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Research Papers

Experimental Characterization of the Vane Heat Flux Under Pulsating Trailing-Edge Blowing

[+] Author and Article Information
J. Saavedra

Maurice J. Zucrow Laboratories
Purdue University,
West Lafayette, IN 47907
e-mail: saavedra@purdue.edu

G. Paniagua

Maurice J. Zucrow Laboratories
Purdue University,
West Lafayette, IN 47907

B. H. Saracoglu

von Karman Institute for Fluid Dynamics,
Rhode Saint Genèse,
Brussels 1640, Belgium

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received September 18, 2016; final manuscript received September 29, 2016; published online February 1, 2017. Editor: Kenneth Hall.

J. Turbomach 139(6), 061004 (Feb 01, 2017) (7 pages) Paper No: TURBO-16-1246; doi: 10.1115/1.4035211 History: Received September 18, 2016; Revised September 29, 2016

The steady improvement of aircraft engine performance has led toward more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency. A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 × 106 and 6 × 106). The steady and unsteady heat flux was retrieved using thin-film two-layered gauges. The complexity of the tests required the development of an original heat transfer postprocessing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient. The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant

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References

Figures

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Fig. 1

(a) Shock interactions over a transonic turbine stage and (b) zoom on the expansion and compression trailing edge waves

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Fig. 2

(a) Double layered domain, (b) airfoil trailing edge and thin film example, and (c) vane and gauge locations

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Fig. 3

(Left) Heat conduction across multilayered substrates, (Right) Heat flux data reduction based on 1D semi-infinite assumption with a Crank Nicolson scheme

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Fig. 4

Validation of the code developed to solve the transient heat transfer

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Fig. 5

Heat transfer coefficient computation based on heat flux data and adiabatic wall temperature definition

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Fig. 6

Unsteady contribution of the heat transfer coefficient characterization

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Fig. 7

Experimental total pressure and temperature

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Fig. 8

Methodology application

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Fig. 9

Criteria application for test window definition

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Fig. 10

(a) and (b) Blow down facility at von Karman Institute, (c) turbine airfoil cooling distribution, (d) test section, and (e) test airfoil

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Fig. 11

Trailing blowing ratio effect on the heat transfer coefficient distribution

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Fig. 12

Experimental (a) and numerical (b) Schlieren images of the test section at Mach 1.1 [7]

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