Abstract

Improving the film cooling technique provides more cooling capacity to withstand the harsh thermal environment in the next generation of gas turbines. A two-dimensional cascade has been designed and constructed in a subsonic wind tunnel in order to investigate the heat transfer of shaped holes over a gas turbine stator. An in-situ calibration technique has been developed to obtain the film cooling performance without disturbing the aerodynamic flow around the vane surface. Subsequently, the cooling performance of two types of shaped holes is measured at different positions over the entire surface. Firstly, a louver scheme was investigated on the convex surface of the suction side and on the concave surface of the pressure side. In addition, a proposed smooth expansion was investigated over the highly curved surface of the leading edge. The location of the hole has a high impact on the cooling performance due to the difference in curvature. The investigated blowing ratios slightly affect the cooling performance of the presented schemes due to the considerable reduction in the jet momentum that impedes the jet lift-off at exit. The shaped holes provide a higher net heat flux reduction compared with the similar cylindrical holes and other shaped holes in the literature. The contribution of this study will help to enhance the cooling performance in the next generation of gas turbines.