Heat Transfer Coefficient and Film-Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade
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Experimental investigations were performed to measure the detailed heat transfer coefficients and film-cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film-cooling holes on the pressure-side near the tip region and the tip surface along the camber line. A hue detection based transient liquid crystal technique was used to measure heat transfer coefficients and film-cooling effectiveness. All measurements were done for the tip gap clearances of 1.0%,1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1 106 and the overall pressure ratio was 1.32. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film-cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film-cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.
