Growth of Secondary Losses and Vorticity in an Axial Turbine Cascade

1988 ◽  
Vol 110 (1) ◽  
pp. 1-8 ◽  
Author(s):  
D. G. Gregory-Smith ◽  
C. P. Graves ◽  
J. A. Walsh
Keyword(s):  
Kinetic Energy ◽  
Flow Visualization ◽  
Secondary Flow ◽  
Surface Flow ◽  
Turbine Cascade ◽  
Streamwise Vorticity ◽  
Blade Passage ◽  
Positive Vorticity ◽  
Passage Vortex

The growth of losses, secondary kinetic energy, and streamwise vorticity have been studied in a high turning rotor cascade. Negative vorticity associated with the passage vortex agreed well with predictions of classical secondary flow theory in the early part of the blade passage. However, toward the exit, the distortion of the flow by the secondary velocities rendered the predictions inaccurate. Areas of positive vorticity were associated with the feeding of loss into the bulk flow and have been related to separation lines observed by surface flow visualization.

Measurements of Endwall Flows in Transonic Linear Turbine Cascades: Part II—High Flow Turning

2010 ◽  
Author(s):  
F. Taremi ◽  
S. A. Sjolander ◽  
T. J. Praisner
Keyword(s):  
Mach Number ◽  
Kinetic Energy ◽  
Flow Visualization ◽  
Flow Separation ◽  
Surface Flow ◽  
Streamwise Vorticity ◽  
Vortical Structures ◽  
Turbine Cascades ◽  
Mach Numbers ◽  
Surface Flow Visualization

An experimental investigation of two low-turning (90°) transonic linear turbine cascades was presented in Part I of the paper. Part II examines two high-turning (112°) turbine cascades. The experimental results include total pressure losses, streamwise vorticity and secondary kinetic energy distributions. The measurements were made using a seven-hole pressure probe downstream of the cascades. In addition to the measurements, surface flow visualization was conducted to assist in the interpretation of the flow physics. The turbine cascades in Part II, referred to as SL1F and SL2F, have the same inlet and outlet design flow angles, but different aerodynamic loading levels: SL2F is more highly loaded than SL1F. The surface flow visualization results show evidence of small flow separation on the suction side of both airfoils. At the design conditions (outlet Mach number ≈ 0.8), SL2F exhibits stronger vortical structures and larger secondary velocities than SL1F. The two cascades, however, produce similar row losses based on the measurements at 40% axial chord lengths downstream of the trailing edge. Additional data were collected at off-design outlet Mach numbers of 0.65 and 0.91. As the Mach number is raised, the cascades become more aft-loaded. The absolute blade loadings increase, but the Zweifel coefficients decrease due to higher outlet dynamic pressures. Both profile and secondary losses decrease at higher Mach numbers; the main vortical structures and the corresponding peak losses migrate towards the endwall, and there are reductions in secondary kinetic energy and exit flow angle variations. The streamwise vorticity distributions show smaller peak vorticities associated with the passage and the counter vortices at higher exit Mach numbers. The corner vortex, on the other hand, becomes more intensified, resulting in reduction of flow overturning near the endwall. The results for SL1F and SL2F are compared and contrasted with the results for the lower turning cascades presented in Part I. The possible effects of suction-surface flow separation on profile and secondary losses are discussed in this context. The current research project is part of a larger study concerning the effects of endwall contouring on secondary losses, which will be presented in the near future.

Effects of an Upstream Cavity on the Secondary Flow in a Transonic Turbine Cascade

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
H. M. Abo El Ella ◽  
S. A. Sjolander ◽  
T. J. Praisner
Keyword(s):  
Mach Number ◽  
Secondary Flow ◽  
Ultra Violet ◽  
Leading Edge ◽  
Surface Flow ◽  
Flow Structures ◽  
Pressure Probe ◽  
Turbine Cascade ◽  
Streamwise Vorticity ◽  
Blow Down

This paper examines experimentally the effects of an upstream cavity on the flow structures and secondary losses in a transonic linear turbine cascade. The cavity approximates the endwall geometry resulting from the platform overlap at the interface between stationary and rotating turbine blade rows. Previous investigations of the effects of upstream cavity geometries have been conducted mainly at low-speed conditions. The present work aims to extend such research into the transonic regime with a more engine representative upstream platform geometry. The investigations were carried out in a blow-down type wind tunnel. The cavity is located at 30 % of axial chord from the leading edge, extends 17 % of axial-chord in depth, and is followed by a smooth ramp to return the endwall to its nominal height. Two cascades are examined for the same blade geometry: the baseline cascade with a flat endwall and the cascade with the cavity endwall. Measurements were made at the design incidence and the outlet design Mach number of 0.80. At this condition, the Reynolds number based on outlet velocity is about 600,000. Off-design outlet Mach numbers of 0.69, and 0.89 were also investigated. Flowfield measurements were carried out at 40 % axial-chord downstream of the trailing edge, using a seven-hole pressure probe, to quantify losses and identify the flow structures. Additionally, surface flow visualization using an ultra-violet reactive dye was employed at the design Mach number, on the endwall and blade surfaces, to help in the interpretation of the flow physics. The experimental results also include blade-loading distributions, and the probe measurements were processed to obtain total-pressure loss coefficients, and streamwise vorticity distributions. It was found that the presence of the upstream cavity noticeably altered the structure and the strength of the secondary flow. Some effect on the secondary losses was also evident, with the cavity having a larger effect at the higher Mach number.

Measurements of Secondary Losses in a Turbine Cascade With the Implementation of Nonaxisymmetric Endwall Contouring

2009 ◽  
Vol 132 (1) ◽  
Author(s):  
D. C. Knezevici ◽  
S. A. Sjolander ◽  
T. J. Praisner ◽  
E. Allen-Bradley ◽  
E. A. Grover
Keyword(s):  
Kinetic Energy ◽  
Axial Flow ◽  
Surface Flow ◽  
Secondary Flows ◽  
Test Facility ◽  
Suction Surface ◽  
Blade Passage ◽  
Low Pressure Turbine ◽  
Endwall Contouring ◽  
Pressure And Velocity Measurements

An approach to endwall contouring has been developed with the goal of reducing secondary losses in highly loaded axial flow turbines. The present paper describes an experimental assessment of the performance of the contouring approach implemented in a low-speed linear cascade test facility. The study examines the secondary flows of a cascade composed of Pratt & Whitney PAKB airfoils. This airfoil has been used extensively in low-pressure turbine research, and the present work adds intrapassage pressure and velocity measurements to the existing database. The cascade was tested at design incidence and at an inlet Reynolds number of 126,000 based on inlet midspan velocity and axial chord. Quantitative results include seven-hole pneumatic probe pressure measurements downstream of the cascade to assess blade row losses and detailed seven-hole probe measurements within the blade passage to track the progression of flow structures. Qualitative results take the form of oil surface flow visualization on the endwall and blade suction surface. The application of endwall contouring resulted in lower secondary losses and a reduction in secondary kinetic energy associated with pitchwise flow near the endwall and spanwise flow up the suction surface within the blade passage. The mechanism of loss reduction is discussed in regard to the reduction in secondary kinetic energy.

Volumetric Velocimetry Measurements of Purge–Mainstream Interaction in a One-Stage Turbine

2021 ◽  
Vol 143 (4) ◽  
Author(s):  
A. J. Carvalho Figueiredo ◽  
B. D. J. Schreiner ◽  
A. W. Mesny ◽  
O. J. Pountney ◽  
J. A. Scobie ◽  
...  
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbines ◽  
Suction Side ◽  
Secondary Flows ◽  
Blade Passage ◽  
One Stage ◽  
Passage Vortex ◽  
Purge Flow ◽  
Secondary Flow Field

Abstract Air-cooled gas turbines employ bleed air from the compressor to cool vulnerable components in the turbine. The cooling flow, commonly known as purge air, is introduced at low radius, before exiting through the rim-seal at the periphery of the turbine discs. The purge flow interacts with the mainstream gas path, creating an unsteady and complex flowfield. Of particular interest to the designer is the effect of purge on the secondary-flow structures within the blade passage, the extent of which directly affects the aerodynamic loss in the stage. This paper presents a combined experimental and computational fluid dynamics (CFD) investigation into the effect of purge flow on the secondary flows in the blade passage of an optically accessible one-stage turbine rig. The experimental campaign was conducted using volumetric velocimetry (VV) measurements to assess the three-dimensional inter-blade velocity field; the complementary CFD campaign was carried out using unsteady Reynolds-averaged Navier–Stokes (URANS) computations. The implementation of VV within a rotating environment is a world first and offers an unparalleled level of experimental detail. The baseline flow-field, in the absence of purge flow, demonstrated a classical secondary flow-field: the rollup of a horseshoe vortex, with subsequent downstream convection of a pressure-side and suction-side leg, the former transitioning in to the passage vortex. The introduction of purge, at 1.7% of the mainstream flowrate, was shown to modify the secondary flow-field by enhancing the passage vortex, in both strength and span-wise migration. The computational predictions were in agreement with the enhancement revealed by the experiments.

Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition

2020 ◽  
Author(s):  
Hoshio Tsujita ◽  
Masanao Kaneko
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Secondary Flow ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Flow Behavior ◽  
Turbine Cascade ◽  
Blade Loading ◽  
Passage Vortex ◽  
Exit Mach Number

Abstract Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.

Effects of an Upstream Cavity on the Secondary Flow in a Transonic Turbine Cascade

2010 ◽  
Author(s):  
H. M. Abo El Ella ◽  
S. A. Sjolander ◽  
T. J. Praisner
Keyword(s):  
Mach Number ◽  
Secondary Flow ◽  
Ultra Violet ◽  
Leading Edge ◽  
Reactive Dye ◽  
Surface Flow ◽  
Flow Structures ◽  
Pressure Probe ◽  
Turbine Cascade ◽  
Blow Down

This paper examines experimentally the effects of an upstream cavity on the flow structures and secondary losses in a transonic linear turbine cascade. The cavity approximates the endwall geometry resulting from the platform overlap at the interface between stationary and rotating turbine blade rows. Previous investigations of the effects of upstream cavity geometries have been conducted mainly at low-speed conditions. The present work aims to extend such research into the transonic regime with a more engine representative upstream platform geometry. The investigations were carried out in a blow-down type wind tunnel. The cavity is located at 30% of axial-chord from the leading edge, extends 17% of axial-chord in depth, and is followed by a smooth ramp to return the endwall to its nominal height. Two cascades are examined for the same blade geometry: the baseline cascade with a flat endwall and the cascade with the cavity endwall. Measurements were made at the design incidence and the outlet design Mach number of 0.80. At this condition, the Reynolds number based on outlet velocity is about 600,000. Off-design outlet Mach numbers of 0.69, and 0.89 were also investigated. Flowfield measurements were carried out at 40% axial-chord downstream of the trailing edge, using a seven-hole pressure probe, to quantify losses and identify the flow structures. Additionally, surface flow visualization using an ultra-violet reactive dye was employed at the design Mach number, on the endwall and blade surfaces, to help in the interpretation of the flow physics. The experimental results also include blade-loading distributions, and the probe measurements were processed to obtain total-pressure loss coefficients, and stream-wise vorticity distributions. It was found that the presence of the upstream cavity noticeably altered the structure and the strength of the secondary flow. Some effect on the secondary losses was also evident, with the cavity having a larger effect at the higher Mach number.

Aerodynamic Aspects of Endwall Film-Cooling

1996 ◽  
Author(s):  
S. Friedrichs ◽  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Static Pressure ◽  
Three Dimensional ◽  
Surface Flow ◽  
Flow Visualisation ◽  
Film Cooling Effectiveness ◽  
Blade Passage ◽  
Endwall Film Cooling

This paper describes an investigation of the aerodynamic aspects of endwall film-cooling, in which the flow field downstream of a large-scale low-speed linear turbine cascade has been measured. The integrated losses and locations of secondary flow features with and without end wait film-cooling have been determined for variations of both the coolant supply pressure and injection location. Together with previous measurements of adiabatic film-cooling effectiveness and surface-flow visualisation, these results reveal the nature of the interactions between the ejected coolant and the flow in the blade passage. Measured hole massflows and a constant static pressure mixing analysis, together with the measured losses, allow the decomposition of the losses into three distinct entropy generation mechanisms: loss generation within the hole, loss generation due to the mixing of the coolant with the mainstream, and change in secondary loss generation in the blade passage. Results show that the loss generation within the coolant holes is substantial and that ejection into regions of low static pressure increases the loss per unit coolant massflow. Ejection upstream of the three-dimensional separation lines on the endwall changes secondary flow and reduces its associated losses. The results show that it is necessary to take the three-dimensional nature of the endwall flow into account in the design of endwall film-cooling configurations.

Effects of Simplified Platform Overlap and Cavity Geometry on the Endwall Flow: Measurements and Computations in a Low-Speed Linear Turbine Cascade

2013 ◽  
Author(s):  
G. D. MacIsaac ◽  
S. A. Sjolander ◽  
T. J. Praisner ◽  
E. A. Grover ◽  
R. Jurek
Keyword(s):  
Recirculation Zone ◽  
Secondary Flows ◽  
Pressure Probe ◽  
Computational Domain ◽  
Turbine Cascade ◽  
Streamwise Vorticity ◽  
Flow Measurements ◽  
Passage Vortex ◽  
Flow Features ◽  
Cfd Analyses

Incorporating the platform overlap and endwall cavity into the early stages of turbine CFD analyses is desirable from the perspective of accurately capturing the near endwall flow features. However, the overlap and cavity geometry increase the complexity of the computational domain making CFD meshes more difficult to generate and the CFD solutions more resource intensive. Thus, geometric approximations are often made to simplify the CFD analysis. This paper examines, experimentally, the secondary flows of a linear turbine cascade with three different platform overlap geometries, two of which incorporate geometric simplifications. These are then compared with the corresponding computations. Experimental measurements were collected using a seven-hole pressure probe at a plane located 40% of the axial chord downstream of the trailing edge. Steady-state computational predictions were performed using ANSYS CFX 12.0 and employed the SST transition turbulence model. The experimental results show that the presence of an upstream rim-seal creates a stronger passage vortex, relative to a flat endwall, resulting in larger integrated losses as well as higher levels of secondary kinetic energy and streamwise vorticity. Subtle differences in the strength of the passage vortex and the associated losses are observed for the simplified geometries in both the measured and predicted results. By examining the details of the cavity flow, a recirculation zone is identified which energizes the formation of the passage vortex. The effect of the recirculation zone may be attenuated or intensified by the rim-seal geometry. The paper concludes by addressing the validity and usefulness of the proposed platform overlap simplifications in design-oriented computations.

Aerodynamic Aspects of Endwall Film-Cooling

1997 ◽  
Vol 119 (4) ◽  
pp. 786-793 ◽  
Author(s):  
S. Friedrichs ◽  
H. P. Hodson ◽  
W. N. Dawes
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Large Scale ◽  
Static Pressure ◽  
Three Dimensional ◽  
Surface Flow ◽  
Film Cooling Effectiveness ◽  
Blade Passage ◽  
Generation Mechanisms ◽  
Endwall Film Cooling

This paper describes an investigation of the aerodynamic aspects of endwall film-cooling, in which the flow field downstream of a large-scale low-speed linear turbine cascade has been measured. The integrated losses and locations of secondary flow features with and without endwall film-cooling have been determined for variations of both the coolant supply pressure and injection location. Together with previous measurements of adiabatic film-cooling effectiveness and surface-flow visualization, these results reveal the nature of the interactions between the ejected coolant and the flow in the blade passage. Measured hole massflows and a constant static pressure mixing analysis, together with the measured losses, allow the decomposition of the losses into three distinct entropy generation mechanisms: loss generation within the hole, loss generation due to the mixing of the coolant with the mainstream, and change in secondary loss generation in the blade passage. Results show that the loss generation within the coolant holes is substantial and that ejection into regions of low static pressure increases the loss per unit coolant massflow. Ejection upstream of the three-dimensional separation lines on the endwall changes secondary flow and reduces its associated losses. The results show that it is necessary to take the three-dimensional nature of the endwall flow into account in the design of endwall film-cooling configurations.

Sign in / Sign up

Export Citation Format

Share Document