Research Papers

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine—Part II: Effect of Cooling Variation on the Vane Airfoil and Inner End Wall

[+] Author and Article Information
Harika S. Kahveci

e-mail: harika.kahveci@ge.com

Charles W. Haldeman

e-mail:  haldeman.5@osu.edu

Randall M. Mathison

e-mail:  mathison.4@osu.edu

Michael G. Dunn

e-mail:  dunn.129@osu.edu
The Ohio State University Gas Turbine Laboratory,
2300 West Case Rd.
Columbus, OH 43235

1Now at GE Energy in Greenville, SC.

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received July 18, 2011; final manuscript received March 25, 2012; published online November 8, 2012. Editor: David Wisler.

J. Turbomach 135(2), 021028 (Nov 08, 2012) (8 pages) Paper No: TURBO-11-1145; doi: 10.1115/1.4006776 History: Received July 18, 2011; Revised March 25, 2012

The impact of film cooling on heat transfer is investigated for the high-pressure vane of a 1-1/2 stage high-pressure turbine operating at design corrected conditions. Cooling is supplied through three independently controllable circuits to holes in the inner and outer end wall, vane leading edge showerhead, and the pressure and suction surfaces of the airfoil, in addition to vane trailing edge slots. Four different overall cooling flow rates are investigated and one cooling circuit is varied independently. All results reported in this part of the paper are for a radial inlet temperature profile, one of the four profiles reported in part I of this paper. Part I describes the experimental setup, data quality, influence of inlet temperature profile, and influence of cooling when compared to a solid vane. This part of the paper shows that the addition of coolant reduces airfoil Stanton number by up to 60%. The largest reductions due to cooling are observed close to the inner end wall because the coolant to the majority of the vane is supplied by a plenum at the inside diameter. While the introduction of cooling has a significant impact on Stanton number, the impact of changing coolant flow rates is only observed for gauges near 5% span and on the inner end wall. This indicates that very little of the increased coolant mass flow reaches all the way to 90% span and the majority of the additional mass flow is injected into the core flow near the plenum. Turning off the vane outer cooling circuit that supplies coolant to the outer end wall holes, vane trailing edge slots, and three rows of holes on the pressure surface of the airfoil, has a local impact on Stanton number. Changes downstream of the holes on the airfoil pressure surface indicate that internal heat transfer from the coolant flowing inside the vane is important to the external heat transfer, suggesting that a conjugate heat-transfer solution may be required to achieve good external heat-transfer predictions in this area. Measurements on the inner end wall show that temperature reduction in the vane wake due to the trailing edge cooling is important to many points downstream of the vane.

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Fig. 1

Location of inner end wall heat-flux gauges

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Fig. 2

Inlet temperature profile shapes for cooling comparison runs

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Fig. 3

Spanwise averaged Stanton number distribution

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Fig. 4

Net Stanton reduction due to cooling variation for airfoil surface

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Fig. 5

Stanton number distribution for inner end wall for (a) no-cooling and (b) nominal cooling case (geometry not to scale)

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Fig. 6

Net Stanton reduction for nominal cooling versus the (a) low-cooling and (b) high-cooling cases (geometry not to scale)

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Fig. 7

Stanton Number reduction for suction surface plotted by X/D relative to (a) no cooling and (b) nominal cooling

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Fig. 8

Comparison of heat flux and Stanton number for vane outer coolant modulation

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Fig. 9

Influence of vane trailing edge cooling on inner end wall Stanton number




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