Physical Mechanisms Investigation of Sharkskin-Inspired Compressor Cascade Based on Large Eddy Simulations

2021 ◽  
pp. 1-26
Author(s):  
Zhihui Li ◽  
Yan Jin ◽  
Juan Du ◽  
Hongwu Zhang ◽  
Chaoqun Nie
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Simulation Method ◽  
Large Eddy Simulations ◽  
Total Pressure Loss ◽  
Control Mechanisms ◽  
Compressor Cascade ◽  
Boundary Flow ◽  
Cascade Flow ◽  
Large Eddy

Abstract In this paper, the effectiveness of sharkskin-inspired riblets in reducing the aerodynamic loss of compressor cascade flow was investigated by using high-fidelity numerical simulation method. Two key normalized parameters were selected to parameterize various riblet designs, and the corresponding relative change in cascade performance was first investigated based on the uRANS simulations with/without transition model. Then, the large eddy simulations were conducted to investigate the cascade flow with the selected riblet design cases. By comparing the flow resistance, transition positions, vortex formations and turbulence fluctuations of the boundary flow, the flow control mechanisms of the riblets were finally studied. Simulation results show that compared with the prototype case, the total pressure loss can be reduced by up to 20.5% in the fully turbulent environment. This is because the spanwise fluctuation of the turbulent vortices is impeded, and the turbulent vortices are lifted above the riblet tip. However, when considering transition from laminar to turbulent boundary flow, the aerodynamic performance of compressor cascade strongly depends on the riblet position relative to the transition on cascade SS. The total pressure loss can only be reduced by up to 8.1%, and even most riblet designs will degrade the cascade performance. The major reason is that the riblets are located upstream of the transition zone, especially at the small incidence angles. Due to the installation of riblets, the contact area between the laminar flow and the wall surface is increased, and the downstream laminar-to-turbulent transition is promoted.

Physical Mechanisms Investigation of Sharkskin-Inspired Compressor Cascade Based on Large Eddy Simulations

2020 ◽  
Author(s):  
Zhihui Li ◽  
Juan Du ◽  
Hongwu Zhang
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Large Eddy Simulations ◽  
Total Pressure Loss ◽  
Shear Stresses ◽  
Compressor Cascade ◽  
Boundary Flow ◽  
Cascade Flow ◽  
Large Eddy

Abstract In order to survive in a complex environment, nature has produced efficient and versatile resource-rich structures. One of the novel drag reduction designs comes from the efficient movement of sharks through microscope riblets aligned along the flow direction. In this paper, the effectiveness of sharkskin-inspired riblets in reducing the aerodynamic loss of compressor cascade flow was investigated by using high-fidelity numerical simulation method. Two key normalized parameters were selected to parameterize various riblet designs, and the corresponding relative change in cascade performance was first investigated based on the uRANS simulations with/without transition model. Then, the large eddy simulations in conjunction with the wall-adapted local eddy viscosity model were conducted to investigate the cascade flow with the selected riblet design cases. By comparing the flow resistance, transition positions, vortex formations and turbulence fluctuations of the boundary flow, the flow control mechanisms of the riblets were finally studied. Simulation results show that compared with the prototype case, the total pressure loss can be reduced by up to 20.5% in the fully turbulent environment. This is because the spanwise fluctuation of the turbulent vortices is impeded, and the turbulent vortices are lifted above the riblet tip. Low-speed streaks inside the riblet valleys generate relatively low shear stresses, while the high-shear stresses occur only at the riblet tips. However, when considering transition from laminar to turbulent boundary flow, the aerodynamic performance of compressor cascade strongly depends on the riblet position relative to the transition on cascade SS. The total pressure loss can only be reduced by up to 8.1%, and even most riblet designs will degrade the cascade performance. The major reason is that the riblets are located upstream of the transition zone, especially at the small incidence angles. Due to the installation of riblets, the contact area between the laminar flow and the wall surface is increased, and the downstream laminar-to-turbulent transition is promoted.

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.

Impact of a Combination of Micro-Vortex Generator and Boundary Layer Suction on Performance in a High-Load Compressor Cascade

2018 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Chuanle Liu
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Control Methods ◽  
Compressor Cascade ◽  
Boundary Layer Suction ◽  
Micro Vortex Generator

The performance of a compressor cascade is considerably influenced by flow control methods. In this paper, the synergistic effects of combination between micro-vortex generators (MVG) and boundary layer suction (BLS) are discussed in a high-load compressor cascade. Seven cases, which are grouped by a kind of micro-vortex generator and boundary layer suction with three locations, are investigated to control secondary flow effects and enhance the aerodynamic performance of the compressor cascade. The MVG is mounted on the end-wall in front of the passage. The rectangle suction slot with three radial positions is installed on the blade suction surface near the trailing edge. The numerical results show that: at the design condition, the total pressure loss is effectively decreased as well as the static pressure coefficient increase when the combined MVG and SBL method (COM) is used, which is superior to MVG in an aerodynamic performance. At the stall condition, the induced vortex coming from MVG could mix the low-energy fluid and mainstream, which result in the reduced separation, and the total pressure loss decreased by 11.54% when the suction flow ratio is 1.5%. The total pressure loss decreases by 14.59% when the COM control methods are applied.

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.

scholarly journals Investigation of Wash Fluid Preheating on the Effectiveness of Online Compressor Washing in Industrial Gas Turbines

2020 ◽  
Author(s):  
Roupa Agbadede ◽  
Biweri Kainga
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Loss Coefficient ◽  
Industrial Gas Turbines ◽  
Total Pressure Loss Coefficient

Abstract This study presents an investigation of wash fluid preheating on the effectiveness of online compressor washing in industrial gas turbines. Crude oil was uniformly applied on the compressor cascade blades surfaces using a roller brush, and carborundum particles were ingested into the tunnel to create accelerated fouled blades. Demineralized water was preheated to 500C using the heat coil provided in the tank. When fouled blades washed with preheated demineralized and the one without preheating were compared, it was observed that there was little or no difference in terms of total pressure loss coefficient and exit flow angle. However, when the fouled and washed cases were compared, there was a significant different in total pressure loss coefficient and exit flow angle.

Flowfield Investigation of a Compressor Cascade at High Incidence—Part 1: Pneumatic Probe Measurements

2009 ◽  
Author(s):  
Pavlos K. Zachos ◽  
Vassilios Pachidis ◽  
Bernard Charnley ◽  
Pericles Pilidis
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Incidence Angle ◽  
Axial Flow ◽  
Quantitative Measure ◽  
Operating Conditions ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Engine Operation ◽  
Pressure Surface

The performance prediction of axial flow compressors and turbines still relies on the stationary testing of blade cascades. Most of the blade testing studies are done for operating conditions close to the design point or in off-design areas not too far from it. However, blade performance remains unexplored at very far off-design conditions, such as windmilling, characterised by operation under extremely low mass flows and rotational speeds which, in turn, imply highly negative incidence angle values. In this paper, the flow field generated by a 3-dimensional linear compressor cascade at a highly negative incidence angle and zero rotational speed is experimentally investigated using a pneumatic miniature cobra probe. The main objective of the study is to derive the total pressure loss through the blades at such a highly negative incidence angle. An overview of the blade geometry as well as of the experimental facility is given whereas the measurement strategy and the data acquisition technique are also presented. An uncertainty study taking into account the most significant factors affecting the quality of the results has been carried out. As shown by the measurements taken at specific positions downstream of the blades, the flowfield is dominated by highly separated flows on the pressure surface, which contribute to the increased values of the total pressure loss coefficient which seems to be weakly dependent on the inlet Mach number. The quantitative measure of the pressure losses at the extremely negative incidence angle examined can be considered to be a validation platform for correspondent numerical studies of similar flow conditions. Additionally, the experimental results obtained can be used to extend the applicability of the current pressure loss models, increasing the predictive capability of the through flow numerical approaches towards far off-design areas of component or whole engine operation.

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.

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