Aero-thermal performance of transonic high-pressure turbine blade tips

This thesis presents an experimental and theoretical study of unshrouded tips on gas turbine blades with and without winglets. The goal of the winglet tip design is to reduce over-tip leakage flow. To assess the potential use of winglets, their effectiveness in reducing loss and the cooling scheme's effectiveness must be determined under engine-realistic flow conditions. An aero-thermal investigation was performed to answer these questions.
A high-speed linear cascade was commissioned for use under transonic conditions. Once realistic inlet and exit conditions were established, the heat transfer measurement technique used was qualified. An infrared thermography technique was used to obtain a detailed time history of the surface temperature, which was used to reconstruct heat flux and calculate heat transfer coefficients. Several processing techniques were compared to determine that a technique using an impulse response digital filter in the frequency domain has the smallest uncertainty, is the most consistent and is the most computationally efficient.
Heat transfer tests were then performed on a flat blade tip under transonic and low-speed flow conditions, and results showed both qualitative and quantitative differences, highlighting the need for testing at engine-realistic conditions. A heat transfer tip gap survey was then performed on a flat tip, and this agreed well with an independent computational study. Uncooled and cooled winglet tip heat transfer was also examined, showing that the uncooled winglet has nearly 3.2 times the area-averaged heat load of a flat tip and the cooled winglet has nearly 3.4 times the area-averaged heat load of a flat tip. Cooled winglet heat transfer results showed enhanced heat transfer compared to the uncooled winglet and an increase in film cooling effectiveness with a smaller tip gap. Reflecting the complexity of testing at high-speed conditions, these are the first detailed, spatially-resolved heat transfer results on the blade tip under transonic conditions.
Aerodynamic loss measurements were obtained using a specially-designed three-hole probe and a single-hole probe to resolve the near-casing region. The flat tip experimental and computational results agree well. The winglet results showed a decrease in loss with coolant injection.
Finally, a numerical study is carried out to investigate the influence of 2-D corner conduction on the accuracy of heat transfer measurements when using the conventional semi-infinite 1-D assumption. The errors in the conventionally processed heat transfer coefficient are shown to be very high around the corner. A relatively simple and efficient new correction method has been developed and shown to be effective.

Keywords: Spatially-resolved
Tip Gap Survey
Infrared Thermography
2-D Corner Conduction
High-Speed Linear Cascade
Three-hole Probe
Experimental
Net Heat Flux Reduction
Film Cooling Effectiveness
Over-tip Leakage Flow
Aerodynamic Loss
Transonic Flow Conditions
High-Pressure Turbine
Heat Transfer Measurement Techniques
Lateral Corner Correction Coefficient
Impulse Technique
Cooled Winglet
Unshrouded Blade Tip Heat Transfer
Heat Flux Reconstruction
Advisor: He, L
Oldfield, M