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.

Mach Number Distribution and Profile Losses for Low-Pressure Turbine Profiles With High Diffusion Factors

2017 ◽  
Vol 139 (10) ◽  
Author(s):  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Flow Separation ◽  
Total Pressure ◽  
Suction Side ◽  
Low Pressure ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
High Diffusion ◽  
Mach Numbers

The results of this investigation come from two linear cascades at high diffusion factors (DFs). The measurements presented for each low-pressure turbine (LPT) 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. This work focuses on two profiles with a diffusion factor in a range of 0.18≤DF≤0.22, where values in this range are considered as a comparable for the two cascades. Profile A is a front-loaded design and has shown no obvious flow separation on the suction side of the blade. Compared to the design A, design B is a more aft-loaded profile which exhibits flow separation on the suction side for all Reynolds numbers investigated. The integral total pressure losses were evaluated by wake traverses downstream of the airfoil. To determine the isentropic Mach numbers and the character of the boundary layer along the suction side of the profile, the static pressure measurements and traverses with a flattened Pitot probe were carried out. A correlation between the position of maximum Mach number on the suction side and the integral total pressure losses has been successfully established. The results show that the optimum location of peak Mach number to minimize integral total pressure losses is significantly dependent on the Reynolds number. However, the correlation presented in this paper, which is based on the data of the integral total pressure losses of an attached boundary layer, is not able to predict the integral total pressure loss or the location of the maximum Mach number on the suction side of the blade when an open separation bubble occurs.

A Comparison of Three Low Pressure Turbine Designs

2013 ◽  
Vol 135 (5) ◽  
Author(s):  
Christian T. Wakelam ◽  
Martin Hoeger ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Total Pressure ◽  
Maximum Velocity ◽  
Suction Side ◽  
Low Pressure ◽  
Surface Boundary ◽  
Suction Surface ◽  
Pressure Losses ◽  
Low Pressure Turbine

As part of the current research, three low pressure turbine (LPT) geometries—which were designed with a common pitch, axial chord, inlet angle, and exit Mach number and to create the same nominal level of turning—are compared. Each of the LPT cascades was investigated under a range of Reynolds numbers, exit Mach numbers, and under the influence of a moving bar wake generator. Profile static pressure distributions, wake traverses at 5% and 40% axial chord downstream of the trailing edge, and suction side boundary layer traverses were used to compare the performance of the three designs. The total pressure losses are strongly dependent on both the maximum velocity location as well as the diffusion on the suction surface. The importance of the behavior of the pressure surface boundary layer turned out to be negligible in comparison. Cases with equivalent operating Reynolds number and suction side diffusion level are compared in terms of the total pressure losses that are generated. It is shown that a relationship between loss and suction side maximum velocity location exists. An optimum suction side maximum velocity location depends on the Reynolds number, diffusion factor, and wake passing frequency.

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.

The Effect of Pulsed Blowing on the Boundary Layer of a Highly Loaded Low Pressure Turbine Blade

2013 ◽  
Author(s):  
Marion Mack ◽  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Highly Loaded

The performance of the low pressure turbine (LPT) can vary appreciably, because this component operates under a wide range of Reynolds numbers. At higher Reynolds numbers, mid and aft loaded profiles have the advantage that transition of suction side boundary layer happens further downstream than at front loaded profiles, resulting in lower profile loss. At lower Reynolds numbers, aft loading of the blade can mean that if a suction side separation exists, it may remain open up to the trailing edge. This is especially the case when blade lift is increased via increased pitch to chord ratio. There is a trend in research towards exploring the effect of coupling boundary layer control with highly loaded turbine blades, in order to maximize performance over the full relevant Reynolds number range. In an earlier work, pulsed blowing with fluidic oscillators was shown to be effective in reducing the extent of the separated flow region and to significantly decrease the profile losses caused by separation over a wide range of Reynolds numbers. These experiments were carried out in the High-Speed Cascade Wind Tunnel of the German Federal Armed Forces University Munich, Germany, which allows to capture the effects of pulsed blowing at engine relevant conditions. The assumed control mechanism was the triggering of boundary layer transition by excitation of the Tollmien-Schlichting waves. The current work aims to gain further insight into the effects of pulsed blowing. It investigates the effect of a highly efficient configuration of pulsed blowing at a frequency of 9.5 kHz on the boundary layer at a Reynolds number of 70000 and exit Mach number of 0.6. The boundary layer profiles were measured at five positions between peak Mach number and the trailing edge with hot wire anemometry and pneumatic probes. Experiments were conducted with and without actuation under steady as well as periodically unsteady inflow conditions. The results show the development of the boundary layer and its interaction with incoming wakes. It is shown that pulsed blowing accelerates transition over the separation bubble and drastically reduces the boundary layer thickness.

Investigation of Variated Unsteady Inflow Boundary Conditions on the Transition Behavior of a Low Pressure Turbine Cascade Family

2014 ◽  
Author(s):  
Franz F. Blaim ◽  
Roland E. Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Condition ◽  
Transport Equation ◽  
Flow Separation ◽  
Total Pressure ◽  
Low Pressure ◽  
Test Facility ◽  
Number Range ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
Transition Behavior

The objective of this paper is to analyze the influence of incoming periodic wakes, considering the variable width, on the integral total pressure loss for two low pressure turbine (LPT) airfoils. In order to reduce the overall weight of a LPT, the pitch to chord ratio was continuously increased, during the past decades. However, this increase encourages the development of the transition phenomena or even flow separation on the suction side of the blade. At low Reynolds numbers, large separation bubbles can occur there, which are linked with high total pressure losses. The incoming wakes of the upstream blades are known to trigger early transition, leading to a reduced risk of flow separation and hence minor integral total pressure losses caused by separation. For the further investigation of these effects, different widths of the incoming wakes will be examined in detail, here. This variation is carried out by using the numerical Unsteady Reynolds Averaged Solver TRACE developed by the DLR Cologne in collaboration with MTU Aero Engines AG. For the variation of the width of the wakes, a variable boundary condition was modeled, which includes the wake vorticity parameters. The width of the incoming wakes was used as the relevant variable parameter. The implemented boundary condition models the unsteady behavior of the incoming wakes by the variation of the velocity profile, using a prescribed frequenc. TRACE can use two different transition models; the main focus here is set to the γ–Reθt transition model, which uses local variables in a transport equation, to trigger the transition within the turbulence transport equation system. The experimental results were conducted at the high speed cascade open loop test facility at the Institute for Jet Propulsion at the University of the German Federal Armed Forces in Munich. For the investigation presented here, two LPT profiles — which were designed with a similar inlet angle, turning, and pitch are analyzed. However, with a common exit Mach number and a similar Reynolds number range between 40k and 400k, one profile is front loaded and the other one is aft loaded. Numerical unsteady results are in good agreement with the conducted measurements. The influence of the width of the wake on the time resolved transition behavior, represented by friction coefficient plots and momentum loss thickness will be analyzed in this paper.

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.

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 High-Fidelity Simulations of a High-Pressure Turbine Vane Subject to Large Disturbances: Effect of Exit Mach Number on Losses

2021 ◽  
pp. 1-11
Author(s):  
Yaomin Zhao ◽  
Richard Sandberg
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Mach Number ◽  
Strong Shock ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Loss Analysis ◽  
Turbine Efficiency ◽  
Turbulence Condition ◽  
Exit Mach Number

Abstract We report on a series of highly resolved large-eddy simulations of the LS89 high-pressure turbine (HPT) vane, varying the exit Mach number between Ma=0.7 and 1.1. In order to accurately resolve the blade boundary layers and enforce pitchwise periodicity, we for the first time use an overset mesh method, which consists of an O-type grid around the blade overlapping with a background H-type grid. The simulations were conducted either with a synthetic inlet turbulence condition or including upstream bars. A quantitative comparison shows that the computationally more efficient synthetic method is able to reproduce the turbulence characterictics of the upstream bars. We further perform a detailed analysis of the flow fields, showing that the varying exit Mach number significantly changes the turbine efficiency by affecting the suction-side transition, blade boundary layer profiles, and wake mixing. In particular, the Ma=1.1 case includes a strong shock that interacts with the trailing edge, causing an increased complexity of the flow field. We use our recently developed entropy loss analysis (Zhao and Sandberg, GT2019-90126) to decompose the overall loss into different source terms and identify the regions that dominate the loss generation. Comparing the different Ma cases, we conclude that the main mechanism for the extra loss generation in the Ma=1.1 case is the shock-related strong pressure gradient interacting with the turbulent boundary layer and the wake, resulting in significant turbulence production and extensive viscous dissipation.

scholarly journals High-Fidelity Simulations of a High-Pressure Turbine Vane Subject to Large Disturbances: Effect of Exit Mach Number on Losses

2020 ◽  
Author(s):  
Yaomin Zhao ◽  
Richard D. Sandberg
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Mach Number ◽  
Strong Shock ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Loss Analysis ◽  
Turbine Efficiency ◽  
Turbulence Condition ◽  
Exit Mach Number

Abstract We report on a series of highly resolved large-eddy simulations of the LS89 high-pressure turbine (HPT) vane, varying the exit Mach number between Ma = 0.7 and 1.1. In order to accurately resolve the blade boundary layers and enforce pitchwise periodicity, we for the first time use an overset mesh method, which consists of an O-type grid around the blade overlapping with a background H-type grid. The simulations were conducted either with a synthetic inlet turbulence condition or including upstream bars. A quantitative comparison shows that the computationally more efficient synthetic method is able to reproduce the turbulence characterictics of the upstream bars. We further perform a detailed analysis of the flow fields, showing that the varying exit Mach number significantly changes the turbine efficiency by affecting the suction-side transition, blade boundary layer profiles, and wake mixing. In particular, the Ma = 1.1 case includes a strong shock that interacts with the trailing edge, causing an increased complexity of the flow field. We use our recently developed entropy loss analysis (Zhao and Sandberg, GT2019-90126) to decompose the overall loss into different source terms and identify the regions that dominate the loss generation. Comparing the different Ma cases, we conclude that the main mechanism for the extra loss generation in the Ma = 1.1 case is the shock-related strong pressure gradient interacting with the turbulent boundary layer and the wake, resulting in significant turbulence production and extensive viscous dissipation.

A Comparison of Three Low Pressure Turbine Designs

2012 ◽  
Author(s):  
Christian T. Wakelam ◽  
Reinhard Niehuis ◽  
Martin Hoeger
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Total Pressure ◽  
Maximum Velocity ◽  
Suction Side ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Pressure Losses ◽  
Pressure Surface

As part of the current research, three LPT geometries — which were designed with a common pitch, axial chord, inlet angle, and exit Mach number and to create the same nominal level of turning — are compared. Each of the LPT cascades was investigated under a range of Reynolds numbers, exit Mach numbers, and under the influence of a moving bar wake generator. Profile static pressure distributions, wake traverses at 5% and 40% axial chord downstream of the trailing edge and suction side boundary layer traverses were used to compare the performance of the three designs. The total pressure losses are strongly dependant on both the maximum velocity location as well as the diffusion on the suction surface. The importance of the behavior of the pressure surface boundary layer turned out to be negligible in comparison. Cases with equivalent operating Reynolds number and suction side diffusion level are compared in terms of the total pressure losses that are generated. It is shown that a relationship between loss and suction side maximum velocity location exists. An optimum suction side maximum velocity location depends on the Reynolds number, diffusion factor, and wake passing frequency.

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