FILM-COOLING EFFECTIVENESS FROM SHAPED FILM COOLING HOLES FOR A GAS TURBINE BLADE

摘要

The effect of film cooling holes placed along the span of a fully-cooled high pressure turbine blade in a stationary, linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Six rows of compound angled shaped film cooling holes are provided on the pressure side while four such rows are provided on the suction side of the blade. The holes have a laidback and fan-shaped diffusing cross-section. Another three rows of holes are drilled on the leading edge to capture the effect of showerhead film coolant injection. The film cooling hole arrangement simulates a typical film cooled blade design used in stage 1 rotor blades for gas turbines used for power generation. A minimum blowing ratio is defined for each film hole row and tests are performed for 1.0x, 1.33x, 1.67x, 2.0x and 2.67x of this minimum value. Tests are performed for an inlet Mach number of 0.36 with a corresponding exit Mach number of 0.51. The flow remains subsonic in the throat region. The corresponding free stream Reynolds number, based on the axial chord length and the exit velocity, is 1.3 million. Turbulence intensity level at the cascade inlet is 5% with an integral length scale of around 5cm. Results show that varying blowing ratios can have a significant impact on film-cooling effectiveness distribution. Large spanwise variations in effectiveness distributions are also observed. Results also show that the effectiveness magnitudes from superposition of effectiveness data from individual rows are comparable with that from full coverage film cooling.