Investigation of aerothermal performance of blade tip with conical holes in transonic flow

The blade tip with conical holes is proposed in the current paper to improve the film cooling performance on the turbine blade tip. The film cooling performance of the blade tip with cylindrical holes is studied as a reference case. The pressure sensitive paint technique is adapted to obtain the tip film cooling effectiveness in transonic flow. The numerical simulation is used to obtain the tip flow field and aerodynamic loss. The cascade exit Mach number and inlet Reynolds number are 1.05 and 370000, respectively. The experiment and calculation are carried out at two tip clearance gaps of 0.7% and 1.5% and four mass-flow-ratios. The results show that at low mass flow ratios, the film cooling effectiveness of the blade tip with conical holes is significantly higher than that of the blade tip with cylindrical holes at the tip mid-chord region. Increasing the tip clearance gap from 0.7% to 1.5% shows a negative effect on the film cooling performance of the blade tip with conical holes in the tip mid-chord region, particularly for the low mass flow ratio. The pitch-wise averaged aerodynamic loss coefficient of the blade tip with cylindrical holes is lower than that of the blade tip with conical holes near the shroud.