Most turbine inlet flows resulting from the combustor exit are nonuniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identified from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingement locations along the stagnation line.
Combustor Turbine Interface Studies—Part 2: Flow and Thermal Field Measurements
Contributed by the International Gas Turbine Institute and presented at the International Gas Turbine and Aeroengine Congress and Exhibition, Amsterdam, The Netherlands, June 3–6, 2002. Manuscript received by the IGTI January 16, 2002. Paper No. 2002-GT-30527. Review Chair: E. Benvenuti.
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Colban , W. F., Lethander , A. T., Thole, K. A., and Zess, G. (April 23, 2003). "Combustor Turbine Interface Studies—Part 2: Flow and Thermal Field Measurements ." ASME. J. Turbomach. April 2003; 125(2): 203–209. https://doi.org/10.1115/1.1561812
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