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Purpose The application of reduced order models (ROMs) in the aerodynamic/aeroelastic analysis of long-span bridges, unlike the aeronautical structures, has not been extensively studied. ROMs are computationally efficient techniques, which have been widely used for predicting unsteady aerodynamic response of airfoils and wings. This paper aims to discuss the application of a reduced order computational fluid dynamics (CFD) model based on the eigensystem realization algorithm (ERA) in the aeroelastic analysis of the Great Belt Bridge (GBB). Design/methodology/approach The aerodynamic impulse response of the GBB section is used to construct the aerodynamic ROM, and then the aerodynamic ROM is coupled with the reduced DOF model of the system to construct the aeroelastic ROM. Aerodynamic coefficients and flutter derivatives are evaluated and compared to those of the advanced discrete vortex method-based CFD code. Findings Results demonstrate reasonable prediction power and high computational efficiency of the technique that can serve for preliminary aeroelastic analysis and design of long-span bridges, optimization and control purposes. Originality/value The application of a system identification tool like ERA into the aeroelastic analysis of long-span bridges is performed for the first time in this work. Authors have developed their earlier work on the aerodynamic analysis of long-span bridges, published in the Journal of Bridge Engineering, by coupling the aerodynamic forces with reduced DOF of structural system. The high computational efficiency of the technique enables bridge designers to perform preliminary aeroelastic analysis of long-span bridges in less than a minute.

Curvilinear stiffened panels are being developed for aerospace structures and other applications. This paper presents an experimental and numerical study into the vibration response of three curvilinear stiffened square plates clamped at the edges. The experimental modal analysis was performed using laser Doppler vibrometry (LDV) and an impact hammer test. Two of the three plates was constructed with curvilinear stiffener geometry, while the third was a straight stiffened plate with the stiffeners oriented at an angle to the edge of the plate. Even though the natural frequencies of the plates are similar, the different stiffening patterns was sufficient in providing unique directional properties and therefore distinct mode shapes. The numerical modal analysis of the plates was performed using finite element analysis (FEA). A comparison of the experimental and numerical results was carried out in terms of natural frequencies and mode shapes, using relative differences and modal assurance criterion (MAC), respectively. The experimental and numerical results were in good agreement for all the three plates. The difference between experimental and numerical natural frequencies was typically less than 5% and the diagonal MAC values typically ranged from 0.8 to 1. This is in line with previously published results in the literature. The reasons for differences between the experimental and numerical results, and the practical significance of the findings, are also discussed.

An accurate aeroelastic assessment of powered HALE aircraft is of paramount importance considering that their behaviour contrasts the one of conventional aircraft mainly due to the use of high aspect-ratio wings with distributed propulsion systems. This particular configuration shows strong dependency of the wing natural frequencies to the propulsion distribution and operating conditions. Numerical and experimental investigations are carried out to better understand the behaviour of flexible wings, focusing on the effect of distributed electric propulsion systems. Several configurations are investigated, including a single propulsion system using an engine pod (a weight with embedded electric motor, a propeller, and the wing-attached structure) installed at selected spanwise positions, and configurations with two and three propellers.

Panels of reentry vehicles are subjected to a wide range of flow conditions during ascent and reentry phases. The flow can vary from subsonic continuum flow to hypersonic rarefied flow with wide ranging dynamic pressure and associated aerodynamic heating. One of the main design considerations is the assurance of safety against panel flutter under the flow conditions characterized by sever thermal environment. This paper deals with supersonic/hypersonic flutter analysis of panels exposed to a temperature field. A 3-D rectangular plate element of variable thickness based on absolute nodal coordinate formulation (ANCF) has been developed for the structural model and subjected to an assumed thermal profile that can result from any residual heat seeping into the metallic panels through the thermal protection systems. A continuum mechanics approach for the definition of the elastic forces within the finite element is considered. Both shear strain and transverse normal strain are taken into account. The aerodynamic force is evaluated by considering the first-order piston theory to linearize the potential flow and is coupled with the structural model to account for pressure loading. A provision is made to take into account the effect of arbitrary flow directions with respect to the panel edges. Aerothermoelastic equations using ANCF are derived and solved numerically. Values of critical dynamic pressure are obtained by a modal approach, in which the mode shapes are obtained by ANCF. A detailed parametric study is carried out to observe the effects of different temperature loadings, flow angle directions, and aspect ratios on the flutter boundary.

A reduced-order finite-element model suitable for Progressive Failure Analysis (PFA) of composite structures under dynamic aeroelastic conditions based on a Thin-Walled Beam (TWB) formulation is presented. Validation of the PFA-TWB against an integrated PFA model based on a shell formulation and implemented in the commercial software tool GENOA is conducted for static load conditions. A helicopter blade made from composite material and previously used in literature for the discussion of damage propagation is used as the reference case. The failure criteria for the different layers of the composite material used in the PFA-TWB model have been formulated in analogy with the corresponding criteria implemented in the shell formulation. Comparisons between the predictions of both models for progressively increasing load have been conducted in terms of the cumulative overall damage volume in the thin-walled structure, the layer-resolved cumulative damage volume, as well as through spatially resolved damage maps for both models. A strikingly similar damage topology has been found from both models up to load values close to final failure, in spite of the restraining assumptions of the TWB formulation. In terms of damage volume the PFA-TWB models predicts slightly higher values which can be traced back to the inevitable differences in the failure criteria formulation in the one-dimensional and the shell model, respectively. It is shown that a good agreement with the predictions of the shell model in terms of the cumulative damage volume is obtained if the strength values of the composite material are adjusted upwards in a uniform manner by about 10%. Considering the common safety factors usually applied in the design process of composite material the agreement of the TWB and the shell model in terms of damage propagation is considered excellent, allowing for the PFA-TWB to be used in systematic design studies.

This paper deals with supersonic/hypersonic flutter analysis of panels. A 3-D rectangular plate element of variable thickness based on absolute nodal coordinate formulation (ANCF) has been developed for the structural model. A continuum mechanics approach for the definition of the elastic forces within the finite element is considered. Both shear strain and transverse normal strain are taken into account. The aerodynamic force is evaluated by considering the first-order piston theory to linearize the potential flow and is coupled with the structural model to account for pressure loading. A provision is made to take into account the effect of arbitrary flow directions with respect to the panel edges. Aeroelastic equations using ANCF are derived and solved numerically. Values of critical dynamic pressure are obtained by a modal approach, in which the mode shapes are obtained by ANCF. A detailed parametric study is carried out to observe the effects of different flow angle directions and aspect ratios on the flutter boundary.

The prediction of damage progression in composite wind turbine blades, especially under dynamic aeroelastic conditions, is usually a cumbersome multi-step process with significant manual user intervention. In this paper a novel approach is presented where the different components of this process – dynamical structural analysis under varying aerodynamic and deterministic loads, and damage progression – are integrated into one reduced-order model capable of predicting the occurrence and progression of damage in real time. Key to this integration is the use of an effective one-dimensional model of the turbine blade known as thin-wall beam model, which allows for the reconstruction of a three-dimensional stress field of a volume given by the blade. This stress field can then be used to assess damage and locally modify the structural properties to account for the presence of damage, leading to a reduced load carrying capacity. The model was previously tested in its components, demonstrating a good agreement of the predicted structural and static damage progression behaviour compared to detailed high-order finite-element models of the same blade. Once validated, the model was applied to severe load cases and the potential for real-time predictions of damage progression was demonstrated.

This paper presents a numerical validation of a thin-walled beam (TWB) finite element (FE) model of a realistic wind turbine rotor blade. Based on the theory originally developed by Librescu et al. and later extended to suit FE modelling by Phuong, Lee and others, this computationally efficient yet accurate numerical model is capable of capturing most of the features found in large blades including thin-walled hollow cross section with variable thickness along the section's contour, inner reinforcements, arbitrary material layup and non-linear anisotropic fibre-reinforced composites; the present application is, for the time being, restricted to linearity. This one-dimensional (1D) FE model allows retaining information of different regions of the blade's shell and therefore approximates the behaviour of more complex three-dimensional (3D) shell or solid FE models more accurately than typical 1D FE beam models. A 9.2 m rotor blade, previously reported in specialized literature, was chosen as a case study to validate the static and dynamic behaviour predicted by a TWB model against an industry-standard 3D shell model built in a commercial software tool. Given the geometric and material complexities involved, an excellent agreement was found for static deformation curves, as well as a good prediction of the lowest frequency modes in terms of resonance frequencies, mode shapes and frequency response functions; the highest (sixth) frequency mode shows only a fair agreement as expected for an FE model. It is concluded that despite its simplicity, a TWB FE model is sufficiently accurate to serve as a design tool for the recursive analyses required during design and optimization stages of wind turbines using only readily available computational tools

This paper presents the nonlinear analysis of functionally graded curved panels under high temperature supersonic gas flows. The aerothermoelastic governing equations are determined via Hamilton's variational principle. The von Karman nonlinear strain–displacement relations are used to account for large deflections. The material properties are assumed to be temperature-dependent and varying through the thickness direction according to a power law distribution in terms of the volume fractions of the constituent components. The panel is assumed to be infinitely long and simply supported. The Galerkin method is applied to convert the partial differential governing equation into a set of ordinary differential equations and the resulting system of nonlinear equations is solved through a numerical integration scheme. The effects of volume fraction index, curved panel height-rise, and aerodynamic pressure, in conjunction with the applied thermal loading, on the dynamical behavior of the panel are investigated. Regular and chaotic motions regime are determined through bifurcation analysis using Poincaré maps of maximum panel deflection, panel time history, phase-space and frequency spectra as qualitative tools, while Lyapunov's exponents and dimension are used as quantitative tools.

The robust aeroelastic control of a three-degrees-of-freedom (3DOF) linear and non-linear wing-flap system under sliding mode control (SMC) and operation in supersonic flight speeds is presented. Open- and closed-loop aeroelastic responses to blast and sonic-boom excitation in the wing-flap system with uncertainty, as well as flutter analysis, are investigated along with the implementation of a Sliding Mode Observer (SMO). The effectiveness of control in reducing the amplitude of oscillation and preventing flutter instability is demonstrated.

Aeroelastic instabilities are dangerous phenomena, where aerodynamic load interacting with the inertia and elastic structural loads can induce catastrophic failures. In this paper the effects of aerodynamic nonlinearities as well as coupled plunging/pitching structural concentrated cubic type and freeplay nonlinearities in the dynamic of a two-dimensional double-wedge airfoil immersed in supersonic/hypersonic flow has been examined. The unsteady nonlinear aerodynamic force and moment on the airfoil are evaluated using the Piston Theory Aerodynamics modified to take into account the effect of the airfoil thickness. The resulting aeroelastic equations are numerically integrated to obtain time responses and to investigate the dynamic instability of the lifting surface under various initial displacement conditions. Results of the complex nonlinear aeroelastic system are presented in the form of bifurcation diagrams constructed from the response amplitude for various types of the system nonlinearity. It is shown that there exist regions, in which the system exhibit Limit Cycle Oscillations (LCOs), strongly dependent on the initial conditions of the aeroelastic system. Concentrated structural nonlinearities, that are freeplays and cubic type nonlinearities, can have significant effects on the flutter behavior and can cause large amplitude oscillations at lower airspeeds than for a linear system. It is also shown that larger amplitude LCOs occur when a pitching freeplay is considered, as compared with the case when a plunging freeplay is taken into account.

In this paper, the problem of implications of time delay feedback control of a two-dimensional supersonic lifting surface on the flutter boundary and on its character, that is, benign or catastrophic, is addressed. In this context, the structural and aerodynamic nonlinearities are included in the aeroelastic governing equations. The model and the associated theory are developed for linear and nonlinear plunging and pitching full-state proportional and velocity feedback controls. Center manifold reduction and normal form theory are applied to investigate the stability in the post-flutter flight speed regimes. Numerical simulations are carried out to determine the implications of time delay in the considered controls, but are restricted to the cases of proportional feedback control and no structural damping.

The aim of this paper is to develop a new concept of unconventional airship based on morphing a lenticular shape while preserving the volumetric dimension. Lenticular shape is known to have relatively poor aerodynamic characteristics. It is also well known to have poor static and dynamic stability after the certain critical speed. The new shape presented in this paper is obtained by extending one and reducing the other direction of the original lenticular shape. The volume is kept constant through the morphing process. To improve the airship performance, four steps of morphing, starting from the lenticular shape, were obtained and compared in terms of aerodynamic characteristics, including drag, lift and pitching moment, and stability characteristics for two different operational scenarios. The comparison of the stability was carried out based on necessary deflection angle of the part of tail surface. The comparison results indicated that new shape concept possesses much better aerodynamic and stability characteristics and could be used for detailed optimisation studies.

This paper introduces the Seabus SB-8, a new Wing-In-Ground-Effect (WIGE) craft designed for 8 - 10 passengers. The craft will be used for fast transportation across Port Phillip Bay in Melbourne, Australia. With a cruise speed of about 140 km/hr, it

raditional User/Maintenance Manuals provide useful information when dealing with simple machines. However, when dealing with complex systems of systems and highly miniaturized technologies, like UAVs, or with machines with millions of parts, a commercial aircraft is a case in point, new technologies taking advantage of Augmented Reality can rapidly and effectively support the maintenance operations. This paper presents a User/Maintenance Manual based on Augmented Reality to help the operator in the detection of parts and in the sequence to be followed to assemble/disassemble systems and subsystems. The proposed system includes a handheld device and/or an head mounted display or special goggles, to be used by on-site operators, with software management providing data fusion and overlaying traditional 2D user/maintenance manual information with an augmented reality software and appropriate interface. This device is connected by internet to a maintenance centre located in the aircraft manufacturer facilities. The on-site operator can directly access to multimedia content and historical information and can be helped or guided remotely by expert engineers residing at the manufacturer company offices. This resource may exploit Computer Aided Design and Product Data Management PDM remote facilities to prepare additional and specific 3D graphic content, supported also by a video and audio streaming from the camera and microphone of the on-site operator's handheld device. The proposed solution has revealed a number of significant advantages compared to the currently used operations: there is no need for preparing animations and graphic content for all the required maintenance sequences. The expert engineers and designers can both be involved directly in the maintenance tasks, a useful mean of feedback to evaluate the design for further projects or for project improvement. Additionally, the sensitive data is not shared outside the company since data is transmitted for visual display but it is stored on a secured location.

The implementation of Synthetic Jet Actuators (SJAs) on Unmanned Aerial Vehicles (UAVs) provides a safe test-bed for analysis of improved performance, in the hope of certification of this technology on commercial aircraft in the future. The use of high resolution numerical methods (i.e. CFD) to capture the details of the effects of SJAs on flows and on the hosting lifting surface are computationally expensive and time-consuming, which renders them ineffective for use in real-time flow control implementations. Suitable alternatives include the use of Reduced Order Models (ROMs) to capture the lower resolution overall effects of the jets on the flow and the hosting structure. This research paper analyses the effects of SJAs on aircraft wings using a ROM for the purpose of determining the unsteady aerodynamic forces modified by the presence of the SJAs. The model developed is a 3D unsteady panel code where the jets are represented by source panels. This code has the ability to model thick 3D trapezoidal wings with sweep and dihedral, accepting any aerofoil coordinate file. Multiple rows of SJAs can also be applied where desired. In order to validate the panel code, comparisons are made with past experimental data and other numerical methods like the vortex lattice method and CFD. There is avenue for further work in implementing a robust control architecture around SJA technologies, which in the future could serve as replacements for traditional hinged flight control surfaces and flutter suppression solutions.

The maturity reached in the development of Unmanned Air Vehicles (UAVs) systems is making them more and more attractive for a vast number of civil missions. Clearly, the introduction of UAVs in the civil airspace requiring practical and effective regulation is one of the most critical issues being currently discussed. As several civil air authorities report in their regulations “Sense and Avoid” or “Detect and Avoid” capabilities are critical to the successful integration of UAV into the civil airspace. One possible approach to achieve this capability, specifically for operations beyond the Line-of-Sight, would be to equip air vehicles with a vision-based system using cameras to monitor the surrounding air space and to classify other air vehicles flying in close proximity. This paper presents an image-based application for the supervised classification of air vehicles. First, several vehicle images, taken from different points of view, are transformed using a descriptor of salient features as to build the five-class database used to train the classification algorithm. Then, the latter compares the descriptor of a vehicle image taken from a random point of view to records in the database. With a positive match, the vehicle will be assigned to one of the following classes: a) civil transport aircrafts, b) military aircrafts, c) general aviation aircrafts, d) helicopters, and e) airships/hot air balloons. The paper provides a possible layout for the algorithm implementation and presents the outcome of several tests performed to evaluate its efficiency and possible exploitation. Indications useful to further studies are presented to help future researches.

In this paper the finite element model of an Unmanned Aerial Vehicle is updated by using experimental data coming from a standard ground vibration test in order to improve the numerical-experimental correlation. A sensitivity-based updating methodology that iteratively minimizes a residual vector, defined on the modal parameters (e.g. natural frequencies and mode shapes), is considered to identify the unknown values of the updating parameters. The structure under investigation is the Clarkson University Golden Eagle UAV. An initial numerical model of the structure is obtained by assembling the individual components previously updated which included wings, fuselage, horizontal tail, vertical tails and tail booms. As a result the identification procedure shifts its focus on the joints between UAV elements which could not be modeled accurately in earlier investigations.

The model of a geometrically nonlinear wing hosting piezoelectric patches with the dual purpose of suppressing aeroelastic vibration and harvesting vibrational energy is presented in this paper. The nonlinearities are introduced in order to consistently reproduce the behavior of the flexible structure, since moderate to large displacements can occur in response of external loading conditions. A nonlinear shear underfomable 3-D Euler-Bernoulli beam theory is used to model the displacements field and structural nonlinearities up to the third order are retained in the model of a straight untapered composite wing. A linear indicial functional representation of the unsteady aerodynamic loads in an incompressible flowfield is adopted. The extended Hamilton principle is used to derive the aeroelastic equations of motion. The composite cantilever wing includes two piezoelectric elements, perfectly bonded on its lower and upper longitudinal surfaces in the proximity of the wing root, and electrically connected by a resistive load, functioning as energy harvesting devices. During the state of deformation the piezoelectric components induce electric charges to be stored for future use as a supplementary power source. The piezoelectric layers also function as damping elements with desired load alleviation properties. The effectiveness of such a solution, both in terms of the amount of energy harvested and load alleviation characteristics, for a well defined wing configuration have been evaluated in this paper. Numerical results and discussions are followed by pertinent conclusions and directions for future work.

The innovative highly flexible wings made of extremely light structures, yet still capable of carrying a considerable amount of non- structural weights, requires significant effort in structural simulations. The complexity involved in such design demands for simplified mathematical tools based on appropriate nonlinear structural schemes combined with reduced order models capable of predicting accurately their aero-structural behaviour. The model presented in this paper is based on a consistent nonlinear beam-wise scheme, capable of simulating the unconventional aeroelastic behaviour of flexible composite wings. The partial differential equations describing the wing dynamics are expanded up to the third order and can be used to explore the effect of static deflection imposed by external trim, the effect of gust loads and the one of nonlinear aerodynamic stall. As to provide a rationale evaluation of the important nonlinear contributions in aeroelastic wing simulations, the aeroelastic governing equations in linear, quadratic, and cubic forms, are considered and compared. Critical conditions will be investigated. The governing equations will be solved based on a discretization technique, along with Galerkin's method starting with a reduced order procedure and passing through a multi-modal approach. Three degrees of freedom in edgewise, flapwise, and torsion, are needed to describe efficiently the dynamics of the wing. Interesting design indicators will be highlighted.

The flight simulation of airships and hot air balloons usually considers the envelope geometry as a fixed shape, whose volume is eventually reduced by ballonets. However, the dynamic pressure or helium leaks in airships, and the release of air to allow descent in hot air balloons can significantly change the shape of the envelope leading to potential dangerous situations. In fact, in case of semi-rigid and non-rigid airships a reduction in envelope internal pressure can reduce the envelope bending stiffness leading to the loss of the typical axial-symmetric shape. For hot air balloons thing goes even worse since the lost of internal pressure can lead to the collapsing of the balloon shape to a sort of vertically stretched geometry (similar to a torch) which is not able to sustain the attached basket and its payload. These effect should be considered in simulations, however to compute in real time the envelope shape with Finite Element Methods is a complex and demanding task due to the high deformations, complex fabric model, and wrinkling effects. A possible solution to overcome this problem is to apply a Cloth Simulation Technique (CST) to the prediction of the envelope behaviour. This paper describes how such a model can be implemented for airship envelops and hot air balloons shape predictions. Appropriate algorithms have been developed in Matlab® and validation test have been conducted. Results show that this model can provide qualitatively good results, in agreement with the experience and the physics of the problem.

Flow control over aerodynamic shapes in order to achieve performance enhancements has been a lively research area for last two decades. Synthetic Jet Actuators (SJAs) are devices able to interact actively with the flow around their hosting structure by providing ejection and suction of fluid from the enclosed cavity containing a piezo-electric oscillating membrane through dedicated orifices. The research presented in this paper concerns the implementation of zero-net-mass-flux SJAs airflow control system on a NACA0015, low aspect ratio wing section prototype. Two arrays with each 10 custom-made SJAs, installed at 10% and 65% of the chord length, make up the actuation system. The sensing system consists of eleven acoustic pressure transducers distributed in the wing upper surface and on the flap, an accelerometer placed in proximity of the wing c.g. and a six-axis force balance for integral load measurement. A dSPACE TM hardware connected to the software environment Matlab/Simulink® and dSPACE Control-Desk® complete the test architecture. Wind tunnel experiments, on the uncontrolled wing (actuators off), are primarily performed for system identification purpose. The open-loop control operation (actuators on but no feedback) of the wing is implemented and tested, obtaining a stall delay of about 2.8 degrees of AOA. Furthermore, a closed-loop strategy, based on the wing upper surface mean pressure chord-wise distributions signature is adopted to characterize the forthcoming boundary layer detachment. This allows for triggering the controller in stall proximity only, for energy saving purpose. Pertinent results and discussion are provided along with concluding remarks and prospects for future research.

The study of the magnetoelastic vibrations of a flat plate immersed in a uniform applied external magnetic field is presented. Kirchhoff’s plate theory and the model of a perfect conductive medium are used. The conditions for the existence of localized bending vibrations in the vicinity of the free edge of the plate are established. It is shown that the localized vibrations can be detected and eventually can be eliminated by means of an applied magnetic field.

This paper concerns the active aeroelastic control of 3-D flap-wing systems exposed to gust and/or sonic boom in an incompressible flow field. This is achieved via implementation of a robust estimation capability (sliding mode observer: SMO), and of the use of the deflected flap as to suppress the flutter instability or enhance the subcritical aeroelastic response to gust or blast loads. To this end, a control methodology using LQG(Linear Quadratic Gaussian) in conjunction with SMO is implemented, and its performance toward suppressing flutter and reducing the vibrational level in the subcritical flight speed range is demonstrated. Moreover, its performances are compared with the ones provided via implementation of conventional LQG with Kalman filter.

Thisworkwas partially supported byKorea Research Foundation Grant (KRF-2004-003-D00067). Dong-Hyun Kim would like to acknowledge the support.

This paper concerns the active aeroelastic control of 2-D wing-flap systems operating in an incompressible flowfield and exposed to a blast pulse. The goal is to implement an active flap control capability to suppress the flutter instability and enhance the subcritical aeroelastic response to time-dependent external pulses. To this end, a combined control law is implemented and its performances toward suppressing flutter and reducing the vibrational level in the subcritical flight speed range is demonstrated.

A study of the effects of time delay on the linear/nonlinear feedback control of two-dimensional lifting surfaces in an incompressible flowfield is presented. Specifically, the case of a one-degree-of-freedom system is considered in detail, and in that context, both the structural and the unsteady aerodynamics models are assumed to be linear. The study of the stability/instability behavior of nonlinear feedback time-delay closed-loop aeroelastic systems

Determination of the nature of the critical flutter boundary (benign/catastrophic) and its control constitute important issues that can be addressed within the nonlinear formulation of lifting surface theory. The main attention of this paper consists in the development of a computational approach enabling one to get a better understanding on time-delayed dynamics as applied to this important aeroelastic problem, and more specifically, to two-dimensional supersonic lifting surfaces. The analysis is based on the reduction of the infinite-dimensional problem to one described on a two-dimensional center manifold. Results presenting the implication of the linear/nonlinear timedelayed feedback control on two-dimensional supersonic lifting surfaces are addressed, and pertinent conclusions are drawn.

An analysis of the flutter and postflutter behavior of infinitely long flat panels in a supersonic/hypersonic flowfield exposed to a high-temperature field is presented. In the approach to the problem, the thermal degradation of thermoelastic characteristics of the material is considered. A third-order piston theory aerodynamic model in conjunction with the von Karman nonlinear plate theory is used to obtain the pertinent aerothermoelastic governing equations. The implications of temperature, thermal degradation, and of structural and aerodynamic nonlinearities on the character of the flutter instability boundary are analyzed. As a byproduct, the implications of the temperature on the linearized flutter instability of the system are discussed. The behavior of the structural system in the vicinity of the flutter boundary is studied via the use of an encompassing methodology based on the Lyapunov First Quantity. Numerical illustrations, supplying pertinent information on the implications of the temperature field and of the thermal degradation are presented, and pertinent conclusions are outlined.

Determination of the subcritical aeroelastic response to arbitrary time-dependent external excitation and determination of the flutter instability of open/closed-loop two-dimensional nonlinear airfoils constitute the main topics. To address these problems, Volterra series and indicial aerodynamic functions are used, and, in the same context, the pertinent aeroelastic nonlinear kernels are determined. Flutter instability predictions obtained within this approach compared with their counterparts generated via the frequency eigenvalue analysis and via experiments reveal excellent agreements. Implications of a number of important parameters characterizing the lifting surface and control law on the aeroelastic response/flutter are discussed, and pertinent conclusions are outlined.

The present paper deals with a study of the benign and catastrophic characters of the flutter instability boundary of 2-D lifting surfaces in a supersonic flow field. The objectives of this work are: (i) to contribute to a better understanding of the implications of aerodynamic and physical non-linearities on the character of the flutter boundary and (ii), to outline the effects exerted in the same respect by some important parameters of the aeroelastic system. With the aim of addressing this problem, the method based on the First Liapunov Quantity is used to study the bifurcational behavior of the aeroelastic system in the vicinity of the flutter boundary. The expected outcomes of this study are: (a) to greatly enhance the scope and reliability of the aeroelastic analysis and design criteria of advanced aircraft and, (b) to provide a theoretical basis for the analysis of more complex non-linear aeroelastic systems.

A comprehensive aeroelastic model developed toward investigating the static divergence, e utter, and dynamic aeroelastic response of composite aerovehicle wings to sharp-edged gust and blast loads in supersonic e owe eld is presented. The aerovehicle wings are modeled as an anisotropic composite thin-walled beam structure featuring circumferentially asymmetric stiffness lay up that generates preferred elastic couplings. A number of nonclassical effects, such as transverse shear, warping restraint, and three-dimensional strain effects, are incorporated in the structural model. Based on the concept of two-dimensional indicial functions considered in conjunction with the aerodynamicstriptheoryextendedtothree-dimensionalwingmodel,theunsteadyaerodynamicloadsinsupersonic e ows are derived. The effect of elastic tailoring and the implications of transverse shear, warping restraint on divergence and dynamic response of selected wing cone gurations are investigated, and pertinent conclusions are outlined. Nomenclature AR = wing aspect ratio, L=b a.s/ = geometric quantity; see Eq. (2) and Fig. 3 aij = one-dimensional global stiffness coefe cients a1 = undisturbed speed of sound b;d = semichord and semidepth of the beam normal cross section, respectively bi = inertia coefe cient Eij = Young’ s modulus of orthotropic materials in the material coordinate system h.s/ = wall thickness as the function of the midline

A number of issues related to the flutter and postflutter of two-dimensional supersonic lifting surfaces are addressed. Among them there are the 1) investigation of the implications of the nonlinear unsteady aerodynamics and structural nonlinearities on the stable/unstable character of the limit cycle and 2) study of the implications of the incorporation of a control capability on both the flutter boundary and the postflutter behavior. To this end, a powerful methodology based on the Lyapunov first quantity is implemented. Such a treatment of the problem enables one to get a better understanding of the various factors involved in the nonlinear aeroelastic problem, including the stable and unstable limit cycle. In addition, it constitutes a first step toward a more general investigation of nonlinear aeroelastic phenomena of three-dimensional lifting surfaces.

A theoretical investigation of the flutter and postflutter of infinitely long thin-walled circular cylindrical panels in a supersonic/hypersonic flowfield is presented. In this context, third-order piston theory and shockwave aerodynamics are used in conjunction with the geometrically nonlinear shell theory to obtain the pertinent aeroelastic governing equations. The effects of in-plane edge restraints and small initial geometric imperfections are also considered in the model. The objective is twofold: 1) to analyze the implications of nonlinear unsteady aerodynamics and structural nonlinearities on the character of the flutter instability boundary and 2) to outline the effects played, in the same respect, by a number of important geometrical, physical, and aerodynamic parameters characterizing the aeroelastic system. As a by-product of this analysis, the implications of these parameters on the linearized flutter instability behavior of the system are captured and emphasized. The behavior of the aeroelastic system in the vicinity of the flutter boundary is studied via the use of an encompassing methodology based on the Lyapunov first quantity. Numerical illustrations, supplying pertinent information on the implications of geometric and aerodynamic nonlinearities, as well as of other effects, such as curvature and thickness ratios, on the flutter instability and on the character of the flutter boundary are examined, and pertinent conclusions are outlined.

This paper deals with the generation and use of proper aerodynamic indicial functions toward the aeroelastic formulation of two-dimensional lifting surfaces in the subsonic compressible, linearized transonic, supersonic and hypersonic flight speed regimes. The indicial function approach enables one to treat in an unified way (i.e. in the time and frequency domains) the subcritical aeroelastic response and the flutter instability of lifting surfaces. Validations of the aerodynamic model are documented and excellent agreements are reported. In addition, closed form solutions and aerodynamic derivatives for different flight speed regimes are obtained; comparisons, and results displaying the aeroelastic response to blast loads are presented, and pertinent conclusions are outlined.

This paper addresses the problem of the determination of the subcritical aeroelastic response and flutter instability of nonlinear two-dimensional lifting surfaces in an incompressible flow-field via indicial functions and Volterra series approach. The related aeroelastic governing equations are based upon the inclusion of structural and damping nonlinearities in plunging and pitching, of the linear unsteady aerodynamics and consideration of an arbitrary time-dependent external pressure pulse. Unsteady aeroelastic nonlinear kernels are determined, and based on these, frequency and time histories of the subcritical aeroelastic response are obtained, and in this context the influence of the considered nonlinearities is emphasized. Conclusions and results displaying the implications of the considered effects are supplied.

A unified approach of stability and aeroelastic response of swept aircraft wings in an incompressible flow is developed. To this end, the indicial function concept in time and frequency domains is used, and on this basis the flutter instability and subcritical aeroelastic response to arbitrary time-dependent external excitation are analyzed. In addition, an original representation of the aeroelastic response in the phase space is presented, and the implications of the related results toward determining the flutter instability in flight are emphasized. Validations of selected results against the theoretical and experimental predictions are supplied, and pertinent conclusions are outlined.

Atmospheric nucleation is now recognized to be an important source of ambient particles. In this study, ground–based measurements using a tower were used to observe new particle formation in the Morgan Monroe State Forest (MMSF) in Southwestern Indiana in May 2008. Nucleation was observed at MMSF on a number of days through examination of the particle size distributions. Most of these events were nucleation and growth events that are typical of regional nucleation phenomena. The particle size and sulfuric acid concentration data were used to investigate the mechanism for the observed nucleation events. Four of the ten observed nucleation events were clearly the result of activation of pre–existing clusters. The others seem likely to be the result of classical ternary nucleation.