Characterising Internal Heat Transfer in Thermal Protection Systems

Thermal protection systems (TPS) are employed for spacecraft to survive high temperature conditions during atmospheric re-entry. For space shuttle type re-entries, the use of ceramic tiles shield the payload from exposure to these high heat fluxes. Recent research into the use of low-density materials, such as alumina foams, brings its own scientific challenges, of which understanding internal heat transfer is one. To this end, the exact 3D geometry of their complex porous structures, before and after plasma torch heating, is obtained by tomography and used in direct pore-level simulations to numerically calculate their effective heat transfer properties. Morphological characterisation is conducted via two-point correlation functions and mathematical morphology operations. Porosity and hydraulic pore diameter are seen to increase from the pre-heating (virgin) to the post-heating (charred) sample. Collision-based Monte Carlo methods are then used for radiative heat transfer characterisation. A decrease in extinction coefficient is noted between the virgin and charred samples. Both samples exhibit a large backward scattering peak for diffusely reflecting surfaces. 1 General Introduction Aerodynamic heating during hypersonic atmospheric re-entry is a chief constraint for spacecraft design and relates to the hot gas in the flowfield surrounding the vehicle. The high temperature in the boundary layer adjacent to the vehicle surface is due to internal viscous effects slowing the entering high kinetic energy hypersonic flow, dissipating and transforming it into internal energy of the gas [1]. In the case of low Earth orbit (LEO) re-entries, heat is transferred to the vehicle predominantly via thermal conduction (dependent on the temperature gradient in the gas at the wall and often called convective heating) and to a lesser extent, radiation. To survive and ensure the safety of the payload, LEO re-entry vehicles are equipped with re-usable Thermal Protection Systems (TPS) which insulate the exterior using a material with near-zero thermal conductivity. In general, re-usable materials such as reinforced carbon-carbon (RCC) are used for the most heat exposed components of the TPS. They consist of a carbon-carbon composite with a triple pre-pyrolysed resin and a silicon carbide (SiC) coating to prevent oxidation during re-entry [2]. More than fifty years since man first exited the Earth’s atmosphere and was brought back safely, TPS sizing still involves significant uncertainties. Large margins are therefore applied to its design, which increase structural and fuel weight and decrease useable payload size. Optimising TPS design is thus imperative. State of the art research studies, amongst other options, low-density TPS materials which bring new scientific challenges of which internal heat