The Effective Positions to Inject Water Into the Cascade of Compressor

2012 ◽  
Author(s):  
Jie Wang ◽  
Qun Zheng ◽  
Lanxin Sun ◽  
Mingcong Luo
Keyword(s):  
Kinetic Energy ◽  
Total Pressure ◽  
Lower Energy ◽  
Water Injection ◽  
Adverse Pressure Gradient ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Pressure Losses ◽  
Ansys Cfx ◽  
Highly Loaded

Generally, droplets are injected into air at inlet or interstage of a compressor. However, both cases did not consider how to utilize the kinetic energy of these moving droplets. Under the adverse pressure gradient of compressor, the lower energy fluids of blade surfaces and endwalls boundary layers would accumulate and separate. Kinetic droplets could accelerate the lower energy fluids and eliminate the separation. This paper mainly investigate the effective positions where to inject water and how to utilize the droplets’ kinetic energy. Four different injecting positions, which located on the suction surface and endwall, are chosen. The changes of vortexes in the compressor cascade are discussed carefully. In addition, the influences of water injection on temperature, total pressure losses and Mach number are analyzed. Numerical simulations are performed for a highly loaded compressor cascade with ANSYS CFX software.

Effects of Air Injection on Performance of Highly-Loaded Compressor Cascades

2007 ◽  
Author(s):  
Yanping Song ◽  
Huanlong Chen ◽  
Fu Chen ◽  
Zhongqi Wang
Keyword(s):  
Performance Improvement ◽  
Adverse Pressure Gradient ◽  
Air Injection ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Capacity Enhancement ◽  
Straight Blade ◽  
Flow Capacity ◽  
Blade Cascade ◽  
Highly Loaded

The effects of air injection on the performance of highly-loaded straight blade compressor cascade and compound lean blade cascade were investigated numerically. Air injection was implemented via the hole/slot penetrating through the pressure and suction surfaces under the effect of the pressure difference between the two sides. Various injection configurations including one-hole, multi-hole and slot configurations were studied in the straight blade cascade first to find the optimum injection configuration in this case. Then the mechanism of the injection slot on the performance improvement of the straight blade cascade was discussed in detail. The results show that air injection provides the low-momentum fluid near the suction surface with kinetic energy, to enhance its ability to withstand the adverse pressure gradient within the compressor cascade. Among the hole/slot configurations, the slot configuration (the radial width of 4.0mm) has the most favorable results in through-flow capacity enhancement and total loss reduction, and is used in the compound lean cascade study. The effects of air injection on the performance improvement in the compound lean blade cascade are less significant than those in the straight blade cascade though the compound lean blade cascade with the injection slot has the most improvement in the aerodynamic performance when the slot is placed at an appropriate location in this study.

Flow control using unsteady pulsed holed suction with different excitation models in a highly loaded compressor cascade

2017 ◽  
Vol 232 (6) ◽  
pp. 593-607 ◽  
Author(s):  
Hongxin Zhang ◽  
Shaowen Chen ◽  
Yun Gong ◽  
Songtao Wang
Keyword(s):  
Flow Control ◽  
Pressure Loss ◽  
Total Pressure ◽  
Control Flow ◽  
Total Pressure Loss ◽  
Control Technique ◽  
Compressor Cascade ◽  
Pressure Losses ◽  
Optimal Excitation ◽  
Highly Loaded

Unsteady pulsed holed suction as a new unsteady flow control technique is first proposed. Unsteady excitation models of four different waveforms (Waveforms 1, 2, 3, and 4) based on unsteady pulsed holed suction are investigated to analysis comparatively the control effects of flow separations in a certain highly loaded compressor. Some related unsteady aerodynamic parameters such as excitation frequency and excitation location are studies. The unsteady pulsed holed suctions of the four different modes (Waveforms 1, 2, 3, and 4) all effectively control flow separations. Their optimum frequencies are all an integer multiple of the natural frequency of vortex shedding. And their excitation locations gaining positive effect and optimal excitation locations are both same. The optimal excitation location is near the separation point of upper endwall in unexcited case. But, they show markedly different performances in reducing the total pressure losses. The unsteady pulsed holed suction of Waveform 3 shows greater advantage at different excitation frequencies and excitation locations. The optimum result is obtained by the unsteady pulsed holed suction of Waveform 3. The total pressure loss is reduced by 16.8%. Simultaneously, the unsteady pulsed holed suctions of the four different modes all can provide better effects than the steady constant holed suction in reducing the total pressure loss with the same suction-to-inlet time-averaged suction flow ratio ms. Especially at ms = 0.29%, for the steady constant holed suction, it is too small to effectively control flow separation, and consequently the total pressure loss are increased by 8.3%. However, for the unsteady pulsed holed suctions of Waveforms 2 and 3, the total pressure losses are reduced by 9.1% and 4.3%, respectively.

Corner Separation Control in a Highly Loaded Compressor Cascade Using Plasma Aerodynamic Actuation

2012 ◽  
Author(s):  
Yun Wu ◽  
Xiao-hu Zhao ◽  
Ying-hong Li ◽  
Jun Li
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Separation Control ◽  
Compressor Cascade ◽  
Control Effect ◽  
Suction Surface ◽  
Corner Separation ◽  
Pressure Loss Coefficient ◽  
Highly Loaded

Corner separation, which forms over the suction surface and endwall corner of a blade passage, causes significant total pressure loss in highly loaded compressors. Plasma flow control, based on the plasma aerodynamic actuation, is a novel active flow control technique to improve aircrafts’ aerodynamic characteristics and propulsion efficiency. This paper reports computational and experimental results on using three types of plasma aerodynamic actuation (PAA) to control the corner separation in a highly loaded, low speed, linear compressor cascade. Reynolds-Averaged Navier-Stokes simulations were performed to optimize the PAA arrangement. The PAA was generated by a microsecond or nanosecond dielectric barrier discharge in wind tunnel experiments. The total pressure loss coefficient distribution was adopted to evaluate the corner separation control effect. The control effect of pitch-wise PAA on the endwall, in terms of relative reduction of the pitch-wise averaged total pressure loss coefficient in the wake, is much better than that of stream-wise PAA on the suction surface. When both pitch-wise PAA on the endwall and stream-wise PAA on the suction surface are turned on simultaneously, the control effect is the best among all three types of PAA. The main effect of pitch-wise PAA on the endwall is to inhibit the crossflow from neighboring pressure surface to the suction surface, whilest the main effect of stream-wise PAA on the suction surface is to inhibit the boundary layer accumulation and separation. Compared to microsecond discharge PAA, nanosecond discharge PAA is more effective at higher freestream velocity. The mechanisms for nanosecond discharge and microsecond discharge PAA are different for corner separation control.

Numerical Investigation of Passive Flow Control Using Wavy Blades in a Highly-Loaded Compressor Cascade

2018 ◽  
Author(s):  
Bo Wang ◽  
Yanhui Wu ◽  
Kai Liu
Keyword(s):  
Numerical Investigation ◽  
Flow Structure ◽  
Cross Flow ◽  
Leading Edge ◽  
Control Flow ◽  
Streamwise Vortices ◽  
Compressor Cascade ◽  
Control Effect ◽  
Suction Surface ◽  
Highly Loaded

Driven by the need to control flow separations in highly loaded compressors, a numerical investigation is carried out to study the control effect of wavy blades in a linear compressor cascade. Two types of wavy blades are studied with wavy blade-A having a sinusoidal leading edge, while wavy blade-B having pitchwise sinusoidal variation in the stacking line. The influence of wavy blades on the cascade performance is evaluated at incidences from −1° to +9°. For the wavy blade-A with suitable waviness parameters, the cascade diffusion capacity is enhanced accompanied by the loss reduction under high incidence conditions where 2D separation is the dominant flow structure on the suction surface of the unmodified blade. For well-designed wavy blade-B, the improvement of cascade performance is achieved under low incidence conditions where 3D corner separation is the dominant flow structure on the suction surface of the baseline blade. The influence of waviness parameters on the control effect is also discussed by comparing the performance of cascades with different wavy blade configurations. Detailed analysis of the predicted flow field shows that both the wavy blade-A and wavy blade-B have capacity to control flow separation in the cascade but their control mechanism are different. For wavy blade-A, the wavy leading edge results in the formation of counter-rotating streamwise vortices downstream of trough. These streamwise vortices can not only enhance momentum exchange between the outer flow and blade boundary layer, but also act as the suction surface fence to hamper the upwash of low momentum fluid driven by cross flow. For wavy blade-B, the wavy surface on the blade leads to a reduction of the cross flow upwash by influencing the spanwise distribution of the suction surface static pressure and guiding the upwash flow.

Investigation Into the Effects of Hub Rotation on the Hub Leakage Flow of Cantilever Stator in a Transonic Axial Compressor

2020 ◽  
Author(s):  
Botao Zhang ◽  
Bo Liu ◽  
Xin Sun ◽  
Hang Zhao
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Root ◽  
Flow Loss ◽  
Total Pressure Ratio

Abstract In order to explore the similarities and differences between the flow fields of cantilever stator and idealized compressor cascade with tip clearance, and to extend the cascade leakage model to compressors, the influence of stator hub rotation to represent cascade and cantilever stator on hub leakage flow was numerically studied. On this basis, the control strategy and mechanism of blade root suction were discussed. The results show that there is no obvious influence on stall margin of the compressor whether the stator hub is rotating or stationary. For rotating stator hub, the overall efficiency is decreased while the total pressure ratio is increased. At peak efficiency point and near stall point, the efficiency is reduced by about 0.43% and 0.34% individually, while the total pressure ratio is enlarged by about 0.23% and 0.27%, respectively. The gap leakage flow is promoted due to stator hub rotation, and the structure of the leakage vortex is weakened obviously. In addition, the hub leakage flow originating from the blade leading edge of rotating hub may contribute to double leakage near the trailing edge of the adjacent blade. However, the leakage flow directly out of the blade passage with stationary stator hub. The stator root loading and strength of the leakage flow increase with the rotation of the hub, and the leakage vortex is further away from the suction surface of the blade and is stretched to an ellipse closer to the endwall under the shear action. The rotating hub makes the flow loss near the stator gap increase, while the flow loss in the upper part of the blade root is decreased. Meanwhile, the total pressure ratio in the end area is increased. Blade root suction of cantilever stator can effectively control the hub leakage flow, inhibit the development of hub leakage vortex, and improve the flow capacity of the passage, thereby reducing the flow loss and modifying the flow field in the end zone.

The compound bowing design in a highly loaded linear cascade with large turning angle

2020 ◽  
Vol 234 (16) ◽  
pp. 2323-2336
Author(s):  
Xingxu Xue ◽  
Songtao Wang ◽  
Lei Luo ◽  
Xun Zhou
Keyword(s):  
Numerical Simulation ◽  
Reynolds Number ◽  
Pressure Gradient ◽  
Side Effect ◽  
Adverse Pressure Gradient ◽  
Suction Surface ◽  
Turning Angle ◽  
Linear Cascade ◽  
Velocity Triangle ◽  
Highly Loaded

Numerical simulation was carried out to study the influences of blade-bowing designs based on a highly loaded cascade with large turning angle, while the compound bowing design showed much lower endwall loss than the conventional design in this study. Generally, it showed that the increased turning angle would strengthen the adverse pressure gradient on the suction surface, so the side effect of negative blade bowing angle would be enhanced because of the reduced flow filed stability near suction–endwall corner. However, the positive corner bowing angle that applied in the compound bowing design would enhance the flow field stability near the suction–endwall corner by adjusting spanwise pressure gradient and velocity triangle, so the side effect of negative blade bowing angle would be suppressed and lead to weaker secondary flow. In detail, the blade bowing angle (as well as the corner bowing angle in the conventional bowed cascades) was varied from −5° to −30° in this study, while the reductions of the loss coefficient in the compound bowed cascades were about 0.662.16 times higher (the absolute differences were about 0.0067 0.0097) than the corresponding conventional bowed cascades. Moreover, the Reynolds number and Mach number at the outlet plane were kept at 2.4 × 105 and 0.6, respectively, during the bowing design to ensure the comparability.

scholarly journals Effects of Surface Roughness, Freestream Turbulence, and Incidence Angle on the Performance of a 2-D Compressor Cascade

1990 ◽  
Author(s):  
W. C. Elrod ◽  
P. I. King ◽  
E. M. Poniatowski
Keyword(s):  
Surface Roughness ◽  
Total Pressure ◽  
Incidence Angle ◽  
Two Dimensional ◽  
Freestream Turbulence ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Linear Cascade ◽  
Pressure Loss Coefficient ◽  
Total Pressure Loss Coefficient

The effects of surface roughness, freestream turbulence, and incidence angle on the performance of a two-dimensional compressor cascade were investigated. The test section consisted of seven NACA 65-A506 airfoils arranged in a linear cascade. Four different surface roughness conditions were applied to the first 25 percent chord on the suction surface of each of the five middle blades in the cascade. Freestream turbulence levels of approximately one and seven percent were used. Incidence angles of −3, zero and +3 degrees were investigated. Of the three parameters tested, freestream turbulence exerted the largest influence on blade performance. The total pressure loss coefficient increased with increased roughness and was reduced for large turbulence. Changes in flow incidence had a lesser effect on the performance of the blade.

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.

Transonic Aerodynamic Losses Due to Turbine Airfoil, Suction Surface Film Cooling

1999 ◽  
Vol 122 (2) ◽  
pp. 317-326 ◽  
Author(s):  
D. J. Jackson ◽  
K. L. Lee ◽  
P. M. Ligrani ◽  
P. D. Johnson
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Surface Film ◽  
Suction Surface ◽  
Pressure Losses ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Mixing Losses ◽  
Aerodynamic Losses

The effects of suction surface film cooling on aerodynamic losses are investigated using an experimental apparatus designed especially for this purpose. A symmetric airfoil with the same transonic Mach number distribution on both sides is employed. Mach numbers range from 0.4 to 1.24 and match values on the suction surface of airfoils from operating aeroengines. Film cooling holes are located on one side of the airfoil near the passage throat where the free-stream Mach number is nominally 1.07. Round cylindrical and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and a correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical holes are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil, as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane. [S0889-504X(00)02202-9]

scholarly journals Transonic Aerodynamic Losses due to Turbine Airfoil, Suction Surface Film Cooling

1999 ◽  
Author(s):  
D. J. Jackson ◽  
K. L. Lee ◽  
P. M. Ligrani ◽  
P. D. Johnson ◽  
F. O. Soechting
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
Surface Film ◽  
Suction Surface ◽  
Pressure Losses ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Mixing Losses ◽  
Aerodynamic Losses

The effects of suction surface film cooling on aerodynamic losses are investigated using an experimental apparatus designed especially for this purpose. A symmetric airfoil with the same transonic Mach number distribution on both sides is employed. Mach numbers along the airfoil surface range from 0.4 to 1.24 and match values on the suction surface of airfoils from operating aeroengines. Film cooling holes are located on one side of the airfoil near the passage throat where the freestream Mach number is nominally 1.07. Round cylindrical, and conical diffused film cooling hole configurations are investigated with density ratios from 0.8 to 1.3 over a range of blowing ratios, momentum flux ratios, and Mach number ratios. Also included are discharge coefficients, local and integrated total pressure losses, downstream kinetic energy distributions, Mach number profiles, and n correlation for integral aerodynamic losses as they depend upon film cooling parameters. The contributions of mixing and shock waves to total pressure losses are separated and quantified. These results show that losses due to shock waves vary with blowing ratio as shock wave strength changes. Aerodynamic loss magnitudes due to mixing vary significantly with film cooling hole geometry, blowing ratio, Mach number ratio, and (in some situations) density ratio. Integrated mixing losses from round cylindrical boles are three times higher than from conical diffused holes, when compared at the same blowing ratio. Such differences depend upon mixing losses just downstream of the airfoil as well as turbulent diffusion of streamwise momentum normal to the airfoil symmetry plane.

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