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Research Papers

Fully Cooled Single Stage HP Transonic Turbine—Part II: Influence of Cooling Mass Flow Changes and Inlet Temperature Profiles on Blade and Shroud Heat-Transfer

[+] Author and Article Information
C. W. Haldeman, M. G. Dunn, R. M. Mathison

Gas Turbine Laboratory, Ohio State University, Columbus, OH 43235

J. Turbomach 134(3), 031011 (Jul 15, 2011) (11 pages) doi:10.1115/1.4002968 History: Received June 30, 2010; Revised July 01, 2010; Published July 15, 2011; Online July 15, 2011

A fully cooled transonic high-pressure turbine stage is utilized to investigate the combined effects of turbine stage cooling variation and vane inlet temperature profile on heat transfer to the blades with the stage operating at the proper design corrected conditions. For this series of experiments, both the vane row and the blade row were fully cooled. The matrix of experimental conditions included varying the cooling flow rates and the vane inlet temperature profiles to observe the overall effect on airfoil heat-transfer. The data presented in Part I focused on the aerodynamics of the fully cooled turbine for a subset of the cases investigating two vane inlet temperature profiles (uniform and radial) and three different cooling levels (none, nominal, and high) for the high Reynolds number condition. This part of the paper focuses on the time-average heat-flux measurements on the blade and shroud region for the same cooling mass flow rates and vane inlet temperature profiles. The cooling effects are shown to be small and are centered primarily on the suction side of the airfoil. This relatively small influence is due to the ratio of the cooling gas to metal temperature being closer to 1 than the design value would dictate. The vane inlet temperature profile effects are more dominant, and using a net Stanton number reduction factor to compare the cases, an effect on the order of about 0.25 is demonstrated. This effect is due primarily to the change in the reference temperature used for the Stanton number calculation. The differences due to profile effects are small but observable toward the trailing edge of both the blade and rotor shroud. This data set forms an excellent baseline for heat-flux calculations, as the variation in the main input conditions are well documented and do not produce large changes in the heat-flux. It provides insight into the flow physics of an actual engine and guidelines about proper normalization of variables for a cooled turbine stage, supporting further development of computational heat-flux modeling techniques.

Copyright © 2012 by American Society of Mechanical Engineers
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References

Figures

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Figure 1

Heat-flux gauges at 50% and 90% span

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Figure 2

Schematic of blade internal cooling passages

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Figure 3

Photograph of recessed tip instrumented

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Figure 4

Stationary shroud thin-film heat-flux gauges

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Figure 5

Temperature profile shape for uniform and radial runs (all cooling levels)

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Figure 6

Schematic of profile and mass flow problem

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Figure 7

Effect of cooling levels on 50% span for uniform profiles

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Figure 8

Net Stanton number reduction for uniform runs, 50% span

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Figure 9

50% heat-flux and pressure uniform profile, suction side (expanded view)

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Figure 10

Build 1 data, uniform profile, 50% span

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Figure 11

Blade coolant temperature (build 1 versus build 2)

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Figure 12

Radial and uniform heat-flux, different spans

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Figure 13

Uniform cooling cases: rotor shroud

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Figure 14

Shroud cooling and profile effects

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Figure 15

Unsteady data, 59.1% axial chord, shroud

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Figure 16

Variation due to profile and cooling on shroud, using rake temperature at 97.5% span as reference

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Figure 17

Profile effects, blade 50% span, using 50% reference temperature from rakes

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Figure 18

Blade Nusselt numbers, 50% span

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