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Research Papers

An Investigation of Coolant Within Serpentine Passages of a High-Pressure Axial Gas Turbine Blade

[+] Author and Article Information
Jeremy Nickol

Mechanical and Aerospace Engineering Department,
The Ohio State University,
2300 West Case Road,
Columbus, OH 43235-7531
e-mail: nickoljb@gmail.com

Randall Mathison

Mechanical and Aerospace Engineering Department,
The Ohio State University,
2300 West Case Road,
Columbus, OH 43235-7531
e-mail: mathison.4@osu.edu

Michael Dunn

Mechanical and Aerospace Engineering Department,
The Ohio State University,
2300 West Case Road,
Columbus, OH 43235-7531
e-mail: dunn.129@osu.edu

Jong Liu

Honeywell International,
111 S. 34th Street,
Phoenix, AZ 85034-2181
e-mail: jong.liu@honeywell.com

Malak Malak

Honeywell International,
111 S. 34th Street,
Phoenix, AZ 85034-2181
e-mail: malak.malak@honeywell.com

1Present address: Laboratory for Energy Conversion, ETH-Zurich, Zurich 8092, Switzerland.

2Corresponding author.

Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received January 26, 2017; final manuscript received February 15, 2017; published online April 11, 2017. Editor: Kenneth Hall.This research was partially funded by the Aviation Applied Technology Directorate under Agreement No. W911W6-08-2-0011. The U.S. Government is authorized to reproduce and distribute reprints for Government purposes notwithstanding any copyright notation thereon.

J. Turbomach 139(9), 091006 (Apr 11, 2017) (8 pages) Paper No: TURBO-17-1016; doi: 10.1115/1.4036109 History: Received January 26, 2017; Revised February 15, 2017

Cooling flow behavior is investigated within the multiple serpentine passages with turbulators on the leading and trailing walls of an axial gas turbine blade operating at design-corrected conditions with accurate external flow conditions. Pressure and temperature measurements at midspan within the passages are obtained using miniature butt-welded thermocouples and miniature Kulite pressure transducers. These measurements, as well as airfoil surface pressure field data from a full computational fluid dynamics (CFD) simulation, are used as boundary conditions for a model that provides quantitative values of film-cooling blowing ratio for each film-cooling hole on the blade. The model accounts for the continuously changing cross-sectional area and shape of the channels, frictional pressure loss, convective heat transfer from the solid portion of the blade, massflow reduction as coolant bleeds out through film-cooling or impingement holes, compressibility effects, and the effects of blade rotation. The results of the model provide detailed coolant ejection information for a film-cooled rotating turbine airfoil operating at design-corrected conditions and also account for the highly variable freestream conditions on the airfoil. While these values are commonly known for simpler experimental geometries, they have previously either been unknown or estimated crudely for full-stage experiments of this nature. The better-quantified cooling parameters provide a bridge for better comparison with the wealth of film-cooling work already reported for simplified geometries. The calculation also shows the significant range in blowing ratio that can arise even among a single row of cooling holes associated with one of the turbulated passages, due to significant changes in both coolant and local freestream massfluxes.

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References

Figures

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Fig. 1

Photo of the rotor blade suction and pressure sides for a blade with static pressure instrumentation

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Fig. 2

Schematic of internal passages

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Fig. 3

Stencil of numerical mesh

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Fig. 4

Full computational domain of the accompanying CFD model (not to scale)

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Fig. 5

Normalized internal pressure with data labeled for the 6.85% coolant case, arrows showing bulk flow direction for each passage

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Fig. 6

Global blowing ratio versus span for each of the seven cooling rows for 6.85% cooling

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Fig. 7

Normalized airfoil pressure at the ejection location of each cooling hole

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Fig. 8

Local blowing ratio versus span for each of the seven cooling rows for 6.85% cooling

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Fig. 9

Local freestream massflux at each cooling hole outlet, normalized by the stage-average massflux

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Fig. 10

Average local blowing ratio for each cooling row at various cooling flow rates

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Fig. 11

Average isentropic coolant ejection Mach number for each cooling row at various cooling flow rates

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Fig. 12

Fraction of maximum flow through a cooling hole as a function of the isentropic ejection Mach number

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