Thermal protection analysis of Mars-earth return vehicles

Abstract

An analysis is performed to determine the aerothermodynamic heating and thermal protection materials requirements of a typical fast manned Mars mission. Entry conditions represent those consistent with an early manned mission having a total trip time of approximately 14-16 months. A coupled flowfield-radiation stagnation region heat transfer viscous shock layer (VSL) calculation was performed to ascertain the various shock layer heating mechanisms and the resultant surface heat fluxes for a generic shaped Mars/Earth return capsule. Computations were performed using the VSL computer code, RASLE. The ablation and thermal response of four different candidate ablating heat shield materials was determined using results from the RASLE code in conjunction with the detailed (one-dimensional) charring ablator-thermal response code CMA. Based on these results, the Earth return capsule thermal protection system (TPS) mass fraction was estimated for each material. A general conclusion is that attractive TPS mass fractions of 10-15% are possible with these candidate systems but that a vehicle and mission dependent choice must be made between heat shield total surface recession and initial TPS mass fraction.