Recent Advances in the Study of Film Cooling on the Gas Turbine Blade

2014 ◽  
Vol 971-973 ◽  
pp. 143-147 ◽  
Author(s):  
Ping Dai ◽  
Shuang Xiu Li
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Performance ◽  
Leading Edge ◽  
Development Trend ◽  
Research Progress ◽  
Gas Turbine Blade ◽  
New Generation

The development of a new generation of high performance gas turbine engines requires gas turbines to be operated at very high inlet temperatures, which are much higher than the allowable metal temperatures. Consequently, this necessitates the need for advanced cooling techniques. Among the numerous cooling technologies, the film cooling technology has superior advantages and relatively favorable application prospect. The recent research progress of film cooling techniques for gas turbine blade is reviewed and basic principle of film cooling is also illustrated. Progress on rotor blade and stationary blade of film cooling are introduced. Film cooling development of leading-edge was also generalized. Effect of various factor on cooling effectiveness and effect of the shape of the injection holes on plate film cooling are discussed. In addition, with respect to progress of discharge coefficient is presented. In the last, the future development trend and future investigation direction of film cooling are prospected.

Application of local indentations for film cooling of gas turbine blade leading edge

2016 ◽  
Vol 23 (5) ◽  
pp. 713-720
Author(s):  
V. Yu. Petelchyts ◽  
A. A. Khalatov ◽  
D. N. Pysmennyi ◽  
Yu. Ya. Dashevskyy
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Gas Turbine Blade ◽  
Blade Leading Edge

Film cooling modeling of a gas turbine blade by considering different injection holes with and without opening angles through CFD

2021 ◽  
Vol 15 (1) ◽  
pp. 7637-7647
Author(s):  
E. Hosseini
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
High Performance ◽  
Turbine Blades ◽  
Flow Characteristics ◽  
Opening Angle ◽  
Inlet Temperature ◽  
Cooling Efficiency ◽  
Gas Turbine Blade

One way to achieve high performance in the gas turbine is to increase the inlet temperature of the turbine. Different cooling techniques have been carried out in order to protect the turbine blades which have been exposed to such high temperatures. Film cooling as an essential cooling method needs to be enhanced to meet the challenging demand. The purpose of the present research is to analyze the film cooling performance over a NACA 0012 gas turbine blade using six different injection holes with and without opening angles, separately through Computational Fluid Dynamics (CFD). 2D Reynolds-Averaged Navier-Stokes (RANS) equations are implemented to consider the heat transfer and flow characteristics by using CFD code Ansys Fluent v16. The flow is considered as steady, turbulent, and incompressible. The RANS equation is solved with the finite-volume method for obtaining solutions. The simulation results revealed that the k-ω SST turbulence model is suitable for simulating the flow characteristics and analyzing the performance of film cooling over the blade. Also, the opening angle has a significant effect on increasing the cooling efficiency for the upper blade surface. The highest value of cooling efficiency is obtained by the injection hole with an opening angle of 15° and height of D. In this configuration, the coolant injected from hole provides better cooling coverage for the entire blade which increases the cooling effectiveness.

Convective Heat Transfer Through Film Cooling Holes of a Gas Turbine Blade Leading Edge

2005 ◽  
Author(s):  
Elon J. Terrell ◽  
Brian D. Mouzon ◽  
David G. Bogard
Keyword(s):  
Heat Transfer ◽  
Convective Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Convective Heat ◽  
Film Cooling ◽  
Leading Edge ◽  
Gas Turbine Blade ◽  
Film Cooling Holes ◽  
Blade Leading Edge

Studies of film cooling performance for a turbine airfoil predominately focus on the reduction of heat transfer to the external surface of the airfoil. However, convective cooling of the airfoil due to coolant flow through the film cooling holes is potentially a major contributor to the overall cooling of the airfoil. This study used experimental and computational methods to examine the convective heat transfer to the coolant as it traveled through the film cooling holes of a gas turbine blade leading edge. Experimental measurements were conducted on a model gas turbine blade leading edge composed of alumina ceramic which approximately matched the Biot number of an engine airfoil leading edge. The temperature rise in the coolant from the entrance to the exit of the film cooling holes was measured using a series of internal thermocouples and an external traversing thermocouple probe. A CFD simulation of the model of the leading edge was also done in order to facilitate the processing of the experimental data and provide a comparison for the experimental coolant hole heat transfer. Without impingement cooling, the coolant hole heat transfer was found to account for 50 to 80 percent of the airfoil internal cooling, i.e. the dominating cooling mechanism.

Prediction of Film Cooling Effectiveness on a Gas Turbine Blade Leading Edge Using ANN and CFD

2018 ◽  
Vol 35 (2) ◽  
pp. 101-111 ◽  
Author(s):  
J. O. Dávalos ◽  
J. C. García ◽  
G. Urquiza ◽  
A. Huicochea ◽  
O. De Santiago
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Hole Diameter ◽  
Relative Importance ◽  
Gas Turbine Blade ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio

Abstract In this work, the area-averaged film cooling effectiveness (AAFCE) on a gas turbine blade leading edge was predicted by employing an artificial neural network (ANN) using as input variables: hole diameter, injection angle, blowing ratio, hole and columns pitch. The database used to train the network was built using computational fluid dynamics (CFD) based on a two level full factorial design of experiments. The CFD numerical model was validated with an experimental rig, where a first stage blade of a gas turbine was represented by a cylindrical specimen. The ANN architecture was composed of three layers with four neurons in hidden layer and Levenberg-Marquardt was selected as ANN optimization algorithm. The AAFCE was successfully predicted by the ANN with a regression coefficient R2<0.99 and a root mean square error RMSE=0.0038. The ANN weight coefficients were used to estimate the relative importance of the input parameters. Blowing ratio was the most influential parameter with relative importance of 40.36 % followed by hole diameter. Additionally, by using the ANN model, the relationship between input parameters was analyzed.

Film cooling optimization on leading edge gas turbine blade using differential evolution

2018 ◽  
Vol 233 (5) ◽  
pp. 1656-1666 ◽  
Author(s):  
Juan C García ◽  
José O Dávalos ◽  
Gustavo Urquiza ◽  
Sergio Galván ◽  
Alberto Ochoa ◽  
...  
Keyword(s):  
Differential Evolution ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Gas Turbine Blade ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Input Parameters ◽  
Artificial Neural

This article reports the optimization of film cooling on a leading edge of a gas turbine blade model, with showerhead configuration, it is based on five input parameters, which are hole diameter, hole pitch, column holes pitch, injection angle, and velocity at plenum inlet. This optimization increased the Area-Averaged Film Cooling Effectiveness [Formula: see text] and reduced the consumption of coolant flow. Differential Evolution assisted by artificial neural networks was used as optimization algorithm. Reynolds Averaged Navier–Stokes computations were carried out to getting the net database and to evaluate the optimized models predicted by artificial neural network. The results show an effective increment of [Formula: see text] by 36% and a mass flow reduction by 66%. These results were reached by means of a better distribution of cooling flow at blade surface as function of the input parameters. To assure the reliability of the numerical model, particle image velocimetry technique was used for its validation.

Novel Technology for Gas Turbine Blade Effusion Cooling

2006 ◽  
Author(s):  
Lorenzo Battisti ◽  
Roberto Fedrizzi ◽  
Giovanni Cerri
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Cooling System ◽  
Turbine Blades ◽  
Innovative Technology ◽  
Gas Turbine Blade ◽  
Combustion Chambers ◽  
Reliable Technique

Gas turbine combustion chambers and turbine blades require better cooling techniques to cope with the increase in operating temperatures with each new engine model. Current gas turbine inlet temperatures are approaching 2000 K. Such extreme temperatures, combined with a highly dynamic environment, result in major stress on components, especially combustion chamber and blades of the first turbine stages. A technique that has been extensively investigated is transpiration cooling, for both combustion chambers and turbine blades. Transpiration-cooled components have proved an effective way to achieve high temperatures and erosion resistance for gas turbines operating in aggressive environments, though there is a shortage of durable and proven technical solutions. Effusion cooling (full-coverage discrete hole film cooling), on the other hand, is a relatively simpler and more reliable technique offering a continuous coverage of cooling air over the component’s hot surfaces. This paper presents an innovative technology for the efficient effusion cooling of the combustor wall and turbine blades. The dedicated electroforming process used to manufacture effusive film cooling systems, called Poroform®, is illustrated. A numerical model is also presented, developed specifically for designing the distributions of the diameter and density of the holes on the cooled surface with a view to reducing the metal’s working temperature and achieving isothermal conditions for large blade areas. Numerical simulations were used to design the effusive cooling system for a first-stage gas turbine blade. The diameter, density and spacing of the holes, and the adiabatic film efficiency are discussed extensively to highlight the cooling capacity of the effusive system.

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