Aerothermal study of high-pressure turbine rotor tips and casing at engine representative conditions with high-speed rotation

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. The Oxford Turbine Research Facility (OTRF) is a high-speed rotating transient test facility that allows unsteady aerodynamics and heat transfer measurements, at engine representative conditions. This thesis presents an experimental investigation, of the HP turbine rotor tip and casing, performed in the OTRF. The single-stage HP turbine consisted of cooled vanes (featuring film and trailing edge slot cooling), and uncooled rotor blades. Three different tip designs (two squealers and one flat) were tested, at two tip gaps, and with two stage inlet temperature profiles that included a spatially uniform total temperature profile and a radially distorted total temperature profile representative of the HP turbine inlet in a modern turbofan engine. The unsteadiness in the heat transfer data, was found to significantly rise with the distorted temperature profile, and has been quantified at the stage inlet. The increased unsteadiness causes complications in the heat transfer data analysis and hence, new processing techniques were developed. A detailed investigation, including a parametric study and uncertainty analysis, for heat transfer calculations from thin-film gauges has been presented in the thesis. Time-averaged and time-resolved static pressure and heat transfer measurements were acquired at the rotor casing. In addition, rotor surface heat transfer data was acquired in the near-tip region using thin-film gauges and thermocouples. Comparisons have been drawn across different tip designs, tip gaps and inlet temperature profiles. The hot gas migration effect, responsible for the extensive heat load on the rotor tip, has been examined. The casing thin- film gauge measurements from the OTRF, were found to show an excellent match with the Rolls-Royce engine thermal paint measurements. In parallel, a numerical investigation of the experimental test cases, was conducted using unsteady Computational Fluid Dynamics (CFD) predictions. The experimental results from the present study, provided a unique dataset to validate and understand the deficiencies in the present CFD models.