Stall Behaviour in a Mixed-flow Compressor with Axial Slot Casing Treatment

Casing treatment is an effective technique in extending stall margin of axial and centrifugal compressor.However,its impacts on the stall behaviour of mixed-flow compressor are still not completely understood until now.To conquer this issue,unsteady full-annulus simulations were conducted to investigate the stall mechanism of a mixed-flow compressor with and without axial slot casing treatment(ASCT).The circumferential propagating speed of spike inception resolved by the numerical approach is 87.1%of the shaft speed,which is identical to the test data.The numerical results confirmed that the mixed-flow compressor fell into rotating stall via spike-type with and without ASCT.The flow structure of the spike inception was investigated at 50%design rotational speed.Instantaneous static pressure traces extracted upstream of the leading edge had shown a classic spiky wave.Furthermore,it was found that with and without ASCT,the mixed-flow compressor stalled through spike with the characteristic of tip leakage spillage at leading edge and tip leakage backflow from trailing edge,which is different from a fraction of the centrifugal compressor.The resultant phenomenon provides conoborating evidence for that unlike in axial-flow compressor,the addition of ASCT does not change the stall characteristics of the mixed-flow compressor.The flow structure that induced spike inception with ASCT is similar to the case with smooth casing.In the throttling process,tip leakage flow vortex had been involved in the formation of tornado vortices,with one end at the suction side,and the other end at the casing-side.The low-pressure region relevant to the downward spike is caused by leading-edge separation vortex or tornado vortex.The high-pressure region relevant to the upward spike is induced by blockage from the passage vortex.These results not only can provide guidance for the design of casing treatment in mixed-flow compressor,but also can pave the way for the stall waring in the highly-loaded compressors of next-generation aeroengines.     

零间隙压气机失稳机理的全通道数值模拟研究

为了探究零间隙压气机流动失稳机理,采用全通道非定常数值模拟方法研究了一台零间隙斜流压气 机转子的失稳机理,数值模拟过程中在转子出口施加了随时间动态变化的背压模拟压气机转子节流,非定常 数值计算结果表明零间隙斜流压气机转子仍然表现为典型突尖流动失稳特征。通过详细地分析斜流压气机 转子节流过程中不同阀系数对应的压气机内部流场结构,结果表明:尽管零间隙斜流压气机无叶顶泄漏特 征,但随着对压气机节流,转子叶片尾缘率先出现流动分离,进一步节流,尾缘流动分离表现为一方面在周向 范围加剧,另一方面分离点逐渐向上游移动,造成通道严重堵塞,最终引发相邻叶片通道尾缘回流和叶片前 缘流动溢出进而诱发叶片通道内部出现径向涡结构,从而形成压气机突尖失速先兆。     

基于弯曲叶片的氦气轴流压气机优化设计

针对高负荷氦气压气机中角区分离、叶顶泄漏严重带来的效率损失问题,以单级氦气压缩机为研究对象,利用CFD方法,分析了不同弯曲角度下氦气压气机内部的角区损失和叶顶泄漏损失,并优化了现有五级轴流氦气压气机。结果表明：叶片正弯会增加端区处的静压,减少角区分离,进而降低角区损失;对动叶而言,在设计攻角下正弯也会增加前缘损失;动叶叶顶反弯使泄漏流远离下一个叶片的压力面,而合适的反弯角度可以降低叶顶泄漏量;选取合适的弯曲角度使五级轴流压气机设计点效率提高1.85%。     

Effects of axial non-uniform tip clearances on aerodynamic performance of a transonic axial compressor

This paper presents a numerical investigation of effects of axial non-uniform tip clearances on the aerodynamic performance of a transonic axial compressor rotor （NASA Rotor 37）. The three-dimensional steady flow field within the rotor passage was simulated with the datum tip clearance of 0.356 mm at the design wheel speed of 17188.7 rpm. The simulation results are well consistent with the measurement results, which verified the numeri- cal method. Then the three-dimensional steady flow field within the rotor passage was simulated respectively with different axial non-uniform tip clearances. The calculation results showed that optimal axial non-uniform tip clearances could improve the compressor performance, while the efficiency and the pressure ratio of the com- pressor were increased. The flow mechanism is that the axial non-uniform tip clearance can weaken the tip leak- age vortex, blow down low-energy fluids in boundary layers and reduce both flow blockage and tip loss.     

Flowfield and heat transfer past an unshrouded gas turbine blade tip with different shapes

This paper describes the numerical investigations of flow and heat transfer in an unshrouded turbine rotor blade of a heavy duty gas turbine with four tip configurations. By comparing the calculated contours of heat transfer coefficients on the flat tip of the HP turbine rotor blade in the GE-E³ aircraft engine with the corresponding experimental data, the κ-ω turbulence model was chosen for the present numerical simulations. The inlet and outlet boundary conditions for the turbine rotor blade are specified as the real gas turbine, which were obtained from the 3D full stage simulations. The rotor blade and the hub endwall are rotary and the casing is stationary. The influences of tip configurations on the tip leakage flow and blade tip heat transfer were discussed. It’s showed that the different tip configurations changed the leakage flow patterns and the pressure distributions on the suction surface near the blade tip. Compared with the flat tip, the total pressure loss caused by the leakage flow was decreased for the full squealer tip and pressure side squealer tip, while increased for the suction side squealer tip. The suction side squealer tip results in the lowest averaged heat transfer coefficient on the blade tip compared to the other tip configurations.     

Numerical simulation of tip clearance flow passive control in axial turbine

This paper focuses on the effects of five different passive turbine tip clearance flow control methods on the tip clearance flow physics, which consists of a partial suction side squealer tip, a double squealer tip, a pressure side tip shelf with inclined squealer tip on a double squealer tip, a tip platform extension edge in pressure side and in suction side respectively. A pressure-correction based, 3D Reynolds-averaged Navier-Stokes equations CFD code with Reynolds Stress Model was adopted. The variable specific heat was considered. The detailed tip clearance flow field with different squealer rims was described with the streamline and the velocity vector. Accordingly, the mechanisms of five passive controls were elucidated; the effects of the passive controls on turbine efficiency and tip clearance flow field were illuminated. The results showed that the secondary flow loss near the outer casing including the tip leakage losses and the passage vortex losses could be reduced in all the five passive control methods. The turbine efficiency could be increased via the rational passive turbine tip clearance flow control. The Improved PS Squealer had the best effect on turbine efficiency, and the efficiency increased by 0.215%.     

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.     

Effects of Width Variation of Pressure-Side Winglet on Tip Flow Structure in a Transonic Rotor

Tip leakage flow has become one of the major triggers for rotating stall in tip region of high loading transonic compressor rotors.Comparing with active flow control method,it’s wise to use blade tip modification to enlarge the stable operating range of rotor.Therefore,three pressure-side winglets with the maximum width of 2.0,2.5 and 3.0 times of the baseline rotor,are designed and surrounded the blade tip of NASA rotor 37,and the three new rotors are named as RPW1,RPW2,and RPW3 respectively.The numerical results show that the width of pressure-side winglet has significant influence on the stall margin and the minimum throttling massflow of rotor,while it produces less effect on the choking massflow and the peak efficiency of new rotors.As the width of the pressure-side winglet increases from new rotor RPW1 to RPW3,the strength of leakage massflow has been attenuated dramatically and a reduction of 20%in leakage massflow rate has appeared in the new rotor RPW3.By contrast,the extended blade tip caused by winglet has not introduced much more aerodynamic losses in tip region of rotor,and the new rotors with different width of pressure-side winglet have the similar peak efficiency to the baseline.The new shape of the leakage channel over blade tip which replaces of the static pressure difference near blade tip has dominated the behavior of the leakage flow in tip gap.As both the new aerodynamic boundary and throat in tip gap have reshaped by the low-velocity flow near the solid wall of extended blade tip,the discharging velocity and massflow rate of leakage flow have been suppressed obviously in new rotors.In addition,the increasing inlet axial velocity at the entrance of new rotor has increased slightly as well,which is attributed to the less blockage in the tip region of new rotor.In consideration of the increased inlet axial velocity and the weakened leakage flow,the new rotor presents an appropriately linear increase of the stall margin when the width of pressure-side winglet increases,and has a nearly 15%increase in new rotor RPW3.     

来流边界层对跨声速压气机转子流场影响的研究

为了研究来流边界层对跨声速压气机转子气动性能及流场的影响,针对Rotor37进行了不同来流边界层进口条件下的跨声速压气机流场数值模拟。结果表明:来流边界层引起其内部的激波结构变化,进而影响60%叶高以上流场,造成该展向范围内的流量分布发生再分配;在来流边界层具有相同的厚度时,总压亏损越大,以60%～90%叶高激波损失为主体的附加损失越高;来流边界层弱化了叶尖泄漏涡系的强度,通过同时改变叶尖负荷和叶尖泄漏流来源流体能量影响泄漏强度,进而影响泄漏涡系的形成和发展。     

Numerical investigation of the unsteady tip leakage flow and rotating stall inception in a transonic compressor

<正>It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception.Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor.For steady analysis,the predicted results agree well with the experimental data for the estimation of compressor rotor global performance.For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage.On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale-spike type rotating stall inception at blade tip region.It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased,the rotating stall cell was further developed in the blade passage.     

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.     

A similitude method and the corresponding blade design of a low-speed large-scale axial compressor rotor

A similitude method to model the tip clearance flow in a high-speed compressor with a low-speed model is presented in this paper. The first step of this method is the derivation of similarity criteria for tip clearance flow, on the basis of an inviscid model of tip clearance flow. The aerodynamic parameters needed for the model design are then obtained from a numerical simulation of the target high-speed compressor rotor. According to the aerodynamic and geometric parameters of the target compressor rotor, a large-scale low-speed rotor blade is designed with an inverse blade design program. In order to validate the similitude method, the features of tip clearance flow in the low-speed model compressor are compared with the ones in the high-speed compressor at both design and small flow rate points. It is found that not only the trajectory of the tip leakage vortex but also the interface between the tip leakage flow and the incoming main flow in the high-speed compressor match well with that of its low speed model. These results validate the effectiveness of the similitude method for the tip clearance flow proposed in this paper.     

机匣喷气量对涡轮间隙流动控制的影响

采用数值模拟方法研究机匣喷气量大小对涡轮间隙流动控制的影响。结果显示,在10％轴向弦长位位置喷气时,增大喷气量,喷气在间隙内轴向上影响范围增大,对间隙流阻塞作用增加,间隙涡出现位置推迟。同时减小了间隙涡、上通道涡区熵增,尤其是上通道涡区损失大幅减小,并减弱机匣喷气引起的气流偏转不足／过偏现象。叶顶压力面附近由间隙流动引起的低压区减小,并向叶片尾缘移动。但由于喷气量增大使得动叶输出功率下降,使得涡轮效率降低。     

Experimental and numerical investigation on compressor cascade flows with tip clearance at a low Reynolds number condition

High flow rate aeroengines typically employ axial flow compressors,where aerodynamic loss is predominantly due to secondary flow features such as tip leakage and corner vortices.In very high altitude missions,turbomachinery operates at low density ambient atmosphere,and the recent trend toward more compact engine core inevitably leads to the reduction of blade size,which in turn increases the relative height of the blade tip clearance.Low Reynolds number flowfield as a result of these two factors amplifies the relative importance of secondary flow effects.This paper focuses on the behavior of tip leakage flow,investigating by use of both experimental and numerical approaches.In order to understand the complex secondary flow behavior,cascade tests are usually conducted using intrusive probes to determine the loss.However relatively few experimental studies are published on tip leakage flows which take into account the interaction between a rotating blade row and its casing wall.Hence a new linear cascade facility has been designed with a moving belt casing in order to reproduce more realistic flowfield as encountered by a rotating compressor row.Numerical simulations were also performed to aid in the understanding of the complex flow features.The experimental results indicate a significant difference in the flowfield when the moving belt casing is present.The numerical simulations reveal that the leakage vortex is pulled by the shearing motion of the endwall toward the pressure side of the adjacent blade.The results highlight the importance of casing wall relative motion in analyzing leakage flow effects.     

Research Progress of Tip Winglet Technology in Compressor

In the present study,the research progress of tip winglets that control tip clearance leakage flow in compressors is reviewed.Firstly,the effects of tip leakage flow on the aerodynamic performance of the compressor are presented.Subsequently,the development of tip winglet technology is reviewed.Next,a series of studies on compressor tip winglet technology are conducted.Besides,the effects of tip winglets on the aerodynamic performance of rectangular cascades of low-speed and high-subsonic compressors,subsonic compressor rotor and transonic compressor rotor are discussed,respectively,and the control effect of tip winglet technology combined with tip groove design on tip leakage is investigated.Lastly,the subsequent development direction and research prospect of compressor tip winglet technology are presented.     

Influence of tip clearance on performance of a contra-rotating fan

Influences of tip clearance on the tip flow and associated loss mechanism in a contra-rotating axial flow fan has been studied in the paper, based on three dimensional numerical results. The results with different tip clearance are compared in terms of stage efficiency, relative total pressure loss coefficient, flow angle. It is found that the efficiency of the contra-rotating fan changes almost linearly with increment of the tip clearance, however, efficiency of the rear rotor is observed to decrease more dramatically than that of the forward rotor given same tip clearance variation. The analysis on the flow structure indicates that the tip region flow field is qualitatively similar in both rotors. However, with the same clearance value, the leakage flow in the rear rotor is effected by a tip leakage vortex of greater intensity caused by relative loading levels and the inter rotor interaction.     

间隙变化对轴流压气机转子近失速工况叶顶流场的影响研究

为研究间隙变化对轴流压气机转子近失速工况下叶顶流场结构的影响,以轴流压气机转子Rotor37为研究对象,对其叶顶流场进行定常和非定常的数值模拟。计算结果表明：随着叶顶间隙的减小,压气机的总压比和等熵效率均有所提高,稳定运行范围扩大;2倍设计间隙下,叶尖泄漏涡经激波作用后发生膨胀破碎,堵塞来流通道,诱发压气机堵塞失速;0.5倍设计间隙下,吸力面流动分离加剧,发生回流,部分回流与来流在压力面前缘上游发生干涉,进口堵塞加剧,致使部分来流从前缘溢出,导致压气机叶尖失速;不同间隙下压气机失速过程的主导因素不同,大间隙下失速由叶尖泄漏涡破碎的非定常波动引起,小间隙下失速主要由流动分离引发的周期性前缘溢流所主导。     

多级轴流压气机模化设计中尺寸缩放对气动性能影响的数值研究

为了研究几何尺寸模化缩放及叶尖间隙对多级轴流压气机气动性能及内部流动的影响,采用Numeca程序对17级轴流压气机开展了数值计算。结果表明：在80%及100%等高转速条件下压气机效率随着模化比例增大而增大,而在50%转速下模化缩放对压气机效率的影响较小。相对于原型压气机,模化放大时,压气机前8级单级压比均有所降低,而后8级压比均提高;模化缩小时,压气机的变化规律则相反。随着压气机几何尺寸的增大,静叶叶根和叶尖区域的总压恢复系数显著提高。同时,叶片叶尖泄漏流区域的熵增减少,从而使各级效率均有所提升。缩放模化中,随着叶尖间隙的增大,泄漏流增多,恶化了动叶叶尖附近的流动分离,降低了动叶后50%弦长区域的相对马赫数,同时扩大了静叶上端壁的流动分离,使压气机效率降低。     

Study on heat transfer of leakage flow in tooth tip clearance for single screw compressor based on fuzzy contourlet finite element method

Heat transfer of leakage flow in tip clearance will reduce the working performance, and therefore it is necessary to study it in depth. To improve the computing effectiveness of heat transfer of leakage flow in the tooth tip clearance of the single-screw compressor, the fuzzy contourlet finite element model is constructed by combining the contourlet finite element method and fuzzy finite element method. First, the related research progresses are summarized. Second, heat transfer model of leakage flow in the tip clearance of single-screw compressor is constructed. Third, the fuzzy contourlet wavelet finite element model is established. Finally, heat transfer simulation of leakage flow in the tip clearance of single-screw compressor is performed, and the variables with fuzziness are chosen, which are transformed to random variables based on information entropy theory. The comparing analysis among simulation analysis and test results is performed, and results show that the contourlet finite element method has highest computing precision and efficiency. In addition, the temperature of leakage flow in the tip clearance (L1) is also obtained, and results show that the fuzzy contourlet wavelet finite element method can effectively obtain the temperature distribution rules with fuzziness.     

带小翼肋条的涡轮叶尖泄漏流场的数值模拟

对叶尖吸力面带小翼肋条的某一轴流转子叶尖间隙泄漏流场进行了数值研究,分析了在不同肋条宽度下泄漏流场细节,并对涡轮效率进行了计算.结果表明:涡轮叶尖单吸力边小翼肋条总体上减小叶尖表面压差,使得吸力面后半部分泄漏流速度减小,从而减小泄漏流动损失,但会增大通道内流动损失,使涡轮转子效率下降;小翼肋条宽度有一个最佳值,小间隙下增大肋条宽度使得涡轮转子效率降低,大间隙下增大肋条宽度却使得涡轮转子效率提高;吸力边小翼肋条改变了叶尖吸力边附近的流场,对压力边附近泄漏流动结构影响不大.