transmission frequency, aperture, and tracking capability; while in space issues of orbit selection, launch vehicle choice, and satellite design will be addressed. Possible objectives of such a mission will be discussed, and a brief look at costs and time schedules included. Recommendations will be made with respect to these issues, and areas which need further study will be pointed out. It should be emphasized that this paper does not pretend to present a complete, workable design. The exact form of an actual Earth to Space power beaming experiment of this type will depend on the specific scientific objectives to be fulfilled and on the budgetary constraints that the program must operate within. This paper merely points out the benefits of the approach and the issues that will need to be addressed to bring the concept into reality. Facilities There are two broad categories of facilities which would be appropriate as radiators for an Earth to space power beaming experiment: military and civil. The military systems of interest are the large radar systems used during the past 40 years as a part of ballistic missile early warning systems. European systems with similar intent exist, but are smaller and less powerful than their American counterparts and hence are not considered here. Little is known of the corresponding Russian systems, but from the sketchy information available from Jane's Defence Data it seems that they would be largely inappropriate for power beaming. The only serious Russian candidate for the type of mission considered in this paper was destroyed in 1992 as part of a US-FSU arms reduction pact. Relevant civil systems are radar systems used for astronomical observations and spacecraft tracking. Table 1 shows some of the vital statistics of the major US military and civilian radar systems. In a brief survey of the table, the facilities at Arecibo Observatory in Puerto Rico stand out from the rest. Its 305 meter dish provides more than 40 times the radiator area of the next largest system, its S-band radar uses a frequency five times higher than most of the military radars, and it is capable of continuous power beaming. It’s closest competitor, COBRA DANE, can generate with its pulses a power density only 17% of Arecibo's, with an average of 1%. In addition to a lower power density, COBRA DANE'S remote polar location, northward facing, and low elevation detract from its utility in a power beaming experiment. In contrast, Arecibo has a low-latitude location and points toward the sky, not the horizon.
Arecibo has two large radar systems suitable for power beaming demonstrations. A third system can be used for ionospheric heating experiments but is not really suitable for power transmission. These radars were used to beam power into the upper atmosphere (but not through it) in 1977 during the Ionosphere / Microwave Beam Interaction Study. [8] The first of these radars, which transmits at 430 MHz, has a peak power of about 2 MW. a 6% duly cycle, and an antenna with 61.5 db of gain. The second radar, which may be of most interest for power beaming, transmits at a frequency of 2.38 GHz with a power level of 400 kW and a continuous duty cycle. It’s antenna has a gain of 71 db. The 2.38 GHz radar at Arecibo is probably the most powerful continuous wave system on Earth at the present time. [7] The facility itself centers around a dish 305 meters in diameter, the largest in the world. Due to its large size, and the fact that it takes up an entire small valley as shown in Figure 1, the dish itself cannot be moved. Pointing is achieved by moving the transmitting antenna which hangs over the dish suspended by a network of cables. This imposes two effective limitations on the use of the facility. The first of these is that it can point at most 20° from zenith. This is a fairly restrictive limit on the area of the sky which can be covered, but it should be noted that Arecibo itself lies at a latitude of 18.3°, and so some equatorial orbits could be covered, though not GEO as will be shown later. The major problem with the facility is that it cannot track quickly; it can follow planets but not satellites. The transmitters focus at infinity only, but this should not be a significant problem.
All of these numbers would seem to make Arecibo an ideal facility for the type of demonstration envisioned except for the problem with tracking. Figure 2 shows the tracking rate necessary for satellites at various orbital heights. For a satellite moving at about 8 km/sec in a 1000 km orbit, tracking on the order of 0.5°/sec is necessary. Tracking at this rate is not within Arecibo’s nominal capabilities, but it is conceivable that some creative modifications could give the facility a limited tracking capability. One possibility for this sort of modification makes use of the fact that Arecibo has two different modes in which it can move its antenna. One of these is a finely controlled motion used for tracking planets and other celestial bodies; the other is a slew mode used for rapid, but less controlled adjustments to pointing. By transmitting while in slew mode, it might be possible to track the satellite as it went overhead. Unfortunately, the slew rate is not very fast: it can move at 24°/min in the azimuthal direction but only 1.85°/min across the zenith. This is not fast enough to track a satellite in a low orbit, though it might be used for one at 5,000 km altitude or greater. Even in those situations, however, it is unclear whether the control level in slew mode would be good enough to keep a satellite within the transmission beam spot. Still, it does seem possible that the slew systems could be modified to give better performance in specific situations. For the remainder of this paper, design trade-offs will focus around the use of Arecibo as the primary transmission source. This does not mean, however, that the military systems described above will have no role. For the best of these systems the average power received by the satellite would be on the order of 1% of that received
from an Arecibo transmission. In their pulsed mode some of the military radars can achieve power densities closer to what Arecibo can generate, but only for extremely short periods. This would be inadequate for most demonstration purposes, but might still be high enough to take useful scientific data. In addition, use of military facilities for a peaceful power transmission experiment might have positive public relations effects that would make their use more attractive. Orbital Considerations Totally aside from the issue of what ground facilities could be used, it is worth remembering that there are two parts to any beamed power demonstration: a transmitter and a receiver. For this particular demonstration, the receiver is a small spacecraft. To determine the characteristics and cost of the receiving spacecraft, it is necessary to take a look at the impact of various orbit choices on the mission. On first consideration, one would desire either a geostationary or low equatorial orbit. A GEO orbit would overcome the tracking problem, while a lower equatorial orbit would pass within Arecibo’s arc several times a day. Figure 3 shows the problem with this concept. Arecibo can only see objects along the equator if they are over 70.000 km away. Note that BMEWS and COBRA DANE radars would also be unable to follow equatorial orbits due to both latitude and facing. Since we can't have an equatorial orbit, and several potentially useful military radars are located at high latitudes, a relatively high inclination orbit would seem to be in order.
An orbit is desired with inclination above Arecibo’s 18.3° and accessible with minimal AV. The orbit need not pass over Arecibo every day, but as much regularity as possible would be desired for scheduling and facilitating possible tracking systems modifications. The exact orbital parameters depend on the results of several trade-offs described below, but the height and inclination should be calculated with a view toward making the orbital period (taking into account the regression of the nodes) an integral fraction of a sidereal day. The choice of orbital height of the satellite is closely related to the specific objectives of its mission. As shown in Figure 4, as orbital height increases, the time that the satellite will be in the beam increases while the power received decreases. A second trade-off is made when deciding which of the radars to use. If the 430 MHz beam is used instead of the 2.38 GHz beam, the amount of power received goes down, but the time in the beam is increased by about a factor of six. Several other factors help to determine the orbit selected for the spacecraft. For a small satellite with a relatively large collector such as the one considered here, drag is a very significant parameter in choosing the orbital height. In an orbit up to 400-500 km, the satellite’s lifetime would be relatively short. For altitudes of 700 km or greater, drag ceases to be a significant problem and larger structures can be employed. Another factor which tends to push the satellite toward a higher altitude is that drag forces at lower altitudes tend to disrupt any attempt to use gravity gradient stabilization, necessitating a more complex and expensive control system. Finally, one must consider the issue of orbital debris, perhaps a serious problem if an inflatable reflector as described below is to be used. From this perspective, an orbit of around 1200 km would be desirable. A high altitude such as this would also give the advantage of having a transit time through the 430 MHz beam of a little over a second. Unfortunately, the trade-offs described above are largely irrelevant for a satellite demonstration aimed at costs under $10 million. To launch a satellite within that sort of budget there are only a few possible vehicles. Of these, the major candidates would appear to be the space shuttle launched “Get-Away Special” (GAS) and the Ariane ASAP ring. As both of these systems are severely weight and volume restrained,
carrying a propulsion system for anything other than attitude control would drastically reduce the usable system mass. Thus the orbit of the satellite is basically determined by the orbit of the main spacecraft. In the case of Ariane ASAP launches it is possible to achieve high inclination orbits, and some small amount of phase adjustment may be used to increase the number of passes over the target radar or radars. In the case of the GAS, the orbit is that of the shuttle. Figure 5 shows the frequency with which typical orbits of these types might pass within view of Arecibo. Looking at these orbits, and noting that for the ERS-1 example the altitude is 785 km, it seems clear that an ASAP launch is dictated. Mission Objectives To choose a proper vehicle configuration within the limitations of such a small spacecraft we must decide exactly what we want the demonstration to achieve. To do this, three major aspects of the demonstration’s mission must be considered: collection of scientific data, demonstration of the reception of useful power levels, and publicity. One possibility is to orient the mission mainly around low cost and collection of scientific data. Here the main objective would be collection of data regarding beam scattering, sidelobe strength, frequency dispersion, atmospheric absorption under different weather conditions, and rectenna efficiency. Excellent information would be
gained about rectenna performance over an extended period in the hostile conditions of the orbital environment. For this type of mission the level of power received could probably generate useful scientific data even if the power received was only on the order of milliwatts. However, the lower the power levels received, the more tenuous the connection with the high level power beaming which the experiment is supposed to model. Keeping that in mind, one should remember the objective of showing that beamed power can be received over long distances at power levels high enough to be useful. The problems with reception of milliwatt power levels have already been discussed. With a single watt of received power, a small transponder could be powered. If power on the order of 10-100 watts was received, a small light bulb could be operated for the time the satellite was in the beam. If the step was made to kilowatts of power received, almost any space system now in operation could be powered. The third mission objective that must be addressed is that of publicity, a factor often ignored or downplayed in pure scientific missions. A highly visible
demonstration can attract f unding both for itself and for future experiments in the same area. For instance, on the single watt level, a transponder could send out a “beep” or Morse-code signal saying something like “power received” strongly enough for amateur radio operators to pick up. On the next higher power level, a light bulb could serve as a flash bright enough to illuminate a logo (“Eat at Joe’s,” for instance) printed on the receiver long enough for a photograph to be taken. At the kilowatt level, a light could be set to flash on and off (in Morse code, perhaps) and could even be visible from Earth with a small telescope. The determining factor for all of these possibilities is, of course, the area of the microwave collector. The equipment necessary for monitoring the received power levels and frequency dispersion and transmitting the data back to Earth can easily be made compact and lightweight, so it does not impose any real restrictions. For reception on the order of a single watt, the rectenna area need only be about 4 m2. Such a rectenna could be easily deployed from a small package using existing, well- tested technology. To get the larger receiver area necessary for demonstrations on the 100 watt scale, more innovative types of deployable structures would have to be used. Reception on the kilowatt scale with a collector deployed from a microsatellite is probably not feasible in the near term. Vehicle Configuration The choice of an ASAP platform puts rather stringent restrictions on the size, shape, and mass of the receiving satellite. The ASAP ring lies on top of the Ariane H10 upper stage, and is capable of carrying up to six separate payloads of up to 50 kg each. Each of these positions can accommodate a payload with dimensions equal to a 45 cm cube, though exceptions are sometimes made allowing the payload’s height to be up to 60 cm or so depending on the nature of the mission’s main payload. Individual payloads on the ring can be connected to each other by wires. [9] Working within the above constraints, a two-section spacecraft is envisioned. One position on the ASAP platform would be taken up by the main satellite equipment, including sensors, data-handling equipment, a transponder for data transmission, an independent power supply, and whatever else is necessary for the demonstration chosen. The other position would be connected to the main satellite bus by wires strung along the ASAP ring, and would be used to house an inflatable reflector. When deployed, the spacecraft would look something like the one shown in Figure 6. The inflatable reflector would be transparent on one side with a reflective parabolic inner surface. Connected by wires to the main bus about 12.5 m away, the satellite would be gravity-gradient stabilized and would always point toward Earth. Oscillations could be damped out using small RCS thrusters. Microwaves transmitted from Arecibo would be collected by the reflector and focused on a small rectenna on the surface of the main spacecraft. There are several advantages which come from the use of such an inflatable reflector to collect the beamed power. The main three are in mass, volume, and area. Using an inflatable rigidifiable antenna allows one to get the largest surface area
reflector into the smallest packaging volume at the least cost in mass. 1986 estimates place the mass of a 10 meter inflatable antenna at about 60 kg, and volume requirement at about 0.3 cubic meters. The use of a reflector also allows a two-part satellite structure which can be gravity-gradient stabilized, and enables the use of a small rectenna placed physically near to the satellite’s other electronics and measuring instruments. 110,11] There are also difficulties associated with the design type chosen. Parabolic reflectors are intended to focus radiation coming in directly along the boresight. This is fine for the cases where the satellite is coming in directly above Arecibo, but in most instances the satellite will appear somewhere else within the 20° from zenith toward which Arecibo can direct its transmission. To use the parabolic reflector properly in these situations, the satellite would have to align itself with its axis along the path of the incoming radiation. That, however, would require the satellite to give up its simple gravity gradient stabilized control system for a more complex, massive, and expensive system. An alternative to this would be to use the guide wires connecting the reflector to the satellite bus to tilt the reflector toward the incident radiation. Figure 7 shows an example of this situation. The incident radiation comes in at an angle but is still focused onto the rectenna surface. The change in inertia caused by tilting the reflector is not enough to affect the directions of the satellite’s principal axes significantly, so the overall spacecraft maintains the same orientation as before. Figure 8 shows the same situation with a more three-dimensional perspective. It is important to note that though the incident radiation is all concentrated onto the rectenna plate it is not all concentrated onto a point as it is when the incident radiation comes in along the boresight. Figure 9 is a close-up view of the situation in Figure 8, showing the pattern of light incident on the rectenna plate. The concentrated radio waves form a clear clover-like pattern which varies greatly in intensity with location. This non-uniform intensity pattern will impose difficulties on rectenna design and may result in lowered efficiency. In addition to the variation in intensity, however, the pattern shown in Figure 9 also implies a variation in phase. The special geometry allowing a parabolic reflector to focus light in phase only works when the light is incident along the boresight. For off- boresight geometries, the individual light rays have to travel differing distances and hence come in out of phase. This phase shift effect is highly dependent on the specific geometry of the satellite and can cause drops in efficiency of 50% or more. In addition, it is possible that the difficulty of predicting the exact extent of this effect would obscure measurements taken by the scientific instruments. Further investigation needs to be done to determine exactly what the effects of such an off-axis reflector based system would be. There is one further difficulty entailed by this reflector-based design. One degree or better pointing accuracy would be required to ensure that the incoming radiation was focused properly on the rectenna. This pointing requirement would also have to be achieved before the satellite enters the transmission beam, as the total time in the beam is only one or two tenths of a second. This does impose further restrictions
on the satellite design, but not insurmountable ones. A link with the satellite could readily be established some time before the actual pass over Arecibo, giving the onboard control systems plenty of time to adjust.
Mass restrictions should not be a problem; designs for a 10 tn inflatable rigidifiable antenna for the QUASAT VLB1 mission quote a total mass of 42.05 kg, which includes the main chamber torus, pressurization subsystem, and stowage elements. [10] As the reflector is being used to collect rather than transmit, less accuracy is needed in manufacturing and the overall antenna structure can be simplified, so it may be possible to fit an even larger structure within the mass limits.
Packing constraints would probably prevent the use of a larger reflector if an ASAP launch was used, however. Volume vs. reflector diameter estimates put the required volume for packaging a 10 meter reflector at about 0.3 cubic meters, too large to fit in an ASAP fairing. The simplifications inherent in use of the reflector for reception rather than transmission may enable more efficient packing and bring the volume for a 10 meter reflector within the ASAP constraints. Table 2 shows some of the operating statistics for a satellite of the type described above flying in the 785 km ERS-1 orbit.
Power received averages about 60 W over the diameter of the first minimum for the 2.38 GHz, and 10 W for the 430 MHz radar. Times of passage are about 0.1 and 0.6 seconds, respectively. The 60 W figure is enough to power a photograph flash as described above. A corporate or departmental logo could be printed on the outer surface of the reflector with something transparent to microwaves but opaque to visible light. The choice of the logo to print on the reflector could be determined by a competitive bidding process with the proceeds going to help finance the project. Program Costs The objective of this design example was to demonstrate trans-atmospheric microwave beaming as well as microwave reception in space for under $10 million. Program costs would include satellite construction, satellite launch, use of the Arecibo radar systems, ground station operations for data reception from the satellite, data processing costs, and staff salaries. Fabrication of the inflatable reflector would probably constitute the largest program cost. Recent estimates place the cost of a 10 meter inflatable antenna at about $6 million. [11] Taking into account the structural simplification of the intended receiving reflector as opposed to a transmitting antenna, one could use that figure as the cost for the entire satellite. The launch costs for an entire ASAP ring of 4 or 6 microsats have been quoted at around $1 million. These costs are somewhat variable, and universities have been known to get much cheaper launches. So one can guess that the cost to launch a satellite taking two of the positions would be around $500,000. Operating costs for Arecibo would not be high; with an average encounter rate of 5 passes/inonth over mission life of about two years, 120 data collection runs could be made. According to Mike Sulzer of Arecibo, "It is a little hard for me to estimate the cost of running the 2380 MHz radar since we do not charge (although we do recover certain exceptional expenses in infrequent cases). I would guess several thousand dollars an hour if you need a number. It really is not practical to think about seconds or minutes; you need at least two hours to get things set up and going." [7] So at $5,000/hour for 120 2-hour periods, the cost would at worst be $1.2 million, and at best would be free. With another couple of hundred thousand dollars thrown in for ground station costs and staff salaries, the total program cost might come to about $8 million. So even if the cost of producing the inflatable was substantially greater than expected, the program would still meet the $10 million target. Time-Table It is believed that this program could be carried out over a period of 5 years starting in 1994. Design would begin in 1994 and last throughout the year. Hardware development would begin in 1995, and actual manufacturing of the satellite systems in early 1996. Testing would be started in mid-1997 while manufacturing of some of the
subsystems was still in progress. Transport of the finished spacecraft to the launch site would take place in the end of 1998 for a launch in early 1999. Most of the necessary equipment would be fairly easy to design and construct; the only question would be procurement of the inflatable reflector. Meeting the 1999 launch date would be essential, as the last ASAP launches are planned for sometime in that period, after which launches will be carried out mainly by the Ariane-5. An Ariane-5 ASAP ring is planned and could probably be used to launch a similar experiment (possibly even one of greater size), but the time-frame for such a system is presently unknown. Alternative Possibilities The design presented above was oriented at producing a quick and dirty solution to the problem. As an alternative, or a follow-up mission, it might be interesting to try a similar experiment on a slightly larger scale. Using a dedicated launch of a Pegasus or Delta to put a larger inflatable collector into a higher orbit has very interesting possibilities. Echo 1 and 2, launched in 1960 and 1964, were aluminized mylar balloons of 30.5 and 40 meters in diameter, respectively. They were the first man-made objects in space to be visible from Earth with the naked eye. Over thirty years later it should surely be possible to produce an inflatable reflector of 100 or even 200 meters in diameter. Such a large satellite could receive tens of kilowatts of power in a relatively debris-free 1200 km polar orbit and really give some interesting insights into the problems of transmitting and receiving high power levels. Or if distance beaming or long-term continuous beaming was the technology that needed to be demonstrated, such a huge structure could receive significant levels of power even at GEO. At the other end of the spectrum, it has been suggested that some military spy satellites may already have on board equipment capable of receiving and measuring incident microwave power. A program using these existing assets instead of requiring launch of a separate satellite would cost next to nothing. References [1] AKIBA, R., K MIURA, M HINADA, H MATSUMOTO, AND N KAYA (1993) ISY-METS Rocket Experiment, ISAS Report #652, Kanagawa ISSN #0285-6808 [2] NAGATOMO. N. AND I. KIYOHIKO (1991), An Evolutionary Satellite Power System for Inter-national Demonstration in Developing Nations 2nd International Symposium SPS 91: “Power from Space”, Paris/ Gif-sur-Yvette. [3] GLASER, P, (1991) The Solar Power Satellites Option Re-examined 2nd International Symposium SPS91: “Powerfrom Space”, Paris/ Gif-sur-Yvette. [4] STEIN, KENNETH J., “BMEWS Update Program Progresses at Thule Site”, A WST, pp 90-91, August 20, 1984.
[5] BROOKNER, E (1981) “A Review of Array Radars"’, Microwave Journal, Vol. 24 No. 10. [61 Pave Paws, BMEWS Radar Site Updates Will Broaden Missile Threat Coverage, AWST, pp. 52-54, December 9, 1985. [7] SULZER, MICHAEL (1992) Personal Contact. [8] DUNCAN, LEWIS M., AND WM. E. GORDON (1977) Final Report: Ionosphere/ Microwave Beam Interaction Study NASA Contract Number NAS9-15212. [9] ARIANESPACE (1990) A.S.A.P. User’s Manual. [10]BERNASCONI, M. C. (1984) Large Spaceborne Antenna Reflectors Using Inflatable Space Rigidised Structures Workshop on Mechanical Technology for Antennas (ESA). [11] SUE, M.K., ed., Personal Access Satellite System (PASS) Study, Fiscal Year 1989 Results, NASA CR-188808, N92-10116, JPL, 1990. [ 12|BERNASCONI, M.C., E. PAGANA, AND G. G. REIBALD1. (1987) Large Inflatable, Space-Rigidized Antenna Reflectors: Land Mobile Services Development, Paper No. IAF 87-315, 38th IAF Congress. [13] SPACE MISSION ANALYSIS AND DESIGN (1991) Wertz, James R. and Wiley J. Larson (Eds) (Boston, Kluwer). [14] ROTH, K.R. et. al. (1989) The Kiernan Reentry Measurements System on Kwajalein Atoll Lincoln Laboratory Journal, Vol. 2, No. 2, pp. 247-75. [15] JANE’S C3I SYSTEMS (1990) Peter Rackham (Ed), Jane’s Defence Data. [16] SPACE SOLAR POWER PROGRAM (1992) ISU 1992 Summer Session Final Design Project Report. [17] ZORPETTE, GLENN, “Kwajalein's new role: radars for SDI”, IEEE Spectrum, pp. 64-69, March 1989. [18] WILKES, OWEN, MEGAN VAN FRANK, AND PETER HAYES (1991) Chasing Gravity’s Rainbow: Kwajalein and US Ballistic Missile Testing, (Canberra, Nautilus Pacific Research).
An Evolutionary Satellite Power System for International Demonstration in Developing Nations MAKOTO NAGATOMO*, KIYOHIKO ITOH** SUMMARY: The ISAS solar power satellite working group is working on a concept of an SPS strawman model for demonstration of electric power supply to customers at the earliest opportunity. The SPS is modularized, so that each unit can be launched by a commercial launcher to an equatorial low earth orbit where it is assembled automatically. The satellite can supply electric power by microwave to rectennas at every pass. Based on this model, technological and programmatic characteristics of a small SPS are discussed. Introduction The SPS Reference System was designed for the SPS Concept Development and Evaluation Program (CDEP) of the U.S.A., to study the wide range of problems anticipated when such a system is introduced as a national power system of the world's largest industrialized country. As a result, many understandings have been obtained, and uncertain issues needing to be studied further have been indicated. The study concluded that the system was technically feasible and further study is required [1 and 2], The ISAS Solar Power Satellite Working Group is interested in a much smaller SPS as a strawman model study, which will be introduced and discussed in this paper. At first, We will explain briefly about the working group as the background of the strawman model study. ISAS SPS Working Group ISAS, which is responsible for implementation of the Japanese space science program, provides researchers of universities and government agencies with the convenience of nationwide research activities. A working group of the organization usually functions as a group to develop a concept of and conduct definition studies on a space mission. The purpose of the SPS working group established in 1987 was different from the other working groups, since it never intended to plan an SPS project for ISAS but, to remain as a research group to investigate the feasibility of "power from space" which was shown by the CDEP and the Reference system. The Scope of the SPS Working Group is described as follows: "The solar power satellite was proposed (by Glaser) to solve future problems caused by activities of human beings on the global scale [3], The research areas of SPS *Professor, Space Power Systems Section, Institute of Space and Astronautical Science, 3-1-1 Yoshinodai, Sagamihara 229, Japan, Fax number for the section: +81-427-59-4239 **Professor, Hokkaido University, North 13, West 8, Kita-ku, Sapporo 060 Japan
are concerned with not only technology and engineering, but also big problems such as "large scale project", "global energy production" and "exploitation of extraterrestrial resources", as well as economical problems such as "large scale and long-term investment" and "risk analysis". The earth environment, which will be related to SPS, is already a societal issue of global scale. Many research areas are beyond the capacity of the working group. The primary objective of the working group is not to plan an operational SPS, such as the DOE/ NASA Reference System. As indicated by the US study, research is most important at the present. In this respect, the members of the working group should find a common interest in connection with the future of SPS, and make plans for space experiments along with a scenario of research and development for SPS. It should be emphasized that the development of high power technology for space use is not the main objective of the research". The working group is divided into thirteen subgroups by specialized research fields. Nine of them arc concerned with studies on SPS subsystems and technologies. Other four subgroups are for studies on the effects of interaction of the SPS operation upon the environment. The individual research fields of the thirteen subgroups are listed as follows; Study on Subsystems and Technologies Microwave transmission Microwave reception Large space structures Guidance and control Laser technology Photovoltaic technology Thermodynamic power generator Propulsion Space robotics Study on Environmental Interaction Effects Spacecraft environment Space electromagnetic environment Communication systems Biology and ecology It is noted that the working group is not a task force type of organization that should be functionally organized to accomplish a certain duty given from outside, but a group of researchers who are interested in various aspects of SPS. Thus, the subgroups have been voluntarily organized by researchers with common interests in each, and are coordinated as a working group. One of the research areas is concerned with the strawman model concept.
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