A detailed aero performance measurement program utilizing fully cooled engine hardware (high-pressure turbine stage) supplied by Honeywell Aerospace Advanced Technology Engines is described. The primary focus of this work was obtaining relevant aerodynamic data for a small turbine stage operating at a variety of conditions, including changes in operating conditions, geometry, and cooling parameters. The work extraction and the overall stage performance for each of these conditions can be determined using the measured acceleration rate of the turbine disk, the previously measured moment of inertia of the rotating system, and the mass flow through the turbine stage. Measurements were performed for two different values of tip/shroud clearance and two different blade tip configurations. The vane and blade cooling mass flow could be adjusted independently and set to any desired value, including totally off. A wide range of stage pressure ratios, coolant to free stream temperature ratios, and corrected speeds were used during the course of the investigation. A combustor emulator controlled the free stream inlet gas temperature, enabling variation of the temperature ratios and investigation of their effects on aero performance. The influence of the tip/shroud gap is clearly seen in this experiment. Improvements in specific work and efficiency achieved by reducing the tip/shroud clearance depend upon the specific values of stage pressure ratio and corrected speed. The maximum change of 3%–4% occurs at a stage pressure ratio and corrected speed greater than the initial design point intent. The specific work extraction and efficiency for two different blade tip sets (one damaged from a rub and one original) were compared in detail. In general, the tip damage only had a very small effect on the work extraction for comparable conditions. The specific work extraction and efficiency were influenced by the presence of cooling gas and by the temperature of the cooling gas relative to the free stream gas temperature and the metal temperature. These same parameters were influenced by the magnitude of the vane inlet gas total temperature relative to the vane metal temperature and the coolant gas temperature.
Aeroperformance Measurements for a Fully Cooled High-Pressure Turbine Stage
Contributed by the International Gas Turbine Institute (IGTI) of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received July 5, 2012; final manuscript received May 23, 2013; published online September 26, 2013. Editor: David Wisler.
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Haldeman, C., Dunn, M., Mathison, R., Troha, W., Vander Hoek, T., and Riahi, A. (September 26, 2013). "Aeroperformance Measurements for a Fully Cooled High-Pressure Turbine Stage." ASME. J. Turbomach. March 2014; 136(3): 031001. https://doi.org/10.1115/1.4024777
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