Detailed experimental investigations have been performed to measure the heat transfer and static pressure distributions on the rotor tip and rotor casing of a gas turbine stage with a shroudless rotor blade. The turbine stage was a modern high pressure Rolls-Royce aero-engine design with stage pressure ratio of 3.2 and nozzle guide vane (ngv) Reynolds number of 2.54E6. Measurements have been taken with and without inlet temperature distortion to the stage. The measurements were taken in the QinetiQ Isentropic Light Piston Facility and aerodynamic and heat transfer measurements are presented from the rotor tip and casing region. A simple two-dimensional model is presented to estimate the heat transfer rate to the rotor tip and casing region as a function of Reynolds number along the gap.
An Investigation on Turbine Tip and Shroud Heat Transfer
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 December 12, 2001. Paper No. 2002-GT-30554. Review Chair: E. Benvenuti.
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Chana, K. S., and Jones, T. V. (August 27, 2003). "An Investigation on Turbine Tip and Shroud Heat Transfer ." ASME. J. Turbomach. July 2003; 125(3): 513–520. https://doi.org/10.1115/1.1575253
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