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.

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]

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.

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

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.

About the Distribution of the Mach Number and Losses Under Consideration of the Diffusion Factor: Profile Distribution and Losses for Low Pressure Turbine Profiles With High Diffusion Factors Under Steady Flow Conditions

2016 ◽  
Author(s):  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbulence Intensity ◽  
Flow Separation ◽  
Total Pressure ◽  
Suction Side ◽  
Diffusion Factor ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
Exit Mach Number

The results of this investigation consist of two linear cascades at high diffusion factors. The present measurements for each low pressure turbine profile were conducted at midspan under a range of Reynolds- and exit Mach numbers. The exit Mach number was varied in a range covering low subsonic up to values where a transonic flow regime on the suction side of the blade could be expected. The variation of the exit Mach number was also used to create different locations of the maximum Mach number and to evaluate the resulting total pressure losses. This work focuses on two profiles with a diffusion factor in a range of 0.18 ≤ DF ≤ 0.22, which is considered as a comparable level for the two cascades. The profile A is a front-loaded design and has shown no obvious flow separation on the suction side of the blade. Compared to the profile A the design B is a more aft-loaded profile which indicates flow separation on the suction side for all investigated Reynolds numbers. The integral total pressure losses were evaluated by wake traverses downstream of the profile. To determine the isentropic Mach numbers and the character of the boundary layer along the suction side of the profile, static pressure tappings and measurements with a flattened Pitot probe were carried out. Numerical studies were also conducted to investigate further the influence of a reduced turbulence intensity on the boundary layer of the suction side of design B. The results show that the optimum of the integral total pressure losses are significantly dependent on the Reynolds number. Therefore a correlation between the maximum Mach number on the suction side and the integral total pressure losses has been successfully established. A significant change of the turbulence intensity at the inlet of the cascade leads to shift of the location of the maximum Mach number. It also results in an equivalent change of the total pressure losses, which has been predicted by the trend line. However, the trend lines, which are based on the data of the integral total pressure losses of an attached boundary layer, are not able to predict the integral total pressure loss or the location of the maximum Mach number on the suction side of the blade since an open separation bubble occurs.

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.

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.

Aerothermal Performance of a Cooled Winglet at Engine Representative Mach and Reynolds Numbers

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
D. O. O’Dowd ◽  
Q. Zhang ◽  
L. He ◽  
B. C. Y. Cheong ◽  
I. Tibbott
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
High Speed ◽  
Tip Clearance ◽  
Test Facility ◽  
Pressure Probe ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Loss Coefficients

This paper presents an experimental investigation of the aerothermal performance of a cooled winglet tip under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27 × 106, based on axial chord). Spatially resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.

scholarly journals Aerodynamic Performance of a Transonic Turbine Blade Passage in Presence of Upstream Slot and Mateface Gap With Endwall Contouring

2014 ◽  
Author(s):  
Sakshi Jain ◽  
Arnab Roy ◽  
Wing Ng ◽  
Srinath Ekkad ◽  
Andrew S. Lohaus ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Film Cooling ◽  
Operating Conditions ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Exit Mach Number ◽  
The Heat Transfer Coefficient

The present article investigates mixed out aerodynamic loss coefficient measurements for a high turning, contoured endwall passage under transonic operating conditions in presence of upstream purge slot and mateface gap. The upstream purge slot represents the gap between stator-rotor interface and the mateface gap simulates the assembly feature between adjacent airfoils in an actual high pressure turbine stage. While the performance of the mateface and upstream slot has been studied for lower Mach number, no studies exist in literature for transonic flow conditions. Experiments were performed at the Virginia Tech’s linear, transonic blow down cascade facility. Measurements were carried out at design conditions (isentropic exit Mach number of 0.88, design incidence) without and with coolant blowing. Upstream leakage flow of 1.0% coolant to mainstream mass flow ratio (MFR) was considered with the presence of mateface gap. There was no coolant blowing through the mateface gap itself. Cascade exit pressure measurements were carried out using a 5-hole probe traverse at a plane 1.0-Cax downstream of the trailing edge. Spanwise measurements were performed to complete the entire 2D loss plane from endwall to midspan, which were used to plot pitchwise averaged losses for different span locations and loss contours for the passage. Results reveal significant reduction in aerodynamic losses using the contoured endwall due to the modification of flow physics compared to a non-contoured planar endwall. The heat transfer experiments, designed to find the heat transfer coefficient and the film cooling effectiveness are described in detail in a separate paper [1].

The Effect of Mach Number on the Loss Generation of LP Turbines

2012 ◽  
Author(s):  
Raúl Vázquez ◽  
Diego Torre
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
High Speed ◽  
State Of The Art ◽  
Pressure Coefficient ◽  
Airfoil Surface ◽  
High Lift ◽  
Pressure Losses ◽  
Coefficient Distribution ◽  
The Impact

The effect of Mach number on the loss generation of Low Pressure (LP) Turbines has been investigated experimentally in a pair of turbine high-speed rigs. Both rigs consist of a rotor-stator configuration. All the airfoils are high lift, high aspect ratio and high turning blades that are characteristic of state of the art LP Turbines. Both rigs are identical with exception of the stator. Two sets of stators have been manufactured and tested. The aerodynamic shape of both stators have been designed in order to achieve the same spanwise distribution of Cp (Compressible Pressure coefficient) over the airfoil surface, each one to its corresponding design Mach number (0.61 and 0.88 respectively). The aim of this experiment is to obtain the sensitivity of profile and endwall losses to Mach number by means of a back-to-back comparison between both sets of airfoils. Because the two sets of stators maintain the same pressure coefficient distribution, Reynolds number and velocity triangles, each one to its corresponding design Mach number; one can state that the results are only affected by the compressibility. Experimental results are presented and compared in terms of area average, radial pitchwise average distributions and exit plane contours of total pressure losses. To complete the paper, the impact of the results on the design of LP Turbines is discussed and presented.

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