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Spacecraft for long duration deep space missions will need to be designed to survive micrometeoroid bombardment of their surfaces some of which may actually be punctured. To avoid loss of the entire mission the damage due to such punctures must be limited to small, localized areas. This is especially true for power system radiators, which necessarily feature large surface areas to reject heat at relatively low temperature to the space environment by thermal radiation. It may be intuitively obvious, that if a space radiator is composed of a large number of independently operating segments, such as heat pipes, a random micrometeoroid puncture will result only in the loss of the punctured segment, and not the entire radiator. Due to the redundancy achieved by independently operating segments, the wall thickness and consequently the weight of such segments can be drastically reduced.

This report discusses application of a new lightweight carbon-carbon (C-C) space radiator technology developed under the NASA Civil Space Technology Initiative (CSTI) High Capacity Power Program to a 20 kWe lunar based power system. This system comprises a nuclear (SP-100 derivative) heat source, a Closed Brayton Cycle (CBC) power conversion unit with heat rejection by means of a plane radiator. The new radiator concept is based on a C-C composite heat pipe with integrally woven fins and a thin walled metallic liner for containment of the working fluid. Using measured areal specific mass values (1.5 kg/m2) for flat plate radiators, comparative CBC power system mass and performance calculations show significant advantages if conventional heat pipes for space radiators are replaced by the new C-C heat pipe technology.

The problem of determining equilibrium temperatures for re-radiating surfaces in space vacuum was analyzed and the resulting mathematical relationships were incorporated in a code to determine space sink temperatures in the solar system. A brief treatment of planetary atmospheres is also included. Temperature values obtained with the code are in good agreement with available spacecraft telemetry and meteorological measurements for Venus and Earth. The code has been used in the design of space power system radiators for future interplanetary missions.

An efficient finite difference (FD) computational code has been developed for the analysis and design of circular sector radiators for linear alternator output “Free Piston Stirling Engine” space power systems utilizing a radioisotope Pu-238 General Purpose (GPHS) heat source. The code calls on a subroutine developed by the author to solve the second order, fourth degree, ordinary differential equation (ODE) of a fin (extended heat transfer surface) radiating to the space environment. Although the code was originally written for a rectangular coordinates system, it was transcribed into polar (cylindrical) coordinates for the present application. The circular sector radiator panel analyzed has an embedded heat pipe at an arbitrary radial location conducting cycle reject heat from the Stirling engine cold end to the radiator.

This report discusses the design implications for spacecraft radiators made possible by the successful fabrication and proof-of-concept testing of a graphite-fiber-carbon-matrix composite (i.e., carbon-carbon (CC)) heat pipe. The prototype heat pipe, or space radiator element, consists of a C-C composite shell with integrally woven fins. It has a thin-walled furnace-brazed metallic (Nb-1%Zr) liner with end caps for containment of the potassium working fluid. A short extension of this liner, at increased wall thickness beyond the C-C shell, forms the heat pipe evaporator section which is in thermal contact with the radiator fluid that needs to be cooled. During the fabrication process the C-C shell condenser section was exposed to an atomic oxygen (AO) ion source for a total AO fluence of 4x1020 atoms/cm2, thereby raising its surface emissivity for heat radiation to a value of 0.85 to 0.90 at design operating temperatures of 700 to 800 K.