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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.

The “Rigidization-on-Command”™ (ROC™) resin development has focused on the development of resin systems that use UV light cure for rigidization. Polymeric sensitizers have been incorporated into the resin formulations to promote cure using Pen-Ray lamps and UV light-emitting diodes (LED's). Formulations containing the polymeric sensitizers were examined by FTIR and DSC. Complete cure was observed after 15 min. exposure with the Pen-Ray lamps. Performance of the Pen-Ray lamps and UV LEDs was thoroughly characterized. Thermal models were developed to optimize the performance of the of the MLI insulation thermal oven used for orbital cure of the boom. Results show that -12°C is the lowest temperature required for cure of the ROC™ resin systems.

The deployment and curing of multi-layer film structures for inflatable and large Gossamer-type space structures requires accurate modeling of the temperature distributions within and among the different layers of the structure. The Sinda Temperature Translator (STT) was developed to specifically provide the application of accurate thermal load profiles on each component of thin film multi-layer space inflatable structures during the performance of multidiscipline analyses. STT provides a methodology to accurately map the temperatures obtained during transient thermal orbital analyses (finite difference methods) to structural finite element analyses (FEA) models. The derivation of the equations are for the translation of temperatures from the thermal models such as SINDA/G, which places the temperatures at the centroid of the elements, to traditional structural models, such as Cosmos/M and NASTRAN that places the temperatures at the corners of the element.

Large space-deployed antennas are of interest in the NASA, military, and commercial sectors for a variety of applications that include communications, long baseline interferometry, microspacecraft, and space-based radar. A need exists for a controlled, clean rigidization technology to harden inflatable spacecraft once they have achieved the required shape. This study addressed the space environment for typical orbits, development of UV curing cationic epoxy resin systems, mechanical properties of UV cured composites, and fabrication of demonstration tubes using the photocurable resin technology. Transient thermal analyses were run on a candidate tube configuration to determine the power required for internal UV lamps to initiate cure (15.50 W/m2) and the temperature range of the thermal processing windows.

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.

A transient thermal analysis of a Carbon/Carbon (C/C) Mars Entry Non-Ablative Aeroshell Assembly was performed to determine the maximum temperatures it would reach during a Mars entry. The purpose of this thermal analyses was to (1) determine the maximum temperatures of the 5 layers and the close-out which make up the aerothermal shield and (2) to transmit these temperatures from SINDA/G finite difference format to finite element format in COSMOS/M structures/dynamic models using Technical Alliance Group (TAG) developed SINDA/ G temperature translator software (STT).

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 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.

A thermal engineering analysis of the Mars Pathfinder MOx chemical cell was performed to determine the feasibility of raising the temperature of the soil and air cells to operating conditions using passive heating in the Mars environment. A critical 10° C rise in the soil sensor above the ground temperature is needed to activate the MOx chemical cell. Little relative spacecraft power is available to heat the instrument to its desired surface operating temperature. Realistic analytical bounds of thermal performance of the MOx chemical cell were predicted using a multi-discipline approach that consisted of materials, thermal, and structural analyses. The models accounted for solar heating, conduction to the ground, radiation to space, convection to the Mars atmosphere, and spacecraft power.