Measurement of Aerodynamic Losses for Turbine Airfoil Cascades With Varying Pitch, Operating Under Transonic Conditions

2012 ◽  
Author(s):  
Santosh Abraham ◽  
Kapil Panchal ◽  
Srinath V. Ekkad ◽  
Wing Ng ◽  
Andrew S. Lohaus ◽  
...  
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Leading Edge ◽  
Experimental Results ◽  
Turning Angle ◽  
Pressure Losses ◽  
Turbine Cascades ◽  
Exit Mach Number ◽  
End Wall ◽  
Secondary Flow Field

The paper presents detailed experimental results of the midspan total pressure losses, secondary flow field, and static pressure measurements on two linear, high-turning turbine cascades at transonic conditions. The airfoils in the two cascades being studied are identical and their aerodynamic loading levels are varied by increasing the pitch of one cascade by 25% with respect to the other. The turbine cascades are referred to as B1-SP and B1-IP. Cascade B1-IP, with its increased pitch, has a Zweifel coefficient increased by 25%. The airfoils have a turning angle of ∼127°. Measurements are made at design and off-design conditions, at exit Mach numbers ranging from 0.71 to 0.95. The exit span of the airfoils are increased relative to the inlet span with the intention of obtaining a ratio of inlet Mach number to exit Mach number that is representative to that encountered in a real engine. This results in one end wall diverging from inlet to exit at a 13 degree angle, which simulates the required leading edge loading as seen in an engine. The objective of this study is to investigate the variation in airfoil loading distribution and the effect it has on aerodynamic performance in terms of pressure losses. Detailed loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 Cax and 1.0 Cax locations downstream of the trailing edge. Results from 3D viscous numerical simulations have been used to assist the interpretation of experimental results.

Annular Cascade Testing of Turbine Nozzles at High Exit Mach Numbers

1986 ◽  
Vol 108 (3) ◽  
pp. 313-320 ◽  
Author(s):  
R. G. Williamson ◽  
S. H. Moustapha
Keyword(s):  
Mach Number ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Three Dimensional Flow ◽  
Pressure Losses ◽  
Dimensional Flow ◽  
Pressure Distributions ◽  
Exit Mach Number ◽  
Annular Cascade ◽  
End Wall

This paper presents detailed information on the three-dimensional flow field in a realistic low aspect ratio, high turning nozzle vane design which incorporates end-wall contouring and which has been tested over a range of exit Mach number from subsonic up to the design value at mean section of 1.15. The experimental results, in the form of nozzle surface pressure distributions as well as surveys of pressure losses and flow angles at exit, are compared with those calculated by a three-dimensional flow analysis. The effects of exit Mach number on the measured nozzle performance are also presented.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

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 Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Experimental Investigation on the Annular Sector Cascade of a High Endwall-Angle Turbine

2016 ◽  
Author(s):  
Weiliang Fu ◽  
Jie Gao ◽  
Chen Liang ◽  
Fukai Wang ◽  
Qun Zheng ◽  
...  
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Incidence Angle ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Control Program ◽  
Turbine Cascade ◽  
Cascade Flow ◽  
Annular Sector ◽  
The Way

The flow in high endwall-angle turbine is complex, and it is different from the ordinary turbine flow in characteristics. In order to study the flow field characteristics of high endwall-angle turbines, the annular sector cascade experimental study of high endwall-angle turbines is carried out. The blade is studied experimentally in the form of annular sector cascade. The cascade includes 7 blades, and makes up 6 flow passages, in order to simulate full cascade flow. The experimental Mach number is adjusted by the way of changing inlet total pressure, and the Mach number influence (0.7, 0.8 and 0.9) on annular sector cascade flow is studied. Based on it, the inlet incidence angle (−15°, −7.5°, 0°, 7.5° and 15° )is changed with the way of changing sector straight pipes upstream of the cascade, and its influence on turbine flow fields is studied at the Mach number of 0.8. Here, five-hole probes are used to measure aerodynamic parameters distributions downstream of the cascade, and static pressure taps are positioned on the blade surface to measure surface static pressure distribution. The auto-traversing system and pressure sensors were operated by a self-compiled program based control program. The results indicate that there are two passage vortices inside the turbine cascade flow passage under the high Mach number condition, and the passage vortex near the high endwall-angle region is bigger. As Mach number increases, the passage vortices inside turbine cascade passage will become strong, and moves towards the blade mid-span. Besides, it is shown that the way of changing sector straight pipes can achieve the variation of inlet incidence angles. And, the blade profile with big leading-edge radius has good design and off-design performance. Detailed results and analyses are presented in the paper.

The Effect of Turning Angle on the Loss Generation of LP Turbines

2017 ◽  
Author(s):  
Diego Torre ◽  
Guillermo Garcia-Valdecasas ◽  
David Cadrecha
Keyword(s):  
High Speed ◽  
Pressure Coefficient ◽  
Experimental Results ◽  
Cfd Simulations ◽  
Airfoil Surface ◽  
High Lift ◽  
Turning Angle ◽  
Pressure Losses ◽  
Coefficient Distribution ◽  
Inlet Angle

The effect of turning angle on the loss generation of Low Pressure (LP) Turbines has been investigated experimentally in a couple of turbine high-speed rigs. Both rigs consisted of a rotor-stator configuration. All the airfoils are high lift and high aspect ratio blades that are characteristic of state of the art LP Turbines. Both rigs are identical with exception of the stator. Therefore, two sets of stators have been manufactured and tested. The aerodynamic shape of both stators has been designed in order to achieve the same spanwise distribution of Cp (Pressure coefficient) over the airfoil surface, each one to its corresponding turning angles. Exit angle in both stators is the same. Therefore the change in turning is obtained by a different inlet angle. The aim of this experiment is to obtain the sensitivity of profile and endwall losses to turning angle 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 Mach number, each one to its corresponding velocity triangles, one can state that the results are only affected by the turning angle. Experimental results are presented and compared in terms of area average, radial pitchwise average distributions and exit plane contours of total pressure losses. CFD simulations for the two sets of stators are also presented and compared with the experimental results.

Production and Development of Secondary Flows and Losses in Two Types of Straight Turbine Cascades: Part 1—A Stator Case

1987 ◽  
Vol 109 (2) ◽  
pp. 186-193 ◽  
Author(s):  
A. Yamamoto
Keyword(s):  
Secondary Flow ◽  
Total Pressure ◽  
Experimental Information ◽  
Secondary Flows ◽  
Accumulation Process ◽  
Flow Angle ◽  
Turning Angle ◽  
Pressure Losses ◽  
Flow Vectors ◽  
Turbine Cascades

The present study intends to give some experimental information on secondary flows and on the associated total pressure losses occurring within turbine cascades. Part 1 of the paper describes the mechanism of production and development of the loss caused by secondary flows in a straight stator cascade with a turning angle of about 65 deg. A full representation of superimposed secondary flow vectors and loss contours is given at fourteen serial traverse planes located throughout the cascade. The presentation shows the mechanism clearly. Distributions of static pressures and of the loss on various planes close to blade surfaces and close to an endwall surface are given to show the loss accumulation process over the surfaces of the cascade passage. Variation of mass-averaged flow angle, velocity and loss through the cascade, and evolution of overall loss from upstream to downstream of the cascade are also given. Part 2 of the paper describes the mechanism in a straight rotor cascade with a turning angle of about 102 deg.

Midspan Flow-Field Measurements for Two Transonic Linear Turbine Cascades at Off-Design Conditions

2001 ◽  
Author(s):  
D. B. M. Jouini ◽  
S. A. Sjolander ◽  
S. H. Moustapha
Keyword(s):  
Flow Field ◽  
Turbine Blade ◽  
Density Ratio ◽  
Leading Edge ◽  
Field Measurements ◽  
Reynolds Numbers ◽  
Experimental Results ◽  
Tabular Form ◽  
Data Set ◽  
Turbine Cascades

The paper presents detailed mid-span experimental results from two transonic linear turbine cascades. The blades for the two cascades were designed for the same service and differ mainly in their leading-edge geometries. One of the goals of the study was investigate the influence of the leading-edge metal angle on the sensitivity of the blade to positive off-design incidence. Measurements were made for incidence values of −10.0°, 0.0°, +4.5°, +10.0°, and +14.5° relative to design incidence. The exit Mach numbers varied roughly from 0.5 to 1.2 and the Reynolds numbers from about 4×105 to 106. The measurements include the midspan losses, blade loadings and base pressures. In addition, the axial-velocity-density ratio (AVDR) was extracted for each operating point The AVDR was found to vary from about 0.98 at −10.0° of incidence to about 1.27 at +14.5°. Thus, the data set also provides some evidence of the influence AVDR on axial turbine blade performance. Detailed experimental results for turbine blade performance at off-design incidence are very scarce in the open literature, particularly for transonic conditions. Among other things, the present results are intended to expand the database available in the open literature. To this end, the key aerodynamic results are presented in tabular form, along with the detailed geometry of the cascades. The results could be used in the development of new or improved correlations for use in the early stages of design. They could also be used to evaluate the ability of current CFD codes to capture reliably the variation in losses and other aerodynamic quantities with variations in blade incidence.

Measurement of the Profile Drag of Compressor and Turbine Cascades and the Effect of Wakes in Exciting Vibration

1953 ◽  
Vol 57 (515) ◽  
pp. 722-725 ◽  
Author(s):  
J. M. Stephenson
Keyword(s):  
Mach Number ◽  
Compressible Flow ◽  
Total Pressure ◽  
Static Pressure ◽  
Error Curve ◽  
Exciting Vibration ◽  
Turbine Cascades ◽  
Simplified Procedure

The Melvill Jones equation for the profile drag of a single aerofoil is adapted to the case of an aerofoil in cascade, where the static pressure may be permanently raised (compressor), or lowered (turbine). A simplified procedure for measuring the drag is then described, assuming that the total pressure wake has the form of an error curve. A table of multiplying factors is given, for compressible flow up to an outlet Mach number of 0·9. Many published measurements of cascade drag have ignored this factor, with a consequent error of up to 25 per cent.

Numerical Modeling of Heat Transfer and Pressure Losses for an Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2009 ◽  
Vol 1 (2) ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; a flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25%, 2.0%, and 2.75% of the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the “sweet spot” matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.

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