Aerothermal study of high-pressure turbine rotor tips and casing at engine representative conditions with high-speed rotation - Singh, D (2022)
Modern aircraft engines operate at high turbine entry temperatures to increase the cycle efficiency. The combustors employ intricate cooling techniques, to survive such high temperatures, resulting in distorted temperature profiles at the combustor exit. Further, extensive platform cooling, upstream of the Nozzle Guide Vanes (NGVs), results in additional radial distortions in the stage inlet temperature profile. Consequently, a rotating High-Pressure (HP) turbine blade is subjected to high cyclic thermal and mechanical loads. Shroudless turbines experience the highest heat load on the rotor tip, which makes it one of the life-limiting components for an engine. As the rotor tip starts to erode, the tip clearance increases which enhances the over-tip leakage flow, resulting in reduced stage efficiency and ultimately, the engine lifetime. Since the 1970s, extensive research has been undertaken to understand the over-tip leakage flow, and develop mitigation strategies such as novel tip designs, to minimize the losses and heat load. However, majority of the present findings are based on linear cascade or low-speed rotational studies, due to the high cost of the experiments and difficulty in instrumentation at engine conditions. Recent studies (mostly numerical) have shown the importance of testing at engine representative conditions, specifically targeting the transonic tip flow in a high-speed rotational environment, which is essential for an accurate understanding.