Abstract
The influence of endwall film cooling on the aerodynamic performance of a linear blade cascade employing an asymmetric contoured endwall is measured. Cylindrical coolant holes are strategically located on the contoured endwall to provide full coverage film cooling. The endwall contour is varied along the pitch direction, and is elevated near the pressure side and depressed near the suction surface. The profile height also varies in the axial direction from the inlet to exit. Measurements of total pressure, vorticity, and velocity components are obtained at different axial locations inside the passage for six inlet blowing ratios ranging from 1.0 to 2.4. The results are compared with the measured data on the contoured endwall without any film cooling flow (uncooled case). All the tests are performed in a low speed cascade facility employing a scaled up two-dimensional blade profile of the GE-E-3 turbine rotor section. The Reynolds number based on the cascade inlet velocity and blade actual chord is 2.30x 10(5). The results near the leading edge show that the suction side-leg vortex on the uncooled endwall weakens when the film cooling jets are employed. The axial vorticity and velocity vectors near the exit plane indicate that the passage vortex is located nearer to the mid-pitch location and higher above the endwall for the film-cooled endwall than for the uncooled contoured endwall. While the overall total pressure loss coefficient across the passage decreases at the high blowing ratios compared to the uncooled case, the overall cascade loss increases with the blowing ratio.