Numerical investigations into heat transfer and film cooling effect on a transonic blade endwall

摘要

Highlights?Heat transfer/film cooling on endwall with different hole-arrangements is studied.?Effect of blowing ratio on heat transfer/film-cooling on endwall is investigated.?Aerodynamic performance in cascade with different film cooling is also analyzed.?Predicted Stanton number on endwall compares favorably with the experimental data.?Thermal load at three high heat transfer areas are reduced by the hole-arrangements.AbstractThe RANS (Reynolds Averaged Navier-Stokes) equations were numerically solved to investigate the heat transfer and film cooling effect on the first stage rotor blade endwall in NASA transonic gas turbine. To reduce the thermal load on the upstream of leading edge, shoulder area and corner area of endwall, different film cooling strategies were presented to evaluate the Stanton number, film cooling effectiveness on endwall and also the aerodynamic performance in cascade at three blowing ratios. Through the film cooling upstream leading edge, the area-averagedSton blade endwall is reduced by 1.4% compared with no film cooling case. Through the shoulder area film cooling, the area-averagedStnumber on endwall is reduced by 1%, and the area-averaged film cooling effectiveness is increased by 44.4% atM?=?0.5. With double hole-arrays at the corner area, the area-averagedStnumber on the endwall can be reduced by 1–2%, and the area-averaged film cooling effectiveness can be improved by 10.2–13.6%. The energy loss coefficient in the cascade is almost not affected by the blowing ratio for the leading edge and shoulder area film cooling configurations, while it is sensitive to the blowing ratio for the corner area film cooling.]]>