Two-Dimensional Inlet Temperature Profile Attenuation in a Turbine Stage PUBLIC ACCESS

[+] Author Affiliations
Daniel J. Dorney, Roger L. Davis

United Technologies Research Center, East Hartford, CT

Om P. Sharma

Pratt & Whitney, East Hartford, CT

Paper No. 91-GT-406, pp. V001T01A104; 8 pages
  • ASME 1991 International Gas Turbine and Aeroengine Congress and Exposition
  • Volume 1: Turbomachinery
  • Orlando, Florida, USA, June 3–6, 1991
  • Conference Sponsors: International Gas Turbine Institute
  • ISBN: 978-0-7918-7898-9
  • Copyright © 1991 by ASME


Experimental evidence has shown that hot streaks which are emitted from the combustors of gas turbines are often largely responsible for the burning of first stage turbine blades. Designers have attempted to counteract the effects of these hot streaks through the use of complex internal and film cooling schemes. Unfortunately, due to the lack of accurate predictive tools which account for temperature non-uniformities in the gas path, as well as a lack of detailed understanding of the physical mechanisms which control the migration and accumulation of the hot streak gases, turbine blade “hot spots” still occasionally occur. In an effort to increase understanding of the interaction mechanisms between combustor hot streaks and turbine blade heat transfer, a numerical investigation has been conducted to determine if a two-dimensional solution procedure can accurately predict rotor airfoil surface heating for flows which include planar hot streaks. A two-dimensional Navier-Stokes analysis is used to predict unsteady viscous flow through a 1-stator/1-rotor configuration with a planar hot streak introduced at the stator inlet. Comparison of the predicted results with a new experimental data set demonstrates that the two-dimensional numerical procedure can be used to accurately predict time-averaged rotor pressure surface temperatures for flows which include planar hot streaks.

Copyright © 1991 by ASME
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