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The thermal design of unmanned satellites and manned spacecraft require the knowledge of heat conduction and radiation of complex geometrical shapes. These complex shapes are usually made up of the more common geometries such as flat rectangular plates, flat polygon plates, triangular plates, cones, disks, parabolas, spheres, cylinders and rectangular boxes known as the nine primitive geometries. The heat transfer conductances have been derived for all the above geometries including circumferential, longitudinal and radial conductances for the non-flat plate type geometries. This paper will present the derivation of the equations for circumferential, longitudinal and radial heat transfer conductance for a right circular thin shell cone or a segment of the cone. A thin shell cone is one in which the radius to thickness ratio is greater than 10.

The conductive heat transfer across two contacting surfaces is difficult to predict because the surfaces will always have imperfections such as roughness, voids and non-flatness. These imperfections occur on flat plates, honeycomb panels or any mating surfaces that transfers heat across their interfaces. These imperfections are hard to characterize experimentally and will reduce the conductive heat transfer between the surfaces and thereby degrade the thermal performance. Quantitative understanding of the phenomenon of surface contact improves the ability to accurately predict temperatures across thermal boundaries. This is important for optimizing the thermal performance of components such as radiators and electronic subassemblies of spacecraft and instruments where size and weight are a premium. The purpose of this paper is to show what parameters, such as contact pressure, surface roughness, flatness, and bolt spacing, are involved in predicting the conductive contact coefficient.

Candidate Aeroshell Test models composed of a quasi-isotropic Carbon/Carbon(C/C) front face sheet (F/S), eggcrate core, C/C back F/S, Carbon Aerogel insulation, C/C radiation shield and the C/C close-out were constructed based on the analytical temperature predictions presented in Part One of this work[1]. The analytical results obtained for a simulated Mars entry of a 2.9 meter diameter cone shaped Carbon-Carbon Aeroshell demonstrated the feasibility of the design. These results showed that the maximum temperature the front F/S reached during the decent was 1752 °C with the resulting rear temperature reaching 326 °C in the thermal model. Part Two of this work documents the thermal modeling and correlation for the Mars Aeroshell test sample and fixture. A finite difference, SINDA/G, thermal math model of the test fixture and sample was generated and correlated to data from an arc jet test conducted at the NASA Ames Research Center's interactive heating facility.

This paper presents the derivation of the equations for circumferential, longitudinal and radial heat transfer conductance for a right circular thin shell parabola or a segment of the parabola. A thin shell parabola is one in which the radius to thickness ratio is greater than 10. The equations for the surface area of a parabola or of a parabolic segment will also be derived along with the equation to determine the location of the Centroid. The surface area is needed to determine the radial conductance in the parabola or parabolic segment and the Centroid is needed to determine the heat transfer center of the parabola or parabolic segment for circumferential and longitudinal conductance. These equations can be used to obtain more accurate results for conductive heat transfer in parabola which is a curved spacecraft components.

The overall goal of this program was the development of a 10 m. diameter, self-deployable antenna based on an open-celled rigid polyurethane foam system. Advantages of such a system relative to current inflatable or self-deploying systems include high volumetric efficiency of packing, high restoring force, low (or no) outgassing, low thermal conductivity, high dynamic damping, mechanical isotropy, infinite shelf life, and easy fabrication with methods amenable to construction of large structures (i.e., spraying). As part of a NASA Phase II SBIR, Adherent Technologies and its research partners, Temeku Technologies, and NASA JPL/Caltech, conducted activities in foam formulation, interdisciplinary analysis, and RF testing to assess the viability of using open cell polyurethane foams for self-deploying antenna applications.