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

Xing Yang, Zhenping Feng, Terrence W. Simon

Research output: Contribution to journalArticlepeer-review

1 Scopus citations

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.

Original languageEnglish (US)
Pages (from-to)1959-1981
Number of pages23
JournalAeronautical Journal
Volume123
Issue number1270
DOIs
StatePublished - Dec 1 2019

Bibliographical note

Funding Information:
This work was primarily supported by the National Natural Science Foundation of China (Grant No. 51336007 and No. 51876156) and China Postdoctoral Science Foundation (Grant No. 2019M653621). The authors would also like to acknowledge Minnesota Supercomputing Institute (MSI) for providing computational resources to conduct this study.

Publisher Copyright:
© Royal Aeronautical Society 2019.

Keywords

  • Aero-engine turbine endwall
  • Conjugate heat transfer
  • Film cooling
  • Impingement heat transfer
  • Numerical simulations
  • Thermal barrier coating

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