A Study of Film Cooling Using Angled Slots with Impinging Feed Holes on the Suction Side of a First Stage Turbine Vane PDF Download
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Author: Frederick Todd Davidson
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Author: Frederick Todd Davidson
Publisher:
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Category :
Languages : en
Pages :
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Author: S. Stephen Papell
Publisher:
ISBN:
Category : Cooling
Languages : en
Pages : 32
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Author: James W. Guantner
Publisher:
ISBN:
Category : Turbines
Languages : en
Pages : 36
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Author: Isabella Gayoso
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ISBN:
Category :
Languages : en
Pages : 0
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In this experiment, measurements of the overall cooling effectiveness for a film cooled turbine vane airfoil in a high-speed cascade were obtained using infrared thermography. The vane used was the NASA C3X with impingement holes (showerhead cooling) and convective cooling holes on both the suction and pressure side. This work was done in the Mechanical Engineering Department's Experimental and Computational Convection Lab and used the high-speed cascade capability of the lab. The rationale for conducting this work was to obtain experimental data on film cooling effectiveness in a turbine vane in engine-like conditions at transonic speeds. Previous work has been done at subsonic speeds, but few pieces of literature examine this parameter at transonic speeds. The data can then be used to validate or compare to CFD models and to better understand what happens to the vane temperature distribution during engine operation. This understanding could inform the design of film cooling holes to reduce thermal strain "hot spots" which lead to failure of the vane. The results showed that trends for values of overall film effectiveness were as expected in this experiment, such as increases in blowing ratio correlating to increases in overall film effectiveness. However, the blowing ratios used in this study were not as high as values studied previously, indicating a need for more data on overall film effectiveness at transonic speeds.
Author: Herbert J. Gladden
Publisher:
ISBN:
Category : Turbines
Languages : en
Pages : 36
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Author: James W. Gauntner
Publisher:
ISBN:
Category : Turbines
Languages : en
Pages : 36
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Author: Marcia Inez Ethridge
Publisher:
ISBN:
Category :
Languages : en
Pages : 250
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Author:
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Category :
Languages : en
Pages : 0
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An experimental study was conducted in a simulated three vane linear cascade to determine the effects of surface roughness and film cooling on the heat transfer coefficient distribution in the region downstream of the first row of suction side coolant holes. Suction side film cooling was operated in the range 0 less than M less than 1.4. The showerhead was tested at M(sub sh) = 1.6. In addition to the completely smooth condition, simulated airfoil roughness was used upstream of the coolant holes, downstream of the coolant holes, and both upstream and downstream of the coolant holes. Two levels of mainstream turbulence intensity were tested. The heat transfer measurements were conducted by application of a uniform heat flux in the region downstream of the coolant holes. The resulting surface temperature distributions were measured with infrared thermography. Because the upstream region was unheated, the influence of film cooling on the heat transfer coefficient was due to only to hydrodynamic effects and not thermal effects. The coolant to mainstream density ratio of the majority of the experiments was unity; however, a single experiment was conducted at a density ratio of DR = 1.6 to determine how the coolant to mainstream density ratio affects heat transfer. Net heat flux reduction calculations were performed by combining the heat transfer coefficient measurements of the present study with adiabatic effectiveness measurements of a separate study. In order to gain insight into the hydrodynamics that affect the heat transfer, boundary layer measurements were conducted using hot-wire anemometry.
Author: Christopher Yoon
Publisher:
ISBN:
Category :
Languages : en
Pages : 126
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Film cooling is often used for turbine airfoil cooling, and there are numerous studies of the performance of a single row of holes. Typically, blades and vanes in gas turbine engines have multiple rows of holes that interact. Consequently, there is a need to develop techniques to predict film cooling performance with multiple rows of holes. One of the method is to superposition single row cooling effectiveness to predict combined effectiveness. Although there have been many studies of superposition techniques with multiple rows of cylindrical holes, there have been very few in which shaped holes were used with a typical turbine airfoil model. In this study, film effectiveness was measured on the suction side of a turbine blade model using two rows of 7-7-7 shaped holes, with pitch to diameter ratio of 6, and the two rows were more than 40 diameters apart. Measurements were made with each row operating independently, which provided the experimental data for superposition predictions. These predictions were evaluated with effectiveness measurements with both rows operational. For these combined row tests, two different upstream blowing ratios and a wide range of downstream blowing ratios were selected. The superposition predictions were reasonably accurate when the upstream blowing ratio was high with a corresponding smaller film effectiveness downstream (due to jet separation). However, when the upstream coolant holes were operated at optimum blowing ratio with maximum film effectiveness downstream, the superposition analysis predicted film effectiveness levels slightly lower than actual levels. These results indicate that there was an interaction between jets that resulted in higher film effectiveness than what the superposition method had predicted
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 231
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Book Description
An experimental study was conducted in a simulated three vane linear cascade to determine the effects of surface roughness and film cooling on the heat transfer coefficient distribution in the region downstream of the first row of suction side coolant holes. Suction side film cooling was operated in the range 0 less than M less than 1.4. The showerhead was tested at M(sub sh) = 1.6. In addition to the completely smooth condition, simulated airfoil roughness was used upstream of the coolant holes, downstream of the coolant holes, and both upstream and downstream of the coolant holes. Two levels of mainstream turbulence intensity were tested. The heat transfer measurements were conducted by application of a uniform heat flux in the region downstream of the coolant holes. The resulting surface temperature distributions were measured with infrared thermography. Because the upstream region was unheated, the influence of film cooling on the heat transfer coefficient was due to only to hydrodynamic effects and not thermal effects. The coolant to mainstream density ratio of the majority of the experiments was unity; however, a single experiment was conducted at a density ratio of DR = 1.6 to determine how the coolant to mainstream density ratio affects heat transfer. Net heat flux reduction calculations were performed by combining the heat transfer coefficient measurements of the present study with adiabatic effectiveness measurements of a separate study. In order to gain insight into the hydrodynamics that affect the heat transfer, boundary layer measurements were conducted using hot-wire anemometry.
