Experimental Modeling and Measurement of Film Cooled Turbine Tip Shroud Heat Transfer Performance PDF Download

Author: Louis William Flamm
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ISBN:
Category : Heat
Languages : en
Pages : 334

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Author: Christopher Douglas Eick
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ISBN:
Category : Gas-turbines
Languages : en
Pages : 150

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Author: Massachusetts Institute of Technology. Gas Turbine Laboratory
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Category : Gas-turbines
Languages : en
Pages : 29

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Author: Ali A. Ameri
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Category :
Languages : en
Pages : 14

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Author: Srinath Varadarajan Ekkad
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ISBN:
Category : Heat
Languages : en
Pages : 248

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Author: Brett James Stanewich
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ISBN:
Category : Gas-turbines
Languages : en
Pages : 266

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Author: Jacob Ward Harral
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ISBN:
Category : Gas-turbine industry
Languages : en
Pages : 248

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Abstract: Increasing the efficiency in gas turbine engines requires constant improvement in the design tools currently available to the industry. One area for potential increases in efficiency deals with the film-cooling effectiveness in the high-pressure turbine section of the engine and the push to increase the temperature at the inlet of the turbine. Modeling of film-cooling effectiveness for incorporation into advanced CFD codes to be used for film effectiveness predictions and subsequent design of advanced engines is currently a major activity within the engine community. For the codes to be implemented as design tools one must gain confidence in their validity. One method that has been used for this purpose is to compare predictions obtained using these codes with experimental results obtained under as realistic conditions as is possible within the confines of controlled laboratory experiments. Under support of the NASA/DoD URETI, the OSU GTL has undertaken the task of performing detailed surface-pressure and surface heat-transfer measurements on the vane surfaces, on the blade surfaces, and on the stationary shroud of a fully cooled high-pressure turbine stage operating at design corrected conditions. Several significant changes have been made to the OSU Gas Turbine Laboratory blowdown turbine facility and to the operating mode of that facility in order to make film effectiveness measurements. One of the major facility changes was the incorporation of a coolant gas supply system (LCF) into the facility. The major changes in operating mode involved operating in blowdown mode instead of shock tube mode. In order to achieve this major change in operating procedure, it was necessary to incorporate a resistance heater into the rig just ahead of the high-pressure turbine vane inlet so that a resistance heater instead of the reflected shock could heat the test gas. The next major task was to sequence the main test gas flow with the coolant gas flow so that one could achieve the proper flow physics. This thesis will focus on the operation and integration of the LCF into the blowdown facility and on the experimental results acquired during the initial film-cooling experiment. Operation of the LCF is divided into three distinct areas: fast acting valve operation and sequencing with the main facility fast acting valve, cooling cycles, and facility controls. Successful integration of the LCF has been achieved and will be illustrated by the results of the initial film-cooling experiment. Through these experimental results and accompanying uncertainty analysis conducted as part of this thesis significant knowledge has been gained and will be applied to future film-cooling measurement programs. With the demonstrated successful operation of the OSU turbine test facility in conjunction with the LCF, the OSU GTL is capable of conducting the critical experiments necessary to provide critical verification information for ongoing film effectiveness modeling and CFD code development.

Author: Diganta P. Narzary
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ISBN:
Category :
Languages : en
Pages :

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Modern gas turbine engines require higher turbine-entry gas temperature to improve their thermal efficiency and thereby their performance. A major accompanying concern is the heat-up of the turbine components which are already subject to high thermal and mechanical stresses. This heat-up can be reduced by: (i) applying thermal barrier coating (TBC) on the surface, and (ii) providing coolant to the surface by injecting secondary air discharged from the compressor. However, as the bleeding off of compressor discharge air exacts a penalty on engine performance, the cooling functions must be accomplished with the smallest possible secondary air injection. This necessitates a detailed and systematic study of the various flow and geometrical parameters that may have a bearing on the cooling pattern. In the present study, experiments were performed in three regions of a non-rotating gas turbine blade cascade: blade platform, blade span, and blade tip. The blade platform and blade span studies were carried out on a high pressure turbine rotor blade cascade in medium flow conditions. Film-cooling effectiveness or degree of cooling was assessed in terms of cooling hole geometry, blowing ratio, freestream turbulence, coolant-to-mainstream density ratio, purge flow rate, upstream vortex for blade platform cooling and blowing ratio, and upstream vortex for blade span cooling. The blade tip study was performed in a blow-down flow loop in a transonic flow environment. The degree of cooling was assessed in terms of blowing ratio and tip clearance. Limited heat transfer coefficient measurements were also carried out. Mainstream pressure loss was also measured for blade platform and blade tip film-cooling with the help of pitot-static probes. The pressure sensitive paint (PSP) and temperature sensitive paint (TSP) techniques were used for measuring film-cooling effectiveness whereas for heat transfer coefficient measurement, temperature sensitive paint (TSP) technique was employed. Results indicated that the blade platform cooling requires a combination of upstream purge flow and downstream discrete film-cooling holes to cool the entire platform. The shaped cooling holes provided wider film coverage and higher film-cooling effectiveness than the cylindrical holes while also creating lesser mainstream pressure losses. Higher coolant-to-mainstream density ratio resulted in higher effectiveness levels from the cooling holes. On the blade span, at any given blowing ratio, the suction side showed better coolant coverage than the pressure side even though the former had two fewer rows of holes. Film-cooling effectiveness increased with blowing ratio on both sides of the blade. Whereas the pressure side effectiveness continued to increase with blowing ratio, the increase in suction side effectiveness slowed down at higher blowing ratios (M=0.9 and 1.2). Upstream wake had a detrimental effect on film coverage. 0% and 25% wake phase positions significantly decreased film-cooling effectiveness magnitude. Comparison between the compound shaped hole and the compound cylindrical hole design showed higher effectiveness values for shaped holes on the suction side. The cylindrical holes performed marginally better in the curved portion of the pressure side. Finally, the concept tip proved to be better than the baseline tip in terms of reducing mainstream flow leakage and mainstream pressure loss. The film-cooling effectiveness on the concept blade increased with increasing blowing ratio and tip gap. However, the film-coverage on the leading tip portion was almost negligible.

Author:
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ISBN:
Category :
Languages : en
Pages : 0

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This report results from a contract tasking Imperial College of Science. Technology and Medicine as follows: The project will involve the quantification of the flow and surface characteristics of turbine blade tip cooling geometry with velocity and turbulence distributions relevant to the gas turbine designer. The experimental program will provide measurements of the film cooling adiabatic effectiveness and heat transfer coefficient in a simulated blade tip. with injection from the pressure surface side near the tip. The program will be progressive and interactive. starting with a single row of film cooling holes with injection from the pressure surface side near the tip. moving towards other representative injection configurations: slot injection on the tip surface and groove-tip geometry. The project will supply validation data for the 3D Navier-Stokes solver Glenn-HT. currently used by NASA. and it will also provide insight into the performance of advanced cooling geometry configurations. The experiments will make use of liquid crystal thermography to obtain the heat transfer data. The data acquisition method corresponds to the steady-state technique with the use of wide band liquid crystals. It requires a reduced number of experiments when compared with narrow band crystals and thermocouples, and provides a high degree of spatial resolution and reduced uncertainty level. It will be accompanied by data from small-diameter thermocouples. hot wires and pressure transducers.

Author: Jungho Choi
Publisher:
ISBN:
Category :
Languages : en
Pages :

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This experimental study contains two points; part 1 - turbine blade heat transfer under low Reynolds number flow conditions, and part 2 - trailing edge cooling and heat transfer. The effect of unsteady wake and free stream turbulence on heat transfer and pressure coefficients of a turbine blade was investigated in low Reynolds number flows. The experiments were performed on a five blade linear cascade in a low speed wind tunnel. A spoked wheel type wake generator and two different turbulence grids were employed to generate different levels of the Strouhal number and turbulence intensity, respectively. The cascade inlet Reynolds number based on blade chord length was varied from 15,700 to 105,000, and the Strouhal number was varied from 0 to 2.96 by changing the rotating wake passing frequency (rod speed) and cascade inlet velocity. A thin foil thermocouple instrumented blade was used to determine the surface heat transfer coefficient. A Liquid crystal technique based on hue value detection was used to measure the heat transfer coefficient on a trailing edge film cooling model and internal model of a gas turbine blade. It was also used to determine the film effectiveness on the trailing edge. For the internal model, Reynolds numbers based on the hydraulic diameter of the exit slot and exit velocity were 5,000, 10,000, 20,000, and 30,000 and corresponding coolant-to-mainstream velocity ratios were 0.3, 0.6, 1.2, and 1.8 for the external models, respectively. The experiments were performed at two different designs and each design has several different models such as staggered / inline exit, straight / tapered entrance, and smooth / rib entrance. The compressed air was used in coolant air. A circular turbulence grid was employed to upstream in the wind tunnel and square ribs were employed in the inlet chamber to generate turbulence intensity externally and internally, respectively.