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.

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

1986 ◽  
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°. A full representation of superimposed secondary flow vectors and loss contours is given at serial fourteen traverse planes located throughout the cascade, which 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°.

Migration of Combustor Exit Profiles Through High Pressure Turbine Vanes

2007 ◽  
Author(s):  
M. D. Barringer ◽  
M. D. Polanka ◽  
J. P. Clark ◽  
P. J. Koch ◽  
K. A. Thole
Keyword(s):  
High Pressure ◽  
Secondary Flow ◽  
Total Pressure ◽  
Secondary Flows ◽  
Working Fluid ◽  
Outer Diameter ◽  
High Pressure Turbine ◽  
Flow Angle ◽  
Radial Temperature ◽  
Turbine Vane

The high pressure turbine stage within gas turbine engines is exposed to combustor exit flows that are nonuniform in both stagnation pressure and temperature. These highly turbulent flows typically enter the first stage vanes with significant spatial gradients near the inner and outer diameter endwalls. These gradients can result in secondary flow development within the vane passage that is different than what classical secondary flow models predict. The heat transfer between the working fluid and the turbine vane surface and endwalls is directly related to the secondary flows. The goal of the current study was to examine the migration of different inlet radial temperature and pressure profiles through the high turbine vane of a modern turbine engine. The tests were performed using an inlet profile generator located in the Turbine Research Facility (TRF) at the Air Force Research Laboratory (AFRL). Comparisons of area-averaged radial exit profiles are reported as well as profiles at three vane pitch locations to document the circumferential variation in the profiles. The results show that the shape of the total pressure profile near the endwalls at the inlet of the vane can alter the redistribution of stagnation enthalpy through the airfoil passage significantly. Total pressure loss and exit flow angle variations are also examined for the different inlet profiles.

Low-Speed Annular Cascade Tests of an Ultra-Highly Loaded Turbine With Tip Clearance: Part 1 — Near Design Incidence

1999 ◽  
Author(s):  
A. Yamamoto ◽  
E. Outa
Keyword(s):  
Flow Visualization ◽  
Total Pressure ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Flow Structures ◽  
Turning Angle ◽  
Pressure Losses ◽  
Blade Tip ◽  
Annular Cascade ◽  
Highly Loaded

Annular cascade tests were carried out to study the performance of an ultra-highly loaded turbine cascade (UHLTC) with a design turning angle of 160 deg. The UHLTC is for applications to future high-temperature gas turbine engines. This paper describes details of the secondary flows and the associated total pressure losses of the UHLTC obtained at a test incidence of −2.7 deg. The cascade flows were measured with a small five-hole Pitot probe located at 21 traverse measurement planes upstream, inside and downstream of the UHLTC. From the measurements, detailed flow structures and the loss evolution process were analyzed. Flow visualization tests were also carried out to see more details of the flows on the blade surfaces, on the endwalls and in the blade tip gap. Various flow separations and various small vortices associated with the passage and leakage vortices, such as corner vortices and edge vortices, separation bubbles, and the associated reverse flows, were seen. These were clarified from various flow lines showing separation, attachment/reattachment and division of each flow. The results obtained from the flow visualization were compared with those from the traverse measurements. Large total pressure losses occur inside the cascade passage as well as downstream of the cascade. Various strong passage vortices, strong leakage vortex, strong swirling flows upstream and downstream of the cascade and their associated various flow separations, are the main causes of the loss generation. The coefficient of total pressure loss generated inside the cascade was 0.28 at the test near-design incidence. The actual turning angle of the flow from the cascade inlet and the cascade outlet was 146 deg. Some schematic drawings of the flow structures in the present UHLTC were also given. The basic flow structures did not differ significantly from those seen in the conventional cascades with much smaller turning angles, except for stronger passage vortices, larger internal loss and larger downstream mixing loss due to the very high turning angle of the UHLTC.

Application of Endwall Contouring to Transonic Turbine Cascades: Experimental Measurements at Design Conditions

2011 ◽  
Author(s):  
F. Taremi ◽  
S. A. Sjolander ◽  
T. J. Praisner
Keyword(s):  
Kinetic Energy ◽  
Flow Visualization ◽  
Flow Regime ◽  
Total Pressure ◽  
Surface Flow ◽  
Secondary Flows ◽  
Experimental Results ◽  
Pressure Losses ◽  
Turbine Cascades ◽  
Endwall Contouring

An experimental investigation of the endwall flows in two transonic linear turbine cascades was presented at the 2010 ASME Turbo Expo (GT2010–22760). Endwall contouring was subsequently implemented in these cascades to control the secondary flows, and reduce the total pressure losses. The current paper presents experimental results from these cascades to assess the effectiveness of endwall contouring in the transonic flow regime. The experimental results include blade loadings, total pressure losses, streamwise vorticity and secondary kinetic energy distributions. In addition, surface flow visualization results are presented in order to interpret the endwall limiting streamlines within the blade passages. The flat-endwall and contoured-endwall cascades produce very similar midspan loading distributions and profile losses, but exhibit different secondary flows. The endwall surface flow visualization results indicate weaker interaction between the secondary flows and the blade suction surface boundary layers in the contoured cascades. Overall, the implementation of endwall contouring results in smaller and less intense vortical structures, and the reduction of the associated secondary kinetic energy (SKE) and exit flow angle variations. However, the mass-averaged losses at the main measurement plane, located 40% axial chord lengths downstream of the cascade (1.4CX), do not corroborate the numerically predicted improvements for the contoured cascades. This is in part attributed to slower mixing rates of the secondary flows in the compressible flow regime. The mass-averaged results at 2.0CX, on the other hand, show smaller losses for the contoured cascades associated with smaller SKE dissipation downstream of the cascades. Accordingly, the mixed-out row losses also show improvements for the contoured cascades.

Migration of Combustor Exit Profiles Through High Pressure Turbine Vanes

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka ◽  
J. P. Clark ◽  
P. J. Koch
Keyword(s):  
High Pressure ◽  
Secondary Flow ◽  
Total Pressure ◽  
Secondary Flows ◽  
Working Fluid ◽  
Outer Diameter ◽  
High Pressure Turbine ◽  
Flow Angle ◽  
Radial Temperature ◽  
Turbine Vane

The high pressure turbine stage within gas turbine engines is exposed to combustor exit flows that are nonuniform in both stagnation pressure and temperature. These highly turbulent flows typically enter the first stage vanes with significant spatial gradients near the inner and outer diameter endwalls. These gradients can result in secondary flow development within the vane passage that is different than what classical secondary flow models predict. The heat transfer between the working fluid and the turbine vane surface and endwalls is directly related to the secondary flows. The goal of the current study was to examine the migration of different inlet radial temperature and pressure profiles through the high turbine vane of a modern turbine engine. The tests were performed using an inlet profile generator located in the Turbine Research Facility at the Air Force Research Laboratory. Comparisons of area-averaged radial exit profiles are reported as well as profiles at three vane pitch locations to document the circumferential variation in the profiles. The results show that the shape of the total pressure profile near the endwalls at the inlet of the vane can alter the redistribution of stagnation enthalpy through the airfoil passage significantly. Total pressure loss and exit flow angle variations are also examined for the different inlet profiles.

Application of Endwall Contouring to Transonic Turbine Cascades: Experimental Measurements at Design Conditions

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
Farzad Taremi ◽  
Steen A. Sjolander ◽  
Thomas J. Praisner
Keyword(s):  
Kinetic Energy ◽  
Flow Visualization ◽  
Flow Regime ◽  
Total Pressure ◽  
Surface Flow ◽  
Secondary Flows ◽  
Pressure Losses ◽  
Turbine Cascades ◽  
Endwall Contouring ◽  
Surface Flow Visualization

An experimental investigation of the endwall flows in two high-turning turbine cascades was presented by Taremi et al. (2010, “Measurements of Endwall Flows in Transonic Linear Turbine Cascades: Part II—High Flow Turning,” ASME Conf. Proc., GT2010-22760, pp. 1343–1356). Endwall contouring was subsequently implemented in these cascades to control the secondary flows and reduce the total pressure losses. The current paper presents experimental results from these cascades to assess the effectiveness of endwall contouring in the transonic flow regime. The results include blade loadings, total pressure losses, streamwise vorticity and secondary kinetic energy distributions. In addition, surface flow visualization results are presented in order to interpret the endwall limiting streamlines within the blade passages. The flat-endwall and contoured-endwall cascades produce very similar midspan loading distributions and profile losses, but exhibit different secondary flows. The endwall surface flow visualization results indicate weaker interaction between the secondary flows and the blade suction surface boundary layers in the contoured cascades. Overall, the implementation of endwall contouring results in smaller and less intense vortical structures, and the reduction of the associated secondary kinetic energy (SKE) and exit flow angle variations. However, the mass-averaged losses at the main measurement plane, located 40% axial chord lengths downstream of the cascade (1.4CX), do not corroborate the numerically predicted improvements for the contoured cascades. This is in part attributed to slower mixing rates of the secondary flows in the compressible flow regime. The mass-averaged results at 2.0CX, on the other hand, show smaller losses for the contoured configurations associated with smaller SKE dissipation downstream of the cascades. Accordingly, the mixed-out row losses also show improvements for the contoured cascades.

Effect of the turning angle on the flow and performance characteristics of long S-shaped circular diffusers

2003 ◽  
Vol 217 (1) ◽  
pp. 29-41 ◽  
Author(s):  
R B Anand ◽  
L Rai ◽  
S N Singh
Keyword(s):  
Total Pressure ◽  
Static Pressure ◽  
Area Ratio ◽  
Secondary Flows ◽  
Performance Characteristics ◽  
Pressure Recovery ◽  
Recovery Coefficient ◽  
Turning Angle ◽  
And Performance ◽  
Pressure Recovery Coefficient

The effect of the turning angle on the flow and performance characteristics of long S-shaped circular diffusers (length-inlet diameter ratio, L/Di = 11:4) having an area ratio of 1.9 and centre-line length of 600 mm has been established. The experiments are carried out for three S-shaped circular diffusers having angles of turn of 15°/15°, 22.5°/22.5° and 30°/30°. Velocity, static pressure and total pressure distributions at different planes along the length of the diffusers are measured using a five-hole impact probe. The turbulence intensity distribution at the same planes is also measured using a normal hot-wire probe. The static pressure recovery coefficients for 15°/15°, 22.5°/22.5° and 30°/30° diffusers are evaluated as 0.45, 0.40 and 0.35 respectively, whereas the ideal static pressure recovery coefficient is 0.72. The low performance is attributed to the generation of secondary flows due to geometrical curvature and additional losses as a result of the high surface roughness (~0.5 mm) of the diffusers. The pressure recovery coefficient of these circular test diffusers is comparatively lower than that of an S-shaped rectangular diffuser of nearly the same area ratio, even with a larger turning angle (90°/90°), i.e. 0.53. The total pressure loss coefficient for all the diffusers is nearly the same and seems to be independent of the angle of turn. The flow distribution is more uniform at the exit for the higher angle of turn diffusers.

Second Vane Total Pressure Loss Due to Endwall Iceform Contouring

2008 ◽  
Author(s):  
Ronald S. LaFleur
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Drag ◽  
Pressure Coefficient ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Shape ◽  
Exit Plane

The iceformation design method generates an endwall contour, altering the secondary flows that produce elevated endwall heat transfer load and total pressure losses. Iceformation is an analog to regions of metal melting where a hot fluid alters the isothermal surface shape of a part as it is maintained by a cooling fluid. The passage flow, heat transfer and geometry evolve together under the constraints of flow and thermal boundary conditions. The iceformation concept is not media dependent and can be used in analogous flows and materials to evolve novel boundary shapes. In the past, this method has been shown to reduce aerodynamic drag and total pressure loss in flows such as diffusers and cylinder/endwall junctures. A prior paper [1] showed that the Reynolds number matched iceform geometry had a 24% lower average endwall heat transfer than the rotationally symmetric endwall geometry of the Energy Efficiency Engine (E3). Comparisons were made between three endwall geometries: the ‘iceform’, the ‘E3’ and the ‘flat’ as a limiting case of the endwall design space. This paper adds to the iceformation design record by reporting the endwall aerodynamic performances. Second vane exit flow velocities and pressures were measured using an automated 2-D traverse of a 1.2 mm diameter five-hole probe. Exit plane maps for the three endwall geometries are presented showing the details of the total pressure coefficient contours and the velocity vectors. The formation of secondary flow vortices is shown in the exit plane and this results in an impact on exit plane total pressure loss distribution, off-design over- and under-turning of the exit flow. The exit plane contours are integrated to form overall measures of the total pressure loss. Relative to the E3 endwall, the iceform endwall has a slightly higher total pressure loss attributed to higher dissipation of the secondary flow within the passage. The iceform endwall has a closer-to-design exit flow pattern than the E3 endwall.

scholarly journals Numerical and Experimental Analysis of Secondary Flow in Modern State-of-the-Art Low Pressure Guide Vane Rows

1995 ◽  
Author(s):  
Ernst Lindner
Keyword(s):  
Aspect Ratio ◽  
Secondary Flow ◽  
Total Pressure ◽  
Guide Vane ◽  
Numerical Results ◽  
Flow Behaviour ◽  
Inlet Guide Vane ◽  
Streamwise Vorticity ◽  
Pressure Losses ◽  
Turbine Guide Vane

To enhance the performance of the inlet guide vane and the annular duct of a jet engine, a detailed investigation of annular cascades with two different types of turbine guide vane rows is made. The first one is a leaned guide vane with an aspect ratio of two and a half and a transition duct ahead of the vane. To avoid the losses associated to the decelerating transition duct an alternative vane is designed and investigated with the same inlet and exit conditions. In this case the chord of the vane is increased to the effect that the vane begins immediately at the enterance of the diverging annulus and so a continuously accelerated flow is achieved. To maintain a good performance for this configuration a bowed-type vane with an aspect ratio of one is designed. The aim of the investigation is to obtain detailed informations on the secondary flow behaviour with particular regard to the development of the total pressure losses and the streamwise vorticity of the vortices inside and behind the blade rows. In the first step a three-dimensional, structured, explicit finite-volume flow-solver with a k–ε turbulence model is validated against the measurements, which were made in cross-sections behind the blades. Having proved that the numerical results are very close to the experimental ones, the secondary flow behaviour inside and behind the blade rows is analysed in the second step. By calculating the streamwise vorticity from the numerical results the formation of horse-shoe vortex, passage-vortex and the trailing edge vortex shed is investigated. The differences of the vortical motion and the formation of the total pressure losses between the two configurations of turbine guide vane rows are discussed.

Production and Development of Secondary Flows and Losses in Two Types of Straight Turbine Cascades: Part 2—A Rotor Case

1987 ◽  
Vol 109 (2) ◽  
pp. 194-200 ◽  
Author(s):  
A. Yamamoto
Keyword(s):  
Pressure Gradient ◽  
Production Process ◽  
Experimental Analysis ◽  
Significant Contribution ◽  
Adverse Pressure Gradient ◽  
Secondary Flows ◽  
Suction Surface ◽  
Turning Angle ◽  
Detailed Mechanism ◽  
Turbine Cascades

Part 1 of this paper [1] presents the detailed mechanism of secondary flows and the associated losses occurring within a straight stator cascade with a relatively low turning angle of about 65 deg. The significant contribution of secondary flows on the loss production process was shown only near the blade suction surface downstream from the cascade throat (Z/Cax = 0.74) in which regional flows decelerated due to adverse pressure gradient. In the second part, the same experimental analysis is applied to a straight rotor cascade with a much larger turning angle of 102 deg. Flow surveys were made at 12 traverse planes located throughout the rotor cascade. The larger turning results in a similar but much stronger contribution of the secondary flows to the loss developing mechanism. Evolution of overall loss starts quite early within the cascade, and the rate of the loss growth is much larger in the rotor case than in the stator case.

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