Conference Dates

June 22-27, 2014


The purpose of this paper is to find the necessary cooling mass flow for a defined wall thickness distribution, which allows a selective cooling for an acceptable temperature range of the sharp leading edge of an atmospheric re-entry vehicle. Due to the angle of attack during the re-entry flight the pressure at the top side is lower than the pressure at the bottom side. However, the highest heat load occurs at the stagnation point at which also the maximum pressure is effective. The efficiency of a transpiration cooling depends on the mass flow rate of the coolant. The cooling mass flow is therefore determined by the pressure difference between the ambient and reservoir pressure. Thus, the coolant mass flow increases if the pressure difference between the reservoir and ambient pressure increases. If so a vehicle dives into earth atmosphere, the maximum coolant mass flow is expected in a higher atmosphere. In particular, the cooling effect on the surface caused by the different pressures between top side and bottom side will be considered in more detail. These different pressures are resulting from oblique shocks caused by an angle of attack with α=5°. A possibility to compensate this effect is to adapt the wall thickness so that the coolant mass flow is constant over the transpiration cooled surface, or higher on the hot side. For this purpose the numerical code Heat Exchange Analysis for Transpiration-cooling Systems (HEATS) will be adapted to compute the temperature distribution for a defined transpiration cooled leading edge geometry and trajectory. The results are showing, for a constant mass flow for top side and bottom side, a temperature difference of about 500K at the surface, only caused by the angle of attack. In addition, the impact to the reservoir pressure is higher than 0.5bar, if the cooling mass flow is assumed as constant.