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Research Papers

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

[+] Author and Article Information
R. M. Mathison

Gas Turbine Laboratory, Ohio State University, 2300 West Case Road, Columbus, OH 43235mathison.4@osu.edu

C. W. Haldeman

Gas Turbine Laboratory, Ohio State University, 2300 West Case Road, Columbus, OH 43235haldeman.5@osu.edu

M. G. Dunn

Gas Turbine Laboratory, Ohio State University, 2300 West Case Road, Columbus, OH 43235dunn.129@osu.edu

J. Turbomach 134(3), 031014 (Jul 15, 2011) (12 pages) doi:10.1115/1.4003173 History: Received August 17, 2010; Revised August 23, 2010; Published July 15, 2011; Online July 15, 2011

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 uncooled 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 © 2012 by American Society of Mechanical Engineers
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References

Figures

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Figure 1

Coolant-supply pressure history for a typical run with nominal flow rates

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Figure 2

Overview of stage instrumentation

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Figure 3

Uniform inlet temperature profiles

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Figure 4

Stanton number distribution for uniform inlet temperature and no cooling (Run 47)

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Figure 5

Net Stanton Number Reduction, uniform profile

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Figure 6

Effect of cooling on pressure-side Stanton number distribution for uniform inlet temperature profile

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Figure 7

Effect of cooling on suction-side Stanton number distribution for uniform inlet temperature profile

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Figure 8

Influence of cooling on Stanton number for the blade platform with uniform inlet temperature profile

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Figure 9

Fluid temperature above the blade platform for cooling variation with uniform inlet temperature profile

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Figure 10

Influence of cooling on Stanton number for the blade angel wing region with uniform inlet profile

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Figure 11

Influence of cooling on Stanton number for the vane side of the purge cavity for uniform inlet profile

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Figure 12

Influence of cooling on Stanton number for outer shroud and tip regions for uniform inlet profile

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Figure 13

Radial profiles, different cooling levels

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Figure 14

Stanton number distribution on airfoil for radial inlet temperature profile without cooling (Run 22)

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Figure 15

Influence of cooling on pressure surface Stanton number distribution for radial inlet temperature profile

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Figure 16

Influence of cooling on the suction surface Stanton number for radial inlet temperature profile

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Figure 19

Influence of cooling on Stanton number, outer shroud and blade tip, and radial inlet temperature profile

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Figure 20

Normalized adiabatic wall temperature for uniform inlet profile and nominal cooling conditions

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Figure 21

Calculation of adiabatic wall temperature

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Figure 22

Normalized adiabatic wall temperature, radial profile, and nominal cooling conditions

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Figure 17

Influence of cooling flow on Stanton number for blade platform with radial inlet temperature profile

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Figure 18

Fluid temperature above blade platform for cooling variation runs with radial inlet temperature profile

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