Experimental investigation of dust-pan shaped Hole film cooling characteristics on pressure side of a turbine blade in a linear transonic cascade

An experimental research of film cooling performance of three single dust-pan shaped hole rows in different positions of a turbine blade was carried out in the short-duration transonic linear cascade at stationary condition, which can model realistic engine aerodynamic conditions. The effects of inlet Reynolds number (Re_in=2.5×10⁵~7.5×10⁵), isentropic exit Mach number (Ma_is=0.71~0.91) and coolant blowing ratio (M=0.8~2.6) on film cooling effectiveness are investigated. Three single hole rows are located at 11.7%, 36.3% and 55.6% relative arc on the pressure sides of three enlarged blade models respectively. The adiabatic film cooling effectiveness are derived from the surface temperatures based on transient heat transfer measurement method. The results show that in the range of blowing ratios studied in the present paper, for location 3 the cooling effectiveness decreases a lot with blowing ratio increasing due to the lift-off of coolant at high blowing ratios, while for location 1 and 2, the film cooling effectiveness increases with blowing ratio increasing, because the strong favorable pressure gradient and high concave curvature near the leading edge lead to a good attachment of coolant on the surface. At M≤1.0 conditions, the film cooling effectiveness of location 1 and 2 is lower than that of location 3, which reflects that strong favorable pressure gradient and high concave curvature weaken film cooling performance at low blowing ratio conditions, while the effect is opposite when M is greater than 1.0. For location 1, the highest general cooling performance is obtained at Re_in=2.5×10⁵ condition, and for location 2, the change of Rein has different effects on cooling effectiveness in different regions. In the range of Ma_is studied in this paper, the change of Ma_is has little effect on film cooling effectiveness.