According to Emmons's classic theory, compressor stall is triggered by high incidence in one certain blade passage and its propagation is driven by flow diversion toward adjacent passages. In this theory, blades are assumed to be axisymmetric around the annulus. In practice, the assembly usually causes the non-axisymmetric distribution of the blades. The focus of this paper is to answer how mistuned blades affect stall inception and whether even can mitigate its propagation. This paper conducts detailed unsteady measurements on an axial compressor with mistuned rotor blades, where the stagger angle is of the priority of blades mistuning. The results indicate that blade loading around the annulus varies and correlates obviously with the stagger angle. Owing to the difference in blade loading, the flow in three specific blades (with higher stagger angle) shows less unsteadiness, compared to that of other blade passages. For the base configuration (the error of stagger angle within 1°), stall inception revolution does not show any difference. Of interest, when the error increases to 4°, disturbances surprisingly occur and propagate around the annulus at the stable operation condition near the stall boundary, which is triggered by the mistuned blade and will decay within one revolution in the rotor frame of reference. The phenomenon demonstrates the suppression effect of other blades with changed stagger angle on the propagation of stall inception. With the compressor throttled further, full rotating stall eventually occurs once disturbances can propagate around the entire annulus. Therefore, the aerodynamic mistuning has a clear effect on the stall inception and propagation, which can be further discovered for the stall control and stability design.

A quantitative model to predict the boundary of instability of axial compressors based on their maximum loading capability is proposed in this paper, which is an improved version of the classic method of stalling pressure rise. The original model correlates the maximum pressure rise of a compressor to a characteristic geometric parameter, which is an analogue of the normalized length of diffusion in two-dimensional diffusers. However, the influence of the aspect ratio of the passage is overlooked in this analogy, which leads to significant discrepancies in its predictions for compressors, especially those with varying blade aspect ratios. Our model contains two improvements to address this issue. The first involves refining in the definition of the normalized length of diffusion, whereas the second introduces a supplementary correction factor for the aspect ratio of the blades. Nearly 20 low-speed compressor configurations, with variations in solidity, aspect ratio, tip clearance, and axial spacing, were tested to develop the proposed model. It can reduce error in the predicted stalling static pressure rise from 10% to 5%. Experimental data robustly verify the accuracy of our model, making it a more reliable predictive tool of instability boundary in the preliminary design of axial compressors.

Abstract This paper applies a body force model developed recently to investigate the interaction between total temperature distortion and a multistage fan. The off-design performance of the fan shows the reasonable predicting accuracy and supports the present model is applicable for high-speed multistage machines. The transfer behaviors of 90° steady-state circumferential total temperature distortion as well as combined total pressure and total temperature distortion in the multistage environment are captured successfully by the model. The mechanism of the phase shift of the high temperature sector is discussed by the model to advance the understanding of the total temperature distortion problem. The results reveal that the large-scale flow feature of total temperature distortion in the multistage environment can be capably quantified by the present body force model with the acceptable computational consumption.

针对常规可移动式插板产生的总压畸变稳态分量和动态分量比值较为固定,无法可调,不能真实反映不同进气条件下的复杂总压畸变的动稳态比例的现状,提出了一种改进型插板,采用对常规插板开孔和开齿等方法,通过不同的参数化设计,通过更换插板在发动机进口截面产生不同稳态畸变和动态畸变比值的总压畸变,采用分离涡模拟的方法对改进型插板进行了数值仿真,结果表明:在边缘开孔或开齿会改变插板后涡结构,但不会改变动稳态畸变比例.在插板上均匀开孔或开齿,既会改变板后涡结构,也会降低动稳态畸变比例,而且随着孔齿数量或尺寸的增加,分离区得到更多的射流能量,动稳态畸变比例降低.变齿数插板畸变比例可在0.14～0.50范围内变化.

The compressor is a critical component that determines the aerodynamic stability of an aero-engine. Total pressure inlet distortion decreases the thrust and shrinks the stability margin, thus inducing severe performance degradation or even flameout. Generally, tip air injection is used to reduce the adverse influence of total pressure inlet distortion on the aerodynamic stability. In the present work, an experimental investigation on the effects of tip air injection on the stability of a two-stage low-speed axial compressor with total pressure inlet distortion was carried out. A flat baffle generated the total pressure distortion at the inlet of the compressor. The stall margin of the compressor was reduced significantly by the total pressure distortion. When the dimensionless insertion depth of the flat baffle was 0.45, the stall margin decreased to 11.4%. Under the total pressure inlet distortion, tip air injection effectively improved the distortion resistance capability of the compressor. The circumferential layout of the nozzle played a critical role in the stability expansion effect of tip air injection under the inlet flow condition of the total pressure distortion. The modal wave disturbance was likely to occur in the distortion-affected region (the low-pressure region and the mixing region). Tip air injection did not inhibit the generation of the modal wave but restrained the development of the modal wave into the stall cell. It improved the low-speed compressor's tolerance to the modal wave and allowed a higher amplitude modal wave to occur.

为了研究进气旋流畸变对压气机性能和稳定性的影响,设计了一种能够产生典型对涡与整体涡可变弯度叶片式旋流畸变发生器.结合正交仿真试验设计方法,分别以对涡旋流强度为优化目标和以对涡旋流强度、整体涡旋流强度和整体涡总压恢复系数为综合优化目标对旋流畸变发生器的几何参数,包括叶片稠度、叶片数量以及轮毂比等进行气动优化设计,并采用CFD数值模拟仿真研究了旋流畸变发生器生成旋流特征.经过单指标优化分析,旋流畸变发生器生成对涡旋流强度最高可达24.60°,整体涡旋流强度最高可达38.73°.经过多指标综合优化,旋流畸变发生器生成对涡和整体涡的总压恢复系数分别提高了 4.26％和3.57％.叶片式旋流畸变发生器设计具有结构简单、操作方便、试验周期短等优点,并具有较好的工程应用性.

This paper presents a satisfactory numerical strategy to reliably evaluate the three-dimensional large-scale flow feature of multistage axial compressors in response to complex swirl distortion with acceptable computational cost. Under the theoretical framework of the body force method, the guide vanes of a swirl distortion generator and the multiple blade rows of a two-stage low-speed axial compressor are described by distributed source terms instead of a complex body-fitted grid approach. The key flow structure of the paired swirl generated by the swirl generator and the main distributions of flow angle at the rotor outlet of the first stage captured by the model agree well with experimental results, demonstrating the effectiveness of the numerical strategy. Additionally, the interaction process between the steady-state paired swirl and the compressor is clearly revealed by the study. The intensity of the swirl distortion can be greatly reduced after passing through the axial compressor. However, the swirl has a significant impact on the local blade loading of the first stage, which induces the mass flux nonuniformity as well as total pressure and total temperature distortion. The combined total pressure and total temperature distortion is significantly attenuated near tip and slightly enhanced near hub as it moves through the second stage.

A distortion generator equipped with a motor-activated movable flat baffle was installed just upstream of a rectangular plenum entrance to investigate the effects of inlet total pressure distortion on the stability and performance of an auxiliary power unit (APU). Experiments and numerical simulations on a direct connect scale inlet model of the APU were carried out to obtain a quantitative relationship between the insertion depth of the flat baffle in the flow stream and the total pressure distortion intensity and region. In the experiments, the blocking coefficient and total pressure distortion coefficient were controlled by adjusting the insertion depth of the flat baffle and the mass flow. In the simulations, detailed flow field was analyzed based on the detached-eddy simulation (DES) method. The results show that the pressure distribution of the distorted flow on the aerodynamic interface plane (AIP) can be divided into a high-pressure region, a transition region, and a low-pressure region. The area affected by the distorted flow was larger than the inserting area of the flat baffle. That area was more related to the relative blocking coefficient, and less affected by the mass flow. The total pressure distortion coefficient had a linear relationship with the mass flow rate and is positively correlated with the relative blocking coefficient. As the relative blocking coefficient increased to a certain value, an exponential growth in the total pressure distortion coefficient occurred, and consequently, the flow field distortion was intensified. In the flow field, a pair of corner vortices were formed at the corner between the flat baffle and the bottom wall of the inlet pipe, and a large separation zone was formed behind the flat baffle and exhibits certain unsteady characteristics.

为研究叶尖泄漏流对稳定性的影响,发展了一种叶尖泄漏涡模型,并且在由课题组开发的TU-SIAC(three dimensional and unsteady stall inception analysis code)程序中实现.该程序将转/静子叶排模化为三维激盘,并在无叶区求解三维非定常欧拉方程,黏性的影响通过特性曲线体现,因此对计算资源的要求较低.数值模拟的结果表明:在叶尖间隙的影响下,压气机的性能和稳定裕度均有所下降.叶尖间隙不改变失速先兆的类型,但使得失速团旋转速度加快,周向尺寸减小.所发展的叶尖涡模型能够预测叶尖间隙尺寸对压气机稳定边界与失速起始过程的影响,从而为在设计初期考虑压气机的气动稳定性、并且优化压气机设计,提供了一种较为实用的方法.

In order to detect the aerodynamic instability of a multistage axial compressor more accurately and earlier, the harmonic Fourier mean amplitude analysis method and heterotopic variance analysis method are developed. The dynamic instability prediction performance of the two methods is studied on a low-speed and a high-speed two-stage axial compressor. The harmonic Fourier mean amplitude analysis method is suitable for predicting the aerodynamic instability of a multistage axial compressor in the form of a rotating stall. Compared with the traditional harmonic Fourier analysis methods, the harmonic Fourier mean amplitude analysis method can capture the detail of the pressure signal more accurately and it can effectively prevent instability misjudgment. The heterotopic variance analysis method is developed based on the conventional variance analysis method, and it can be used to distinguish whether the compressor is in the rotating stall or the surge state. The heterotopic variance analysis method can predict the aerodynamic instability ahead of the harmonic Fourier mean amplitude analysis method, and fewer circumferential measuring points were employed. The layout of the measuring points also influences the detection of the aerodynamic instability of the compressor. The aerodynamic instability of the high-speed axial compressor can be predicted earlier by employing measuring points at the compressor outlet.

为进一步加深对压气机静子角区分离流动非定常性和湍流特性的认识,以某台用于低速模拟的多级低速轴流压气机的第3级改型静子为研究对象,采用延迟脱落涡模拟(DDES)方法进行详细的数值研究.结果表明:压气机静子角区分离流动受到叶片几何参数和来流条件的综合影响;通道涡是引起静子角区分离的主要涡系结构,其发展过程伴随着大小尺度发簪涡的交替出现;通道涡、发簪涡和尾迹脱落涡的相互作用是引起静子角区分离流动非定常性的主要来源;静子角区存在高度各向异性和能量反传的湍流特征;角区分离区的频率幅值强于主流区,振幅较大的区域位于500Hz以下的低频区.

In order to improve the performance of high-loaded compressor stator with large camber turning angle, a stator cascade blade with tubercle leading edge was designed based on the wavy leading edge of humpback's flipper, and computational fluid dynamics simulation was carried out. The results show that the tubercle leading edge can effectively improve the aerodynamic performance of the stator cascade at high attack angles, and the total pressure loss coefficient can be reduced by 26.46%. The main reason why the tubercle structure improves the performance is that it makes the radial displacement of airflow appear as a butterfly-like structure at the leading edge of the blade, which restrains the occurrence and development of airflow separation. By comparing the performance of the blades with full-span and part-span tubercle leading edges, it is considered that the tubercle leading edge of 80% blade height in the mainstream region can improve the aerodynamic performance of cascade better, while the one of 50% blade height has the worst effect.

To investigate the effect of high temperature steam ingestion on the aerodynamic stability of a multistage axial compressor, a two-stage low-speed axial compressor was studied, and full-annulus steady-state and unsteady-state numerical simulations were carried out. The effect of the high temperature steam mass fraction and the distribution of steam at the inlet boundary on the aerodynamic stability of a two-stage low-speed axial compressor was investigated. From the simulation results, we found that high temperature steam ingestion has an adverse effect on the low-speed axial compressor. The larger the steam mass fraction is, the greater the impact of the steam ingestion on the stability boundary and stall margin will be. When the steam mass fraction is equal to 0.35 and 0.7%, the stability margin decreases from 36.07 to 29.72% and 28.05%, respectively. The distribution of steam at the inlet boundary will change the performance and stability. When the steam ingestion range is less than 90°, the steam ingestion area increases and the stability margin will decrease gradually. After 90°, the stability margin is almost unchanged. The difference between the calculated and experimental values of the stability margin reduction caused by steam ingestion is 0.87%. In addition, with the ingestion of high temperature steam, the blockage in the corresponding passages is intensified and the loss is increased, which leads to the occurrence of the stall in advance. It is evident that steam ingestion has a significant impact on compressor stability, ensuring that the steam mass fraction and steam ingestion range are close to the actual value.

In order to investigate the influence of steam ingestion on the aerodynamic stability of a two-stage low-speed axial-flow compressor, multiphase flow numerical simulation and experiment were carried out. The total pressure ratio and stall margin of the compressor was decreased under steam ingestion. When the compressor worked at 40% and 53% of the nominal speed, the stall margin decreased, respectively, by 1.5% and 6.3%. The ingested steam reduced the inlet Mach number and increased the thickness of the boundary layer on the suction surface of the blade. The low-speed region around the trailing edge of the blade was increased, and the flow separation region of the boundary layer on the suction surface of the blade was expanded; thus, the compressor was more likely to enter the stall state. The higher the rotational speed, the more significant the negative influence of steam ingestion on the compressor stall margin. The entropy and temperature of air were increased by steam. The heat transfer between steam and air was continuous in compressor passages. The entropy of the air in the later stage was higher than that in the first stage; consequently, the flow loss in the second stage was more serious. Under the combined action of steam ingestion and counter-rotating bulk swirl distortion, the compressor stability margin loss was more obvious. When the rotor speed was 40% and 53% of the nominal speed, the stall margin decreased by 6.3% and 12.64%, respectively.

The aim of this article mainly lies in two aspects. The first is to investigate the effect of inlet swirl distortion on the performance and stability of a low-speed compressor experimentally. The second is to quantify swirl pattern revolution through the compressor and find out background causes of the change in compressor performance. Swirl distortion makes the leading-edge incidence opposite between tip and hub regions, compared to that of clean flow. And the compressor performance change is ultimately determined by these two aspects. Results indicate that negative bulk swirl improves pressure rise, and the effect is on the contrary to the positive bulk swirl. Under the condition of paired swirl, pressure rise also presents a reduction. All these three types of swirl have little effect on the stall boundary. Although swirl distortion shows clear recovery at rotor exit, downstream components still work at off-design conditions due to the induced nonuniformity in axial velocity and total pressure.

In order to simulate the stall boundary of a compressor more accurately under steam ingestion condition, a two-dimensional stability analysis model is developed. The model simulates compressor flow by solving compressible, two-dimensional, unsteady Euler equations with appropriate source terms. Normally, viscosity is considered for compressor flow when compressor characteristics, expressed by the Navier-Stokes equations or experiments, are used to calculate the source terms, and engineering applications pay more attention to compressor characteristics rather than detailed flow field. Therefore, the control equations are assumed to be non-viscous in this work. A corrected correlation method that does not need numerous computations under different proportions of steam is proposed at present. The method reflects the effects of gas properties on the compressor characteristic lines — a quality absent in traditional correlation method. It is most significant for the paper to adopt the proposed corrected correlation to reduce the number of independent parameters that describe compressor characteristics under different proportions of steam compared with the traditional correlation. Then the source terms associate with compressor inlet parameters by adopting the corrected correlation method to establish a modified model. The model is used to calculate the stall boundary under steam ingestion, and the results are compared with clean condition to investigate the effects of steam properties on stall boundary. As more steam is ingested, the corrected similar flow increases at the stall point while the total pressure ratio decreases when total temperature distortion remains constant. If you want to describe compressor characteristics under ingesting other gases, just modify these parameters related to gas properties in the corrected correlation.

基于数值缩放理念,以一台大涵道比涡扇发动机为研究载体,将三维彻体力模型与二维多子平行发动机模型进行有机结合,初步搭建了适用于分析复杂进气畸变对航空发动机整机流场特性影响的多维耦合计算模型.利用该多维耦合模型定量分析了稳态周向总压畸变及插板式总压畸变进气下发动机内部的流场特性.计算结果表明即便是进口单纯的方波周向总压畸变进气,在经过大涵道比风扇转子作用后在下游内外涵进口的畸变流场特征也将具有显著差异;在内涵进口形成的总压/总温复合畸变在四级增压级中均得到不同程度的衰减;插板式总压畸变进气下风扇转子进口近轮毂处受周向静压梯度驱动将产生一定程度旋流,并导致风扇转子出口轮毂处形成低压区.

Inlet swirl distortion will degrade compressor performance and aerodynamic stability. In order to investigate its mechanism, a blade type swirl distortion generator was designed, which is competent to produce both of the twin swirl and bulk swirl. Experiment was carried out using this swirl generator. A twin swirl case was simulated numerically using steady-state and unsteady calculation methods, respectively. The results show that bulk swirl may not have the same effect as pre-swirl inlet condition, since the center of the vortex is not on the axis of rotation, causing more complicated results. All forms of the swirl will decrease the total pressure ratio and efficiency, but have different influences on the stall margin and stable region, which is analyzed in detail. Under the settings of the experiment, the co-rotating part of the twin swirl has a small impact on the compressor, and that of the counter-rotating part is strong, which causes suction surface boundary layer separation of some of the blade roots. This is the reason for the reduction of the aerodynamic stability of the compressor. The stall margins given by the steady-state and unsteady-state solutions differ significantly, and the unsteady results are closer to the experimental results. The main reason causing compressor instability is also analyzed in detail through numerical simulation. The findings of this study are novel and may provide a new perspective into the mechanism of swirl distortion effects.

研究蒸汽吸入和旋流畸变对压气机气动稳定性的影响,以某两级低速轴流压气机为试验研究对象,建立压气机蒸汽吸入和旋流畸变试验台.试验结果表明:压气机吸入蒸汽引起入口总压和总温畸变,转速增加,总压畸变强度增大,总温畸变强度减小.蒸汽吸入导致总压比和稳定裕度降低,当压气机转速为600r/min和800 r/min时,稳定裕度分别降低1.5％和6.3％.蒸汽吸入和反向整体涡旋流畸变两者共同作用下稳定裕度下降更为显著,当转速为600 r/min和800 r/min时,稳定裕度分别降低6.30％和12.64％.对试验压气机进行蒸汽吸入数值模拟研究,结果表明:压气机转速为600 r/min和800 r/min时,稳定裕度分别降低3.21％和8.12％.