Numerical simulation of influence of leading edge film-cooling on aerodynamic parameters in turbine cascade

Finite difference scheme with three-order precision and TVD property, arbitrarily-curved surface technique and multi-block computational method were employed. The film-cooling flow field in turbine cascade with three rows of coolant holes at leading edge was calculated and its influence on aerodynamic parameters was discussed under two blowing ratios. The coolant ejection near blade surface was illustrated to show the difference of cooling film formation on the stagnation region, pressure and suction surfaces of the leading edge. The result shows that inlet mass flux and local mach number are both reduced appreciably as coolant flux increases. Profile pressure is not changed obviously in the whole, with the exception in the local region near the coolant hole with a pressure variation, which is greater on the suction side. The adiabatic effectiveness on the stagnation line and suction surface are higher than that on pressure surface. The discipline of coolant flow on the pressure side of leading edge is very complicated.