Experimental Investigation of Interaction of Shock Heated Test Gases with 7.25 μm Carbon Fibres in a Shock Tube

Abstract

Designing thermal protection system (TPS) is a challenging task for the success of re-entry space vehicles. The TPS may be ablative or non-ablative heat shield material which is subjected to very high thermal stress at the time of re-entry. The thermal behavior of thick ablating material is strongly related to the flight environment such as the impact pressure, enthalpy of the gas and heat transfer rate from shock layer. The function of the ablative heat shield is to reduce the temperature in the shock layer. Ablation causes the TPS layer to char and sublimate through the process of pyrolysis [1]. The pyrolysis gas produced in the reaction zone of ablator interior is injected into the shock layer and is expected to reduce the net heat flux near the surface. The composite materials using ceramic matrix with high temperature carbon fibers or carbon/carbon have been used for various applications for over three decades. Carbon composite maintains their strength in inert atmospheres up to about 2500 K [2]. The mechanical properties of the composites are affected by damage induced during the oxidation tests [3]. The modern composite materials are available in carbon/carbon composites, ceramic matrix composites made with silicon carbide, silicon nitride and alumina fibers.