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.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2008 ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient ◽  
Aerodynamic Losses

The performance of suction-side gill region film cooling is investigated using the University of Utah Transonic Wind Tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77 to 1.99 similar to values present in operating gas turbine engines. Presented are local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated Aerodynamic Losses (IAL) increase anywhere from 4 to 45 percent compared to a smooth blade with no film injection. The performance of each hole type depends upon the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL aerodynamic losses are generally present at any particular wake location for the round radial RR or shaped axial SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA shaped axial holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

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.

scholarly journals Showerhead Film Cooling Performance of a Turbine Vane at High Freestream Turbulence in a Transonic Cascade

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Shakeel Nasir ◽  
Trey Bolchoz ◽  
Wing-Fai Ng ◽  
Luzeng J. Zhang ◽  
Hee Koo Moon ◽  
...  
Keyword(s):  
Reynolds Number ◽  
Nusselt Number ◽  
Mach Number ◽  
Film Cooling ◽  
Boundary Layer Transition ◽  
Suction Surface ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

This paper experimentally investigates the effect of blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane. The vane midspan was instrumented with single-sided platinum thin film gauges to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Three different exit Mach numbers of Mex = 0.57, 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 9.7 × 105, 1.1 × 106 and 1.5 × 106, respectively—were tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0, 1.5, 2.0, and 2.5 and a density ratio of DR = 1.3. Nusselt number and adiabatic film cooling effectiveness distributions were presented on the vane surface over a range of s/C = −0.58 on the pressure side to s/C = 0.72 on the suction side of the vane. The primary effects of coolant injection were to augment the Nusselt number and reduce the adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at all exit Mach number conditions. In general, an increase in blowing ratio (BR = 1.5 to 2.5) showed noticeable Nusselt number augmentation on pressure surface as compared to suction surface at exit Mach 0.57 and 0.75; however, the Nusselt number augmentation for these blowing ratios was found to be negligible on the vane surface for exit Mach 1.0 case. At exit Mach 1.0, an increase in blowing ratio (BR = 1.5 to 2.5) was observed to have an adverse effect on the adiabatic effectiveness on the pressure surface but had negligible effect on suction surface. The effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition in the region s/C = 0.28 to s/C = 0.45 at all blowing ratio and exit Mach number conditions. An increase in Reynolds number from exit Mach 0.76 to 1.0 increased heat transfer levels on the vane surface at all blowing ratio conditions. A large increase in Reynolds number adversely affected adiabatic effectiveness on the pressure surface at all blowing ratio conditions. On the suction surface, a large increase in Reynolds number also affected adiabatic effectiveness in the favorable pressure gradient and boundary layer transition region.

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.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Influence of Unsteady Wake With Trailing Edge Coolant Ejection on Turbine Blade Film Cooling

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Shiou-Jiuan Li ◽  
Akhilesh P. Rallabandi ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Surface Film ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Blowing Ratio ◽  
Unsteady Wake

Detailed film cooling effectiveness distributions along a modeled turbine rotor blade under the combined effects of an upstream trailing edge unsteady wake with coolant ejection are presented using the pressure sensitive paint (PSP) mass transfer analogy method. The experiment is conducted in a low speed wind tunnel facility with a five blade linear cascade. The exit Reynolds number based on the axial chord is 370,000. Unsteady wakes and trailing edge coolant jets are produced by a spoked wheel-type wake generator with hollow rods equipped with several coolant ejections from holes. The coolant-to-mainstream density ratios for both the blade and trailing edge coolant ejection range from 1.5 to 2.0 for simulating realistic engine conditions. Blade blowing ratio studies are 0.5 and 1.0 on the suction surface and 1.0 and 2.0 on the pressure surface. The trailing edge jet blowing ratio and Strouhal numbers are 1.0 and 0.12, respectively. The results show that the unsteady wake reduces the overall film cooling effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances the overall effectiveness. The results also show that the overall filming cooling effectiveness increases by using heavier coolant for trailing edge ejection and for blade surface film cooling.

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 Turbine Nozzle Film Cooling Study Using the Pressure Sensitive Paint (PSP) Technique

1999 ◽  
Author(s):  
Luzeng Zhang ◽  
Michael Baltz ◽  
Ram Pudupatty ◽  
Michael Fox
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
High Speed ◽  
Surface Film ◽  
Pressure Sensitive Paint ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Sensitive

The use of pressure sensitive paint (PSP) to measure film cooling effectiveness on a turbine nozzle surface was demonstrated in a high speed wind tunnel. Film cooling effectiveness was measured from a single row of holes located on a turbine vane suction surface with a shaped exit. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Three blowing ratios were studied for each of the five freestream conditions: a reference condition, a reduced and an increased Reynolds number condition, and a reduced and an increased Mach number condition. The freestream turbulence intensity was kept at 12.0% for all the tests. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. The film effectiveness increased with blowing ratio for all the freestream conditions. The effects of secondary flow and freestream Mach number and Reynolds number on turbine nozzle suction surface film cooling are also discussed.

scholarly journals Showerhead Film Cooling Performance of a Turbine Vane at High Freestream Turbulence in a Transonic Cascade

2008 ◽  
Author(s):  
S. Nasir ◽  
T. Bolchoz ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon ◽  
...  
Keyword(s):  
Reynolds Number ◽  
Nusselt Number ◽  
Mach Number ◽  
Film Cooling ◽  
Boundary Layer Transition ◽  
Suction Surface ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

This paper experimentally investigates the effect of blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane. The vane midspan was instrumented with single-sided platinum thin film gauges to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Three different exit Mach numbers of Mex = 0.57, 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 9.7 × 105, 1.1 × 106 and 1.5 × 106, respectively—were tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0, 1.5, 2.0, and 2.5 and a density ratio of DR = 1.3. Nusselt number and adiabatic film cooling effectiveness distributions were presented on the vane surface over a range of s/C = −0.58 on the pressure side to s/C = 0.72 on the suction side of the vane. The primary effects of coolant injection were to augment the Nusselt number and reduce the adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at all exit Mach number conditions. In general, an increase in blowing ratio (BR = 1.5 to 2.5) showed noticeable Nusselt number augmentation on pressure surface as compared to suction surface at exit Mach 0.57 and 0.75; however, the Nusselt number augmentation for these blowing ratios was found to be negligible on the vane surface for exit Mach 1.0 case. At exit Mach 1.0, an increase in blowing ratio (BR = 1.5 to 2.5) was observed to have an adverse effect on the adiabatic effectiveness on the pressure surface but had negligible effect on suction surface. The effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition in the region s/C = 0.28 to s/C = 0.45 at all blowing ratio and exit Mach number conditions. An increase in Reynolds number from exit Mach 0.76 to 1.0 increased heat transfer levels on the vane surface at all blowing ratio conditions. A large increase in Reynolds number adversely affected adiabatic effectiveness on the pressure surface at all blowing ratio conditions. On the suction surface, a large increase in Reynolds number also affected adiabatic effectiveness in the favorable pressure gradient and boundary layer transition region.

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