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Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine: Part I—Influence of Cooling Variation

[+] Author Affiliations
R. M. Mathison, C. W. Haldeman, M. G. Dunn

The Ohio State University, Columbus, OH

Paper No. GT2010-22713, pp. 271-284; 14 pages
doi:10.1115/GT2010-22713
From:
  • ASME Turbo Expo 2010: Power for Land, Sea, and Air
  • Volume 4: Heat Transfer, Parts A and B
  • Glasgow, UK, June 14–18, 2010
  • Conference Sponsors: International Gas Turbine Institute
  • ISBN: 978-0-7918-4399-4 | eISBN: 978-0-7918-3872-3
  • Copyright © 2010 by ASME

abstract

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the un-cooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton Number distribution on the blade.

Copyright © 2010 by ASME

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