be made between the two competing technologies in terms of mass and cost. In addition, the use of the Energia launch vehicle imposes a number of requirements which are as follows: • Payload mass = 15 tons into polar orbit • Fairing diameter = 6.1 m (dynamic envelope), 6.7 m (static envelope) • Fairing height = 42 m • The spacecraft must be designed to fit within these constraints. It has been attempted to use the same platform design as the baseline option whenever this does not comprise the performance of this option. For example the SDS satellite needs a large radiator to reject waste heat and the photovoltaic system does not. This approach allows a reduction in the overall cost and in the schedule risk and possibly would allow for the integration of the SDS as the power source for the baseline option. Spacecraft Configuration Options The spacecraft design is determined by the selected Solar Dynamic System and in particular the large solar concentrators required to collect the necessary heat energy. An area of approximately 280 m^ is required for the solar concentrators to generate 120 kW to the satellite. An important consideration is that the satellite should be symmetrical. Otherwise the spacecraft would have to compensate for external torque's that are created by solar radiation. Solar Concentrators Concepts Since clearly the concentrator is too large to be launched in its deployed state, some method of stowing the concentrator is required. Unfurl able mesh antennas as used on the NASA's TDRSS satellite cannot be used to collect solar radiation since they work at a different wavelength. Inflatable space rigidified antennas are a possible candidate as concentrators, using an aluminized coating to reflect the solar power to the receiver. However, it will be very difficult to achieve the very high precision surfaces which are required to obtain the high concentration ratio necessary. Therefore inflatable space rigidified antennas have been rejected as a solar concentrator for this design. Another possibility is to have folded mirrors which unfold in orbit but these require a mechanism to unfold the mirror which may increase the risk. Therefore, for the solar concentrators, a petal fold out design is proposed using overlapping mirrors which are deployed in orbit. Small stub separators are used to keep the mirrors separate in the stowed configuration thereby protecting the mirrors. Possible Spacecraft Configurations A number of schemes are possible to mount the concentrators onto the platform. One scheme is to mount two reflectors, each with a diameter of -13 m to the platform on opposite walls of the spacecraft, i.e., side-mounted, each supplying one receiver. Another is to mount a single concentrator on top of the platform. This scheme would use an Cassegrain type configuration with a primary and secondary mirror to focus solar heat back onto the receivers. Since only one concentrator is used the diameter of the dish is larger at -19 m. Schematics of both designs are shown in Figure 10.3.10. Figure 10.3.10 Possible Deployed Satellite Configurations
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