FE Analysis of an ablative Thermal Protection Systems for re-entry vehicles

The TPS (Thermal Protection System) is a subsystem providing protection and isolation from aerodynamic heating. The aerodynamic heating, investigated in this paper, is related to the hypersonic speed reached by space vehicles during the atmospheric re-entry phase. Thermal protection systems can be divided into two wide categories: reusable and ablative. The reusable thermal protection systems do not experience mass and composition variation of the constituent materials during the exposure to the aero-thermic environment. Ablative TPS are characterized by surface material removal including chemical reactions, vaporization and erosion phenomena, able to guarantee the resistance to the thermal loads thanks to the phase change and the consequent ablative mass loss. The consequence of all these phenomena is a thickness reduction of the ablative material and a shape variation of the heat shield. The objective of this work is to present a numerical procedure based on ablative decomposition simulation able to compute the instantaneous material recession rate and the in-depth temperature response of the TPS. This procedure gives the capability to evaluate the amount of material required as insulation to keep the bond-line temperature within the mission requirements. To this end, in order to appreciate the material ablation on the heat shield, an advanced finite element model has been implemented. The ablative material investigated in this work is PICA (Phenolic Impregnated Carbon Ablator), made of a thermosetting resin which acts as a matrix for carbon fibers. The adopted numerical test case for the performed analyses can be considered representative of the Stardust probe.