Authors: Marie-Claude Druguet
Addresses: Laboratoire IUSTI/UMR CNRS 7343, Aix-Marseille Université, 5, rue Enrico Fermi – Technopole de Château-Gombert, 13453 Marseille Cedex 13, France
Abstract: This paper presents a numerical study of the effects of thermal boundary conditions on the shock layer flow over a spacecraft entering a planetary atmosphere at hyper-velocity, and on the resulting convective heat fluxes to its heat shield surface. This study is motivated by the fact that spacecraft heat shields are covered with materials that ablate to reduce the spacecraft wall temperature. The ablation is not taken into account, but the effects of changes in the temperature of the vehicle surface are thoroughly studied. The results show that the flow field solutions within the shock layer and the convective heat fluxes are sensitive to the wall thermal boundary conditions. A computation of the flow field solution for an adiabatic wall is also performed. In this case, the gas recovery temperature along the wall increases as the gas moves away from the stagnation point. Computation convergence is also analysed.
Keywords: aerodynamics; hypersonic flow; atmospheric entry; hypersonic shock layer; boundary conditions; thermal boundary layer; convective heat flux; blunt bodies; spacecraft entry; heat shield surface; ablation; spacecraft heat shields; wall temperature; space vehicles.
International Journal of Aerodynamics, 2012 Vol.2 No.2/3/4, pp.316 - 338
Published online: 31 Oct 2014 *Full-text access for editors Access for subscribers Purchase this article Comment on this article