Experimental Investigation of Effects of Leading-Edge Tubercles on Compressor Cascade Performance

2014 ◽  
Author(s):  
M. C. Keerthi ◽  
Abhijit Kushari ◽  
Ashoke De ◽  
Arun Kumar
Keyword(s):  
Incidence Angle ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Stall Margin ◽  
Axial Compressors ◽  
Pressure Loss Coefficient ◽  
Lift And Drag

In the present study, the effectiveness of passive structures called tubercles on an axial compressor blade row is studied experimentally. Tubercles are the modifications to the leading edge of an airfoil in the form of blunt wave-like serrations. Although several studies on the effect of tubercles on isolated blades are available in literature, detailed study of their effect on a cascade of blades, such as in the case of an axial flow turbo-machine is lacking. Such an application in an axial compressor will result in a significant increase in the stall margin. Presently, experiments have been performed on a linear compressor cascade with a blade height of 0.15 m and mean chord of 0.06 m, on a NACA 65209 airfoil profile. The plain and modified blades are fabricated using rapid prototyping to ensure conformity to the required geometry. The cascade is designed in such a way that the incidence (angle of attack) and the stagger can be changed easily. The measurements are taken at the exit plane using a five-hole Pitot probe to obtain three-components of velocity and static pressure data over fine measurement grids. The effect is determined in terms of lift and drag coefficients, lift-to-drag ratio and total pressure loss coefficient. Experiments have been carried out for different pitch and amplitude (serration depth) of tubercles to understand their effect. The stall incidence angle for the best performing blade is found to increase up to 8.6° from that of the unmodified blade of 6.0°. Application of such structures in axial compressor blades may well be adequate to prevent stalling in axial compressors over a wide operating range.

Effect of Various Trench Designs on Axial Compressor Blade Tip Aerodynamics

2019 ◽  
Author(s):  
Ashwin Ashok ◽  
Patur Ananth Vijay Sidhartha ◽  
Shine Sivadasan
Keyword(s):  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Horseshoe Vortex ◽  
Tip Clearance ◽  
Compressor Cascade ◽  
Tip Leakage Vortex ◽  
Stall Margin ◽  
Casing Treatment ◽  
Tip Leakage

Abstract Tip clearance of axial compressor blades allows leakage of the flow, generates significant losses and reduces the compressor efficiency. The present paper aims to discuss the axial compressor tip aerodynamics for various configurations of tip gap with trench. The various configurations are obtained by varying the clearance, trench depth, step geometry and casing contouring. In this paper the axial compressor aerodynamics for various configurations of tip gap with trench have been studied. The leakage flow structure, vorticity features and entropy generations are analyzed using RANS based CFD. The linear compressor cascade comprises of NACA 651810 blade with clearance height varied from 0.5% to 2% blade span. Trail of the tip leakage vortex and the horseshoe vortex on the blade suction side are clearly seen for the geometries with and without casing treatments near the stalling point. Since the trench side walls are similar to forward/backing steps, a step vortex is observed near the leading edge as well as trailing edge of the blade and is not seen for the geometry without the casing treatment. Even though the size of the tip leakage vortex seams to be reduces by providing a trench to the casing wall over the blade, the presence of additional vortices like the step vortex leads to comparatively higher flow losses. An increase in overall total pressure loss due to the application of casing treatment is observed. However an increase in stall margin for the geometries with casing is noted.

Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl

1987 ◽  
Vol 109 (1) ◽  
pp. 142-148 ◽  
Author(s):  
C. H. Law ◽  
A. J. Wennerstrom
Keyword(s):  
Static Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Compressors ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Pressure Distributions ◽  
Blade Section ◽  
Flow Compressor

A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performance of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.

Effect of undulating blades on highly loaded compressor cascade performance

2020 ◽  
pp. 095765092090782
Author(s):  
Lirong Su ◽  
Xiaoqing Qiang ◽  
Tan Zheng ◽  
Jinfang Teng
Keyword(s):  
Incidence Angle ◽  
Leading Edge ◽  
Control Technique ◽  
Maximum Magnitude ◽  
Streamwise Vortices ◽  
Compressor Cascade ◽  
Pressure Loss Coefficient ◽  
Positive Effects ◽  
The Impact ◽  
Highly Loaded

Humpback whale’s flipper with leading-edge tubercles has been attracting aerodynamic and hydrodynamic researchers’ attentions by its stall-delayed characteristics. Inspired by this, the undulating configuration is used in a highly loaded compressor cascade as a new type of passive flow control technique. A new model of undulating compressor blade is studied in this paper. To investigate the effect of the undulating configuration on cascade performance without the impact of endwall, steady Reynolds-averaged Navier–Stokes simulations of infinite-span cascades are carried out with and without undulations at an inlet Mach number of 0.5. A parametric study is performed to conclude that, with a suitable wavelength, the undulating blade could achieve a rise in diffusion capacity, accompanied by 12.9% reduction in total pressure loss coefficient at a post-stall incidence angle of 8°, whereas it produces negligible impact in cascade performance at 0° incidence angle. Flow visualization further reveals that wavelength is a crucial parameter, determining the spanwise space for the formation of streamwise vortices. Undulating blades could produce positive effects with maximum magnitude when the counter-rotating streamwise vortices take dominant position along span with an appropriate size.

Impact of Vortex Produced by a Novel Curve-Micro Vortex Generator on Secondary Flow in Compressor Cascade

2017 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Lanpan Li ◽  
Jinhua Lang
Keyword(s):  
Secondary Flow ◽  
Incidence Angle ◽  
Axial Compressor ◽  
Vortex Generator ◽  
Leading Edge ◽  
Suction Side ◽  
Chord Length ◽  
Research Experience ◽  
Compressor Cascade ◽  
Micro Vortex Generator

The performance of compressor cascade is considerably influenced by secondary flow. An extensive experimental study of vortex generator (VG) applied on axial compressor was conducted by many scholars, in order to control these effects. Particularly, MVG is one of the hot researches in present to restrain secondary flow. On the foundation of research experience finished by the former scholars, a new Curve-micro vortex generator (C-MVG) was proposed in this paper. In order to investigate the effect of C-MVG on secondary flow in low-Mach number cascade, the present was carried out on a high-loaded axial compressor cascade with incoming flow of Ma<0.3. The experiment of baseline was conducted at a low speed (incompressible) cascade wind tunnel. The C-MVGs were placed on the end-wall at a distance of 7% chord length ahead of passage and a pitch distance of 26 mm from the leading edge of suction side. 8 cases with different spacing and θVGs were calculated. The height of all the C-MVGs were 5 mm and each case was comprised of 3 vanes. At design and stall incidence angle (−1 deg and 8 deg), the total pressure loss coefficient averaged by mass-flow (Loss) in the outlet was analyzed with numerical method of k-omega turbulence model. Different combinations of C-MVGs were compared. Results show that the Loss in 140% axial chord length (Ca) after leading edge was increased on design condition. At 8 deg incidence angle, all cases could delay the inception of separation and decrease loss. The case VGθ3 showed the highest loss reduction benefit of 7.3%, which indicated that C-MVGs could control the large separation area effectively.

Influence of key geometric parameters on the aerodynamic performance of a supersonic compressor cascade

2017 ◽  
Vol 232 (12) ◽  
pp. 2338-2348 ◽  
Author(s):  
Cui Cui ◽  
Zhenggui Zhou ◽  
Jinhuan Zhang ◽  
Sheng Tao
Keyword(s):  
Mach Number ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Simulation Method ◽  
Wave Pattern ◽  
Aerodynamic Performance ◽  
Compressor Cascade ◽  
Stall Margin ◽  
Edge Radius

The shocks in a supersonic/transonic axial compressor can increase the pressure ratio and cause flow losses. Therefore, it is essential to organize the shock wave pattern in the flow passage to reduce these losses. This study uses a numerical simulation method to study the influences of the leading-edge radius, cascade solidity, and pre-compression on the aerodynamic performance of a supersonic cascade. The cascade is designed using the pre-compression method to reduce shock losses; the inlet Mach number is 2.0 and the total pressure ratio is approximately 3.4. The results indicate that the cascade efficiency and stall margin decrease with an increase in the leading-edge radius; however, when the leading-edge radius is less than 0.1 mm, the influences of its change decrease. As cascade solidity increases, the stall margin first increases and then decreases. The larger the degree of pre-compression, the smaller the Mach number in front of the first oblique passage shock and the higher the efficiency; however, a large pre-compression effect can cause the ending normal shock to move upstream, decreasing the stall margin.

scholarly journals Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl

1986 ◽  
Author(s):  
C. Herbert Law ◽  
Arthur J. Wennerstrom
Keyword(s):  
Static Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Compressors ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Pressure Distributions ◽  
Blade Section ◽  
Flow Compressor

A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performances of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.

Application of Casing Circumferential Grooves for Improved Stall Margin in a Transonic Axial Compressor

2002 ◽  
Author(s):  
D. C. Rabe ◽  
C. Hah
Keyword(s):  
Incidence Angle ◽  
Axial Compressor ◽  
Numerical Procedure ◽  
Leading Edge ◽  
Navier Stokes ◽  
Stall Margin ◽  
Casing Treatment ◽  
Transonic Compressor ◽  
Flow Mechanisms ◽  
Groove Configurations

Experimental and numerical investigations were conducted to study the fundamental flow mechanisms of circumferential grooves in the casing of a transonic compressor and their influence on compressor stall margin. Three different groove configurations were tested in a highly loaded transonic compressor. Experimental results show that circumferential grooves increase the stall margin of the compressor at the tested operating condition. Grooves with a much smaller depth than conventional designs are shown to be similarly effective in increasing the stall margin. Steady-state Navier-Stokes analyses were performed to study flow structures associated with each casing treatment. The numerical procedure calculates the overall effects of the circumferential grooves correctly. Detailed investigation of calculated flow fields indicates that losses are generated by interaction between the main passage flow and flow exiting the grooves. The grooves increase the stall margin by reducing the flow incidence angle on the pressure side of the leading edge, despite an overall increase in the endwall boundary layer thickness. This is due to complex interaction of the main passage flow with the additional radial and tangential flows created by the grooves.

Three-Dimensional Flows and Loss Reduction in Axial Compressors

1987 ◽  
Vol 109 (3) ◽  
pp. 354-361 ◽  
Author(s):  
Y. Dong ◽  
S. J. Gallimore ◽  
H. P. Hodson
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Tracer Gas ◽  
Suction Surface ◽  
Axial Compressors ◽  
Small Clearance ◽  
Oil Flow ◽  
Stator Blade ◽  
Flow Compressor

Measurements have been performed in a low-speed high-reaction single-stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface–hub corner separations, their associated loss mechanisms, and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualizaton indicated that the leakage reduced the extensive suction surface–hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.

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.

A Proposal for Passive Wall Treatment Applied in High Performance Axial Flow Compressors

2020 ◽  
Author(s):  
Rubén Bruno Díaz ◽  
Jesuino Takachi Tomita ◽  
Cleverson Bringhenti ◽  
Francisco Carlos Elizio de Paula ◽  
Luiz Henrique Lindquist Whitacker
Keyword(s):  
Stokes Equations ◽  
Axial Compressor ◽  
Axial Flow ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Smooth Wall ◽  
Tip Leakage Flow ◽  
Viscous Effects ◽  
Stall Margin ◽  
Tip Leakage

Abstract Numerical simulations were carried out with the purpose of investigating the effect of applying circumferential grooves at axial compressor casing passive wall treatment to enhance the stall margin and change the tip leakage flow. The tip leakage flow is pointed out as one of the main contributors to stall inception in axial compressors. Hence, it is of major importance to treat appropriately the flow in this region. Circumferential grooves have shown a good performance in enhancing the stall margin in previous researches by changing the flow path in the tip clearance region. In this work, a passive wall treatment with four circumferential grooves was applied in the transonic axial compressor NASA Rotor 37. Its effect on the axial compressor performance and the flow in the tip clearance region was analyzed and set against the results attained for the smooth wall case. A 2.63% increase in the operational range of the axial compressor running at 100%N, was achieved, when compared with the original smooth wall casing configuration. The grooves installed at compressor casing, causes an increase in the flow entropy generation due to the high viscous effects in this gap region, between the rotor tip surface and casing with grooves. These viscous effects cause a drop in the turbomachine efficiency. For the grooves configurations used in this work, an efficiency drop of 0.7% was observed, compared with the original smooth wall. All the simulations were performed based on 3D turbulent flow calculations using Reynolds Averaged Navier-Stokes equations, and the flow eddy viscosity was determined using the two-equation SST turbulence model. The details of the grooves geometrical dimensions and its implementation are described in the paper.

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