Space Power Technological, Economic and Societal Issues in Space Systems Development Volume 7 Numbers 3/4 1988 -KWfaM— publishing company
SPACE POWER Published under the auspices of the SUNS AT Energy Council EDITOR Andrew Hall Cutler, Space Studies Institute Space Power is an 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 resource utilization, space manufacturing, and other areas related to the development and use of space for the long-term benefit of humanity. Papers should be of general and lasting interest and should be written so as to make them accessible to technically educated professionals who may not have worked in the specific area discussed in the paper. Editorial and opinion pieces of approximately one journal page in length will occasionally be considered if they are well argued and pertinent to the content of the journal. Submissions should represent the original work of the authors and should not have appeared elsewhere in substantially the same form. Proposals for review papers are encouraged and will be considered by the Editor on an individual basis. Editorial Correspondence: Dr Andrew Hall Cutler can be reached by telephone at (619) 284-2779, and his address is 3030 Suncrest No. 214, San Diego, CA 92116, USA. Dr Cutler should be consulted to discuss the appropriateness of a given paper or topic for publication in the journal, or to submit papers to it. Questions and suggestions about editorial policy, scope and criteria should initially be directed to him, although they may be passed on to an Associate Editor. Details concerning the preparation and submission of manuscripts can be found on the inside back cover of each issue. Business correspondence, including orders and remittances to subscriptions, advertisements, back numbers and offprints, should be addressed to the publishers: Carfax Publishing Company, P.O. Box 25, Abingdon, Oxfordshire 0X14 3UE, United Kingdom. The journal is published in four issues which constitute one volume. An annual index and titlepage is bound in the December issue. ISSN 0951-5089 Cover: An artist's concept of the enhanced configuration of the permanently manned Space Station, produced by Rockwell International. The enhanced configuration includes an upper and lower keel for attaching external payloads, a 50 kilowatt solar dynamic system mounted on the ends of the transverse boom, a servicing bay and a co-orbiting platform (not pictured). Reproduced by Courtesy of NASA, Washington, DC, USA. © 1988, SUNSAT Energy Council
SPACE POWER Volume 7 Numbers 3/4 1988 Nobuhiro Tanatsugu & Yutaka Momose. Conceptual Design of a Solar Dynamic Power Generation System for a Space Experiment 227 Douglas R. Sparks. Fabrication of Large Photovoltaic Arrays in Space from Lunar Materials 235 Samiran Das & I. Selvaraj. Solar Array Mechanisms for Indian Satellites— APPLE, IRS and INSAT-IITS 247 Liu Ling-Hui. Development and Flight Performances of China's Satellites' Power Systems 261 Gert Eggers. Advanced Power Supply and Distribution Systems for COLUMBUS 267 Gordon R. Woodcock. Economic Potentials for Extraterrestrial Resources Utilization 293 S. Chandrashekar. An Approach to Selecting Appropriate Treaty Limits on Laser Antisatellite Weapons 311 Kelly Parks & Bevin McKinney. Mars on $800 000 a Day 327 Albert A. Harrison. Beyond Earthnocentrism: Anthropology on the High Frontier 345 Isabelle Durand-Zaleski. Man in Space: A Survey of the Medical Literature 353 William E. MacDaniel. Scenario for Extraterrestrial Civilization 365
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Conceptual Design of a Solar Dynamic Power Generation System for a Space Experiment NOBUHIRO TANATSUGU & YUTAKA MOMOSE Summary This paper presents an outline of a conceptual design of a small solar power generation system for a space experiment using a Stirling engine as a thermodynamic power converter. This system produces 3 kWe of rated electric power continuously in low Earth orbit. This power plant has the following two features. (1) It employs a thermal energy storage system on the hot side of a Stirling engine in order to work continuously in the course of eclipse as well as insolation. Latent heat of molten LiF is utilized for thermal energy storage. (2) The Stirling engine consists of one displacer driven by an electric motor and two free power pistons moving in the opposite direction from each other. This leads to a reduction of weight, size and vibration. A preliminary study of the LiF unit and the linear induction generator have been almost completed. A protomodel of the Stirling engine generator was integrated in 1987 and tested from 1988. Experiment in orbit is scheduled for 1994 using the Space Flyer Unit. Introduction The Japanese Experiment Module (JEM) attached to the US space station and the unmanned Space Flyer Unit (SFU) are planned to be in operation by the mid-1990s. In such space activities, a larger amount of electric power will be required for both housekeeping and miscellaneous experiments. In a large power generation plant, more benefits come from a Solar Dynamic Power System due to its higher conversion efficiency. The higher conversion efficiency of the power plant in utilizing solar energy can reduce the area of the solar collector and therefore the orbital decay due to atmospheric drag, especially in low earth orbit. This saves reboost propellant for the space station. When more than 100 kWe is required, nuclear power plants should also be considered. The Institute of Space and Astronautical Science has a plan to experiment with a solar dynamic power system in space on the SFU. In this mission the electric power produced is supplied to the MPD thruster and the SFU is boosted to an orbit about 100 km higher. The electric power requirement for the MPD thruster experiment is 3 kWe at 330 V ac and 60 Hz in continuous operation. Nobuhiro Tanatsugu, Associate Professor, Institute of Space and Astronautical Science, 4-6-1, Komaba Meguro-Ku, Tokyo 153, Japan. Yutaka Momose, Aisin Seiki Co. Ltd, 2nd R&D Center, 80, Kowari, Minaminakane-cho, Nishio-City, Aichi Pref., Japan. A version of this paper was presented at the sixth ISAS Space Energy Symposium, 12-13 March 1987.
A non-lubricated free-piston type of Stirling engine and linear type of induction generator are used for energy conversion from thermal to mechanical and from mechanical to electrical energy respectively. Two free power pistons are arranged in such a manner that they undergo reciprocal motion in opposite directions and therefore vibration is reduced to less than 20 microns net movement. The thermal storage system, employed on the hot side of the Stirling engine, stores thermal energy received from the solar collector in the course of insolation and then supplies it to the Stirling engine during eclipse. This system moderates the variation of system temperature caused by the more frequent repetition of insolation and eclipse in low earth orbit than on the ground, and it also makes possible continuous power operation. The conceptual design of a small Stirling engine operator to be used for a space experiment on the SFU is presented in this paper. Fig. 1 shows an artist's concept of an SFU flight operating with the solar thermodynamic power generator. Concept of a Solar Dynamic Power Generation System for a Space Experiment Solar Dynamic Power Generation System with Free-Piston Stirling Engine As shown in Fig. 2, Solar energy collected by the solar collector is changed to thermal energy on the inner surface of the receiving cavity and transferred to the hydrogen gas working fluid of the Stirling engine through the thermal storage system and then used for expansion work in the Stirling cycle. Two free power pistons are arranged in such a manner that they move reciprocally in opposite directions from each other. The
reciprocal motion of free pistons coincides with that of a displacer piston which is driven by an electric motor, forcing the reciprocal frequency of the power pistons to maintain a constant rated value. In the present design, the frequency is set at 60 Hz. The linear induction generator consists of permanent magnets on the moving power pistons and winding coils on the outer circumference of the cylinder. AC current is induced in the stationary coils by the reciprocal movement of the magnet. This linear induction generator is now designed by Toyota Central Research & Development Laboratories, Inc. The details of the Stirling engine generator are illustrated in Fig. 3.
A Cassegrain type solar collector is employed in the present system. Three subsystems are provided for emergency use. The first one is a heater in the cooling loop to prevent cooling water from freezing and the second one is the dummy load to regulate the load power, and the last one is a vent valve for the working fluid. Table I indicates the size, weight, output power and durability of the primary components. Thermal Storage System Various types of energy storage systems (thermal energy, electrical energy and mechanical energy storage) have been proposed in order to supply continuous power during eclipse. In the present system, a thermal energy storage system is employed as shown in Fig. 4(a). Thermal energy is stored by the latent heat of a molten salt such as LiF or LiH. The advantage of this system is that the output power of the engine can be half compared to that of other systems shown in Figs. 4(b) and 4(c) and so its weight and size can be kept to a minimum. In both the electrical energy storage system with a battery and the mechanical energy storage system with a flywheel, the engine has to produce, during insolation, the excess electric power necessary during eclipse. In the system using a flywheel, auxiliary subsystems such as a clutch and a speed change- over transmission have to be provided. The characteristics of latent heat storage materials (LiF and LiH) appropriate for the present system are shown in Table II. LiH is easily decomposed and the hydrogen decomposed transpires out of the space through a metal wall of the container. On the other hand, although LiF has the disadvantage of relatively low latent heat, it is a stable and harmless material. However, LiF has very corrosive characteristics against a metal and also a large volume change occurs in the phase transition from solid to liquid. The latter problem is reduced by mixing MgF2* and by applying a bellows mechanism as shown in Fig. 5. It is also necessary to find the best configuration of the thermal storage system in order to optimize heat transfer. Performance of the Free-Piston Stirling Engine Fig. 6 shows the heat flow for the present solar thermodynamic power generation system at rated operation. The efficiency of the Stirling engine generator is around 31% and overall system efficiency is around 25%, assuming 81% for the solar collector. The efficiency of the linear induction generator is 87%.
Control of the Free-Piston Stirling Engine Generally speaking, the control of a free-piston type of Stirling engine is quite difficult due to the fact that the strength of the gas spring set between a displacer and a power piston is varied in response to a stroke of the power piston and also the temperature itself. In the present system, in order to solve these problems, the output variation caused by an input and/or a load condition is trimmed by using a dummy load. This control system and feature are shown in Fig. 7. The dummy load is always adjusted so as to keep the output power at the requested value.
Control of System Shut Down There are many causes for the shut-down of a power system and a power system for space use has to be shut down safely in every case, especially in emergencies before fatal damage extends to other systems. The system shut-down can be performed by stopping the power input and purging the stored energy (thermal and pressure energy) in the system. The power input can be stopped by pointing the solar collector away
from the sun. The thermal energy stored in the thermal storage system is eliminated by a forced cooling system to the radiator and the pressure energy is eliminated by purging a working fluid through a non-propulsive vent valve. The most preferable shut-down procedure is carried out in accordance with each shut-down situation. Nominal shut-down and non-emergency shut-down are performed so that the system can restart easily and so no working fluid is vented. In case of emergency shut-down and complete shut-down for system retrieval by Shuttle, all the stored energy is eliminated so the system cannot restart. Table III shows the control procedures for emergency shutdown. Conclusions We have proposed a space experiment using a solar thermodynamic power generation system (3 kWe) with a Stirling cycle engine in the second mission of the SFU in 1994. Conceptual design of the flight model for the space experiment is now in progress. We can utilize several useful technologies on the Stirling engine accumulated by the many projects which have been carried out for terrestrial use in Japan. We have to design the Stirling engine taking into consideration both the space environment and the space operation acceptable by SFU and the launch and recovery system. Lubrication, vibration and cooling of the Stirling engine and safe operation are especially important items to be considered. A protomodel of the Stirling engine was integrated in 1987 and tested from 1988. REFERENCE [1] Abe, Y. (1986) Review on Space Solar Dynamic Power System and Thermal Energy Storage, Bui. Electrotech. Lab., 50, 12, pp. 41-62.
Fabrication of Large Photovoltaic Arrays in Space from Lunar Materials DOUGLAS R. SPARKS Summary The manufacturing of large silicon photovoltaic arrays, suitable for use in solar power satellites, from lunar materials is discussed in detail. The system constraints include: the use of lunar materials only; processes must be compatible with zero-gravity, therefore excluding open liquids; all processes should be based on existing technologies and if possible should be continuous for high throughput, as well as automated and easily maintainable to minimize the need for human presence. Purification of materials is covered as is the growth of single and polycrystalline silicon. Processing steps such as doping, gettering, lithography, deposition of metallic and insulating films as well as wirebonding techniques are covered. The possibility of manufacturing highly efficient pointcontact silicon photovoltaics (>27% efficiency) with the aforementioned processing constraints is covered. Introduction Due to simplicity of design and safety considerations, solar power systems are the primary power sources mentioned in recent studies of space habitation and industrialization [1-6]. As has been pointed out in a recent study by Space Research Associates for the Space Studies Institute [6], high initial costs have been the major obstacle to the development of Solar Power Satellite (SPS) systems. It was also pointed out in that study that by utilizing lunar materials instead of transporting materials from earth, the cost of such a project could be significantly reduced. Of the many solar power conversion systems surveyed, the silicon photovoltaic system utilized the highest percentage of lunar material (virtually 99%). Recent advances with point contact silicon solar cells [7], as well as over 20 years of silicon processing experience, make this system perhaps the most likely candidate for use in future large-scale space power generating systems. In this paper the fabrication of silicon photovoltaics in space will be covered. Approximately 2000 tons of silicon, excluding scrap, are used each year to produce semiconductor devices [8], Of this amount only 8% is used for discrete devices, which include solar cells. Based on current estimates [6, 9, 10], from 6000 to 15 000 tons of silicon will be required to fabricate a 10 GW SPS. The use of SPS as a significant terrestrial power source will require roughly three to eight times the amount of device grade silicon as is currently being produced. As has been mentioned in a previous paper [11], the implementation of the SPS concept will require a dramatic shift in Douglas R. Sparks, GM Hughes Electronic Corp., Kokomo, IN 46901, USA.
emphasis in semiconductor device technology. The economic growth of a space-based economy centred around the SPS may be limited by the fabrication and maintenance of solar panels. Previous studies [3, 6] have concluded that the large amount of material required for the fabrications of SPS dictates the use of extraterrestrial materials. The primary components of the lunar soil are known to be A1203/10.3-27.2%, CaO/9.7-15.8%, FeO/5.2-11.3%, MgO/5.8-11.3%, SiO2/39.9-48.1%, and TiO2/0.5-9.4% [12, 13], To reduce the complexity and weight of the processing equipment and to become compatible with a zero-gravity environment, open liquids should not be required. The process should also be able to function in vacuum. The human element is one of the most costly in any space programme. To minimize the number of people required to produce the solar arrays, a continuous, automated process which is easily maintained should be employed. Purification of Materials The first step in the production of silicon devices is the purification of the silicon and other materials used in its production (processing gases, liquids and metallization and insulating materials). The two primary types of impurities which affect silicon solar cells are dopant atoms (boron, aluminum, phosphorous) and transition metals which can reduce the minority carrier lifetime and hence reduce the efficiency of the cells. Phinney et al. [14] proposed that slagging and vacuum distillation be used initially to separate silicon from the other components of the lunar soil. Based on similar systems these two techniques could reduce the levels of impurities down to 0.3-2.0%. The aluminothermic reduction of raw silicon dioxide has recently been utilized by Dietl and Holm [15] to obtain high purity silicon for the fabrication of solar cells. The material produced in this manner has a very low boron content (<lppm) which is very significant because the large segregation coefficient of boron in silicon makes it difficult to obtain high resistivity silicon using a non-chlorine-based purification technique. Since aluminum is readily available from lunar soil this purification technique could be adaptable to space processing. Zone refining is another means of obtaining high purity silicon [16]. As a material freezes, it usually rejects impurities into the remaining melt. The ratio of the impurity concentration incorporated into the solid, to the impurity concentration in the melt, is termed the segregation coefficient. The more impurity that is rejected, the smaller the segregation coefficient. Table I gives the segregation coefficients of several impurities in silicon [17, 18], This phenomenon is used in the zone-refining process [19] to purify silicon as well as many other materials. In this process, the silicon ingot is passed slowly through a succession of heating zones, each of which causes local melting of the silicon for a short distance along the axis of the ingot. Commercial units with up to 12 zones are not uncommon [19], As the ingot moves through each zone, the corresponding molten section, in effect, sweeps the entire length of the crystal. The impurities are collected at the end of the ingot, and this section is usually removed. Table I shows that all elements except for boron, phosphorus and arsenic will easily be reduced to very low levels using this technique. The importance of the aluminothermic reduction technique to reduce levels of boron can be seen here, since zone refining does very little to remove boron. Boron is present in the lunar soil in concentrations of 2.0 to 39.0 ppm. Utilizing a 12 zone, zone refiner aluminothermically reduced quartz could yield silicon with resistivities as high as 5-6 ohms-cm. Such silicon resistivities have
yielded cells with diffusion lengths in excess of 200 um, a length which exceeds the typical cell thickness quoted for use in silicon-based SPS [6]. In addition to silicon a wide variety of other materials can be purified by zone refining. Iron, titanium, aluminum oxide and titanium oxide are materials found in the lunar soil which have been purified using zone refining [19]. Based on the eutectic temperatures found in the phase diagrams of metals [20], the maximum impurity concentration over which zone refining will be effective can be determined, and is given below, Induction, electron beam, and focused radiant heating are the most common techniques currently used to melt materials. It may be possible to used solar concentrators as a heat source in space. While zone refining promises to be a means by which high purity silicon can be produced, the contamination of silicon by incidental contact with transition metal implements must also be avoided. Recent work has shown that common waferhandling equipment is capable of introducing substantial amounts of transition metals to silicon [21, 22]. Transition metals diffuse so fast in silicon that a few seconds at temperatures in excess of 500C can lead to contamination through out a 625 um thick slice [23]. Fortunately these metals can be gettered just as rapidly to inactive areas of
the device when mechanical damage [23, 24], laser damage [23, 25], ion damage [26], or a heavily doped phosphorus layer or polysilicon layer is applied [27, 28], The formation of oxygen precipitates has also been used to getter metals [29]. This method can only be used when sufficient oxygen is present in the silicon. The oxygen usually comes from the wall of the quartz crucible used to hold the molten silicon during crystal growth. In the gettering process metal atoms become trapped at dislocations and grain boundaries by the reduction in lattice strain associated with such a reaction. Such gettering results in the improvement of the minority carrier lifetime [25], An external gettering step can be located at any step in the process provided a heat treatment is employed to allow the metal atoms to diffuse to the gettering site. Crystal Growth The most common method currently used for the growth of silicon crystal is the Czochralski method, in which a crystal is pulled from an open melt. The individual wafers from which devices are made are obtained by slicing the crystal, then lapping and polishing the slices. This method is not suitable for space applications because it employs an open melt, liquids are required for slicing, lapping and polishing and it is very wasteful since a great deal of the raw silicon is lost during the slicing and lapping operations. Material experiments performed during the Skylab missions revealed that the crystal surface facets produced in zero-gravity appear to be atomically smooth [30]. Carruthers [31] has pointed out that it may be possible to perform direct semiconductor processing on such a flat surface without further lapping and polishing. Almost 30 years of research and development has gone into the fabrication of solar cells on ribbon crystals which stems from even earlier work dating back to the 1930s [32] on the growth of thin monocrystalline plates. Only a few techniques satisfy the space fabrication criteria used in this paper. They include the float zone technique which has already been demonstrated for silicon on the space shuttle [33], the sintering of polycrystalline rods, the pulling or extruding of crystal from an enclosed melt, and combinations of the above. As has been pointed out in an earlier study, Stepanov's method appears to be the best candidate for space applications. In Stepanov's method the molten silicon is held in a piston-like chamber, as shown in Fig. 1 [11], the silicon is forced out from a specially shaped orifice to produce a ribbon of silicon. It is likely that the walls of the chamber will be made of quartz which will introduce oxygen into the silicon and allow for internal gettering later on in the process. In addition to the growth of free silicon crystal, films of silicon can be epitaxially grown on a variety of substrates. Most techniques employ silicon chlorides to deposit the films, but a few methods such as molecular beam epitaxy, ion cluster beam epitaxy, and ion beam epitaxy do not require liquids or gases. In the past few years several large conferences have been held in silicon molecular beam epitaxy [34, 35], and so it is expected that technological advancements will continue in this field. A variety of methods which apply silicon films on substrates made of steel, quartz, graphite and polycrystalline silicon have been developed [36-40]. These methods may find application in space although the efficiencies of solar cells fabricated on such substrates have been relatively low [36, 40],
Radiation Damage in Space The bulk properties of silicon can be dramatically affected by radiation damage. Study of the effects of radiation damage on silicon solar cells has been going on since the first satellites were put into orbit. The various radiation sources: solar wind, Van Allen belts, and cosmic rays, have been categorized in an early work published by the American Society for Testing and Materials [41]. Controlled studies using energetic electrons, ions and neutrons have provided even greater insight into the problem of radiation damage of semiconductors. The effect of energetic particles on silicon has perhaps been best characterized as the result of the study of the ion implantation of silicon. It is known that at higher energies ions undergo electronic interactions resulting in little structural damage to the silicon lattice. At lower acceleration energies or as the ion decelerates nuclear interactions occur and the displacement of atoms from their original lattice sites occurs. As the mass of the incident ion increases the damage for a given acceleration energy increases. As the energy increases the depth of damage increases. Since only a threshold energy of <50 eV is required to displace a silicon atom to a stable interstitial site accelerations of several keV to MeV will result in the displacement of thousands of lattice atoms per incoming ion. Electron, neutron and ion damage reduces electron and hole mobility and as a result electrical conductivity. Energy levels are introduced into the band gap of silicon which can lead to compensation and possibly type conversion as well as reduce minority carrier lifetime. Reductions in minority carrier lifetime and carrier mobility will both reduce solar cell efficiency. Damage is generally classified into two broad categories: light and heavy damage. Light damage is characterized by discrete, spatially isolated damage regions and heavy damage is characterized by a continuous amorphous region of damage. The significant difference between these two categories of damage lies in the ease with which this damage can be annealed out. The amorphous region produced by heavy damage will regrow via solid phase epitaxy on the underlying silicon substrate. This regrowth occurs at moderate temperatures (550-650 C) [42]. Light radiation damage is much more difficult to anneal out, and as Table II indicates temperatures as high as 750 C are required to eliminate all the deep levels associated with such damage
[42]. Depending on the mass of the incident particle light doses are generally considered to be less than 1 X 10E15 ions/cm2 [42]. Table III gives the typical proton doses per year and the energies associated with such proton experienced in near Earth space [41], It is apparent from the data shown in Table III that light damage is to be expected in space over a relatively short time period (<1 year). Passive solar annealing, lithium doping and quartz shielding have been means used in the past to reduce the impact of radiation damage on solar cells. For an SPS glass shielding and periodic annealing will most likely be utilized [6]. Anneals in excess of 750 C place special material constraints on the system. Silicon Processing Once the silicon crystal of the desired shape is obtained, a variety of processing steps such as doping, insulator deposition or film growth, metallization and the patterning of the aforementioned steps are performed. Dopants in the part per million to part per billion range are used to change the electrical characteristics of semiconductors. There are two types of dopants: n-type (P, As, Sb) and p-type (B, Al, Ga). Dopants are commonly introduced during crystal growth, or diffused into the silicon from a gas, by gas transport from a dopant oxide or nitride, from a spun-on source, a deposited dope oxide film, by alloying with a dopant metal (Al, Sb), or by ion implantation. If imported gases and liquids are excluded only the alloying and ion implantation techniques can be utilized in space. Alloying is one of the oldest doping techniques used in the semiconductor industry [43]. Aluminum, a common element in the lunar soil, alloys readily with silicon. Selective alloy doping can be obtained by evaporation of Al through a solid mask. Ion
implantation requires a high vacuum in order to prevent collisions between ions and gas molecules, and hence is ideally suited for space applications. An ion implanter consists of an ion source, a magnetic mass analyser that filters unwanted species of ions from the beam, an acceleration, focusing and deflection system, and a target holder. Several solid source systems have been constructed [44, 45]. Pattern definition can be obtained by use of a solid mask as described by Sacher [46] or by controlling the beam to write the desired pattern [47], The implanted ions must be thermally activated, this is usually done at around 900 C [42]. In a continuous space system, a rod- or wire-fed source system could be employed to reduce machine down-time. To keep different areas electrically separated an insulating layer on top of the silicon is needed. Such insulating layers, usually oxides, can also be employed to shield the cells from radiation. Generally a thermally grown oxide layer is used. Thermally grown oxide can be produced from lunar materials but there is no way of etching through to the silicon without the importation photoresists and etchants as well as the aid of gravity, all of which violate the system constraints. Oxide films can be evaporated [48]. Table IV shows the temperatures at which the vapor pressure of common lunar oxides equals 10~2 torr. Silicon monoxide, the most volatile of those shown is produced by heating a mixture of silicon and silicon dioxide under vacuum [49]. Containment of oxides during evaporation is usually done in a refractory metal crucible (Ta, Mo, W) or in an oxide crucible (zirconia). An electron beam or resistance heated crucible is used to obtain such high temperatures. By using solid masks patterning in the dielectric film can be obtained as well. The choice of oxide system to use in solar cells will depend on the electrical, mechanical, and optical properties desired. Electrically the oxides of silicon, aluminum, calcium and magnesium are insulators while those of iron and titanium are semiconductors [50]. Mechanically, a close match between the coefficient of expansion of the oxide and silicon as well as the metal used as a conductor is desirable. The expansion coefficient can be varied by compositional changes in the glass produced [51]. Optically the glass should transmit a high percentage of the incident solar radiation, quartz is perhaps the best at transmitting a high amount of solar radiation. Evaporation of metal can be employed for metallization of the solar cells. Of the possible conductors available (Al, Mg, Fe, Ti, various silicides, and doped polycrystalline silicon) Al and Mg can be eliminated if anneals as high as 750 C must be used to minimize the effect of radiation damage. Titanium and titanium silicides are commonly used in the integrated circuit industry [52] and would be an excellent choice for a single crystal solar cell system. If polycrystalline silicon is used for the photovoltaic system iron and titanium may present a problem with reduced carrier diffusion lengths if repeated high temperature anneals are employed [53]. Solid masking could also be
utilized for pattern definition. For high throughput a spooled, wire-fed evaporator unit could be employed [53], The electrical bond at the metallized pad of a solar cell is the interface between the solar cell and the transmission line. This interconnect is often a major problem in microelectronic systems. For space utilization where the solar cells are being periodically annealed solder wire bonding will not be considered. This still leaves a variety of techniques, which include: thermocompression bonding, ultrasonic wire bonding and welding utilizing lasers and electron beams [54], Thermocompression bonding is further divided into three types: Ball bonding, stitch bonding and wedge bonding. Thermocompression bonding is a process in which both heat and pressure are used to make a diffusion-bonded connection. Ball bonding can most probably be eliminated from consideration since a hydrogen flame is used to sever the wire and to form a metal ball at the end of the severed wire. In wedge bonding the wire is placed under a wedge-shaped tungsten carbide tool prior to the application of force and heat. The wedge tool is internally heated and special scissors are used to cut the wire so this technique is space compatible. Stitch bonding is a series of thermocompression bonds made with the same wire without breaking the wire. The wire is fed through a heated capillary. This technique could also be space compatible. Ultrasonic die bonding is employed to break down the oxide layer grown on the metal pad. Such an oxide layer may not be present in vacuum processed solar cells. Both thermocompression and ultrasonic bonding may be difficult to use with iron and titanium metallized solar cells. Laser or electron-beam welding may be more suitable for such relatively hard metals. If the pads are not annealed with the rest of the solar cell, aluminum may be employed as a wire material. The use of aluminum wire for bonding is a mature technology [55]. All wire-bonding techniques which use coiled wire are very easily automated. Fabrication of Point Contact Solar Cells in Space The fabrication of a very high efficiency silicon solar cell using point-contact technology has recently been announced by Swanson [7], Efficiencies as high as 27.5% have been obtained. Three main design features are responsible for the cell's high efficiencies. First, the front surface through which light enters is textured to trap light in the cell. This texturing is accomplished by chemical etching. Second, all electrical contacts are on the back of the cell so that no metal blocks the incoming light. Third, the current is drawn off through a series of tiny holes instead of from an entire
backside contact. Fig. 2 shows a cross-sectional diagram of such a cell. Swanson believes this design can be further optimized to 29% efficiency. Since most recent SPS studies quote silicon solar cell efficiencies of 14.34-15.75% [6] the use of high efficiency point-contact solar cells would result is considerable weight and hence cost savings. It may be possible to fabricate such a solar cell system in space using the techniques arrived at in this paper. The most difficult step would be the fabrication of an array of holes in an insulating layer. This is typically done using a photoresist mask and a wet chemical etch, both of which violate our special fabrication constraints.
However, by using a two step masked oxide evaporation, as shown in Fig. 3(a) and (b), this task can be accomplished. In the first step, a solid mask consisting of fine parallel wires is employed to deposit strips of oxide on the backside of the silicon. This step is shown in Fig. 3(a). Next a similar mask with wires perpendicular to the already deposited oxide strips is used to apply another group of oxide strips. This will result in an array of silicon contacts as shown in Fig. 3(b). Prior to this oxide deposition step polycrystalline silicon is deposited on the frontside to provide texture. Concurrently with either of the backside oxide deposition steps a frontside oxide deposition is employed to reduce radiation damage. After the contact holes are fabricated a solid mask is used to ion implant first the p-type contacts and then the n-type contacts as shown in Fig. 3(c). This implant is followed by an anneal to eliminate the ion damage. After this doping step a solid mask can be utilized to deposit the metal film. This metallization step is shown in Fig. 3(d). Wire bonding and mounting follow the metal deposition. All these steps can be accomplished with the restraints imposed by this study. Conclusions The fabrication of a variety of different types of silicon solar cells in space with lunar material, no imported gases or chemicals, in zero-gravity and utilizing existing technologies was shown to be feasible. Processing steps such as crystal growth, zone refining, molecular beam epitaxy, ion implantation, alloying, evaporation of metals and oxides and wire bonding were shown to be space compatible. Due to the requirement of annealing radiation damage it was proposed that titanium or possibly iron or a silicide of either, be used as the primary cell conductor. It may be possible to fabricate highly efficient point-contact solar cells in space with the constraints imposed in this study. REFERENCES [1] Glaser, P.E. (1968) Power from the Sun: its future, Science, 162, 857. [2] O'Leary, B. (1977) Mining the Apollo and Amor Asteroids, Science, 197, 363. [3] O'Neill, G.K. (1978) The low (profile) road to space manufacturing, Astronaut. Aeronaut. 16, 24. [4] Koelle, H.H. (1982) Preliminary analysis of a baseline system model for lunar manufacturing, Acta Astronautica, 9, 401. [5] Sparks, D.R. (1986) Recovery of asteroidal metals for terrestrial utilization, Acta Astronautica, 13, 101. [6] DuBose, P. et al. (1986) Solar power satellite built of lunar materials, Space Power, 6, 1. [7] Swanson, R.M. (1986) IEEE Spectrum, 23, 24. [8] Hutchenson, G.D. (1985) Semiconductor demand and wafer consumption, Semiconductor Inti., 8, 43. [9] Piland, R.O. (1978) The solar power satellite concept evaluation programme, in: Billman, K.W. (ed.) Radiation Energy Conservation in Space, AIAA, 61, pp. 3. [10] Glaser, P.E. (1977) The potential of satellite solar power, Proc. IEEE, 65, 1162. [11] Sparks, D.R. (1987) The large-scale manufacturing of electronic and electrical components in space, Acta Astronautica, 15, 239. [12] Adler, L, Trombka, J., Gerard, J., Blodget, H., Eller, E., Yin, L. Lamother,
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Solar Array Mechanisms for Indian Satellites, APPLE, IRS and INSAT-IITS SAMIRAN DAS & I. SELVARAJ Summary Large area rigid panel deployable and trackable solar arrays are widely used in present-day operational satellites. Three solar array mechanisms of this type for Indian spacecraft are described, one of which has already undergone flight verification. The other two mechanisms are of higher complexity and being readied for launch in the immediate future, as essential for providing India's space services in communication and remote sensing fields. Design approach, test programme and implications of test modelling towards achieving the design goal are discussed. Performance characteristics of the solar array mechanisms achieved after flight verification and qualification tests are also highlighted in the paper. 1. Introduction To meet the ever increasing power needs of spacecraft for various applications, the deployable solar array is the commonly adopted design strategy. Here the fundamental requirement of a reliable mechanism to deploy the arrays needs particular emphasis. This paper describes the solar array mechanisms of Indian Space Research Organisation's (ISRO) three satellites, one of which is already launched successfully for demonstrating space communications capability. The other two are scheduled to be launched shortly also under ISRO's communications and remote sensing programmes. The mechanisms have grown in complexity for a 2-panel system in APPLE (Ariane Passenger Payload Experiment), to a 6-panel system in IRS (Indian Remote Sensing Satellite) and then to a more complex one-sided panel system with balancing sail boom in INSAT-IITS (Indian National Satellite). This paper provides the details on how the system design has been evolved capable of meeting the individual mission requirements, the testing and qualification features, problems encountered during the qualification phase, along with flight performance of one of the mechanisms. 2. Solar Array Mechanism for APPLE The first Indian 3-axis body stabilized geostationary communication satellite, APPLE, was an experimental venture. With 670 Kg as lift-off mass and 380 Kg as dry mass the satellite's power requirements were derived from its two driven and deployable solar arrays, which folded and deployed in accordion fashion. Samiran Das, Vikram Sarabhai Space Centre, Trivandrum, India. I. Selvaraj, ISRO Satellite Centre, Bangalore, India.
The mechanism design progressed in a series of iterations with new inputs from analysis and tests being incorporated progressively, while keeping in pace with spacecraft's milestones. The studies conducted were on materials for thermal behaviour pattern, theoretical simulation of deployment dynamics, structural dynamics and correlations with experimental data. Being the first in the series of ISRO's communication satellites, many uncertainties and associated interface changes were to be absorbed in the design. 2.1 Design The array sizing was determined on the basis of generating 280 watts of power. A total array area of 2.86 m2 shared equally by two panels of 1.1 X 1.3 m each was specified. A single step deployment was adopted, by which, on initiation of a pyrotechnic device, all the clamp-down latches got released and the spring powered array and yoke deployed automatically (Fig. 1). The closed table loop control device (CCL), shown in Fig. 2, comprising of a coordinating rope with pulley sets, ensured the mechanism was a single degree of freedom system and also provided redundancy for spring torque. The pyrotechnic device was in itself a redundant system, with two opposing guillotines facing each other, and each of them powered by a separate dimple motor with an individual electrical connection. The 1 Amp—1 watt NO-FIRE and 5 Amp ALL-FIRE characteristics were the currently accepted pyrotechnic safety and reliability standards. The thermal protection system design was based on the predicted temperature extremes. Temperature compensation springs were introduced along all wire rope interconnections. The mechanism parts and the exposed surfaces of array were provided with thermal protection device in the form of thermal paint, low emissivity tape and high reflectivity tape, as per design.
2.2 Testing and Qualification The qualification was demonstrated by conducting elaborate test runs in component level, system level and finally in spacecraft level. 2.2.1 Functional Tests The deployment of the solar array was tested in a zero-g rig incorporating gravity compensation device, for measurement of critical parameters of latchup shock, final angular velocity and deployment time scatter. The test method was based on the principle that if the hinge axes were aligned vertically plane motions were aligned horizontally and zero friction was imparted at plane motion, then zero gravity simulation and measurement accuracy would be optimum. The rigid fixture in the rig (Fig. 3) which ensured all the alignments and the set of clustered air bearing pads moving on a precision granite table provided the zero-g compensation.
Air drag bearing friction and harness bending torque that influence the simulation were included in the theoretical modelling of deployment, and a correlation was obtained with test performance. 2.2.2 Thermo Vacuum Endurance Test. Some of the performance degradations expected in the environment were increased co-efficient of friction and increased harness bending torque at low temperatures, increased friction torque in vacuum, thermal distortion, figure and other effects. A full-scale model of the mechanism with spacecraft simulated side wall was used as the test article and the predicted extreme in- orbit operating conditions were duplicated, as shown in Fig. 4. 2.2.3 Vibration Test. The initial vibration test of the solar array mechanism along with a fully fledged spacecraft revealed a problem of high amplification factor, causing cell damage and mechanical interference with the nearest interfaces. The clampdown latches were then modified with spring dampers and the set-up was vibrated again. A good improvement was indicated when the transmissibility got reduced by 50% which is illustrated in Fig. 5(A) and (5B) and Table I. 2.2.5 Spacecraft Level Tests. A zero-g fixture was developed with a vertical suspension type support for conducting a partial deployment test with the solar array mounted on to the spacecraft. With its own articulate mechanism the fixture enables observation of the release of latches and first motion of array. 2.2.6 Summary of Achieved Mechanism Characteristics. Deployment time = 4.5 Sec Latchup shock = 3-5 “G” Mass —Array substrate = 3.100 Kg each panel —Cell network = 2.400 Kg each panel —Deployment mechanism = 2.675 Kg each wing —Total (2 array wings) = 16.350 Kg
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