轴流压气机非轴对称轮毂造型对间隙流控制研究

对带静子间隙的单级轴流压气机进行了全三维数值模拟,流场分析表明静子间隙泄漏流在通道端壁区引起较大的流动分离。为控制流动损失,对静子轮毂进行了非轴对称造型,造型后的压气机总性能得到改善。流场分析表明:非轴对称端壁造型改变了壁面静压分布,改善了间隙泄漏流在通道内的流动结构,消除了通道出口处的回流损失区,使压气机总压比增加,等熵效率提高0.9%。     

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。     

叶顶间隙对跨音速离心压气机气动性能影响分析

基于控制变量法对某跨音速离心压气机进行数值模拟,研究了叶轮尾缘叶间隙改变对其气动性能的影响。仅改变该离心叶轮的尾缘叶顶间隙,在设计转速下进行全三维黏性数值模拟,对相关气动参数进行分析。计算结果表明,相较于小流量工况,尾缘叶顶间隙的改变对离心压气机大流量工况的气动性能影响更大;在设计流量下,离心叶轮的压比、效率与叶轮尾缘出口间隙大小之间具有一定的线性关系,随着叶尖间隙增大,叶轮叶尖泄漏流的强度明显增强,导致叶轮的增压能力下降。     

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。     

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。     

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。     

动叶间隙对压气机气动性能的影响

为研究动叶间隙大小对压气机性能及流动的影响,以一台高亚声速一级半压气机级为研究对象,在设计间隙、0. 5倍、1. 5倍及2倍设计间隙下进行定常三维数值模拟。计算结果表明,随着间隙增大,压气机效率及总压比下降。大间隙下泄漏流增强,导致动叶叶尖及其下游区域损失增加,压气机转子效率随间隙增大而线性下降;同时,泄漏流的增强也恶化了动静叶匹配,导致静叶上端壁产生额外损失,压气机级效率下降幅度大于转子效率。     

叶型探针对跨声压气机性能影响的数值模拟

为揭示叶型探针对跨声压气机气动性能影响的物理机理,采用三维数值模拟方法,通过构建带多点实体探头的1.5级压气机计算模型,详细研究了设计转速下压气机在安装探头前后性能特性与内部流场的变化规律。结果表明:级间局部静叶安装探头后,压气机下堵点流量减小0.1%,最高效率降低0.4%,失速点流量增大0.15%;由压气机工作点改变所引起的静叶攻角变化是影响叶片表面两侧探头绕流尺度的重要因素,并且探头绕流影响程度与静叶气动负荷存在较高的关联度;探头大尺度绕流加剧下游转子叶背根部附面层径向迁移与分离是导致该型压气机气动失稳提前的主要原因。     

离心压气机轮缘压力及间隙泄漏非轴对称现象研究

蜗壳通常被设计成螺旋状结构,其几何结构的周向不对称诱发蜗壳内部流场周向分布不均匀现象,会对压气机内部流场造成显著影响.采用试验和数值结合的方法,对两种流量工况下离心压气机内部的非轴对称流动特性进行研究.结果表明:压气机轮缘静压在周向上的分布具有与蜗壳内部相同特征的非轴对称形式;下游流场畸变在向上游逆向传播过程中对上游的影响逐渐减弱;叶轮出口轮缘周向的高压区传播至叶轮进口时,与叶轮出口高压区的周向位置存在相位差.在不同流量工况下,轮缘周向静压分布不均导致叶顶间隙泄漏流在周向的分布存在差异.在小流量工况下,泄漏流量呈现先减小后增大再减小的波动分布;在大流量工况下,泄漏流量呈现先增大后持续减小分布.     

雷诺数对压气机性能影响的数值模拟

采用加入AGS转捩模型的Spalart—Allmaras湍流模型对低雷诺数条件下NASA37跨音速压气机转子内部流场进行数值模拟。分析了雷诺数对NASA37压气机转子内部流场特性和性能的影响,探讨了雷诺数对压气机气动性能的影响机理和影响规律。结果表明,随着雷诺数的降低,叶片表面吸力面附面层分离增大,压气机的效率和增压比都逐渐下降,稳定工作范围减小,特别是当雷诺数低于临界雷诺数时,压气机的工作性能和稳定性都发生明显下降。     

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。     

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.     

Performance improvements of a subsonic axial-flow compressor by means of a non-axisymmetric stator hub end-wall

This paper deals with the application of a non-axisyrmnetric hub end-wall on the stator of a single stage high subsonic axial-flow compressor. In order to obtain a state-of-the-art stator non-axisymmetric hub end-wall con- figuration fulfilling the requirements for higher efficiency and total pressure ratio, an automated multi-objective optimizer was used, in conjunction with 3D-RANS-flow simulations. For the purpose of quantifying the effect of the optimal stator non axis-symmetric hub contouring on the compressor performance and its effects on the sub- sonic axial-flow compressor stator end-wall flow field structure, the coupled flow of the compressor stage with the baseline, axisymmetric and the non-axisynunetric stator hub end-wall was simulated with a state-of-the- art multi-block flow 3D CFD solver. Based on the CFD simulations, the optimal compressor hub end-wall con- figuration is expected to increase the peak efficiency by approximately 2.04 points and a slight increase of the to- tal pressure ratio. Detailed analyses of the numerical flow visualization at the hub have uncovered the different hub flow topologies between the cases with axisymmetric and non-axisymmetric hub end-walls. It was found that that the primary performance enhancement afforded by the non-axisymmelric hub end-wall is a result of the end-wall flow structure modification. Compared to the smooth wall case, the non-axisymmetric hub end-wall can reduce the formation and development of in-passage secondary flow by aerodynamic loading redistribution.     

Influence of Hub Contouring on the Performance of a Transonic Axial Compressor Stage with Low Hub-Tip Ratio

The rotor blade height with low hub-tip ratio is relatively longer,and the aerodynamic parameters change drastically from hub to tip.Especially the organization of flow field at hub becomes more difficult.This paper takes a transonic 1.5-stage axial compressor with low hub-tip ratio as the research object.The influence of four types of rotor hub contouring on the performance of transonic rotor and stage is explored through numerical simulation.The three-dimensional numerical simulation results show that different hub contourings have obvious influence on the flow field of transonic compressor rotor and stage,thus affecting the compressor performance.The detailed comparison is conducted at the rotor peak efficiency point for each hub contouring.Compared with the linear hub contouring,the concave hub contouring can improve the flow capacity,improve the rotor working capacity,and increase the flow rate.The flow field near blade root and efficiency of transonic rotor is improved.The convex hub contouring will reduce the mass flow rate,pressure ratio and efficiency of the transonic rotor.Full consideration should be given to the influence of stator flow field by hub contouring.     

单级轴流压气机周向槽与梯状间隙结构的对比

在验证了计算模型的可靠性后,对单级轴流亚音压气机进行了71%设计转速下,实壁机匣,周向槽机匣与梯状间隙结构工作特性的数值模拟。结果表明,周向槽机匣与梯状间隙结构都能在一定范围内起到扩稳作用。周向槽的作用更多地体现在对叶片前缘叶顶泄漏涡的控制上,因此更适于设计在叶片通道的前半段。梯状间隙结构在叶片通道前半段与后半段都能起到扩稳作用,只是施于叶片前缘时,压气机效率损失较大,而施于叶片尾缘时,可以减小叶片尾迹分离区的面积,对提高压气机效率有益。     

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.     

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.     

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 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.