Abstract
The present study investigated the cooling performance of the second stage turbine nozzle vane using a transonic linear turbine cascade test facility. For experiments, 500 hp or 2,250 hp compressor and electric heaters were used for mainstream supply and 50 hp compressor for secondary flow supply and a single nozzle vane with adjusted sidewalls on each side of the nozzle was installed in the test section. The test nozzle model having complex cooling configuration was made of Inconel 718 using metal 3D printing technique. The overall cooling effectiveness of the cooled turbine nozzle vane was measured using two Infrared cameras. Four different sets of experiments were conducted under two different mainstream conditions, subsonic and transonic conditions and the corresponding Reynolds numbers of the mainstream are 6.5x10SUP5/SUP and 1.3x10SUP6/SUP based on chord length, respectively. The results showed that overall cooling effectiveness on the nozzle vane surface is dominated by internal heat convection in a rib-roughened surface, film cooling flow around the leading edge region and heat conduction around the hole and pin-fins. When comparing the results at different inlet Reynolds numbers of the internal cooling passage, the internal cooling configuration has more dominant effect on overall cooling performance than film cooling. Therefore it is obvious that overall cooling effectiveness as well as film cooling effectiveness should be taken into account to assess exact cooling performance and characteristics on turbine nozzle vane cooling design.
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