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Research Papers

Multirow Film Cooling Performances of a High Lift Blade and Vane

[+] Author and Article Information
S. Naik

e-mail: shailendra.naik@power.alstom.com

M. Schnieder

Alstom,
Brown Boveri Strasse,
Baden 5401, Switzerland

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received May 15, 2013; final manuscript received July 6, 2013; published online September 27, 2013. Editor: Ronald Bunker.

J. Turbomach 136(5), 051003 (Sep 27, 2013) (8 pages) Paper No: TURBO-13-1076; doi: 10.1115/1.4025274 History: Received May 15, 2013; Revised July 06, 2013

This paper investigates the aerodynamic and film cooling effectiveness characteristics of a first stage turbine high lift guide vane and its corresponding downstream blade. The vane and blade geometrical profiles and operating conditions are representative of that normally found in a heavy-duty gas turbine. Both the vane and the blade airfoils consist of multirow film cooling holes located at various axial positions along the airfoil chord. The film cooling holes are geometrically three-dimensional in shape and depending on the location on the airfoil, they can be either symmetrically fan shaped or nonsymmetrically fan shaped. Additionally the film cooling holes can be either compounded or in-line with the external flow direction. Numerical studies and experimental investigations in a linear cascade have been conducted at vane and blade exit isentropic Mach number of 0.8. The influence of the coolant flow ejected from the film cooling holes has been investigated for both the vane and the blade profiles. For the nozzle guide vane, the measured film cooling effectiveness compared well with the predictions, especially on the pressure side. The suction side film cooling effectiveness, which consisted of two prethroat film rows, proved very effective up to the suction side trailing edge. For the blade, there was a reasonable comparison between the measured and predicted film cooling effectiveness. Again the blade prethroat fan shaped cooling holes proved very effective up to the suction side trailing edge. For the vane, the impact of varying the blowing ratios showed a strong variation in the film cooling effectiveness on the pressure side. However, on the blade, the effect of varying the blowing ratio had a greater impact on the suction side film effectiveness compared to the pressure side.

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Figures

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Fig. 2

Vane linear cascade test rig [7]

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Fig. 1

High-speed four-passage blade linear cascade

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Fig. 3

Vane cascade (a) CFD domain, (b) mesh

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Fig. 13

Measured film effectiveness against correlation for the last row on the suction side

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Fig. 16

Comparison of measured and predicted film cooling effectiveness at nominal blowing ratio

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Fig. 14

Close-up of the film effectiveness at the trailing edge pressure side bleed

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Fig. 15

TLC response contours from the blade film cooling

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Fig. 4

Vane isentropic Mach number

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Fig. 5

Oil flow distribution on (a) vane suction side, (b) detailed view of suction side

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Fig. 6

Overview of blade (a) 3D CFD model, (b) mesh

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Fig. 7

Blade predicted and measured isentropic mach numbers at 50% span

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Fig. 8

Oil flow visualization on blade suction side

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Fig. 9

Influence of incidence angle on blade mach numbers

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Fig. 10

TLC response of the vane film effectiveness with different camera views (a) on pressure side, (b) on suction side

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Fig. 11

Measured film effectiveness for the vane at nominal blowing rate, compared with prediction

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Fig. 12

Effect of the blowing rate on measured film effectiveness

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Fig. 17

Comparison of measured film cooling effectiveness at nominal, low and high blowing ratios

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Fig. 18

Qualitative comparisons of film cooling effectiveness (a) measurements, (b) predictions

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