3-D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall

2002 ◽  
Vol 124 (3) ◽  
pp. 358-366 ◽  
Author(s):  
Anton Weber ◽  
Heinz-Adolf Schreiber ◽  
Reinhold Fuchs ◽  
Wolfgang Steinert
Keyword(s):  
Transonic Flow ◽  
Numerical Study ◽  
Reverse Flow ◽  
Boundary Layer Separation ◽  
Ring Vortex ◽  
Blade Surface ◽  
Compressor Cascade ◽  
Complex Flow ◽  
Flow Angle ◽  
Blade Passage

An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with a low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of 1.9×106. Detailed flow visualizations on the surfaces and five-hole probe measurements inside the blading and in the wake region showed clearly a three-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3-D shock system at blade passage entrance. The experimental data have been used to validate and improve the 3-D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distributions on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results. An analysis of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow near the walls is seen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the sidewall are in qualitatively good agreement to numerical surface streaklines. A considerable improvement of the numerical results could be achieved by a gradual grid refinement, especially in the corner region and by successive code development.

3D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall

2001 ◽  
Author(s):  
Anton Weber ◽  
Heinz-Adolf Schreiber ◽  
Reinhold Fuchs ◽  
Wolfgang Steinert
Keyword(s):  
Transonic Flow ◽  
Numerical Study ◽  
Reverse Flow ◽  
Boundary Layer Separation ◽  
Ring Vortex ◽  
Blade Surface ◽  
Compressor Cascade ◽  
Complex Flow ◽  
Flow Angle ◽  
Blade Passage

An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with a low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of 1.9×106. Detailed flow visualizations on the surfaces and 5-hole probe measurements inside the blading and in the wake region showed clearly a 3-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3D shock system at blade passage entrance. The experimental data has been used to validate and improve the 3D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distributions on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results. An analysis of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow near the walls is seen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the sidewall are in qualitatively good agreement to numerical surface streaklines. A considerable improvement of the numerical results could be achieved by a gradual grid refinement especially in the corner region and by successive code development.

Numerical Analysis of Turbomachinery Flows With Transitional Boundary Layers

2002 ◽  
Author(s):  
P. De Palma
Keyword(s):  
Boundary Layers ◽  
Transonic Flow ◽  
Stokes Equations ◽  
Numerical Study ◽  
Three Dimensional ◽  
Boundary Layer Separation ◽  
Transition Model ◽  
Compressor Cascade ◽  
Complex Flow ◽  
Flow Through

This paper provides a numerical study of the flow through two turbomachinery cascades with transitional boundary layers. The aim of the present work is to validate some state-of-the-art turbulence and transition models in complex flow configurations. Therefore, the compressible Reynolds-averaged Navier–Stokes equations, with an Explicit Algebraic Stress Model (EASM) and k − ω turbulence closure, are considered. Such a turbulence model is combined with the transition model of Mayle for separated flow. The space discretization is based on a finite volume method with Roe’s approximate Riemann solver and formally second-order-accurate MUSCL extrapolation with minmod limiter. Time integration is performed employing an explicit Runge–Kutta scheme with multigrid acceleration. Firstly, the computations of the two- and three-dimensional subsonic flow through the T106 low-pressure turbine cascade are briefly discussed. Then, a more severe test case, involving shock-induced boundary-layer separation and corner stall is considered, namely, the three-dimensional transonic flow through a linear compressor cascade. In the present paper, calculations of such a transonic flow are presented, employing the standard k − ω model and the EASM, without transition model, and a comparison with the experimental data available in the literature is provided.

Non-Axisymmetric Endwall Contouring in a Linear Compressor Cascade

2020 ◽  
Author(s):  
Yuchen Ma ◽  
Jinfang Teng ◽  
Mingmin Zhu ◽  
Xiaoqing Qiang
Keyword(s):  
Pressure Loss ◽  
Reverse Flow ◽  
Measurement Techniques ◽  
Suction Side ◽  
Limiting Factor ◽  
Blade Surface ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Endwall Contouring ◽  
Linear Compressor Cascade

Abstract Modern axial compressors are designed to be highly loaded in terms of aerodynamics, which can lead to challenges of increasing the compressor efficiency. Losses associated with secondary flow effects are well known to be the major limiting factor of improving the compressor performance. In this study, non-axisymmetric endwall contouring in a linear compressor cascade was generated through the optimization process. Combined with numerical simulation, wind tunnel tests on linear cascades with flat and contoured endwall were performed with various measurement techniques at the design and off-design conditions. The simulation results show that optimal endwall design can provide 3.08% reduction of the total pressure loss at the design condition. The reduction of pressure loss obtained is mainly below 24%span with the size of the high loss region being effectively reduced. At off-design condition, the numerical benefit of the endwall contouring is found less pronounced. The discrepancy is spotted between simulation and experiments. The experimental pressure loss reduction is mainly below 18% at ADP. And the pressure loss for the CEW increases greatly at offdesign condition in experiments. Flow patterns revealed by numerical simulations show that the separation on the blade surface is mitigated with focus point disappearing, and reverse flow on the endwall near the suction side corner is moved away from the blade surface. CFD analysis indicates that the altered pressure distribution on the endwall accelerates the flow at the suction side corner and moves the reverse flow core further downstream. The weakened interaction between the corner vortex and tornado-like vortex from the endwall near the suction side corner is the main control mechanism of the CEW. The performance improvement in the linear compressor is mainly gained from it.

Rotating Mechanism of Diffuser Stall in a Centrifugal Compressor With Vaneless Diffuser

2021 ◽  
Author(s):  
Nobumichi Fujisawa ◽  
Yuki Agari ◽  
Yoshifumi Yamao ◽  
Yutaka Ohta
Keyword(s):  
Boundary Layer ◽  
Mass Flow ◽  
Centrifugal Compressor ◽  
Reverse Flow ◽  
Boundary Layer Separation ◽  
Vaneless Diffuser ◽  
Low Velocity ◽  
Flow Angle ◽  
Discharge Flow ◽  
Layer Separation

Abstract The rotating mechanism of diffuser stall in a centrifugal compressor with a vaneless diffuser is investigated via experimental and computational analyses. Diffuser stall is generated as the mass flow rate decreases, and it rotates at 25%–30% of the impeller rotational speed. First, a diffuser stall cell emerges at 180° from the cutoff by the hub-side boundary layer separation. Subsequently, the diffuser stall cell develops further owing to boundary layer separation accumulation and an induced low-velocity area. The low-velocity region forms a blockage across the diffuser passage span. The diffuser stall cell expands owing to the boundary layer separations that occurred on the shroud and hub wall by turns. Finally, the diffuser stall cell vanishes when it passes the cutoff because mass flow recovery occurred. Furthermore, the static pressure ahead of the rotating stall decreases because of the merging of the impeller discharge flow and the reverse flow from the casing. Accordingly, a reverse flow occurred owing to the evolution of the separation vortex at the diffuser exit. In addition, the flow angle decreases by the merging of the impeller discharge flow and reverse flow from the casing. Therefore, boundary layer separations start occurring on the shroud and hub wall ahead of the stall cell. The rotating mechanism of the diffuser stall is induced by the reverse flow development and a decrease in the flow angle ahead of the stall cell.

Numerical Study of the Effect of Secondary Vortex on Three-Dimensional Corner Separation in a Compressor Cascade

2016 ◽  
Vol 33 (1) ◽  
Author(s):  
Yangwei Liu ◽  
Hao Yan ◽  
Lipeng Lu
Keyword(s):  
Numerical Study ◽  
Three Dimensional ◽  
Horseshoe Vortex ◽  
Detached Eddy Simulation ◽  
Compressor Cascade ◽  
Complex Flow ◽  
Suction Surface ◽  
Corner Separation ◽  
Delayed Detached Eddy Simulation ◽  
Passage Vortex

AbstractThe complex flow structures in a linear compressor cascade have been investigated under different incidences using both the Reynolds-averaged Navier–Stokes (RANS) and delayed detached eddy simulation (DDES) methods. The current study analyzes the development of horseshoe vortex and passage vortex in a compressor cascade based on DDES results and explores the effect of the passage vortex on corner separation using the RANS method. Results show that the effect of horseshoe vortex on three-dimensional corner separation is weak, whereas the effect of passage vortex is dominant. A large vortex breaks into many small vortices in the corner separation region, thereby resulting in strong turbulence fluctuation. The passage vortex transports the low-energetic flow near the endwall to the blade suction surface and enlarges corner separation in the cascade. Hence, total pressure loss increases in the cascade.

Comparison of Theoretical and Experimental Data for an Oscillating Transonic Compressor Cascade

1999 ◽  
Author(s):  
Volker Carstens ◽  
Stefan Schmitt
Keyword(s):  
Experimental Data ◽  
Transonic Flow ◽  
Operating Conditions ◽  
Experimental Investigations ◽  
Test Case ◽  
Flow Conditions ◽  
Compressor Cascade ◽  
Blade Passage ◽  
Aerodynamic Damping ◽  
Transonic Compressor

Numerical and experimental results are compared for a compressor cascade performing harmonic oscillations in transonic flow. The flow field was calculated by a Q3D Navier Stokes code, the basic features of which are the use of an upwind flux difference scheme for the convective terms, the implementation of an effective one-equation turbulence model and the use of deforming multi-block grids. The experimental investigations were performed in an annular cascade windtunnel where unsteady blade pressures were measured for two different operating conditions of the cascade. The present data were all obtained for tuned torsional modes where the blades performed pitching oscillations with the same frequency and amplitude, but with a constant interblade phase angle. In the first test case the steady flow around the blades was purely subsonic. For the second test case the compressor cascade was run under transonic flow conditions where a normal shock in the front part of the blades’ suction side is followed by a blade passage shock. It becomes apparent that under subsonic flow conditions the predicted aerodynamic damping coefficients are in resonable agreement with the experimental data, although the numerical pressure amplitudes are much higher than the measured ones. In transonic flow significant discrepancies between computed and experimentally determined pressure amplitudes are observed, whereas the accuracy of the pressure phase prediction is comparable to the subsonic test case. Another important result of these investigations is that oscillations of the blade passage shock lead to strong variations of the local aerodynamic damping of the blades, but do not significantly change the global damping coefficient of the tested compressor cascade.

Experimental and Numerical Investigation of the Flow Field in the Radial Inlet of a Centrifugal Compressor

2012 ◽  
Author(s):  
Fenghui Han ◽  
Datong Qi ◽  
Jiajian Tan ◽  
Li Wang ◽  
Yijun Mao
Keyword(s):  
Flow Field ◽  
Centrifugal Compressor ◽  
Numerical Study ◽  
Internal Flow ◽  
Structure Design ◽  
Outlet Section ◽  
Complex Flow ◽  
Flow Angle ◽  
Flow Phenomena ◽  
Probe Calibration

This paper presents an experimental and numerical study of the flow field in a typical geometry of a centrifugal compressor radial inlet. A five-hole probe system, which makes the probe calibration and the data acquisition automatically controlled by computer, was developed and used to measure the pressure, velocity and flow angle distributions inside the radial inlet. The testing portion consists of the entrance and exit of the radial inlet, the outlet section of the suction nozzle and the exit of the plenum, including 46 sampling holes and 923 measuring points. In parallel with the experiment, a computational analysis was also carried out to simulate the internal flow of the radial inlet with a commercial CFD Code. The numerical results are compared with the experimental data. It shows a good agreement between CFD and the measurement on most sections. Based on the experiment and simulation, this study reveals the detailed flow conditions in a radial inlet, which helps to figure out how the complex flow pattern in a radial inlet forms and develops as well as the influences on the downstream components. It yields an improved understanding of the principle of flow phenomena in radial inlets, and gives recommendations for optimizing the structure design of the radial inlet of centrifugal compressors.

Performance and Boundary Layer Development of a High Turning Compressor Cascade at Sub- and Supercritical Flow Conditions

2012 ◽  
Author(s):  
Christoph Bode ◽  
Dragan Kožulović ◽  
Udo Stark ◽  
Heinz Hoheisel
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Density Ratio ◽  
Boundary Layer Separation ◽  
Experimental Results ◽  
Numerical Code ◽  
Compressor Cascade ◽  
Boundary Layer Development ◽  
Boundary Layer Interaction ◽  
Inlet Angle

Based on current numerical investigations, the present paper reports on new Q2D midspan-calculations and results for the well known high turning (Δβ = 50°) supercritical (Ma1 = 0.85) compressor cascade V2. A Q2D treatment of the problem was chosen in order to avoid the difficult modelling of the porous endwalls in a corresponding 3D approach. All simulations were done with the RANS solver TRACE of the DLR Cologne in combination with modified versions of the Wilcox turbulence model and Langtry/Menter transition model. Existing experimental Q2D midspan-results for the V2 compressor cascade were used to demonstrate the improved ability of the numerical code to determine performance characteristics, blade pressure and Mach number distributions as well as boundary layer parameter and velocity distributions. The loss characteristics show minimum loss regions when plotted against inlet angle or axial velocity density ratio. Within these regions, increasing with decreasing Mach number, the experimental results were adequately predicted. Outside these regions it turned out difficult to reproduce the experimental results due to increasing boundary layer separation. Furthermore, the prediction quality was very good for subsonic conditions (Ma1 = 0.60) and still reasonable for supercritical conditions (Ma1 = 0.85), where shock/boundary layer interaction made the prediction more difficult.

scholarly journals Numerical Study of Centrifugal Fan with Slots in Blade Surface

2015 ◽  
Vol 126 ◽  
pp. 588-591 ◽  
Author(s):  
Rui Rong ◽  
Ke Cui ◽  
Zijun Li ◽  
Zhengren Wu
Keyword(s):  
Numerical Study ◽  
Centrifugal Fan ◽  
Blade Surface
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