Influence of Boundary-Layer Turbulence on Effusion Cooling at Mach 2.67

Future rocket-nozzle extensions have to be thermally protected by a film of cooling gas. The hot supersonic boundary-layer flow may be fully turbulent despite a strong favorable pressure gradient and strongly cooled walls. In contrast to previous investigations, where a laminar boundary-layer state was considered, the present work deals with the effects of a turbulent flow. The cooling film is generated by wall-normal cooling-gas injection and, hence, the term effusion cooling is used. The investigations of a flat-plate boundary-layer flow with Mach number 2.67 and zero pressure gradient are performed using direct numerical simulations. Air is employed as hot and cooling gas and is injected into the adiabatic turbulent boundary layer through a discrete slit, which is infinite in spanwise direction. The results are compared with a corresponding laminar case. Despite a fuller mean-flow profile in the turbulent case, keeping the cooling gas closer to the wall if it were truly steady, the averaged turbulent cooling effectiveness decreases significantly due to an increased mixing of the hot flow and the cooling gas.