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.

EFFECT OF AXIAL SPACING ON COMPRESSOR STABILITY

Experimental studies are carried out at a low speed axial compressor with five different rotor/stator gaps. Analysis of the effect of axial spacing of two successive blade rows on the measured mean flow coefficient at stall inception and on the flow range of compressor under multi-cell rotating stall operating conditions proves that the stator can suppress the flow disturbance in the compressor and strengthen the stability of the compressor. Experimental data show that the stall flow coefficient decreases by reducing the axial spacing of successive blade rows. Moreover, by reducing the axial spacing, the stall pattern transition pace from multi-cell stall to single-cell stall can be shifted. And the compressor directly slips into single-cell stall at 21.0% CR axial spacing. By analyzing the pressure fluctuation closed to the surge line, it can be known that there exists an eigenfrequency where the amplitude of the oscillating pressure suddenly and dramatically increases as the compressor runs close to the surge line and this pressure disturbance is relevant to the compressor instability.

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.

Study on a High Precision Inter-stage Flow Field Test System for Axial Compressors

The accurate parameters measurement of the flow field between the stages for axial compressors is a significant demand.This paper proposes an axial compressor inter-stage flow field high-precision test system,which mainly consists of a probe motion scanning mechanism,fully automated test control software,and data processing methods.Iterative correction is applied to the original readings obtained from the scanning tests to enhance testing accuracy.Using this test system,detailed tests are conducted on a 1.5-stage subsonic axial compressor under different operating conditions.The test results effectively captured the impact of surface roughness and tip clearance variations on compressor performance.The distribution characteristics of parameters measured in inter-stage sections can characterize the effects of blade wake area and changes in aerodynamic performance at different blade heights.The developed test system can be extended to multi-stage compressors.

Numerical Simulation on Unsteady Flow Mechanism of a 1.5-Stage Axial Transonic Compressor

In this paper,a numerical simulation method is used to calculate a 1.5-stage axial transonic compressor to explore its unsteady flow mechanism.The performance curve is compared with the experimental data to verify the calculation method with a high numerical accuracy,which shows that the unsteady calculation has good reliability.According to the analysis of the data from the monitoring points under the near-stall condition,the unsteady disturbances originate from the tip region of blade and perform the strongest at the blade pressure surface with a broadband characteristic.Further analysis is conducted by combining with the characteristics of the transient flow field at the tip of blade.The results show that the unsteady pressure fluctuations are caused by the migration of the new vortex cores.These new vortex cores are generated by the breakdown of leakage vortex in the downstream,which is induced by the leakage vortex and shock wave interference.Moreover,the relationship between the unsteady flow characteristics and the working conditions is also studied.The leakage vortex intensity and the shock wave strength gradually increase with the decrease of flow rate.When the combination of the leakage vortex intensity and shock wave strength reaches the first threshold,a single frequency of unsteady disturbances appears at the blade tip.When the combination of the leakage vortex intensity and shock wave strength reaches the second threshold,the frequency of unsteady disturbances changes to a broadband.

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.

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.

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.

轴流压气机串列叶栅前后叶型位置匹配特性研究

采用数值和试验相结合的方法研究了高负荷轴流压气机串列叶栅前后叶型位置的匹配特性,详细分析了周向节距比和轴向重叠度2个参数对叶栅性能的影响规律和机理。结果表明,串列叶栅性能受周向节距比的影响较轴向重叠度更为显著,且在周向节距比的最优设计空间内,串列叶栅性能不再随2个参数的改变发生显著变化。周向节距比增大使缝隙流道收敛度和缝隙射流加速比增大,改变了前叶尾缘压力面和后叶前缘吸力面附近的流动特性,使前叶尾缘负荷增加,后叶前缘负荷和攻角降低,进而减弱了后叶吸力面分离。当周向节距比增大到缝隙流道闭合边界附近时,缝隙射流动量的显著降低导致分离损失显著增大。最后,采用叶栅风洞吹风试验,验证了数值计算得到的部分典型周向节距比和轴向重叠度匹配方案研究结果的可靠性。

Experimental and computational investigation on flow characteristics of rotating cavities with high inlet pre-swirl axial throughflow

As an integral part of the internal air system of aero-engines,the axial throughflow of the cooling air can interact with the cavity flow between the rotating compressor disks,forming a threedimensional,unsteady,and unstable flow field.The flow characteristics in an engine-like rotating multi-stage cavity with throughflow were investigated using particle image velocimetry,flow visualization technology and three-dimensional unsteady Reynolds-Averaged Navier-Stokes (RANS)simulations.The focus of current research was to understand the distribution of the mean swirl ratio and its variation with a wide range of non-dimensional parameters in the co-rotating cavity with high inlet pre-swirl axial throughflow.The maximum axial Reynolds number and rotational Reynolds numbers could reach 4.41×10^(4)and 1.24×10^(6),respectively.The velocity measurement results indicate that the mean swirl ratio is greater than 1 and decreases with an increase in the radial position.The flow structure is dominated by the Rossby number,and two different flow patterns (flow penetration and flow stratification) are identified and confirmed by flow visualization images.In the absence of buoyancy,the flow penetration caused by the precession of the throughflow makes it easier for the throughflow to reach a high radius region.Satisfactory consistency of results between measurements and numerical calculations is obtained.This study provides a theoretical basis and data support for toroidal vortex breakdown,which is of practical significance for the design of high-pressure compressor cavities.

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。

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。

Blade bowing effects on radial equilibrium of inlet flow in axial compressor cascades

The circumferentially averaged equation of the inlet flow radial equilibrium in axial compressor was deduced. It indicates that the blade inlet radial pressure gradient is closely related to the radial component of the circumferential fluctuation（CF） source item. Several simplified cascades with/without aerodynamic loading were numerically studied to investigate the effects of blade bowing on the inlet flow radial equilibrium. A data reduction program was conducted to obtain the CF source from three-dimensional（3D） simulation results. Flow parameters at the passage inlet were focused on and each term in the radial equilibrium equation was discussed quantitatively. Results indicate that the inviscid blade force is the inducement of the inlet CF due to geometrical asymmetry. Blade bowing induces variation of the inlet CF, thus changes the radial pressure gradient and leads to flow migration before leading edge（LE） in the cascades. Positive bowing drives the inlet flow to migrate from end walls to mid-span and negative bowing turns it to the reverse direction to build a new equilibrium. In addition, comparative studies indicate that the inlet Mach number and blade loading can efficiently impact the effectiveness of blade bowing on radial equilibrium in compressor design.

Relationship between the Flow Blockage of Tip Leakage Vortex and its Evolutionary Procedures inside the Rotor Passage of a Subsonic Axial Compressor

The near casing flow fields inside the rotor passage of a 1.5 stage axial compressor with different blade-loading levels and tip gap sizes were measured by using stereoscopic particle image velocimetry(SPIV). Based on a carefully defined blockage extracting method, the variations of blockage parameter inside the blade passage were analyzed. It was found that the variation of blockage parameter appeared as a non-monotonic behavior inside the blade passage in most cases. This non-monotonic behavior became much more remarkable as the blade loading increases or mass flow rate decreases.The variations of the blockage parameter inside the blade passage had close relation to the evolutionary procedures of the tip leakage vortex(TLV). The destabilization of the TLV caused a rapid increasing of the blockage parameter. After the TLV lost the features of a concentrated streamwise vortex,the blockage parameter usually got a peak value. And then, because of the intense turbulent mixing between the TLV low momentum flow and its surrounding flows, the flow deficit inside the TLV recovered.

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.

Experimental Investigation of Effects of Suction-Side Squealer Tip Geometry on the Flow Field in a Large-scale Axial Compressor using SPIV

It is well known that the non-uniform tip geometry is a promising passive flow control technique in turbomachinery.However,detailed investigation of its effects on the unsteady flow field of turbomachinery is rare in the existiug hteratures.This paper presents an experimental investigation of effects of suction side squealer tip configuration on both the steady and unsteady flow field of an isolated compressor rotor.The flow field at 10%chord downstream from the trailing edge was measured using a mini five-hole probe.Meanwhile,the unsteady flow field inside the passage was investigated using stereo particle image velocimetry(SPIV).The steady results show that the SSQ tip configuration exerts positive effect on the static pressure rise performance of this compressor,and the radial equilibrium at the rotor outlet is obviously rearranged.The SSQ tip configuration would create a stronger tip leakage vortex at the formation phase,and it experiences a faster dissipation process around the rear chord.Also,the splitting process of the tip leakage vortex is severer,which is the main cause of the relatively higher probability of the presence of the streamwise reverse flow.The quantitatively analysis of the tip leakage vortex indicates that the velocity loss inside the blockage region is direct response of the evolutionary procedure of the tip leakage vortex.It keeps increasing until the end of the splitting process.Although the blockage coefficient grows sustainably,the velocity loss will reduce once the turbulent mixing procedure is dominant.

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.