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Conjugate heat transfer modeling of a turbine vane endwall with thermal barrier coatings

Published online by Cambridge University Press:  19 July 2019

Xing Yang Open the ORCID record for Xing Yang [Opens in a new window] ,
Zhenping Feng Open the ORCID record for Zhenping Feng [Opens in a new window]  and
Terrence W. Simon
Xing Yang
Affiliation:
Xi’an Jiaotong UniversityShaanxi Engineering Laboratory of Turbomachinery and Power Equipment Institute of Turbomachinery, School of Energy and Power Engineering Xi’an, Shaanxi China
Zhenping Feng
Affiliation:
Xi’an Jiaotong UniversityShaanxi Engineering Laboratory of Turbomachinery and Power Equipment Institute of Turbomachinery, School of Energy and Power Engineering Xi’an, Shaanxi China
Terrence W. Simon
Affiliation:
University of Minnesota Department of Mechanical Engineering Minneapolis, MN USA
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Abstract

Advanced cooling techniques involving internal enhanced heat transfer and external film cooling and thermal barrier coatings (TBCs) are employed for gas turbine hot components to reduce metal temperatures and to extend their lifetime. A deeper understanding of the interaction mechanism of these thermal protection methods and the conjugate thermal behaviours of the turbine parts provides valuable guideline for the design stage. In this study, a conjugate heat transfer model of a turbine vane endwall with internal impingement and external film cooling is constructed to document the effects of TBCs on the overall cooling effectiveness using numerical simulations. Experiments on the same model with no TBCs are performed to validate the computational methods. Round and crater holes due to the inclusion of TBCs are investigated as well to address how film-cooling configurations affect the aero-thermal performance of the endwall. Results show that the TBCs have a profound effect in reducing the endwall metal temperatures for both cases. The TBC thermal protection for the endwall is shown to be more significant than the effect of increasing coolant mass flow rate. Although the crater holes have better film cooling performance than the traditional round holes, a slight decrement of overall cooling effectiveness is found for the crater configuration due to more endwall metal surfaces directly exposed to external mainstream flows. Energy loss coefficients at the vane passage exit show a relevant negative impact of adding TBCs on the cascade aerodynamic performance, particularly for the round hole case.

Keywords

Aero-engine turbine endwallConjugate heat transferImpingement heat transferFilm coolingThermal barrier coatingNumerical simulations
Type
Research Article
Information
The Aeronautical Journal , Volume 123 , Issue 1270 , December 2019 , pp. 1959 - 1981
Copyright
© Royal Aeronautical Society 2019 

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