| description abstract | The effect of film-hole geometry and angle on turbine blade leading edge film cooling has been experimentally studied using the pressure sensitive paint technique. The leading edge is modeled by a blunt body with a semicylinder and an after-body. Two film cooling designs are considered: a heavily film cooled leading edge featured with seven rows of film cooling holes and a moderately film cooled leading edge with three rows. For the seven-row design, the film holes are located at 0 deg (stagnation line), ±15 deg, ±30 deg, and ±45 deg on the model surface. For the three-row design, the film holes are located at 0 deg and ±30 deg. Four different film cooling hole configurations are applied to each design: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. Testing was done in a low speed wind tunnel. The Reynolds number, based on mainstream velocity and diameter of the cylinder, is 100,900. The mainstream turbulence intensity is about 7% near of leading edge model and the turbulence integral length scale is about 1.5 cm. Five averaged blowing ratios are tested ranging from M=0.5 to M=2.0. The results show that the shaped holes provide higher film cooling effectiveness than the cylindrical holes, particularly at higher average blowing ratios. The radial angle holes give better effectiveness than the compound angle holes at M=1.0–2.0. The seven-row film cooling design results in much higher effectiveness on the leading edge region than the three-row design at the same average blowing ratio or same amount coolant flow. | |
