Research Papers: Gas Turbines: Heat Transfer

Computational Study of the Effects of Shock Waves on Film Cooling Effectiveness

[+] Author and Article Information
C. X.-Z. Zhang

 Concordia University, Montréal, QC, H3G 1M8, Canada

I. Hassan1

 Concordia University, Montréal, QC, H3G 1M8, CanadaIbrahimH@alcor.concordia.ca


Corresponding author.

J. Eng. Gas Turbines Power 131(3), 031901 (Feb 10, 2009) (10 pages) doi:10.1115/1.3026568 History: Received April 13, 2008; Revised July 09, 2008; Published February 10, 2009

The performance of a louver cooling scheme on a transonic airfoil has been studied numerically in this paper. Film cooling holes are located near the passage throat. The Mach number at the location of the jet exit is close to unity. A comparison of film cooling effectiveness between numerical prediction and experimental data for a circular hole shows that the numerical procedures are adequate. In addition to the shock-wave effects and compressibility, curvature effect was also studied by comparing cooling effectiveness on the airfoil surface with that on a flat plate. Substantially higher cooling effectiveness for the louver cooling scheme on the airfoil was predicted at blowing ratios below 1 in comparison to other cooling configurations. At higher blowing ratios than 2 the advantages of the louver cooling scheme become less obvious. It was also found that for the same cooling configuration the cooling effectiveness on the transonic airfoil is slightly higher than that on a flat plate at moderately low blowing ratios below 1. At high blowing ratios above 2 when the oblique shock becomes detached from the leading edge of the hole exits, dramatic reduction in cooling effectiveness occurs as a result of boundary layer separation due to the strong shock waves. A coolant-blockage and shaped-wedge similarity was proposed and found to be able to qualitatively explain this phenomenon satisfactorily.

Copyright © 2009 by American Society of Mechanical Engineers
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Figure 1

Geometries of the cooling schemes (a. b. c. reproduced from (18))

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Figure 2

Test section and computational domain with boundary conditions

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Figure 3

Multi-block structured mesh

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Figure 4

Contours of Mach number on the cross section without film cooling hole

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Figure 5

Mach number distribution around the airfoil surface without film cooling hole

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Figure 6

Laterally averaged η for different cooling configurations

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Figure 7

Predicted laterally averaged η at high blowing ratios

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Figure 8

Curvature effect at different blowing ratios

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Figure 9

Contours of cooling effectiveness

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Figure 10

Velocity profiles at different locations along the centerline

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Figure 11

Mach number distribution for the circular hole case at different blowing ratios

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Figure 12

Mach number distribution for the louver scheme at different blowing ratios

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Figure 13

Contours of Mach number on iso-surfaces of constant temperature at m=1

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Figure 14

Schematic of shock wave structures at different blowing ratios

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Figure 15

Static pressure distribution around the jet exit at m=3



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