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Research Papers: Gas Turbines: Turbomachinery

Impact of Wall Temperature on Turbine Blade Tip Aerothermal Performance

[+] Author and Article Information
Q. Zhang

University of Michigan-Shanghai
Jiao Tong University,
Joint Institute,
Shanghai Jiao Tong University,
Shanghai, China
e-mail: QZhang@sjtu.edu.cn

L. He

Department of Engineering Science,
University of Oxford,
Oxford, UK
e-mail: Li.He@eng.ox.ac.uk

1Corresponding author.

Contributed by the Heat Transfer Committee of ASME for publication in the JOURNAL OF ENGINEERING FOR GAS TURBINES AND POWER. Manuscript received October 17, 2013; final manuscript received October 29, 2013; published online January 2, 2014. Editor: David Wisler.

J. Eng. Gas Turbines Power 136(5), 052602 (Jan 02, 2014) (9 pages) Paper No: GTP-13-1378; doi: 10.1115/1.4026001 History: Received October 17, 2013; Revised October 29, 2013

Currently the aerodynamics and heat transfer over a turbine blade tip tend to be analyzed separately with the assumption that the wall thermal boundary conditions do not affect the over-tip-leakage (OTL) flow field. There are some existing correlations for correcting the wall temperature effect on heat transfer when scaled to engine realistic conditions. But they were either developed to account for the temperature dependence of fluid properties largely empirically, or based on a boundary-layer model. It would be difficult (if not impossible) to define a boundary layer in many parts of a realistic blade passage with marked three-dimensional (3D) end wall and secondary flows (including those within a blade tip and around it). The questions to be asked here are: is the OTL aerodynamics significantly affected by the wall thermal condition? And if it is, how can we count this effect consistently in turbine blade tip design and analysis using modern CFD methods? In the present study the problem has been examined for typical high-pressure turbine blade tip configurations. An extensively developed RANS code (HYDRA) is employed and validated against the experimental data from a high speed linear cascade testing rig. The numerical analysis reveals that the wall–gas temperature ratio could greatly affect the transonic OTL flow field and there is a strong two-way coupling between aerodynamics and heat transfer. The feedbacks of the thermal boundary condition to aerodynamics behave differently at different flow regimes over the tip, clearly indicating a highly localized dependence of the convective heat transfer coefficient (HTC) upon wall temperatures. This implies that to use HTC for blade metal temperature predictions without resorting a fully conjugate solution, the temperature dependence needs to be corrected locally. A nonlinear correction approach has been adopted in the present work, and the results demonstrate its effectiveness for the transonic turbine tip configurations studied.

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Figures

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Fig. 1

Flat, winglet tip geometries and computational domain employed in the present study

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Fig. 2

Structured grids for the flat tip and winglet tip

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Fig. 3

Tip HTC distributions (a) experiment, (b) CFD prediction, with a tip gap clearance of 1.5% [19]

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Fig. 4

Comparisons of tip HTC distributions between results for TR = 0.92 and TR = 0.63 (G/S = 1.5%)

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Fig. 5

Comparisons of tip HTC distributions between results for TR = 0.92 and TR = 0.63 (G/S = 1.0%)

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Fig. 6

Mach number, density, and total temperature distributions on cut plane A (subsonic region) for TR = 0.63

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Fig. 7

Static pressure and HTC distributions for tip surface on cut plane A (subsonic region)

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Fig. 8

Mach number distributions on cut plane B (transonic region) for (a) TR = 0.92 and (b) TR = 0.63

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Fig. 9

Density distributions on cut plane B (transonic region) for (a) TR = 0.92 and (b) TR = 0.63

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Fig. 10

Surface static pressure and wall heat flux along tip wall distance of cut plane B (transonic region) with different gas–wall temperature ratios (dash line indicates a shift of shock wave footing)

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Fig. 11

HTC distributions for tip surface along tip wall distance of cut plane B (transonic region)

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Fig. 12

Mass flux ratio distributions along the curve length of the suction side edge for cases with TR = 0.92 and 0.63

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Fig. 13

Comparisons of winglet tip HTC distributions between results for TR = 0.92 and TR = 0.63 (G/S = 1.0%)

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Fig. 14

Heat flux variations with gas–wall temperature ratio at location x/L = 0.33 of cut plane B (symbols: direct CFD solutions; red dash line: three-point correction)

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Fig. 15

Comparisons of heat flux results for TR = 0.63 calculated from different approaches along tip wall distance of cut plane B (transonic region)

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Fig. 16

Heat flux distributions for a winglet tip obtained by different approaches (TR = 0.50)

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Fig. 17

Percentage difference of heat flux relative to the direct solution obtained by two approaches (TR = 0.50)

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