叶顶微射流对小型高负荷离心压气机性能及流场结构的影响

以某小型高速离心压气机为研究对象,采用数值方法研究了微射流对压气机性能和叶轮叶顶流场结构的影响。研究结果表明：射流为1%设计流量时,失速裕度能够提高3.12%,稳定工作范围拓宽28.17%;在设计点,原型离心压气机叶顶来流马赫数达1.8以上,叶顶存在复杂激波/间隙泄漏流干扰,工作稳定性较差,微射流改变了“λ”状的激波结构,使前缘激波的强度减弱,后掠角度减小,并且降低了叶顶的负荷水平;微射流能够抑制间隙泄漏流的周向运动,并削弱激波/间隙泄漏流之间的相互作用,间隙泄漏涡不易发生破裂、溃散,极大增强了压气机工作的稳定性。     

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。     

端壁运动对自发射流抑制叶尖泄漏影响的数值研究

通过数值手段分析了端壁运动对带有叶尖自发射流的叶栅流场特性、叶顶间隙泄漏流量特性及叶片周向载荷分布特性的影响。结果表明:端壁运动的反向刮削作用会显著影响间隙附近流场,随着端壁运动速度增大,泄漏涡和上通道涡会逐渐向叶片吸力面靠近,同时,泄漏涡强度逐渐削弱,而上通道涡强度逐渐增强,平均总压损失增大;端壁运动对叶尖自发射流抑制泄漏流的效果有放大作用;端壁运动会显著改变叶尖吸力面附近静压系数的分布,使其具有后加载的特性。     

轮毂泄漏流对跨声速压气机转子性能影响的数值研究

为揭示转子前缘轮毂间隙泄漏流对高负荷压气机气动性能影响的物理机制，采用轮毂间隙边界条件模化处理方法，开展了轮毂泄漏流对跨声速压气机转子性能影响的三维定常数值模拟，分析了不同轮毂泄漏流量下压气机轮毂壁面流场结构与流态变化特征。研究结果表明：轮毂泄漏流会恶化压气机流通能力，影响程度随着泄漏量增加而逐渐增大。在近峰值效率工况下，当泄漏流量达到0.50%时，压气机流量约减小0.74%。当轮毂泄漏流达到一定强度后，反而呈现出部分正面效果，使得压气机压比或效率得到一定程度改善。轮毂泄漏流通过影响轮毂壁面流场结构空间分布来对压气机气动性能施加影响，尤其是鞍点的位置决定着轮毂间隙下游回流区和顺流区的影响范围以及轮毂壁面横向潜流强度。     

超跨声速风扇中叶尖间隙影响的两种机制

以带进口导叶的某跨声速高负荷高速单极风扇为研究对象,以转子尖部叶栅前缘半径量化叶尖间隙,相对叶尖间隙范围为0～4。采用NUMECA软件,基于S-A湍流模型进行数值模拟。计算结果表明：相对叶尖间隙为0.1时风扇综合性能最佳;尖部叶栅通道内激波后有附面层分离的低速流与泄漏流两类流动;当相对叶尖间隙在0～4范围内变动时,附面层分离逐渐被抑制,泄漏流逐渐增强;当相对叶尖间隙小于1时附面层分离的低速流占主导;当相对叶尖间隙大于1时泄漏流占主导;当相对叶尖间隙等于1时激波后的流动从附面层分离低速流过渡到泄漏流状态。     

封严篦齿对轴流压气机涡系结构与流动损失的影响

为了解决压气机级间泄漏与二次流流动问题,航空发动机轴流压气机静叶根部与转子之间通常采用篦齿进行封严。为研究封严篦齿泄漏流对压气机性能的影响,基于某轴流压气机建立了带封严篦齿真实结构的几何模型,采用三维数值模拟的方法,研究了篦齿泄漏流对某轴流压气机主流涡系结构和流动损失的影响,并探究了其影响机理。结果表明：封严篦齿泄漏流使压气机的压比和效率都有不同程度的下降;篦齿泄漏会增强上游转子叶根吸力面的尾缘角区涡和静子叶根吸力面的马蹄涡,并使设计工况的上游转子和静子的流动损失分别增大3.1%和13.1%;静子叶根后附面层低能流体被抽吸,改善了下游流场,使下游转子流动损失减小2.4%;在近喘振点,由于压气机内流场恶化严重,篦齿泄漏带来的流场变化并不显著,泄漏流对主流影响小。     

端壁相对运动对压气机叶栅间隙流场影响的数值模拟

压气机端壁与叶片间的相对运动是影响叶顶间隙气流流动的重要因素.采用数值模拟的方法考察了端壁运动对不同叶顶间隙压气机叶栅内三维流场的影响.结果表明:端壁相对运动改变了叶栅间隙流场结构,叶栅通道内出现向相邻叶片压力面运动的刮削泄漏涡,上通道涡及叶顶分离涡受到抑制,叶尖负荷增大,间隙泄漏流量增加,叶栅总损失由于叶顶区掺混损失减少而减少.     

带有自发射流的涡轮叶顶间隙流场PIV测量

为了探究叶尖射流对涡轮叶栅流场特性的影响,搭建了一个小尺度低速叶栅风洞实验台,利用粒子成像测速(PIV)技术对带有自发射流的涡轮叶顶间隙流场进行了直接测量,获得了低雷诺数(Re=6.46×103~3.23×104)下射流孔附近的流动图像及速度测量结果,展示了叶顶间隙内层流和紊流2种流态下自发射流与泄漏流的相互作用过程,揭示了低雷诺数工况下(涵盖层流到紊流的转捩)叶尖射流抑制泄漏流的作用机理及影响因素,并对叶尖射流尾迹中出现的类卡门涡街的涡分布现象进行了探讨.结果表明:叶尖射流的引入在泄漏流抑制方面取得一定收益,但同时也进一步加剧了叶顶间隙流动的复杂性.     

汽轮机叶顶间隙内非定常流动的数值分析

为了分析叶顶间隙泄漏涡的影响范围、运行轨迹和强度的变化规律,以某汽轮机高压级为研究对象,采用SSTκ-ω湍流模型,应用PISO算法对叶项间隙内的非定常流动进行了数值模拟．结果表明：叶顶间隙泄漏流是有规律的周期性的非定常流动,泄漏涡的影响范围、运行轨迹和强度随时间和叶顶间隙的变化而变化;泄漏流对主流的影响呈现出从弱到强、再从强到弱的周期性变化规律;叶顶间隙泄漏涡在丁／4时刻的强度和影响范围均达到最大,在T／2时刻,静叶脱落涡和动叶吸力面前部的泄漏涡混合形成新的涡系,而动叶吸力面后部的泄漏涡却与其边界层的脱涡混合,离开吸力面．     

带叶尖自发射流的平面叶栅流场特性数值研究

通过数值求解三维定常黏性雷诺时均N-S方程,获得了叶尖单孔自发射流条件下的叶栅流场,分析了自发射流与泄漏流的相互作用,比较了有、无自发射流条件下叶尖泄漏及载荷分布,探讨了端壁相对运动速度对自发射流抑制泄漏流有效性的影响.结果表明:叶尖射流在其下游形成的扇形低速区占据了部分泄漏流通道,由此对泄漏流产生一定的抑制作用,同时对叶尖吸力面载荷分布产生影响;与无叶尖射流相比,当进口雷诺数同为5.764×105时,叶尖自发射流的存在使泄漏比相对值降低5.42%,单个叶片载荷增加1.41%.     

Numerical study on the influence mechanism of inlet distortion on the stall margin in a transonic axial rotor

A numerical study is conducted to investigate the influence of inlet flow condition on tip leakage flow (TLF) and stall margin in a transonic axial rotor.A commercial software package FLUENT,is used in the simulation.The rotor investigated in this paper is ND_TAC rotor,which is the rotor of one-stage transonic compressor in the University of Notre Dame.Three varied inlet flow conditions are simulated.The inlet boundary condition with hub distortion provides higher axial velocity for the incoming flow near tip region than that for the clean inflow,while the incoming main flow possesses lower axial velocity near the tip region at tip distortion inlet boundary condition.Among the total pressure ratio curves for the three inlet flow conditions,it is found that the hub dis-torted inlet boundary condition improves the stall margin,while the tip distorted inlet boundary condition dete-riorates compressor stability.The axial location of interface between tip leakage flow (TLF) and incoming main flow (MF) in the tip gap and the axial momentum ratio of TLF to MF are further examined.It is demonstrated that the axial momentum balance is the mechanism for interface movement.The hub distorted inflow could de-crease the axial momentum ratio,suppress the movement of the interface between TLF and MF towards blade leading edge plane and thus enhance compressor stability.     

Experimental Investigation of Aerodynamic Performance due to Blade Tip Clearance in a Gas Turbine Rotor Cascade

This study examines how the complex flow structure within a gas turbine rotor affects aerodynamic loss. An unshrouded linear turbine cascade was built, and velocity and pressure fields were measured using a 5-hole probe. In order to elucidate the effect of tip clearance, the overall aerodynamic loss was evaluated by varying the tip clearance and examining the total pressure field for each case. The tip clearance was varied from 0% to 4.2% of blade span and the chord length based Reynolds number was fixed at 2×10⁵. For the case without tip clearance, a wake downstream of the blade trailing edge is observed, along with hub and tip passage vortices. These flow structures result in profile loss at the center of the blade span, and passage vortex related losses towards the hub and tip. As the tip clearance increases, a tip leakage vortex is formed, and it becomes stronger and eventually alters the tip passage vortex. Because of the interference of the secondary tip leakage flow with the main flow, the streamwise velocity decreases while the total pressure loss increases significantly by tenfold in the last 30% blade span region towards the tip for the 4.2% tip clearance case. It was additionally observed that the overall aerodynamic loss increases linearly with tip clearance.     

跨音轴流压气机动叶的三维弯掠设计研究

对一单级跨音轴流压气机中的动叶分别进行了前掠和正弯设计的参数研究,并根据研究得到的弯、掠动叶气动性能变化规律对动叶进行了前掠和正弯联合的三维设计,同时对动叶中部截面的叶型进行了二维设计以弥补弯掠动叶中部性能的降低.最终设计的跨音级性能显著提高,级最大效率提高3%,失速裕度提高40%,同时压比有所增加.数值计算结果表明,前掠和正弯叶片都可以使叶顶激波位置移向下游,降低激波强度,减轻叶顶激波与边界层和泄漏涡的作用.弯掠动叶控制激波强度和端壁流动的能力更加突出.     

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.     

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.     

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.     

Entropy Generation Analysis in a Mixed-Flow Compressor with Casing Treatment

Casing treatments(CT) can effectively extend compressors flow ranges with the expense of efficiency penalty. Compressor efficiency is closely linked to loss. Only revealing the mechanisms of loss generation can design a CT with high aerodynamic performance. In the paper, a highly-loaded mixed-flow compressor with tip clearance of 0.4 mm was numerically studied at a rotational speed of 30,000 r/min to reveal the effects of axial slot casing treatment(ASCT) on the loss mechanisms in the compressor. The results showed that both isentropic efficiency and stall margin were improved significantly by the ASCT. The local entropy generation method was used to analyze the loss mechanisms and to quantify the loss distributions in the blade passage. Based on the axial distributions of entropy generation rate, for both the cases with and without ASCT, the peak entropy generation rate increased in the rotor domain and decreased in the stator domain during throttling the compressor. The peak entropy generation in rotor was mainly caused by the tip leakage flow and flow separations near the rotor leading edge for the mixed-flow compressor no matter which casing was applied. The radial distributions of entropy generation rate showed that the reduction of loss in the rotor domain from 0.4 span to the rotor casing was the major reason for the efficiency improved by ASCT. The addition of ASCT exerted two opposite effects on the losses generated in the compressor. On the one hand, the intensity of tip leakage flow was weakened by the suction effect of slots, which alleviated the mixing effect between the tip leakage flow and main flow, and thus reduced the flow losses; On the other hand, the extra losses upstream the rotor leading edge were produced due to the shear effect and to the heat transfer. The aforementioned shear effect was caused by the different velocity magnitudes and directions, and the heat transfer was caused by temperature gradient between the injected flow and the incoming flow. For case with smooth casing(SC), 61.61% of the overall loss arose from tip leakage flow and casing boundary layer. When the ASCT was applied, that decreased to 55.34%. The loss generated by tip leakage flow and casing boundary layer decreased 20.54% relatively by ASCT.     

Effects of probe support on the flow field of a low-speed axial compressor

This paper presents an investigation on the effect of probe support on the flow field of an axial compressor.The experiment is carried out in a large-scale low-speed research compressor.A cylindrical probe support intruding to 50% blade span was installed at 50% chord upstream from the rotor leading edge.The region from 5° to 32° off the probe support in the direction of rotation at the rotor outlet was measured with a 5-hole probe and a high-response total pressure probe.The experiment is performed at both near-design and near-stall points.The measuring results of 5-hole probe and high-response total pressure probe indicate that the probe blockage effect is different at different blade spans.The wake of the probe support weakens the leakage vortex intensity at the tip region,leading to greater total pressure rise.At near-design condition,the presence of probe support has a negative effect on the region from 75% to 92% span,while improves the flow field below 75% span.At near stall condition,the probe support has a negative effect on the region from 70% to 90% span,and almost has no influence on the flow field below 70% span.