The solar collector panels are supported by nonconducting as well as conducting structures which carry the power to the microwave generators via the central masts. Dielectric materials are used for the continuous support structure which is transparent to the microwave beam. Rotary joints are provided at the perimeter of the central mast to allow rotation of the microwave transmitting antenna. These joints are the only major continuously active components in an otherwise passive satellite. The solar cells will actually be based on the "roll out" blanket design of current technology but will incorporate improved fabrication techniques, a substantial reduction in cell thickness, and the use of solar concentrators to reduce weight from the currently attainable 30 Ib/kW (14 kg/kW) to about 3 Ib/kW (1.4 kg/kW). Solar concentrators with Kapton film mirrors coated to reflect solar radiation onto the solar cells and to filter undesirable portions of the solar spectrum are designed to reduce the area requirements for the solar cells and their weight and cost. Exhibit 3 indicates the arrangement of the solar cell arrays and concentrating reflectors. A concentration factor (n) of 2 will reduce the efficiency of an 18% silicon solar cell to about 14% at the operating temperature when heat rejecting coatings are used for the solar cell array. The exposure of the arrays to the space environment is projected to result in a logarithmic degradation of silicon solar cells, with a 6% loss of the original efficiency after the first five years. Micrometeroids are projected to impact 1% of the solar cells during a 30-year operational lifetime. A reaction control system based on the use of ion engines (Argon is the candidate propellant) will be required to keep the satellite in the appropriate orbit and to assure that the solar collector panels point towards the sun to within one degree, while the microwave antenna is directed towards the receiving antenna on earth. To achieve the desired stationkeeping and attitude control, about 100,000 lb (45,000 kg) of propellant will be required per year, depending upon specific orbital characteristics .
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