INVESTIGATING MULTISTAGE AXIAL FLOW COMPRESSOR INSTABILITIES

A numerical methodology for investigating compressor instabilities in a multistage environment is presented. The method is based on a stage-by-stage dynamic compression model and considers air compressibility explicitly throughout the compressor. It involves discretizing the compression system into distinct elements and a use of the local elemental characteristic of mean performance. The models are presented in both nonlinear and linearized forms. The linearised form permits well surge condition prediction for multistage axial compressors, while the non-linear form is able to investigate the growth of local flow disturbances, and helps to develop practical control strategy. Validations were carried out using the data from several aircraft engine compressors. A good experiment-model consistency is achieved.

Mechanism of Affecting the Performance and Stability of an Axial Flow Compressor with Inlet Distortion

The purpose of the investigation is to obtain the mechanism of affecting the performance and stability of an axial flow compressor with two kinds of multi distortion area.In the investigation,the compressor inlet distortion is made by the distortion generator,which is a column and is placed in the compressor rotor passage upstream.The five-passages unsteady numerical results show that the compressor performances for two kinds of multi distortion area are both lower than that for the clean inlet,and the compressor peak efficiencies for inlet distortions of 6×12 degrees and 6×24 degrees are about 3.5%,and 9.2%lower than that for the clean inlet,respectively.Further,two kinds of multi distortion area both reduce the compressor stability,and the stall margin improvements of–2.93%and–6.11%are provided by inlet distortions of 6×12 degrees and 6×24 degrees,respectively.The flow field analyses show that the flow conditions of the rotor tip inlet upstream deteriorate after the application of multi distortion area.The values of axial velocity and flow angle of incoming flows in some regions,which are near the shroud,become small due to the adverse effect made by the inlet distortion.So,the circulation capacity of the regions becomes low,and inlet blockages appear in some rotor tip passages.The bigger the distorted range is,the greater inlet blockages and corresponding flow losses are.When multi distortion area changes from 6×12 degrees to 6×24 degrees,there are obvious separations of the boundary layer near the blade suction surface in some blade tip passages,and the serious separations prevent the incoming flows from entering the blade tip passage.As a result,overflows occur at the leading edge of the blade tip.Furthermore,as the number of the distortion generator increases by a factor of one,the effects of inlet distortion on the amplitudes of the air relative velocity and static pressure in the rotor inlet upstream at the distortion generator passing frequency of 1076.5 Hz increase by a factor of about one.

Experimental and numerical study of tip injection in a subsonic axial flow compressor

Parametric study of tip injection was implemented experimentally on a subsonic axial flow compressor to understand the underlying flow mechanisms of stability improvement of the compressor with discrete tip injection.Injector throat height varied from 2 to 6 times the height of rotor tip clearance,and circumferential coverage percentage ranged from 8.3% to 25% of the annulus.Static pressure fluctuations over the rotor tip were measured with fast-response pressure transducers.Whole-passage time-accurate simulations were also carried out to help us understand the flow details.The combinations of tip injection with traditional casing treatments were experimentally studied to generate an engineering-acceptable method of compressor stall control.The results indicate that the maximum stability improvement is achieved when injectors are choked despite their different sizes.The effect of circumferential coverage percentage on compressor stability depends on the value of injector throat height for un-choked injectors,and vice versa.Tip blockage in the blade passage is greatly reduced by the choked injectors,which is the primary reason for stability enhancement.The accomplishment of blockage diminishment is maintained in the circumferential direction with the unsteady effect of tip injection,which manifests as a hysteresis between the recovery of tip blockage and the recovery of tip leakage vortex.The unsteady effect is primarily responsible for the effectiveness of tip injection with a partial circumferential coverage.Tip injection cannot enhance the stability of the rotor with axial slots significantly,but it can improve the stability of the rotor with circumferential grooves further.The combined structure of tip injection with circumferential grooves is an alternative for engineering application.

Numerical Investigation of Inlet Distortion on an Axial Flow Compressor Rotor with Circumferential Groove Casing Treatment

On the base of an assumed steady inlet circumferential total pressure distortion, three-dimensional time-dependent numerical simulations are conducted on an axial flow subsonic compressor rotor. The performances and flow fields of a compressor rotor, either casing treated or untreated, are investigated in detail either with or without inlet pressure distortion. Results show that the circumferential groove casing treatment can expand the operating range of the compressor rotor either with or without inlet pressure distortion at the expense of a drop in peak isentropic efficiency. The casing treatment is capable of weakening or even removing the tip leakage vortex effectively either with or without inlet distortion. In clean inlet circumstances, the enhancement and forward movement of tip leakage vortex cause the untreated compressor rotor to stall. By contrast, with circumferential groove casing, the serious flow separation on the suction surface leads to aerodynamic stalling eventually. In the presence of inlet pressure distortion, the blade loading changes from passage to passage as the distorted inflow sector is traversed. Similar to the clean inlet circumstances, with a smooth wall casing, the enhancement and forward movement of tip leakage vortex are still the main factors which lead to the compressor rotor stalling eventually. When the rotor works trader near stall conditions, the blockage resulting from the tip leakage vortex in all the passages is very serious. Especially in several passages, flow-spillage is observed. Compared to the clean inlet circumstances, circumferential groove casing treatment can also eliminate the low energy zone in the outer end wall region effectively.

Unsteady Flow Field under Surge and Rotating Stall in a Three-stage Axial Flow Compressor

The unsteady flow structure between rotor blade-to-blade passages in a three-stage axial flow compressor is experimentally investigated by detailed measurements of unsteady performance characteristics,casing wall pressure fluctuations and their wavelet analyses.The main feature of the test compressor is a capacity tank facility connected in series to the compressor outlet in order to supply compression and/or expansion waves from downstream of the compressor.Research attention is focused on the post-stall characteristics of the surge and rotating stall which occur simultaneously.The influence of the compressor operating point on the unsteady performance curve shows that the surge cycle changes irregularly depending on the steady-state resistance characteristics,and the results of the wavelet analyses of the wall pressure fluctuations suggest that the surge cycle may selectively be determined by the rotating stall cell structure within the rotor cascade.

Unsteady Behavior of Surge and Rotating Stall in an Axial Flow Compressor

Unsteady behaviors as well as unsteady cascade flow fields of a single-stage axial flow compressor were experimentally investigated by detail measurements of unsteady performance characteristics and casing wall pressure and internal flow velocity fluctuations. The main feature of the test compressor is a capacity tank connected directly to the compressor outlet in series through slits and a concentric duplex pipe, and also jet nozzles in order to inject compressed air toward the rotor tip region. Research attention is focused on the post-stall characteristics of surge and rotating stall which occur simultaneously. When the compressor was connected to the capacity tank, surge was generated with rotating stall in accordance with the capacitance increment of whole compressor system. The surge behavior changed irregularly with throttling valve installed behind the compressor, and several types of surge cycles were observed. In addition, the surge cycle changed by jet injection to the rotor tip region. The results suggested that the blockages of the cascade flow which were generated by a stall cell play an important role in deciding the surge behaviors.

Coexisting Phenomena of Surge and Rotating Stall in an Axial Flow Compressor

The unsteady inner flow structure of a single-stage axial flow compressor under the coexisting conditions of surge and rotating stall was experimentally investigated via detailed measurements of the unsteady characteristics and the internal flow velocity fluctuations. The main relevant feature of the tested compressor is a shock tube with a capacity tank connected in series to the compressor outlet through slits and a concentric duplex pipe: surge and rotating stall can both be generated by connecting the shock tube. Research attention is focused on the unsteady behavior of a rotating stall during the surge cycle. The size of the rotating stall cell during the recovery process of an irregular surge cycle was experimentally determined by the circumferential flow velocity fluctuations ahead of the rotor blade. The results suggested that the size of the rotating stall cell at the switching point of the performance curve between large and small cycles is considered to be the key parameter in determining the following surge cycle. In addition, the surge cycle is largely influenced by the unsteady behavior of the rotating stall cell.

Modeling of the Double Leakage and Leakage Spillage Flows in Axial Flow Compressors

A model to predict the double leakage and tip leakage leading edge spillage flows was developed.This model was combined by a TLV trajectory model and a TLV diameter model and formed as a function of compressor one-dimensional design parameters,i.e.the compressor massflow coefficient,Ф and compressor loading coefficient,ψ,and some critical blade geometrical parameters,i.e.blade solidity,σ,stagger angle,βS,blade chord length,C,and blade pitch length,S.By using this model,the double leakage and tip leakage leading edge spillage flow could be predicted even at the compressor preliminary design process.Considering the leading edge spillage flow usually indicates the inception of spike-type stall,i.e.the compressor is a tip critical design,this model could also be used as a tool to choose the critical design parameters for designers.At last,some experimental data from literature was used to validate the model and the results proved that the model was reliable.

Stall Inception Process and Prospects for Active Hub-Flap Control in Three-Stage Axial Flow Compressor

使用活跃中心拍动的可能性控制方法，它是为一台单个阶段的压缩机的一个证明旋转货摊控制方法，到 3 阶段轴的压缩机试验性地被讨论，在复杂旋转货摊开始进程被观察的地方。研究压缩机是 3 阶段并且能为转子片和定子翼改变 stagger 角度背景。十六个转子片 / 定子翼配置模式被变化测试蹒跚为定子翼的角度。由表面压力变化的测量，货摊开始进程被调查，测量压力变化数据为旋转货摊被用作一个预兆的信号。货摊察觉系统在一台单个阶段的压缩机用于活跃中心拍动控制的试验性的结果表演能有用地与一个更复杂的货摊开始过程在一台 3 阶段压缩机被用于那。

A Simple and Unified Model for Spanwise Mixing in Multistage Axial Flow Compressors

A basic equation system for meridional throughflow fileds in multistage axial flow compressors has been deduced,containing many unknown correlation terms,which describe different kinds of spanwise mixing mechanism in a unified form.The equation system shows that spanwise mixing of meridional flows in compressors is attributed to three kinds of mechanism:molecular motion,turbulent diffusion,and circumferential non-uniformities,the last of which includes secondary flow effects and others,Therefore the equation system unifles the two models for spanwise mixing coefficients defined and introduced into the basic equation system,a novel,much simpler equation system,without additional unknown correlation terms included,has been obtained.This novel equation system makes throughflow computations including mixing far easier for multistage compressors .It has been rigorously shown that these apparent mixing coefficients contain full information of all the three kinds of mixing mechanism,so that the simpler equation system can also be taken as a unified model for meridional flows with all the dinds of the mixing effects.Calculations of the flow through multistage machines have been made by incorporating the new model into a streamline curvature throughflow calculation method and the improved agreement with experimental data has been obtained.It is believed that the simpler equation system can be apphed to the flows not only in subsonic but in transonic and supersonic compressors if an appropriate model is propsed for the apparent mixing coeffcients.

Investigation of Stator Wake Variability in Single Stage Axial-Compressor

In this paper, we investigate the stator wake variability along the axial direction under different rotating speeds and mass flow conditions, in the terms of wake-core position, wake depth, wake width and wake velocity gradient. Experimental study is carried out in a single stage axial compressor by using Hot-wire Anemometer (HWA). Simultaneously, the numerical calculation is brought out in the same experiment rig. Firstly, we analyze the time-averaged velocity and radial vorticity at different axial positions, and explore the wake variability along the axial direction. Then, we present the effects of operation conditions on the wake-core position, and find out the inlet Mach number and incident angle are both the key influence factors of the wake-core position. Finally, we summarize four preliminary conclusions of the wake variability.

NEW SCALING METHOD FOR COMPRESSOR MAPS USING AVERAGE INFINITESIMAL STAGE

The estimation of the precise performance of existing multistage axial-flow compressors of gas turbine engines is fast becoming a great concern, as the use of gas turbine engines in the power generation and in the military industry increases, in order to reduce the analysis performance error of the traditional scaling method, a new scaling method for estimating the characteristics of multistage axial flow compressors is proposed. This novel method is based on experimental and partial data provided by engine manufacturers. Taking the effect of density-change into account, we introduce the average infinitesimal stage concept, and thereby divide the compression process into an infinite number of infinitesimal processes corresponding to infinitesimal stages. Subsequently, we adopt the corrected Reynolds analogy method for compressible flow calculation in order to ensure much better compliance with the similarity criterion, Validation checks show that the proposed method has enough precision to predict the off-design performance characteristics of multistage axial flow compressors.

基于压气机三维彻体力模型的叶尖射流扩稳方法

为了分析转子叶尖射流对多级轴流压气机的扩稳作用,研究了轴流式压气机三维黏性彻体力模型,通过对ANSYS CFX软件的二次开发,求解带动量源项和能量源项的N-S(Navier-Stokes)方程,实现无叶片条件下的压气机全通道数值模拟。基于模型计算的Stage 35单级压气机压比特性与实验值误差不大于0.8%,失稳点流量比实验值大1.2%,对端壁边界层流动能够准确模拟。基于模型计算了无射流和有射流条件下的多级压气机特性和失稳边界,并分析了叶尖射流扩稳机理。结果表明,模型能够模拟叶尖射流对多级压气机的扩稳作用,在100%设计转速下,设计点流量的4.6%的射流流量使稳定裕度增加了7.06%,射流能够消除叶尖区域的堵塞流动,从而扩大了压气机的稳定工作裕度。

多级高负荷轴流压气机喘振特征分析

针对某十级高负荷轴流压气机,在第一级和第九级机匣壁面上布置动态压力传感器,通过减小排气阀门开度的方式进行逼喘试验,获得了不同转速下多级高负荷轴流压气机喘振时的动态压力数据,采用时序、FFT结合滤波分析方法研究了多级高负荷轴流压气机在不同转速下的喘振特征。结果表明:在高转速下,该压气机喘振模式为深度喘振,喘振引起的高温气体倒流使得入口总温瞬时升高约80 K,喘振由出口级的模态波扰动导致流动堵塞引起,模态波频率约为16%~20%转频,喘振时入口级工作于压升特性增长阶段而出口级工作于压升特性顶点位置;在低转速下,该压气机喘振模式为经典喘振,未发生气体倒流,喘振主要由尖脉冲扰动引起,压气机前后级工作于特性线的增长阶段。高转速喘振恢复阶段,伴随着明显的旋转失速和喘振特征,而低转速下在喘振的第一个周期内以喘振特征为主,之后喘振频率的幅值波动迅速减小,旋转失速频率的幅值逐渐增加并占主导。高转速下喘振恢复时间约4.1 s,而低转速下喘振恢复的时间为0.56 s。

In uence of Endwall Boundary Layer Suction on the Flow Fields of a Critically Loaded Di usion Cascade

Boundary layer suction is an e ective method used to delay separations in axial compressors. Most studies on bound?ary layer suction have focused on improving the performance of compressors,whereas few studies investigated the influence on details of the flow fields,especially vortexes in compressors. CFD method is validated with experi?mental data firstly. Three single?slot and one double?slot endwall boundary layer suction schemes are designed and investigated. In addition to the investigation of aerodynamic performance of the cascades with and without suction,variations in corner open separation,passage vortex,and concentration shedding vortex,which are rarely seen for the flow controlled blades in published literatures,are analyzed. Then,flow models,which are the ultimate aim,of both baseline and aspirated cascades are established. Results show that single?slot endwall suction scheme adjacent to the suction surface can e ectively remove the corner open separation. With suction mass flow rate of 0.85%,the overall loss coe cient and endwall loss coe cient of the cascade are reduced by 25.2% and 48.6%,respectively. Besides,this scheme increases the static pressure rise coe cient of the cascade by 3.2% and the flow turning angle of up to 3.3° at 90% span. The concentration shedding vortex decreases,whereas the passage vortex increases. For single?slot suction schemes near the middle pitchwise of the passage,the concentration shedding vortex increases and the passage vortex is divided into two smaller passage vortexes,which converge into a single?passage vortex near the trailing edge section of the cascade. For the double?slot suction scheme,triple?passage vortexes are presented in the blade passage. Some new vortex structures are discovered,and the novel flow models of aspirated compressor cascade are proposed,which are important to improve the design of multi?stage aspirated compressors.

The Numerical Simulation of Gas Turbine Inlet-Volute Flow Field

The structural and aerodynamic performance of the air inlet volute has an important influence on the performance of the gas turbine. On one hand, it requires the airflow flowing through inlet volute as even as possible, in order to reduce the pressure loss, to avoid a decrease in the effective output power and an increase of the fuel consumption rate of the internal combustion engine which indicate the inefficiency of the entire power unit;On the other hand, it requires the size of the inlet volute to be as small as possible in order to save mounting space and production costs. The thesis builds the structure model and develops flow fields numerical simulation of several different sizes of the inlet volutes. Further, the unreasonable aerodynamic structure is improved according to the flow field characteristics and thereby, a better aerodynamic performance of the inlet volute is obtained.

Effects of circumferential configuration concerning splitter blade on the aerodynamic performance in a transonic axial fan

For a certain type of transonic axial fan, the flow field of a fan rotor with splitter blade was computed by numerical simulation, and the shape of the rotor was modified. The effects of different circumferential distributions concerning the splitter cascades upon the aerodynamic performance were investigated. The studies show that the optimum splitter cascade is not very close to the suction side of main blade. The load between the main blade and the splitter blade can be soundly distributed in terms of the adjustment of circumferential position of the splitter blade. The best aerodynamic performance can be successfully obtained according to the optimum shape of the expanding fluid channel reasonably formed by the splitter blade and the main blade.

轴流压缩机空腔碳纤维叶片防冲击设计

为适应我国航空航天器的发展要求,风洞得到快速发展,风洞规模和性能指标逐渐提高,以使气动试验更加接近真实状态。叶片是保证主压缩机性能的核心部件,碳纤维材质的叶片结构简单、质量轻,可满足大直径、大功率、高转速运行使用的需求。但是在使用过程中,由于各种原因叶片存在一定的碰撞风险,需要对轴流压缩机空腔碳纤维叶片在低温、潮湿环境下气流入口方向的凝结冰粒,或者在运转过程中异物随着气流撞击叶片造成的损伤进行研究,为改进叶片结构设计提供理论依据。

采用动态交互作用分析的叶轮机械优化算法研究

针对当前基于代理模型的优化算法因遭遇“维度灾难”而难以开展叶轮机械高维优化的难题,提出了自适应厘清变量间关系的动态聚合全局优化算法。该算法能在不消耗额外样本的条件下获取设计变量间的交互信息,并基于该信息一方面将高维问题分解为多个低维问题,实现高维问题的高效优化,另一方面为后续知识挖掘提供信息基础。在完成典型30维函数算例测试后,利用所提出的算法完成了包含28个设计变量的Rotor 37叶栅气动优化与设计空间知识挖掘。结果表明,优化后,Rotor 37叶栅设计工况效率相对参考设计提高了1.69%,且变工况性能相对参考设计亦有不同程度提高。在获得最优解的同时,所提出的算法还厘清了Rotor 37叶栅设计空间各变量间交互作用关系。进一步,结合算法所揭示的变量交互作用进行知识挖掘与CFD验证,证明对分离线附近型线进行微调可有效降低激波与边界层相互作用,从而提高跨声速叶栅的气动性能。由此,所提出的算法在求解高维大资源叶轮机械优化设计问题上的适用性和有效性得到验证。