time more data this rate is still in the range of any network as a MIL STD 1553B already used in space. Phase computation could be done in two ways: using the global attitude of the total array, (through an interferometric measurement of a reference signal beamed from the ground), or using an individual measurement at each element level. The latter solution allows a better shaping of the beam, and takes into account the effect of the atmospheric path on each individual element. However, it will also need some additional hardware and a heavier computational load. There is not enough information available to select one of these two solutions. It is probable that an intermediate approach could be successful, relying on interferometric measurement to determine the attitude and measuring at ground level the shape of the beam, then sending back to the spacecraft a correction signal. Rectenna Considerations Strictly speaking, the rectenna is not a part of the spacecraft design. Nevertheless it is impossible to ignore the constraints induced on (and by) this part of the system. Using a high altitude orbit and a relatively low size transmitting antenna presents the main problem of the energy density at ground level. The rectifying elements, based on diodes, exhibits an energy level threshold under which their efficiency decreases quickly. Using high orbits and transmitting continuously does not allow to provide an energy density above the threshold level. To overcome this point, one can emit in bursts to increase the instantaneous value or use concentrators. Using bursts seems very attractive but requires a specific electronics on board, very similar to the power stages of the emitting part of a radar. This kind of technology is not yet mature at this power level for space applications and could be a possible threat in GEO due to EMI induced perturbations. Moreover it would need an extra mass to be flown leading to an increase in cost for launch and development. The concentrators, usually of parabolic shape, should track the spacecraft. For kilometers wide rectenna this solution seems barely practical due to the number or size of these concentrators. The concentrators used should thus be mechanically static. This requirement is easy to fulfill for the GEO orbit. For a different orbit it is generally not feasible. However, for an equatorial orbit, a set of parabolic cylinders with their axis parallel to the ground track of the spacecraft and featuring rectifying elements on their focus plane could be used. Due to the short wavelength, a metallic lattice would act as a good reflector allowing to keep the cost of the system to be lower. Thermal Control During Mission Phases The basic assumption is to use the structure of the antenna itself to dissipate the excessive heat. With the solid angle of view of the earth being low compare to that of the black space, it can be considered for a first estimation that the antenna radiates as a black body at 4 K. The total amount of heat radiated is given by the black body equation weighted by the emissivity of the radiators. Q = RsE(Tr4-Ts^\ where Q is the amount of energy, R the area of the radiators, s the Planck constant, E the emissivity, Tr the absolute temperature of the radiator, and Ts the absolute temperature of the surrounding space. Considering that the solar arrays are self radiating, the energy lost in the chain from the solar arrays to the rectenna is at the maximum 13 MW x 0.15 x (1-0.0864/. 15) = 8.27 kW. Assuming a temperature of the radiator of 373 K, which is a low value when considering tube temperature, and a space temperature of 4 K, the area needed is well below the total surface of the antenna. When the antenna is facing the sun, the solar arrays are no longer powered, drastically reducing the cooling requirement. The other elements of the spacecraft not in contact with this face could radiate directly to space through the triangular open surfaces in the north and south directions. Propulsion Subsystem Ulis section discusses the propulsion needs of the MW class demonstrator. From these requirements, an outline of the optimal system is devised, accounting for all stages of the mission. The two main mission phases are the construction phase in LEO and the operational phase in 36000 km orbit (worst case scenario compare to 20309 km). Of these, the latter must be the design driver, that is, the phase of the mission which the propulsion subsystem design is biased towards. The tasks of the propulsion system for the mission phases identified above are: LEO construction phase: attitude control and drag compensation
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