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Turbine Endwall Film Cooling With Pressure Side Radial Holes

[+] Author Affiliations
Yang Zhang, Xin Yuan

Tsinghua University, Beijing, China

Paper No. GT2013-95273, pp. V03BT13A049; 11 pages
  • ASME Turbo Expo 2013: Turbine Technical Conference and Exposition
  • Volume 3B: Heat Transfer
  • San Antonio, Texas, USA, June 3–7, 2013
  • Conference Sponsors: International Gas Turbine Institute
  • ISBN: 978-0-7918-5515-7
  • Copyright © 2013 by ASME


A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the nozzle guide vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the side gill pressure region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with radial cylindrical holes on the pressure side.

The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. Four rows of staggered radial film-cooling holes were placed at the pressure side gill region. The diameter of the cylindrical holes was 1 mm and the length was 5 d, with a hole space of 6 d. The spanwise angle of the cooling holes was 35 ° and the radial angle was 90 °. Three blowing ratios were chosen as the test conditions in the experiment, M = 0.7, M = 1.0 and M = 1.3. The film-cooling effectiveness was probed using PSP (pressure sensitive painting) technology and the post processing was performed by means of a mass and heat transfer analogy.

Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface near the pressure side gill region increased, with the highest parameter at X/Cax = 0.3; 2) a double-peak cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the pressure side radial cooling holes was apparent on the endwall surface near the gill region, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the pressure film cooling could only be detected in the downstream area of the endwall at the higher blowing ratio.

Copyright © 2013 by ASME



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