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.

NUMERICAL INVESTIGATION OF AXIAL FLOW COMPRESSOR CASING TREATMENT

Using an inviscid model with inlet total pressure gradient and a J. D. Denton scheme, this paper for the first time numerically solves the 3D flow field of compressor casing treatment, and also explores some boundary singularities and numerical stability. Agreement is attained in qualitative explanations of some, casing treatment test results and its mechanism.

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.

低雷诺数下轴流压气机转子3维流动特性数值模拟

为了探究轴流压气机转子在低雷诺数下的性能及3维流动特性,对跨声速转子压气机Rotor37在地面和高空下的气动性能和内部流场进行了数值模拟。结果表明:在20 km高空下,压气机在近失速点的总压比和绝热效率分别降低了8.0%和5.3%,喘振裕度降低了3.1%;在低雷诺数下,转子吸力面上的流动分离位置会提前掺混且更剧烈,导致分离附面层增加,同时由于吸力面附面层较厚,激波与边界层的作用加强,减小叶表附面层内低能流体向上运动的压力分量,低能流体在离心力的作用下更容易向叶顶附近迁移,这可能是触发转子工作失稳的重要因素;在低雷诺数下,叶片根部吸力面在84%轴向弦长处发生完全分离并形成角涡,在其向叶顶移动的过程中形成脱落涡,角涡和脱落涡面积明显增加,这是引起转子低雷诺数工作效率降低的重要原因之一。

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。

轴流压气机叶尖三维非定常流动的本征正交分解

为进一步认识跨声速轴流压气机叶尖流场的非定常流动特征,文章针对NASA Rotor 35的非定常数值计算结果,通过本征正交分解法(Proper Orthogonal Composition,POD)对近失速工况下的叶尖非定常流动进行分析。结果表明:本征正交分解法可以将叶尖非定常流动按能量高低进行模态分解。并通过分析各模态系数频谱特性,捕捉到了叶尖流场非定常流动的波动频率及对应模态在叶片前缘沿周向移动的大尺度扰动结构。该方法使用叶尖流场波动频率对应模态进行流场重构,揭示了流场非稳定流动的特征结构。

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.

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.

多级轴流压气机末级串列静子三维优化设计与分析

为充分挖掘串列叶片对压气机扩稳增效的潜能,在对多级轴流压气机末级静子串列改型的基础上,选取不同叶高处前后排叶片相对位置、负载分配、弯掠特征及端壁轮廓等25个参数为设计变量,开展某工业用6.5级轴流压气机末级串列静子叶片与端壁的一体化优化设计研究,并采用随机森林方法对串列叶片几何参数进行特征重要性分析,提炼出串列参数设计指导。结果表明:优化后压气机末级在设计工况、近喘工况和近堵工况下的绝热效率分别提升4.81%、3.37%和8.195%;对于末级串列静子,最优串列参数随叶高变化而变化;调整前后叶片的弦长比与弯角比以实现串列叶片负荷前加载、调整叶片后掠与反弯以降低叶片端壁处负荷、设置叶中部分更大的周向交错度使缝隙区流体充分加速,均能有效降低末级串列叶片流动损失;串列叶片缝隙区尺寸,尤其是周向交错度,较其他参数对串列叶片性能的影响更为显著。研究成果应对压气机三维串列叶片设计具有参考价值。

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.

Compressor Map Prediction by Neural Networks

This paper presents a study where artificial neural networks are used as a curve fitting method applying measured data from an axial compressor test rig to predict the compressor map. Emphasis is on models for prediction of pressure ratio, compressor mass flow and mechanical efficiency. Except for evaluation of interpolation and extrapolation capabilities, this study also investigates the effect of the design parameters such as number of neurons and size of training data. To reduce the effect of noise, the auto associative neural network has been applied for noise filtering of the data from the parameters used to calculate the efficiency. In summary, the results show that artificial neural network can be used for compressor map prediction, but it should be emphasized that the selection of data normalisation scale is crucial for the model where compressor mass flow is predicted. Furthermore, it is shown that the AANN （auto associative neural network） can be used to the reduce noise in measured data and thereby enhance the quality of the data.

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。

某重型燃气轮机压气机动叶弯曲变形故障分析

针对国内某热电厂首次发生的燃气轮机压气机第二级动叶顶部弯曲变形故障,从压气机进出口压力、叶片通道温度、排气温度、进口导叶(IGV)阀开度、轴承振动、余热锅炉出口烟气流量等多个监测参数入手,多参数协同分析可能造成故障的原因。在燃气轮机运行过程中,压气机出现进口静压增大0.0012 MPa,出口压力减小0.0782 MPa,叶片通道温度增#大14.09℃,排气温度增大6.12℃,2轴承振动增大10.69μm,余热锅炉出口烟气流量减少了0.25×106 Nm3/s等异常现象。揭示了轴流式压气机第二级动叶叶顶弯曲变形机理,判断动叶叶顶局部弯曲与压气机旋转失速现象密切相关。若出现压气机进口压力增大,压气机出口压力减小,叶片通道温度及排气温度升高,轴承振动增大,余热锅炉出口烟气流量减小等参数的异常变化,可初步判断压气机发生旋转失速。

基于EMD-样本熵的轴流压气机喘振分析

轴流压气机作为燃气轮机安全运行的重要设备装置,它的稳定性直接影响燃气轮机的输出效率。因此如何快速准确判断并及时规避轴流压气机的喘振工况是燃气轮机优化设计的重点。本文提出一种基于轴流压气机已有传感器处理提取特征值后对喘振工况进行判断的新方法。首先对压气机的一维传感信号进行EMD分解升维后进行数据降维至数据本质维数上,再利用样本熵处理后提取特征信号从而对喘振工况进行判定。基于三级轴流压气机喘振进行喘振的诱导试验,实现压气机的喘振工况并且利用EMD-样本熵的方法对压气机的级间压力、出口压力进行分析对比。试验结果表明,此方法提取的特征值可以有效区分出压气机正常与喘振的工况,验证了方法的可行性,为轴流轴流机的故障预测提供基础。

双喷气自循环机匣处理对对转压气机性能的影响

为发展适用于对转压气机变转速比工况下的扩稳技术,基于某两级对转压气机设计出一种新型双喷气自循环机匣处理结构,能够实现在不同转速比下提高对转压气机的失速裕度。在设计转速比下研究了轴向喷气位置对自循环机匣处理扩稳能力的影响,并选取扩稳性能最佳的双喷气自循环机匣处理方案探究其在变转速比工况下的扩稳潜力。研究结果表明,在设计转速下,双喷气自循环机匣处理主要改善后排转子的叶顶流场,失速裕度最大提高8.40%;在非设计转速比下,双喷气自循环机匣处理可将失速裕度提升8.71%~9.74%。双喷气自循环机匣处理的双射流将叶顶泄漏流和主流的交界面推向下游,同时改善了两转子的叶顶流场品质,从而推迟失速的发生。

轴流压气机跨音叶型前缘优化设计

叶型前缘的几何外形直接影响轴流压气机的综合气动性能,为了研究前缘椭圆度的最佳区间变化规律,对跨音叶型Stator35叶中截面进行了优化设计。基于参数化方法和元启发式算法建立了叶型前缘椭圆度的寻优模型,探讨了圆形和椭圆形前缘叶型的气动性能,对比分析了多种椭圆度前缘叶型的流场结构,并对椭圆前缘三维叶片的几何外形进行了气动寻优。结果表明:椭圆度前缘叶型在全工况范围内降低了总压损失系数,拓宽了3°攻角,提升了落后角和折转角性能;跨音叶型椭圆前缘设计存在最佳椭圆度区间[0.8,0.9],并且转捩点位置距前缘距离与分离泡长度负相关;优化后设计点处椭圆前缘叶片的绝热效率提高了0.4%,且在全流量范围内绝热效率均高于原始叶片,喘振裕度从9.8%提升到12.4%。

某轴流压缩机可转导叶驱动机构设计

以轴流压缩机可转导叶驱动机构为研究对象,本文指出可转导叶驱动机构是静叶可调轴流压缩机的重要组成部件之一,其作用是在机组启停或变工况时驱动调节静叶转角以满足机组运行要求,驱动机构的可靠运行对于保护机组安全运行、提高效率至关重要。针对某轴流压缩机,设计了一套驱动机构方案,建立了其力学模型,计算了驱动机构所需的驱动力,为液压设备选型提供了依据。