Numerical Simulations of Leading Edge Film Cooling Flows for Gas Tubine Airfoils PDF Download

Author: Cheryl A. Martin
Publisher:
ISBN:
Category :
Languages : en
Pages : 442

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Author: Xuezhi Zhang
Publisher:
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Category :
Languages : en
Pages : 0

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This work presents the performance of a louver film-cooling scheme under different operating conditions. The louver cooling scheme consists of a bend by which the coolant going through the flow passage is redirected from vertical to horizontal direction before being injected into the mainstream through an expanded exit. Not only is the momentum of the coolant converted to the mainstream direction, but it is also reduced by the expanded exit before injection. The impingement of the coolant on the blade surface inside the bend also enables further cooling on the targeted surface. The louver cooling scheme was tested under a variety of conditions, from a flat plate to airfoils, from low speed incompressible flows to transonic flows, from a stationary airfoil to a rotating airfoil, and from the leading edge to the middle of an airfoil. Unsteady analysis using a DES (Detached Eddy Simulation) model was also carried out to evaluate its ability to accurately simulate film cooling by comparing with steady state analysis. In general, the louver cooling scheme has been proved to provide enhanced cooling protection to the targeted surface in comparison with other cooling schemes in all conditions tested. At low speed incompressible flow conditions, a higher blowing ratio led to a higher cooling effectiveness. At transonic flow conditions, a moderately higher blowing ratio also proved helpful with a higher cooling effectiveness. Very high blowing ratios, however, proved to be detrimental to the cooling performance since strong detached shock wave structures due to high blowing ratios caused boundary layer separation, rendering the coolant virtually ineffective. The rotation of blade was found to have a significant impact on the level of cooling effectiveness at the leading edge of an airfoil. With regard to the cooling performance, blowing ratio was the dominant factor at low rotational speeds and the rotational speed was the dominant factor at high blowing ratios for circular holes. For the louver scheme as jet liftoff was avoided, effectiveness increased with rotating speed. Results also showed that, unsteady analysis was not significantly more accurate than steady analysis. The unsteady analysis did capture the coolant lateral spreading better, with a high cost of computing, however. Results in this work show that shock waves encountered on transonic airfoils had a significant impact on film cooling effectiveness on any shaped holes. Therefore, experimental data obtained under low speed test should be used with great caution in real design of turbine blade cooling. There are fundamental differences in film cooling between at the leading edge and elsewhere on an airfoil in that a slight incidence shifting due to turbine rotating speed may cause a sudden decrease in cooling effectiveness level at high blowing ratios for circular hole. This could lead to a catastrophic failure if the blade is already in a weak and stressed state. Using of shaped holes with expanded exits may prevent this from happening.

Author: A. Chernobrovkin
Publisher:
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Category :
Languages : en
Pages : 0

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Presented at the International Gas Turbine & Aeroengine Congress & Exhibition, Stockholm, Sweden, Jun 2-5, 1998.

Author: Bengt Sundén
Publisher: Witpress
ISBN:
Category : Medical
Languages : en
Pages : 544

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This title presents and reflects current active research on various heat transfer topics and related phenomena in gas turbine systems. It begins with a general introduction to gas turbine heat transfer, before moving on to specific areas.

Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781723439759
Category :
Languages : en
Pages : 178

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The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils. Hylton, L. D. and Nirmalan, V. and Sultanian, B. K. and Kaufman, R. M. Unspecified Center EQUIPMENT SPECIFICATIONS; FILM COOLING; HEAT TRANSFER; LEADING EDGES; STRUCTURAL DESIGN; VANES; AIRCRAFT ENGINES; CASCADE FLOW; DATA PROCESSING; GAS TURBINES; HIGH TEMPERATURE; PARAMETERIZATION; TWO DIMENSIONAL FLOW...

Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 34

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Author: Jason B. Blitz
Publisher:
ISBN:
Category : Gas-turbines
Languages : en
Pages : 384

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Author:
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Category :
Languages : en
Pages : 14

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Author: Michael W. Cruse
Publisher:
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Category :
Languages : en
Pages : 0

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Presented at the International Gas Turbine & Aeroengine Congress & Exhibition, Orlando, FL, Jun 2 - Jun 5, 1997.

Author: Chaitanya D Ghodke
Publisher: SAE International
ISBN: 0768095026
Category : Technology & Engineering
Languages : en
Pages : 238

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Gas turbines play an extremely important role in fulfilling a variety of power needs and are mainly used for power generation and propulsion applications. The performance and efficiency of gas turbine engines are to a large extent dependent on turbine rotor inlet temperatures: typically, the hotter the better. In gas turbines, the combustion temperature and the fuel efficiency are limited by the heat transfer properties of the turbine blades. However, in pushing the limits of hot gas temperatures while preventing the melting of blade components in high-pressure turbines, the use of effective cooling technologies is critical. Increasing the turbine inlet temperature also increases heat transferred to the turbine blade, and it is possible that the operating temperature could reach far above permissible metal temperature. In such cases, insufficient cooling of turbine blades results in excessive thermal stress on the blades causing premature blade failure. This may bring hazards to the engine's safe operation. Gas Turbine Blade Cooling, edited by Dr. Chaitanya D. Ghodke, offers 10 handpicked SAE International's technical papers, which identify key aspects of turbine blade cooling and help readers understand how this process can improve the performance of turbine hardware.