Optimum Design of Ablative Thermal Protection Systems for Atmospheric Entry Vehicles
CHALLENGE - The Thermal Protection System (TPS) provides spacecraft entering the atmosphere, thermal insulation from the aero-thermodynamic heating. The design of such a subsystem is very critical, considering that its damage can lead to a catastrophic failure of the whole entry system, in particular if ablative materials are considered. In order to design an ablative TPS, in fact, a reliable numerical procedure, able to compute surface recession rate, pyrolysis and internal temperature histories under severe heating conditions, is necessary.
SOLUTION - The TPS needs to be sized to effectively shield the spacecraft from the high heat fluxes acting during the atmospheric entry phase. At the same time, its weight has to be minimized yet should be able to guarantee suitable protection. This article aims to describe an optimization procedure for the design of ablative heat shields by implementing a FEM based model for the thermal analysis of ablative TPSs . The numerical method is applied to the ablative TPS of the hypersonic reentry capsule Stardust. A genetic optimization procedure has been applied to the thermal protection shield in order to evaluate the minimum weight for the capsule TPS able to guarantee an effective thermal protection. This numerical model has been implemented in the optimization tool modeFRONTIER, obtaining the configuration with the lowest ablative volume, while observing the temperature conditions of the bondline.
BENEFITS - The numerical analysis consists in an FE model allowing to estimate the surface and bond-line temperatures and the final thickness of the ablative shield. Thanks to an optimization algorithm, the lowest ablative volume configuration, respecting the imposed temperature constraint, has been found. By comparing initial and best configurations, a 46% volume reduction resulted.