Full-coverage film cooling for a turbine blade with axial-shaped holes
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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 showerhead region, while axial laid-back fan-shaped 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 an 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 the 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 the average blowing ratio. The presence of a stationary upstream wake can be very detrimental to the film effectiveness on the blade surface depending on the wake phase location. Compared with the case of no showerhead injection, the spanwise 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 to the compound angle-shaped holes and the axial-shaped holes, whereas on the suction side, the compound angle-shaped holes provide better effectiveness than the axial-shaped holes due to jet deflection and an expanded hole breakout area. Copyright 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
