Effect of a reduction in blade thickness on performance of a single-stage 20.32-centimeter mean-diameter turbine
National Aeronautics and Space Administration, 1973 - Technology & Engineering - 24 pages
As part of a program to reduce the manufacturing costs of a small gas-turbine engine, the turbine blading was reduced in thickness to facilitate coining. Tests were made to determine the effect of this modification on turbine performance. The working fluid was air at nominal inlet total conditions of 535 deg F and 20.0 psia. Performance results are presented and compared for four stator-rotor combinations in terms of equivalent torque, mass flow, and efficiency at equivalent design speed and at inlet-total to exit-static pressure ratios of 1.8 to 3.8
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aerodynamic parameters axial chord Axial-Flow blade rows blade thickness change in blade conditions of 297.2 configurations investigated Design value dynamometer equivalent design speed equivalent mass flow equivalent torque exit-static pressure ratios flow angle four configurations inlet total conditions inlet-total to exit-static Kofskey Lewis Research Center m/sec mean diameter mean section Milton G NASA NASA TN nominal inlet total original and thin original blade original rotor blade Original stator outer walls percent performance results pressure-surface diffusion Pressure-surface velocities rotor configurations rotor exit shown in figure shows SINGLE-STAGE static efficiency static-pressure ratio station stator and rotor stator-exit static pressure substantial reduction suction-surface diffusion taps equally spaced thermocouples thickness from leading thin blade profiles thin rotor blade thin stator blades three blade sections tip clearance tip sections total efficiency total pressure trailing edge Transonic Turbine Efficiency turbine performance turbine-inlet turbofan U.S. standard sea-level Variation velocity diagrams work-speed parameter YL Yu YL YU YL Yu