多级轴流压气机内复杂流动结构的实验和数值研究

发布时间：2018-08-28 07:18

【摘要】：端壁区内复杂的粘性流动严重降低多级轴流压气机的气动性能,现有研究表明,包括叶尖间隙流以及二次流在内的端壁损失通常占压气机内流动总损失的50%~70%,因此深入认识多级轴流压气机内特别是端壁区的流场结构及其主要物理机制对提升高压压气机乃至航空发动机的气动性能有着重要的现实意义。鉴于高压压气机后面级尺寸小、转速高,开展详细流场测量技术难度大、成本高、风险大且周期长,低速模拟实验得到了一定的应用,但迄今为止国内在该领域开展的实验研究仍相当匮乏,极大地限制了我国航空发动机高压压气机设计能力的提升。本文以丰富全面的实验测量为主,定常/非定常CFD计算为辅,针对两台用于某验证机高压压气机低速模拟实验的四级压气机以及某叶尖临界型转子A和非叶尖临界型转子B,开展了压气机气动性能、内部流动结构的研究,获得了压气机内部的主要流动结构,包括:类重复级的流动特征,转子叶顶泄漏流及其与主流的交界面随流量的减小而逐渐与前缘齐平的流动规律,静子通道内叶片附面层沿程发展及在两端吸力面角区形成的一对严重影响其性能的角涡,具有较大弓形的静子随流量的减小在吸力面叶中位置首先出现回流等,揭示了多级轴流压气机内复杂定常/非定常流动相互作用的物理机制及高流动损失的来源,并基于对原型压气机内部复杂流动结构的认识,采用先进叶片设计技术改善了原型压气机轮毂区的流动,提高了压气机的气动性能,为我国当代高性能航空发动机高压压气机的设计提供了必要的支持。本文主要包括以下三部分研究工作:第1部分基于“模拟准则”参数,开展了低速原型4级类重复级轴流压气机气动设计及其内部详细流动结构的实验研究。气动设计包括模拟目标的建立、低速模拟级总体参数确定、S2气动设计、重复级及类重复级叶型设计,将类重复级设计以及进口堵塞纳入了低速模拟压气机的设计中,完善了低速模拟设计体系,有效提高了低速压气机设计的可靠性和低速模拟的精度。分析了实验中各主要气动参数的不确定度及影响不确定度的主要因素,流量系数和总总压升系数的不确定度分别为0.42%和0.4%。通过通道间、转子和静子叶片通道内、静子表面静压分布及转子叶顶区流场测量完整地呈现了整个模拟级内部的详细流场结构,包括类重复级流动特征,叶顶泄漏流及其与主流交界面随流动工况变化的规律,静子叶片通道内形成的端壁角区对涡等。研究表明,原型压气机第3级转子轮毂区叶片负荷偏高且存在着一定的流动分离,同时静子轮毂也出现了较大的流动堵塞、分离及更高的总压损失。第2部分基于对原型压气机内部复杂流动结构的认识,开展了原型压气机第3级叶片改进设计及改进设计压气机性能及详细内部流动结构的实验研究。针对模拟级转、静子开展了积叠规律的参数化研究,明确了端部弯角的选择以减小轮毂区的流动损失,改进方案包含如下要素:转子“J”型积叠,增大的静子几何进口角,静子“前加载”技术以及更大的弓形积叠。实验结果表明,四级压气机效率提升了约1个百分点,总总压升系数提高了约1.4%,第3级总总压升系数提高了约10%,失稳流量和原型基本一致。改进设计实验的流场结构反映出了新的三维叶片设计特征,结合CFD计算结果研究了新的叶片造型改善压气机内流动结构及气动性能的主要物理机制。对级间的非定常测量结果进行频谱分析、系综平均等处理后发现,上游转子尾迹脱落引起静子尾迹区两侧叶片通过频率(fBPF)具有更高的能量及总压均方根值,静子附面层与更前面级静子附面层的相互作用使高阶谐波能量增强,且这种现象在靠近叶尖及压气机出口时更明显,在第3级出口叶尖区域2fBPF的能量幅值达到fBPF的8倍,该现象对于压气机内气动噪音和振动的控制均具有重要的意义。第3部分针对某叶尖失速型转子A和某非叶尖失速型转子B叶顶区域开展了详细的数值研究。研究揭示了叶尖泄漏涡的主要流动结构,对于非叶尖临界型转子B,在间隙高度62.5%以下的叶尖区域内,从叶尖间隙前缘流出的流体会卷吸成叶尖泄漏涡,且随着间隙高度的增加,其占据的叶尖泄漏涡的位置由内而外,而62.5%间隙高度以上的流动则并不卷吸入叶尖泄漏涡内,而主要表现为二次泄漏。在近失速工况下,叶尖临界型转子A叶顶会出现自激非定常性,这种现象的出现与否强烈依赖于叶尖间隙尺寸,而与上游附面层厚度关系不大;在叶顶区形成两个堵塞区,其中叶尖二次涡和破碎的叶尖泄漏涡的相互作用形成的大堵塞区会产生大的流动损失,将直接决定转子的气动稳定性。通过有效的流动控制方法消除或至少减弱该二次涡的强度将有助于拓宽该转子的稳定工作范围。
[Abstract]:Complex viscous flow in the end-wall region seriously degrades the aerodynamic performance of a multistage axial compressor. Existing studies have shown that the loss of the end-wall, including tip clearance flow and secondary flow, usually accounts for 50%-70% of the total loss of flow in the compressor. Therefore, a thorough understanding of the flow field structure in the multistage axial compressor, especially in the end-wall region, as well as its main physical properties is given. The mechanism has important practical significance for improving the aerodynamic performance of high-pressure compressors and even aero-engines. In view of the small size and high speed of the rear stage of high-pressure compressors, it is difficult to carry out detailed flow field measurement technology, high cost, high risk and long cycle, low-speed simulation experiments have been applied to a certain extent, but so far domestic research in this field has been carried out. Experimental studies are still scarce, which greatly limits the improvement of design capability of high-pressure compressors for aero-engines in China. In this paper, based on abundant and comprehensive experimental measurements, steady/unsteady CFD calculations are supplemented by two four-stage compressors used in low-speed simulation experiments of high-pressure compressors for a certain validator and a blade-tip critical rotor A and A. A non-tip critical rotor B was developed to study the aerodynamic performance and internal flow structure of a compressor. The main flow structures in the compressor were obtained, including the flow characteristics of the quasi-repetitive stage, the tip leakage flow and its interface with the mainstream gradually flattened with the leading edge as the flow rate decreased, and the blade boundary in the stator passage. A pair of angular vortices formed along the layer and in the corner region of the suction surface at both ends seriously affect the performance of the compressor. The stator with a larger bow first appears to be recirculated in the suction surface blade with the decrease of flow rate. The physical mechanism of complex steady/unsteady flow interaction and the source of high flow loss in a multistage axial compressor are revealed. Based on the understanding of the complex flow structure inside the prototype compressor, advanced blade design technology is adopted to improve the flow in the hub region of the prototype compressor and improve the aerodynamic performance of the compressor, which provides the necessary support for the design of the high-pressure compressor of the contemporary high-performance aeroengine in China. Based on the "simulation criteria" parameters, the aerodynamic design and detailed flow structure of a low-speed prototype 4-stage repetitive axial compressor were studied experimentally. The aerodynamic design includes the establishment of simulation objectives, the determination of the overall parameters of the low-speed simulation stage, the S2 aerodynamic design, the repetitive stage and the quasi-repetitive stage blade design, and the quasi-repetitive stage design. The inlet blockage was incorporated into the design of the low-speed simulated compressor, and the low-speed simulated design system was improved. The reliability and accuracy of the design of the low-speed compressor were effectively improved. The uncertainty of the main aerodynamic parameters and the main factors affecting the uncertainty were analyzed. The static pressure distribution on the stator surface and the measurement of the rotor tip flow field in the rotor and stator blade passages show the detailed flow field structure in the whole simulation stage, including the characteristics of the quasi-repetitive stage flow, the variation of the tip leakage flow and its interface with the main flow conditions. The results show that the blade load in the hub region of the third stage rotor of a prototype compressor is high and there is a certain flow separation. At the same time, the stator hub also has a large flow blockage, separation and higher total pressure loss. Part 2 is based on the recognition of the complex flow structure in the prototype compressor. In order to reduce the flow loss in the hub region, a parameterized study of stacking rule was carried out on the stator for the simulated stage rotation. The selection of the end bending angle was made clear to reduce the flow loss in the hub region. The improvement scheme includes the following elements: rotor "J" The experimental results show that the efficiency of the four-stage compressor is increased by about one percentage point, the total pressure rise coefficient is increased by about 1.4%, the total pressure rise coefficient of the third stage is increased by about 10%, and the unstable flow rate is basically the same as that of the prototype. The structure reflects the new three-dimensional blade design characteristics, and the main physical mechanism of improving the flow structure and aerodynamic performance of the compressor by using the new blade shape is studied with CFD calculation results. The blade passage frequency (fBPF) has higher energy and root mean square of total pressure. The interaction between stator boundary layer and stator boundary layer in the preceding stage enhances the high-order harmonic energy. This phenomenon is more obvious near the blade tip and compressor outlet. The energy amplitude of 2fBPF in the blade tip region of the third stage is eight times that of fBPF. Aerodynamic noise and vibration control in a compressor are of great importance. Part 3 is a detailed numerical study of the tip region of a blade-tip stall rotor A and a non-tip stall rotor B. The study reveals the main flow structure of tip leakage vortices. For non-tip critical rotor B, the blade with a clearance height of 62.5% or less has been investigated. In the tip region, the fluid flowing from the leading edge of the tip clearance will entrap into the tip leakage vortex. With the increase of the tip clearance height, the tip leakage vortex occupies the position from inside to outside, while the flow above 62.5% clearance height does not entrap into the tip leakage vortex, but mainly presents secondary leakage. Self-excited unsteadiness occurs at the tip of rotor A blade, which is strongly dependent on tip clearance size and has little to do with the thickness of upper boundary layer. Two blockage zones are formed at the tip of rotor A blade, in which the large blockage zone formed by the interaction of tip secondary vortex and broken tip leakage vortex will produce a large flow loss, which will be directly related to the thickness of upper boundary layer. The aerodynamic stability of the rotor is determined. Eliminating or at least weakening the strength of the secondary vortex by effective flow control method will help to widen the stable operating range of the rotor.
【学位授予单位】：南京航空航天大学
【学位级别】：博士
【学位授予年份】：2015
【分类号】：V233

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