Vortex Control Over End Wall Flow in Compressor Cascades

1990 ◽  
Author(s):  
Yan-Ping Tang ◽  
Mao-Zhang Chen
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Control Techniques ◽  
Vortex Control ◽  
New Concepts ◽  
Wall Flow ◽  
Linear Compressor Cascade ◽  
End Wall

Three methods of vortex control over the end wall flow in compressor cascades have been investigated experimentally. The total pressure loss at the exit of a linear compressor cascade is reduced 6.5%, 10.5% and 26.5% respectively by these methods for different incidences over a range of moderate-high values. The physics of these methods has been discussed and some new concepts of vortex control techniques in compressor cascades have been proposed.

Investigation on Performance of Compressor Cascade with Tubercle Leading Edge Blade

2020 ◽  
Vol 37 (3) ◽  
pp. 295-303 ◽  
Author(s):  
Tu Baofeng ◽  
Zhang Kai ◽  
Hu Jun
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Large Scale ◽  
Design Method ◽  
Leading Edge ◽  
Experimental Tests ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Linear Compressor Cascade

AbstractIn order to improve compressor performance using a new design method, which originates from the fins on a humpback whale, experimental tests and numerical simulations were undertaken to investigate the influence of the tubercle leading edge on the aerodynamic performance of a linear compressor cascade with a NACA 65–010 airfoil. The results demonstrate that the tubercle leading edge can improve the aerodynamic performance of the cascade in the post-stall region by reducing total pressure loss, with a slight increase in total pressure loss in the pre-stall region. The tubercles on the leading edge of the blades cause the flow to migrate from the peak to the valley on the blade surface around the tubercle leading edge by the butterfly flow. The tubercle leading edge generates the vortices similar to those created by vortex generators, splitting the large-scale separation region into multiple smaller regions.

Effect of End-Wall Vortex Generator Jet Parameters on Flow Control in High Subsonic Compressor Cascade

2018 ◽  
Author(s):  
Cong Chen ◽  
Jianyang Yu ◽  
Fu Chen
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Control Effect ◽  
Pressure Loss Coefficient ◽  
Vortex Generator Jet ◽  
End Wall ◽  
Total Pressure Loss Coefficient

In order to explore the control mechanism of vortex generator jet, which is located in the passage (PVGJ), on the separation flow, the influence of the pitch angle, skew angle, locations and jet-to-inflow ratio are studied using numerical methods in a high subsonic compressor cascade. The changing of the flow pattern is also analyzed in detail. The results show that the control effect of the end-wall vortex generator jet located in the passage is better than the leading edge one and the aerodynamic performance is effectively improved. The maximum total pressure loss coefficient decreases by 12% and the static pressure coefficient increases by 5.2% while the jet-to-inflow ratio is only 0.3%. The control effect is sensitive to the change of jet parameters. When 0 deg < β < 80 deg, 20 deg < α < 50 deg,, x < 0.5B, y < 0.15t, the vortex generation jet could acquire an ideal control effect. As the jet mass increases, the total pressure loss coefficient gradually reduces. The VGJ prevent separation mainly by bringing high momentum fluid into the near wall region and by promoting momentum transport through turbulent mixing in previous studies. Both the LVGJ and PVGJ mainly take advantage of jet vortex to prevent the cross flow from interacting with the suction side boundary layer.

Effect of the end-wall vortex generator jets on the performance of a high subsonic compressor cascade

2018 ◽  
Vol 233 (3) ◽  
pp. 324-336
Author(s):  
Cong Chen ◽  
Huaping Liu ◽  
Fu Chen
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Numerical Study ◽  
Vortex Generator ◽  
Total Pressure Loss ◽  
Experimental Result ◽  
Compressor Cascade ◽  
Design Point ◽  
Vortex Generator Jets ◽  
End Wall

This paper presents a numerical and experimental result of the end-wall vortex generator jets for controlling corner separation and enhancing the aerodynamic performance in a high subsonic (Ma = 0.7) compressor cascade. The experiments were carried out on a compressor cascade at design point ( i = 0°) and off-design points ( i = −2°, 2°, and 4°). At design point, the total pressure loss coefficient could be reduced up to 12.1%.With the increase in the incidence, the control effect is enhanced first and then reduced. The maximum total pressure loss reduction is up to 14.6% when the incidence is 2°. The numerical study is further conducted to analyze the flow pattern and the vortex structure. The jet vortex is formed downstream of the jet hole using the vortex generator jets, the cross flow on the end wall is also suppressed.

scholarly journals Wake Mixing Improvements in a Linear Compressor Cascade With Crenulated Trailing Edges

1990 ◽  
Author(s):  
J. L. Veesart ◽  
P. I. King ◽  
W. C. Elrod ◽  
A. J. Wennerstrom
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Trailing Edge ◽  
Turbine Engine ◽  
Pressure Surface ◽  
Trailing Edges ◽  
Linear Compressor Cascade

Trailing edge crenulations offer one possible way of energizing the trailing edge wakes from a gas turbine engine compressor blade by creating small vortices as a result of the pressure differential between the suction surface and pressure surface. The effect of crenulated trailing edges on wake dissipation and mixed-out total pressure loss in a linear, subsonic, compressor cascade was investigated for three low aspect ratio blade configurations: one with no crenulations and two others with large and small crenulation patterns, respectively. The effect of crenulations was to improve the wake mixing and reduce the velocity deficit. larger crenulations dissipated the wake most rapidly, and both crenulation configurations offered an improvement in total pressure loss and some improvement in flow turning.

Experimental and Numerical Investigations of Pressure Loss and 3-D Flow Separations in a Linear Compressor Cascade

2019 ◽  
Author(s):  
Saeed A. El-Shahat ◽  
Hesham M. El-Batsh ◽  
Ali M. A. Attia ◽  
Guojun Li ◽  
Lei Fu
Keyword(s):  
Flow Field ◽  
Flow Separation ◽  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Axial Compressor ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Loss Mechanism ◽  
Linear Compressor Cascade

Abstract Flow separation is a major parameter affecting the compressor performance. It reduces the compressor efficiency, limits static pressure rise capability and contributes to instability in compressors. In applied research, there is a lack of understanding of the nature and mechanism of the three-dimensional (3-D) flow separation in the axial compressor especially on the juncture of the endwall and blade corner region. In the present study, the 3-D flow field in an axial compressor cascade has been studied experimentally as well as numerically. For the experimental study part, a linear compressor cascade has been installed in an open loop wind tunnel. The experimental data was acquired for a Reynolds number Rec = 2.98 × 105 based on the blade chord and the inlet flow conditions. The total pressure loss progress through the blade passage has been measured by using calibrated five and seven-hole pressure probes connected to ATX sensor module data acquisition system. The static pressure distribution on the endwall has been measured employing static pressure taps connected to digital micromanometers. To investigate the loss mechanism through the cascade, the total pressure loss coefficient has been calculated from the measured data. The computational fluid dynamics (CFD) study of the flow field was performed to gain a better understanding of the flow features. Two turbulence models, Spalart-Allmaras (S-A) and shear stress transport SST (k-ω) were used. From both parts of study, the flow field development and total pressure loss progress through the cascade have been investigated and compared. Moreover, the received data demonstrated a good agreement between the experimental and computational results. The predicted flow streamlines by numerical calculations showed regions characterized by flow separation and recirculation zones that could be used to enhance the understanding of the loss mechanism in compressor cascades. All measurements taken by 5-hole and 7-hole pressure probes have been analyzed and compared. It was found that their readings were almost the same and there are no excellences for using 7-hole probe. Furthermore S-A turbulence model calculations showed more consistencies with experimental results than SST (k-ω) model.

Optimization study on the influence of little blades’ spatial position on a compressor cascade performance

2021 ◽  
pp. 095441002110443
Author(s):  
Shan Ma ◽  
Xiaolin Sun
Keyword(s):  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Spatial Position ◽  
Total Pressure Loss ◽  
Field Analysis ◽  
Stable Operation ◽  
Compressor Cascade ◽  
Optimal Position ◽  
End Wall

To reveal the importance of little blades’ spatial position to improve the cascade performance at different condition, the pitchwise and axial direction of the little blades on the end-wall are adopted as the optimization variables to complete a double-objective optimization. Meanwhile, the three-dimensional flow field characteristics of the cascade with and without little blades are analyzed comparatively. The study found that as the optimal solutions are obtained at the three bigger incidences (3°, 5°, and 7°), the optimal position is always close to the leading edge of blade and far away from the blade suction surface, and the more intuitive design suggestions are given in this article. Moreover, at the near design conditions (−1°, 0°, and 1°), little blades increase the total pressure loss and reduce the static pressure, which are considered unsuitable for improving the cascade performance. If the stable operation range are the main performance indicators, the optimization of the little blades’ spatial position should be completed at the near stall condition (7° incidence). If the conditions with mid-range incidences (2°< i <5°) are the main performance index, the parameter optimization of little blades should be achieved at 5°. Based on the further flow field analysis of the optimization results obtained at 3°, 5°, and 7° incidences (named Opt_Act3, Opt_Act5, and Opt_Act7), the induced vortices resist the effect of axial reverse pressure gradient and pass through the blade passage, which is the main reason for the total pressure loss reduction. Appropriate spatial position of little blades not only strengthens the capability to prevent the low-energy fluids accumulating in the corner region near the end-wall, but exhibits sufficient advantage to weaken the boundary layer.

Endwall Contouring and Fillet Design for Reducing Losses and Homogenizing the Outflow of a Compressor Cascade

2014 ◽  
Author(s):  
Oliver Reutter ◽  
Stefan Hemmert-Pottmann ◽  
Alexander Hergt ◽  
Eberhard Nicke
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Wall Region ◽  
Optimized Design ◽  
Compressor Cascade ◽  
Outflow Region ◽  
Endwall Contouring ◽  
Operating Points

The following paper deals with the development of an optimized fillet and an endwall contour for reducing the total pressure loss and for homogenizing the outflow of a highly loaded cascade with a low aspect ratio. The NACA-65 K48 cascade profile without a fillet and without endwall contouring is used as a basis. Optimizations are performed using the DLR in-house tool AutoOpti and the RANS-solver TRACE. Three operating points at an inflow Mach number of 0.67 with different inflow angles are used to secure a wide operating range of the optimized design. At first only a fillet is optimized. The optimized fillet is small at the leading edge and rather high, wide and thick towards the trailing edge. It reduces the total pressure loss and homogenizes the outflow up to a blade height of 20 %. Following this a combined optimization of the endwall and the fillet is performed. The optimized contour leads to the development of a vortex, which changes the secondary flow in such a way, that the corner separation is reduced, which in turn significantly reduces the total pressure loss up to 16 % in the design operating point. The contour in the outflow region leads to a significant homogenization of the outflow in the near wall region.

Location Effect of Boundary Layer Suction on Compressor Hub-Corner Separation

2014 ◽  
Author(s):  
Ping-Ping Chen ◽  
Wei-Yang Qiao ◽  
Karsten Liesner ◽  
Robert Meyer
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Base Flow ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Corner Separation ◽  
Boundary Layer Suction

The large secondary flow area in the compressor hub-corner region usually leads to three-dimensional separation in the passage with large amounts of total pressure loss. In this paper numerical simulations of a linear high-speed compressor cascade, consisting of five NACA 65-K48 stator profiles, were performed to analyze the flow mechanism of hub-corner separation for the base flow. Experimental validation is used to verify the numerical results. Active control of the hub-corner separation was investigated by using boundary layer suction. The influence of the selected locations of the endwall suction slot was investigated in an effort to quantify the gains of the compressor cascade performance. The results show that the optimal chordwise location should contain the development section of the three-dimensional corner separation downstream of the 3D corner separation onset. The best pitchwise location should be close enough to the vanes’ suction surface. Therefore the optimal endwall suction location is the MTE slot, the one from 50% to 75% chord at the hub, close to the blade suction surface. By use of the MTE slot with 1% suction flow ratio, the total-pressure loss is substantially decreased by about 15.2% in the CFD calculations and 9.7% in the measurement at the design operating condition.

A new curvature-controlled stacking-line method for optimization design of compressor cascade considering surface smoothness

2019 ◽  
Vol 234 (5) ◽  
pp. 1061-1074
Author(s):  
Xinyi Zhang ◽  
Xiaoqing Qiang ◽  
Jinfang Teng ◽  
Wensheng Yu
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Optimization Design ◽  
Parametric Method ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Surface Smoothness ◽  
Line Method ◽  
Radial Curvature ◽  
Bézier Spline

The paper presents an advanced parametric method of blade stacking lines in terms of sweep and lean based on controlled curvature. To the knowledge of the authors, there is no related approach reported in open literature that uses Bezier spline as the radial curvature distribution to improve the smoothness of the blade surface; most previous studies ignored the discontinuous slopes of curvature of the parametric curves. The parametric method called curvature-controlled stacking-line method (CCSLM) is performed by changing the magnitude of the sweep or lean. A fourth Bezier spline is adopted to define the curvature of spanwise stacking line directly ensuring surface smoothness. Then, the redesign cascades are created by sectional profiles stacked along the radial stacking lines which are obtained by twice integrating the Bezier spline. Then, the advanced method is conducted to optimize a high-subsonic controlled diffusion airfoil at design point, where the blade shape is generated in terms of lean. A single-objective optimization is performed using Kriging model and genetic algorithm to optimize total pressure loss, and the optimized geometry is obtained. The optimization results show that the blade design CCSLM has significant effects on the endwall flow vortex as well as radial loading distribution. The reduction of total pressure loss and secondary flow is also observed, and the aerodynamic performance is well improved compared with the original cascade.

Adaptive Detached Eddy Simulation of End-Wall Flow in a Linear Compressor Cascade

2020 ◽  
Author(s):  
Zifei Yin
Keyword(s):  
Pressure Coefficient ◽  
Detached Eddy Simulation ◽  
Compressor Cascade ◽  
Local Flow ◽  
Linear Compressor ◽  
Eddy Simulation ◽  
Wall Flow ◽  
Linear Compressor Cascade ◽  
End Wall ◽  
Des Model

Abstract Delayed detached eddy simulations and wall-modeled eddy simulations using the adaptive DES model were performed to simulate corner separation in the Ecole Centrale de Lyon linear compressor cascade. The adaptive DES model directly uses length scale to define eddy viscosity, which makes it nature to compute the model constant CDES dynamically. The dynamic procedure adapts viscosity to local flow and grid. Delayed detached eddy simulations, with and without the dynamic procedure, were performed to demonstrate the benefit of adapting viscosity to local flow. Recycling method was adopted to generate inflow unsteady turbulent boundary layer for wall-modeled eddy simulations. The wall-modeled eddy simulation showed improvement over delayed-DES, in terms of static pressure coefficient around the blade and total pressure loss at downstream locations.

Sign in / Sign up

Export Citation Format

Share Document