Space Vehicle Heat Shield Having Edgewise Strips of Ablative Material

A heat shield for a space vehicle comprises a plurality of phenolic impregnated carbon ablator (PICA) blocks secured to a surface of the space vehicle and arranged in a pattern with gaps therebetween. The heat shield further comprises a plurality of PICA strips disposed in the gaps between the PICA blocks. The PICA strips are mounted edgewise, such that the structural orientation of the PICA strips is substantially perpendicular to the structural orientation of the PICA blocks

Ablative Thermal Protection System (TPS) Margin Study

Predicting the reliability of the heat shield for crewed space vehicles has been a topic of continuing interest within NASA for many years. The design of a thermal protection system (TPS) is subject to numerous large sources of uncertainty and reliability assessments of such TPS are rare. The proposed talk discusses both the application of Design Of Experiments (DOE) to developing a new arc jet testing campaign for a given TPS and the reliability assessment conducted for the same TPS for a crewed space capsule (similar to Apollo capsules) to withstand re-entry to earth from space. The objectives of the study were to 1) provide recommendations for a planned arc jet testing campaign, 2) determine the design reliability of a proposed TPS, and 3) conduct a sensitivity analysis to determine the effect of input parameters and user choices on the TPS thermal design reliability. The development of a recommended arc jet test matrix employed a combination of techniques based on the analysis of variance (ANOVA) statistical methodology. The techniques employed include DOE, response surface (RS) methodology and uncertainty quantification (UQ). The recommended test matrix consisted of 30 test cases and includes four replicated condition pairs. Randomization was used to establish the test order, the testing facility, and the test sample cut pattern from three lots of material. The resulting arc jet test matrix was a compromise between one derived from statistical DOE techniques and the existing capabilities of arc jet test facilities located at NASA JSC and ARC. Statistical metrics were employed to objectively compare the assessment-derived matrix to an existing testing proposal. Five of the six metrics examined favored the new proposed test plan over the existing proposed test plan; one of the metrics (and, perhaps, the most important) strongly favored the new test plan proposal over the existing test plan. The reliability assessment investigated the sensitivity of reliability estimates to various input parameters, which included multiple studies to examine the total bond line temperature reliability based on 7 body point locations for 2 proposed trajectories. Each body point and trajectory combination was subjected to 5 different combinations of trajectory and aerothermal environment assumptions. The reliability was assessed based on a composite material failure criterion, which associated a greater probability of system failure proportional with the exceedance of an assumed safe bond line temperature limit. The study also investigated the sensitivity of reliability predictions to various input and problem formulation parameters. A large, statistically significant difference was found in the estimated TPS reliability when considering various formulations of the reliability problem, including the use of different failure conditions. The proposed talk will summarize the work performed in these areas and highlight some of the findings and recommendations that emerged from the work

Orbiter Return-To-Flight Entry Aeroheating

The Columbia accident on February 1, 2003 began an unprecedented level of effort within the hypersonic aerothermodynamic community to support the Space Shuttle Program. During the approximately six month time frame of the primary Columbia Accident Investigation Board activity, many technical disciplines were involved in a concerted effort to reconstruct the last moments of the Columbia and her crew, and understand the critical events that led to that loss. Significant contributions to the CAIB activity were made by the hypersonic aerothermodynamic community(REF CAIB) in understanding the re-entry environments that led to the propagation of an ascent foam induced wing leading edge damage to a subsequent breech of the wing spar of Columbia, and the subsequent breakup of the vehicle. A core of the NASA hypersonic aerothermodynamics team that was involved in the CAIB investigation has been combined with the United Space Alliance and Boeing Orbiter engineering team in order to position the Space Shuttle Program with a process to perform in-flight Thermal Protection System damage assessments. This damage assessment process is now part of the baselined plan for Shuttle support, and is a direct out-growth of the Columbia accident and NASAs response. Multiple re-entry aeroheating tools are involved in this damage assessment process, many of which have been developed during the Return To Flight activity. In addition, because these aeroheating tools are part of an overall damage assessment process that also involves the thermal and stress analyses community, in addition to a much broader mission support team, an integrated process for performing the damage assessment activities has been developed by the Space Shuttle Program and the Orbiter engineering community. Several subsets of activity in the Orbiter aeroheating communities support to the Return To Flight effort have been described in previous publications (CFD?, Cavity Heating? Any BLT? Grid Generation?). This work will provide a description of the integrated process utilized to perform Orbiter tile damage assessment, and in particular will seek to provide a description of the integrated aeroheating tools utilized to perform these assessments. Individual aeroheating tools will be described which provide the nominal re-entry heating environment characterization for the Orbiter, the heating environments for tile damage, heating effects due to exposed Thermal Protection System substrates, the application of Computational Fluid Dynamics for the description of tile cavity heating, and boundary layer transition prediction. This paper is meant to provide an overall view of the integrated aeroheating assessment process for tile damage assessment as one of a sequence of papers on the development of the boundary layer transition prediction capability in support of Space Shuttle Return To Flight efforts

3D Printing Heat Shields

Heat shields that protect spacecraft from the heat of entering a planet's atmosphere are labor intensive to manufacture, cost prohibitive and the heat shield design is constrained by the manufacturing process. Little, if any, automation is used to manufacture and install the thermal protection system (TPS) on the vehicles and most heat shields are segmented with many seams and gaps. The goal of this project is to change this paradigm by developing new TPS materials and taking advantage of additive manufacturing technology to automate the processes for making and installing monolithic, ablative TPS on large entry vehicles

Arcjet Tests of Different Gap-Filler Options for the Orion PICA Heatshield

PICA (Phenolic Infiltrated Carbon Ablator) is one of the candidate thermal protection materials for the Orion vehicle. Because PICA is fabricated in blocks, gaps exist between the blocks, similar to the individual ceramic tiles of the Shuttle thermal protection system. The results of this work focus on arcjet test results of different gap-filler options for PICA, performed as part of the Orion TPS Advanced Development Project. The arcjet tests were performed at NASA Ames Research Center on stagnation models 4 inches in diameter at conditions representative of Orion flight conditions for both Lunar and Low Earth Orbit return. Performance of gap-filler options was evaluated based on the extent of backface temperature change, as compared to PICA without gaps, and on the extent of flow penetration into the gap, evident from the gap opening and widening