The utilization of Inlet Guide Vane (IGV) plays a key factor in affecting the instability evolution. Existing literature mainly focuses on the effect of IGV on instability inception that occurs in the rotor region. Ho...The utilization of Inlet Guide Vane (IGV) plays a key factor in affecting the instability evolution. Existing literature mainly focuses on the effect of IGV on instability inception that occurs in the rotor region. However, with the emergence of compressor instability starting from the stator region, the mechanism of various instability inceptions that occurs in different blade rows due to the change of IGV angles should be further examined. In this study, experiments were focused on three types of instability inceptions observed previously in a 1.5-stage axial flow compressor. To analyze the conversion of stall evolutions, the compressor rotating speed was set to 17 160 r/min, at which both the blade loading in the stator hub region and rotor tip region were close to the critical value before final compressor stall. Meanwhile, the dynamic test points with high-response were placed to monitor the pressures both at the stator trailing edges and rotor tips. The results indicate that the variation of reaction determines the region where initial instability occurs. Indeed, negative pre-rotation of the inlet guide vane leads to high-reaction, initiating stall disturbance from the rotor region. Positive pre-rotation results in low-reaction, initiating stall disturbance from the stator region. Furthermore, the type of instability evolution is affected by the radial loading distribution under different IGV angles. Specifically, a spike-type inception occurs at the rotor blade tip with a large angle of attack at the rotor inlet (−2°, −4° and −6°). Meanwhile, the critical total pressure ratio at the rotor tip is 1.40 near stall. As the angle of attack decreases, the stator blade loading reaches its critical boundary, with a value of approximately 1.35. At this moment, if the rotor tip maintains high blade loading similar to the stator hub, the partial surge occurs (0° and +2°);otherwise, the hub instability occurs (+4° and +6°).展开更多
The influence of Impedance Boundary Condition (IBC) on transonic compressors is investigated. A systematic input–output analytical framework is developed, which treats the nonlinearities as unknown forcing terms. The...The influence of Impedance Boundary Condition (IBC) on transonic compressors is investigated. A systematic input–output analytical framework is developed, which treats the nonlinearities as unknown forcing terms. The framework is validated through the experiments of rotating inlet distortion within a low-speed compressor. The input–output method is subsequently applied to transonic compressors, including NASA Rotor37 and Stage35, wherein impedance optimization is studied along with the exploration of its fundamental mechanisms. The IBC is employed to model the effect of Casing Treatment (CT). The optimal complex impedance values are determined through predicted results and tested across a range of circumferential modes and forcing frequencies. The IBC significantly reduces the energy and Reynolds stress gain, notably at the first-order circumferential mode and within the Rotor Rotating Frequency (RRF) range. Output modes reveal that transonic compressors with fine-tuned impedance values exhibit a more confined perturbation distribution and redistribute the perturbations compared to the uncontrolled case. Additionally, the roles of resistance and reactance are elucidated through input–output analysis, and resistance determines the energy transfer direction between flow and pressure waves and modulates the amplitude, whereas reactance modifies the phase relationships and attenuates the perturbations.展开更多
In this paper,a series of flutter simulations are carried out to investigate the effects of split drag rudder(SDR)on the transonic flutter characteristic of rigid NACA 64A010.A structural dynamic model addressing two-...In this paper,a series of flutter simulations are carried out to investigate the effects of split drag rudder(SDR)on the transonic flutter characteristic of rigid NACA 64A010.A structural dynamic model addressing two-degree-of-freedom pitch-plunge aeroelastic oscillations was coupled with the unsteady Reynolds-averaged Navier-Stokes equations to perform flutter simulation.Meanwhile,the influence mechanism of SDR on flutter boundary is explained through aerodynamic work and the correlated shock wave location.The results show that the SDR delays the shock wave shifting downstream,and the Mach number corresponding to reaching freeze region increases as the split angle increases.Therefore,the peak value of aerodynamic moment coefficient amplitude and the sharp ascent process of phase occurs at higher Mach number,which leads to the delay in the occurrence of the transonic dip.Besides,before the transonic dip of airfoil without SDR occurs,the aerodynamic moment phase of airfoil with the SDR decreases slowly due to the decrease in the speed of shock wave moving downstream.This results in an increased flutter speed when employing the SDR before the transonic dip of airfoil without SDR occurs.Meanwhile,the effects of asymmetric split angles on the transonic flutter characteristics are also investigated.Before the transonic dip of airfoil without SDR occurs,the flutter characteristic is dominated by the smaller split angle.展开更多
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...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.展开更多
The application of higher bypass ratios and lower pressure ratios significantly reduces specific fuel consumption with the development of turbofan engines.However,it also increases the risk of flow separation at the i...The application of higher bypass ratios and lower pressure ratios significantly reduces specific fuel consumption with the development of turbofan engines.However,it also increases the risk of flow separation at the intake,leading to severe circumferential non-uniform inlet conditions.This study aimed to present an experimental investigation on instability evolutions of the compressor under circumferential non-uniform inlet conditions.Two stall inceptions regarding the different spatial scales and initial locations were selected to investigate this issue.The experiments were carried out on one tested rig,which the stall inceptions verified with the rotational speeds.At 65%design rotational speed(X),the stall inception was the spike,which was triggered by disturbances within serval pitches scale at the tip.Consequently,the spike-type stall inception was sensitive to circumferential distortion and led to a shrunk stall margin of the compressor.With the rotational speed increasing to 88%X,the stall inception switched to partial surge,which was induced by the flow blockage in the hub region around the full-annular.The results indicated that the partial surge was insusceptible to the circumferential distortion,which caused an extended stall margin with a lower stalled mass flow rate.In summary,the influence of distortion on the stability of the target compressor was found to be determined by the stall inception.展开更多
The pneumatic probe is widely used for contact measurements in turbomachinery flow field research.However,it inevitably interferes with the original flow field,leading to additional errors,particularly in wake flow fi...The pneumatic probe is widely used for contact measurements in turbomachinery flow field research.However,it inevitably interferes with the original flow field,leading to additional errors,particularly in wake flow fields or transonic regions with significant pressure gradients.This study employed Reynolds-Averaged Navier-Stokes delete and high-fidelity numerical simulation to investigate the impact of an inserted pneumatic probe on the wake flow field of a transonic turbine blade and compared it to the baseline flow field.Results indicate that the probe causes the shock waves premature occurrence in the high subsonic wake region near the turbine blade trailing edge.These shock waves affect vortex shedding by thickening the boundary layer near the trailing edge and changing the shedding pattern from high-frequency-low-energy to low-frequencyhigh-energy.In addition,the extra flow loss is incurred,and the blade's heat transfer characteristic is changed.This research provides a reference for testing experiments in complex transonic flow fields,guiding experimental researchers to minimize instrument interference with the original flow field.展开更多
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 t...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.展开更多
This paper is devoted to studying the stability of transonic shock solutions to the Euler-Poisson system in a one-dimensional nozzle of finite length.The background charge in the Poisson equation is a piecewise consta...This paper is devoted to studying the stability of transonic shock solutions to the Euler-Poisson system in a one-dimensional nozzle of finite length.The background charge in the Poisson equation is a piecewise constant function.The structural stability of the steady transonic shock solution is obtained by the monotonicity argument.Furthermore,this transonic shock is proved to be dynamically and exponentially stable with respect to small perturbations of the initial data.One of the crucial ingredients of the analysis is to establish the global well-posedness of a free boundary problem for a quasilinear second order equation with nonlinear boundary conditions.展开更多
An investigation on the ventral diverterless high offset S-shaped inlet is carried out at Mach numbers from 0.600 to 1.534, angles of attack from -4° to 9.4°, and yaw angles from 0° to 8°. Results ...An investigation on the ventral diverterless high offset S-shaped inlet is carried out at Mach numbers from 0.600 to 1.534, angles of attack from -4° to 9.4°, and yaw angles from 0° to 8°. Results indicate： （1） a large region of low total pressure exists at the lower part of the inlet exit caused by the counter-rotating vortices in the S-shaped duct; （2） the performances of the inlet at Mach number 1.000 reach almost the highest, so the propulsion system could work efficiently in terms of aerodynamics; （3） the total pressure recovery increases slowly at first and then remains unvaried as the Mach number rises from 0.6 to 1.0, however, it does in an opposite manner in the conventional diverter-equipped S-shaped inlet; （4） the performances of the inlet are generally insensitive to angles of attack from -4° to 9.4° and yaw angles from 0° to 8° at Mach number 0.850, and angles of attack from -2° to 6° and yaw angles from 0° to 5° at Mach number 1.534.展开更多
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 inter...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.展开更多
A coupled fluid-structure method is developed for flutter analysis of blade vibrations in turbomachinery. The approach is based on the time domain solution of the fluid-structure interaction in which the aerodynamic a...A coupled fluid-structure method is developed for flutter analysis of blade vibrations in turbomachinery. The approach is based on the time domain solution of the fluid-structure interaction in which the aerodynamic and structural equations are marched simultaneously in time. The three-dimensional （3D） unsteady Reynolds average Navier-Stokes （RANS） equations are solved with a multiblock finite volume scheme on dynamic deforming grids to evaluate the aerodynamic force. Dual time-stepping technique and an efficient implicit scheme with multigrid are employed to march the solution in time. The blade vibration is modeled with an aeroelasticity model in which blade motion is computed by linear combination of responses of each mode under unsteady loads. The code is validated in prediction of the unsteady flow flutter behavior of an oscillating cascade and is applied to flutter analysis of a transonic fan at the design speed.展开更多
In transonic flow,buffet is a phenomenon of flow instability caused by shock wave/boundary layer interaction and flow separation.The phenomenon is common in transonic flow,and it has serious impact on the structural s...In transonic flow,buffet is a phenomenon of flow instability caused by shock wave/boundary layer interaction and flow separation.The phenomenon is common in transonic flow,and it has serious impact on the structural strength and fatigue life of aircraft.In this paper,three typical airfoils:the supercritical OAT15A,the high-speed symmetrical NACA64A010,and the thin,transonic/supersonic NACA64A204 are selected as the research objects.The flow fields of these airfoils under pre-buffet and buffet onset conditions are simulated by Unsteady Reynolds Averaged Navier-Stokes (URANS) method,and the mode analysis of numerical results is carried out by Dynamic Mode Decomposition (DMD).Qualitative and quantitative analysis of the shock wave motion,shock wave intensity,shock foot bubble and trailing edge separation,and pressure coefficient fluctuation were performed to attain deep insight of transonic buffet flow features of different airfoils near buffet onset conditions.The results of DMD analysis show that the energy proportion of the steady mode of these airfoils decreases dramatically when approaching the buffet onset angle of attack,while the growth rate of the primary mode increases inversely.It was found that at the onset of buffet,there exist different degrees of merging behavior between shock foot bubble and trailing edge separation during one buffet cycle,and the instability of shock wave and separation induced shear layer are closely related to the merging behavior.展开更多
A Non-Intrusive Reduced-Order Model(NIROM)based on Proper Orthogonal Decomposition(POD)has been proposed for predicting the flow fields of transonic airfoils with geometry parameters.To provide a better reduced-order ...A Non-Intrusive Reduced-Order Model(NIROM)based on Proper Orthogonal Decomposition(POD)has been proposed for predicting the flow fields of transonic airfoils with geometry parameters.To provide a better reduced-order subspace to approximate the real flow field,a domain decomposition method has been used to separate the hard-to-predict regions from the full field and POD has been adopted in the regions individually.An Artificial Neural Network(ANN)has replaced the Radial Basis Function(RBF)to interpolate the coefficients of the POD modes,aiming at improving the approximation accuracy of the NIROM for non-samples.When predicting the flow fields of transonic airfoils,the proposed NIROM has demonstrated a high performance.展开更多
Transition prediction is a hot research topic of fluid mechanics.For subsonic and transonic aerodynamic flows,e^(N) method based on Linear Stability Theory(LST)is usually adopted reliably to predict transition.In 2013...Transition prediction is a hot research topic of fluid mechanics.For subsonic and transonic aerodynamic flows,e^(N) method based on Linear Stability Theory(LST)is usually adopted reliably to predict transition.In 2013,Coder and Maughmer established a transport equation for Tollmien-Schlichting(T-S)instability so that the e^(N) method can be applied to general Reynolds-Average-Navier-Stokes(RANS)solvers conveniently.However,this equation focuses on T-S instability,and is invalid for crossflow instability induced transition which plays a crucial role in flow instability of three-dimensional boundary layers.Subsequently,a transport equation for crossflow instability was developed in 2016,which is restricted to wing-like geometries.Then,in 2019,this model was extended to arbitrarily shaped geometries based on local variables.However,there are too many tedious functions and parameters in this version,and it can only be used for incompressible flows.Hence,in this paper,after a large amount of LST analyses and parameter optimization,an improved version for subsonic and transonic boundary layers is built.The present improved model is more robust and more concise,and it can be applied widely in aeronautical flows,which has great engineering application value and significance.An extensive validation study for this improved transition model will be performed.展开更多
In this paper, a new permeable adaptive integration surface is developed in order to evaluate transonic rotor noise in accordance with FW-H_pds equations(Ffowcs Williams-Hawkings equations with penetrable data surface...In this paper, a new permeable adaptive integration surface is developed in order to evaluate transonic rotor noise in accordance with FW-H_pds equations(Ffowcs Williams-Hawkings equations with penetrable data surface). Firstly, a nonlinear near-field solution is computed on the basis of Navier-Stokes equations, which is developed on moving-embedded grid methodology.The solution calculated through the present CFD method is used as the input for acoustic calculations by FW-H_pds equations. Then, two criteria for constructing integration surfaces are established based on the analysis of the quadrupole source strength and the nonlinear characteristic.A new surface is determined adaptively by the pressure gradient or density in a given flowfield,eschewing the uncertainties associated with determining cylinder-shaped integration surfaces. For varying hover cases, transonic noises are simulated with new integration surfaces for a UH-1 model rotor. Furthermore, numerical results of the new integration surface derived from the density perturbation value conform better to experimental data than results derived from the pressure gradient.Finally, the integration surface given by jrqj being 0.1, which is an applicable criterion obtained from hover cases, is used to predict transonic rotor noise in forward flight. The computational accuracy of the new integration surface method has been validated in predicting transonic rotor noise of an AH-1 model rotor at different advance ratios.展开更多
To simulate the transonic atomization jet process in Laval nozzles,to test the law of droplet atomization and distribution,to find a method of supersonic atomization for dust-removing nozzles,and to improve nozzle eff...To simulate the transonic atomization jet process in Laval nozzles,to test the law of droplet atomization and distribution,to find a method of supersonic atomization for dust-removing nozzles,and to improve nozzle efficiency,the finite element method has been used in this study based on the COMSOL computational fluid dynamics module.The study results showed that the process cannot be realized alone under the two-dimensional axisymmetric,three-dimensional and three-dimensional symmetric models,but it can be calculated with the transformation dimension method,which uses the parameter equations generated from the two-dimensional axisymmetric flow field data of the three-dimensional model.The visualization of this complex process,which is difficult to measure and analyze experimentally,was realized in this study.The physical process,macro phenomena and particle distribution of supersonic atomization are analyzed in combination with this simulation.The rationality of the simulation was verified by experiments.A new method for the study of the atomization process and the exploration of its mechanism in a compressible transonic speed flow field based on the Laval nozzle has been provided,and a numerical platform for the study of supersonic atomization dust removal has been established.展开更多
Current research shows that the traditional shock control bump（SCB） can weaken the intensity of shock and better the transonic buffet performance. The author finds that when SCB is placed downstream of the shock, it...Current research shows that the traditional shock control bump（SCB） can weaken the intensity of shock and better the transonic buffet performance. The author finds that when SCB is placed downstream of the shock, it can decrease the adverse pressure gradient. This may prevent the shock foot separation bubble to merge with the trailing edge separation and finally improve the buffet performance. Based on RAE2822 airfoil, two types of SCB are designed according to the two different mechanisms. By using Reynolds-averaged Navier-Stokes（RANS） and unsteady Reynolds-averaged Navier-Stokes（URANS） methods to analyze the properties of RAE2822 airfoil with and without SCB, the results show that the downstream SCB can better the buffet performance under a wide range of freestream Mach number and the steady aerodynamics characteristic is similar to that of RAE2822 airfoil. The traditional SCB can only weaken the intensity of the shock under the design condition. Under the off-design conditions, the SCB does not do much to or even worsen the buffet performance. Indeed, the use of backward bump can flatten the leeward side of the airfoil, and this is similar to the mechanism that supercritical airfoil can weaken the recompression of shock wave.展开更多
Algebraic methods and rapid deforming techniques are used to generate three-dimensional boundary-fitted dynamic grids for assemblies. The conservative full-potential equation is solved by a time-accurate approximate f...Algebraic methods and rapid deforming techniques are used to generate three-dimensional boundary-fitted dynamic grids for assemblies. The conservative full-potential equation is solved by a time-accurate approximate factorization algorithm and internal Newton iterations. An integral boundary layer method based on the dissipation integral is used to account for viscous effects. The computational results about unsteady transonic forces on wings, bodies and control surfaces are in agreement with experimental data.展开更多
Numerical simulations are performed to study the aeroelastic responses of an elastically suspended airfoil in transonic buffet flow, by coupling the unsteady Reynolds-averaged Navier- Stokes （RANS） equations and str...Numerical simulations are performed to study the aeroelastic responses of an elastically suspended airfoil in transonic buffet flow, by coupling the unsteady Reynolds-averaged Navier- Stokes （RANS） equations and structural motion equation. The current work focuses on the char- acteristic analysis of the lock-in phenomenon. Great attentions are paid to studying the frequency range of lock-in and the effects of the three parameters, namely the structural natural frequency, mass ratio and structural damping, on lock-in characteristic of the elastic system in detail. It is found that when the structural natural frequency is close to the buffet frequency, the coupling fre- quency of the elastic system is no longer equal to the buffet frequency, but keeps the same value as the structural natural frequency. The frequency lock-in occurs and stays present until the structural nature frequency is near the double buffet frequency. It means that the lock-in presents within a broad range, of which the lower threshold is near the buffet frequency, while the upper threshold is near the double buffet frequency. Moreover, the frequency range of lock-in is affected by mass ratio and structural damping. The lower the mass ratio and structural damping are, the wider the range of lock-in will be. The upper threshold of lock-in grows with the mass ratio and structural damping decreasing, but the lower threshold always keeps the same.展开更多
To predict the flutter dynamic pressure of a wind tunnel model before flutter test,an accurate Computational Fluid Dynamics/Computational Structural Dynamics(CFD/CSD)-based flutter prediction method is proposed under ...To predict the flutter dynamic pressure of a wind tunnel model before flutter test,an accurate Computational Fluid Dynamics/Computational Structural Dynamics(CFD/CSD)-based flutter prediction method is proposed under the conditions of a 2.4 m×2.4 m transonic wind tunnel with porous wall.From the CFD simulations of the flows through an inclined hole of this wind tunnel,the Nambu's linear porous wall model between the flow rate and the differential pressure is extended to the porous wall with inclined holes,so that the porous wall can be conveniently modeled as a boundary condition.According to the flutter testing approach for the current wind tunnel,the steady CFD calculation is conducted to achieve the required inlet Mach number.A timedomain CFD/CSD method is then employed to evaluate the structural response of the experimental model,and the critical flutter point is obtained by increasing the dynamic pressure step by step at a fixed Mach number.The present method is applied to the flutter calculations for a vertical tail model and an aircraft model tested in the current transonic wind tunnel.For both models,the computed flutter characteristics agree well with the experimental results.展开更多
基金support of the National Natural Science Foundation of China(No.52322603)the Science Center for Gas Turbine Project of China(Nos.P2022-B-II-004-001 and P2023-B-II-001-001)+1 种基金the Fundamental Research Funds for the Central Universities,Chinathe Beijing Nova Program of China(Nos.20220484074 and 20230484479).
文摘The utilization of Inlet Guide Vane (IGV) plays a key factor in affecting the instability evolution. Existing literature mainly focuses on the effect of IGV on instability inception that occurs in the rotor region. However, with the emergence of compressor instability starting from the stator region, the mechanism of various instability inceptions that occurs in different blade rows due to the change of IGV angles should be further examined. In this study, experiments were focused on three types of instability inceptions observed previously in a 1.5-stage axial flow compressor. To analyze the conversion of stall evolutions, the compressor rotating speed was set to 17 160 r/min, at which both the blade loading in the stator hub region and rotor tip region were close to the critical value before final compressor stall. Meanwhile, the dynamic test points with high-response were placed to monitor the pressures both at the stator trailing edges and rotor tips. The results indicate that the variation of reaction determines the region where initial instability occurs. Indeed, negative pre-rotation of the inlet guide vane leads to high-reaction, initiating stall disturbance from the rotor region. Positive pre-rotation results in low-reaction, initiating stall disturbance from the stator region. Furthermore, the type of instability evolution is affected by the radial loading distribution under different IGV angles. Specifically, a spike-type inception occurs at the rotor blade tip with a large angle of attack at the rotor inlet (−2°, −4° and −6°). Meanwhile, the critical total pressure ratio at the rotor tip is 1.40 near stall. As the angle of attack decreases, the stator blade loading reaches its critical boundary, with a value of approximately 1.35. At this moment, if the rotor tip maintains high blade loading similar to the stator hub, the partial surge occurs (0° and +2°);otherwise, the hub instability occurs (+4° and +6°).
基金co-supported by the National Natural Science Foundation of China(Nos.52325602,52306036 and 52306035)the National Science and Technology Major Project of China(No.Y2022-II-0003-0006 and Y2022-II-0002-0005)+1 种基金the project funded by China Postdoctoral Science Foundation(No.2022M720346)supported by the Key Laboratory of Pre-Research Management Centre of China(No.6142702200101).
文摘The influence of Impedance Boundary Condition (IBC) on transonic compressors is investigated. A systematic input–output analytical framework is developed, which treats the nonlinearities as unknown forcing terms. The framework is validated through the experiments of rotating inlet distortion within a low-speed compressor. The input–output method is subsequently applied to transonic compressors, including NASA Rotor37 and Stage35, wherein impedance optimization is studied along with the exploration of its fundamental mechanisms. The IBC is employed to model the effect of Casing Treatment (CT). The optimal complex impedance values are determined through predicted results and tested across a range of circumferential modes and forcing frequencies. The IBC significantly reduces the energy and Reynolds stress gain, notably at the first-order circumferential mode and within the Rotor Rotating Frequency (RRF) range. Output modes reveal that transonic compressors with fine-tuned impedance values exhibit a more confined perturbation distribution and redistribute the perturbations compared to the uncontrolled case. Additionally, the roles of resistance and reactance are elucidated through input–output analysis, and resistance determines the energy transfer direction between flow and pressure waves and modulates the amplitude, whereas reactance modifies the phase relationships and attenuates the perturbations.
文摘In this paper,a series of flutter simulations are carried out to investigate the effects of split drag rudder(SDR)on the transonic flutter characteristic of rigid NACA 64A010.A structural dynamic model addressing two-degree-of-freedom pitch-plunge aeroelastic oscillations was coupled with the unsteady Reynolds-averaged Navier-Stokes equations to perform flutter simulation.Meanwhile,the influence mechanism of SDR on flutter boundary is explained through aerodynamic work and the correlated shock wave location.The results show that the SDR delays the shock wave shifting downstream,and the Mach number corresponding to reaching freeze region increases as the split angle increases.Therefore,the peak value of aerodynamic moment coefficient amplitude and the sharp ascent process of phase occurs at higher Mach number,which leads to the delay in the occurrence of the transonic dip.Besides,before the transonic dip of airfoil without SDR occurs,the aerodynamic moment phase of airfoil with the SDR decreases slowly due to the decrease in the speed of shock wave moving downstream.This results in an increased flutter speed when employing the SDR before the transonic dip of airfoil without SDR occurs.Meanwhile,the effects of asymmetric split angles on the transonic flutter characteristics are also investigated.Before the transonic dip of airfoil without SDR occurs,the flutter characteristic is dominated by the smaller split angle.
基金co-supported by the Beijing Natural Science Foundation,China(No.3244044)the National Natural Science Foundation of China(No.52022009)+1 种基金the Science Center for Gas Turbine Project of China(No.P2022-A-II-003-001)the Key Laboratory Foundation,China(No.2021-JCJQ-LB-062-0102).
文摘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.
基金support of the National Natural Science Foundation of China(Nos.52322603,51976005,52006002,and 51906005)the Science Center for Gas Turbine Project,China(No.P2022-B-II-004-001)+5 种基金the Advanced Jet Propulsion Creativity Center,AEAC,China(No.HKCX2020-02-013)the National Science and Technology Major Project,China(No.2017-Ⅱ-0005-0018)the Fundamental Research Funds for the Central Universities,China(No.501XTCX2023146001)the Beijing Nova Program,China(No.20220484074)the Beijing Municipal Natural Science Foundation,China(No.3242016)the Collaborative Innovation Center for Advanced Aero-Engines,China。
文摘The application of higher bypass ratios and lower pressure ratios significantly reduces specific fuel consumption with the development of turbofan engines.However,it also increases the risk of flow separation at the intake,leading to severe circumferential non-uniform inlet conditions.This study aimed to present an experimental investigation on instability evolutions of the compressor under circumferential non-uniform inlet conditions.Two stall inceptions regarding the different spatial scales and initial locations were selected to investigate this issue.The experiments were carried out on one tested rig,which the stall inceptions verified with the rotational speeds.At 65%design rotational speed(X),the stall inception was the spike,which was triggered by disturbances within serval pitches scale at the tip.Consequently,the spike-type stall inception was sensitive to circumferential distortion and led to a shrunk stall margin of the compressor.With the rotational speed increasing to 88%X,the stall inception switched to partial surge,which was induced by the flow blockage in the hub region around the full-annular.The results indicated that the partial surge was insusceptible to the circumferential distortion,which caused an extended stall margin with a lower stalled mass flow rate.In summary,the influence of distortion on the stability of the target compressor was found to be determined by the stall inception.
基金supported by the National Science and Technology Major Project(Grant Nos.2017-V-0016-0068,and J2019-V-0017-0112)the National Natural Science Foundation of China(Grant No.51776011).
文摘The pneumatic probe is widely used for contact measurements in turbomachinery flow field research.However,it inevitably interferes with the original flow field,leading to additional errors,particularly in wake flow fields or transonic regions with significant pressure gradients.This study employed Reynolds-Averaged Navier-Stokes delete and high-fidelity numerical simulation to investigate the impact of an inserted pneumatic probe on the wake flow field of a transonic turbine blade and compared it to the baseline flow field.Results indicate that the probe causes the shock waves premature occurrence in the high subsonic wake region near the turbine blade trailing edge.These shock waves affect vortex shedding by thickening the boundary layer near the trailing edge and changing the shedding pattern from high-frequency-low-energy to low-frequencyhigh-energy.In addition,the extra flow loss is incurred,and the blade's heat transfer characteristic is changed.This research provides a reference for testing experiments in complex transonic flow fields,guiding experimental researchers to minimize instrument interference with the original flow field.
基金supported by the National Natural Science Foundation of China (No. 52176039)the National Science and Technology Major Project of China (No. 2017-Ⅱ-0007-0021)
文摘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.
基金supported by the National Natural Science Foundation of China(11871134,12171166)the Fundamental Research Funds for the Central Universities(DUT23LAB303)。
文摘This paper is devoted to studying the stability of transonic shock solutions to the Euler-Poisson system in a one-dimensional nozzle of finite length.The background charge in the Poisson equation is a piecewise constant function.The structural stability of the steady transonic shock solution is obtained by the monotonicity argument.Furthermore,this transonic shock is proved to be dynamically and exponentially stable with respect to small perturbations of the initial data.One of the crucial ingredients of the analysis is to establish the global well-posedness of a free boundary problem for a quasilinear second order equation with nonlinear boundary conditions.
基金National Basic Research Program of China (5130802)
文摘An investigation on the ventral diverterless high offset S-shaped inlet is carried out at Mach numbers from 0.600 to 1.534, angles of attack from -4° to 9.4°, and yaw angles from 0° to 8°. Results indicate： （1） a large region of low total pressure exists at the lower part of the inlet exit caused by the counter-rotating vortices in the S-shaped duct; （2） the performances of the inlet at Mach number 1.000 reach almost the highest, so the propulsion system could work efficiently in terms of aerodynamics; （3） the total pressure recovery increases slowly at first and then remains unvaried as the Mach number rises from 0.6 to 1.0, however, it does in an opposite manner in the conventional diverter-equipped S-shaped inlet; （4） the performances of the inlet are generally insensitive to angles of attack from -4° to 9.4° and yaw angles from 0° to 8° at Mach number 0.850, and angles of attack from -2° to 6° and yaw angles from 0° to 5° at Mach number 1.534.
基金National Natural Science Foundation of China (506460210) Chinese Specialized Research Fund for the Doctoral Program of Higher Education (20060213007)Development Program for Outstanding Young Teachers in Harbin Institute of Technology (HITQNJS.2006.046)
文摘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.
文摘A coupled fluid-structure method is developed for flutter analysis of blade vibrations in turbomachinery. The approach is based on the time domain solution of the fluid-structure interaction in which the aerodynamic and structural equations are marched simultaneously in time. The three-dimensional （3D） unsteady Reynolds average Navier-Stokes （RANS） equations are solved with a multiblock finite volume scheme on dynamic deforming grids to evaluate the aerodynamic force. Dual time-stepping technique and an efficient implicit scheme with multigrid are employed to march the solution in time. The blade vibration is modeled with an aeroelasticity model in which blade motion is computed by linear combination of responses of each mode under unsteady loads. The code is validated in prediction of the unsteady flow flutter behavior of an oscillating cascade and is applied to flutter analysis of a transonic fan at the design speed.
基金supported by the National Natural Science Foundation of China (No. 11802009)
文摘In transonic flow,buffet is a phenomenon of flow instability caused by shock wave/boundary layer interaction and flow separation.The phenomenon is common in transonic flow,and it has serious impact on the structural strength and fatigue life of aircraft.In this paper,three typical airfoils:the supercritical OAT15A,the high-speed symmetrical NACA64A010,and the thin,transonic/supersonic NACA64A204 are selected as the research objects.The flow fields of these airfoils under pre-buffet and buffet onset conditions are simulated by Unsteady Reynolds Averaged Navier-Stokes (URANS) method,and the mode analysis of numerical results is carried out by Dynamic Mode Decomposition (DMD).Qualitative and quantitative analysis of the shock wave motion,shock wave intensity,shock foot bubble and trailing edge separation,and pressure coefficient fluctuation were performed to attain deep insight of transonic buffet flow features of different airfoils near buffet onset conditions.The results of DMD analysis show that the energy proportion of the steady mode of these airfoils decreases dramatically when approaching the buffet onset angle of attack,while the growth rate of the primary mode increases inversely.It was found that at the onset of buffet,there exist different degrees of merging behavior between shock foot bubble and trailing edge separation during one buffet cycle,and the instability of shock wave and separation induced shear layer are closely related to the merging behavior.
基金supported by the National Natural Science Foundation of China(No.11802245).
文摘A Non-Intrusive Reduced-Order Model(NIROM)based on Proper Orthogonal Decomposition(POD)has been proposed for predicting the flow fields of transonic airfoils with geometry parameters.To provide a better reduced-order subspace to approximate the real flow field,a domain decomposition method has been used to separate the hard-to-predict regions from the full field and POD has been adopted in the regions individually.An Artificial Neural Network(ANN)has replaced the Radial Basis Function(RBF)to interpolate the coefficients of the POD modes,aiming at improving the approximation accuracy of the NIROM for non-samples.When predicting the flow fields of transonic airfoils,the proposed NIROM has demonstrated a high performance.
基金supported by the National Science Foundation for Young Scholars of China(No.:11802245)。
文摘Transition prediction is a hot research topic of fluid mechanics.For subsonic and transonic aerodynamic flows,e^(N) method based on Linear Stability Theory(LST)is usually adopted reliably to predict transition.In 2013,Coder and Maughmer established a transport equation for Tollmien-Schlichting(T-S)instability so that the e^(N) method can be applied to general Reynolds-Average-Navier-Stokes(RANS)solvers conveniently.However,this equation focuses on T-S instability,and is invalid for crossflow instability induced transition which plays a crucial role in flow instability of three-dimensional boundary layers.Subsequently,a transport equation for crossflow instability was developed in 2016,which is restricted to wing-like geometries.Then,in 2019,this model was extended to arbitrarily shaped geometries based on local variables.However,there are too many tedious functions and parameters in this version,and it can only be used for incompressible flows.Hence,in this paper,after a large amount of LST analyses and parameter optimization,an improved version for subsonic and transonic boundary layers is built.The present improved model is more robust and more concise,and it can be applied widely in aeronautical flows,which has great engineering application value and significance.An extensive validation study for this improved transition model will be performed.
基金supports of the National Natural Science Foundation of China (Nos. 11272150 and 11572156)the Scientific Research Innovation Program of Jiangsu Province of China (No. KYLX15_0250)
文摘In this paper, a new permeable adaptive integration surface is developed in order to evaluate transonic rotor noise in accordance with FW-H_pds equations(Ffowcs Williams-Hawkings equations with penetrable data surface). Firstly, a nonlinear near-field solution is computed on the basis of Navier-Stokes equations, which is developed on moving-embedded grid methodology.The solution calculated through the present CFD method is used as the input for acoustic calculations by FW-H_pds equations. Then, two criteria for constructing integration surfaces are established based on the analysis of the quadrupole source strength and the nonlinear characteristic.A new surface is determined adaptively by the pressure gradient or density in a given flowfield,eschewing the uncertainties associated with determining cylinder-shaped integration surfaces. For varying hover cases, transonic noises are simulated with new integration surfaces for a UH-1 model rotor. Furthermore, numerical results of the new integration surface derived from the density perturbation value conform better to experimental data than results derived from the pressure gradient.Finally, the integration surface given by jrqj being 0.1, which is an applicable criterion obtained from hover cases, is used to predict transonic rotor noise in forward flight. The computational accuracy of the new integration surface method has been validated in predicting transonic rotor noise of an AH-1 model rotor at different advance ratios.
基金Supported by the National Natural Science Foundation of China (NO: 51704146, 51274116, 51704145).
文摘To simulate the transonic atomization jet process in Laval nozzles,to test the law of droplet atomization and distribution,to find a method of supersonic atomization for dust-removing nozzles,and to improve nozzle efficiency,the finite element method has been used in this study based on the COMSOL computational fluid dynamics module.The study results showed that the process cannot be realized alone under the two-dimensional axisymmetric,three-dimensional and three-dimensional symmetric models,but it can be calculated with the transformation dimension method,which uses the parameter equations generated from the two-dimensional axisymmetric flow field data of the three-dimensional model.The visualization of this complex process,which is difficult to measure and analyze experimentally,was realized in this study.The physical process,macro phenomena and particle distribution of supersonic atomization are analyzed in combination with this simulation.The rationality of the simulation was verified by experiments.A new method for the study of the atomization process and the exploration of its mechanism in a compressible transonic speed flow field based on the Laval nozzle has been provided,and a numerical platform for the study of supersonic atomization dust removal has been established.
文摘Current research shows that the traditional shock control bump（SCB） can weaken the intensity of shock and better the transonic buffet performance. The author finds that when SCB is placed downstream of the shock, it can decrease the adverse pressure gradient. This may prevent the shock foot separation bubble to merge with the trailing edge separation and finally improve the buffet performance. Based on RAE2822 airfoil, two types of SCB are designed according to the two different mechanisms. By using Reynolds-averaged Navier-Stokes（RANS） and unsteady Reynolds-averaged Navier-Stokes（URANS） methods to analyze the properties of RAE2822 airfoil with and without SCB, the results show that the downstream SCB can better the buffet performance under a wide range of freestream Mach number and the steady aerodynamics characteristic is similar to that of RAE2822 airfoil. The traditional SCB can only weaken the intensity of the shock under the design condition. Under the off-design conditions, the SCB does not do much to or even worsen the buffet performance. Indeed, the use of backward bump can flatten the leeward side of the airfoil, and this is similar to the mechanism that supercritical airfoil can weaken the recompression of shock wave.
基金Aeronautical Science Foundation of China (99A52007)
文摘Algebraic methods and rapid deforming techniques are used to generate three-dimensional boundary-fitted dynamic grids for assemblies. The conservative full-potential equation is solved by a time-accurate approximate factorization algorithm and internal Newton iterations. An integral boundary layer method based on the dissipation integral is used to account for viscous effects. The computational results about unsteady transonic forces on wings, bodies and control surfaces are in agreement with experimental data.
基金supported by the project of the National Natural Science Foundation of China(No.11272262)
文摘Numerical simulations are performed to study the aeroelastic responses of an elastically suspended airfoil in transonic buffet flow, by coupling the unsteady Reynolds-averaged Navier- Stokes （RANS） equations and structural motion equation. The current work focuses on the char- acteristic analysis of the lock-in phenomenon. Great attentions are paid to studying the frequency range of lock-in and the effects of the three parameters, namely the structural natural frequency, mass ratio and structural damping, on lock-in characteristic of the elastic system in detail. It is found that when the structural natural frequency is close to the buffet frequency, the coupling fre- quency of the elastic system is no longer equal to the buffet frequency, but keeps the same value as the structural natural frequency. The frequency lock-in occurs and stays present until the structural nature frequency is near the double buffet frequency. It means that the lock-in presents within a broad range, of which the lower threshold is near the buffet frequency, while the upper threshold is near the double buffet frequency. Moreover, the frequency range of lock-in is affected by mass ratio and structural damping. The lower the mass ratio and structural damping are, the wider the range of lock-in will be. The upper threshold of lock-in grows with the mass ratio and structural damping decreasing, but the lower threshold always keeps the same.
基金supported by the National Natural Science Foundation of China(No.11872212)a project funded by the Priority Academic Program Development of Jiangsu Higher Education Institutions。
文摘To predict the flutter dynamic pressure of a wind tunnel model before flutter test,an accurate Computational Fluid Dynamics/Computational Structural Dynamics(CFD/CSD)-based flutter prediction method is proposed under the conditions of a 2.4 m×2.4 m transonic wind tunnel with porous wall.From the CFD simulations of the flows through an inclined hole of this wind tunnel,the Nambu's linear porous wall model between the flow rate and the differential pressure is extended to the porous wall with inclined holes,so that the porous wall can be conveniently modeled as a boundary condition.According to the flutter testing approach for the current wind tunnel,the steady CFD calculation is conducted to achieve the required inlet Mach number.A timedomain CFD/CSD method is then employed to evaluate the structural response of the experimental model,and the critical flutter point is obtained by increasing the dynamic pressure step by step at a fixed Mach number.The present method is applied to the flutter calculations for a vertical tail model and an aircraft model tested in the current transonic wind tunnel.For both models,the computed flutter characteristics agree well with the experimental results.
