Numerical Investigation of the Unsteady Flow in a Transonic Compressor with Curved Rotors

The unsteady 3D flow fields in a single-stage transonic compressor under designed conditions are simulated numerically to investigate the effects of the curved rotors on the stage performance and the aerodynamic interaction between the blade rows. The results show that, compared to the compressor with unurved rotors, the compressor under scrutiny acquires remarkable increases in efficiency with significantly reduced amplitudes of the time-dependent fluctuation. The amplitude of the pressure fluctuation around the stator leading edge decreases at both endwalls, but increases at the mid-span in the curved rotors. The pressure fluctuation near the stator leading edge, therefore, becomes more uniform in the radial direction of this compressor. Except for the leading edge area, the pressure fluctuatinn amplitude declines remarkably in the tip region of stator surface downstream of the curved rotor, but hardly changes in the middle and at the hub.

Wet Compression Performance of a Transonic Compressor Rotor at its Near Stall Point

In order to study the effects of wet compression on a transonic compressor,a full 3-D steady numerical simulation was carried out under varying conditions.Different injected water flow rates and droplet diameters were considered.The effect of wet compression on the shock,separated flow,pressure ratio,and efficiency was investigated.Additionally,the effect of wet compression on the tip clearance when the compressor runs in the near-stall and stall situations was emphasized.Analysis of the results shows that the range of stable operation is extended,and that the pressure ratio and inlet air flow rate are also increased at the near-stall point.In addition,it seems that there is an optimum size of the droplet diameter.

Flow Diagnosis and Optimization Based on Vorticity Dynamics for Transonic Compressor/Fan Rotor

This paper presents two optimized rotors. The first rotor is as part of a 3-blade row optimization (IGV-rotor-stator) of a high-pressure compressor. It is based on modifying blade angles and advanced control of curvature of the airfoil camber line. The effects of these advanced blade techniques on the performance of the transonic 1.5-stage compressor were calculated using a 3D Navier-Stokes solver combined with a vortex/vorticity dynamics diagnosis method. The first optimized rotor produces a 3-blade row efficiency improvement over the baseline of 1.45% while also improving stall margin. The throttling range of the compressor is expanded largely because the shock in the rotor tip area is further downstream than that in the baseline case at the operating point. Additionally, optimizing the 3-blade row block while only adjusting the rotor geometry ensures good matching of flow angles allowing the compressor to have more range. The flow diagnostics of the rotor blade based on vortex/vorticity dynamics indicate that the boundary-layer separation behind the shock is verified by on-wall signatures of vorticity and skin-friction vector lines. In addition, azimuthal vorticity and boundary vorticity flux (BVF) are shown to be two vital flow parameters of compressor aerodynamic performance that directly relate to the improved performance of the optimized transonic compressor blade. A second rotor-only optimization is also presented for a 2.9 pressure ratio transonic fan. The objective function is the axial moment based on the BVF. An 88.5% efficiency rotor is produced.

Numerical Investigation into the Transient Behavior of the Spike-Type Rotating Stall for a Transonic Compressor Rotor

In this paper,a numerical investigation into a spike-type rotating stall process is carried out considering a transonic compressor rotor(the NASA Rotor 37).Through solution of the Unsteady Reynolds-Averaged Navier-Stokes(URANS)equations,the evolution process from an initially circumferentially-symmetric near-stall flow field to a stable stall condition is simulated without adding any artificial disturbance.At the near-stall operating point,periodic fluctuations are present in the overall flow of the rotor.Moreover,the blockage region in the channel periodically shifts from middle span to the tip.This fluctuating condition does not directly lead to stall,while the full-annulus calculation eventually evolves to stall.Interestingly,a kind of“early disturbance”feature appears in the dynamic signals,which propagates forward ahead of the rotor.

Identification of vortices in a transonic compressor flow and the stall process

A novel vortex identification method for the visualization of the flow field is used for the study of the stall process of a transonic compressor.The parameter η4,which is one of the five invariants formed by the stain rate and vorticity tensors from the theory of modern rational mechanics,is found to have good ability to identify vortex stretching and vortex relaxation/breakdown processes,is introduced here to identify the tip leakage vortices.Compare with former generally used DPH(dynamic pressure head) contour,the new method reveals much more flow details which may advance our understanding of the compressor behaviors.The Vortices details are revealed in both peak efficiency and near stall condition.A possible stall process is also suggested based on the vortices analysis.The tip leakage flow from mid-chord,besides leading edge leakage flow,is also considered to play an important role in the stall process.

Design optimization of transonic compressor stage using CFD and response surface model

In order to shorten the design period, the paper describes a new optimization strategy for computationally expensive design optimization of turbomachinery, combined with design of experiment (DOE), response surface models (RSM), genetic algorithm (GA) and a 3-D Navier-Stokes solver(Numeca Fine). Data points for response evaluations were selected by improved distributed hypercube sampling (IHS) and the 3-D Navier-Stokes analysis was carried out at these sample points. The quadratic response surface model was used to approximate the relationships between the design variables and flow parameters. To maximize the adiabatic efficiency, the genetic algorithm was applied to the response surface model to perform global optimization to achieve the optimum design of NASA Stage 35. An optimum leading edge line was found, which produced a new 3-D rotor blade combined with sweep and lean, and a new stator one with skew. It is concluded that the proposed strategy can provide a reliable method for design optimization of turbomachinery blades at reasonable computing cost.

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.

Hybrid URANS/LES Method of Flow Fields in Axial-flow Compressor Rotor Rotor

Accurate and efficient prediction of the aerodynamic performance and flow details of axial-flow com-pressors is of great engineering application value for the aerodynamic design and flow control of axial-flow compres-sors.In this work,a delayed detached eddy simulation method is developed and applied to numerically simulate the tur-bulent channel flow and the aerodynamic performance of NASA Rotor 35.Several acceleration techniques including parallel implementation are also used to speed up the iteration convergence.The mean velocity distribution and Reyn-olds stress distribution in the boundary layer of turbulent channel flow and the aerodynamic performance curve of NASA Rotor 35 are predicted.The good agreement between the present delayed detached eddy simulation results and the available direct numerical simulation results or experimental data confirms the effectiveness of the developed meth-od in the accurate and efficient prediction of complex flow in turbomachinery.

Experimental and Numerical Investigations of Shock-Wave Boundary Layer Interactions in a Highly Loaded Transonic Compressor Cascade

Experimental and numerical investigations were conducted to investigate the variations of shock-wave boundary layer interaction(SBLI) phenomena in a highly loaded transonic compressor cascade with Mach numbers.The schlieren technique was used to observe the shock structure in the cascade and the pressure tap method to measure the pressure distribution on the blade surface.The unsteady pressure distribution on blade surface was measured with the fast-response pressure-sensitive paint(PSP) technique to obtain the unsteady pressure distribution on the whole blade surface and to capture the shock oscillation characteristics caused by SBLI.In addition,the Reynolds Averaged Navier Stokes simulations were used to compute the three-dimensional steady flow field in the transonic cascade.It was found that the shock wave patterns and behaviors are affected evidently with the increase in incoming Mach number at the design flow angle,especially with the presence of the separation bubble caused by SBLI.The time-averaged pressure distribution on the blade surface measured by PSP technique showed a symmetric pressure filed at Mach numbers of 0.85,while the pressure field on the blade surface was an asymmetric one at Mach numbers of 0.90 and 0.95.The oscillation of the shock wave was closely with the flow separation bubble on the blade surface and could transverse over nearly one interval of the pressure taps.The oscillation of the shock wave may smear the pressure jump phenomenon measured by the pressure taps.

Non-synchronous vibration of rotor blade in a six-stage transonic compressor

This paper presents an experimental study on the Non-Synchronous Vibration(NSV)in a six-stage transonic compressor.The first part of the paper describes the NSV phenomenon of Rotor 1,which occurs when both Stator 1(S1)and Stator 2(S2)or S1 only are closed.Detailed measurements and analysis are carried out for the former case through the unsteady wall pressure and the Blade Strain(BS).The spinning mode theory used in the rotor/stator interaction noise is employed to explain the relation between the circumferential wave number of the aerodynamic disturbance and the Nodal Diameter(ND)of the blade vibration.The variations of the vibration amplitudes of different blades and the Inter-Blade Phase Angles(IBPAs)at different moments suggest that the evolution of NSV is a highly nonuniform phenomenon along the circumferential direction.In addition,the difference between the wall-pressure spectra generated by the NSV and the classic flutter has been discussed.In the second part,the variations of aerodynamic loading due to the adjustment of the staggers of the Inlet Guide Vane(IGV),S1 and S2 have been investigated.It is found that closing S1 only can result in a great fluctuation to the performance of the front stages,which might be detrimental to the flow organization and increase the risk of NSV.In contrast,the effect of closing S2 only on the performance of the first two stages appears to be slighter relatively.

An improved deviation model for transonic stages in axial compressors

Deviation model is an important model for through-flow analysis in axial compressors.Theoretical analysis in classical deviation models is developed under the assumption of onedimensional flow,which is controlled by the continuity equation.To consider three-dimensional characteristics in transonic flow,this study proposes an improved theoretical analysis method combining force analysis of the blade-to-blade flow with conventional analysis of the continuity equation.Influences of shock structures on transverse force,streamwise velocity and streamline curvature in the blade-to-blade flow are analyzed,and support the analytical modelling of density flow ratio between inlet and outlet conditions.Thus,a novel deviation model for transonic stages in axial compressors is proposed in this paper.The empirical coefficients are corrected based on the experimental data of a linear cascade,and the prediction accuracy is validated with the experimental data of a three-stage transonic compressor.The novel model provides accurate predictions for meridional flow fields at the design point and performance curves at design speed,and shows obvious improvements on classical models by Carter and C¸etin.

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.

Experimental and Numerical Investigation of Non-synchronous Blade Vibration Excitation in a Transonic Axial Compressor

The complex flow phenomenon of rotating instability(RI) and its induced non-synchronous vibration(NSV) have become a significant challenge as they continuously increase aerodynamic load.This study aims to provide an understanding of the non-synchronous blade vibration phenomenon caused by the rotating instability of a transonic axial compressor rotor.In this case,blade vibrations and non-synchronous excitation are captured by strain gauges and unsteady wall pressure transducer sensors.Unsteady numerical simulations for a full-annulus configuration are used to obtain the non-synchronous flow excitation.The results show that the first-stage rotor blade exhibits an NSV close to the first bending mode;NSV is accompanied by a sharp increase in pressure pulsation;amplitude can reach 20%,and unsteady aerodynamic frequency will lock in a structural mode frequency when the blade vibrates in a large-amplitude motion.The predicted NSV frequency aligns well with the experimental results.The dominant mode of circumferential instability flow structure is approximately 47% of the number blades,and the cell size occupies 2-3 pitches in the circumferential direction.The full-annulus unsteady simulations demonstrate that the streamwise oscillation of the shedding and reattachment vortex structure is the main cause of NSV owing to the strong interaction between the tip leakage and separation vortices near the suction surface.

Effects of Number of Bleed Holes on Shock-Wave/Boundary-Layer Interactions in a Transonic Compressor Stator

An extensive numerical investigation is conducted to characterize the flow separation control in a transonic compressor cascade with a porous bleed.The bleed holes are arranged on the suction surface in a single row,two staggered rows and three staggered rows.For each bleed scheme,five bleed pressure ratios are examined at an inlet Mach number of 1.0.The results indicate that the aerodynamic performance of the cascade is significantly improved by the porous bleed.For the single-row scheme,the maximum reduction in total pressure losses is 57%.For the two-staggered-row and three-staggered-row schemes,there is an optimal bleed pressure ratio of 1.0,and the maximum reductions in total pressure loss are 68% and 75%,respectively.The low loss in the cascade is due to the well-controlled boundary layer.The new local supersonic region created by the bleed hole is the key reason for the improved boundary layer.The vortex induced by side bleeding provides another mechanism for delaying flow separation.Increasing the bleed holes could create multiple local supersonic regions,which reduce the range of the adverse pressure gradient that the boundary layer needs to withstand.This is the reason why cascades with more bleed holes perform better.

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.

叶片破损对压气机性能的影响

为了研究叶片破损对压气机内部流场的影响机理,本文针对带有破损的跨音速压气机叶片进行气动性能变化规律和机理研究。通过NASA rotor 37实验结果验证了计算方法的准确性,利用数值仿真研究了叶片破损对压气性能的影响规律,并对2种典型工况下叶片破损前后的流场结构进行对比分析,探讨引起性能变化的具体原因。研究表明:叶片破损会导致压气机的压比降低0.6631%、绝热效率降低0.7877%,且压气机的稳定裕度降低更为明显。叶尖破损主要对叶顶流场结构产生影响,使叶顶区域泄漏流增强,当其与激波相互作用时,通道内产生了流动堵塞和流动损失。本文研究成果可为实际服役中的航空发动机特性评定提供理论参考。

进气温度畸变对多级跨声速压气机气动性能影响的试验研究

进气温度畸变是严重影响压气机气动稳定性的一个重要因素。设计了一种进气电加热畸变发生器,以4级跨声速压气机试验件为研究对象,开展了不同转速下的压气机进气温度畸变气动性能测试试验。结果表明,进气温度畸变使跨声速压气机特性线向小流量工况偏移,且转速越高,偏移量越大;温度畸变导致总压比和等熵效率显著降低,其中总压比最大降幅为4.2%,等熵效率最大降幅为14.7%。高转速下畸变工况导致机组振动超限,说明进气温度畸变产生的不均匀热负荷对压气机的稳定安全运行产生了负面影响。当电加热段功率固定时,由于进气量随着转速增加而增加,高温区与低温区温差减小,进气总温畸变强度降低。

轮廓度误差对超声速压气机叶栅气动性能的影响

[目的]旨在评估轮廓度误差对压气机气动性能的影响,并为叶片鲁棒性设计提供参考。[方法]建立单峰值轮廓度误差分布数学模型,采用数值模拟方法,研究压力面和吸力面不同轮廓度组合误差对超声速压气机平面叶栅气动性能的影响。[结果]结果表明:吸力面轮廓度误差分布是影响叶栅总压损失的关键因素,随着吸力面轮廓度峰值误差位置向下游移动,总压损失系数逐渐降低;压力面和吸力面误差分布对气流折转角和静压升系数的影响趋势相反。对较低来流马赫数的叶栅,吸力面误差对气流折转角和静压升均起主导作用;对较高来流马赫数的叶栅,压力面误差对气流折转角和静压升影响明显。激波位置和激波强度、激波后扩张通道的流道型线综合决定了叶片表面和叶栅流道内的流动状态,使得近吸力面侧流动损失增大,近压力面侧流动损失减小,其综合效果决定了叶栅损失、气流折转角和静压升的变化。[结论]结果对指导跨声速压气机设计、加工和超差审理均具有重要意义。

Numerical Investigation of the Unsteady Tip Leakage Flow and Rotating Stall Inception in a Transonic Compressor

It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception.Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor.For steady analysis,the predicted results agree well with the experimental data for the estimation of compressor rotor global performance.For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage.On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale-spike type rotating stall inception at blade tip region.It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased,the rotating stall cell was further developed in the blade passage.

An analysis of the Circumferential Grooves Casing Treatment for transonic compressor flow

The stall mechanism of the NASA Rotor 37 is investigated through the analysis of the critical flow structures near the stall under the transonic condition. The performance of the rotor with Circumferential Grooves Casing Treatment (CGCT) is also studied based on the Reynolds-Averaging Navier-Stokes approach. The study finds that stall margin improvement can be achieved without significant penalty on the efficiency for the two CGCT configurations applied. The effects of circumferential grooves on the critical flow structures are studied through the analysis of the tip leakage mass and momentum transport that further re-veal the CGCT mechanism.