The Influence of Multimode Transition Initiated by Periodic Wakes on the Profile Loss of a Linear Compressor Cascade

2000 ◽  
Author(s):  
Reinhold Teusch ◽  
Stefan Brunner ◽  
Leonhard Fottner ◽  
Marius Swoboda
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Laminar Separation ◽  
Blade Row ◽  
Linear Compressor Cascade ◽  
Profile Loss ◽  
Undisturbed Flow

This paper presents results of boundary layer and loss measurements in a high speed cascade wind tunnel on a linear compressor cascade under the influence of unsteady, periodic wakes. The wakes of an upstream blade row were simulated by cylindrical bars moved by a belt mechanism upstream of the cascade. Extensive hot-film array, hot-wire and pressure measurements with variation of steady and unsteady inlet flow conditions have been performed for a better understanding of the transition and loss mechanisms on a blade row interacting with wakes. The incoming wakes are inducing early forced transition in the boundary layer followed in time by calmed regions. Due to its higher shear stress level and its fuller velocity profile, the calmed flow is able to suppress laminar separation bubbles and to delay transition in the region with undisturbed flow between wakes, playing a significant role in the loss generation process. At the investigated low Reynolds number, where the measurements for the steady flow case showed a well-developed laminar separation bubble, reductions of profile loss up to 20% were observed for the measured configuration. In the case of the high Reynolds number, where in undisturbed flow only a small separation bubble was detected, a profile loss rise up to 30% was measured. Beside a better understanding of unsteady flow physics the goal of these basic investigations of unsteady transition is to create a wide database for the improvement of transition modeling in unsteady CFD codes.

A Study of Loss Mechanism in a Linear Compressor Cascade at the Corner Stall Condition

2017 ◽  
Author(s):  
Zhiyuan Li ◽  
Juan Du ◽  
Aleksandar Jemcov ◽  
Xavier Ottavy ◽  
Feng Lin
Keyword(s):  
Boundary Layer ◽  
Secondary Flow ◽  
Main Flow ◽  
Wake Flow ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Corner Separation ◽  
Flow Loss ◽  
Linear Compressor Cascade ◽  
Profile Loss

The loss-generating mechanism of a linear compressor cascade at the corner stall condition was numerically studied in this paper. The hybrid RANS/LES method was used to perform the high-fidelity simulations. By comparing the results captured by SSTDES, DDES, SAS models with the experimental data, the SSTDES model is proven to be more accurate in capturing the detailed flow structure of the corner stall than the other two models. Taking the turbulence dissipation term of SSTDES model into account, the volumetric entropy generation rate and a new dimensionless local loss coefficient are proposed and used to analyze the loss-generating mechanism in this work. It was found that the main flow loss generated in this cascade could be sorted as the wake flow loss, the profile loss, the secondary flow loss and the endwall loss according to their amounts. The corner separation significantly affects the secondary flow loss, wake flow loss and profile loss in the cascade passage. The mixing between the separated boundary layer flow and the main flow, the shear between a tornado vortex and the main flow are the main sources of the secondary flow loss. The wake flow loss is the largest loss source of the cascade, accounting for 41.8% of the total loss. There are two peaks of the wake flow loss along the spanwise direction near the corner stall region. This phenomenon is related to the appearance of large velocity gradient flows when the main flows and the corner separation flows mix together. The profile loss takes up 40.06 % of the total loss. The profile loss intensity in the corner region is lower than the mid blade span due to the interaction of the boundary layer on the suction side with the corner separation.

Effect of Leading Edge Roughness and Reynolds Number on Compressor Profile Loss

2013 ◽  
Author(s):  
Ju Hyun Im ◽  
Ju Hyun Shin ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Leading Edge ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Laser Doppler Velocimeter ◽  
Compressor Cascade ◽  
Edge Roughness ◽  
Profile Loss

An experimental investigation has been conducted to characterize the influence of leading edge roughness and Reynolds number on compressor cascade profile loss. Tests have been conducted in a low-speed linear compressor cascade at Reynolds numbers between 210,000 and 640,000. Blade loading and loss have been measured with pressure taps and pneumatic probes. In addition, a two-component laser-doppler velocimeter (LDV) has been used to measure the boundary layer velocity profiles and turbulence levels at various chordwise locations near the blade suction surface. The “smooth” blade has a centerline-averaged roughness (Ra) of 0.62 μm. The “rough” blade is roughened by covering the leading edge of the “smooth” blade, including 2% of the pressure side and 2% of the suction side, with a 100 μm-thick tape with a roughness Ra of 4.97 μm. At Reynolds numbers ranging from 210,000 to 380,000, the leading edge roughness decreases loss slightly. At Reynolds number of 210,000, the leading edge roughness reduces the size of the suction side laminar separation bubble and turbulence level in the turbulent boundary layer after reattachment. Thus, the leading edge roughness reduces displacement and momentum thicknesses as well as profile loss at Reynolds number of 210,000. However, the same leading edge roughness increases loss significantly for Re = 450,000 ∼ 640,000. At Reynolds number of 640,000, the leading edge roughness decreases the magnitude of the favorable pressure gradient for axial chordwise locations less than 0.41 and induces turbulent separation for axial chordwise locations greater than 0.63, drastically increasing loss. Thus, roughness limited to the leading edge still has a profound effect on the compressor flow field.

Experimental and Numerical Investigation of Transition Effects on a Low Reynolds Number Airfoil

2017 ◽  
Author(s):  
Michael J. Collison ◽  
Peter X. L. Harley ◽  
Domenico di Cugno
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Low Reynolds Number ◽  
Flow Visualisation ◽  
Transition Model ◽  
Laminar Separation ◽  
Oil Flow ◽  
Low Reynolds ◽  
Low Reynolds Number Airfoil

Low speed, small scale turbomachinery operates at low Reynolds number with transition phenomena occurring. In small consumer product applications, high efficiency and low noise are key performance metrics. Transition behaviour will partly determine the state of the boundary layer at the trailing edge; whether it is laminar, turbulent or separated impacts aerodynamic and acoustic performance. This study aimed to evaluate a commercially available CFD transition model on a low Reynolds number Eppler E387 airfoil and identify whether it was able to correctly model the boundary layer transition, and at what expense. CFD was carried out utilising the ANSYS Shear Stress Transport (SST) k-ω γ-Reθ transition model. The CFD progressed from 2D in Fluent v150, through to single cell thickness 3D (pseudo 2D) in CFX v172. An Eppler E387 low Reynolds number airfoil, for which experimental data was readily available from literature at Re = 200,000 was used as the validation case for the CFD, with results computed at numerous incidence angles and mesh densities. Additionally, experimental surface oil flow visualisation was undertaken in a wind tunnel using a scaled E387 airfoil for the zero incidence case at Re = 50,000. The flow visualisation exhibited the expected key features of transition in the breakdown of the boundary layer from laminar to turbulent, and was used as a validation case for the CFD transition model. The comparison between the results from the CFD transition model and the experimental data from literature suggested varying levels of agreement based on the mesh density and CFD solver in the starting location of the laminar separation bubble, with higher disparity for the position of the reattachment point. Whether 2D or 3D, the prediction accuracy was seen to worsen at high incidence angles. Finally, the location of the laminar separation bubble between CFD and oil flow visualisation had good agreement and a set of guidelines on the mesh parameters which can be applied to low Reynolds number turbomachinery simulations was determined.

Large Eddy Simulation of Transitional Flow in a Compressor Cascade

2017 ◽  
Author(s):  
Syed Anjum Haider Rizvi ◽  
Joseph Mathew
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Separation Bubble ◽  
Transitional Flow ◽  
Transitional Region ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Pass Filter ◽  
Low Pass Filter ◽  
Large Eddy

At off-design conditions, when the blade Reynolds number is low, a significant part of the blade boundary layer can be transitional. Then, standard RANS models are unable to predict the flows correctly but explicit transition modeling provides some improvement. Since large eddy simulations (LES) are improvements on RANS, the performance of LES was examined by simulating a flow through a linear, compressor cascade for which experimental data are available — specifically at the Reynolds number of 210,000 based on blade chord when transition processes occur over a significant extent of the suction surface. The LES were performed with an explicit filtering approach, applying a low-pass filter to achieve sub-grid-scale modeling. Explicit 8th-order difference formulas were used to obtain high resolution spatial derivative terms. An O-grid was wrapped around the blade with suitable clustering for the boundary layer and regions of large changes along the blade. Turbulent in-flow was provided from a precursor simulation of homogeneous, isotropic turbulence. Two LES and a DNS were performed. The second LES refines the grid in the vicinity of the separation bubble on the suction surface, and along the span. Surface pressure distributions from all simulations agree closely with experiment, thus providing a much better prediction than even transition-sensitive RANS computations. Wall normal profiles of axial velocity and fluctuations also agree closely with experiment. Differences between LES and DNS are small, but the refined grid LES is closer to the DNS almost everywhere. This monotonic convergence, expected of the LES method used, demonstrates its reliability. The pressure surface undergoes transition almost immediately downstream of the leading edge. On the suction surface there are streaks as expected for freestream-turbulence-induced transition, but spots do not appear. Instead, a separating shear layer rolls up and breaks down to turbulence at re-attachment. Both LES capture this process. Skin friction distribution reveals the transition near the re-attachment to occur over an extended region, and subsequent relaxation is slower in the LES. The narrower transition zone in the DNS is indicative of the essential role of smaller scales during transition that should not be neglected in LES. Simulation data also reveal that an assumption of laminar kinetic energy transition models that Reynolds shear stress remains small in the pre-transitional region is supported. The remaining differences in the predictions of such models is thus likely to be the separation-induced transition which preempts the spot formation.

Shear Layer Development, Separation, and Stability Over a Low-Reynolds Number Airfoil

2018 ◽  
Vol 140 (7) ◽  
Author(s):  
Paul Ziadé ◽  
Mark A. Feero ◽  
Philippe Lavoie ◽  
Pierre E. Sullivan
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Shear Layer ◽  
Separation Bubble ◽  
Low Reynolds Number ◽  
Base Flow ◽  
Mean Velocity ◽  
Laminar Separation ◽  
Low Reynolds ◽  
Layer Development

The shear layer development for a NACA 0025 airfoil at a low Reynolds number was investigated experimentally and numerically using large eddy simulation (LES). Two angles of attack (AOAs) were considered: 5 deg and 12 deg. Experiments and numerics confirm that two flow regimes are present. The first regime, present for an angle-of-attack of 5 deg, exhibits boundary layer reattachment with formation of a laminar separation bubble. The second regime consists of boundary layer separation without reattachment. Linear stability analysis (LSA) of mean velocity profiles is shown to provide adequate agreement between measured and computed growth rates. The stability equations exhibit significant sensitivity to variations in the base flow. This highlights that caution must be applied when experimental or computational uncertainties are present, particularly when performing comparisons. LSA suggests that the first regime is characterized by high frequency instabilities with low spatial growth, whereas the second regime experiences low frequency instabilities with more rapid growth. Spectral analysis confirms the dominance of a central frequency in the laminar separation region of the shear layer, and the importance of nonlinear interactions with harmonics in the transition process.

scholarly journals Flow control in linear compressor cascades by inclusion of suction side dimples at varying locations

2018 ◽  
Vol 232 (6) ◽  
pp. 706-721 ◽  
Author(s):  
Hua-wei Lu ◽  
Yi Yang ◽  
Shang Guo ◽  
Yu-xuan Huang ◽  
Hong Wang ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Separation Bubble ◽  
Pressure Rise ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
Chord Length ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Linear Compressor ◽  
Comparison Results

The flow characteristics and loss behavior over an array of parallel recessed dimples on a high turning linear compressor cascade have been investigated using the Reynolds-averaged Navier–Stokes approach. Steady simulations have been carried out at three dimple locations of 10–32%, 38–60%, 60–82% chord length of suction surface with the inlet Mach number of 0.7. Flow conditions were compared in exit loss coefficient, static pressure rise, streamline patterns, vortex structures, boundary layer parameters, and blade surface pressure between the smooth and the modified cascades. The results indicate that the dimples prior to the separation line report an overall enhancement in the aerodynamic performance in comparison to that of a smooth blade. Symmetric spanwise vortex, which energizes the boundary layer, can roll up inside the dimples. Therefore, the boundary layer with the higher momentum can bear the adverse pressure gradient, which will suppress the flow separation and associated losses. Three dimpled configurations can all eliminate the separation bubble on the suction side, but the dimples located at 60–82% chord length take the negative effect on the aerodynamic performance due to the more chaos condition in the corner separation region. The comparison results also indicate that the optimum location of dimples may exist in front of the separation bubble. Loss reduction of 18.8% and 10.8% can be achieved under the 10–32% c and 38–60% c dimple configurations, respectively.

Effects of Reynolds number and loading distribution on the aerodynamic performance of a high subsonic compressor airfoil

2020 ◽  
Vol 234 (8) ◽  
pp. 1069-1083
Author(s):  
Ming-Yang Wang ◽  
Zi-Liang Li ◽  
Sheng-Feng Zhao ◽  
Yan-Feng Zhang ◽  
Xin-Gen Lu
Keyword(s):  
Reynolds Number ◽  
Separation Bubble ◽  
Aerodynamic Performance ◽  
Generation Process ◽  
Laminar Separation Bubble ◽  
Turbulence Level ◽  
Displacement Effect ◽  
Laminar Separation ◽  
Performance Deterioration ◽  
Profile Loss

The laminar-turbulent transition process on the compressor blade surface is often induced by the laminar separation flow at low Reynolds number ( Re). In the present study, numerical simulations were conducted to investigate the structure of the laminar separation bubble and its effects on the profile loss of a high subsonic compressor airfoil under different Re conditions, and the mechanism for the performance deterioration of compressor airfoil at low Re was clarified. Besides, the airfoil was redesigned to obtain a series of airfoils with different loading distributions, and the aerodynamic performance of these airfoils was compared and analyzed in detail. According to the simulation results, the laminar separation bubble mainly determined the loss generation process of a compressor airfoil. When Re decreased from 12 × 105 to 1.5 × 105, the laminar separation bubble on the suction surface grew thicker and the length was increased by 11.2% of the axial chord. As such, the reversed flow inside the laminar separation bubble became more obvious and the turbulence level downstream of the maximum thickness of laminar separation bubble was increased. Also, the growth in the turbulent boundary layer was enhanced, causing more serious flow blockage and wake mixing. According to the Denton's profile loss model, the larger trailing edge loss caused by the stronger displacement effect of laminar separation bubble was supposed to be the main reason for the performance deterioration of compressor airfoil under low Re conditions. The ultra-front loading distribution for airfoil has the possibility to suppress or even eliminate the negative effect of laminar separation bubble, and the profile loss was decreased by 26.7% at Re = 1.5 × 105; however, the less significant performance improvement was observed at some higher Re. Moreover, the ultra-front loaded airfoil was less sensitive to the inlet turbulence level and the superiority still holds even at some supercritical conditions.

Direct Numerical Simulation of By-Pass and Separation-Induced Transition in a Linear Compressor Cascade

2006 ◽  
Author(s):  
Tamer Zaki ◽  
Paul Durbin ◽  
Jan Wissink ◽  
Wolfgang Rodi
Keyword(s):  
Numerical Simulation ◽  
Boundary Layer ◽  
Direct Numerical Simulation ◽  
Pressure Gradient ◽  
Surface Boundary ◽  
Mean Flow ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Free Stream Turbulence ◽  
Linear Compressor Cascade

Direct Numerical Simulation (DNS) of flow through a linear compressor cascade with incoming free-stream turbulence was performed. On the pressure side, the boundary layer flow is found to undergo by-pass transition: The incident vortical disturbances trigger the formation of elongated boundary layer perturbation jets (or streaks) with amplitudes on the order of 10% of the mean flow. The inception of turbulent spots, which leads to breakdown, is triggered on the backward perturbation jets (negative u-fluctuations). The turbulent patches spread and finally merge into the downstream, fully turbulent region. The suction surface boundary layer is initially subject to a Favorable Pressure Gradient (FPG), followed by a strong Adverse Pressure Gradient (APG). The FPG suppresses the formation of boundary layer streaks. The result is a stabilized boundary layer that does not undergo transition. Farther downstream, the strong APG causes the laminar boundary layer to separate, which is followed by turbulent reattachment.

Unsteady Boundary Layer Development Due to Wake Passing Effects on a Highly Loaded Linear Compressor Cascade

2004 ◽  
Vol 126 (4) ◽  
pp. 493-500 ◽  
Author(s):  
Lothar Hilgenfeld ◽  
Michael Pfitzner
Keyword(s):  
Boundary Layer ◽  
Measurement Techniques ◽  
Unsteady Boundary Layer ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Boundary Layer Development ◽  
Linear Compressor Cascade ◽  
Wake Passing ◽  
Layer Development ◽  
Highly Loaded

The effects of wake passing on boundary layer development on a highly loaded linear compressor cascade were investigated in detail on the suction side of a compressor blade. The experiments were performed in the High Speed Cascade Wind Tunnel of the Institut fuer Strahlantriebe at Mach and Reynolds numbers representative for real turbomachinery conditions. The experimental data were acquired using different measurement techniques, such as fast-response Kulite sensors, hot-film array and hot-wire measurements. The incoming wakes clearly influence the unsteady boundary layer development. Early forced transition in the boundary layer is followed in time by calmed regions. Large pressure fluctuations detectable in the ensemble averaged Kulite data reveal the existence of coherent structures in the boundary layer. Distinct velocity variations inside the boundary layer are amplified when approaching the blade surface. The time–mean momentum thickness values are reduced compared to the steady ones and therefore clarify the potential for a loss reduction due to wake passing effects.

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