• Solar array is assumed to have an efficiency of 0.15 (GaAs technology), • Power conversion efficiency (vacuum tubes and HT power) is assumed to be 0.8 • Antenna efficiency (including phase shifter) is assumed to be 0.9 • Frequency selected is 35 GHz (see transmitter discussion section) • Atmospheric path coefficient is assumed to be 0.8 ( reference, value at 35 GHz for dry air is 0.95) • Total efficiency is 0.0864. With this value it is easy to determine the area The solar array area is: A= 106 / (1300 x 2 x 0.0864) = 8903 m2. • We have selected a basic structure of 100 m x 100 m per face in order to have some margins. PV Material Selection/Suitability Much progress has been made in the past decade regarding the development of efficient photovoltaic arrays which are applicable for use in the space environment. Of these, InP (Indium Phosphide) appears attractive because there is little degradation of an already high efficiency in the space environment. However, the development of 1 hectare or more of these cells might be costly at this point since the manufacturing process is difficult. Another choice includes the thin film a-Si (amorphous silicon) arrays, which can be produced in large quantity. The disadvantages seem to include a relatively low efficiency which has been found to degrade further in a space environment. Whereas a long term commercial project may choose the more expensive option of InP, the choice for a constrained demonstration project might be the thin film a-Si technology or GaAs if economically viable, with the potential of future upgrades if other mission factors (that is, the final orbit selection) permit. It should be noted that the selected orbits are extremely clean. As previously stated, for the 20309 km orbit, the electrons are the main concern even if ions can not be totally neglected. Power Subsystem As the solar arrays are fixed (that is, not mechanically movable with respect to the spacecraft structure), the amount of solar energy collected varies with the relative orientation of the spacecraft toward the sun. The energy goes from zero to a first relative maximum, then reaches a second maximum and repeats symmetrically to the end by reaching zero when the arrays are masked by the antenna as shown in Figures 10.4.6 and 10.4.7. Figure 10.4.6 Angular Convention To simplify the system design, it is chosen to cease the power transmission when the level drops below 80% of the maximum (when the two arrays are symmetrically facing the sun). The fraction of energy lost due to this choice is 21% of the total collected amount. This fraction does not justify increasing significantly the mass of the spacecraft by a battery or a fly wheel so as to store energy on board. Furthermore, this storage would occur during a phase when the power collected varies strongly, thus complicating the power subsystem. With the figures of efficiency used above, the
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