Aerodynamic performances of axial compressors are significantly affected by variation of Reynolds number in aero-engines. In the design and analysis of compressors, previous correction methods for cascades and stages have difficulties in predicting comprehensively Reynolds number effects on airfoils, matching and characteristics curves. This study proposes Re-correction models for loss, deviation angle and endwall blockage based on classical theories and cascade tests, and loss and deviation models show good agreement in test data of NACA65 and C4 cascades. Through-flow method considering Reynolds number effects is developed by integrating the correction models into a verified Streamline Curvature (SLC) tool. A three-stage axial compressor is investigated through SLC and CFD methods from design Reynolds number (Red = 2×106) to low Re = 4×104, and the numerical methods are validated with test data of characteristic curves and spanwise distributions at Red. With Re reduction, SLC method with correction models well predicts variation in overall performances compared with CFD calculations and Wassell’s model. Streamwise and spanwise matching such as total pressure and loss distributions in SLC predictions are basically consistent with those in CFD results at near-stall points under design and low Reynolds numbers. SLC and CFD methods share similar detections of stall risks in the third stage (Stg3), and their analyses of diffusion processes deviate to some extent due to different predictions in separated endwall flow. The correction models can be adopted to consider Reynolds number effects in through-flow design and analysis of axial compressors.

Composite materials are increasingly being applied in spacecraft structures. This paper investigates the nonlinear mechanical behavior of a space deployable composite boom (DCB) during the coiling process. A finite element model is established using ABAQUS software, optimizing complex boundary conditions to analyze their dynamic mechanical performance. Through simulations of DCB coiling process, the evolution of strain energy during coiling is revealed, and it is found that the driving torque stabilizes after 0.5 coiling circle. Further parametric analysis indicates the driving torque decreases gradually with the coiling radius increasing. The research results provide theoretical support for the design of DCBs and their applications in space structures.

The flutter instability characteristics and physical mechanisms of a weakly damped Ziegler double pendulum subjected to the follower type circulatory loading are investigated, while the double pendulum is assumed to be with arbitrary masses, stiffness and damping. Different with the existing mathematical analysis methods, an energy method with clear physical meaning is adopted to deduce the pendulum flutter instability boundary conditions, and to evaluate the corresponding critical parameters in this study. Thus through introducing some ratio parameters of the structural mass, stiffness and damping coefficients, the complex influences of the structural mass, stiffness and damping on the pendulum flutter instability characteristics are discussed in details. The results indicate that, in addition to the well-known counter-intuitive “damping Ziegler Paradox” influence, there also exist the stabilizing and destabilizing influences of the structural mass and stiffness. To clarify the corresponding physical mechanisms, the power flow characteristics on the pendulum flutter instability occurrence are investigated. It is observed that the structural mass and stiffness related powers on each coordinate can constitute the “power exchange twins”, and can cause the related destabilizing or stabilizing influences of the pendulum mass and stiffness, while the “damping Ziegler Paradox” influence can be regulated by the energy transmission efficiency between the stiffness related powers on each coordinate.

Vibration analysis of heated panels in supersonic airflow under aerothermoelastic effects has attracted wide attention, but the high-frequency vibration characteristics have not yet been thoroughly studied. A novel energy finite element method (EFEM), which can consider the aerothermoelastic effects, is proposed to investigate the high-frequency vibration response of heated panels in supersonic airflow. By separating the high-frequency response from the static or low-frequency aerothermoelastic response, a linearized high-frequency forced vibration equation is derived from the nonlinear motion equation of the panel. On this basis, the effects of in-plane and aerodynamic forces on the propagation properties of elastic waves in the panel are analyzed theoretically. By introducing these effects and the equivalent loss factors of elastic waves into EFEM theory, the energy density governing equation and the corresponding energy finite element model are established. Through numerical comparative studies, it is demonstrated that the proposed EFEM can well capture the aerothermoelastic effects and efficiently predict the high-frequency vibration response of heated panels in supersonic airflow. Particularly, the present EFEM is valid for panels in different aerothermoelastic states such as thermal buckling and flutter. In addition, the effects of aerothermoelasticity on the high-frequency vibration response are studied in detail, and the corresponding effect mechanism is revealed.

Civilian aircraft can experience noticeable vibrations in the cockpit and cabin due to mechanical faults during flight. To address this issue, a hybrid approach was utilized to investigate fluid-induced vibration load characteristics in the front landing gear compartment under different hatch opening angles. The results reveal that the root mean square (RMS) of cumulative pressure loads on both small and large hatches under different opening angles is largest at a 15°. For all the simulated cases (0°, 5°, 10°, 15°, 20°), the power spectral density (PSD) results of the chosen monitoring points on the inner wall of the large hatch exhibit broadband frequency characteristics, and the peak PSD values for the chosen monitoring points on the outer wall of the small hatch exhibit a significant concentration of energy at approximately 75 Hz. The peak PSD values for the selected monitoring points on the inner wall of the small hatch demonstrate a more uniform distribution of energy. Utilizing the iso-surface of Q-criterion, spatial streamlines, and streamlines at different cross-sections to analyze flow characteristics, the study investigates the fluctuating load mechanisms of the compartments. The results indicate that unsteady loads stem from the blunt edges of the hatches, which induce unsteady flow and spanwise flow. Geometric gaps between different locations cause flow separation, and the flows inside the compartment exhibit characteristics similar to those of a clean cavity. Furthermore, the mutual interference can be described using circulating flow and spanwise flow, resulting in flow unsteadiness. The flow separation zones enlarge and vortex intensity increases with the increase of the hatch opening angle from 0° to 15°; then, their values decrease as the hatch opening angle increases from 15° to 20°. These variations explain the maximum RMS of cumulative pressure loads at 15°.

In order to analyze the impact of component performance deviation on the transient performance of a low bypass ratio mixed turbofan engine, the acceleration and deceleration process of the engine are calculated based on a turbofan engine model, combined with the Monte Carlo method and component performance deviation model. The distribution rules of key parameters during engine transient process caused by component performance deviations are obtained, and the key parameters affected during the engine transient process are analyzed. The results show that during acceleration process, the surge margin of compressor should be noticed to prevent the surge due to component deviation, while the risk of over temperature and rich extinction should also be noticed. The conclusions can provide support for the design of engine control laws.

The interaction between an elastic structure and electrodynamic shakers commonly exists in Ground Flutter Simulation Tests (GFST) with multi-point excitations, causing a considerable discrepancy between the practical excitation forces and desired ones. To investigate the excitation force characteristics on a cantilever beam excited by a voltage-sourced electrodynamic shaker, the coupled shaker-beam system is modeled to derive the excitation force formula using Hamilton’s principle and Galerkin’s approach. Simulation results using the multi-mode beam model coupled with the shaker model are in good agreement with experimental results, verifying that the proposed multi-mode method can accurately predict the excitation force. Furthermore, parametric studies show that the influence of system parameters on the excitation force is related to the shaker's operating mode. Unlike in current mode of shaker, when the beam resonant frequency approaches the suspension frequency of shaker armature, the variation of excitation force amplitude in voltage mode is no longer minimal. Meanwhile, if the exciting point in the GFST is located far away from the modal node, it is essential to compensate the force because the accuracy of tests can be reduced dramatically. The coupled shaker-beam model proposed in this paper can provide the basis for compensation measures.

Multi-shaker vibration testing has gained increasing interest in recent years as researchers have demonstrated that complicated environmental response can be accurately replicated in the laboratory using multiple electrodynamic shakers. One of the crucial issues to be addressed is the discrepancy between the actual excitation forces and the required ones due to the coupling between the structure and multiple shakers. To analyze the excitation force characteristics, a coupled multi-shaker-cantilever beam system is studied and modeled using Hamiltonian principle and Galerkin's approach, and the excitation forces exerted on the beam are derived. Comparisons of experimental and numerical results indicate that the developed model can accurately predict the excitation forces under multi-point excitation. Furthermore, the relationship between the excitation force Frequency Response Function (FRF) matrix and the dynamic characteristics of the coupled system is revealed, showing that driving point FRF F-ii under multi-point excitation differs from that under single-point excitation. If the model parameters of two shakers are not perfectly identical, there will be an apparent discrepancy in amplitude-frequency characteristic between cross-point FRF F-ij and F-ji (i not equal j), while the phase-frequency characteristics are identical. Besides, when multiple shakers simultaneously drive the structure, the excitation forces are coupled due to structure vibration. Therefore, it is necessary to consider this coupling effect when performing the multi-point excitation for flexible structure. The coupled shaker-beam model proposed in this paper can provide the basis for designing decoupling controller.

In order to evaluate the influence of distributed propulsion, a calculate model of aircraft fuel consumption is established based on the model of propulsion system. The specific fuel consumption of distributed propulsion is calculated to survey the effects of transmission loss and bypass ratio. Fuel consumption is compared with that of a baseline engine to research the effects of additional weight and energy loss from electrical system. Improvement of aerodynamic performance and structure weight due to distributed propulsion are considered in the calculation of fuel consumption to analysis the combined effects of distributed propulsion. The transmission efficiency and power density of the electrical system required to reduce fuel consumption under different income levels are clarified. The results provide references for parameter optimization of distributed propulsion systems and aircraft design.

The aim of this study is to investigate the complex damping influences on the oscillatory/static instability characteristics of heated panels in supersonic airflow. Firstly, by assuming a constant, uniform thermal loading and adopting the piston theory, the panel aeroelastic governing equation is obtained. After deriving the panel buckling and vibration modes, the reduced order model can be built and adopted to investigate the system primary instability in the modal coordinates. Then, introducing the modal damping coefficients ratio eta > 0, the critical parameters of the panel flutter oscillation are theoretically evaluated based on the non-conservative energy balance principle, thus the system instability characteristics can be investigated. The results show that the system oscillatory and static buckling instability characteristics are significantly regulated by the thermal loading and modal damping. For the oscillatory instability, there exists the damping paradox, which is associated with the system energy dissipation efficiency, and can be quantitatively evaluated by the ratio 2 root eta/(1 + eta). The system static buckling instability characteristics is also affected by the modal damping, and this damping destabilization is clarified based on Hamiltonian energy conservation law. The results agree well with that obtained by Routh-Hurwitz criteria, and lead an in-depth understanding of the complex role played by the damping within the non-conservative dissipative systems.

A combined energy method is proposed to investigate the flutter instability characteristics of weakly damped panels in the supersonic airflow. Based on the small damping assumption, the motion governing partial differential equation (PDE) of the panel aeroelastic system, is built by adopting the first-order piston theory and von Karman large deflection plate theory. Then by applying the Galerkin procedure, the PDE is discretized into a set of coupled ordinary differential equations, and the system reduced order model (ROM) with two degrees of freedom is obtained. Considering that the panel aeroelastic system is non-conservative in the physical nature, and assuming that the panel exhibits a single period oscillation on the flutter occurrence, the non-conservative energy balance principle is applied to the linearized ROM within one single oscillation period. The obtained result shows that the ROM modal coordinate amplitudes ratio is regulated by the modal damping coefficients ratio, though each modal damping coefficient is small. Furthermore, as the total damping dissipation energy can be eliminated due to its smallness, the He's energy balance method is applied to the undamped ROM, therefore the critical non-dimensional dynamic pressure on the flutter instability occurrence, and the oscillation circular frequency amplitude relationship (linear and nonlinear form) are derived. In addition, the damping destabilization paradoxical influence on the system flutter instability is investigated. The accuracy and efficiency of the proposed method are validated by comparing the results with that obtained by using Routh Hurwitz criteria.

To investigate the nonlinear aeroservoelastic behaviors of a three-dimensional supersonic aircraft with control fin free-play, a modeling method based on the component-mode synthesis technique is utilized. The most distinctive feature of this method is that the structural nonlinearity can be expressed explicitly in the reduced-order aeroservoelastic model. The unsteady aerodynamic model with discrete gust loads is formulated using the supersonic lifting surface theory and the minimum state approximation. The results validate the feasibility of the reduced-order aeroelastic model for gust response analysis. A hybrid adaptive feedback control algorithm is proposed for the gust load alleviation by integrating the positive position feedback (PPF) and the filtered-x least-mean-square (FxLMS) algorithm. The comparative study demonstrates that the designed PPF-FxLMS algorithm has a better control performance for the alleviation of the pitch attitude angle induced by the gust loads, and it remains effective in the post-flutter regime. Moreover, the effects of free-play nonlinearity on the dynamic behaviors of the aeroservoelastic system are also studied, and numerical results show that the presence of free-play can lead to a larger peak value of gust response. The nonlinear vibration of the control fin induces the aircraft pitch attitude to produce stable limit cycle oscillations with high frequency, and variation in the free-play magnitude significantly influences the gust responses of the nonlinear aeroservoelastic system. (C) 2020 Elsevier Ltd. All rights reserved.

In the commented paper, the authors considered a two-dimensional airfoil with nonlinear stiffness and damping, and derived the amplitude modulation equation (AME) of the system limit cycle oscillation around its bifurcation. Based on the variation equation of the AME, they conducted the stability analysis. Unfortunately, the obtained variation equation associated with the AME is incorrect.

In this paper, an energy method is presented to study the coupled modes type panel flutter stability. The aeroelastic system continuous motion equation is built by adopting the first order piston theory aerodynamic loading. This continuous motion equation can be transformed by applying Galerkin method and constructed into a two-degree-of-freedom reduced order model (2-DOF ROM). Based on this ROM consisting of the first two structural modes, energy method is presented to investigate the system coupled modes type panel flutter stability. The obtained stability condition can be validated by comparing with its counterpart derived by Routh-Hurwitz criteria. Additionally, the critical system parameters can be evaluated and validated. Comparing with previous Ritz averaging method, the relationship between the modal coordinate amplitudes ratio and the modal damping coefficients ratio is believed to be firstly derived. The system damping paradoxical effect on coupled modes type panel flutter stability can be investigated after introducing the parameter, modal damping coefficients ratio. For given system modal damping coefficients, the phase difference and the modal coordinate amplitudes ratio of the first two modal coordinates can be specified. Based on the system power flow equations and the calculated parameters, the energy transfer characteristic between the supersonic airflow and the first two structural modes can be clarified. Such energy transfer procedure can be done within a half of oscillation period to sustain a neutrally stable single period oscillation.

An analysis of the fracture cause for a steam turbine blade root is conducted in this study. During the maintenance of a steam turbine, nineteen IP-6S (Intermediate Pressure 6th Stage) blades with root cracks were detected. The causes of blade cracking are analyzed through macroscopic inspection, magnetic particle testing, room temperature mechanical property testing, micro-hardness testing, chemical analysis, metallo-graphic analysis, scanning electron microscopy and energy spectrum analysis, as the blade static stress is simulated in this paper. The results showed that the chemical composition and mechanical properties of the blade leaves are qualified, and none abnormal corrosive substances are found in the crack initiation part. The numerical simulation result showed that the surface of the first tooth arc is the mostly stressed part of the root, and the stress level can be increased by a loosing installation. The main reason for the root cracking of the blade is that the gap between the blade root and the leaf groove becomes larger, which leads to the changes of the blade root contact mode, then the stress of the first tooth arc was further increased, finally the root and leaf groove wear is aggravated, and fatigue cracking occurs at the stressed first tooth.

This paper focuses on the nonlinear aeroelastic characteristics of a folding wing in the quasi-steady condition (namely at fixed folding angles) and during the morphing process. The structure model of the folding wing is formulated by the Lagrange equations, and the constraint equation is used to describe the morphing strategy. The aerodynamic influence coefficient matrices at several folding angles are calculated by the Doublet Lattice method, and described as rational functions in the Laplace domain by the rational function approximation, and then the Kriging agent model technique is adopted to interpolate the coefficient matrices of the rational functions, and the aerodynamics model of the folding wing during the morphing process is built. The aeroelastic responses of the folding wing with cubic stiffness are simulated, and the results show that the motion types of aeroelastic responses in the quasi-steady condition and during the morphing process are all sensitive to the initial condition and folding angle. During the morphing process, the transition of the motion types is observed. And apart from the period of transition, the aeroelastic response at some folding angles may exhibit different motion types, which can be found from the results in the quasi-steady condition.

This paper details the study of the aeroelastic effect on modal interaction and dynamic behavior of acoustically excited square metallic panels with fully clamped edges using finite element method. The first-order shear deformation plate theory and von Karman nonlinear strain–displacement relationships are employed to consider the structural geometric nonlinearity caused by large vibration deflections. Piston aerodynamic theory and Gaussian white noise are used to simulate the aerodynamic load and the acoustic load, respectively. Motion equations are derived by the principle of virtual work in the physical coordinates and then transformed into the truncated modal coordinates with reduced orders. Runge–Kutta method is employed to obtain the system response, and the modal interaction mechanism is quantitatively valued by the modal participation distribution. Results show that in the pre-/near-flutter regions, in addition to the dominant fundamental resonant mode, the first twin companion antisymmetric modes can be largely excited by the aeroelastic coupling mechanism; thus, aeroelastic modal participation distribution and the spectrum response can be altered, while the dynamic behavior still exhibits linear random vibrations. In the post-flutter region, the dominant flutter motion can be enriched by highly ordered odd order super-harmonic motion occurs due to 1:1 internal resonances. Correspondingly, the panel dynamic behavior changes from random vibration to highly ordered motions in the fashion of diffused limit-cycle oscillations (LCOs). However, this LCOs motion can be affected by the intensifying acoustic excitation through changing the aeroelastic modal interaction mechanism. Accompanied with these changes, the panel can experience various stochastic bifurcations.

Nonlinear aeroelastic behavior of a trapezoidal wing in hypersonic flow is investigated. The aeroelastic governing equations are built by von Karman large deformation theory and the third-order piston theory. The Rayleigh–Ritz approach combined with the affine transformation is formulated and employed to transform the equations of a trapezoidal wing structure, modeled as a cantilevered wing-like plate, into modal coordinates. And then the modal equations are solved by numerical integrations. Several typical cases are studied to validate the capability of the proposed method for linear and nonlinear aeroelastic analysis of trapezoidal cantilever plate in hypersonic flow. The effects of Rayleigh–Ritz mode truncation for various wing-plate geometrical characteristics, i.e., sweep angle of leading edge, taper ratio and span, are examined to determine the appropriate mode number for accurate modeling and fast calculation. Meanwhile, the effects of various geometries of trapezoidal cantilever plates on the flutter stability are investigated. The nonlinear dynamic behaviors of the model with three typical geometries, namely, the rectangular, parallelogram and trapezoidal wing-like plate, are simulated numerically. Furthermore, complex dynamic behaviors are observed and identified via the phase plot, the Poincare map and the largest Lyapunov exponent. The results demonstrate that geometrical parameters of trapezoidal wing have significant effects on the nonlinear aeroelastic behaviors of wing structure. In particular, the evolution processes of chaos exhibit remarkable difference for these three wing configurations.

A nonlinear flutter analysis of viscoelastic heated panels with aerodynamic loading exerting on its both surfaces is presented. The aeroelastic motion equations of such panels can be formulated by using the von Karman large deflection plate theory and piston aerodynamics theory, while the thermal induced membrane force and the Kelvin type viscoelastic damping are taken into account. By using Galerkin method, the continuous partial differential motion equation can be transformed into a set of nonlinear ordinary differential equations with coupled aerodynamic stiffness and aerodynamic/viscoelastic damping terms. By applying Routh-Hurwits criterion, the static divergence stability (buckling) boundary and the elastic/viscoelastic flutter stability boundaries of the panel initial flat equilibrium can be obtained. The obtained linear stability results revealed that the system dynamic bifurcation boundary can be significantly affected by the additive structural viscoelastic damping, and such effect can be enhanced by increasing the dynamic pressure of the external flow exerting on either single panel surface. Additionally, the sum of dynamic pressures exerting on both panel surfaces functions as the dynamic pressure exerting on either single pane surface. The corresponding nonlinear viscoelastic response can be simulated by using the fourth order Runge-Kutta numerical integration method, thus the system bifurcation diagrams with varying dynamic pressures can be obtained. The results revealed that the additive viscoelastic damping may exhibit the paradoxical effect on the system dynamic stability with lower temperature elevation, while the post flutter chaotic motions can be regulated as periodic motions with reduced amplitudes. However, with a higher temperature elevation, the effect of the additive viscoelastic damping can be always stabilizing to both the aeroelastic system stability and the post flutter chaotic motions.

高超声速飞行器壁板在非定常气动力、热载荷和噪声载荷构成的多物理场联合作用下,将表现出复杂的非线性气动弹性声振响应,特别是在颤振临界动压附近,受热载荷以及声载荷作用,壁板表现出复杂的跳变运动.基于von Karman大变形板理论,建立了热-声载荷和气动力共同作用下的壁板运动方程,分析了超声速气流中受热壁板的屈曲变形及热屈曲稳定性,借助势阱概念初步分析了壁板跳变运动产生的机理.通过定义“穿零频次”给出了跳变运动定量的分类方法,并计算得到不同温升和动压情况下,壁板发生跳变运动所对应的临界声压级.结果表明:在颤振临界动压之前,随着动压的增加,受热壁板势阱的深度先增大后减小,且受热壁板的势阱深度随着温升的增加而增大.