Full Coverage Film Cooling for a Turbine Blade with Axial Shaped Holes
Conference Paper
Overview
Identity
Additional Document Info
Other
View All
Overview
abstract
The film cooling effectiveness on a fully film cooled high pressure turbine blade is investigated using the Pressure Sensitive Paint technique. Three rows of radial angled cylindrical holes are arranged in the leading edge region, while axial laidback fanshaped holes are provided on the blade surfaces with four rows on the pressure side and two rows on the suction side. The shaped holes are featured with 10 lateral expansion from the hole centerline and additional 10 forward expansion to the blade surface. The coolant ejects through all the film cooling holes at four average blowing ratios ranging from 0.3 to 1.2. The influence of wake from upstream vane is simulated by placing a periodic set of rods upstream of the test blade. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The Mach numbers at the inlet and the exit are 0.27 and 0.44, respectively, resulting in a blade pressure ratio of 1.14. Results show the film cooling effectiveness increases with increasing of average blowing ratio. The presence of upstream wake rods can be very detrimental to the film effectiveness on the blade surface depending on the wake rod phase location. Compared with the case of no showerhead injection, the spanwsie averaged film cooling effectiveness in the downstream region of the pressure side and suction side rows increases with the showerhead injection. The film effectiveness on the pressure side is comparable for compound angle shaped holes and axial shaped holes; while on the suction side, the compound angle shaped holes provide better effectiveness than the axial shaped holes due to jet deflection and expanded hole breakout area..
