Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, September 1990. / Thesis Advisor(s): Henline, William D. ; Platzer, Max F. Second Reader: Chandrasekhara, M. S. "September 1990." Description based on title screen as viewed on December 21, 2009. DTIC Identifier(s): Ceramic materials, ablative materials, heat shields, reusable equipment, space flight, thermal insulation, atmospheric entry, hypersonic flow, Mars probes, arc heaters, melting, glass, RSI (reusable surface insulaiton), aerodynamic heating, stagnation pressure, mathematical prediction, theses. Author(s) subject terms: Ablative, hypersonic, NASA, Mars, heat shield. Includes bibliographical references (p. 110). Also available in print.

A computational model is developed to study the effects of alumina layer formation on an ablative surface when exposed to high temperature particle laden gas flow. The solidification dynamics i.e., the solid and liquid alumina layer growth rate, and the heat transferred to the ablative surface are investigated. A one-dimensional model is developed taking into consideration the thermal loading, particle loading and the temperature dependence of the thermo-physical properties of alumina. A fully implicit finite volume method is used to solve the coupled set of non-linear heat conduction equations. The solidification interface is tracked using the Lagrangian interpolation technique. The particle mass flux was found to be the major factor affecting the solid layer growth rate. The gas heat flux also has a major effect on the solid growth rate and the heat transferred to the ablative surface, but only for lower particle mass fluxes. On other hand the particle temperature has a linear effect on the solidification dynamics and the heat transferred to the ablative surface for all particle mass fluxes. The heat transferred to the ablative surface is reduced by approximately 39% to 88%, depending on the mass fluxes, due to the formation of the alumina layer. / Graduation date: 2003

Thermal Protection System (TPS) is required to shield an atmospheric entry vehicle against the high surface heating environment experienced during hypersonic flight. There are significant uncertainties in the tools and models currently used for the prediction of entry aeroheating and TPS material thermal response. These uncertainties can be reduced using experimental data. Analysis of TPS ground and flight data has been traditionally performed in a direct fashion. Direct analyses center upon comparison of the computational model predictions to data. Qualitative conclusions about model validity may be drawn based on this comparison and a limited number of model parameters may be iteratively adjusted to obtain a better match between predictions and data. The goal of this thesis is to develop a more rigorous methodology for the estimation of surface heating and TPS material response using inverse estimation theory. Built on theoretical developments made in related fields, this methodology enables the estimation of uncertainties in both the aeroheating environment and material properties from experimental temperature data. Unlike direct methods, the methodology developed here is capable of estimating a large number of independent parameters simultaneously and reconstructing the time-dependent surface heating profile in an automated fashion. This methodology is applied to flight data obtained from thermocouples embedded in the Mars Pathfinder and Mars Science Laboratory entry vehicle heatshields.