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The lunar environment places some unique demands on a thermal management system designed for manned lunar missions. A principal concern is that for many prime base locations the effective thermal sink temperature is often near or above nominal room temperature (25°C). This is due to the fact that a conventional radiator must look at either, or both, the sun and the hot lunar surface. Direct rejection of waste heat at such temperatures is thus impossible, and some alternative approach is needed to enable a sustained mission. This paper presents three such alternative systems: a heat pump assisted central thermal bus; an innovative, selective field-of-view radiator; and use of the lunar regolith as a heat sink. All of these concepts appear feasible, but each has uncertainties associated with its practicality and weight estimate. The heat pump assisted thermal bus appears to be the most viable concept and is discussed herein in some detail.

With the use of single and two-phase heat transport systems for the thermal control of large space facilities, there will be numerous instances where fluid lines will have to traverse joints that either rotate or move in some other manner. Flexible hoses are being considered as one means of traversing these joints. They would be subjected to a variety of stresses and to as many as 150,000 rotational cycles. In order to test the resilience of flexible hoses to bending stress, a test assembly was constructed to determine the number of flexing cycles the hoses could withstand before losing their ability to maintain a constant pressure. Corrugated metal hoses of 1/4, 3/8, and 1/2-inch diameter and teflon hoses of 3/8 (smooth bore) and 1/2-inch (convoluted) diameter were tested at different pressures with nitrogen gas. The metal hoses had lives ranging from 30,000 to 100,000 flexing cycles.

Thermal conductances are computed for relatively simple node geometries that are typical of analytical models of spacecraft. Heat-transfer rates computed with these conductances are compared to heat-transfer rates computed from exact, closed-form mathematical formulas. The comparisons show that most accurate results are obtained with rectangular nodal arrangements and with triangular arrangements with interior angles less than ninety degrees. For rectangles, the conductances obtained from finite-difference, finite-element and centroid methods are identical. For triangles, the finite-element conductances are best. For other quadrilateral arrangements and triangles, the centroid method is the most reliable.

To provide spacecraft thermal design engineers with the data needed to define the earth albedo and emission, we have analyzed data from the Earth Radiation Balance Experiment (ERBE) over a period of 36 months, covering orbital inclinations of 57 and 99 degrees. Design values of the albedo and earth emission can be obtained from this data directly; however, we present an earth-map (zonal) method for computing design values for any orbital parameters. In addition, the zonal method permits less pessimistic assumptions to be made in a proof-of-design simulation.

A design study of pumped two-phase and capillary-pumped thermal control systems (TCS) was conducted on a typical, advanced earth-orbiting spacecraft (AEOS). NASA's Upper Atmosphere Research Satellite (UARS) size and baseline design were chosen as the configuration for AEOS; however, the power requirements were increased and the allowable temperature range was decreased to represent the requirements of a more advanced spacecraft. For peak-power dissipations of 1 or 2 kW, the capillary-pumped system was lightest. At these power dissipations, the line sizes were the smallest available: therefore, there was no net advantage to having a pump. The 91kg weight saving over the 173kg conventional system was due primarily to reduced heater power. However, more convenient component locations also reduced the structural weight significantly.