Control of the Corner Separation in a Linear Cascade by Trailing Gaps

2017 ◽  
Author(s):  
Wenfeng Zhao ◽  
Bin Jiang ◽  
Qun Zheng
Keyword(s):  
Flow Field ◽  
Suction Surface ◽  
Optimal Length ◽  
Corner Separation ◽  
Linear Cascade ◽  
High Loss ◽  
Axial Compressors ◽  
Pressure Surface ◽  
Sst Turbulence Model ◽  
The Stability

Hub corner is the high-loss area in the blade passages of turbo machinery. It is well known that the flow separation and vortex development in this area affects directly not only the energy losses and efficiency, but also the stability of axial compressors. Linear compressor cascades with partial gaps and trailing gaps which can blow away the corner separation by the pressure difference between the suction surface and pressure surface are numerically simulated in this paper. A proposed linear cascade model with gaps has been built. The steady flow field in a linear cascade with different length gaps is studied by numerical simulation of RANS with SST turbulence model and γ-Reθ transition model focusing on the streamline structure between the corner separation vortex and the gap leakage vortex, especially the interaction of the two vertical vortex. When the length of trailing edge gaps is enough (in this paper, the optimal length of the gap is 30% chord), the corner vortex basically disappears completely. At the same time, the mode of flow field changes from the closed separation to the open separation.

Flow Structure and Unsteady Behavior of Hub-Corner Separation in a Stator Cascade of a Multi-Stage Transonic Axial Compressor

2018 ◽  
Author(s):  
Seishiro Saito ◽  
Kazutoyo Yamada ◽  
Masato Furukawa ◽  
Keisuke Watanabe ◽  
Akinori Matsuoka ◽  
...  
Keyword(s):  
Data Mining ◽  
Flow Field ◽  
Unsteady Flow ◽  
Axial Compressor ◽  
Flow Fields ◽  
Suction Surface ◽  
Corner Separation ◽  
High Loss ◽  
Stator Blade ◽  
Transonic Axial Compressor

This paper describes unsteady flow phenomena of a two-stage transonic axial compressor, especially the flow field in the first stator. The stator blade with highly loaded is likely to cause a flow separation on the hub, so-called hub-corner separation. The flow mechanism of the hub-corner separation in the first stator is investigated in detail using a large-scale detached eddy simulation (DES) conducted for its full-annulus and full-stage with approximately 4.5 hundred million computational cells. The detailed analysis of complicated flow fields in the compressor is supported by data mining techniques. The data mining techniques applied in the present study include vortex identification based on the critical point theory and topological analysis of the limiting streamline pattern. The simulation results show that the flow field in the hub-corner separation is dominated by a tornado-type separation vortex. In the time averaged flow field, the hub-corner separation vortex rolls up from the hub wall, which is generated by the interaction between the mainstream flow, the leakage flow from the front partial clearance and the secondary flow across the blade passage toward the stator blade suction side. The hub-corner separation vortex suffers a vortex breakdown near the mid chord, where the high loss region due to the hub-corner separation expands drastically. In the rear part of the stator passage, a high loss region is migrated radially outward by the induced velocity of the hub-corner separation vortex. The flow field in the stator is influenced by the upstream and downstream rotors, which makes it difficult to understand the unsteady effects. The unsteady flow fields are analyzed by applying the phase-locked ensemble averaging technique. It is found from the phase-locked flow fields that the wake interaction from the upstream rotor has more influence on the stator flow field than the shock wave interaction from the downstream rotor. In the unsteady flow field, a focal-type separation also emerges on the blade suction surface, but it is periodically swept away by the wake passing of the upstream rotor. The separation vortex on the hub wall connects with the one on the blade suction surface, forming an arch-like vortex.

Numerical Investigation of a Cantilevered Compressor Stator at Varying Clearance Sizes

2015 ◽  
Author(s):  
Chengwu Yang ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
Shengfeng Zhao ◽  
Junqiang Zhu
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
High Flow ◽  
Leakage Flow ◽  
Streamwise Direction ◽  
Stall Margin ◽  
Axial Compressors ◽  
Pressure Surface ◽  
The Impact ◽  
Highly Loaded

The clearance size of cantilevered stators affects the performance and stability of axial compressors significantly. Numerical calculations were carried out using the commercial software FINE/Turbo for a 2.5-stage highly loaded transonic axial compressor, which is of cantilevered stator for the first stage, at varying hub clearance sizes. The aim of this work is to improve understanding of the impact mechanism of hub clearance on the performance and the flow field in high flow turning conditions. The performance of the front stage and the compressor with different hub clearance sizes of the first stator has been analyzed firstly. Results show that the efficiency decreases as clearance size varies from 0 to 3% of hub chordlength, but the operating range has been extended. For the first stage, the efficiency decreases about 0.5% and the stall margin is extended. The following analysis of detailed flow field in the first stator shows that the clearance leakage flow and elimination of hub corner separation is responsible for the increasing loss and stall margin extending respectively. The effects of hub clearance on the downstream rotor have been discussed lastly. It indicates that the loss of the rotor increases and the flow deteriorates due to increasing of clearance size and hence the leakage mass flow rate, which mainly results from the interaction of upstream leakage flow with the passage flow near pressure surface. The affected region of rotor passage flow field expands in spanwise and streamwise direction as clearance size grows. The hub clearance leakage flow moves upward in span as it flows toward downstream.

scholarly journals Unsteady Flow Field due to Nozzle Wake Interaction With the Rotor in an Axial Flow Turbine: Part I — Rotor Passage Flow Field

1995 ◽  
Author(s):  
Michael A. Zaccaria ◽  
Budugur Lakshminarayana
Keyword(s):  
Flow Field ◽  
Unsteady Flow ◽  
Axial Flow ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Pressure Surface ◽  
Aerodynamic Interaction ◽  
Flow Gradients ◽  
Rotor Flow

The flow field in turbine rotor passages is complex with unsteadiness caused by the aerodynamic interaction of the nozzle and rotor flow fields. The two-dimensional steady and unsteady flow field at midspan in an axial flow turbine rotor has been investigated experimentally using an LDV with emphasis on the interaction of the nozzle wake with the rotor flow field. The flow field in the rotor passage is presented in Part I, while the flow field downstream of the rotor is presented in Part II. Measurements were acquired at 37 axial locations from just upstream of the rotor to one chord downstream of the rotor. The time average flow field and the unsteadiness caused by the wake has been captured. As the nozzle wake travels through the rotor flow field, the nozzle wake becomes distorted with the region of the nozzle wake near the rotor suction surface moving faster than the region near the rotor pressure surface, resulting in a highly distorted wake. The wake is found to be spread out along the rotor pressure surface, as it convects downstream of midchord. The magnitude of the nozzle wake velocity defect grows until close to midchord, after which it decreases. High values of unresolved unsteadiness were observed at the rotor leading edge. This is due to the large flow gradients near the leading edge and the interaction of the nozzle wake with the rotor leading edge. High values of unresolved unsteadiness were also observed near the rotor pressure surface. This increase in unresolved unsteadiness is caused by the interaction of the nozzle wake with the flow near the rotor pressure surface.

Unsteady Flow in Oscillating Turbine Cascade: Part 1 — Linear Cascade Experiment

1996 ◽  
Author(s):  
L. He
Keyword(s):  
Computational Study ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Leading Edge ◽  
Computational Method ◽  
Reattachment Point ◽  
Suction Surface ◽  
Linear Cascade ◽  
Unsteady Pressure ◽  
Pressure Surface

An experimental and computational study has been carried out on a linear cascade of low pressure turbine blades with the middle blade oscillating in a torsion mode. The main objectives of the present work were to enhance understanding of the behaviour of bubble type of flow separation and to examine the predictive ability of a computational method. In addition, an attempt was made to address a general modelling issue: was the linear assumption adequately valid for such kind of flow? In Part 1 of this paper, the experimental work was described. Unsteady pressure was measured along blade surfaces using off-board mounted pressure transducers at realistic reduced frequency conditions. A short separation bubble on the suction surface near the trailing edge and a long leading-edge separation bubble on the pressure surface were identified. It was found that in the regions of separation bubbles, unsteady pressure was largely influenced by the movement of reattachment point, featured by an abrupt phase shift and an amplitude trough in the 1st harmonic distribution. The short bubble on the suction surface seemed to follow closely a laminar bubble transition model in a quasi-steady manner, and had a localized effect. The leading-edge long bubble on the pressure surface, on the other hand, was featured by a large movement of the reattachment point, which affected the surface unsteady pressure distribution substantially. As far as the aerodynamic damping was concerned, there was a destabilizing effect in the separated flow region, which was however largely balanced by the stabilizing effect downstream of the reattachment point due to the abrupt phase change.

scholarly journals Secondary Flow Reduction in a Nozzle Guide Vane Cascade by Non-Axisymmetric End-Wall Profiling

1999 ◽  
Author(s):  
J. Yan ◽  
D. G. Gregory-Smith ◽  
P. J. Walker
Keyword(s):  
Secondary Flow ◽  
Static Pressure ◽  
Guide Vane ◽  
Suction Surface ◽  
Linear Cascade ◽  
Pressure Surface ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
End Wall

A linear cascade of HP steam turbine nozzle guide vanes was designed and built in order to study the effect of a non-axisymmetric profile for the endwall. The profile was designed by using CFD for the purpose of reducing the secondary flow. The method was to use convex curvature near the pressure surface to reduce the static pressure and concave curvature near the suction surface to increase it. Thus the cross passage pressure gradient which drives the secondary flow would be reduced. Detailed investigations of the flow field with a flat end-wall and the profiled end-wall were conducted. The effect of the profiled end-wall on the secondary flow development was determined and also compared with the CFD design predictions. It was found that the secondary loss and secondary kinetic energy were both reduced by about 20% with the shaped endwall, and a more uniform exit flow was also achieved.

Design Optimization of a Blended Blade and Endwall in a Compressor Cascade

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
Jiabin Li ◽  
Lucheng Ji ◽  
Ling Zhou
Keyword(s):  
Pressure Difference ◽  
Axial Position ◽  
Design Parameters ◽  
Support Vector ◽  
Optimization Approach ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Corner Separation ◽  
High Loss ◽  
Transverse Pressure

Abstract The blended blade and endwall (BBEW) contouring technology can adjust the dihedral angle between suction surface and endwall, thus reducing corner separation in compressors. Generally, the design of BBEW relies on the experiences, the effective design results may not be the optimal result. In this paper, an optimization approach based on the genetic algorithm (GA) for feature selection and parameter optimization of support vector machine (SVM) is used to obtain the optimal BBEW parameters in a compressor cascade. Based on the sensitivity analysis of the results, it is found that the maximum blended width and the axial position of the maximum blended width are the two most important design parameters. The experimental results show that the optimal BBEW cascade can stretch the spanwise area of the high loss region, and reduce the maximum value in it. The numerical studies were conducted to analyze the flow mechanism. It is shown that the BBEW cascade has a transverse pressure difference at the axial position of the maximum blended width, and magnitude of the pressure difference in proportion to the maximum blended width. The transverse pressure difference removes the low-energy fluid from the corner to the main flow, thus improving the corner separation.

Experimental and Numerical Investigation on the Aerodynamic Performance of a Compressor Cascade Using Blended Blade and End Wall

2017 ◽  
Author(s):  
Jiabin Li ◽  
Lucheng Ji ◽  
Weilin Yi
Keyword(s):  
Wind Tunnel ◽  
Flow Field ◽  
Turbulence Model ◽  
Numerical Study ◽  
Incidence Angle ◽  
Good Effect ◽  
Compressor Cascade ◽  
Corner Separation ◽  
Sst Turbulence Model ◽  
End Wall

Nowadays, the corner separation, occurring near the corner region formed by the suction surface of blade and end wall, has been an important limitation for the increasing of the aerodynamic loading in the compressor. The previous numerical studies indicate that the Blended Blade and End Wall (BBEW) technology is useful in delaying, or reducing, or even eliminating the corner separation. To further validate the concept, this paper presents combined experimental and numerical investigations on a BBEW cascade and its prototype. Firstly, the NACA65 linear compressor cascade with the turning angle 42 degrees was designed and tested in a low-speed wind tunnel. Then, the cascade with blended blade and end wall design was made and tested in the same wind tunnel. The experimental results show that the design of blended blade and end wall can improve the performance of the cascade when the incidence angle was positive or at the design point, and the total pressure loss coefficient was reduced by 7%–8%. The performance improvement mainly located from 10%–25% span heights. Secondly, based on the experimental data, the numerical study made by our internal code Turbo-CFD shows the difference of the simulation precision of the results, obtained from four different turbulence model after the mesh independence test. The four turbulence model is Spalart-Allmaras model, standard k-ε model, standard k-ω model, and shear stress transport k-ω model. For this case, the SST turbulence model has better performance compared with others. Thirdly, based on the results which were calculated with the turbulence model SST, the effect of the blended blade and end wall design was discussed. The numerical study shows that the design with the blended blade and end wall can have a good effect on the corner flow of the cascade. The strong three-dimensional corner separation, caused by the accumulation of the flow happening at the trail of the suction side was avoided, and the flow losses of the prototype cascade were reduced. Above all, the experiment shows that the design with blended blade and end wall can improve the performance of the cascade. Compared with the experiment data, the SST turbulence model shows the best results of the flow field. Based on the numerical results, the details of the flow field and the effect of the blended blade and end wall design on the corner separation are discussed and analyzed.

The Mechanism of the Flow in the Hub Corner and the Control by Tailing Edge Gaps

2018 ◽  
Author(s):  
Wenfeng Zhao ◽  
Bin Jiang ◽  
Yu Duan ◽  
Zhitao Tian ◽  
Qun Zheng
Keyword(s):  
Flow Field ◽  
Pressure Gradient ◽  
Pressure Ratio ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Turn Angle ◽  
Corner Separation ◽  
Transverse Pressure ◽  
Streamwise Pressure Gradient ◽  
Surface Vortex

High-pressure ratio is one of the important characteristics of the sustainable development of the modern aero-engine compressor components. When the fluid flows through the compressor cascade row, it will be influenced by both the streamwise pressure gradient and the transverse pressure gradient, which will cause hub-corner separation or stall. In this paper, different diffusion factors are chosen for the cascades. Each diffusion factor has different turning angles. The formation mechanism of hub-corner separation is studied under the condition of zero angle of attack. Numerical simulation is used to study the influence of pressure gradient on the flow field in the corner. The scale of the concentrated shed vortex forms in the suction surface increases with the increasing of the transverse pressure gradient during the hub-corner separation. When the streamwise pressure gradient increases, the suction surface vortex forms the corner stall. By reasonable design, the two vortexes can cancel out each other. At this time, the loss of cascades is the minimum. Based on the flow mechanism of the corner separation/stall, the trailing gaps are set on three typical turn angle cascades. The results show that the trailing gaps can control the radial development of the suction surface vortex during the stall and improve flow field. The jet cannot blow the suction side boundary layer away during the corner separation, because the gap does not change the static pressure distribution at the root of the cascade. In a word, the trailing edge gaps can not only inhibit the separation in the hub corner but also have the minimum leakage loss at design point. It can be used as an effective and practical compressor design method.

scholarly journals Unsteady Flow Field Due to Nozzle Wake Interaction With the Rotor in an Axial Flow Turbine: Part I—Rotor Passage Flow Field

1997 ◽  
Vol 119 (2) ◽  
pp. 201-213 ◽  
Author(s):  
M. A. Zaccaria ◽  
B. Lakshminarayana
Keyword(s):  
Flow Field ◽  
Unsteady Flow ◽  
Axial Flow ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Pressure Surface ◽  
Aerodynamic Interaction ◽  
Flow Gradients ◽  
Rotor Flow

The flow field in turbine rotor passages is complex with unsteadiness caused by the aerodynamic interaction of the nozzle and rotor flow fields. The two-dimensional steady and unsteady flow field at midspan in an axial flow turbine rotor has been investigated experimentally using an LDV with emphasis on the interaction of the nozzle wake with the rotor flow field. The flow field in the rotor passage is presented in Part I. while the flow field downstream of the rotor is presented in Part II. Measurements were acquired at 37 axial locations from just upstream of the rotor to one chord downstream of the rotor. The time-averaged flow field and the unsteadiness caused by the wake have been captured. As the nozzle wake travels through the rotor flow field, the nozzle wake becomes distorted with the region of the nozzle wake near the rotor suction surface moving faster than the region near the rotor pressure surface, resulting in a highly distorted wake. The wake is found to be spread out along the rotor pressure surface, as it convects downstream of midchord. The magnitude of the nozzle wake velocity defect grows until close to midchord, after which it decreases. High values of unresolved unsteadiness were observed at the rotor leading edge. This is due to the large flow gradients near the leading edge and the interaction of the nozzle wake with the rotor leading edge. High values of unresolved unsteadiness were also observed near the rotor pressure surface. This increase in unresolved unsteadiness is caused by the interaction of the nozzle wake with the flow near the rotor pressure surface.

Numerical Simulation of Influence with Surface Contamination on Aerodynamic Performance of Dedicated Wind Turbine Airfoil

2013 ◽  
Vol 724-725 ◽  
pp. 572-575
Author(s):  
Pan Wu ◽  
Chun Li ◽  
Zhi Min Li
Keyword(s):  
Numerical Simulation ◽  
Wind Turbine ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Surface Contamination ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Pressure Surface ◽  
Sst Turbulence Model ◽  
Wind Turbine Airfoil

A Numerical simulation on the influence of airfoil surface contamination on the aerodynamic performance of wind turbines was performed. It chose the dedicated wind turbine airfoil as the research object. The k-ω Shear Stress Transmission (SST) turbulence model was selected for CFD calculation. The roughness height which arranged evenly on the airfoil was changed from 0.03mm to 2.0mm to obtain the sensitive roughness. The airfoil was divided into 18 sections for analyzing the effect on the lift & the drag coefficient, due to various locations of sensitive roughness. By comparing the result computed by XFOIL and CFD calculation, it can be known this airfoils sensitive locations in suction surface and pressure surface. The sensitive locations in suction surface were 53% and 92% from the chord line towards the leading edge, while 44% and 88% in pressure surface. The sensitive roughness in sensitive locations delayed the location of the transition point.

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