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Dynamic mode decomposition (DMD) is a powerful data-driven modal decomposition technique that extracts spatiotemporal coherent structures: a useful process in flow diagnostics and future state estimation of complex nonlinear flow phenomena. Transonic shock buffet is a complicated phenomenon, and modal decomposition techniques such as DMD provide significant insight into its complicated flow physics; but, often, flowfield data are corrupted because of various sources of noise due to the presence of outliers or the absence of critical data components. Therefore, noise corruption renders the modal decomposition inaccurate, and thereby not useful. In this paper, two sources of noise have been considered: simple white noise, and complex salt-and-pepper-type spurious noise. Various DMD techniques including standard DMD, forward–backward DMD, total-least-squares DMD, higher-order DMD, and robust DMD have been implemented. Their effectiveness and limitations in countering noise corruption have been investigated systematically. In the case of white noise corruption, forward–backward DMD, total-least-squares DMD, and higher-order DMD capture the buffet frequency and growth rate with sufficient accuracy, whereas the latter outperforms the other two when the noise variance level is above 5%. In the case of spurious noise, robust DMD handles noise corruption efficiently, with surprisingly high values of pixel corruption of up to 30%.

This paper proposes a novel method for an automatic operational modal analysis. An available finite element structural model is used together with sensor data to derive a linear representation of t...

This paper presents a physics-based projection framework for the interfacial treatment of spatially and topologically different grids. The formulation proposes the use of constraint type elements, ...

The paper broadly addresses how Industry 4.0 program drivers will impact maintenance in aviation. Specifically, Industry 4.0 practices most suitable to aeronautical maintenance are selected, and a detailed exposure is provided. Advantages and open issues are widely discussed and case studies dealing with realistic scenarios are illustrated to support what has been proposed by authors. The attention has been oriented towards Augmented Reality and Additive Manufacturing technologies, which can support maintenance tasks and spare parts production, respectively. The intention is to demonstrate that Augmented Reality and Additive Manufacturing are viable tools in aviation maintenance, and while a strong effort is necessary to develop an appropriate regulatory framework, mandatory before the wide-spread introduction of these technologies in the aerospace systems maintenance process, there has been a great interest and pull from the industry sector. Highlights Industry 4.0 practices most suitable to aeronautical maintenance are selected. Advantages and open issues are widely discussed and case studies are illustrated. Augmented Reality can support maintenance tasks. Additive Manufacturing can be useful to produce spare parts. A strong effort is necessary to develop an appropriate aeronautical regulatory framework.

The dynamic instability of electroconductive cylindrical shells interacting with an external magnetic fields is considered in the paper. Two cases of applied external dynamic loads are discussed, namely (i) a harmonic mechanical force, and (ii) a harmonic magnetic field force. Analytical descriptions of the two-dimensional equations and associate conditions of dynamic instability are presented. On the basis of the formulated problems, specific issues related to the dynamic instability of electroconductive cylindrical shells in a magnetic field are offered. The study illustrated the effect of magnetoelastic interaction. Specifically it is shown that there exist a minimum value of the given magnetic field intensity, above which the parametric resonance due to external harmonic force is eliminated; furthermore, in a presence of time-harmonic magnetic field the parametric resonance with a resonance frequency can be generated not only near the first frequency of the external magnetic field (which is equal to the natural frequency of vibrations), but near the double frequency of natural vibrations. When the forced vibrations of conductive shells caused by external forces of non-electromagnetic origin is considered, including the effect of a time-harmonic magnetic field, resonance can occur in the presence of a non-stationary harmonic magnetic field. Results reveal that the rapid increase of the amplitude of vibrations occurs when the frequency of the external magnetic field is in close proximity to the first natural frequency of the magnetoelastic vibrations of the plate, as well as when the frequency of the magnetic field, is equal to half of the shell natural frequency.

Understanding the effect of mechanical uncertainties can play a significant role in design of the nanocomposites. The uncertain natural frequencies of moderately thick doubly-curved functionally graded composite panels reinforced by carbon nanotube (CNT) are investigated. Specifically, doubly-curved shell panels, including spherical, cylindrical and hyperbolic paraboloid panels are examined. To evaluate uncertainty propagation, uncertainty resources including distribution of the CNT through the thickness as well as the mechanical properties of the CNT and polymer matrix are taken into consideration. To assess the propagated uncertainties in the vibrational characteristics of nanocomposite panels, the interval analysis method is employed while the mechanical properties of nanocomposite panels are predicted using the modified rule of mixture method. Based on the comparison between the results of the present study and those reported in the literature, the accuracy of the results is validated. The sensitivity analysis is performed to distinguish the most prominent uncertain variables. Furthermore, numerical results reveal the influences of various uncertainty resources on the upper and lower bounds of uncertain frequencies and uncertainty propagation percent.

Environmental vibrations often exist in a broadband form. To have a robust performance over a wide frequency range, vibration energy harvesters need to be designed to be insensitive to excitation frequencies. In this paper, we propose a novel two-degree-of-freedom (DOF) piezoelectric energy harvester (PEH) with stoppers that introduce nonlinear dynamic interaction between the two DOFs. First, the mechanical model of the 2DOF system with stoppers is developed by emulating the impact behavior as a piecewise linear stiffness and the working principle is explained. Subsequently, the analytical solution of the system with piecewise linear stiffness is derived using the averaging method and the dynamic response of the system is obtained and confirms its wide bandwidth property. Finally, by integrating the mechanical system with a piezoelectric transducer, the energy harvesting performance of the proposed 2DOF PEH with stoppers is numerically evaluated. The open circuit voltage response of the proposed system is compared with that of the conventional linear 2DOF and 1DOF PEHs. A parametric study reveals the effect of the stopper distance on the energy harvesting performance in terms of both the bandwidth and open circuit voltage output. The superiority of the proposed system in terms of both power output and operation bandwidth is demonstrated.

In this paper, a robust output feedback control design is developed for suppression of aeroelastic vibration of a 2-DOF nonlinear wing section system. The aeroelastic system operates in a quasi-steady aerodynamic incompressible flowfield and is actuated using a combination of a leading-edge (LE) and a trailing-edge (TE) flap. By only utilizing measurements of pitching and plunging deflections, an innovative Lyapunov-based procedure is used to design sliding mode control inputs for the LE and TE control surface deflections. The closed-loop system is shown to have semi-global asymptotic stability even in the presence of model uncertainty and unknown external gust loading. Extensive simulation results under a variety of scenarios show the effectiveness of the control strategy.

In the framework of the membrane theory of cylindrical shells, the localised vibration near the shell free edge is considered. This paper presents the development leading to the dispersion equations for finite length shells and for three types of boundary conditions. Based on the attained dispersion equations the localised membrane vibration conditions are analysed and remarks are offered.

The recent development of the Structural Health Monitoring (SHM) framework has required the simultaneous development of the tools fundamentals for its realization. The shape sensing methods and the loads identification ones have established themselves as crucial tools for the monitoring of aerospace structures. The two families of methods have been developed separately, although the possibility to achieve the knowledge of the displacement field and of the external loads together can enable a further progress in the SHM. In this paper, an integrated approach to simultaneously preform loads identification and shape sensing from discrete strain measurements is proposed. The methods is based on a two-steps process. The first step involves the identification of continuously distributed and concentrated external loads from discrete strain measurements. This step is achieved by discretizing the loads with Finite Elements (FE) and by computing the coefficients of influence between the nodal values of the loads and the discrete strain measurements. The second step reduces the shape sensing inverse problem to a simple direct Finite Element Analysis. The loads identified in the previous step are applied to a refined FE model of the structure and the displacement field is easily obtained through a direct FE analysis. This investigation proves that the two-steps method can accurately identify the loads and the displacement field of a wing box subject to an aerodynamic pressure distribution and a set of concentrated forces, when a sufficient number of strain information is available. When the number of discrete strain measurements decreases or the strains are affected by measurement error, the loads are poorly predicted but the method is still capable of an extremely accurate displacements reconstruction.

To generate more power, wind turbine rotors are growing in size and consequently, wind turbine tower are becoming increasingly taller and more flexible. As a result, fluid–structure interaction (FSI) of the flexible tower caused by strong wind is a very important phenomenon, and tower vibration must be carefully considered. In this paper, the physical model of the wind turbine tower is simplified appropriately, and then a multi-body dynamics model of wind turbine tower system is established based on Transfer Matrix Method of Multibody System (MSTMM). Compared with the data from finite element model (FEM) and field tests, the simulation results show that the model has a good accuracy. By coupling the mode shapes with two degrees of freedom (2-DOF) wake oscillator model, the dynamic responses of the flexible tower are computed. The influence of various foundation stiffness and top mass on tower vibration is studied systematically using this model. The results indicate that different boundary conditions can affect the maximum amplitude and displacement along the axis of the tower. This work provides a reference for dynamic modeling and simulation of high-rise flexible structure, and the prediction of the maximum amplitude of the tower vibration, which can be used for aeroelastic control purpose.

Purpose This paper aims to describe a methodology to optimize the trajectory of unconventional airship performing a high-altitude docking manoeuvre. Design/methodology/approach The trajectories are based upon Bezier curves whose control points positions are optimized through particle swarm optimization algorithm. A minimum energy strategy is implemented by considering the airship physical properties. The paper describes the mathematical model of the airships, the trajectories modelling through Bezier’s curves and the optimization framework. A series of test cases has been developed to evaluate the proposed methodology. Findings Results obtained show that the implemented procedure is able to optimize the airship trajectories and to support their in-flight docking; a strong influence of the wind speed and course on the trajectories planning is highlighted. Research limitations/implications The wind speed considered in these simulations depends only on altitude, and gusts effect has been neglected. Practical implications The proposed model can support the study of unconventional airship trajectories and can be useful to evaluate best in-air docking strategies. Originality/value The paper addresses the problem of trajectory optimization for a class of new air vehicles with an heuristic approach.

The vortex induced vibration (VIV) mitigation on a circular cylinder with low mass ratio (i.e. the ratio of structural to displaced fluid mass) parameter and moderate Reynolds numbers through a nonlinear energy sink (NES) is investigated numerically as basis for applications on dynamics of spar platform, marine tunnel or risers used in the ocean engineering industry. The Reynolds-Averaged-Navier–Stokes (RANS) equations and shear Stress transport (SST) k – ω turbulence model are used to calculate the flow field, while a fourth-order Runge–Kutta method is employed for evaluating the nonlinear structure dynamics of flow–cylinder–NES coupled system. The computational model includes an overset mesh solution which can avoid the negative volume grid problem as a dynamic mesh method is involved to deform the domain according to the motion of the fluid–structure interface. The numerical model is validated against experimental data of VIV of an isolated cylinder in uniform current. The study is aimed to investigate the effect of NES parameters, including mass, damping and nonlinear cubic stiffness, and various reduced velocities on the VIV response of the cylinder. The VIV amplitudes’ distribution along the various reduced velocities, trajectories of cylinder motion, hydrodynamic response, and temporal evolution of vortex shedding, are obtained by conducting numerical simulations. Subsequently, it is found that placing a NES with appropriate parameters inside the cylinder, can affect the distribution of the three-branch response and narrows the lock-in range. A good VIV amplitudes’ suppression can be achieved with a large NES mass, small nonlinear cubic stiffness, and the NES damping is higher in the critical range for this kind low mass ratio cylinder structures.

Low-fidelity fluid–structure interaction model of flexible propeller blades is assessed by means of comparison with high-fidelity aeroelastic results. The low-fidelity model is based on a coupled extended blade-element momentum model and non-linear beam theory which were both implemented in Matlab. High-fidelity fluid–structure interaction analysis is based on coupling commercial computational fluid dynamics and computational structural dynamics codes. For this purpose, Ansys CFX® and Ansys Mechanical® were used. Three different flexible propeller blade geometries are considered in this study: straight, backward swept, and forward swept. The specific backward and forward swept blades are chosen due to their aeroelastic response and its influence on the propulsive performance of the blade while a straight blade was selected in order to serve as a reference. First, the high-fidelity method is validated against experimental data available for the selected blade geometries. Then the high- and low-fidelity methods are compared in terms of integral thrust and breaking power as well as their respective distributions along the blades are compared for different advancing ratios. In a structural sense, the comparison is performed by analyzing the blade bending and torsional deformation. Based on the obtained results, given the simplicity of the low-fidelity method, it can be concluded that the agreement between the two methods is reasonably good. Moreover, an important result of the comparison study is an observation that the advance ratio is no longer a valid measure of similarity in the case of flexible propeller blades and the behavior of such blades can change significantly with changing operating conditions while keeping the advance ratio constant. This observation is supported by both high- and low-fidelity methods.

In this paper, a fully adaptive control design is considered for a four-degree-of-freedom aeroelastic system that has structural nonlinearities and operates in an unsteady aerodynamic incompressible flowfield. By using the flap hinge torque of a trailing-edge flap surface in combination with a leading-edge active flap, a closed-loop controller that adapts for uncertainties in the system parameters is designed. An implicit observer is implemented in the control design to compensate for lack of measurements of the lag states that model the unsteady flow. The innovative Lyapunov-based control design procedure results in a partial-state feedback adaptive controller that is globally asymptotically stable. Numerical simulation results show the effectiveness of the control strategy; comparative simulations are run to illustrate the benefit of using twin flaps as opposed to the conventional single flap control design.

International audience; A heaving and pitching two degrees of freedom bridge-deck sectional model apparatus representative of a long-span bridge was designed, built, and tested in the wind tunnel to acquire measurement of aeroelastic derivatives. The main objective of this study was the experimental validation of the procedure to compute the aeroelastic derivatives based on the Iterative Least Square Method identification technique. A good correspondence was found between the flutter derivatives experimentally extracted from the model, and the corresponding reference values, which have been analytically derived from the Theodorsen's theory. A simulation model in MATLAB/Simulink® environment, useful for future control laws applications, was also developed and validated by real experimental results.

The modern development in design of airships and aerostats has led to unconventional configurations quite different from the classical ellipsoidal and spherical ones. This new class of air-vehicles presents a mass-to-volume ratio that can be considered very similar to the density of the fluid displaced by the vehicle itself, and as a consequence, modeling and simulation should consider the added masses in the equations of motion. The concept of added masses deals with the inertia added to a system, since an accelerating or decelerating body moving into a fluid displaces a volume of the neighboring fluid. The aim of this paper is to provide designers with the added masses matrix for more than twenty Lighter Than Air vehicles with unconventional shapes. Starting from a CAD model of a given shape, by applying a panel-like method, its external surface is properly meshed, using triangular elements. The methodology has been validated by comparing results obtained with data available in literature for a known benchmark shape, and the inaccuracies of predictions agree with the typical precision required in conceptual design. For each configuration, a CAD model and a related added masses matrix are provided, with the purpose of assisting the practitioner in the design and flight simulation of modern airships and scientific balloons.

This paper presents the development of indicial functions (IFs) for two-dimensional bridge deck sections. A new set of IFs predicted for the cross section of the Great Belt Bridge (GBB) are discussed. Two approaches, based on the time- and frequency-domain descriptions, are applied. In time domain, IFs are determined by imposing an instantaneous change to a system state variable, e.g., the angle of attack (AOA) of air flow to the bridge section. In frequency domain, indicial representation is derived from the aeroelastic (or flutter) derivatives typically used to define the self-excited harmonic forces in an eigenvalue problem and generated by imposing a sinusoidal motion to the bridge section and performing a sweep in the frequency range of oscillations of the section. IFs are thus evaluated by exploiting the reciprocal relations that exist between them and the aerodynamic derivatives. To determine the aerodynamic response of a bridge cross section due to a step change in the AOA and to calculate the flutter derivatives from sinusoidal oscillations, the meshless discrete vortex method implemented in DVMFLOW ® is adopted. The results from the proposed work can be applied in the development of reduced-order models (ROM) of aerodynamic loads suitable to investigate fluid–structure interaction (FSI) problems associated with practical analysis of wind effects on long-span bridges, including phenomena such as flutter, vortex-induced vibration, buffeting and galloping. The IFs reported in this paper illustrate the importance of flow separation and vortex shedding and their dependence on the magnitude of the change in a state variable such as, in particular, the AOA for the cases reported herein.

The 3D wind and turbulence characteristics of the atmospheric boundary layer experiment (3D Wind) was conducted to evaluate innovative remote sensing and in situ platforms for measurements of wind and turbulence regimes. The experiment is part of a planned series that focuses on quantifying wind and turbulence characteristics at the scales of modern wind turbines and wind farms and was conducted in northern Indiana in May 2012. 3D Wind had the following specific objectives: (i) intercomparison experiments evaluating wind speed profiles across the wind turbine rotor plane from traditional cup anemometers and wind vanes on a meteorological mast and from a tethered balloon, sonic anemometers (mast mounted and on an unmanned aerial vehicle), three vertical-pointing (continuous wave) lidars and a pulsed scanning lidar, and (ii) integrate these measurements and output from 3-km-resolution (over the inner domain) simulations with the Weather Research and Forecasting Model to develop a detailed depiction of the a...

This paper introduces a novel semi-analytical representation of a displacement-based finite element model reduced via nested polynomials obtained through fitting of modal data. This method, termed Poly-SAFE (Polynomial Semi-Analytical Finite Element), is particularly suitable for modelling thin-walled composite structures subject to recursive analyses under varying loads, a common scenario in fluid–structure interaction (FSI) and Progressive Failure Assessment (PFA). The resulting functionals, i.e. polynomials inside polynomials, can be evaluated in an analytical fashion to yield displacements at arbitrary positions not limited to typical finite element grid nodes. These functionals remain virtually load-independent, allowing a Poly-SAFE model to be constructed without previous knowledge of magnitude, direction and location of applied loads, either static or dynamic. In this paper the theoretical framework of the Poly-SAFE method is presented in some detail, followed by an application of the theory to an extruded airfoil-shaped, laminated thin-walled beam subject to static loads. The displacement field captured by the new method is compared to the predictions of its associated finite element model, showing an excellent overall agreement. Finally, the advantages of Poly-SAFE against FE models in specific analyses and contexts are emphasised.

This paper provides insights into the stability of flexible orthotropic rectangular plates in supersonic flow by revealing the amplitude-speed dependency behavior for such panels in precritical and postcritical flight conditions. Flutter and postflutter behaviors are discussed to highlight the important effect of the structural and aerodynamic nonlinearities inherently present for panels in high-speed flows. The influence of transversal shear on stability is also illustrated. Depending on the plate’s geometrical parameters, soft- and hard-type flutter behaviors are possible, and a graphical interpretation is provided along with pertinent conclusions.

The mathematical model and experimental verification of flexible propeller blades are presented in this paper. The propeller aerodynamics model is based on an extended blade-element momentum model, while the Euler–Bernoulli beam theory and Saint–Venant theory of torsion are used to account for bending and torsional deformations of the blades, respectively. The proposed blade-element momentum model extends the standard blade-element momentum theory with the aim of providing a quick and robust model of propeller action capable of treating high-aspect-ratio propeller blades with a blade axis of arbitrary geometry. Based on the proposed mathematical model, a static flexible propeller blade design procedure and its associated analysis algorithm are established. Dynamic aeroelastic phenomena like propeller flutter and divergence are not covered by the presented mathematical model, design, and analysis algorithm. Experimental validation was carried out with an objective of evaluating the performance of the devel...

The nonlinear aeroelastic modeling and the post-flutter behavior of high-altitude long-endurance wings are discussed. A parametric structural model of wings based on an exact kinematic approach is formulated and coupled with an incompressible unsteady aerodynamic model that is obtained via an indicial formulation accounting for viscous effects, including dynamic stall and flow separation. To this end, a modified Beddoes-Leishman model is employed, and the equations of motion, together with the equations governing the aerodynamic states, are obtained via a total Lagrangian formulation. The critical and post-critical dynamic aeroelastic response is evaluated, and the limit cycles occurring in the post-flutter condition past the Hopf bifurcation are studied. Together, with comparisons from the available data of an experimental wing model with tip store, the effects of the unsteady loads and dynamic stall are evaluated and compared with predictions obtained from a model using a classical quasi-steady aerodyna...

The aim of this work is to illustrate a new wing morphing concept and to compare its performance with the conventional wing-flap configuration. While there is no formal definition for the word "morphing", it is usually considered to mean large shape change or transfiguration. In the field of aeronautics, "shape morphing" has been used to identify those aircraft that undergo significant geometrical changes to enhance or adapt to their mission profiles. This investigation aims to find out more about non conventional solutions over the classical wing structural design. The main goal in this research is to replace the ordinary wing structure adopted on a model aircraft with a new morphing wing. In this way, aerodynamic performances during different flight conditions may be improved thanks to the prevention of typical aerodynamic losses caused by geometrical discontinuities in conventional design.

A modern class of micro aerial vehicles (MAVs) uses deformable membrane wings (DMWs), which comprise rigid or rigidizable structural elements that guarantee stiffness, as well as flexible membranes that provide an aerodynamic shape, all in a compact and lightweight form. To contribute to the body of knowledge about DMWs, this paper presents extensive numerical simulations, accompanied by experiments that substantiate the findings pertinent to the aerodynamics and static aeroelastic behavior of these unconventional wings. The performance characteristics of canonical, untapered, and untwisted rectangular DMWs are studied using a parametric physical-based two-dimensional (2D) fluid structure interaction (FSI) model adopting a two-way coupled FSI algorithm. The effect of geometrical, material, and flow properties is investigated, and the results obtained show that the material Young’s modulus is the most influential parameter affecting the aerodynamic characteristics of the studied DMWs, with their hi...

The model-free control of aeroelastic vibrations of a non-linear 2-D wing-flap system operating in supersonic flight speed regimes is discussed in this paper. A novel continuous robust controller design yields asymptotically stable vibration suppression in both the pitching and plunging degrees of freedom using the flap deflection as a control input. The controller also ensures that all system states remain bounded at all times during closed-loop operation. A Lyapunov method is used to obtain the global asymptotic stability result. The unsteady aerodynamic load is considered by resourcing to the non-linear Piston Theory Aerodynamics (PTA) modified to account for the effect of the flap deflection. Simulation results demonstrate the performance of the robust control strategy in suppressing dynamic aeroelastic instabilities, such as non-linear flutter and limit cycle oscillations.

yields a semiglobal asymptotic stability result, leading to rapid suppression of the plunging and pitching motions. Numerical simulation results demonstrate the performance of the multi-input/multi-output continuous robust control toward suppressing aeroelastic vibration and limit cycle oscillations at pre- and postflutter flight-speed regimes, even in the presence of bounded unknown external disturbance.

In this paper, robust adaptive control design problem is addressed for a class of parametrically uncertain aeroelastic systems. A full-state robust adaptive controller was designed to suppress aeroelastic vibrations of a nonlinear wing section. The design used leading and trailing edge control actuations. The full state feedback (FSFB) control yielded a global uniformly ultimately bounded result for two-axis vibration suppression. The pitching and plunging displacements were measurable; however, the pitching and plunging rates were not measurable. Thus, a high gain observer was used to modify the FSFB control design to become an output feedback (OFB) design while the stability analysis for the OFB control law was presented. Simulation results demonstrate the efficacy of the multi-input multi-output control toward suppressing aeroelastic vibrations and limit cycle oscillations occurring in pre- and post-flutter velocity regimes.

Active aeroelastic control is an emerging technology aimed at providing solutions to structural systems that under the action of aerodynamic loads are prone to instability and catastrophic failures, and to oscillations that can yield structural failure by fatigue. The purpose of the aeroelastic control among others is to alleviate and even suppress the vibrations appearing in the flight vehicle subcritical flight regimes, to expand its flight envelope by increasing the flutter speed, and to enhance the postflutter behavior usually characterized by the presence of limit cycle oscillations. Recently adaptive and robust control strategies have demonstrated their superiority to classical feedback strategies. This review paper discusses the latest development on the topic by the authors. First, the available control techniques with focus on adaptive control schemes are reviewed, then the attention is focused on the advanced single-input and multi-input multi-output adaptive feedback control strategies developed for lifting surfaces operating at subsonic and supersonic flight speeds. A number of concepts involving various adaptive control methodologies, as well as results obtained with such controls are presented. Emphasis is placed on theoretical and numerical results obtained with the various control strategies.

The next generation of slender, flexible aircraft wings requires extremely lightweight structures capable of carrying a considerableamountof non-structuralweight.With the increased slenderness and flexibility, possible with the advent of advanced composites, these wings can exhibit aeroelastic instabilities quite different from their rigid counterparts. The design of highly flexible aircraft, such as high-altitude long endurance (HALE) configurations, must include phenomena that are not usually considered in traditional aircraft design, such that an alternative design philosophy has been proposed for this class of vehicles. The discussion in this article is restricted, among the various aeroelastic phenomena, to the flutter condition. Classical procedures usually refer to aero-structural systems where the undeformedstate is taken as the reference point. This is not the case with slender wing configuration where, due to the high structural flexibility, a proper beam model, capable of describing the structural flight deflections, should be adopted. Consequently, the flutter analysis has to be performed considering the deflected state as a reference point. Herein, an approximate procedure is proposed and flutter is evaluated by means of Galerkin's approach applied to the perturbed small motions of the aero-structural system. The effect of typical parameters, including stiffness ratio, mass eccentricity, store pod, deflection amplitude, as well as the wing aspect ratio, are considered. For a simplified wing configuration, comparisons between analytical and experimental findings are presented along with discussions and suggestions for new design criteria. Keywords: advanced aircraft design, linear/non-linear aeroelasticity, flutter, experimental aeroelasticity

Expressions for the bispectral density functions for multi-degree-of-freedom spring–mass–damper systems possessing quadratic nonlinearities and subject to Gaussian excitation are derived. The derivation uses a Volterra-series model for the system response and yields expressions for both auto-spectra, where the output of only one degree-of-freedom is used, and cross-spectra, where the bispectral density contains multiple output-response time series. The proposed formulation is used to identify the presence and location of quadratic nonlinearities in multiple degree-of-freedom systems. Results show that the ability to detect and localize nonlinearity is heavily dependent on which particular bispectral density is utilized.

The aeroelastic modeling and instability of shear deformable swept wings under roll angular velocity is investigated. The structural wing model was originally developed by Librescu and consists of non-classical effects such as warping inhibition and transverse shear flexibility. This model is used to study divergence and flutter instabilities when the aircraft wing is subjected to a roll moment created during a maneuver. The aeroelastic governing equations and boundary conditions are determined via Hamilton’s variational principle. The resulting partial differential equations are transformed into a set of eigenvalue/boundary value equations through the Extended Galerkin approach and solved by numerical integration. The effects of roll angular velocity, sweep angle, and wing aspect ratio on divergence and flutter speed are presented for classic and shear deformable wings. Validations of selected results against the previous publications are also supplied. Results indicate that roll angular velocities have a significant influence on the static and dynamic aeroelastic instability region.

Higher-order spectra have become a useful tool in spectral analysis, particularly for identifying the presence and type of system non-linearity. Two such spectra that have figured prominently in signal processing are the bispectrum and trispectrum. In a previous work, the authors developed an analytical solution for the bispectrum for multi-degree-of-freedom systems. Here this analysis is extended to the trispectrum. Specifically, an expression is developed for the trispectrum of a multi-degree-of-freedom system subject to Gaussian excitation applied at an arbitrary location. The analytical expression is compared to those obtained via estimation using the direct method.

To achieve efficient attenuation of noise and vibration characteristics within the helicopter environment, solid-state actuators are seen as one of the most promising smart technologies. The Smart Hybrid Active Rotor Control System project is expected to demonstrate the ability of smart structure systems, employing multiple active material actuators, sensors, and closed-loop controllers, to reduce simultaneously rotorcraft vibration and noise. Within this project, piezoceramic elements were used as actuators to vary the dry friction and the stiffness of the whole helicopter blade system. This active control concept, named Smart Spring, originated a prototype used to demonstrate the ability to attenuate vibrations. Before testing the Smart Hybrid Active Rotor Control System in operative conditions, the dynamic properties of the Smart Spring installed on a nonrotating helicopter blade are investigated. The effects of the Smart Spring actuator on the modal properties are studied through experimental activities carried out at Carleton University. Furthermore, the capability of the Smart Spring to change the dynamic behavior of the helicopter blade is demonstrated by analyzing the shifts in the modal parameters. Finally, a beam finite element model of the blade, with stiffness and mass properties tuned to the experimental structure, is presented.