Heat Transfer Coefficient and Film-Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

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.