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
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