| dc.description.abstract | Researchers in the field of gas turbines are always interested in gas turbine vane and blade heat transfer and film cooling performance under realistic engine flow conditions. Complicated flow patterns are present around the three-dimensional airfoils. Therefore, it is necessary to carry out testing on a three-dimensional vane and blade in a wind tunnel that closely simulates the real engine flow. Both annular and linear cascades are used to study the film cooling effectiveness on vane surfaces as well as on the blade tip and trailing edge using the pressure sensitive paint (PSP) measurement technique.
The first study is an experimental study of film cooling effectiveness on two real scale turbine vanes in a 3-vane annular cascade. The cascade is connected to a high flow steady compressor to provide the mainstream flow. The inlet velocity is maintained at 35 m/s at the center line of the annular cascade. Two heavily cooled, real scale turbine vanes are tested with cooling holes on the pressure and suction surfaces. Vane 1 has 645 cooling holes distributed around the vane. Vane 2 has an additional two row of holes at the near leading edge SS of the vane. The MFR is varied from 3.12% to 4.82%. Increasing the MFR increases the film effectiveness. The introduction of additional rows of cooling holes resulted in re-distribution of coolant from the pressure surface to the suction surface.
The second study is an experimental study of film cooling effectiveness on the blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The free stream turbulence intensity is measured to be 13.5 % upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a 2-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant ejection scenarios are considered by ejecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. A positive blowing ratio and density ratio effect is witnessed. Introduction of coolant injection also reduces the over-tip flow leakage.
The third study measures the film cooling effectiveness along the trailing edge of a turbine blade. The film effectiveness is measured and analyzed using the pressure sensitive paint (PSP) technique. Two different trailing edge designs are investigated including the standard pressure side cutback and the new alternating discharge design (referred to as a wavy trailing edge design). The alternating discharge design is a new design having a wavy like structure between the pressure and suction surfaces of the trailing edge. The new wavy structure helps discharge the coolant from the trailing edge in a manner so it alternates between the pressure and suction surfaces. Testing is carried out in a five-blade, linear, steady state cascade with inlet and exit Mach numbers of 0.20 and 0.30, respectively. The freestream turbulence intensity is measured to be 10.3% upstream of the blade leading edge. Coolant-to-mainstream mass flow ratios (MFR) of 0.30%, 0.50%, 0.75%, 1.0%, 1.25% and coolant-to-mainstream density ratios (DR) of 1 .0, 1.5 and 2 .0 are examined. The total pressure loss coefficients are also acquired to compare the aerodynamic loss between the two trailing edge designs. The pressure loss coefficient is acquired by traversing a group of pitot static probes across the blade span in a plane downstream of the trailing edge, resulting in a pressure map at the exit plane. A positive MFR and density ratio effect is witnessed with negligible difference in the aerodynamic loss. | |
