Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

Abstract

The film cooling effectiveness on the surface of a high pressure turbine blade is measured using the Pressure Sensitive Paint (PSP). Four rows of fan-shaped, laid-back compound angled cooling holes are distributed on the pressure side while two such rows are provided on the suction side of the blade. The coolant is only injected to either the pressure side or suction side of the blade at five average blowing ratios from 0.4 to 1.5. Presence of wake due to upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate a progressing wake. Effect of wake is recorded at four phase locations with equal intervals along the pitch-wise direction. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively, resulting in a blade pressure ratio of 1.14. Results reveal that the tip leakage vortices and endwall vortices sweep the coolant film on the suction side to the midspan region. The fan-shaped, laid-back compound angled holes produce good coolant film coverage on the suction side except for those regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and effectiveness level is low on the pressure surface. However, the pressure side acquires relatively uniform film coverage with the design of multiple rows of cooling holes. The presence of stationary upstream wake results in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.4 to 1.5 shows steady increase in effectiveness on the pressure side or the suction side for a given wake rod phases locations. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film cooling effectiveness particularly at higher blowing ratios.