Performance Comparison of Different Leading Edge Cooling Layouts Considering Inlet Swirl

Abstract In modern aero-engine and gas turbine, the interaction between the combustor and the turbine has a significant impact on both aerodynamic and cooling performance of the turbine. Due to the interaction between the combustor and the turbine, non-uniform distributions of temperature and velocity can be observed on the inlet surface of the turbine, which notably affect the performance of turbine blade. The primary research interest of this work is to improve the cooling performance by optimizing the layout of the cooling holes according to the researches on the influence of swirl. First, cooling performance comparison of different inlet swirl is conducted using the measurement based on Pressure-Sensitive-Paint (PSP) technology. The swirl generator with the swirl number 0.778 is used to simulate the non-uniform velocity inlet and the related distance of vane was changed. As the swirl which core has the same circumferential position with the leading edge caused the worst influence to film cooling and made the coverage of film become highly asymmetric, then asymmetric cooling hole layouts, which cooling hole distributions driven by sinusoidal function, were studied by numerical methods and find they can clearly improve the cooling performance on swirl condition. Lastly, based on the experimental results and the analyses, a new layout based on the stagnation line is proposed, which can increase the uniformity of film cooling effectiveness and the performance is verified by PSP experiment. The new layout can increase the area-averaged film cooling effectiveness up to 9.7% under high blowing ratio and reduce the standard deviation of film cooling effectiveness to 24.3% under low blowing ratio. Thus changing the layout of film holes especial the leading edge is an efficient and practical method to improve the cooling performance under swirl condition.