of an interconnect surrounded by a set of dielectric surfaces. We show that the drag on such a system may be quite different from the system studied in Ref. [1] and discuss a simple physical model to explain the results. (Paper number IAF-ICOSP89-1-10.) REFERENCES [1] Kuninaka, H. & Kuriki, K. (1987) Numerical analysis of the interaction of a high-voltage solar array with ionospheric plasma, Journal of Spacecraft and Rockets, 24, pp. 512-517. 2. NUCLEAR SPACE POWER SYSTEMS 2-1. Design Considerations for 10 to 1000 kW(e) Nuclear Reactor Power Systems for Space Applications* Louis O. Cropp', Donald R. Gallup', Albert C. Marshall', David T. Furgal2 & Barbara I. McKissock2 ‘Sandia National Laboratories, PO Box 5800, Albuquerque, NM 87185, USA; 2G2 vanTel, Ltd., PO Box 5639, Albuquerque, NM 87185, USA; 'NASA Lewis Research Center, 21000 Brookpark Rd., Cleveland, OH 44135, USA. Masses and radiator areas of typical space nuclear power concepts are estimated as a function of the continuous electrical power required during a ten-year mission. Results are presented as a function of power level in the range of 10 to 1000 kW electrical. Three general reactor types will be discussed: (1) the radiatively cooled Star-C reactor technology with thermionic conversion external to the core; (2) the liquid metal cooled Topaz technology with pin-type thermionic fuel element conversion in the core; and (3) the liquid metal cooled SP100 reactor technology with thermoelectric, Stirling, and Rankine conversion systems. Mass estimates include all satellite subsystems except the payload itself. Area estimates include radiators to dump the waste heat from the reactor, power conversion, power conditioning, power transmission, and reactor control subsystems but not the payload. All system components utilize near-term technology with the exception of the liquid metal Rankine turbine. This conversion technology is further out in time because it requires two-phase flow in a microgravity environment and this raises issues that are not likely to be resolved in the near term. Scaling of the total satellite system's mass and launch and operational envelopes with power level is an extremely important issue since mission power levels are not well known 5-40 years in the future. Any proposed US reactor will require at least five years for development, and the costs will be significant. This means the technology selected will be expected to satisfy US operational needs for many years, perhaps as many as 35. As a result, space reactors must have the flexibility to provide either increased or decreased power in successive designs without excessive increases in specific mass (i.e. mass per unit of electrical power output) or ‘volume envelopes'. Since the various concepts differ greatly in their ability to provide such flexibility and no single concept is perfect for all power levels, realistic appraisals of future power requirements are essential to choosing the correct reactor technology. However, these appraisals must also include requirements for safety, reliability, the ability to survive natural or wartime threat environments in space, etc. In this regard, it is extremely important to recognize that military and civilian requirements are different. How much different depends largely on where the reactor is expected to operate (e.g. earth orbit, lunar surface, Mars exploration, etc.) and the severity of the hostile threat chosen by the military. These system-specific issues are considered, first by investigating the differences and similarities between military and civilian uses and delineating desirable system characteristics for each use. This leads naturally to a discussion of the technologies required for
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