Experimental study of effects of one kind of tangentially non-uniform tip clearance on the flow field at an exit of a compressor cascade passage

This paper presents an experimental investigation of effects of one kind of tangentially non-uniform tip clearance on the flow field at an exit of a compressor cascade passage.The tests were performed in a low-speed large-scale cascade with the uniform tip clearance and the non-uniform clearance.The three-dimensional flow field was measured at the exit at three incidence angles of 0°,5°,and 8° using a mini five-hole pressure probe.The measurement results show that the non-uniform tip clearance can moderate the leakage flow and blow down more low-energy fluids at the tip corner and decrease the accumulation of low-energy fluids which cause the flow blockage in the blade passage.In the meantime,the non-uniform clearance can weaken the tangential migration of the low-energy fluids in the endwall boundary layer and reduce the secondary loss and the flow blockage in the tip region.     

Boundary layer separation control on a highly-loaded, low-solidity compressor cascade

Separated flow can be effectively controlled through the management of blade boundary layer development. Numerical simulations on a highly-loaded, low-solidity compressor cascade indicate that combined blowing and suction flow control technique can significantly improve cascade performance, especially in increasing the cascade loading and static pressure ratio as well as decreasing the loss coefficient. Meanwhile, it is more effective to improve cascade performance by blowing near leading edge on suction surface than suction near trailing edge. Both the locations and flow rates of blowing and suction are major impact factors of this method to cascade performance. Comparing to the baseline, the static pressure ratio increases by 15% and loss coefficient decreases by 80%, with a blowing fraction of 1.7% and a suction fraction of 1.38% of the inlet mass flow.     

Effects of airfoil-probe tubes on the flow field of a compressor cascade

To explore the effects of airfoil-probe tubes and its installment position on the flow field of the compressor cascade,and find out the mechanism that how the airfoil-probes affect the aerodynamic characteristics of the compressor cascade,this paper performed both numerical and experimental works on the same compressor cascade.The experiment mainly focused on the cases of low Mach number (Ma =0.1),and cases with different Mach numbers (0.1,0.3,0.7) and different incidence angles (-5,0,5) are investigated by the numerical method.The case without the airfoil-probe tube was referenced as the baseline,and other three cases with the airfoil-probe tubes installed in different chordwis positions (30％,50％,70％ of the chord length) were studied.The diameter of the airfoil-probe tube is 3ram,which is configured as 300％ amplification of some particular airfoil-probe according to the geometrical similarity principle.The results show that the airfoil-probe tubes have a negative influence on the flow capacity of the cascade at all investigation points.The separations and the large scale streamwise vortices that induced by the airfoil-probe tube on the pressure side cause most the losses at the high Mach number.The influence of the airfoil-probe tube on the flow field in the vicinity of the pressure side surface is local separation at the low Mach number.The airfoil-probe tubes also have a clearly effect on the leakage flow.It decreases the mass flow of the leakage flow and weakens the intensity of the leakage vortex,but enlarges the influence area.The total pressure loss of the case that the tube is installed at the half chordwise position is generally lower than other cases especially at the high Mach number,it can even decrease the losses compared with the basic case.     

Effect of winglet geometry arrangement and incidence on tip clearance control in a compressor cascade

An experimental study is conducted to investigate the influences of blade tip winglet on the flow field of a compressor cascade. The tests are performed in a low speed linear cascade with stationary endwall, with three blade tip configurations, including the baseline tip, the suction-side winglet tip and the pressure-side winglet tip. The flowfield downstream of the cascade is measured using five-hole probe, from which the three-dimensional velocity field, vorticity field and pressure field are obtained. Static pressure measurements are made on the endwall above the blade row using pressure taps embedded in the plywood endwall. All measurements are made at both design and off-design conditions for tip clearance level of about 2 percent of the blade chord. The results revealed the incidence variation significantly affects the secondary flow and the associated loss field downstream of the cascade, where the tip leakage vortex and passage vortex exist as the major contributors on the field. The winglet geometry arrangements can change the trajectory of the tip leakage vortex. The suction-side winglet tip blade provides a lower overall total pressure loss coefficient when compared to the baseline tip blade and pressure-side winglet tip blade at all incidence angles.     

The influence of synthetic jet excitation on secondary flow in compressor cascade

The detailed numerical simulation has been carried out to investigate the effect of synthetic jet excitation on the secondary flow at 5° incidence in a compressor cascade, in which the synthetic jet actuation is equipped on the suction surface. The influence of excitation position, one fixed near the trailing edge and the other fixed a little far from the trailing edge, has also been studied. The results show that unsteady disturbance of desirable synthetic jet effectively enhances the mixing of the fluid inside the separation region, which reduces the vortex intensity and the energy loss, improves the flow status in the cascade, and also suppresses velocity fluctuation near the trailing edge. Additionally, the actuation fixed near the separation region proves to be more effective and exit load distribution is more uniform due to the employment of the synthetic jet.     

Investigation of endwall flow behavior with plasma flow control on a highly loaded compressor cascade

To discover the flow behavior in the endwall region and mechanism of plasma flow control on a highly loaded compressor cascade, distributions of static pressure coefficient, total pressure loss coefficient and streamline pattern were investigated. Results show that cross flow from the pressure surface to neighboring suction surface exists under pitch-wise pressure gradient. The deflected endwall boundary layer flow interacts with the incoming flow, and then both of them leave off the endwall in the form of a span-wise vortex. Effect of angle of attack on static pressure is greater than that of free stream velocity. The distinct variations of total pressure loss with endwall actuations are mainly located within the outer verge of a triangular area with high total pressure loss. Effect of pitch-wise actuation on separated flows is much better than that of stream-wise actuation, and both enhance with the increase of angle of attack and actuation strength. An efficient method for plasma flow control in the endwall region is the increase of actuation strength, such as adjusting discharge voltage or changing plasma power supply.     

Experimental and numerical investigation of the unsteady tip leakage flow in axial compressor cascade

For convenience of both measurement and adjusting the clearance size and incidence, the current research is mainly conducted by experiments on an axial compressor linear cascade. The characteristics and the condition under which the unsteadiness of tip leakage flow would occur were investigated by dynamic measuring in different clearances, inlet velocities and incidences. From the experiment it is found that increasing tip clearance size or reducing rotor tip incidence can affect the strength of the tip clearance flow. Then the experimental results also indicate the tip leakage shows instability in certain conditions, and the frequency of unsteadiness is great influenced by inflow angle. The condition of occurrence of tip leakage flow unsteadiness is when the leakage flow is strong enough to reach the pressure side of the adjacent blade. The main cause of tip leakage flow unsteadiness is the tip blade loading.     

Numerical investigation of aerodynamic performance influenced by circumferentially pre-swirling coming flow on the crossover and de-swirling cascade of a multistage centrifugal compressor

Numerical investigation is implemented on aerodynamic performance inside the crossover and de-swirling cascade of a multistage centrifugal compressor. The emphasis is put on the aerodynamic performance influenced by the circumferentially pre-swirling coming flow. The results indicate that flow separation occurs inside the crossover, and the separation area may be changed with different circumferentially pre-swirling coming flow. Decreased pre-swirling intensity of the coming flow may effectively restrain the flow separation and make the outflow from the crossover more uniform, which helps to improve the aerodynamic performance of the successive de-swirling cascade. The flow inside the de-swirling cascade is a non-uniform swirling flow with large separation. The complex secondary flow occurs along the main flow and experiences a process of generating, developing, dissipating and collapsing. __________ Translated from Power Engineering, 2007, 27(1): 24–28, 49 [译自: 动力工程]     

Effects of a kind of non-smooth blade on the unsteady flow field at the exit of an axial fan

An experimental investigation of effects of a kind of streamwise-grooved blade on the unsteady flow field at an exit of an axial-flow fan was performed. The flow field at 25% chord downstream from the trailing edge at hub was measured using a fast-response five-hole pressure probe at different mass-flow conditions. The unsteady flow of the grooved blades was compared with that of the smooth blades. The measurement results indicate that： （1） the grooved blades restrain the velocity fluctuation and the pressure fluctuation by modulating the blade boundary layers, which contributes to the flow loss reduction in the hub region and in the rotor wake region at the design condition; （2） the stream-wise grooves play an important role in restraining the radial migration in the blade boundary layer and abating the tip flow mixing, which contributes to the flow loss reduction in the tip region at the design condition; （3） at the near stall condition, the grooved surface can not reduce the flow loss, even increase the loss nearby when the separation happens in the blade boundary layer.     

Frequency Analyses for Mechanism Research of Unsteady Flow Inside Three Dimensional Compressor Cascade

Numerical simulations are carried out for unsteady flow field of certain kind of 3D compressor cascade. Emphasis is laid on vortex shedding and frequency analysis in a compressor cascade. Numerical simulations using unsteady Reynolds-averaged viscous turbulent equations are carried out. The results show that the flows in separated areas and wake areas are characterized as periodic or quasi-periodic vortex shedding and the frequencies of vortex shedding vary with incidences and Mach number. At the same Mach number, the frequency of vortex shedding will decrease as the incidence increases. Yet, the frequency will increase as the Mach number increases at the same incidence. In the same computation case, the frequency of vortex shedding will vary along the span of blade. The frequency is smaller at the middle of the span than that at the hub.     

Numerical investigation on the effects of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade

This paper presents a detailed numerical investigation of the influence of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade. The boundary layer suction (BLS) was used to control the location and strength of shock waves, with the aspirated slot locating at 49% chord, where is just downstream of the impingement point of shock wave at the leading edge. The numerical simulation is based on NUMECA, a commercial software, where the cell-centered control volume approach with third-order spatial accuracy is used to solve the 3-D Reynolds-averaged Navier-Stokes equations under the Cartesian coordinate system. Several conclusions can be made through the observation of the numerical results. (1) Multiple shock waves in cascade passage leaded the velocity deficits of boundary layer on suction surface downstream of shock wave, resulting in seriously separated flow on the suction side of blade, especially when the front shock wave is much stronger than the rest of the shocks. (2) BLS with small mass flow rate can not effectively improve the boundary layer. When the impingement point of oblique shock wave coming from cascade leading edge is bled to downstream of the passage shock wave by BLS, the boundary layer flow is greatly improved. However, if the BLS mass flow rate exceeds a critical value (1.2%), the boundary layer downstream of shock wave would separate from suction surface. (3) At the blade mid-span, the aerodynamic performance of compressor blade is improved as BLS mass flow rate increases. The optimum BLS is about 1.2%. Compared with the baseline case, the BLS with flow rate of 1.2% increases the total pressure recovery coefficient by 12%, and decreases diffusion factor by 18% and deviation angle to 7 ° while keeping the pressure rise constant. (4) The three dimensional flow structure of the compressor cascade ranged from 25% span to 75% span was improved greatly with the 1.2% BLS flow rate. However it could not control the development of the corner boundary layer effectively.     

新型极小展弦比叶栅的叶型研究

以数值手段分析研究了具有不同极小展弦比的改型设计的后加载透平叶栅的内部定常三维粘性流动及损失，同时对该透平叶栅在不同出口马赫数和进气角下进行平面叶栅吹风实验。数值分析结果和实验结果表明，后加载叶栅在极小展弦比下能有效控制叶栅内二次流动的形成，具有低的三维叶栅损失。     

A study of the unsteady tip flow field of a transonic compressor

An experimental investigation on the unsteady tip flow field of a transonic compressor rotor has been performed.The casing-mounted high frequency response pressure transducers were arranged along both the blade chord and the blade pitch.The chord-wise ones were used to indicate both the ensemble averaged and time varying flow structure of the tip region of the rotor at different operating points under 95% design speed and 60% design speed.The pitch-wise circumferential transducers were mainly used to analyze the unsteadiness frequency of the tip leakage flow in the rotor frame at the near stall condition.The contours of casing wall pressure show that there were two clear low pressure regions in blade passages,one along the chord direction,caused by the leakage flow and the other along the tangential direction,maybe caused by the forward swept leading edge.Both low pressure regions were originated from the leading edge and formed a scissor-like flow pattern.At 95% design speed condition,the shock wave interacted with the low pressure region and made the flow field unsteady.With the mass flow reduced,the two low pressure regions gradually contracted to the leading edge and then a spike disturbance emerged.     

Experimental Investigation of the Three—Dimensional Flow in an Annular Compressor Cascade at Large Incidence

This paper presents a detailed experimental investigation concerning the influence of blade loading (incidence)on the three-dimensional flow in an annular compressor cascade.The data are acquired at four incidence anglesunder low Mach number and low Reynolds number conditions.Experimental techniquss include the oil-filmvisualization on the profile and the endwall surfaces,the laser-sheet visualization of the flow field inside theblade passage,and the measurement by radial-circumferential traveress using a seven-hole probe.The behaviorand nature of the three-dimensional flow with severe separations inside the blade passage and at the exit areobtained.The distributions of the total pressure loss,static pressure,velocity and outflow angle are also given.These results are valuable for establishing the physical model of the three-dimensional complex flow in axialcompressors and for examining the computational procedures.     

正倾斜叶片压气机叶栅二次流的数值研究

应用Beam-Warming近似隐式因子分解格式以及MML代数湍流模型,采用拟压缩性方法求解雷诺平均拟压缩N－S方程组,对正倾斜叶片压气机叶栅内三维粘性流场进行了数值研究,并与直叶栅进行了对照。结果发现,正倾斜叶栅中上、下通道涡的发生、发展过程与直叶栅存在明显的差异,这导致正倾斜侧二次流减弱,负倾斜侧二次流高损失区扩大,流动状况恶化,叶栅顶部区域的附面层分离发展成一个向叶栅中部扩展的更大的区域。计算与实验结果比较,两者吻合较好。     

Experimental investigation on flow improvement in compressor cascades

There is no general rule in the literature to help choose a correct flow control device for any given case of turbomachinery applications. This suggests individual optimization of flow control devices for each specific case. The objective of this study is to prove experimentally the benefits of passive control methods in improving the compressor performance. This allows to reduce the fuel consumption, leading to energy saving and reduction of atmospheric pollution. Two features have been controlled in this study: flow separation over the blade surfaces and the secondary flow over the cascade endwalls. Vortex generator ribs are tested on the blade suction side for flow reattachment on the blade surfaces, and low‐profile vortex generators are tested on the side walls of the compressor cascade against secondary flow losses. Different vortex generator designs are compared for total pressure recovery, flow turning, boundary layer characteristics, and pressure distributions over the endwalls. Copyright © 2016 John Wiley & Sons, Ltd.     

An Experimental Study on 3—D Flow in an Annular Cascade of High Turning Angle Turbine Blades

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用CFD研究涡轮静叶栅的二次流损失

利用CFD软件Fluent对大转折角涡轮叶栅三维流场进行了数值模拟。采用静叶栅前移动的圆柱列替代上游动叶,发现圆柱尾迹进入叶栅流道的位置不同,对叶栅总压损失有较大影响。同时,通道内逐渐增大的横向压力梯度对二次涡发展产生了显著的影响,引起沿流向叶栅总压损失的急剧增大,认为叶高的减小会极大提高叶栅的二次流损失。     

Numerical analysis of the tip leakage flow field in a transonic axial compressor with circumferential casing treatment

This paper reports on numerical investigations aimed at understanding the influence of
circumferential casing grooves on the tip leakage flow and its resulting vortical structures.The results
and conclusions are based on steady state 3D numerical simulations of the well-known transonic axial
compressor NASA Rotor 37 near stall operating conditions.The calculations carried out on the casing
treatment configuration reveal an important modification of the vortex topology at the rotor tip
clearance.Circumferential grooves limit the expansion of the tip leakage vortex in the direction
perpendicular to the blade chord,but generate a set of secondary tip leakage vortices due to the
interaction with the leakage mass flow.Finally,a deeper investigation of the tip leakage flow is
proposed.     

不同参数端壁翼刀控制二次流机理的实验研究

不同周向和轴向位置的压气机叶栅上安装1／2轴向弦长翼刀的叶栅出口流场测量结果表明,两种方案的叶栅总损失随翼刀周向位置变化的总体趋势是翼刀靠近压力面时叶栅总损失降低。翼刀安装在流道前半部的最佳周向位置是距离吸力面60％相对节距处;安装在流道后半部的翼刀最佳周向位置是距离吸力面80％相对节距处。通过对比初步探讨了翼刀减小二次流损失的机理：一方面通过降低流道内端壁附面层内横向压力梯度,减弱低能流体向吸力面／壁角区的堆积;另一方面是通过产生的反向翼刀涡限制马蹄涡的压力面分支发展,从而减小通道涡的尺寸和强度。