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Manuel Wilhelm
Aerothermal Impact of Low Emission Combustion on the Turbine Blade Tip
A key technology to meet future gas turbine emission requirements has been identified in the concept of low emission, lean-burn combustion. In comparison to conventional combustion technology, lean-burn alters the flow and temperature field at the inlet of the high pressure turbine, with consequences for turbine aerodynamics and thermodynamics.
Within this work, the aerothermal impact of lean-burn inlet conditions on the rotor tip region is experimentally investigated at a large scale turbine rig. At the turbine inlet, a nonreactive combustor simulator creates a lean-burn representative flow field. The rotor tip geometry includes a recessed squealer-type cavity for leakage reduction and a pressure side row of cylindrical film cooling holes. A variety of different experimental techniques is applied, including steady pressure and temperature measurements, five-hole probes, hot-wire anemometry in conjunction with split-fiber probes, time-resolved rotor casing pressure measurements, and pressure sensitive paint. Results are presented for both a clean annulus, axial inflow baseline case, but also for two swirler-to-stator clocking positions, as well as for different rotor tip coolant injection ratios.
It is found that lean-burn inlet conditions cause a mass flow redistribution towards the hub and casing endwalls, an increase in turbulence, and an incidence at the rotor tip. This affects the blade loading and leakage, which in turn impacts the formation of secondary flows. The turbine efficiency is found to decline, losses increase, especially so in the turbine rotor, and cooling coverage is deteriorated at the tip trailing edge. Robustness of the rotor tip towards lean-burn turbine inlet conditions can be improved in the future by decoupling tip leakage from tip loading.
Within this work, the aerothermal impact of lean-burn inlet conditions on the rotor tip region is experimentally investigated at a large scale turbine rig. At the turbine inlet, a nonreactive combustor simulator creates a lean-burn representative flow field. The rotor tip geometry includes a recessed squealer-type cavity for leakage reduction and a pressure side row of cylindrical film cooling holes. A variety of different experimental techniques is applied, including steady pressure and temperature measurements, five-hole probes, hot-wire anemometry in conjunction with split-fiber probes, time-resolved rotor casing pressure measurements, and pressure sensitive paint. Results are presented for both a clean annulus, axial inflow baseline case, but also for two swirler-to-stator clocking positions, as well as for different rotor tip coolant injection ratios.
It is found that lean-burn inlet conditions cause a mass flow redistribution towards the hub and casing endwalls, an increase in turbulence, and an incidence at the rotor tip. This affects the blade loading and leakage, which in turn impacts the formation of secondary flows. The turbine efficiency is found to decline, losses increase, especially so in the turbine rotor, and cooling coverage is deteriorated at the tip trailing edge. Robustness of the rotor tip towards lean-burn turbine inlet conditions can be improved in the future by decoupling tip leakage from tip loading.
Schlagwörter: Gas Turbine; Film Cooling; Swirl; Combustor-Turbine-Interaction; Pressure Sensitive Paint
Forschungsberichte aus dem Institut für Gasturbinen, Luft- und Raumfahrtantriebe
Herausgegeben von Prof. Dr.-Ing. H.-P. Schiffer, Darmstadt
Band 15
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