At the large scale turbine rig (LSTR) at Technische Universität Darmstadt, Darmstadt, Germany, the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with an axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position toward the nozzle guide vane (NGV) leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence, and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 35% and an increase of Nusselt numbers of 10–20% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30% for swirling inflow.

Issue Section:
Research Papers
Keywords:
Computational fluid dynamics, Fan, Fluid dynamics, Heat transfer, Measurement techniques, Compressor, Turbine aerodynamic design, Film cooling, Heat transfer phenomena, Turbine components, Gas turbine engines
Topics:
Boundary layers, Combustion chambers, Coolants, Film cooling, Flow (Dynamics), Heat transfer, Inflow, Nozzle guide vanes, Turbines, Turbulence, Probes, Cooling, Swirling flow, Pressure

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