The thermal efficiency of a gas turbine is largely dependent on the turbine inlet temperature (TIT). Modern gas turbines may operate at temperatures as high as 2000K, which is higher than the melting point of the material in use. Hence, thermal protection of gas turbine hot components is a big challenge. Film cooling is the most common cooling technique adopted in this application. In film cooling, coolant is injected at discrete locations along the metal surface, which forms a layer of cool air immediately over the hot surface, thus, protecting it from direct contact with hot mainstream air. The cooling is strong along the centerline of the hole in the downstream region and rapidly decreases over the span-wise direction. The distributed cooling can result in large thermal gradients, inducing thermal stresses in the material. The region with least cooling may lead to a cascade failure of the blade. Film cooling with backward injection holes has been proved to reduce this effect. In the current work, backward coolant injection scheme is explored under fan-shaped holes numerically. Fluent, a commercial CFD software, is used in the current work for numerical simulations. Effects of blowing ratio, injection angle, and turbulence are considered. Numerical results show that fan-shaped holes are better than simple cylindrical holes in terms of both cooling effectiveness and its uniformity. Numerical results are validated with experimental results. The image of temperature fields on cooling surface is captured with an Infrared camera.

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