Turbine Tip Film Cooling With Fanned Holes

Abstract Improving the efficiency of jet engines is essential to reduce aircraft emissions. This requires increases in High-Pressure Turbine (HPT) entry temperatures, resulting in a rise in thermal loading which is particularly problematic for the HPT blade tip. Hot gas leaks through the clearance gap between the blade and the casing causing high heat fluxes, most notably on the thin squealers which are challenging to adequately cool, expediting degradation and reducing efficiency. This work examines how tip temperatures can be reduced by using fanned holes to cool the blade tip instead of conventional cylindrical holes. Representative linear cascade experiments demonstrate a local 15–45% improvement in cooling film effectiveness across mass flow rates and tip clearances on the critical pressure side rim, and a 10% decrease in heat transfer coefficient across the tip. Numerical computational fluid dynamics simulations complement the experimental work and show that the diffusion of flow by the fanned cooling holes increases the attachment of the coolant to the blade, delays mixing with the mainstream gas and distributes the coolant over a larger surface area. As coolant mass flow increases, the improvement margin diminishes due to greater coolant separation from the blade surface at higher blowing ratios for both fanned and cylindrical cooling holes. Increasing tip clearance results in heightened coolant dilution caused by larger over-tip leakage mass flow rates.