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In this paper, an adaptive fuzzy sliding mode controller utilizing a high-gain observer is proposed to address the aeroelastic instabilities of a novel nonlinear aircraft wing section under unsteady aerodynamics. The controller is designed to effectively mitigate high amplitude oscillations that arise in the system and enable aircraft to operate within an extended flight envelope. Additionally, the proposed controller utilizes output feedback to estimate the states and nonlinear dynamics of the model. The two-degree-of-freedom (2-DOF) nonlinear aeroelastic system describes the pitch and plunge motions of the wing section equipped with trailing and leading edge control surfaces. The resulting aeroelastic model, including the structural stiffness nonlinearities and the unsteady aerodynamic model, is based on Wagner's indicial function. The simulation results demonstrate the effectiveness of the suggested controller in suppressing the flutter phenomenon and improving the flight speed range. Further, the proposed controller accurately estimates the states of the model and successfully drives them to the origin despite the presence of disturbances and uncertainties. [ABSTRACT FROM AUTHOR]

Interactions between carbon nanotubes (CNTs) and fluid flows are central to the operation of several emerging nanotechnologies. In this paper, we explore the fluid-structure interaction of CNT micropillars in wall-bounded shear flows, relevant to recently developed microscale wall shear stress sensors. We monitor the deformation of CNT micropillars in channel flow as the flow rate and wall shear stress are gradually varied. We quantify how the micropillars bend at low wall shear stress, and then will commonly tilt abruptly from their base above a threshold wall shear stress, which is attributed to the lower density of the micropillars in this region. Some micropillars are observed to flutter rapidly between a vertical and horizontal position around this threshold wall shear stress, before settling to a tilted position as wall shear stress increases further. Tilted micropillars are found to kink sharply near their base, similar to the observed buckling near the base of CNT micropillars in compression. Upon reducing the flow rate, micropillars are found to fully recover from a near horizontal position to a near vertical position, even with repeated on–off cycling. At sufficiently high wall shear stress, the micropillars were found to detach at the catalyst particle-substrate interface. The mechanical response of CNT micropillars in airflow revealed by this study provides a basis for future development efforts and the accurate simulation of CNT micropillar wall shear stress sensors. [ABSTRACT FROM AUTHOR]

It is well known that the flutter performance of a section is sensitive to the changing wind angles of attack. Exploring the thin plate's flutter mechanism under different angles of attack is excellent, which helps understand inner flutter characteristics and ensures structural safety. This study investigated flutter derivatives of a thin plate with an aspect ratio of 40 under different wind angles of attack via the forced vibration technique. The otherness of aerodynamic damping and phase lag under different wind angles, which helps in understanding the flutter mechanism, are analyzed using the bimodal-coupled flutter method. It is shown that coupled vertical–torsional flutter dominated the flutter modality under 0° and 3° wind angles of attack. The critical flutter velocity dramatically decreased with increasing wind angles of attack which is attributed to the increasing negative aerodynamic damping induced by coupled self-excited forces and the decreasing positive aerodynamic damping induced by uncoupled self-excited forces. Moreover, the vertical motion lags behind the torsional motion under the 7° angle of attack, which was totally contrary to other angles of attack. Major works in this study reveal the aerodynamic mechanism of the weakened flutter performance of thin plates under large wind angles of attack and provide a reference for the flutter analysis of thin plates. [ABSTRACT FROM AUTHOR]

The self-sustained interactions between a flexible film and periodic vortices epitomize the spirit of fish swimming and flag flapping in nature, involving intricate patterns of flow–structure coupling. Here, we comprehensively investigate the multiple coupling states of a film in the cylinder wake mainly with experiments, complemented by theoretical solutions and nonlinear dynamical analyses. Four regimes of film motion states are identified in the parameter space spanned by the reduced velocity and the length ratio. These regimes are (i) keeping stationary, (ii) deflection flutter, (iii) hybrid flutter and (iv) periodic large-amplitude flapping, each governed by a distinct coupling mechanism, involving regular and irregular Kármán vortices, local instability of the elongated shear layers and 2P mode vortex shedding. The film futtering in regimes (ii) and (iii) is substantiated to be chaotic and bears a resemblance to the 'entraining state' of fish behind an obstacle in the river. The periodic flapping in regime (iv) manifests itself in an amalgam of standing and travelling waves, and has intrinsic relations to the 'Kármán gaiting' of fish in periodic vortices. With the spatiotemporal reconstruction for the periodic flapping, we procure the energy distributions on the film, revealing the energy transfer processes between the film and the large-scale vortices. The findings unequivocally indicate that the flow–structure interaction during the energy-release stage of the film is more intense than that during the energy-extraction stage. Given the similarities, the mathematical and physical methods presented in this work are also applicable to the research on biological undulatory locomotion. [ABSTRACT FROM AUTHOR]

This paper presents a novel framework for the active control of transonic airfoil flutter using synthetic jets through deep reinforcement learning (DRL). The research, conducted in a wide range of Mach numbers and flutter velocities, involves an elastically mounted airfoil with two degrees of freedom of pitching and plunging oscillations, subjected to transonic flow conditions at varying Mach numbers. Synthetic jets with zero-mass flux are strategically placed on the airfoil's upper and lower surfaces. This fluid–structure interaction (FSI) problem is treated as the learning environment and is addressed by using the arbitrary Lagrangian–Eulerian lattice Boltzmann flux solver (ALE-LBFS) coupled with a structural solver on dynamic meshes. DRL strategies with proximal policy optimization agents are introduced and trained, based on the velocities probed around the airfoil and the dynamic responses of the structure. The results demonstrate that the pitching and plunging motions of the airfoil in the limited cycle oscillation (LCO) can be effectively alleviated across an extended range of Mach numbers and critical flutter velocities beyond the initial training conditions for control onset. Furthermore, the aerodynamic performance of the airfoil is also enhanced, with an increase in lift coefficient and a reduction in drag coefficient. Even in previously unseen environments with higher flutter velocities, the present strategy is achievable satisfactory control results, including an extended flutter boundary and a reduction in the transonic dip phenomenon. This work underscores the potential of DRL in addressing complex flow control challenges and highlights its potential to expedite the application of DRL in transonic flutter control for aeronautical applications. [ABSTRACT FROM AUTHOR]

In order to improve the aerodynamic characteristics of the drone wings, CFD method was used to simulate and calculate the lift coefficient, drag coefficient, and lift-drag ratio under different relative inflow velocities, as well as the velocity and pressure fields under different attack angles. Modal calculations were conducted on the wing to obtain the first four modal shapes, providing a basis for analyzing flutter characteristics. An iterative calculation method of incompressible potential flow-boundary layer based on surface element method was combined with the software XFOIL to optimize the airfoil at low wind speeds. The results indicate that the airfoil is susceptible to stall at high angles of attack, with the pressure of the separation flow being nearly equivalent to that at the separation point. Subsequent to separation, there is an increase in differential pressure resistance, resulting in a marked rise in the drag coefficient. At the optimized angle of attack, the lift-drag ratio of the optimized wing increases by 12.58 %, while there is a decrease of 0.084 % in lift coefficient and an increase of 11.21 % in drag coefficient. [ABSTRACT FROM AUTHOR]

The violent vibration of supersonic wings threatens aircraft safety. This paper proposes the strongly nonlinear acoustic metamaterial (NAM) method to mitigate aeroelastic vibration in supersonic wing plates. We employ the cantilever plate to simulate the practical behavior of a wing. An aeroelastic vibration model of the NAM cantilever plate is established based on the mode superposition method and a modified third-order piston theory. The aerodynamic properties are systematically studied using both the timedomain integration and frequency-domain harmonic balance methods. While presenting the flutter and post-flutter behaviors of the NAM wing, we emphasize more on the pre-flutter broadband vibration that is prevalent in aircraft. The results show that the NAM method can reduce the low-frequency and broadband pre-flutter steady vibration by 50%–90%, while the post-flutter vibration is reduced by over 95%, and the critical flutter velocity is also slightly delayed. As clarified, the significant reduction arises from the bandgap, chaotic band, and nonlinear resonances of the NAM plate. The reduction effect is robust across a broad range of parameters, with optimal performance achieved with only 10% attached mass. This work offers a novel approach for reducing aeroelastic vibration in aircraft, and it expands the study of nonlinear acoustic/elastic metamaterials. [ABSTRACT FROM AUTHOR]

This paper presents initial findings from aeroelastic studies conducted on a wing-propeller model, aimed at evaluating the impact of aerodynamic interactions on wing flutter mechanisms and overall aeroelastic performance. The flutter onset is assessed using a frequency-domain method. Mid-fidelity tools based on the time-domain approach are then exploited to account for the complex aerodynamic interaction between the propeller and the wing. Specifically, the open-source software DUST and MBDyn are leveraged for this purpose. The investigation covers both windmilling and thrusting conditions. During the trim process, adjustments to the collective pitch of the blades are made to ensure consistency across operational points. Time histories are then analyzed to pinpoint flutter onset, and corresponding frequencies and damping ratios are identified. The results reveal a marginal destabilizing effect of aerodynamic interaction on flutter speed, approximately 5%. Notably, the thrusting condition demonstrates a greater destabilizing influence compared to the windmilling case. These comprehensive findings enhance the understanding of the aerodynamic behavior of such systems and offer valuable insights for early design predictions and the development of streamlined models for future endeavors. [ABSTRACT FROM AUTHOR]

In this study, stall flutter onset prediction in a transonic compressor is carried out using the (uncoupled) energy method with Fourier transform. As the study is conducted computationally using computational fluid dynamics (CFD)-based simulations, the energy method was employed due to its higher computational efficiency by implementing the one-way FSI (Fluid Structure Interaction) model. The energy method is relatively uncommon for determining the aerodynamic damping and flutter prediction, specifically in blade stall conditions for the 3D blade passages. The NASA Rotor 67 was chosen for the validation of the study due to the availability of a wide range of experimental data. A flutter prediction analysis was performed computationally using CFD for the two-blade passages of the rotor in the peak efficiency and stall regions. Prior to this, the modal analysis on the prestressed blade was conducted, considering the centrifugal effects. The modal analysis provided accurate blade frequency and amplitude, which were the inputs of the flutter analysis. The first three modes of blade resonance were studied with a range of nodal diameters within near-peak efficiency and stall regions. The energy method implemented in this study for the flutter analysis was successfully able to predict the aerodynamic damping coefficients of the first three modes for a range of nodal diameters from the periodic-unsteady solution of the defined blade oscillation within the regions of interest (peak efficiency and stall point). The results of the study confirm the rotor blade's stability within the near-peak region and, most importantly, the prediction of the flutter onset in the stall region. The study concluded that the computationally inexpensive and time-efficient energy method is capable of predicting the stall flutter onset. In the future, further validations of the energy method and investigations related to flow mechanism of stall flutter onset are planned. [ABSTRACT FROM AUTHOR]

In this study, we conducted an aeroelastic stability analysis of a laminated composite magnetorheological fluid (MRF) sandwich panel in supersonic airflow for varying yawed angles. The aeroelastic equations for a rectangular sandwich panel (MRF core layer and composite cross‐ply laminate constraining and host layers) were established using a MIN3 plate element. Aerodynamic forces for different yawed angles would result in different coupled flutter boundaries of the panel. The first‐order piston theory with a flow‐yawed angle was employed. The flutter dynamic pressure was obtained through eigenvalue analysis. The effects of various parameters such as the magnetic field intensity, MRF core and constraining layer thicknesses, ply orientation, and yawed flow angle on the flutter dynamic pressure were studied for the simple‐ and fixed‐support boundary conditions. Our results demonstrated that the flutter dynamic pressure of the laminated composite MRF sandwich panel (i) increased for increasing magnetic field intensity and constraining layer thickness; (ii) initially decreased and then increased with the increasing MRF core thickness; and (iii) was strongly influenced by the ply orientation and yawed flow angle. [ABSTRACT FROM AUTHOR]

Numerical simulations of a T-tail flutter in subsonic and transonic flow regimes were conducted using a 3D Navier-Stokes code which considers the in-plane motion of the horizontal tail plane (HTP). In our subsonic flutter simulations, we confirmed that the flutter velocity of the T-tail is dependent upon the mean angle of attack of the HTP, and also confirmed that the results show a satisfactory agreement with those predicted using a lifting-surface theory. Thereby we were able to verify the reliability and accuracy of the 3D Navier-Stokes code. In transonic simulations, the sharp transonic-dip phenomena of the dynamic pressure of flutter and limit cycle oscillation (LCO) were predicted. It can be shown that the characteristics of LCO and flutter are considerably different depending on Mach number and the angle of attack of the HTP. As a possible cause of these phenomena, we examined the behavior of the shock wave generated on the upper or lower surface of the HTP, which varies strongly depending on the Mach number and the angle of attack of the HTP. [ABSTRACT FROM AUTHOR]

Stall flutter is a serious threat to the operational integrity in turbomachinery, particularly in the final stage rotors of steam turbines and in compressors. Although computer science has developed rapidly and much of the research can be carried out using numerical tools, the simulation of some phenomena, such as stall flutter, is still very challenging and needs to be supported by experimental data. This paper presents an innovative experimental linear blade cascade design with five prismatic blades with pitch degrees of freedom, designed to be operated in a low subsonic wind tunnel. The geometry of the blade cascade was chosen on the basis of the experimental and numerical tests to allow stall flutter initiation. New suspension, measurement and electromagnetic excitation systems were developed and experimentally tested to allow accurate measurement of aerodynamic damping during controlled flutter tests. The novelty of the experimental blade cascade is the possibility of single pulse excitation of the blades. The cascade can be brought to the edge of stability by adjusting the angle of attack and flow velocity, and then the pulse can be used to induce stall flutter. Measurement of both mechanical and flow characteristics, also demonstrated in this paper, will provide data for in-depth analysis of stall flutter initiation and propagation. [ABSTRACT FROM AUTHOR]

In order to reveal the influence of initial excitation on the bifurcation phenomenon of bridge decks, a new perspective of flow characteristics is developed based on the computational fluid dynamics numerical simulation method. Then, the bifurcation mechanism of vortex-induced vibration (VIV) response and nonlinear flutter response of the H-shaped section is investigated. The results show that when the wind speed is 2 m/s, under a small torsional excitation of 0.5°, the flow field of the H-shaped section will develop into the vortex shedding mode of the vertical vibration, resulting in vertical VIV. However, while under a large excitation of 6°, the flow field will directly transform into the vortex shedding mode of the torsional vibration, resulting in torsional VIV. Therefore, the bifurcation phenomenon of the VIV response is observed. When the wind speed is 4 m/s, the H-shaped section exhibits a nonlinear flutter limit cycle oscillation under a large excitation of 8°, but its response can be ignored under a small excitation of 0.5°. This phenomenon is attributed to the significant change in the transition of the vortex shedding mode from a small amplitude to a stable large amplitude, and the flow field lacks enough energy to complete the transition of the vortex shedding mode, resulting in the bifurcation phenomenon of the nonlinear flutter response. When the wind speed is 3.0 m/s, the large excitation will change the vortex shedding frequency of the new H-shaped section, resulting in the torsional VIV. [ABSTRACT FROM AUTHOR]

The use of experimental data is a relevant topic in aeroelasticity. A new aircraft prototype design and certification involve both flutter analysis and tests. The latter is essential to assess a proper finite element (FE) model for the aeroelastic analysis. The current work perfectly fits in this framework, presenting a novel methodology exploiting data from ground vibration testing (GVT) for flutter calculations. The first application relies on the only experimental modal parameters employed to perform aerodynamic flutter calculation. This method does not necessitate a consistent structural model, allowing fast and reliable computation of the flutter clearance just based on GVT results. This application is especially valuable when it is not possible for a structural dynamic model development, e.g., in the absence of a sufficient amount of technical data. The method also allows combinations of experimental and numerical mode shapes to perform flutter calculations. All or some modal parameters can be incorporated in the standard numerical flutter model, adjusting the computed numerical modes, or adding/deleting a few. The method allows a real true correlation between the GVT results and the numerical model, which is usually very difficult to obtain. This work represents the first case based on a combination of experimental and numerical results, paving the way to various possible levels of hybridization to perform aeroelastic analyses. [ABSTRACT FROM AUTHOR]

One of the novel applications of nano-scale structure is in the field of sensors in supersonic flow conditions. In this regard, a cylindrical nano-scale structure made of two-dimensional functionally graded (2D-FG) material in two-phase supersonic flow condition is considered in this study to explore the state of instability and flutter. It is obvious from the problem definition that several aspects from different fields have to be incorporated into the formulation of the problem. First, the small scale of the structure enforced the utilization of size-dependent theories of elasticity. Here, the nonlocal strain gradient theory is utilized to relate the stress and strain fields together. Moreover, the strain field of the cylindrical structure is computed from assumed displacement fields from higher-order shear deformation theory (HSDT). On the other hand, the equations of motion and boundary conditions are obtained using Hamilton's principle considering the variational of all the work and energy components. The external energy of the system comes from the pressure loading from the aerodynamic supersonic flow. The validation and convergence examination of the presented method is conducted and a detailed parametric study is performed to find the most important parameter affecting the stability of the nanopipe structure. It is revealed that the Mach number and gas volume fraction of two-phase supersonic flow is extremely influential in determining flutter in the nanopipes. [ABSTRACT FROM AUTHOR]

The aerothermoelastic characteristics of curved fiber composite laminated panels subjected to supersonic airflow are proposed. Based on the refined zig-zag theory, the refined zig-zag theory is adopted to describe the panel and the quasi-steady first-order piston theory is utilized for calculating the aerodynamic force. With considering uniformly distributed temperatures as thermal loads throughout the panel thickness, a finite element method is employed to solve the discretization equation of aerothermoelastic motion using the complex mode method. The validity of this aerothermoelastic model is substantiated through a meticulous comparison of computed results with existing solutions documented in the literature. The maximum error of the present critical flutter dynamic pressure of the curvilinear fiber composite laminates is 0.16% comparing with the literature. Furthermore, different plate theories are investigated to explore aerothermoelastic stability, including flat, critical buckling, and limit cycle oscillation for curvilinear fiber composite laminated panels. Detailed discussions are provided on how curvilinear fiber angle and temperature rise influence natural frequencies and flutter characteristics of composite laminates. [ABSTRACT FROM AUTHOR]

This paper studies the instability of a non-ideal pipe system conveying fluid. Unlike ideal modeling, three types of possible non-ideal factors have been considered in the present analysis. The pipe has a crack and is subjected to boundary relaxation, and the fluid in the pipe is viscous. It is an interesting instability problem of a fluid-structure system with multiple forms of imperfections. It gains motivation from our previous study and aims at further essential results for comprehensively understanding the nature of such a problem. In order to incorporate multiple imperfections, a new data-driven approach is developed to model the non-ideal pipe-fluid system. Then the influences of the support stiffness, fluid–solid mass ratio, crack, and fluid viscosity on the pipe instability are systematically discussed. Results show that the pipe exhibits flutter-divergence conversion at low mass ratios and flutter-mode transition at high mass ratios. Due to the two types of instability transitions, the critical flow speed could experience multiple sudden jumps as the support stiffness changes. With the imperfections of crack and viscous flow, the pipe's instability behavior tends to be more complicated, but the inducing mechanisms for the sudden changes are straightforward. They are instability-type conversion, instability-mode transition, and laminar-turbulent flow transition. [ABSTRACT FROM AUTHOR]

Understanding settling motion of coral grains is important in terms of protection of coral reef systems and resotoration of the associated ecosystems. In this paper, a series of laboratory experiments was conducted to investigate the settling motion, using optical microscopy to measure shape parameters of coral grains and the particle-filtering-based object tracking to reconstruct the three-dimensional trajectory. Three characteristic descent regimes, namely, tumbling, chaotic and fluttering, are classified based on the three-dimensional trajectory, the spiral radius variation and the velocity spectrum. It is demonstrated that if one randomly picks up one coral grain, then the probabilities of occurrence of the three regimes are approximately $26\,\%$ , $42\,\%$ and $32\,\%$ , respectively. We have shown that first, the dimensionless settling velocity generally increases with the non-dimensional diameter and Corey shape factor and second, the drag coefficient generally decreases with the Reynolds number and Corey shape factor. Based on this, the applicability of existing models on predicting settling velocity and drag coefficient for coral grains is demonstrated further. Finally, we have proposed extended models for predicting the settling velocity. This study contributes to better understanding of settling motion and improves our predictive capacity of settling velocity for coral grains with complex geometry. [ABSTRACT FROM AUTHOR]

This paper introduces a novel numerical method for investigating the dynamic stability of a flutter panel exposed to a supersonic gas flow and a fluctuating axial excitation force. Initially, the system equation of motion was derived using Lagrange’s equation, where the first two modes are coupled Mathieu–Hill equations with damping, constituting a system of linear second-order differential equations with periodically variable coefficients. Subsequently, a new numerical method was proposed to analyze the dynamic stability of coupled Mathieu–Hill equations with damping. This method involves breaking down an arbitrary parametric load into discrete segments to approximate the variable excitation function using a step function. The system responses of each segment are then accumulated in matrix form. The proposed numerical method proves particularly effective for dynamic systems whose parameters cannot be treated as small. In practical application, the method allows the construction of instability regions corresponding to natural frequencies, subharmonics, and combination frequencies. Dynamic stability diagrams were generated based on dynamic pressure ratio, air/panel density ratio, Mach number, panel thickness—length ratio, and excitation frequency. The results demonstrated general agreement with those obtained through Hsu’s perturbation method, however, our numerical results have proven more accurate. The paper concludes by offering suggestions for suppressing panel flutter through appropriate parameter combinations. [ABSTRACT FROM AUTHOR]

Modal parameter estimation of turbulence response is an important aspect of flutter test data processing. Using the modal parameter estimation results of turbulence response and the damping extrapolation curve, the flutter boundary of aircrafts can be predicted. However, owing to the randomness of atmospheric turbulence excitation, modal parameter estimation of the turbulence response has a challenge. This study analyses the turbulence response and calculates the corresponding impulse response using the seq-to-seq self-attention generative network. The encoder performs feature compression of the turbulence response, whereas the multi-head self-attention structure in the middle layer extracts the time series features. Up-sampling is performed based on the decoder to obtain the impulse response, and modal parameter estimation of the turbulence response is achieved. The self-attention generative model is validated using simulation data and the turbulence response is obtained by the wind tunnel test and the flutter flight test. The results demonstrate the feasibility and engineering applicability of the seq-to-seq generative model. [ABSTRACT FROM AUTHOR]

The consideration of aeroelasticity is essential during the development phase of axial flow compressors, and its implications should be factored into the operational planning of gas turbines. While synchronous vibration can typically be mitigated during the blade design stage, the complete avoidance of non-synchronous vibration remains a challenge, prompting ongoing research efforts for predictive solutions. The study utilized an industrial gas turbine axial compressor with 1.5-stage blades for aerodynamic performance and flutter assessments. The calculated results were comprehensively compared with those of 1.5-stage scaled rig test. The comparative analysis demonstrated a strong alignment between predictions regarding aerodynamic performance and the presence or absence of flutter. Furthermore, unsteady flutter calculations were conducted for cases both with and without flutter, allowing for a detailed analysis of the factors contributing to flutter occurrence. Through this investigation, the study established methodologies for aeroelastic design and evaluation, along with a proposed approach for preventing flutter generation. [ABSTRACT FROM AUTHOR]

In this study, the dynamical behavior of spinning pipes conveying fluid of axially variable density is investigated. First, based on Hamilton's principle, the coupled governing equations for flexural vibration of the pipe system are derived. Then, the motion equations are truncated by using the Galerkin method. As a result, the discretized motion equations as well as the eigenfrequency equations of the system are obtained. The natural frequencies, divergence, and flutter instability thresholds of the fluid–structure interaction system are acquired by computing the complex frequencies in the first two modes of the system. Also, a comparative study is conducted to validate the accuracy of the present model and solution approach. Finally, the effects of main parameters, such as spinning velocity, flow velocity, mass ratio, and fluid density gradient parameter, on the vibration and stability of the pipe system are evaluated. The results show that the stability of the pipe system is dominated by the mass ratio and the fluid density gradient parameters, while the spinning velocity mainly affects the natural frequency of the system. [ABSTRACT FROM AUTHOR]

Stall flutter is a large-amplitude, flow-induced limit cycle oscillation that poses a risk of structural failure of wings. In this paper, a numerical investigation is conducted to explore the mitigation of stall flutter using active surface morphing with a single-mode harmonic motion at a Reynolds number of 1000. The effects of oscillation position and phase on mitigation are evaluated based on the net energy gain of the airfoil over one oscillation cycle. The results show the effectiveness of active surface morphing in mitigating stall flutter, achieving a maximum of 46% reduction in energy obtained by the airfoil from the flow. The control effect is primarily attributed to the decrease in the aerodynamic moment peak during the pitch-up phase and the increase in the aerodynamic moment trough during the pitch-down phase. The former is essentially due to the large-scale clockwise vortex induced by the skin bulging in the pitch-up phase. The latter is fundamentally caused by the weakening of the trailing-edge vortex and the strengthening of the second leading-edge vortex. Finally, maps depicting the energy exchange between the flow and the airfoil system are presented to further demonstrate the control effectiveness of active surface morphing. [ABSTRACT FROM AUTHOR]

Sandwich structures with lattice cores are novel, lightweight composite structures and are widely used in the aerospace industry. Besides, the aeroelastic behavior of sandwich panels in a supersonic flow regime still needs to be thoroughly studied. This work investigates the supersonic flutter of a sandwich panel whose core is topology-optimized. A finite element model of a sandwich panel based on the layerwise theory, coupled with the first-order piston theory, is presented. The sandwich panel core is assessed using a topology optimization approach with flutter loading constraints. The subsequent analytical homogenization scheme is developed to provide the equivalent mechanical properties of the topology-optimized panel. The modeling approach is fully validated, and the results demonstrate that the sandwich panel is capable of enlarging the flutter-free operational flight range when compared with other conventional panel designs. A parametric analysis of the topology-optimized sandwich panel regarding the critical flutter conditions is performed. [ABSTRACT FROM AUTHOR]

In this paper, the vibration and stability of spinning viscoelastic tapered shafts with axial motion are investigated. The model of the shaft system is developed by incorporating the spinning motion, axial motion, axial load, viscoelasticity, and geometric properties. The coupled governing equations of motion of the spinning shaft system are obtained by applying Hamilton's principle. By employing the Laplace transform and Galerkin discretization method, the natural frequency and modal damping are computed. Also, the critical divergence axial and spinning velocities, as well as the flutter axial velocity, are determined. As a result, the divergence and flutter instability conditions of the system are examined. The effects of the main parameters, such as taper ratio, axial load, and viscoelasticity of material, on the vibration behavior of the shaft system are evaluated. The results show that compared to the homogeneous spinning shaft system with uniform cross-section, the vibration and stability of the viscoelastic tapered shaft system undergo an essential evolution. It is further demonstrated that the axial load and viscoelasticity are also the key parameters governing the stability of the system. [ABSTRACT FROM AUTHOR]

For the doctoral study on the influence of the aircraft external stores over the aeroelasticity behavior of the IAR-99 HAWK, we have developed a new Finite Elements Model (FEM) with new stateof-the-art methods and technologies available at the end of 2020. After creating a new Finite Elements Model (FEM) of the aircraft IAR-99 HAWK using special elements, the team decided to validate the results by performing a free-free vibrations analysis on this model. In this presentation we have described the finite elements used and their peculiarities, the finite elements model obtained for the aircraft IAR-99 HAWK and the results of the free-free vibrations analysis obtained in the empty equipped configuration. The results of the theoretical free-free vibrations obtained for the IAR-99 HAWK in the empty equipped configuration are presented and compared to those obtained from ground tests, thus confirming the accuracy of the new Finite Element Model (FEM). The new Finite Element Model (FEM) of the IAR-99 HAWK will be versatile enough to provide insight into the structural behavior of the aircraft, regardless of the phenomenon for which it is analyzed: static stress, aerodynamic simulations, flutter, etc. [ABSTRACT FROM AUTHOR]

A promising mathematical model is considered for studying the aeroelastic stability of sandwich shells supported by an internal hollow elastic cylinder and annular ribs. Using the relations derived, the influence of thickness of the inner hollow cylinder, the length of the shell, and the shear modulus of core on the critical air flow speed leading to the occurrence of flutter was studied. Based on the calculations performed, recommendations were developed for choosing parameters of the sandwich composite structure considered. The mathematical model presented makes it possible to expand the range of topical scientific and applied problems that can be solved in the field of aeroelastic stability of sandwich shells. [ABSTRACT FROM AUTHOR]

Bioroughness plays an important role in modifying the velocity and sediment flux near the riverbed. It is therefore pertinent to study the influence of benthic fauna on the bed forms. To this end, large-eddy simulations are performed to investigate the influence of the arrays of mounds and their density on the turbulence characteristics in an open-channel flow. The simulated distributions of the time-averaged streamwise velocity and the turbulence intensity are in good agreement with the experimental data. Four numerical simulations are performed with varying streamwise spacings of mounds. Details of the time-averaged and instantaneous flow velocities are analyzed by multiple visualization methods, and the effects of the bioroughness density on the equivalent roughness height and the Darcy–Weisbach friction factor are quantified. The time-averaged flow in the wake of the mounds is characterized by a symmetric pair of vortices. The mounds behave like bluff bodies, increasing the riverbed roughness and heterogeneity in the flow environment. An increase in mound density is to promote the development of secondary currents and to increase the dispersive stress near the bed. The peaks of the Reynolds shear stress distributions decrease in both the streamwise and vertical directions for the high-density case due to a blockage effect. The instantaneous flow features, in the form of various turbulence structures, are generated near the top edge and the wake zone of mounds. The spacing between low-speed streaks decreases with an increase in equivalent roughness height. Multifrequency behavior that is observed is a result of shear layer roll-up from the edges of mounds and the flapping of wake. Finally, two formulas for equivalent roughness height and Darcy–Weisbach friction factor are proposed involving the bioroughness density and height. The findings demonstrate the effects of the bioroughness on the near-bed turbulence characteristics and sediment stability. Practical Applications: Bed forms produced by aquatic benthos are usually of small scale but with a large abundance. We performed a flume experiment as well as large-eddy simulations (a high-fidelity numerical simulation approach) to answer whether these microtopographies are important when studying flow and transport processes near the riverbeds. Four numerical simulations are performed with varying mound density. It was found that the microtopographies created by benthos indeed change the near-bed turbulence characteristics and sediment stability. Specifically at high abundances, benthos-induced bed forms could raise the bed roughness and the friction factor eightfold and threefold, respectively. This suggests the necessity of a modified roughness height formula and a friction coefficient formula based on the benthos abundance, which has been carried out in this study. Thus, engineers who are concerned with the flow and sediment transport in habitats of high-abundance benthos (e.g., wetlands and estuaries) may prefer to use these modified formulas and to develop a more accurate model of sediment transport. [ABSTRACT FROM AUTHOR]

The aeroelastic stability of the tail is significantly challenged by flutter instability. Skin and spars strongly affect flutter speed due to their torsional and bending stiffness, respectively. C-section spars are primarily utilized in composite structures due to their straightforward manufacturing process. This research aims to investigate the impact of the position and orientation of a laminated composite C-spar on the flutter speed of the airfoil section, utilizing a two-degree-of-freedom flutter method. The position of the C-spar varies between 10% and 50% of the chord length from the leading edge of the airfoil section, while the orientation of the C-spar with respect to the leading edge or trailing edge is also examined. To ensure comparability, the elastic section modulus and mass of the composite spar are maintained nearly constant. When it comes to the structural design process, one of the key challenges is determining the flutter and divergence speeds. In a novel approach, Finite Element Method (FEM) is utilized to calculate the torsional and bending stiffness values. This method provides a more accurate and efficient way to evaluate these important parameters. The results indicate that the location design of the C-spar exerts a more substantial influence on the flutter speed than the orientation of the spar. Furthermore, it is crucial to consider the nonlinear effects of the spar's position and direction in comprehending the aeroelastic instability of the aircraft tail. Additionally, the study found that the addition of a spar to a hollow section of the V-tail does not significantly enhance aeroelastic behavior. Only a modest increase of approximately 20% in flutter speed was observed. The primary effect of the spar lies in the bending stiffness, which does not lead to a substantial increase in flutter speed. Moreover, while flutter occurs before divergence, there can be a considerable distance between the respective speeds. Moving the spar from the leading edge to the mid-chord can reduce this margin, potentially compromising stability. Results show when the C-par position is close to the center of the airfoil, the flutter and divergence speed increase. [ABSTRACT FROM AUTHOR]

Active aerodynamic flaps can effectively improve the aerodynamic stability of bridges; however, the determination of optimal control parameters often requires a large number of experiments. This study proposes a method for determining the optimal control parameters of active flaps based on the surrogate model and computational fluid dynamics (CFD) simulation technology. The computational fluid dynamics method is used to calculate the flutter derivatives of the deck–flap system under a small number of control parameters, and then the surrogate model is used to extend the flutter derivatives to the design domain of entire control parameters. Furthermore, the changing trend of flutter performance under flap control parameters can be obtained. This method can avoid complex and expensive active control wind tunnel tests and obtain more accurate optimal control parameters at less cost, which is helpful for the convenient application of active control technology in actual engineering. [ABSTRACT FROM AUTHOR]

As the size of wind turbine blades increases, the flexibility of the blades increases. In actual operation, airflow flow can cause aerodynamic elastic instability of the blade structure. Long blades may experience coupled mode flutter due to the bending torsion coupling effect, leading to blade failure. Based on Euler Bernoulli beam theory combined with Theodorsen non directional aerodynamic loads, a blade flutter characteristic equation is established through finite element method. Taking NREL 5 MW wind turbine blades as an example, analyze the influence of parameter changes in different regions of the blades on flutter characteristics. Research has found that paramter changes in the tip region of blade have the greatest impact on flutter characteristics. The vibration frequency shows an overall upward trend with the increase of waving stiffness and torsional stiffness. The flutter velocity of the three regions tends to stabilize as the bending stiffness decreases. The blade flutter speed increases with the increase of torsional stiffness. The radius of gyration is inversely proportional to the flutter frequency and flutter velocity. The impact of centroid offset on blade structure flutter frequency is minimal, but the centroid offset in the tip region has a greater impact on flutter velocity. Increasing the torsional frequency can prevent coupled mode flutter and provide a theoretical basis for blade flutter prevention design. [ABSTRACT FROM AUTHOR]

In this paper, the flutter analysis is studied for a cylindrical sandwich shell subjected to external supersonic fluid flow. The sandwich shell consists of a re-entrant auxetic honeycomb (AH) core made of functionally graded materials (FGMs) which is covered with two polymeric face layers enriched with either carbon nanotubes (CNTs), graphene nanoplatelets (GNPs), or graphene oxide powders (GOPs). The volume fraction (percentage) of the ceramic phase in the functionally graded auxetic honeycomb (FGAH) core varies from zero at the inner surface to one at the outer one based on either power-law function (P-FGM), exponential function (E-FGM), or sigmoid function (S-FGM). It is assumed that the nanofillers are distributed uniformly inside the face layers. The first-order shear deformation theory (FSDT) and the piston theory are employed to provide the mathematical models of the shell and the aerodynamic pressure, respectively. The governing equations and boundary conditions are derived via Hamilton's principle. An exact solution is performed in the circumferential direction via harmonic trigonometric functions (sine and cosine) and an approximate solution is performed in the axial direction utilizing the differential quadrature method (DQM). The natural frequencies and corresponding damping ratios are attained through the presented semi-analytical solution and the impacts of several factors on the natural frequencies and the critical aerodynamic pressure of the shell are examined. These factors the type of the FGM, the material index, and inclined angle of the cells in the FGAH core, the thickness of the FGAH core, the type and mass fraction of the nanofillers in the nanocomposite face layers, and boundary conditions. [ABSTRACT FROM AUTHOR]

• The flutter behavior of variable angle tow composite laminated plates in supersonic flow is theoretically studied. • Finite element simulations are conducted and successfully validate the developed theoretical model. • The effect of elastic foundations and laminate layouts on the flutter behavior of the composite plate are revealed. • This work can provide a reference for the dynamic design of the VAT composite laminates in the aerospace engineering. A theoretical model is developed to investigate the aerothermoelastic flutter behavior of variable angle tow (VAT) composite laminated plates on elastic foundation in supersonic flow with general boundary conditions. The equations of motion of the VAT composite laminated plate are established by adopting the first-order shear deformation theory and the supersonic piston theory. The composite laminated plate is considered to be lying on an elastic foundation (such as a thermal protection shield) on one side and exposed to supersonic airflow on the other side. The displacement components of the structure are expanded into Chebyshev polynomials multiplied by the boundary characteristic functions to satisfy the geometric or natural boundary conditions. To validate the proposed theoretical model, a finite element (FE) model is also established by ABAQUS, and the numerical results fit well with that of the developed theoretical model. The effects of the boundary conditions, thermal loads, elastic foundations and laminate layouts on the flutter characteristics of the VAT composite plate are revealed and discussed in detail. This work can provide a reference for the dynamic design of the VAT composite laminates in the aerospace engineering. [ABSTRACT FROM AUTHOR]

The paper aims to perform numerical simulations of flapping flight using Discrete Vortex Method (DVM) scheme. The scheme is suitable for moderately low Reynolds number (< ∼ 1 0 5) and computationally less expensive as the flow domain doesn't need to be discretized at each time step. Flapping of a one-dimensional (1D) flexible filament in a two-dimensional (2D) inviscid flow is simulated. The effect of bending by varying the wing shape along the spanwise length on aerodynamic performance was studied. It is observed that compared to rigid wings, bending wings are found to be better in generating lift. The effect of various bending wing configurations (F 1 , F 2 , F 3 and F 4) and different methods of imposing the bending (W 1 , W 2 and W 3) is studied. It was demonstrated that applying wing bending only during the downstroke phase (W 2) is more effective than imposing bending throughout the flapping cycle (W 1). Moreover, an effort is made to replicate the bending configuration observed in manta ray fish (W 3) to investigate its impact on flow domain characteristics, lift, and thrust forces. Furthermore, the inclusion of a winglet is found to significantly enhance lift generation. In addition to the study of bending effects on aerodynamic performance, the study also seeks to emulate a unique aspect of bat flight kinematics, specifically the dynamic variation in wing span length during flapping. In a comparative analysis of two span length variation strategies, it is discerned that exclusively varying span length during the upstroke phase is the optimal approach for achieving increased lift generation. The study highlights the crucial role of wing bending and span length modulation in achieving elevated lift forces while simultaneously reducing drag. These findings are seen as holding significant promise for the design and optimization of Micro Air Vehicles (MAVs) utilizing flapping-based lift generation mechanisms, contributing substantially to the identification of optimal parameters for enhancing MAVs' aerodynamic performance and operational efficiency. [ABSTRACT FROM AUTHOR]

The paper investigates the phenomenon of the aeroelastic response of flexible thin airfoils under various angles of attack (AOAs) and flow velocities through wind tunnel experiments and numerical simulations. The vibration modal characteristics are explored, including vibration frequencies, amplitudes, modal transition, and instantaneous characteristics. Vibration is directly measured using non-contact laser sensors, and the numerical model is appropriately configured to simulate the fluid–structure interaction (FSI) problem under large deformation. Experiments cover a range of AOAs (1 ° – 20 °) and incoming velocities (from 10 to 73 m / s), with dynamic responses measured using four laser sensors. Both average and instantaneous modal response features are analyzed, revealing multi-modal characteristics as velocity and AOA increase. The vibration mode transitions from pure bending to higher-order modes as incoming velocity increases. Specifically, at higher velocities and increased AOA, the high-order vibration component shifts from bending-torsional coupled mode to pure-torsional mode. Comparison of vibration frequencies between experimental measurements and finite element method simulations highlights significant shifts, particularly in the pure-torsional mode. Furthermore, employing commercial software ANSYS CFX and ANSYS Mechanical, a two-way three-dimensional FSI model successfully replicates flutter boundaries observed experimentally at 1 ° AOA and approximately incoming velocity 73 m/s. This FSI model is extended to simulate the multi-modal vibrations at 15 ° AOA, yielding insights into flow phenomena contributing to multi-modal vibration at this AOA. An explanation is provided for the multi-modal vibration phenomenon observed in the experiments based on the above insight. Finally, the differences between the experimental and numerical simulations are speculated upon. [ABSTRACT FROM AUTHOR]

The onset of flow-induced oscillations in valve-like configurations remains not completely understood, despite the wide relevance in fluid transport across human physiology and various industrial applications. The present work explores the excitation mechanisms of self-sustained oscillation with key operating parameters in a general-purpose configuration by means of high-fidelity simulations. The investigation is carried out with a partitioned framework that resolves the fluid field by a finite-difference fractional step scheme, discretizes the structural domain via an isogeometric method, and considers an immersed boundary forcing through the interpolation/spreading kernel built by moving-least squares. Our findings confirm the onset of flapping motion in valvular shells, jointly influenced by geometric parameters, structural properties, and flow conditions. Specifically, at a Reynolds number (Re) of 800 and shell aspect ratio of 1.0, a critical reduced velocity exists at around 6, bifurcating static and periodic oscillation modes. After this criterion, flexible shells flutter in the third-plate-mode natural frequency, with oscillation amplitudes approaching an asymptotic value, coupled with intensified vortex shedding, as the reduced velocity increases. Re mainly imparts a destabilizing effect on the fluid-shell system; a lower Re suppresses flow-induced vibrations through viscous dissipation, while a higher Re introduces three-dimensional complexities, asymmetrical oscillations, and quasi-periodicity in the flapping dynamics, especially within the critical regime of reduced velocity. The impact of shell aspect ratio is intricate; in contrast, the case with an aspect ratio of 1.3 displays more intensive flapping motion compared to the reference case of 1.0, whereas further increasing to 1.6 mainly shows stabilizing effects in the shell dynamics. [ABSTRACT FROM AUTHOR]

Labyrinth seals are commonly used in turbomachinery in order to control leakage flows. Flutter is one of the most dangerous potential issues for them, leading to High Cycle Fatigue (HCF) life considerations or even mechanical failure. This phenomenon depends on the interaction between aerodynamics and structural dynamics; mainly due to the very high uncertainties regarding the details of the fluid flow through the component, it is very hard to predict accurately. In 2014, as part of the E-Break research project funded by the European Union (EU), an experimental campaign regarding the flutter behaviour of labyrinth seals was conducted at "Centro de Tecnologias Aeronauticas" (CTA). During this campaign, three realistic seals were tested at different rotational speeds, and the pressure ratio where the flutter onset appeared was determined. The test was reproduced using a linearised uncoupled structural-fluid methodology of analysis based on Computational Fluid Dynamics (CFD) simulations, with results only in moderate agreement with experimental data. A procedure to adjust the CFD simulations to the steady flow measurements was developed. Once this method was applied, the matching between flutter predictions and the measured data improved, but some discrepancies could still be found. Finally, a set of simulations to retain the influence of the external cavities was run, which further improved the agreement with the testing data. [ABSTRACT FROM AUTHOR]

The vibrations of plate structures placed in a supersonic flow was considered. The undisturbed fluid flow was parallel to the plate. This type of problem is especially important in the aerospace industry, where it is named panel flutter. It has been noticed for a long time that panel flutter may be problematic at high speeds. In this article, two specific problems were treated: in the first one, the plate was in the form of an infinite strip and the flow was in the direction of its finite length. Rigid walls indefinitely extended from the sides of the plate. In the second problem, the plate was a finite rectangle and the flow was parallel to one of its sides. The rest of the plane of the rectangle was again rigid. The first problem was a limiting case of the second problem. The flow was modeled by piston theory, which assumes that the fluid pressure on the plate is proportional to its local slope. This approximation is widely used at high speeds (supersonic speeds in the range of M > 1), and reduces the interaction between the fluid flow and the vibrations of the plate to an additional term in the vibration equation. The resulting problem can be solved by assumed mode methods. In this study, the solution was also found by using the collocation method. The contribution of this study is the correlation between the flutter velocity and the other parameters of the plate. The main result is the flutter velocity of the free fluid flow under which the plate vibrations become unstable. Finally, simple expressions are proposed between the various non-dimensional parameters that allows for the quick estimation of the flutter velocity. These simple expressions were deduced by least squares fits to the computed flutter velocities. [ABSTRACT FROM AUTHOR]

Understanding aeroelastic issues related to isolators is pivotal for the structural design and flow control of scramjets. However, research on fluid–structure interactions (FSIs) between thin-walled structures and the isolator flow remains limited. This study delves into the FSIs between thin-walled panels and the isolator flow, as characterized by an oblique shock train, by quantitatively analyzing 11 flow parameters assessing the structural response, separation zones, shock structures, flow symmetry, and performance. The results reveal that an FSI triggers panel flutter under oblique shock train conditions, with the panel shapes exhibiting a combination of first- and second-mode responses, peaking at 0.75 of the panel length. Compared to rigid wall conditions, isolators with a flexible panel at the bottom wall experience downstream movement of the separation zones and shock structures, reduced flow symmetry, and minor changes in performance. Transient fluctuations occur due to the panel flutter. Two flexible panels at the top and bottom walls have a comparatively lesser influence on the averaged parameters but exhibit more violent transient fluctuations. Furthermore, the FSI effects under oblique shock train conditions are contrasted with those under normal shock train conditions. The flutter response under normal shock train conditions is more pronounced, with a larger amplitude and higher frequency, driven by the heightened participation of the first-mode response. The effects of FSIs under normal shock train conditions on the averaged parameters are the opposite (with a larger influence) to those under oblique shock train conditions, with significantly more drastic transient fluctuations. Overall, this study sheds light on the complex and substantial influence of FSIs on the isolator flow, emphasizing the necessity of considering FSIs in future isolator design and development endeavors. [ABSTRACT FROM AUTHOR]

Recent trends in aeroelastic analysis have shown a great interest in understanding the role of shock boundary layer interaction in predicting the dynamic instability of aircraft structural components at supersonic and hypersonic flows. The analysis of such complex dynamics requires a time-accurate fluid-structure interaction solver. This study focuses on the development of such a solver by coupling a finite-volume Navier-Stokes solver for fluid flow with a finite-element solver for structural dynamics. The coupled solver is then verified for the prediction of several panel instability cases in 2D and 3D uniform flows and in the presence of an impinging shock for a range of subsonic and supersonic Mach numbers, dynamic pressures, and shock strengths. The panel deflections and limit cycle oscillation amplitudes, frequencies, and bifurcation point predictions were compared within 10 % of the benchmark results; thus, the solver was deemed verified. Future studies will focus on extending the solver to 3D turbulent flows and applying the solver to study the effect of turbulent load fluctuations and shock boundary layer interactions on the fluid-structure coupling and structural dynamics of 2D panels. [ABSTRACT FROM AUTHOR]

This study relates the structural deformation and the unsteady aerodynamics involved for lift production in bioinspired nanocomposite flapping wings. The sectional chord deformation effects on theoretically supported aerodynamic lift are addressed. Digital image correlation setup and L-shape mechanism are developed and synchronized for wing deformation and aerodynamic forces measurement respectively. DIC system with estimated resolution of 0.159 mm/pixel and L-shape mechanism with net uncertainty of ± 0.72 % verify the measurement accuracy for low Reynolds number flapping wing experiments and alike miniature aero-structures. The flapping wing shape is investigated at 25%, 50% and 75% wing-span to study the effect of sectional chord deformation on lift components obtained from Theodorsen theory. Circulatory lift is dominant at mid-downstroke while Non-circulatory lift is dominant at the beginning of flapping cycle due to inertial effects. [ABSTRACT FROM AUTHOR]

The self-sustained interactions between a flexible film and periodic vortices epitomize the spirit of fish swimming and flag flapping in nature, involving intricate patterns of flow-structure coupling. Here, we comprehensively investigate the multiple coupling states of a film in the cylinder wake mainly with experiments, complemented by theoretical solutions and nonlinear dynamical analyses. Four regimes of film motion states are identified in the parameter space spanned by the reduced velocity and the length ratio. These regimes are (i) keeping stationary, (ii) deflection flutter, (iii) hybrid flutter and (iv) periodic large-amplitude flapping, each governed by a distinct coupling mechanism, involving regular and irregular Kármán vortices, local instability of the elongated shear layers and 2P mode vortex shedding. The film futtering in regimes (ii) and (iii) is substantiated to be chaotic and bears a resemblance to the 'entraining state' of fish behind an obstacle in the river. The periodic flapping in regime (iv) manifests itself in an amalgam of standing and travelling waves, and has intrinsic relations to the 'Kármán gaiting' of fish in periodic vortices. With the spatiotemporal reconstruction for the periodic flapping, we procure the energy distributions on the film, revealing the energy transfer processes between the film and the large-scale vortices. The findings unequivocally indicate that the flow-structure interaction during the energy-release stage of the film is more intense than that during the energy-extraction stage. Given the similarities, the mathematical and physical methods presented in this work are also applicable to the research on biological undulatory locomotion. [ABSTRACT FROM AUTHOR]

The aeroelasticity of a panel in the presence of a shock is a fundamental issue of great significance in the development of hypersonic vehicles. In practical engineering, cavity pressure emerges as a crucial factor that influences the nonlinear dynamical characteristics of the panel. This study focuses on the aeroelastic bifurcation of a flexible panel subjected to both cavity pressure and oblique shock. To this end, a computational method is devised, coupling a high-fidelity reduced-order model for unsteady aerodynamic loads with nonlinear structural equations. The solution is meticulously tracked by continuous calculations. The obtained results indicate that cavity pressure plays a pivotal role in determining the bifurcation and stability characteristics of the system. First, the system exhibits hysteresis behaviour in response to the ascending and descending dynamic pressures. The evolution of hysteresis behaviour originates from the phenomenon of cusp catastrophe. Second, variations in cavity pressure induce three types of bifurcation phenomena, exhibiting characteristics akin to supercritical Hopf bifurcation, subcritical Hopf bifurcation and saddle-node bifurcation of cycles. The system's response at the critical points of these bifurcations manifests as long-period asymptotic flutter or explosive flutter. Lastly, the evolution of the dynamical system among these three types of bifurcations is an important factor contributing to the discrepancies observed in certain research results. This study enhances the understanding of the nonlinear dynamical behaviour of panel aeroelasticity in complex practical environments and provides new explanations for the discrepancies observed in certain research results. [ABSTRACT FROM AUTHOR]

Aeroelastic flutter is a dynamically complex phenomenon that has adverse and unstable effects on elastic structures. It is crucial to better predict the phenomenon of flutter within the scope of aircraft structures to improve the design of their wings. This review aims to establish fundamental guidelines for flutter analysis across subsonic, transonic, supersonic, and hypersonic flow regimes, providing a thorough overview of established analytical, numerical, and reduced-order models as applicable to each flow regime. The review will shed light on the limitations and missing components within the previous literature on these flow regimes by highlighting the challenges involved in simulating flutter. In addition, popular methods that employ the aforementioned analyses for optimizing wing structures under the effects of flutter--a subject currently garnering significant research attention--are also discussed. Our discussion offers new perspectives that encourage collaborative effort in the area of computational methods for flutter prediction and optimization. [ABSTRACT FROM AUTHOR]

In this study, models for the wake capture lift and drag force coefficients of a hovering flapping wing were presented using numerical fluid dynamics simulation to improve the blade element theory. The investigated wing is inspired by the fruit fly and has combined flapping and pitching movements. The effect of changing the wing acceleration time at the start (Δ τ A ) and end of the half cycle (Δ τ D ), as well as the Reynolds numbers in the range of 136–6800, on wake capture using the Taguchi orthogonal array test design, is investigated using the numerical fluid dynamics method, and the values of average lift and drag coefficients due to wake capture are obtained. These force coefficients were applied to linear and nonlinear regression methods to obtain the mathematical model, and a model for its changes was extracted. Finally, to obtain the instantaneous coefficients, the extracted models were placed in the normal distribution function, and the final instantaneous model was obtained. Examining the verification cases of the application of these wake capture relationships with the blade element theory, as well as the effects of translational force, rotational force and added mass force, revealed that this developed theory is capable of correctly predicting the wake capture force's initial peak. The force coefficient trend in the final quasi-steady model with wake capture is similar to the computational fluid dynamics (CFD) results, according to a qualitative examination of the lift and drag force coefficients in a half cycle. This demonstrates a significant result: this theory, which is divided into four parts: transnational, rotational, added mass and wake capture, adequately covers the general physics of these complex movements. [ABSTRACT FROM AUTHOR]

As a representative model in the investigation into fluid–structure coupling, a flexible film interacting with the wake of a circular cylinder involves intricate patterns of both solid motion and fluid evolution. Recent investigations have found that the interactions could be either periodic or aperiodic in experiments; the latter, however, is often overlooked. In this work, the irregular aperiodic flutter of a flexible film behind a circular cylinder is investigated experimentally. It is determined that the irregular flutter intermittently exhibits transient quasi-periodic mode and aperiodic mode. The former is accompanied by the large-scale vortices alternatively formed against the film surfaces, while the latter is associated with vortices formed downstream of the film's trailing edge, so that the whole film is enveloped by the extended shear layers. In order to separate the data individually pertaining to each of the two modes from the whole dataset, a motion-mode recognition method is proposed, and then conditional statistics of flow fields are achieved. The quasi-periodic mode corresponds to more intense velocity fluctuations in the wake, while in the aperiodic mode, the observed localized instability of shear layers induces an increase in the local streamwise velocity fluctuation. [ABSTRACT FROM AUTHOR]

An arbitrary Lagrangian-Eulerian lattice Boltzmann flux solver (ALE-LBFS) coupled with the mode superposition method is proposed in this work and applied to study two- and three-dimensional flutter phenomenon on dynamic unstructured meshes. The ALE-LBFS is applied to predict the flow field by using the vertex-centered finite volume method with an implicit dual time-stepping method. The convective fluxes are evaluated by using lattice Boltzmann solutions of the non-free D1Q4 lattice model and the viscous fluxes are obtained directly. Additional fluxes due to mesh motion are calculated directly by using local conservative variables and mesh velocity. The mode superposition method is used to solve for the dynamic response of solid structures. The exchange of aerodynamic forces and structural motions is achieved through interpolation with the radial basis function. The flow solver and the structural solver are tightly coupled so that the restriction on the physical time step can be removed. In addition, geometric conservation law (GCL) is also applied to guarantee conservation laws. The proposed method is tested through a series of simulations about moving boundaries and fluid–structure interaction problems in 2D and 3D. The present results show good consistency against the experiments and numerical simulations obtained from the literature. It is also shown that the proposed method not only can effectively predict the flutter boundaries in both 2D and 3D cases but can also accurately capture the transonic dip phenomenon. The tight coupling of the ALE-LBFS and the mode superposition method presents an effective and powerful tool for flutter prediction and can be applied to many essential aeronautical problems. [ABSTRACT FROM AUTHOR]

Recent studies in turbomachinery have shown that the phase of acoustic wave reflection within an intake can have either positive or negative effects on the aeroelastic stability of fan rotor blades. However, the typical flow structures, such as the shock wave, within rotor blade passages with acoustic wave reflection remain unclear. The aim of this research was to address this gap by investigating how these flow structures impact blade aeroelastic stabilities with acoustic wave reflections. The focus of this study was the NASA Rotor 67 blade with an extended intake. Moreover, a bump is incorporated on the shroud at different distances from the fan to reflect acoustic waves of varying phases. Utilizing the energy method, variations in the aerodynamic work density on blade surfaces were calculated under different phases of reflected acoustic waves. Analysis indicates that the spatial position of the shock wave undergoes periodic changes synchronized with the phase of acoustic reflection, marking the first instance of such an observation. This synchronization is identified as the primary factor causing variations in the aeroelastic stability of blades due to acoustic wave reflection, contributing to a deeper understanding of the mechanism behind acoustic flutter. The acoustic–vortex coupling at the blade tip leads to unpredictable variations in unsteady pressures on the blade suction surface, although its effect on blade aeroelastic stabilities is relatively limited compared to that of the shock wave. [ABSTRACT FROM AUTHOR]

When aiming to cut down on the emission of nitric oxides by gas turbine engines, it is advantageous to have them operate at low combustion temperatures. This is achieved by lean premixed combustion. Although lean premixed combustion is a proven and promising technology, it is also very sensitive to thermoacoustic instabilities. These instabilities occur due to a coupling between the unsteady heat release rate of the flame and the acoustic field inside the combustion chamber. In this paper, this coupling is investigated in detail. Two acoustic design parameters of a swirl-stabilized pressurized preheated air (300 °C)/natural gas combustor are varied, and the occurrence of thermoacoustic limit cycle oscillations is explored. The sensitivity of the acoustic field as a function of combustion chamber length (0.9 m to 1.8 m) and reflection coefficient (0.7 and 0.9) at the exit of the combustor is investigated first using a hybrid numerical and analytical approach. ANSYS CFX is used for Unsteady Reynolds Averaged Navier-Stokes (URANS) numerical simulations, and a one-dimensional acoustic network model is used for the analytical investigation. Subsequently, the effects of a change in the reflection coefficient are validated on a pressurized combustor test rig at 125 kW and 1.5 bar. With the change in reflection coefficient, the combustor switched to limit cycle oscillation as predicted, and reached a sound pressure level of 150 dB. [ABSTRACT FROM AUTHOR]