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.

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.

scholarly journals Numerical investigation about backflow of film cooling in static turbine blade leading edge

2019 ◽  
Vol 11 (11) ◽  
pp. 168781401988581
Author(s):  
Chao Gao ◽  
Haiwang Li ◽  
Huimin Zhou ◽  
Yiwen Ma ◽  
Ruquan You
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Hole Diameter ◽  
Computational Method ◽  
Coverage Area ◽  
Slowing Down ◽  
Film Cooling Holes ◽  
Blade Leading Edge

In this article, film cooling characteristics, especially the phenomenon of backflow for the straight turbine blade leading edge, are investigated. Shear stress transport k-ω turbulence model and structured grids are employed to assure the accuracy of the simulation, and the computational method is verified by the available experimental data. The influences of blow ratio, hole diameter, and the spacing between holes in each row are analyzed. The formation mechanism of backflow is discussed to prevent it from happening or relieve the degree of backflow, thereby to improve the cooling efficiency. The results showed that backflow can be avoided by adjusting the structure and the layout of film cooling holes. With increase in blow ratio, the cooling film becomes more obvious at first and then fades gradually for departing from the blade surface. The jet flow is influenced by the total pressure ratio between coolant cavity and surface of blade leading edge. Smaller film hole diameter and larger hole spacing makes it easier to eject coolant and form continuous film by slowing down the pressure in the cavity. Increasing ratio of hole spacing to hole diameter ( p/ d) can effectively prevent backflow, whereas larger p/ d also makes the film coverage area smaller.

Numerical Study on Influencing Factors of Film Cooling on Turbine Blade Leading Edge Under Rotating State

2018 ◽  
Author(s):  
Yiwen Ma ◽  
Haiwang Li ◽  
Meisong Yang ◽  
Min Wu ◽  
Huimin Zhou
Keyword(s):  
Heat Transfer ◽  
High Temperature ◽  
Turbine Blade ◽  
Rotation Number ◽  
Film Cooling ◽  
Leading Edge ◽  
Heat Transfer Characteristics ◽  
Transfer Characteristics ◽  
Flow And Heat Transfer ◽  
Blade Leading Edge

Engine turbine blades operate at a high speed of rotation and are subjected to high temperature and pressure prevailing gas from the combustion chamber, making the working condition very harsh. In particular, the leading edge of the blade, which is directly subjected to high-temperature gas impacts, is the hottest part of the turbine. Therefore, it is of great importance to improve the protection of the blade leading edge and enhance the understanding of this part of the flow field and temperature field. This paper will focus on the phenomenon of wake deflection and study the film cooling characteristics of the turbine blade under rotating condition. The characteristics of pressure surface and suction surface of the blade are verified by numerical simulation. The contents cover the influence of the film hole diameter, pitch, blowing ratio, rotation number and the development process, the film cooling efficiency on the outflow of coolant film. The result shows that Coriolis force, centrifugal force and secondary flow induced by rotation will change the mainstream flow along the blade, which will lead to changes of pattern concerning the development of the film on the blade surface. In the process of wake development, deflection occurs in different directions at different positions, and the greater the rotation number is, the more obvious the degree of deflection will be. Studying the model with film holes on the leading edge of the blade, these phenomena can be observed along the downstream on the pressure and suction surfaces. Also, models with film holes independently set on the pressure and suction surfaces can be used as proof of these features. At the same time, this paper studies the flow and heat transfer characteristics of the leading-edge gas film under rotating condition and focuses on the influence of rotation on the outflow and the development processes of the wake. The gas film cooling models in rotating state of different film hole diameters and film hole radial spacing will also be compared to further understand the flow and heat transfer characteristics of film cooling on the leading edge of the blade.

Net Heat Flux Reduction and Overall Effectiveness for a Turbine Blade Leading Edge

2005 ◽  
Author(s):  
Brian D. Mouzon ◽  
Elon J. Terrell ◽  
Jason E. Albert ◽  
David G. Bogard
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Cooling Performance ◽  
Net Heat Flux ◽  
Laboratory Models ◽  
Overall Effectiveness ◽  
Blade Leading Edge

The external cooling performance of a film cooled turbine airfoil can be quantified as a net reduction in heat transfer relative to the turbine airfoil without film cooling. This quantification is generally accomplished by using measurements of the adiabatic effectiveness and the change in heat transfer coefficients (hf/h0) for the film cooled surface to determine the net heat flux reduction (Δqr). Although measurement of Δqr for laboratory models give an indication of the ultimate film cooling performance, this does not show how much the surface temperature of the airfoil is reduced by film cooling. Measurement of scaled surface temperatures can be accomplished by using laboratory models constructed so that the Biot number is matched with that of the actual airfoil. These measurements provide a scaled temperature distribution on the airfoil that is referred to as the overall effectiveness, φ. For the current study, measurements of Δqr and φ have been made for a simulated turbine blade leading edge. The simulated leading edge incorporated shaped coolant holes, and had three rows of coolant holes. Improvements due to the shaped holes were determined by comparisons with previously measured round hole configurations. Spatially distributed hf/h0 show increases of 5% to 15% for M = 1.0 and 10% to 30% for M = 2.0. Results show that local variation in Δqr much greater than variation in φ, but laterally averaged Δqr distributions are reasonable predictors of the laterally averaged φ distributions.

Predictions of Cooling From Dirt Purge Holes Along the Tip of a Turbine Blade

2003 ◽  
Author(s):  
E. M. Hohlfeld ◽  
J. R. Christophel ◽  
E. L. Couch ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip ◽  
Flow Through ◽  
Film Cooling Holes

The clearance gap between the tip of a turbine blade and its associated shroud provides a flow path for leakage from the pressure side of the blade to the suction side. The tip region is one area that experiences high heat transfer and, as such, can be the determining factor for blade life. One method for reducing blade tip heat transfer is to use cooler fluid from the compressor, that exits from relatively large dirt purge holes placed in the tip, for cooling purposes. Dirt purge holes are typically manufactured in the blade tip to extract dirt from the coolant flow through centrifugal forces such that these dirt particles do not block smaller diameter film-cooling holes. This paper discusses the results of numerous computational simulations of cooling injection from dirt purge holes along the tip of a turbine blade. Some comparisons are also made to experimental results in which a properly scaled-up blade geometry (12X) was used to form a two-passage linear cascade. Computational results indicate that the cooling achieved through the dirt purge injection from the blade tip is dependent on the gap size as well as the blowing ratio. For a small tip gap (0.54% of the span) the flow exiting the dirt purge holes act as a blockage for the leakage flow across the gap. As the blowing ratio is increased for a large tip gap (1.63% of the span), the tip cooling increases only slightly while the cooling to the shroud increases significantly.

Film Cooling Parameter Waveforms on a Turbine Blade Leading Edge Model With Oscillating Stagnation Line

2016 ◽  
Vol 138 (7) ◽  
Author(s):  
James L. Rutledge ◽  
Tylor C. Rathsack ◽  
Matthew T. Van Voorhis ◽  
Marc D. Polanka
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Stagnation Line ◽  
Inverse Heat Transfer ◽  
Adiabatic Effectiveness ◽  
Blade Leading Edge

It is necessary to understand how film cooling influences the external convective boundary condition involving both the adiabatic wall temperature and the heat transfer coefficient in order to predict the thermal durability of a gas turbine hot gas path component. Most studies in the past have considered only steady flow, but studies of the unsteadiness naturally present in turbine flow have become more prevalent. One source of unsteadiness is wake passage from upstream components which can cause fluctuations in the stagnation location on turbine airfoils. This in turn causes unsteadiness in the behavior of the leading edge coolant jets and thus fluctuations in both the adiabatic effectiveness and heat transfer coefficient. The dynamics of h and η are now quantifiable with modern inverse heat transfer methods and nonintrusive infrared thermography. The present study involved the application of a novel inverse heat transfer methodology to determine time-resolved adiabatic effectiveness and heat transfer coefficient waveforms on a simulated turbine blade leading edge with an oscillating stagnation position. The leading edge geometry was simulated with a circular cylinder with a coolant hole located 21.5 deg downstream from the leading edge stagnation line, angled 20 deg to the surface and 90 deg to the streamwise direction. The coolant plume is shown to shift in response to the stagnation line movement. These oscillations thus influence the film cooling coverage, and the time-averaged benefit of film cooling is influenced by the oscillation.

Heat Transfer and Effectiveness on the Film Cooled Tip and Inner Rim Surfaces of a Turbine Blade

2010 ◽  
Author(s):  
Jun Su Park ◽  
Dong Hyun Lee ◽  
Hyung Hee Cho ◽  
Dong-Ho Rhee ◽  
Shin-Hyung Kang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
High Heat ◽  
Chord Length ◽  
Tip Clearance ◽  
Film Cooling Effectiveness ◽  
High Heat Transfer ◽  
Film Cooling Holes

Detailed heat/mass transfer coefficients and film-cooling effectiveness were measured on the tip and inner rim surfaces of a rotor blade with a squealer rim. The blade was a two-dimensional version of a modern first-stage gas turbine rotor blade with a squealer rim. The experimental apparatus was equipped with a linear cascade of three blades, the axial chord length (Cx) of which was 237 mm with a turning angle of 126°. The mainstream Reynolds number based on the axial chord was 1.5×105. The turbulence intensity level at the cascade inlet was approximately 12%. Measurements were made at three different rim heights (H) of about 3%, 6%, and 9% of the axial chord length. The tip clearance (C) ranges were 1–3% of the axial chord length. Also, three different types of blade tip surfaces were equipped with a single row of film-cooling holes along the camber line, near the pressure and the suction side rim. In particular, a coolant was injected at an incline of 45° from near the suction side film cooling holes. The film cooling experiments were done with a fixed tip clearance and rim height at 1% and 6% of the axial chord length. The blowing rate was fixed at 1.5. High heat transfer rates were observed near the leading edge on the tip surface in some cases, due to the reattachment of tip leakage flow. The peak values moved toward the suction-side edge, and the magnitude and area of high heat transfer increased near the leading edge as the tip clearance increased. The heat transfer decreased on the tip surface with increases in the rim height. In the film-cooling cases, the high heat transfer and film-cooling effectiveness region appeared near the film-cooling holes.

Film Cooling Parameter Waveforms on a Turbine Blade Leading Edge Model With Oscillating Stagnation Line

2015 ◽  
Author(s):  
James L. Rutledge ◽  
Tylor C. Rathsack ◽  
Matthew T. Van Voorhis ◽  
Marc D. Polanka
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Stagnation Line ◽  
Inverse Heat Transfer ◽  
Adiabatic Effectiveness ◽  
Blade Leading Edge

It is necessary to understand how film cooling influences the external convective boundary condition involving both the adiabatic wall temperature and the heat transfer coefficient in order to predict the thermal durability of a gas turbine hot gas path component. Most studies in the past have considered only steady flow, but studies of the unsteadiness naturally present in turbine flow have become more prevalent. One source of unsteadiness is wake passage from upstream components which can cause fluctuations in the stagnation location on turbine airfoils. This in turn causes unsteadiness in the behavior of the leading edge coolant jets and thus fluctuations in both the adiabatic effectiveness and heat transfer coefficient. The dynamics of h and η are now quantifiable with modern inverse heat transfer methods and non-intrusive infrared thermography. The present study involved the application of a novel inverse heat transfer methodology to determine time resolved adiabatic effectiveness and heat transfer coefficient waveforms on a simulated turbine blade leading edge with an oscillating stagnation position. The leading edge geometry was simulated with a circular cylinder with a coolant hole located 21.5° downstream from the leading edge stagnation line, angled 20° to the surface and 90° to the streamwise direction. The coolant plume is shown to shift in response to the stagnation line movement. These oscillations thus influence the film cooling coverage and the time-averaged benefit of film cooling is influenced by the oscillation.

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