Space Power Technological, Economic and Societal Issues in Space Systems Development Volume 8 Numbers 1/2 1989 —corfox— publi/hing compony
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 © 1989, SUNSAT Energy Council 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.
SPACE POWER Volume 8 Numbers 1/2 1989 Papers from the IAF International Conference on Space Power, Cleveland, OH, USA, 5-7 June 1989 SOLAR PHOTOVOLTAIC SYSTEMS & TECHNOLOGY 1-1. E. L. Ralph. Photovoltaic Space Power History and Perspective 3 1-2. Gary F. Turner & Stephen C. DeBrock. Large Solar Array Design 11 1-5. P. A. Iles, Y. C. M. Yeh & F. Ho. Gallium Arsenide Technologies in Photovoltaic Conversion 23 1-8. Geoffrey A. Landis, Sheila G. Bailey & Dennis J. Flood. Advances in Thin-Film Solar Cells for Lightweight Space Photovoltaic Power 31 1-9. H. Kuninaka, Y. Nozaki & K. Kuriki. High Voltage Solar Array Interacting with Ionospheric Plasma 51 NUCLEAR SPACE POWER SYSTEMS 2-7. J. R. Lance & J. W. H. Chi. Space Nuclear Power Systems for Extraterrestrial Basing 69 ENERGY STORAGE 3-2. S. Weingartner, J. Blumenberg & F. Lindner. Space Power Thermal Energy Storage: Planned Experiments for Phase Change Material Tests in Microgravity 83 SOLAR DYNAMIC POWER FOR SPACE 4-1. Thomas L. Labus, Richard R. Secunde & Ronald G. Lovely. Solar Dynamic Power for Space Station Freedom 97 NUCLEAR SPACE POWER TECHNOLOGIES 5-3. J. F. Wett, J. W. H. Chi & J. M. Livingston. NERVA-Derivative Reactor Technology—A National Asset for Diverse Space Power Applications 115 5-4. William B. Harper Jr., Anthony Pietsch & William G. Beggenstoss. The Future of Closed Brayton Cycle Space Power Systems 125 5-7. Jack G. Slaby. Free-piston Stirling Technology for Space Power 137 POWER CONSERVATION, CONTROL & CONDITIONING 6-5. Praveen Jain, J. Bottrill & Af. Tanju. Considerations of Power Conversion Techniques in Future Space Applications 149 ADVANCED SOLAR SPACE POWER SYSTEMS 7-4. Jerry Grey & Lucien Deschamps. Central-Station Electric Power for Spacecraft 179 ADVANCED NUCLEAR SPACE POWER SYSTEMS CONCEPTS 8-4. AL L. Ramalingam & AL Morgan. Optimization of Lanthanum Hexaboride Electrodes for Maximum Thermionic Power Generation 199
8-5. Elliot B. Kennel, Mark S. Perry & Brian D. Donovan. Reliability and Single Point Failure Design Considerations in Thermionic Space Nuclear Power Systems 219 8-7. Nils J. Diaz, Samim Anghaie, Edward T. Dugan & Isaac Maya. Ultrahigh Temperature Vapor Reactor and Magneto Conversion for Multi-megawatt Space Power Generation 225 8-8. J. Norwood Jr., J. Nering, B. C. Maglich & C. Powell. A High Specific Power Aneutronic Space Reactor 237
1-1. Photovoltaic Space Power History and Perspective E. L. RALPH Summary Photovoltaic space power systems have been the preferred technology for most spacecraft missions. This is due to a combination offactors. Of primary importance is the high reliability afforded by the large redundant matrix of solid state photovoltaic devices that make up a solar array. Of course, the low cost of a solar power system relative to the complete spacecraft is also an important factor. Progress over the past 30 years has continually improved the conversion efficiences, decreased system weights, and reduced costs, so that alternate technologies always have a moving target to surpass if they are to be competitive. This paper describes the progress that has been made and some advances that are likely to be achieved in the future. Background and History Since its invention in 1954, the solar cell has played an important part in the development of space power systems. The first applications for solar cells were in toys and portable radios, but in 1958 space exploration changed everything. Solar power systems using photovoltaic devices and associated battery storage provided an excellent source for spacecraft power needs. These systems utilized a solid state device which was inherently reliable, readily manufacturable, easily qualifiable, scaleable to the desirable power levels, and relatively inexpensive. These attributes made photovoltaic power systems the work horse of the spacecraft industry. Research and development efforts over the past 30 years have been aimed primarily at increasing solar cell conversion efficiency, increasing battery specific power, decreasing array weight, decreasing costs and improving radiation resistance. In addition, much work has been accomplished in improving reliability, increasing the power levels, and developing innovative solutions to specific mission requirements. Solar cell efficiency has consistently increased over the years as shown in Fig. 1. The first solar cells used on the Vanguard Satellite in 1958 were silicon cells of only about 5% air mass zero (AMO) efficiency and were P on N (N doped base) type devices which were very sensitive to high energy particle radiation. After the Van Allen radiation belt was discovered by the Explorer 1 satellite a more radiation resistant solar cell was desired. Switching to a N on P (P doped base) type cell design provided significant improvements in radiation resistance, so most satellites after 1962 utilized silicon devices of this type. About this same time gridded front contacts were E. L. Ralph, Hughes Aircraft Company, PO Box 92919, Los Angeles, CA 90009, USA. Paper number IAF-ICOSP89-1-1.
developed, which substantially reduced the distributed series resistance and increased efficiencies to 12% AMO. Over the following years (see Table I), a series of developments improved silicon cell AMO efficiencies incrementally to the present level of about 15%. Base resistivities were increased to 10 ohm cm to improve radiation resistance; junction depth was decreased to increase short wavelength response, thus increasing efficiency and also increasing radiation resistance; multilayer antireflection coatings and textured surfaces were developed to increase efficiency; back surface fields were added to reduce the rear surface minority carrier recombination rate thus increasing efficiency (unfortunately this gain was lost after extensive radiation exposure for 200 //m thick cells); cells were made much thinner (~62 /zm thick) which decreased weight and gained back the increased efficiency provided by the back surface field even after typical orbital radiation exposure levels; larger area cells were designed and manufactured using mechanized processes that reduced costs substantially; and back surface reflective contacts were developed that rejected the unusable long wavelength light, thus increasing efficiency by reducing cell operating temperatures in orbit. Gallium arsenide solar cells have also been developed over the years but have only recently been manufactured in the USA for space use. Single junction GaAs devices typically have AMO efficiencies of about 19% while dual junction GaAs on germanium devices have shown efficiencies as high as 22%.
Similarly battery developments have decreased the weights and increased capacity thus making solar power systems more competitive (see Table II). In the early 1960s NiCd cells had only 5 Ahr capacity and produced about 4.5 Whr/Kg (2 Whr/lb) while today NiCds are 50 Ahr and about 22 Whr/Kg (10 Whr/lb). Similarly the number of discharge cycles has been increased substantially so that missions could be increased from 3 years to 15 years today. Solar cell array structure development efforts have also progressed dramatically over the years. Early solar arrays were made from aluminum honeycomb panels and had specific power outputs of about 11 W/Kg (5 W/lb). More recently, lighter weight honeycomb panels covered with graphite/Kapton or Kevlar facesheets have been used to make large area cylindrical or flat plate arrays of about 22 W/Kg (10 W/lb). In 1971, the Hughes FRUSA flexible roll up array was placed into orbit demonstrating that lightweight solar cell arrays of about 66 W/Kg (30 W/lb) could be stored in a small volume. More recently, in 1984 Lockheed demonstrated the SAFE lightweight flexible array on a shuttle flight. This array utilized a foldable Kapton film substrate that was extended and held in position using a extendable truss beam. This design also demonstrated 66 W/Kg (30 W/lb) technology and the feasibility of stowing large area solar arrays in a small volume. This paved the way for missions with very high power levels, to be provided by solar array modules of this type. The present US space station power system design relies heavily on flexible array technology. Design Criteria A block diagram of a typical solar power system along with the key design factors influencing performance is shown in Fig. 2. Battery storage is typically needed to provide power during eclipse and when peak loads are encountered. Sizing of the solar
array is determined by balancing the average load (Kw hr) requirements against the array generation capacity while considering the efficiency factors of the battery, power electronics and distribution system. Battery sizing (Kw hr) is determined by the peak power (Kw) requirements and the maximum depth of discharge desired. Although high performance and reliability are key factors desired in a solar power system, cost is still the deciding factor in most cases. There are exceptions for some missions where cost may not be a major consideration but this is usually related to cases where there is some other restraining design factor such as a limited array area or a weight restriction. In these cases, the higher cost of an advanced technology, such as a higher efficiency GaAs solar cell, can be justified if it allows power growth in existing spacecraft designs. In these cases the cost of the increased performance is offset by the cost saving of using an existing design versus redesign and requalification. Another way that one can justify the higher cost of an advanced technology is to provide weight savings or reduced drag area which can be balanced against launch costs or fuel expenditures. These types of trade-offs however, are often difficult to quantify and require multidisciplinary analysis, so a solar power system design will usually be selected based on performance and cost. Solar Array Technology Silicon solar cells in the range of 14 to 16% AMO efficiency are typically used in most missions today. However, there are many opportunities for improved solar cell performance or weight reduction using new materials or cell designs (see Table III). Conventional single junction GaAs solar cells of 19% AMO efficiency are now available in production quantities, but they are about two to three times the cost and about twice the weight of silicon solar cells. This limits their use to only a few applications where the improved output performance justifies their use.
In order to minimize these cost and weight penalties, GaAs cells are being developed that are deposited on large area lower cost germanium substrates. These substrates can be made very thin to reduce the weight penalty. This type of cell design can be made in a dual tandem junction structure thus resulting in about 22% AMO efficiencies. Pilot production facilities are being installed now so these cells and the associated performance gains should be available within the next few years. Another advanced solar cell type that is being studied extensively is the indium phosphide (InP) cell. This material has the advantage that radiation degradation can be minimized because the crystal damage sites can be annealed out at room temperatures or at slightly elevated temperatures. Also, the theoretical efficiencies are high (similar to GaAs) with laboratory cells demonstrating about 19% AMO efficiency at this time. Material costs are very high and in limiting supply at this time so cost reduction developments, similar to those being worked on for GaAs cells, must be achieved before this cell type will receive extensive application. Further increases in solar cell efficiency are still possible by utilizing multiple band-gap devices. For instance, research is being conducted to develop a triple junction multi band-gap cell utilizing an AlGaAs, GaAs, InGaAs structure which theoretically has an AMO efficiency of 30%. As in the case above, there will need to be a great deal of cost reduction development done before this type cell becomes a competitive alternative. An alternative approach toward developing an advanced solar cell with specific mission benefits is to utilize an ultra thin film design. Amorphous silicon (ctSi) and copper indium diselenide (CuInSe2) solar cells are examples that have been extensively researched and have been manufactured in quantity for terrestrial markets. These devices have the advantages of being very low cost and can be extremely lightweight since the active device need be only a few microns thick. A major disadvantage today is that large area production type cells are only about 5% AMO efficiency. Laboratory devices can be made in the 10% efficiency range and multiple junction designs theoretically could reach greater than 15% efficiency. There are many unanswered concerns about the radiation resistance and space-worthiness of these devices, but the ultra lightweight and simple compact storage capability provide potential solutions to many space mission needs of the future. The projected power-to- weight ratios for a 10% efficiency thin film cell blanket is expected to reach about 350 watts per kilogram. Conceptually values as high as 5000 watts per kilogram are projected. A summary of the effect of some of the above solar cell advances on the complete solar array properties can be seen in Table IV. Today's typical rigid panel using thick silicon solar cells is used as a baseline for comparison of weight and area parameters. Advanced cell designs combined with advanced lightweight fold-up wing designs (typical of the space station design), provide a comparison of the advances in array technology expected in the near future.
Electrochemical Storage Technology Major advances have been made in improving the number of cycles and the depth of discharge of space qualified rechargeable batteries over the years. The specific weights have been substantially decreased as well. As is the case with solar cell technology, this trend is expected to continue (see Table V). Nickel Cadmium (NiCd) cells have been the most commonly used spacecraft batteries. Today, typical NiCd cells are supplied with specific weights of 17.5 Whr/Kg (8 Whr/lb) for GEO and about 8.8 Whr/Kg (4 Whr/lb) for LEO at an 80% depth of discharge for GEO missions and about 40% depth of discharge for LEO missions. New developments of the Super NiCd cells provide increased cycling capability so that 15-year missions can be accommodated. Within the past 15 years, nickel hydrogen (NiH2) cells have been developed and many new spacecraft batteries utilize this cell design. The major advantage of this cell is that the specific weights are about 33 Whr/Kg (15 Whr/lb) GEO and about 22 Whr/Kg (10 Whr/lb) LEO and the depth of discharge can be increased to at least 90%. A more advanced cell being developed today is the sodium-sulfur cell. This cell should be able to provide over 153 Whr/Kg (70 Whr/lb) initially and eventually
above 220 Whr/Kg (100 Whr/lb) for missions lasting up to 15 years in earth orbit. This type of battery will make large solar power systems in the 50 Kw range feasible, and indeed attractive in the future. Summary From the previous discussion it can be seen that improvements in solar cell efficiency, solar array specific power, and battery specific energy have been significant and are expected to continue for a long time. Figure 3 summarizes some of these expectations for several solar power system parameters. The present state of the art is represented by technologies that have been utilized in flight. Specific power values of solar cell arrays today, as represented by the FRUSA and SAFE arrays, are at 66 W/Kg. This can be increased to 120 W/Kg using the SEP array design with 30% multi band-gap solar cells, or even to 350 W/Kg using thin film solar cell technology. Specific energy values for storage subsystems are at 30 Whr/Kg for NiH2 cells today, and this can be expected to go to 150 Whr/Kg with NaS technology and 400 Whr/Kg with regenerative fuel cells. Consequently, the complete photovoltaic power system specific power of about 7 W/Kg today for rigid silicon solar cell panels and NiCd batteries, can now be increased to 22 W/Kg using silicon cells on flexible panels along with NiH2 batteries, and in the near future this can be increased to 50 W/Kg using 30% efficiency multi band-gap solar cells on a flexible array design with NaS batteries. These advances will make it possible for photovoltaic system power levels to increase from the present day 1 to 10 KWe levels to 25 KWe level anticipated for
several near-term missions and even to the 100 KWe levels planned for the space station. With steady progress in PV power system technology, there is good reason to believe that almost any spacecraft mission being planned can be effectively and economically accomplished using a PV system. Early conceptual design studies of the space solar power station (SSPS) showed that even a 15 GWe PV design was feasible. In conclusion, it appears that the anticipated improvements in PV technology will provide a viable space power system option for most missions being planned and can provide much larger power levels than considered previously.
1-2. Large Solar Array Design GARY F. TURNER & STEPHEN C. DEBROCK Summary and Introduction In the early 1970s, in conjunction with the initiation of the first Space Station solar array technology studies, a set of basic design decisions were made on the best generic approach to the design of solar arrays in sizes from tens to hundreds of kilowatts. Considerations were maximum potential for low weight, low launch package volume, adequate deployed structural stiffness and damping, at least a ten-year temperature cycle life, and adaptability to low-cost automated assembly. In the intervening 18 years between then and now, a number of programs contributed to the development of this basic design concept and the associated manufacturing processes. These developments form the foundation for the current Space Station Freedom Solar Array and platform designs and for two operational satellite programs. These current designs are unchanged in basic concept from the original studies in 1972, but they now reflect the mature technology which has been developed and proven out since then. To manufacture and test these systems, an integrated manufacturing, assembly and test facility has been constructed and is in operation with many automated features, including cell welding and substrate manufacture. Basic Design Concept for Large Area Solar Arrays In 1970, Johnson Space Center sponsored the Large Space Station Technology Evaluation Contract with Lockheed Missiles and Space Company. The program was conducted in two phases: (1) a trade study of all relevant structural and electrical design approaches which might be employed in a large 150 kw solar array design; and (2) the design and construction of a ground demonstration unit employing the selected design approaches. Figure 1 shows the ground deployment test unit during extension tests. Gary F. Turner, Manager, Aerospace Products, Lockheed Missiles & Space Co., Inc., Sunnyvale, CA 94089, USA. Stephen C. DeBrock, Manager, Power Systems Work Pkg 4, Space Station Program, Lockheed Missiles & Space Co., Inc., Sunnyvale, CA 94089, USA. Paper number IAF-ICOSP89-1-2.
The basic design decisions made are summarized in Table I and typical applications are illustrated in Fig. 2. For the structure, we selected a flat-fold packaging rather than the drum-type ‘window shade' approach which several investigators had developed in smaller sizes. Analyses showed that in these large sizes, with these orbital loads, the tubular-type deployment booms were completely inadequate, and we therefore selected a truss-type boom. The Astromast was selected from among the truss booms studied because it was completely retractable and capable of sustaining load while partially deployed—a feature which many of the candidates did not have. The selection of this particular combination of packaging and deployment rather than the ‘drum' type allowed us to decouple the mechanics of the packaging and deployment hardware, resulting in a much less complex system. The selection of solar module design approach represented an even more radical departure from conventional designs in that it depended on a wraparound solar cell. Although a few were made under this initial contract, none had yet been developed with acceptable performance, nor had the associated welding processes required to perform automated panel assembly been developed or demonstrated. Some work had been done on copper printed circuit substrates, but this had not been applied to welded solar array assemblies. Most important, the ability of the proposed assembly techniques to stand up to the packaged launch environment and the tens of thousands of thermal cycles experienced in low earth orbit had not been demonstrated. Nonetheless, the concept had promise and, more important, the commitment of the associated NASA and contractor organizations to follow through on development and demonstration of the key technologies.
Concept Evolution and Hardware Development The technology currently being employed in the Space Station Freedom solar array was the result of a large number of contributing efforts from three NASA Centers and the in-house efforts of Applied Solar, Spectrolab, Able Engineering, Dupont, and Lockheed. The contributing contract efforts and the major development milestones are shown in Fig. 3, starting with the Space Station Solar Array Technology Development previously discussed. After this effort was concluded with a successful ground demonstration, the Marshall Space Flight Center initiated a long-term program which was to culminate 12 years later in the first flight demonstration of this technology in August of 1984 on Space Shuttle. Early emphasis in the MSFC projects was on the printed circuit substrate panel design and fabrication, which had received only minor effort on the previous contract. The main thrust of these programs was the development of welding techniques as well as associated inspection and repair methodology. These programs also included the building and temperature cycle testing of sample panel assemblies. They established the foundation for the process development which culminated in 1982 in a proven weld process with long-time temperature cycle life being demonstrated by test both at
the Marshall Center and at Lockheed. Similar samples were also employed at Lewis Research Center for plasma testing. In 1975, the first system application for this technology, the Solar Electric Propulsion (SEPS) program appeared probable enough to warrant the planning of a full-scale development and ground demonstration program, and the Marshall Center undertook such a program, with the original goal being ground test only. The ground
testing was carried out, and the decision was then made to rework the hardware for a Space Shuttle flight test, and this project was initiated. In parallel with the SEPS and the flight experiment, two other important applications emerged—the 25 Kw Power Module (Marshall Space Flight Center) and the Shuttle Power Extension Package (PEP) initiated by Johnson Space Center. Solar array system designs for all three were undertaken under contract to both JSC and MSFC. Each of these system configurations is depicted in Fig. 4, along with an advanced geostationary platform which is still being studied. These design efforts, directed towards four completely different vehicle flight configurations, demonstrated the versatility of the basic solar array design approach adopted years before. Deployed, the systems resembled those shown in Fig. 4. Packaged for launch, they were configured as shown in Fig. 5. With respect to solar array blanket design, mast design (except size), and blanket packaging design, these systems were identical, differing only in the way in which the blanket packages are deployed after launch—a relatively minor mechanism problem. In addition to the design efforts, these programs also sponsored a series of low ‘g' deploy/retract experiments on the KC-135 which validated the fold-up and deployment concepts being proposed, and provided a baseline of data for future designs. The PEP program also led to another important development—that of large area (5.9 X 5.9 cm) wraparound solar cells for use with these systems. This was a three- center cooperative program with JSC and LeRC sponsoring and directing the cell development and MSFC acting as integrator to insert the cells into the Solar Array Flight Experiment. This program proved for the first time the feasibility of manufacturing large area solar cells in production quantities. Lockheed constructed test panels with these cells, and the Marshall Center incorporated these panels into the SAFE program. The cells performed as predicted in the space environment after surviving launch and multiple deployments on STS 41-D in August of 1984 (Fig. 6).
Objectives of the solar array orbital flight test were to demonstrate the readiness of large-area, lightweight photovoltaic technology, to demonstrate the deployment, retraction, and restowage of the array; and to study dynamic behaviour of large, flexible space structure. In parallel with the various hardware development activities there was, of course, extensive effort expended in supporting analytical tools—chiefly deployed dynamics and thermal models. These models were also validated by the SAFE experiment, resulting in an established set of flight test-validated design tools for structures of this type, such as space station. Flight data correlated extremely well with the outputs of both models (Fig. 7). In addition to providing technical data, the SAFE also served as a demonstration which provided confidence to other Lockheed programs to employ these types of systems, and indeed two programs are now in production. These programs continue to provide further technology input to the Space Station Freedom Program. In addition, the Industrial Space Facility, a commercial space lab venture, adopted this design as their baseline and funded design efforts, atomic oxygen testing of cell assemblies, and an important flat-pack bypass diode development. Also, as the problem of Kapton erosion by atomic oxygen became apparent from JSC shuttle flight data, Lewis Research Center began working on an SiO2 coating method to combat the problem. This led to the Photovoltaic Array Environmental Protection (PAEP) program which is directed at evaluating various coatings for use on
space station and other programs. The SiO2 approach is currently baselined for both the Space Station Freedom and the ISF programs. It can be seen from this discussion that the technology we now are applying for space station was the result not of a single effort, but of a multi-NASA center, multicontractor effort over a period of many years—all predicated on a set of technology development goals set in the early 1970s and pursued diligently since that time. Current Space Station Freedom Design The space station solar array consists of eight solar array wings which support a 75 Kw bus with a 187.2 Kw power output at the four-year design point. The solar array wing, shown in Fig. 8, is composed of two split blankets with a central mast—similar in concept to the Power Module array wing previously discussed. Each blanket is approximately 4.7 X 32.6 metres—virtually identical to the SAFE flight unit which was 4.3 X 32 metres. The wing is stowed for ease of Shuttle packaging as shown in Fig. 8. This is also an approach with a previous design heritage from the MSFC Power Module Solar Array studies. The array is extended by a 0.8 metre diameter coilable longeron mast, giving the assembly adequate stiffness to provide a 0.1 Hz minimum deployed natural frequency—considerably higher than the 0.04 Hz frequency demonstrated on SAFE.
The electrical design uses a large area 8 X 8 cm wrapthrough contact silicon solar cell (Fig. 2). The cell is 0.2 mm (8 mils) thick, has a gridded back-surface contact for added thermal performance, and is covered with a 0.125 mm (5 mils) thick ceria stabilized microsheet cover. As is the case with all Lockheed silicon solar cell array designs, parallel gap welding is used to bond the cell to a copper printed circuit interconnect system. The Kapton film is coated with SiO2 for AO protection prior to starting the fabrication of the flexible substrate. A series string of 400 cells provides a nominal 160 V de output. Figure 9 shows two development space station panels (one circuit) completely assembled. The polar and co-orbiting platform solar arrays are shorter than space station (25 and 6.1 metres), but employ the same electrical panels with masts similar in design to space station, but smaller in diameter. Extensive use of common hardware is planned. The development test program concentrates on aspects of the design which are different than SAFE. This includes the larger mast, atomic oxygen protection (which is also being addressed under the separate LeRC PAEP contract), and the larger wrapthrough solar cell. Given the strong design and test heritage of this system, there is high confidence that this system will prove a reliable, cost-effective power source for space station. Solar Array Manufacturing With the advent of two using programs and the prospects for others such as Space Station, CDSF, and other future government programs, Lockheed has invested in a modern, highly automated integrated manufacturing, assembly, and test facility for large area solar arrays. This 40,000 square foot facility (Fig. 10) has the capability of
processing large area solar arrays from incoming raw materials and parts through final system deployment test. Two key elements of the manufacturing process, the roll laminator which assembles the Kapton, copper and photo resist preparatory to photo etching the printed circuit and the automated weld station, are shown in Figs 11 and 12. The weld station operates automatically with an X/Y table used to position the welds and a closed loop control system which terminates each weld at a preset weld temperature. Weld parameters are stored on tape for future reference should later inspections reveal anomalies. The operators load the panels for welding and observe the weld operation through a closed circuit TV, overriding the operation if they observe any abnormal occurrences such as misalignment of the weld head. This facility, now in full production on two programs, provides us with the capability to produce in a reasonable manufacturing cycle the over 1300 solar panels required for the space station, plus the requirements of the related platforms in a reliable, repeatable process. Conclusion The series of developments discussed herein, the result of long-term cooperation between three NASA centers and a number of contractors, has resulted in a solid technology base for the Space Station Freedom solar array as well as for numerous other present and future large solar array applications. Although many individual
projects and contracts contributed, perhaps the most important contribution of all was the initial NASA technology study which laid out a well thought-out set of technology goals and plans to form a foundation for the work that followed. ACKNOWLEDGEMENT Acknowledgement should be given to the many people at NASA JSC, LeRC, and MSEC as well as the employees at Lockheed, ASEC, Spectrolab, Able and ASTRO who contributed so much to these cooperative efforts. More specifically, special acknowledgement is deserved by Jim Cioni of Johnson Space Center and Jim Miller of Marshall Space Flight Center for having the vision to risk failure and explore these new approaches and the tenacity to see them through.
1-5. Gallium Arsenide Technologies in Photovoltaic Conversion P. A. ILES, Y. C. M. YEH & F. HO Summary This paper describes how GaAs technology has been applied to manufacture high efficiency solar cells for use on spacecraft. Improved substrates, layer growth and cell fabrication processes have been combined to demonstrate that GaAs cells can provide some attractive options for spacecraft power designers and can be supplied at the required production levels. To penetrate this highly conservative market has required steady development for about ten years, and this development work is demonstrating promise for wider use of GaAs-based cells. Introduction For over 30 years, Silicon (Si) solar cells have provided electrical power for most spacecraft. Cell performance has increased steadily, and the cells have met increasing demands on output and reliability. GaAs cells have advantages over Si cells, including higher efficiency (around 25% more) and increased resistance to damage from particle impact. Unlike Si cells, there is no need to trade off high efficiency for high radiation resistance. In addition, GaAs cells have less fall-off in power when operating at high temperatures. GaAs solar cell arrays have been used on some Soviet satellites since the 1960s. Also in the 1960s RCA workers successfully made GaAs cells, but the difficulties in controlling GaAs crystal growth and cell processing did not warrant the additional costs required for manufacture. Milestones in Technology Development Advances in the 1970s In this period, work at IBM, Rockwell and Carnegie-Mellon University led to significant improvements in the performance and understanding of GaAs cells. In the late 1970s, US Air Force (USAF) contract work at Hughes Research Laboratories using liquid phase epitaxy to grow the GaAs layers demonstrated that efficiencies up to 19% could be obtained, that scale-up was possible, and that GaAs cells could meet most of the array requirements. P. A. Iles, Y. C. M. Yeh and F. Ho, Applied Solar Energy Corporation, City of Industry, CA 91749, USA. Paper number IAF-ICOSP89-1-5.
Air Force MANTECH Contract (‘High Efficiency GaAs Cells') In the period 1982 to 1986, under an Air Force Manufacturing Technology contract, Applied Solar Energy Corporation (ASEC) demonstrated producibility by delivering 5500 cells with average efficiency of 17% (AMO). These cells met the basic space qualification tests. This program showed that production rates of 1000 2 X 2 cm cells per week could be maintained. This work also demonstrated that high throughput, metal organic chemical vapor deposition (MOCVD) reactors could be used to deposit the active GaAs (and AlGaAs) layers on GaAs substrates. The program also developed the necessary infrastructure, mainly by providing several sources of suitable GaAs substrates and CVD sources. The GaAs substrate development was interesting, leading to steady supplies of substrates tailored specifically for space cells. The substrates were grown with low dislocation density in near-rectangular shaped boats (typically 2.5 X 4.5 cm2 or 4.5 X 4.5 cm2 in size). The shape reduced wastage and led to substrates with costs per unit area less than half the costs at the time the program started. Parallel Production Efforts on GaAs Cells Before the MANTECH Program was completed (around 1984) but triggered by the results achieved to that time, an active space-cell program was begun at ASEC. This production program for 2x4 cm2 cells accelerated the build-up of the production efforts and by imposing specific user-needs, extended the range of the qualification tests. The array methods were extended to handle and bond cells in array assembly, and the GaAs cells were modified to withstand extended operation under reverse bias conditions, a possibility if parts of the array were shadowed. The scaled-up MOCVD effort also required increased attention to safety, because large quantities of toxic gases (mainly arsine, AsH3) were used. Extensive monitoring equipment and personnel protection procedures were installed. In this period, Government safety regulations were also intensified. Around 1986, Mitsubishi produced 50,000 2x2 cm2 GaAs/GaAs cells for use on a Japanese communication satellite. Increased Efficiency GaAs Cells In the early 1980s, groups at Varian, Lincoln Laboratories, Spire and Mitsubishi Corporation increased GaAs cell efficiency over 20% (up to 21.5% AMO). Most of these groups used MOCVD to deposit the GaAs layers. At this stage, GaAs cells 10-12 mils thick could be successfully processed under production conditions and interconnected into arrays. It was realized that to be competitive in future arrays, the weight of the GaAs cells must be reduced. Tests with thinner GaAs substrates were only moderately successful, and the easy cleavage of GaAs indicated that severe breakage problems could be expected if large area, thin cells were needed. Air Force Program (‘Rugged GaAs Solar Cells') In 1984, the Air Force recognized the need for lightweight GaAs cells and awarded a contract to evaluate GaAs cells grown on germanium (Ge) substrates. The atomic lattice spacing and thermal expansion coefficient of Ge are closely matched to those of
GaAs. These factors are favourable for growing good quality epitaxial layers of GaAs on Ge. Also Ge is stronger than GaAs and the higher strength can be exploited to form larger area, thinner cells. In the long term, Ge substrates will cost less than GaAs substrates, especially for larger area slices. Previous work at Jet Propulsion Laboratory, Lincoln Laboratories and Spire Corporation had shown that good quality GaAs cells could be epitaxially grown on Ge substrates. The AF contract extended this previous work and led to GaAs/Ge cells, up to 4x4 cm2 in area and 3-4 mils thick, with efficiency at about 17%. The layer growth and processing of those cells were similar to the methods used for GaAs/GaAs cells. Interim Re-evaluation of GaAs/Ge Cells As the GaAs/Ge cell work progressed, the MOCVD growth conditions which appeared to give the best quality GaAs layers on Ge substrates led to cells which had higher Voc than GaAs cells, and usually lower curve fill factors (CFF). The enhanced Voc was generated at the GaAs/Ge interface, and operated in cascade (series) with the GaAs PN junction. At the time, it was believed that the additional voltage (up to 200 mV) would lead to efficiencies greater than those obtained for GaAs/GaAs cells, expecially if the CFF values could be increased. However, further tests of these cells consisting of a GaAs cell grown on active-Ge substrates showed some possible problems. The temperature coefficient of Voc for the cascade cells was higher, up to twice the values typical of GaAs cells, and this offset a major advantage of GaAs cells. During measurements of the temperature coefficients, an older two-light solar simulator was used. The spectral output for this simulator was closely matched to the true AMO spectrum, and the enhanced Voc cells showed ‘kinked' I-V curves (Fig. 1). The kinks were attributed to insufficient current generation and collection at the GaAs/Ge interface, the generation resulting from absorption of the near infrared region of the AMO spectrum which is transmitted through the GaAs layers (i.e. wavelengths above 0.9 /zm). If the test spectrum was richer in nearinfrared wavelengths than the true AMO spectrum, no kinks were observed. In Fig. 1, the Hoffman simulator was close to AMO, the XT-10 simulator was red-rich. When the cascaded cells operated under the true AMO spectrum, near maximum power conditions, because the current generated near the GaAs/Ge interface was lower than the current collected at the GaAs PN junction, the GaAs/Ge interface was driven into reverse bias operation. The kink resulted from superposition of the reverse I-V characteristic of the interface and the forward I-V characteristic of the GaAs PN junction. The tests with different simulators were confirmed by testing at high altitudes by NASA-Lewis aeroplane flights, and it was concluded that the high efficiencies obtained for the active-Ge structures were mainly caused by red-rich test conditions. The red- rich simulators had good spectral match for the wavelengths used to test silicon or GaAs cells, which only extend to 1.2 /zm or 0.9 /zm respectively. On revising theoretical estimates of the maximum current available for collection at the GaAs/Ge interface (after passing through the GaAs layers), it was concluded that even for ideal conditions, current matching for the interface and the PN junction would be difficult to obtain under the true AMO spectrum. Tests were made to shift the anti-reflective coating reflectance minima, to reduce the reflectance of the near infrared wavelengths, but they did not lead to current
matching unless the GaAs PN junction current was significantly decreased. This option is self-defeating, because the efficiency of the cascade cell derives mostly from the GaAs cell. Parallel work on cascade cells consisting of an optimized Ge PN junction cell, connected by a tunnel diode to a top cell had led to selection of an AlGaAs top cell where the Al content was 8-10%, to ensure current matching for both cells. Even if current matching was possible, growth on Ge with no intentional junction would require control of surface states and/or impurity inter-diffusion at the GaAs/Ge interface during the growth of the GaAs layers. Also, retention of the active-Ge structure would require adjustment in users' simulators, and in cell testing correlation. For these reasons, ASEC decided to pursue a passive-Ge structure, where the additional photovoltage at the GaAs/Ge interface is suppressed. There are possible methods to achieve passivation after layer growth (involving buried, shorting contacts or deep ion implantation) but these would make processing more difficult and could introduce other problems. Tests were made to change the MOCVD growth conditions especially for the early GaAs layers formed near the Ge surface. In effect, the ‘optimized' procedures developed in the previous work were reversed. At the time, it was not known if it would still be possible to grow high quality GaAs layers under these modified conditions.
After several months' work, successful results were obtained. The growth conditions were changed to lower temperatures, and slower growth rates, along with some changes in substrate preparation. The quality of the GaAs layers has improved steadily and it is now possible to grow high efficiency GaAs cells on the passive-Ge substrate. These passive-Ge cells do not have excess Voc, although the Voc values and other I-V parameters are typical of the values obtained for high efficiency GaAs/GaAs cells. Also, these cells have the same low temperature coefficient as GaAs/GaAs cells. Despite the presence of a low density of growth defects not observed for growth on GaAs substrates, the GaAs/Ge cell performance has equalled that of co-grown and coprocessed GaAs/GaAs cells. The best cells have exceeded 20% (Table I). These same tests also showed that similar efficiencies were obtained for all four 2x2 cm2 cells grown on the 4.5X4.5 cm2 substrates used (Table II). This indicates that with an optimized grid design (a minor modification), the same high efficiencies can be projected for 4 X 4 cm2 GaAs/Ge cells. Air Force MANTECH Program (‘Rugged, Thin GaAs Solar Cells') In 1988, the Air Force awarded a MANTECH program to develop and produce high efficiency GaAs cells on thin Ge substrates. The goals were 18-20% efficiency, areas 4x4 cm2 to 6x6 cm2, and thickness 3-4 mils. The program also called for demonstration of panel technology for these cells, using either welding or soldering. Although the program was planned when active-Ge substrate cells were considered promising, when ASEC received the contract, it was decided to use the passive-Ge substrate approach. The good results reported in the previous section are encouraging for the MANTECH Program. Production of GaAs/Ge Cells The production effort outlined above was satisfactorily completed, the total output for the deliverables being over 35 Kw, with average cell efficiences of 16.5-17%. For follow-on production runs, GaAs/Ge cells were promising, and cells with the passive-Ge structure are undergoing qualification testing, with production to follow validation that all qualification requirements are met.
Current Trends Although the production run involved ‘frozen' procedures, tests showed that it was possible to achieve efficiencies comparable to those reported above for cells grown in the large-scale production reactors and processed with standard production methods. Figure 2 shows the I-V curve for a 2 X 2 cm2 GaAs/GaAs cell with efficiency around 21.5% (AMO). The results shown in Tables I and II and in Fig. 2 suggest that with some further fine-tuning, the GaAs cells on passive-Ge substrates can reach even higher efficiencies than achieved to date. For production runs, either of GaAs/GaAs or GaAs/Ge cells, increased median efficiency values are projected, first to around 18.5%, with chance of further increase in future runs. When these higher efficiencies are obtained for larger area, thinner cells with the operating advantages of GaAs/GaAs cells, it is apparent that GaAs/Ge can offer advantages to designers of advanced arrays. Either more power can be obtained from the same area or for the same power, the arrays can have 20-30% lower area. Although Ge and GaAs are both more than twice as dense as Si, at the array level, 3-4 mil thick GaAs/Ge cells can be shown to have advantages in watts per kilogram and watts per square metre. Even though the cost of GaAs cells is higher than the cost of Si cells, the cost impact is reduced when overall system costs are assessed. For these reasons, GaAs/Ge cells are now competitive in many applications where
Si cells have dominated. As production experience with GaAs-based cell grows, further cost reduction can be projected. Also, more complex cell designs, especially with wraparound or wrapthrough contacts will become available. The same processing has also been applied to make GaAs cells with very high efficiency under high concentrations (over 24% at 100-times concentrations of AMO sunlight, and over 27% at 500 X terrestrial sunlight). Several spacecraft array designs use concentrators operating between 20 X and 100 X, and plans to use these arrays are based on use of GaAs cells. There is also parallel development of optimized cascade cells, some based on GaAs/Ge technology, and these are being evaluated for higher power output in both flat-plate and concentrator arrays for future spacecraft applications. Finally, high efficiency (18.5% AMO) has been obtained for GaAs cells grown on Si substrates, and these cells have advantages in reduced weight. Present tests are evaluating whether these efficiencies can be obtained over large areas, and also whether the cells can withstand the temperature cycling experienced in most space applications. If GaAs/Si cells prove to be effective, they will add considerable choice in GaAs- related technology for the array designers. Conclusions This paper has traced the evolution of GaAs cell designs for space use, and has shown how the cell performance and production experience have been steadily increased to meet an increasing number of array demands. This evolution has taken about ten years, and this appears to be a typical cycle for the application of advanced technology to provide spacecraft power. The engineering community responsible for maintaining and extending technology for space power is considered to be conservative. Much of this conservatism stems from a good record of success in designing and manufacturing arrays which have usually exceeded their predicted period of operation. Array designers believe that detailed attention to many reliability criteria during cell manufacture and assembly into arrays is an essential feature ensuring the good record. This period, during which possible advantages in applying GaAs technology to space power were assessed and incorporated into the design options available to array designers for advanced missions, has provided valuable experience to both cell and array producers. Further advances can be expected.
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1-8. Advances in Thin-film Solar Cells for Lightweight Space Photovoltaic Power GEOFFREY A. LANDIS, SHEILA G. BAILEY & DENNIS J. FLOOD Summary The present status and current research directions of photovoltaic arrays as primary power systems for space are reviewed. There have recently been great advances in the technology of thin-film solar cells for terrestrial applications. In a thin-film solar cell the thickness of the active element is only a few microns; transfer of this technology to space arrays could result in ultra low-weight solar arrays with potentially large gains in specific power. Recent advances in thin-film solar cells are reviewed, including polycrystalline copper-indium selenide (CuInSe^ and related I-III-VI(2) compounds, poly crystalline cadmium telluride and related II-VI compounds, and amorphous silicon:hydrogen and alloys. The best experimental efficiency on thin-film solar cells to date is 12% AMO for CuInSe2. This efficiency is likely to be increased in the next few years. The radiation tolerance of thin-film materials is far greater than that of single crystal materials. CuInSe2 shows no degradation when exposed to 1-MeV electrons. Experimental evidence also suggests that most or all of the radiation damage on thin-film materials can be removed by a low temperature anneal. The possibility of all thin-film cascade multibandgap solar cells is discussed, including the trade-offs between monolithic and mechanically stacked cells and voltage-matched versus current-matched configurations, and remaining problems to be solved. The best current efficiency for a cascade cell is 12.5% AMO for an amorphous silicon on CuInSe2 multibandgap combination. Higher efficiencies are expected in the future. For several missions, including solar-electric propulsion, a manned Mars mission, and lunar exploration and manufacturing, thin- film photovoltaic arrays may be a mission-enabling technology. 1. Introduction In this paper we discuss the development of photovoltaic arrays beyond the next generation, particularly looking at the potentials of thin-film polycrystalline and amorphous cells. We discuss two important figures of merit, efficiency (i.e. what fraction of the incident solar energy is converted to electricity), and specific power (power to weight ratio). Geoffrey A. Landis, Sheila G. Bailey and Dennis J. Flood, NASA Lewis Research Center, MS 302-1, Cleveland, OH 44142, USA. Paper number IAF-ICOSP89-1-8.
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