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.

Novel Gas Turbine Blade Leading Edge Cooling Configuration Using Advanced Double Swirl Chambers

2015 ◽  
Author(s):  
Karsten Kusterer ◽  
Gang Lin ◽  
Takao Sugimoto ◽  
Dieter Bohn ◽  
Ryozo Tanaka ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Leading Edge ◽  
Gas Turbine Blade ◽  
Stagnation Line ◽  
Impingement Cooling ◽  
Cooling Technology ◽  
Cooling Air ◽  
Blade Leading Edge

The Double Swirl Chambers (DSC) cooling technology, which has been introduced and developed by the authors, has the potential to be a promising cooling technology for further increase of gas turbine inlet temperature and thus improvement of the thermal efficiency. The DSC cooling technology establishes a significant enhancement of the local internal heat transfer due to the generation of two anti-rotating swirls. The reattachment of the swirl flows with the maximum velocity at the center of the chamber leads to a linear impingement effect on the internal surface of the blade leading edge nearby the stagnation line of gas turbine blade. Due to the existence of two swirls both the suction side and the pressure side of the blade near the leading edge can be very well cooled. In this work, several advanced DSC cooling configurations with a row of cooling air inlet holes have been investigated. Compared with the standard DSC cooling configuration the advanced ones have more suitable cross section profiles, which enables better accordance with the real blade leading edge profile. At the same time these configurations are also easier to be manufactured in a real blade. These new cooling configurations have been numerically compared with the state of the art leading edge impingement cooling configuration. With the same configuration of cooling air supply and boundary conditions the advanced DSC cooling presents 22–26% improvement of overall heat transfer and 3–4% lower total pressure drop. Along the stagnation line the new cooling configuration can generate twice the heat flux than the standard impingement cooling channel. The influence of spent flow in the impinging position and impingement heat transfer value is in the new cooling configurations much smaller, which leads to a much more uniform heat transfer distribution along the chamber axial direction.

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

Effect of Film Cooling Arrangement on Impingement Heat Transfer on Turbine Blade Leading Edge

2013 ◽  
Author(s):  
Li Yang ◽  
Rui Kan ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Stagnation Zone ◽  
Intermediate Area ◽  
High Heat Transfer ◽  
Hole Position ◽  
Film Cooling Holes ◽  
Blade Leading Edge

Impingement cooling plays an important role in gas turbine blade leading edge where proper heat transfer distribution is needed for extremely high and nonuniform thermal load. A 2/3 cylinder leading edge model with 3 arrays of film cooling holes was investigated with 8 film cooling arrangements. The impingement parameters and the jet Reynolds number were kept the same for the 8 configurations. The transient liquid crystal (TLC) measurement was applied on heat transfer coefficient on the leading edge. A 3D numerical method with the SST k-ω model was verified by experimental data, which shows a heat transfer error less than 15%. The film suction creates both local heat transfer enhancement and limit effect to wall jets. The hole position of film cooling holes significantly affects the shape of high heat transfer area and cooling of the intermediate area. The array angle of film cooling holes affects the spread of heat transfer laterally. The Nu in stagnation zone decreases with the increase of array angle of film cooling holes. Smaller pitch of film cooling holes helps decrease the size of fountaining flow and heat transfer valley. The Nu in stagnation zone increases with the decrease of pitch of film cooling holes. The hole position of x0/P = 0.125 is recommended for the best cooling performance in the intermediate area. The configurations with θ = 13 or P/pf = 3 work best in this study.

Leading Edge Cooling of a Gas Turbine Blade With Double Swirl Chambers

2014 ◽  
Author(s):  
Karsten Kusterer ◽  
Gang Lin ◽  
Dieter Bohn ◽  
Takao Sugimoto ◽  
Ryozo Tanaka ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Maximum Velocity ◽  
Leading Edge ◽  
Inlet Temperature ◽  
Gas Turbine Blade ◽  
Impingement Cooling ◽  
Swirl Flows ◽  
Blade Leading Edge

The gas turbine blade leading edge area has locally extremely high thermal loads, which restrict the further increase of turbine inlet temperature or the decrease of the amount of coolant mass flow to improve the thermal efficiency. Jet impingement heat transfer is the state of the art cooling configuration, which has long been used in this area. In the present study, a modified double swirl chambers cooling configuration has been developed for the gas turbine blade leading edge. The double swirl chambers cooling (DSC) technology is introduced by the authors and comprises a significant enhancement of heat transfer due to the generation of two anti-rotating swirls. In DSC cooling the reattachment of the swirl flows with the maximum velocity at the middle of the chamber leads to a linear impingement effect, which is most suitable for the leading edge cooling for a gas turbine blade. In addition, because of the two swirls both suction side and pressure side of the blade near the leading edge can be very well cooled. In this work, a comparison among three different internal cooling configurations for the leading edge (impingement cooling, swirl chamber and double swirl chambers) has been investigated numerically. With the same inlet slots and the same Reynolds number based on hydraulic diameter of channel the DSC cooling shows overall higher Nusselt number ratio than that in the other two cooling configurations. Downstream of the impingement point, due to the linear impingement effect, the DSC cooling has twice the heat flux in the leading edge area than the standard impingement cooling channel.

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