The steady improvement of aircraft engine performance has led towards more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency.
A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 and 6 million). The steady and unsteady heat flux was retrieved using thin-film 2-layered gauges. The complexity of the tests required the development of an original heat transfer post-processing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient.
The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant.