The application of transpiration cooling as a thermal protection system for hypersonic vehicles

This thesis investigates the application of transpiration cooling to reduce surface heat transfer on hypersonic vehicles. Such vehicles experience high peak heat fluxes and time-integrated heat loads throughout the ascent, re-entry, and cruise phases. The most significant heat fluxes occur at the stagnation point of sharp leading edges, regions of laminar-turbulent transition, or shock-wave boundary layer interactions. Whilst previous and current hypersonic vehicles employ ablation for many of these regions, such thermal protection systems are neither reusable nor suitable for all scenarios. In this work, the performance of transpiration cooling is assessed for a generic flight vehicle. Subsequently, the application of transpiration cooling to laminar and turbulent flows and shock-wave boundary layer interaction regions is assessed experimentally.
For the first time, the additional benefit of transpiration cooling to the underlying substructure is evaluated numerically. The low fidelity PIRATE code for evaluating transpiration-cooled systems is extended to account for quasi-two-dimensional lateral heat conduction effects, enabling very fast calculations of the two-dimensional transient temperature response of a transpiration-cooled thermal protection system. To solve for the transpiration-cooled outer wall and a two-dimensional solid substructure, PIRATE has been coupled with the commercial finite element package COMSOL, enabling the modelling of the longer-duration thermal effects of the integrated heat load over a flight trajectory. Transpiration cooling using helium coolant has been applied to a wing leading edge with an aluminium substructure. Carbon–carbon ceramic composite and the Ultra-High-Temperature Ceramic (UHTC) zirconium diboride (ZrB2) were chosen as candidate materials. The substructure temperature history for the Space Shuttle re-entry trajectory was obtained, showing that transpiration cooling can lead to a 35% reduction in peak substructure temperature and a 65% decrease in thermal gradients.
A novel heat transfer measurement method is developed to enable measurements of surface heat transfer in transpiration-cooled porous injectors in short-duration hypersonic facilities. Traditional surface heat flux methods such as discrete thin film gauges or thermocouples are not suitable for modern micro-porous materials with pore sizes of the order of 10 microns. High-speed infrared thermography was employed to measure the surface temperature of the porous material. A bespoke calibration system was developed, and the method was verified experimentally in the Oxford High-Density Tunnel on a flat-faced hemispherical probe, with a porous Alumina injector. Nitrogen, air, argon, krypton, and helium injection gases were used with mass flow rates ranging from 0.01-0.235 kgs-1m-2. It was found that the Stanton number reduction matched to within 10% of both CFD results and correlations. The direct measurements provided by this technique enable the acquisition of spatially resolved heat transfer on a porous surface with mass injection.
After the development of the heat transfer measurement technique, this was applied to two different scenarios. The first concerns the need for detailed local heat transfer information on and in the vicinity of the porous injector to design a transpiration-cooled system. In this work, experiments were conducted in the Oxford High-Density Tunnel at Mach 6.1 in both laminar and turbulent regimes. Spatially resolved 2D surface heat transfer measurements were acquired by imaging directly on and downstream of two micro-porous transpiration-cooled injectors (METAPOR CE170 and Zirconia) using high-speed infrared thermography. It was found that a modification to existing relations from film theory successfully correlates the stream-wise heat transfer distribution on the injector for different blowing rates of nitrogen and helium injection at both laminar and turbulent regimes. A key result is that helium performs much better than reported in previous experiments. Finally, the downstream thermal effectiveness is characterised for turbulent flows and an analytical correlation is proposed.
In the final component of this thesis, transpiration cooling is applied to mitigate the peak heat fluxes caused by shock-wave boundary layer interactions. Experiments were again conducted in the Oxford High-Density Tunnel at Mach 6.1 in both laminar and turbulent regimes where a 10 degree shock generator created a strong oblique shock wave which impinged onto a transpiration-cooled micro-porous injector. For the laminar condition, due to the strength of the incident shock, a laminar-transitional shock-wave-boundary-layer interaction region was formed with peak heating over 50 times greater than the nominal laminar level. Both nitrogen and helium were used as coolants and relatively low injection rates were sufficient to greatly reduce the heat transfer downstream of shock interaction for both regimes. The experimental data are correlated, and both fully turbulent and laminar-transitional cases display a similar trend. Empirical fits are proposed which may be used for initial systems design.

Keywords: transpiration cooling
aerothermodynamics
heat transfer
hypersonics
infrared thermography
Advisor: McGilvray, M
Hermann, T