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m tThe effect of non-equilibrium flow chemistry on the surface temperature distribution over the forebody heat shield on the Aeroassisted Flight Experiment ( A m ) vehicle was investigated using a reacting boundary layer code entitled "Boundary Layer Integral Method Procedure with Kinetics" (BLIMF'K). Computations were performed by using boundary layer edge properties determined from global iterations between the boundary layer code and flow field solutions from a viscous shock layer (VSL) and a full Navier-Stokes (NS) solution. Surface temperature distribution over the AFE heat shield was calculated for two flight conditions during a nominal AFE trajectory.This study indicates that the surface temperature distribution is sensitive to the non-equilibrium chemistry in the shock layer. Heating distributions over the AFE forebody calculated using non-equilibrium edge properties in BLIMPK were similar to values calculated using the VSL program. u u M n P 9 R S 1 Nomenclature -Base diameter, m -Mach Numbcr ~ Distance along stagnation point stream line, cm -Pressure, atm -Heat transfer rate, W/cm2 -Radius, m -Surface distance from stagnation point, m ~ Time, scc *Copyright QtW1 byUlc Americanlnsututeof Acronautics and Astronautics, Inc. No copyright is asscncd in the United Slatcs under Titlc 17. U.S. Code. The U.S. Government has a royalty-free liccnsc to excrcisc all rights under thc copyright claimed hcrcin for Government purposes. All othcr rights arc rescrvcd by !he copyright owcr. tRcsearch Scicntist. tRescarch Specialist. v T -Temperature, K V, -Frce stream velocity, km/s 7 -Spccific Heat ratio,C,,/C. 7 .~ -Recombination coefficient for nitrogen 70 -Recombination coefficient for oxygen A -Bow shock wave stand-off distance, cm pm -Free stream density, gm/cm3 Subscripts n -reference to nosetip 7 -radative property o -Stagnation point w -Wall condition 00 -Free stream property IntroductionThe Aeroassist Flight Experiment (AFE) is being developed by NASA lo provide flight data necessary for the design of future Aeroassisl Space Transfer Vehicles (ASP'S). These vehicles will carry payloads between geosynchronous and low earth orbits.' An asymmeiric elliptical blunt cone has been chosen as the carrier vehicle for the flight experiment. During the aerobraking maneuver, the AFE will penetrate the upper regions of the earth's atmosphere at or near geosynchronous return velocities. Behind the bow shock wave, high temperature non-equilibrium flow is present and both oxygen and niuogen molecules are fully dissociated and some ionization occurs.The heat transfer rate to the forebody heat shield consists of radiant, sensible and chemical heating. The amount of chemical heating generated at the heat shield results from the diffusion of atomic oxygen and nitrogen to the surface, followed by subsequent recombination and energy release. The energy release is dircctly dependent on the catalytic efficiency of the surface material cov-1 Downloaded by UNIVERSITY OF MICHIGAN on July 20, 2015 | http://arc.aiaa.org | cring thc heat shic...

An aerocapture trajectory specific to the Mars 2001 mission scenario is presented with results of aerothermal and heatshield thermal protection system (TPS) mass estimates that are required to protect the spacecraft orbiter payload. Specifically, a point design of the Mars aerocapture vehicle was selected and, using an overshoot lifting trajectory, a full 3D aerothermal flowfield was simulated for selected trajectory points. The computational flowfield models were then directly used in a TPS mass sizing algorithm that estimated the overall TPS mass of the aerocapture vehicle forebody and afterbody heatshield and the thickness distributions. This analysis methodology has never been previously attempted for a 3D flowfield in a Mars aerocapture trajectory, and the results presented in this paper represent NASA-Ames' current capabilities to support the complex simulation challenges associated with aerocapture maneuvers at Mars. (Author) AbstractAn aerocapture trajectory specific to the Mars 2001 mission scenario is presented with results of aerothermal and heatshield thermal protection system (TPS) mass estimates that are required to protect the spacecraft orbiter payload. Specifically, a point design of the Mars aerocapture vehicle was selected and using an overshoot, lifting trajectory, a full 3D aerothermal flowfield was simulated for selected trajectory points. The computational flowfield models were then directly used in a TPS mass sizing algorithm that estimated the overall TPS mass of the aerocapture vehicle forebody and afterbody heatshield and the thickness distributions. This analysis methodology has never been previously attempted for a 3D flowfield in a Mars aerocapture trajectory and the results presented in this paper represent NASA Ames' current capabilities to support the complex simulation challenges associated with aerocapture maneuvers at Mars. Nomenclature L/D -lift-to-drag ratio m/C D A -ballistic coefficient, kg/m 2 -entry mass, kg -heating rate (non-ablating), W/cm--heat load (non-ablating), J/cm 2 -freestream Reynolds number, based on vehicle diameter s -distance from stagnation point, m V crd -entry velocity relative to atmosphere, km/s m q Q Re

AIAA 34th Aerospace Sciences Meeting and Exhibit, Reno, NV Jan 15-18, 1996Methodology and numerical procedures developed at NASA Ames for full-body thermal protection system sizing are documented. The core procedure is a robust implicit solver for one dimensional transient heat conduction in multilayer stacks of reusable thermal protection and structural materials. The solver includes an arbitrary number of material layers, contact resistances between materials, temperature and pressure dependent material properties, general boundary-condition options, and self-adaptive time stepping. A menu driven Graphical User Interface is used to couple the solver with an Access-to-Space material database of candidate thermal protection and structural materials and with a thermal-environment database obtained from trajectory-based full-body Navier-Stokes computations of the external flowfield, to analyze and visualize the time-dependent thermal response of the thermal protection system, and to create and modify material mapping data files. Examples of thermal protection system sizing are presented, including results for a single-stage-to-orbit rocket vehicle concept. (Author) AbstractMethodology and numerical procedures developed at NASA Ames for full-body thermal protection system sizing are documented. The core procedure is a robust implicit solver for one dimensional transient heat conduction in multilayer stacks of reusable thermal protection and structural materials. The solver includes an arbitrary number of material layers, contact resistances between materials, temperature and pressure dependent material properties, general boundary-condition options, and selfadaptive time stepping. A menu driven Graphical User Interface is used to couple the solver with an Access-toSpace material database of candidate thermal protection and structural materials and with a thermal-environment database obtained from trajectory-based full-body Navier-

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