10.4 Megawatt Class Demonstration This chapter presents a conceptual design for a 1 MW commercial precursor. The main departure from similar concepts is the LEO assembly using EVA operations rather relying on robotics. The structure retained is a prismatic one with 100 m vertices. The sun energy conversion process is based upon GaAs solar cells and the energy is beamed at 35 GHz to keep the antenna dimensions in the frame of the spacecraft. It is also shown that high orbits, GSO or near GSO are of interest for operational reasons. The mass assumption shows clearly the dominant importance of the antenna subsystem. To put the design in perspective, a set of alternative concepts are summarized: the Japanese SPS 2000 project, and two generic planar platforms in addition to the precursor concept. In addition, future trends are mentioned showing that integrated array technology could drastically change the structure of the spacecraft. A schedule for realization is provided, consistent with the rest of the project. The study is closed by a summary and a set of conclusions. 10.4.1 Constraints The following constraints are imposed on the system design in order to limit the cost. One or two Energia class launchers and one space transportation system (STS) assembly flight using manned operations are used to perform assembly in low earth orbit. Those classes of launchers have been retained as basis for launch and assembly. It does not mean necessarily that they will still be available or be the cheapest launchers in the future. The main point is to have payload and trajectory references with which to develop the design. As much as possible, proven or highly cost effective technology should be selected. The total target cost of the system is to be in the range of 1 billion US dollars, and is erected by an international crew. Orbit Choice The choice of an orbit is driven primarily by mission considerations. In this case the main mission is to beam power to ground at optimum efficiency in order to provide a practical quantity of energy for use on Earth. This means that orbits leading to a short time of visibility are not well suited as long as no efficient device is available to store large amount of energy provided in a short time. Since this kind of technology is not foreseen in the near future, it is then necessary to consider higher orbits, up to GEO, to allow for an adequate time of visibility. A brief listing of the approximate visibility windows is given in Table 10.4.1. _________________Table 10.4.1 Visibility for Various Orbits________________ Orbit Type Frequency of Pass Duration LEO 300 km 28.5 Infrequent Minutes Sun Synchronous Infrequent Minutes Equatorial 1000 km 90 minutes Minutes* Equatorial 20000 km 12 hours 6.4 hours* Equatorial 36000 km Continuous Continuous ♦Assumes rectenna can receive a signal from 30 degrees above horizon. Since the system is required to be operational for many years, and in order not to use many spacecraft (at least in early phases), it is of some interest to avoid any seasonal effect. This consideration precludes elliptical orbits for a single spacecraft, that is, the orbit should be circular. The visibility time does not favor polar orbits for which the combined rotations of spacecraft and earth result in a very short time over the same point. Furthermore, as long as the elements are launched from the ground, plane changes to the polar orbit are costly. In addition, direct launch is not feasible within the study constraints due to the manned assembly limitation.
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