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Simulating interactions involving compressible fluids with potential shock waves and highly deformable structures demands the precise capture of flow dynamics and structural responses. This coupling must effectively manage substantial deformations while adapting to changes in the fluid domain's topology. This study focuses on the rhoSonicFsiFoam solver, developed using the open-source computational fluid dynamics (CFD) software, OpenFOAM. The computational approach addresses aeroelasticity, fluid compressibility and advanced coupling. Two tests were performed, with the first involving the beating of a thin plate in supersonic flow and the second featuring the interaction of a shock wave with a deformable plate. In the flutter test, simulation results intimately matched the theoretical Mach number of 2.2677, indicating a strong alignment between the numerical results of this present study ( M ∞ num ∈ 2.26 , 2.27 ) and theoretical expectations. Additionally, simulation results from the computational approach closely align with theoretical projections for plate fluttering and correlate well with experimental and numerical outcomes from the T80 shock tube. This project aims to unravel fluid–structure interaction by combining a robust solution method with the latest advancements in CFD. Leveraging this platform, the computational approach accurately captures aeroelastic phenomena, including shock waves and deformations during interactions with compressible fluid and deformable structures. Implications span aerospace, defense and energy sectors. In conclusion, this study underscores the pivotal role of advanced simulation techniques in comprehending intricate interactions involving compressible fluids and structures. [ABSTRACT FROM AUTHOR]

A new central result that gives the necessary and sufficient conditions for two n by n skew-symmetric matrices and one symmetric matrix to be simultaneously quasi-diagonalized by a real orthogonal congruence is proved. Based on this result, the decomposition of linear multi-degree-of-freedom dynamical systems with gyroscopic, circulatory, and potential forces is investigated through a real linear coordinate transformation generated by an orthogonal matrix. Several sets of conditions, applicable to real-life structural and mechanical systems arising in aerospace, civil, and mechanical engineering, under which such a coordinate transformation exists are found, thereby allowing these systems to be decomposed into independent, uncoupled subsystems, each with a maximum of two degrees of freedom. The conditions are expressed in terms of the coefficient matrices of the system. A specific form for the circulatory (gyroscopic) matrix is posited, and when the gyroscopic (circulatory) matrix is simple--a situation that commonly appears in real-life applications--it is shown that just a single necessary and sufficient condition is required for the decomposition of the multi-degree-of-freedom system. Numerical examples are provided throughout to demonstrate the analytical results. [ABSTRACT FROM AUTHOR]

The source and doublet panel method (SDPM) developed by Morino in the 1970s can model unsteady compressible ideal flow around wings and bodies. In this work, the SDPM is adapted to the calculation of aeroelastic solutions for wings. A second-order nonlinear version of Bernoulli's equation is transformed to the frequency domain and written in terms of the generalized mode shapes and displacements. It is shown that the pressure component at the oscillating frequency is a linear function of the generalized displacements, velocities, and accelerations and can therefore be used to formulate a linear flutter problem. The proposed approach has several advantages over the usual doublet lattice method (DLM) approach: the exact geometry is modeled (including thickness, camber, and twist effects), the motion of the surface can be represented using all six degrees of freedom, the pressure calculation is of higher order, and the aerodynamic mass, damping, and stiffness terms are calculated explicitly. The complete procedure is validated using experimental data from the weakened AGARD 445.6 wing, a NACA 0012 rectangular wing with pitch and plunge degrees of freedom, and an experimental model of a T-tail, yielding flutter predictions that lie closer to the experimental observations than those obtained from the DLM, particularly in the case of the T-tail. [ABSTRACT FROM AUTHOR]

Introduction: The application of artificial intelligence, particularly neural network models, has expanded throughout various fields of academic and industrial studies, recently. This research focuses to investigate aeroelastic phenomena and predict flutter speeds without expensive computational and experimental methods utilizing these algorithms, specifically the artificial neural network (ANN). Methods: Fundamentally, the approach involves an ANN algorithm to navigate the complexities of aeroelastic systems, achieved by layering multiple ANNs. The network's neurons effectively interpret diverse flight data by this structure. The study also incorporates state-space representation and Theodorsen's unsteady aerodynamic theory to create a comprehensive dataset on aeroelastic flutter speed across different wing parameters. In order to forecast the flutter speeds, a feed-forward neural network (FFNN) model, which is an approach of ANN method, containing sigmoid hidden neurons and a linear output neuron has been proposed for the conditions above. Results: The proposed model accomplished a regression value of 0.92 for flutter speeds according to the experimental findings. As presented, the suggested FFNN model is successful and can be employed to predict the flutter speed estimation. The findings of the FFNN structure are validated with the previous work for aeroelastic flutter speed analysis in the literature. Discussion: Finally, the findings indicate that the FFNN model formulated in this study is highly accurate in predicting flutter speeds with the accordance of numerical results of aeroelastic structure modeled with unsteady aerodynamic theory. In addition, the correctness of the predicted flutter speeds demonstrates that analyzing without expensive tests and high computational costs is in the near future. [ABSTRACT FROM AUTHOR]

Abstract Flutter is an extremely significant academic topic in both aerodynamics and aircraft design. Since flutter can cause multiple types of phenomena including bifurcation, period doubling, and chaos, it becomes one of the most unpredictable instability phenomena. The complexity of modeling aeroelasticity of high flexibility wings will be substantially simplified by investigating the prospect of system identification techniques to forecast flutter velocity. Therefore, a novel neural network (NN)-based method for aeroelastic system identification is proposed. The proposed NN-based approach constructs an NN framework of high flexibility wings flutter models with different materials and sizes, which can effectively predict the flutter velocity of flexible wings. The accuracy of the method is demonstrated by comparing with the simulation results.

Hypersonic aircraft encounter challenges in adapting to complex external environments during missions. Conventional fixed-wing aircraft have limited optimal aerodynamic performance under specific conditions. In contrast，morphing aircraft can adapt to various mission environments by altering their aerodynamic shape to control their flight characteristics. Based on numerical simulations，the changes in aerodynamic characteristics of a hypersonic folding-wing aircraft before and after wing folding are analyzed in this paper for revealing the effects of wing folding on aerodynamic performance. Meanwhile，the influence of aeroelasticity in the wing on the aerodynamic characteristics of the aircraft during the wing folding process is discussed. The results show that the wing folding process has a little effect on the drag coefficient but significantly affects the lift coefficient and lift-to-drag ratio. Moreover，due to the presence of shock waves at the wingtips，stress is mainly concentrated in these areas，resulting in a relatively small impact of aeroelasticity on the aircraft's aerodynamic behavior.

Based on the extended potential and piston aerodynamics theories, buckling and vibration analysis of sandwich cylindrical shells subjected to external airflow in both subsonic and supersonic regimes is presented. Sandwich cylindrical shells are considered to be made of same homogenous face sheets and an anisogrid composite lattice core with hexagonal cells. An equivalent single layer (ESL) theory in the framework of classical Love's shell theory is used to formulate the global mechanical behavior of such shells. According to the extended potential and piston aerodynamic theories, aerodynamic pressure distributions are determined for steady, inviscid, irrotational compressible, and external airflow. The closed–form solutions of critical upstream pressure and natural frequency are developed for different boundary conditions using the Galerkin method. Finally, case studies are presented to compare between results of extended potential and piston theories and investigate the effects of upstream speed, geometric ratios, boundary conditions, and lattice specifications on various response quantities. Findings indicate that the differences in results of extended potential and piston theories are less than 3% in the supersonic regime with 2.3 ≤ M ∞ ≤ 2.7 for intermediate-length shells with a wide range of R. The piston theory is inadequate for analyzing aeroelasticity behavior of very long thin sandwich cylindrical shells. [ABSTRACT FROM AUTHOR]

A solution for passive aeroelastic control is presented and tested experimentally on a bidimensional wing setup. The solution consists of a flap integrated in the wing which acts as a secondary mass damper and absorbs energy when aeroelastic instability is encountered. This device is passive, which makes it safe in emergency cases and adds little mass to the system in the fundamental case presented in this paper. If installed on actual aircraft wings, the flap-NES would not substitute but coexist with classic active control methods, adding virtually zero mass to the existing system. Using a flap placed in the airflow enables the control system to benefit from aerodynamic damping with a behavior dependent on wind speed. The device can either have a linear or nonlinear stiffness. It is shown that both options absorb energy from the main system, that is, the wing. The flap acts as a TMD (tuned mass damper) in the linear case and as an NES (nonlinear energy sink) in the nonlinear case. The nonlinear solution not only absorbs more energy at given wind speed but it also is a more suited solution as the damper follows the wind speed dependent frequencies due to its nonlinear feature. In this work, the flap-TMD and the flap-NES are tested experimentally on a bidimensional wing which presents classic flutter. [ABSTRACT FROM AUTHOR]

This paper investigates an output-based approach for flutter bifurcation analysis of single-bay panels in supersonic flow. The approach leverages bifurcation forecasting, a class of methods to predict bifurcation diagrams using prebifurcation output data. This work is the first study into this approach applied to panel limit-cycle oscillations, building on previous efforts focused on geometrically nonlinear wings and propeller-nacelle systems. The study uses output data from transient simulations of single-bay panels at as few as two preflutter dynamic pressures, which are selected using an eigenvalue-based criterion that ensures consistent prediction accuracy across panel configurations. The approach captures the bifurcation type and amplitude variation of limit-cycle oscillations around the flutter point for a variety of materials, boundary conditions, thermal loads, and cross-stream curvatures. This approach can facilitate nonintrusive panel limit-cycle oscillation analyses for parametric studies and design. [ABSTRACT FROM AUTHOR]

The accurate prediction of the aeroelastic behavior of turbomachinery for aircraft propulsion poses a difficult yet fundamental challenge, since modern aircraft engines tend to adopt increasingly slender blades to achieve a higher aerodynamic efficiency, incurring an increased aeroelastic interaction as a drawback. In the present work, we present a reduced order model for flutter prediction in axial compressors. The model exploits the aerodynamic influence coefficients technique with the adoption of a broadband frequency signal to compute the aerodynamic damping for multiple reduced frequencies using a single training simulation. The normalized aerodynamic work is computed for a single oscillation mode at three different vibration frequencies, comparing the outputs of aerodynamic input/output models trained with a chirp signal to those from single-frequency harmonic simulations. The results demonstrate the ability of the adopted model to accurately and efficiently reproduce the aerodynamic damping at multiple frequencies and arbitrary nodal diameters with a single simulation. [ABSTRACT FROM AUTHOR]

Objective: In this paper, the nonlinear flutter of the wing is investigated under the influence of aerodynamic control surfaces. Methods: The wing aerodynamic loads are determined using Peter's unsteady aerodynamic model, and the aerodynamic loads of the control surface are added with quasi-steady relations in the interior of the equations. The governing aeroelastic equations are presented in the structure of fully intrinsic and these equations are discretized using the finite difference method. Results: The effects of the presence of an aerodynamic control surface have been investigated based on the analytical-experimental relationships and considering the nonlinear effects of high control surface deflections. Furthermore, investigation of the effects of some important parameters such as deflections, location, chord size, and length of the control surface on the speed and frequency of flutter instability, is another achievement of this article. Conclusions: The results show that based on aeroelastic considerations, the deflection angle of the control surface has an important effect on the aeroelastic stability. Also, by bringing the control surface closer to the wing tip, increasing the thickness ratio and the chord ratio in accordance with other effective parameters, flutter suppression can be caused. [ABSTRACT FROM AUTHOR]

To enable wind energy to surpass fossil fuels, the power-to-cost ratio of wind turbines must be competitive. Increasing installation capacities and wind turbine sizes indicates a strong trend toward clean energy. However, larger rotor diameters, reaching up to 170 m, introduce stability and aeroelasticity concerns and aerodynamic phenomena that cause noise disturbances. These issues hinder performance enhancement and social acceptance of wind turbines. A critical aerodynamic challenge is flow separation on the blade's suction side, leading to a loss of lift and increased drag, ultimately stalling the blade and reducing turbine performance. Various active and passive flow control techniques have been studied to address these issues, with passive techniques offering the advantage of no external energy requirement. High-lift devices, such as leading-edge slats, are promising in improving aerodynamic performance by controlling flow separation. This study explores the geometric parameters of slats and their effects on wind turbine blades' aerodynamic and acoustic performance. Using an adequate turbulence model at Re = 106 for angles of attack from 14° to 24°, 77 slat configurations were evaluated. Symmetric slats showed superior performance at high angles of attack, while slat chord length was inversely proportional to aerodynamic improvement. A hybrid method was employed to predict noise, revealing slat-induced modifications in eddy topology and increased low- and high-frequency noise. This study's main contribution is correlating slat-induced aerodynamic improvements with their acoustic effects. The directivity reveals a 10–15 dB reduction induced by the slat at 1 kHz, while the slat induces higher noise at higher frequencies. [ABSTRACT FROM AUTHOR]

The aeroelastic responses and nonlinear behaviors of a two-dimensional panel impinged by an oscillating Mach stem shock are investigated through theoretical analysis. Through the nonlinear descriptors, such as Poincaré maps and Largest Lyapunov exponents, the panel with oscillating shock impingement is found to exhibit multiple responses, including single/multi-periodic limit cycle oscillation, quasi-periodic motion, and chaotic motion. Without altering the in-plane force, which is the principal source of structural nonlinearity, the shock oscillation complicates the nonlinear behaviors of the panel. With shock oscillation, the original divergence instability is transformed into post-divergence limit cycle oscillation, and the flutter response exhibits rich nonlinear characteristics. The effect of initial shock impingement location, shock oscillating amplitude, and shock oscillating frequency are disclosed through the bifurcation diagram, which significantly influences the nonlinear characteristics of the panel response. By reasonably adjusting the shock oscillating parameters, unpredictable nonlinear behaviors, especially chaotic motions, can be avoided. [ABSTRACT FROM AUTHOR]

Transonic aeroelasticity remains a significant challenge in aerospace. The coupling mechanism of aeroelastic problems involving the coexistence of fluid modes and multiple structural modes still needs further investigation. For this purpose, we analysed the dynamic characteristic of a two-degree-of-freedom (2DOF) NACA0012 airfoil in pre-buffet flow. First, we constructed an aeroelastic reduced-order model, which can represent near-unstable transonic flow using the dominant fluid mode. Then, the flutter mechanism was investigated by studying the main eigenvalues of the model that vary with the natural pitching frequency. The results revealed that the existence of the fluid mode transitions the transonic flutter type from coupled-mode flutter to single-DOF (SDOF) flutter, which leads to a reduction in the flutter boundary. Under the effect of the fluid mode, the system produces six aeroelastic phenomena at different structural natural frequencies, including SDOF heaving/pitching flutter, heaving/pitching instability within coupled-mode flutter, forced vibration and stable state. Moreover, we identified two types of SDOF flutter in the 2DOF system. The first type corresponds to the traditional SDOF flutter, where the coupling of other modes has a small impact on the system's stability in most cases. However, within specific ranges of natural frequencies, this type of SDOF flutter may disappear due to coupling with other modes. The second type of SDOF flutter is characterized by strong coupling dominated by the unstable mode. It arises from the interaction among the flow, heaving and pitching modes, and does not manifest in the absence of any of these modes. [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]

Mesh deformation technology is widely used in aerodynamic applications like unsteady flow, aeroelasticity, and aerodynamic shape optimization because of its low computational costs and consistent mesh connectivity. In order to raise deformed mesh quality and improve efficiency, a new mesh deformation method based on quaternion and displacement normal propagation (named QN method) is introduced in this paper. The boundary points propagate their displacements composed of translational vectors and quaternions to corresponding volume points along the normal direction under the control of the damping function, which preserves the mesh shape and guarantees the quality near boundaries, including orthogonality and normal size. It can also prevent the volume points from being interfered by other less relevant boundary points, so dealing with complex displacement fields effectively. In addition, it avoids complicated matrix and interpolation operations, thus saving lots of computational costs. For mesh with complex topology, a hybrid method combining the QN method and the radial basis function method (RBF method) is investigated to broaden application scenarios, which are applied to the mesh with normal correspondence inside the boundary layer and the mesh outside, respectively. Benefitting from effective handling for the near-wall elements by the QN method, the RBF interpolation part in the hybrid method requires minor support points to carry out valid large deformation, improving the deformation efficiency greatly compared to the individual RBF method. Five typical test cases with different deformation modes and mesh characteristics are implemented, showing better performance of the proposed method in deformed mesh quality and deformation efficiency. [ABSTRACT FROM AUTHOR]

The in-plane folding wing is one of the important research directions in the field of morphing or bionic aircraft, showing the unique application value of enhancing aircraft maneuverability and gust resistance. This article provides a structural realization of an in-plane folding wing and an aeroelasticity modeling method for the folding process of the wing. By approximating the change in structural properties in each time step, a method for calculating the structural transient response expressed in recursive form is obtained. On this basis, an aeroelasticity model of the wing is developed by coupling with the aerodynamic model using the unsteady panel/viscous vortex particle hybrid method. A wind-tunnel test is implemented to demonstrate the controllable morphing capability of the wing under aerodynamic loads and to validate the reliability of the wing loads predicted by the method in this paper. The results of the gust simulation show that the gust scale has a significant effect on the response of both the open- and closed-loop systems. When the gust alleviation controller is enabled, the peak bending moment at the wing root can be reduced by 5.5%∼47.3% according to different gust scales. [ABSTRACT FROM AUTHOR]

A forced response induced by inlet distortion and blade row interactions can lead to high-cycle fatigue failure, especially when the unsteady excitation frequency is close to the natural vibration frequency of a blade. This paper presents the study of forced response sensitivity analysis and minimization using a full-viscosity unsteady discrete adjoint method. The goal is to improve the aeroelastic performances of turbomachinery blades and simultaneously constrain/improve aerodynamic performances. The decoupled modal reduction method is used to compute the forced response, which is considered the objective function in adjoint-based design optimization. To analyze aeroforcing and aerodamping flow and adjoint fields efficiently, the harmonic balance method and its adjoint counterpart developed by an automatic differentiation tool are applied. Two cases--the NASA Rotor 67 with inlet distortion and a three-row configuration with multiple fundamental frequencies in the second row--are used to demonstrate the effectiveness of the aerodynamic and aeroelastic coupled design optimization system developed in this work. For the latter case, the Fourier-transform-based method that first decomposes and then matches the time and space modes at two sides of an interface is used for interface coupling. The almost-periodic Fourier transform method is used to determine the time instances for cases with multiple fundamental frequencies. [ABSTRACT FROM AUTHOR]

The growing demand for increasing the engine power of a liquid rocket is driving the development of high-power De-Laval nozzles, which is primarily achieved by increasing the expansion ratio. A high-expansion-ratio for De-Laval nozzles can cause flow separation, resulting in unsteady, asymmetric forces that can limit nozzle life. To enhance nozzle performance, various separation control methods have been proposed, but no methods have been fully implemented thus far due to the uncertainties associated with simulating flow phenomena. A numerical study of a high-area-ratio rocket engine is performed to analyze the aeroelastic performance of its structure under flow separation conditions. Based on numerical methodology, the flow inside a rocket nozzle (the VOLVO S1) is analyzed, and different separation patterns are comprehensively discussed, including both free shock separation (FSS) and restricted shock separation (RSS). Since the location of the flow separation point strongly depends on the turbulence model, both the single transport equation and two-transport-equation turbulence models are simulated, and the findings are compared with the experimental results. Therefore, the Spalart–Allmaras (SA) turbulence model is the ideal choice for this rocket nozzle geometry. A wavelet is used to analyze the amplitude frequencies from 0 to 100 Hz under various pressure fluctuation conditions. Based on a clear understanding of the flow field, an aeroelastic coupling method is carried out with loosely coupled computational fluid dynamics (CFD)/computational structural dynamics (CSD). Some insights into the aeroelasticity of the nozzle under separated flow conditions are obtained. The simulation results show the significant impact of the structural response on the inherent pressure pulsation characteristics resulting from flow separation. [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]

This work concerns the dynamic similarity criteria of various phenomena occurring in the aerodynamics of buildings and structures, originally derived from the ratios of forces and force moments affecting these phenomena. This paper is a continuation of [12], which addresses the foundations of dynamic similarity criteria formulated in this manner. At the end of [12], an authorial method and procedure for determining dynamic similarity criteria in fluid-solid interaction issues are presented. This method serves as the basis for the formulations and considerations of dynamic similarity criteria discussed further for various practical problems encountered in simple cases of building and structure aerodynamics, including self-exciting vibrations and wind-induced vibrations. [ABSTRACT FROM AUTHOR]

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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 state-of-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.

Elements of powerful steam turbines are subjected to significant unsteady loads, in particular the rotor blades of the last stages. These loads, in some cases, can cause self-excited oscillations, which are extremely dangerous and have a negative impact on the efficiency and service life of the blade cascade. Therefore, when designing new or modernizing existing steam turbine stages, it is recommended to study the aeroelastic characteristics of the blades. The conditions for the occurrence of self-excited oscillations are influenced by both the geometric characteristics and the alloy from which the blade is made. To determine the effect of the blade material on the aeroelastic behaviour, a numerical analysis of the aeroelastic characteristics of the last stage blades made of steel and titanium alloy was performed. For the analysis, the method of simultaneous modeling of unsteady gas flow through the blade cascades and elastic vibrations of the blades (coupled problem) was used, which allows obtaining the amplitude-frequency spectrum of the interaction of unsteady loads and blade vibrations. The paper presents the results of numerical analysis for harmonic oscillations with a given amplitude and a given inter-blade phase angle, as well as for the regime of coupled vibration of blades under the action of unsteady aerodynamic forces. The dependences of the aerodamping coefficient on the inter-blade phase angle and the distribution of the coefficient along the blade are presented. The results of modeling the coupled vibration of the blades for the first six natural forms are presented in the form of a time-evolving displacement of the blade peripheral section, as well as forces and moments acting on the peripheral section. The corresponding amplitude-frequency spectra of displacements and loads in the peripheral section are also presented. The analysis of the results showed an insignificant difference in the characteristics of the proposed blade materials. For the first natural form of blade oscillations, the possibility of self-excited oscillations was found, and for the second form, there are conditions for the appearance of stable self-oscillations.

The dynamic behaviour of a spring-mounted symmetrical NACA0012 wing in a freestream flow of air is studied in the pre-stall region, over 0° to 12° angles of incidence. The primary aim of this work is for use within the automotive sector to reduce drag and fuel emissions. However, this work will also be of interest in the motorsport sector to improve performance, and also have some applications within the aerospace and renewable energy sectors. The general operation of the concept has previously been verified at these low angles in the pre-stall region with that of a theoretical estimation using finite and infinite wings. This paper provides a numerical solution of the same problem and is compared with the previous experimentation. At these low angles, the computations yield a dynamic response settling into a static equilibrium. The stable solutions match the start of a steady regime well, when compared with the experiment. The trends are also comparable with the experiment, but the velocities at which they occur are underestimated in the computation. The computations demonstrate a drag reduction of 59% when compared to a fixed wing, whereas the lift remains stable at a near constant value with increasing wind speed. Thence, downforce is maintained whilst drag is reduced, which will facilitate higher speeds on the straight whilst maintaining vehicle direction stability. Limitations to this proof-of-concept work are highlighted and future development work is suggested to achieve even further increases in performance.

This paper presents a CFD/CSD coupling method for aeroelastic simulation of folding tail morphing aircraft. The unsteady aerodynamic analysis is based on an in-house computational fluid dynamics (CFD) solver for the Euler equations, and emphasis is made on developing an efficient dynamic mesh method for the tail’s hybrid fold motion/elastic vibration deformation. The structural dynamic analysis is based on the computational structural dynamics (CSD) technique for solving the structural equation of motion in modal space. The aeroelastic coupling was achieved through successive iterations of CFD and CSD computations in the time domain. An adaptive multi-functional morphing aircraft allowing tail fold motion was selected to be studied. By using the developed method, aeroelastic simulation and mechanism analysis for fixed configurations at different folding angles and for variable configurations during the folding process were performed. The influence of folding rate on tail aeroelasticity and its influence mechanism were also analyzed.

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]

This study employs a high-fidelity numerical approach to simulate fluid–structure interaction phenomena for the dynamic response of flexible hyperelastic morphing wing structures under low aerodynamic loads. The computations are performed using the open-source solids4Foam toolbox, employing a partitioned two-way fluid–structure interaction approach with a finite volume solver for both fluid and solid. The considered morphing wing is divided into a flexible and a rigid segment, with the flexible segment featuring a 60% chord length and being made of a hyperelastic rubber-like material. The concept of damping is incorporated into the solid momentum balance equation as a virtual force that opposes the velocity of the structure. Damping is employed to disperse energy from the system, hence mitigating the oscillations and reducing computational time. To understand morphing wing aerodynamics and aeroelasticity behavior, a series of tests are conducted at low and medium Reynolds numbers, specifically 2 × 10 5 and 5 × 10 5 . The results show that, for low Reynolds number, the morphing structure has a negligible impact on aerodynamic behavior. However, at higher Reynolds numbers, morphing results in improved aerodynamic efficiency at low angles of attack. Overall, the study highlights the aero-structural behavior of hyperelastic morphing wings and their potential for developing efficient and adaptive wing structures, highlighting their promise for future aircraft design innovations. [ABSTRACT FROM AUTHOR]

The reliability of a pitch-plunge hypersonic airfoil in random fluctuating flow with both cubic and freeplay nonlinearity is examined. The Hopf bifurcation and dynamic responses of the hypersonic airfoil are performed. To analyze the reliability, the effects of stochasticity on the dynamic behaviors of the hypersonic airfoil model are discussed in detail. Several unwanted phenomena that result in the failure of the airfoil structure are induced by random fluctuations. Subsequently, the reliability of the airfoil model is defined and analyzed according to the first passage failure criteria. The effects of different parameters on the reliability are investigated. Furthermore, a nonlinear energy sink is introduced to suppress the vibration of the airfoil and enhance the reliability. Two-dimensional reliability regions of the airfoil model are given to provide the safety parameter region. The results show that the reliability of the airfoil model is significantly improved with the nonlinear energy sink. This work will provide new insights into the safety design of hypersonic aircraft. [ABSTRACT FROM AUTHOR]