Experimental Investigation of Sweeping Jet Film Cooling in a Transonic Turbine Cascade

2020 ◽  
Vol 142 (4) ◽  
Author(s):  
Mohammad A. Hossain ◽  
Munevver E. Asar ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Experimental Investigation ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Suction Surface ◽  
Cooling Performance ◽  
Aerodynamic Loss ◽  
Comparison Results ◽  
Shaped Hole

Abstract The sweeping jet (SJ) film cooling hole has shown promising cooling performance compared to the standard shaped hole in low-speed conditions. The present work demonstrates the first attempt of SJ film cooling at an engine relevant Mach number. An experimental investigation was conducted to study the SJ film cooling on a nozzle guide vane suction surface. A well-established additive manufacturing technique commonly known as stereolithography (SLA) was utilized to design a transonic, engine representative vane geometry in which a row of SJ holes was used on the vane suction surface. Experiments were performed in a linear transonic cascade at an exit Mach number of 0.8 and blowing ratios of BR = 0.25–2.23. The measurement of heat transfer was conducted with the transient IR method, and the convective heat transfer coefficient (HTC) and adiabatic film cooling effectiveness were estimated using a dual linear regression technique (DLRT). Aerodynamic loss measurements were also performed with a total pressure Kiel probe at 0.25Cax downstream of the exit plane of the vane cascade. Experiments were also conducted for a baseline-shaped hole (777-hole) for a direct comparison. Results showed that the SJ hole has a wider coolant spreading in the lateral direction near the hole exit due to its sweeping motion that improves the overall cooling performance particularly at high blowing ratios (BR > 1). Aerodynamic loss measurement suggested that the SJ hole has a comparable total pressure loss to the 777-shaped hole.

Experimental Investigation of Sweeping Jet Film Cooling in a Transonic Turbine Cascade

2019 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Munevver E. Asar ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Experimental Investigation ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Suction Surface ◽  
Cooling Performance ◽  
Aerodynamic Loss ◽  
Comparison Results ◽  
Shaped Hole

Abstract The sweeping jet (SJ) film cooling hole has shown promising cooling performance compared to the standard shaped hole in low-speed conditions. The present work demonstrates the first attempt of sweeping jet film cooling at an engine relevant Mach number. An experimental investigation was conducted to study the sweeping jet film cooling on a nozzle guide vane suction surface. A well-established additive manufacturing technique commonly known as Stereolithography (SLA) was unitized to design a transonic, engine representative vane geometry in which a row of SJ holes was used on the vane suction surface. Experiments were performed in a linear transonic cascade at an exit Mach number of 0.8 and blowing ratios of BR = 0.25–2.23. The measurement of heat transfer was conducted with the transient IR method and the convective heat transfer coefficient (HTC) and adiabatic film cooling effectiveness were estimated using a dual linear regression technique (DLRT). Aerodynamic loss measurements were also performed with a total pressure Kiel probe at 0.25Cax downstream of the exit plane of the vane cascade. Experiments were also conducted for a baseline shaped hole (777-hole) for a direct comparison. Results showed that the SJ hole has a wider coolant spreading in the lateral direction near the hole exit due to its sweeping action that improves the overall cooling performance particularly at high blowing ratios (BR>1). Aerodynamic loss measurement suggested that the SJ hole has a comparable total pressure loss to the 777-shaped hole.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

Heat Transfer Coefficient Measurements of Film-Cooling Holes With Expanded Exits

1998 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Film Cooling ◽  
Cylindrical Hole ◽  
Superposition Method ◽  
Transfer Coefficients ◽  
Cooling Performance ◽  
Heat Transfer Coefficients ◽  
Film Cooling Holes ◽  
Injection Location

Detailed measurements of heat transfer coefficients in the nearfield of three different film-cooling holes are presented. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fan-shaped and a laidback fanshaped hole). They were tested over a range of blowing ratios M = 0.25…1.75 at an external crossflow Mach number of 0.6 and a coolant-to-mainflow density ratio of 1.85. Additionally, the effect of the internal coolant supply Mach number is addressed. Temperatures of the diabatic surface downstream of the injection location are measured by means of an infrared camera system. They are used as boundary conditions for a finite element analysis to determine surface heat fluxes and heat transfer coefficients. The superposition method is applied to evaluate the overall film-cooling performance of the hole geometries investigated. As compared to the cylindrical hole, both expanded holes show significantly lower heat transfer coefficients downstream of the injection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the injected coolant than the fanshaped hole which leads to lower laterally averaged heat transfer coefficients. Coolant passage crossflow Mach number affects the flowfield of the jet being ejected from the hole and, therefore, has an important impact on film-cooling performance.

Transonic Aerodynamic Losses Due to Turbine Airfoil, Suction Surface Film Cooling

1999 ◽  
Vol 122 (2) ◽  
pp. 317-326 ◽  
Author(s):  
D. J. Jackson ◽  
K. L. Lee ◽  
P. M. Ligrani ◽  
P. D. Johnson
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Surface Film ◽  
Suction Surface ◽  
Pressure Losses ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Mixing Losses ◽  
Aerodynamic Losses

The effects of suction surface film cooling on aerodynamic losses are investigated using an experimental apparatus designed especially for this purpose. A symmetric airfoil with the same transonic Mach number distribution on both sides is employed. Mach numbers range from 0.4 to 1.24 and match values on the suction surface of airfoils from operating aeroengines. Film cooling holes are located on one side of the airfoil near the passage throat where the free-stream Mach number is nominally 1.07. Round cylindrical and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and a correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical holes are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil, as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane. [S0889-504X(00)02202-9]

scholarly journals Transonic Aerodynamic Losses due to Turbine Airfoil, Suction Surface Film Cooling

1999 ◽  
Author(s):  
D. J. Jackson ◽  
K. L. Lee ◽  
P. M. Ligrani ◽  
P. D. Johnson ◽  
F. O. Soechting
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Surface Film ◽  
Suction Surface ◽  
Pressure Losses ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Mixing Losses ◽  
Aerodynamic Losses

The effects of suction surface film cooling on aerodynamic losses are investigated using an experimental apparatus designed especially for this purpose. A symmetric airfoil with the same transonic Mach number distribution on both sides is employed. Mach numbers along the airfoil surface range from 0.4 to 1.24 and match values on the suction surface of airfoils from operating aeroengines. Film cooling holes are located on one side of the airfoil near the passage throat where the freestream Mach number is nominally 1.07. Round cylindrical, and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and n correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical boles are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane.

Film Cooling Performance of Waist-Shaped Slot Holes

2010 ◽  
Author(s):  
Cun-liang Liu ◽  
Hui-ren Zhu ◽  
Jiang-tao Bai ◽  
Du-chun Xu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Thermal Protection ◽  
Upstream Region ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Shaped Hole ◽  
The Heat Transfer Coefficient

Film cooling performance of a new shaped hole: waist-shaped slot hole is studied in this paper. Experimental measurement and numerical simulation are carried out to investigate the film cooling character and physics of this new shaped hole. And comparisons between the waist-shaped slot hole and two kinds of console holes are also performed. Both the cooling effectiveness distribution and the heat transfer coefficient distribution of the waist-shaped slot hole are similar with those of the console hole with large divergence angle because of the effect of the waist-shaped slot hole’s structure. The middle constriction structure of the waist-shaped slot hole and the coupled vortices make jets from the waist-shaped slot holes produce higher cooling effectiveness in the midspan region between adjacent holes. And also due to the effect of the middle constriction structure, the heat transfer coefficient of the waist-shaped slot hole is very high in the upstream midspan region. However, the heat transfer coefficient in the downstream midspan region is lower than that in the region near the hole centerline because of the effect of the coupled vortices. The waist-shaped slot holes provide the surface with very good thermal protection, especially in the upstream region. Although the console holes with small exit-entry area ratio provide better thermal protection than the waist-shaped slot holes due to small turbulence intensity, the flow resistance characteristic of the waist-shaped slot hole is much better.

Sweeping Jet Film Cooling on a Turbine Vane

2018 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Lucas Agricola ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Convective Heat Transfer Coefficient ◽  
Total Pressure Loss ◽  
Pressure Probe ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Shaped Hole

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Experimental Investigation of Film Cooling Performance for a New Shaped Hole at the Leading Edge

2014 ◽  
Author(s):  
Tarek Elnady ◽  
Ibrahim Hassan
Keyword(s):  
Heat Transfer ◽  
Experimental Investigation ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Local Heat Transfer ◽  
Stagnation Line ◽  
Cooling Performance

An experimental investigation has been performed to study the film cooling performance of a smooth expansion exit at the leading edge of a gas turbine vane. A two-dimensional cascade has been employed to measure the cooling performance of the proposed expansion using the transient Thermochromatic Liquid Crystal technique. One row of cylindrical holes, located on the stagnation line, is investigated with two expansion levels, 2d and 4d, in addition to the standard hole. The air is injected at 90° and 60° inclination angle relative to the vane surface at four blowing ratios ranging from 1 to 2 at a 0.9 density ratio. The Mach number and the Reynolds number based on the cascade exit velocity and the axial chord are 0.23 and 1.4E5, respectively. The detailed local heat transfer coefficient over both the pressure side and the suction side are presented in addition to the lateral-averaged normalized heat transfer coefficient. The proposed expansion provides a lower heat transfer coefficient compared with the standard cylindrical hole over the investigated blowing ratios. Combining the heat transfer coefficient with the corresponding cooling effectiveness, previously presented, the smooth expansion shows a significant reduction in the heat load with more uniform distribution of the coolant over the leading edge region. The strong confrontation between the coolant jet and the mainstream, in case of 90° injection, yields a strong dispersion of the coolant with higher heat transfer coefficient and high thermal load over the vane surface.

Sweeping Jet Film Cooling on a Turbine Vane

2019 ◽  
Vol 141 (3) ◽  
Author(s):  
Mohammad A. Hossain ◽  
Lucas Agricola ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Convective Heat Transfer Coefficient ◽  
Total Pressure Loss ◽  
Pressure Probe ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Shaped Hole

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (SH) (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole, while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Aero-Thermal Optimization and Experimental Verification for the Discrete Film Cooling of a Turbine Airfoil

2013 ◽  
Author(s):  
Carole El Ayoubi ◽  
Othman Hassan ◽  
Wahid Ghaly ◽  
Ibrahim Hassan
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Pareto Front ◽  
Optimal Solutions ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Flow Parameters ◽  
Turbine Vane

The optimization aims to maximize the film cooling performance while minimizing the corresponding aerodynamic penalty. The film cooling performance is assessed using the adiabatic film cooling effectiveness, while the aerodynamic penalty is measured with a mass-averaged total pressure loss coefficient. Two design variables are selected; the coolant to mainstream temperature ratio and total pressure ratio. Two staggered rows of discrete cylindrical film cooling holes on the suction surface of a turbine vane are considered. The effect of varying the coolant flow parameters on the adiabatic film cooling effectiveness and the aerodynamic loss is analyzed using the optimization method and three-dimensional Reynolds-averaged Navier-Stokes (RANS) simulations. The CFD predictions of the adiabatic film cooling effectiveness and aerodynamic performance are assessed and validated against corresponding experimental measurements. The optimal solutions are reproduced in the experimental facility and the Pareto front is substantiated with experimental data. A non-dominated sorting genetic algorithm (NSGA-II) is coupled with an artificial neural network (ANN) to perform a multiple objective optimization of the film coolant flow parameters on the suction surface of a high pressure gas turbine vane. The numerical predictions are employed to construct the artificial neural network that produces low-fidelity predictions of the objectives during the optimization. The Pareto front of optimal solutions is generated by the optimization methodology.

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