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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...

This article provides new insights into the aerothermoelastic stability of thin plates. Particularly, the issue of loss of stability of an isotropic plate-strip of constant thickness immersed in a supersonic flow field and subjected to a variable temperature field through the thickness is examined. Using the basic principles of the theory of aerothermoelasticity of isotropic bodies, the theories of flexible panels, and the linear law of temperature field through the thickness of the panel, the stability equations and associated boundary conditions are obtained. As expected, the coefficients of the aerothermoelastic governing equations depend on the thermal load, and consequently the panel-flutter critical speed depends on temperature. The model takes into account quadratic and cubic aerodynamic non-linearities as well as cubic geometric non-linearities. Due to the inhomogeneity of the temperature field distribution across the thickness plate buckling instability occurs. This instability accounts for the d...

Hybrid systems consisting of a single or multiple primary sources along with a storage component are used in renewable energy production due to the wide disparity between the intermittent power generated and the power demand. Finding the optimal size of a hybrid system with no loss of power supply (LPS) is of utmost concern when considering the levelized cost of energy (LCE) of the project over its lifecycle. In this study, an optimization routine employing a search algorithm is developed to find the minimum LCE that meets zero LPS in the context of data center application. To this end, a system model is developed by integrating basic models of the subsystems. The system model is then used to investigate two different loading cases, 1) a case where the demand cannot be controlled as in the case of power demand in a residential network, and 2) a case where the demand can be controlled up to certain limits, as in the case of power demand in a data center or a data center network. Various types of controllable power demand scenarios are studied. The results demonstrate a significant reduction in the life cycle costs of the system when controllable power demand is considered, especially in regions where wind and solar resources are relatively low.

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.

This paper describes the features of AMCOMP, a CAD environment conceived to manage the computation of the Added Masses of bodies moving in an infinite fluid. A methodology already published in literature has been implemented in a CAD environment and its features are exploited to perform the evaluation of the Added Masses. These include the importing of the model in STL file format and the computation of the complete Added Mass matrix and non-dimensional terms. The CAD has been validated by computing the Added Masses values of bodies whose exact formulation was found in literature. The validated tool has been applied to more complex case studies where the Added Masses values were guessed using empirical laws or the equivalent ellipsoid approximation. Significant differences have been found between approximated methods and computations on real shapes. It is shown that the precision in the Added Masses computation and the computational time depend on the meshing quality of the model and power of the PC on which the software runs. Several tools have been merged to obtain a model useful for the Added Masses computation and to assess the errors arising from using approximated formulas instead of the real shape of the body.

This article evaluates the amount of energy that can be extracted from a gust using an aeroelastic energy harvester composed of a flexible wing with attached piezoelectric elements. The harvester operates in a subcritical flow region. It is modeled as a linear Euler–Bernoulli beam sandwiched between two piezoceramics. The extended Hamilton’s principle is used to derive the harvester’s equations of motion and an eigenfunction expansion is used to form a three-degree-of-freedom reduced-order model. The degrees of freedom retained in the model are two flexural degrees for the in-plane and out-of-plane displacements, and a torsional degree for the rotational displacement. Wagner and Küssner functions are used to represent the unsteady aerodynamic and gust loading, respectively. The amount of energy extracted from the system is then compared for two different deterministic gust profiles, 1-COSINE and two sharp-edged gusts forming a square gust, for various magnitudes and durations. The results show that the harvester is able to extract more energy from the square gust profile, although for both profiles the harvester extracts more power after the gust has subsided.

This paper presents a study to design, analyze and optimize an airfoil trailing edge, i.e., shape morphing of the airfoil trailing-edge topology. The primary idea behind morphing is to improve the wing performance for different flight conditions. Modern aircrafts are designed for unique operating conditions. In order to obtain the best configuration, a dynamic optimization algorithm has been developed based on a Multi-swarm Particle Swarm Optimization algorithm (MPSO), a population-based stochastic optimization algorithm inspired by the social interaction among insects or animals. However, with respect to aircraft design and in the context of computational fluid dynamics (CFD), function evaluations are computationally expensive; typically requiring large computational grids to obtain a reasonable representation of the flow-field. In this paper, the developed MPSO algorithm is combined with a Kriging surrogate representation of the objective space, to alleviate the computational effort. The topology of the trailing edge is defined and characterized by four control points. Two different hypothetical mission profiles are analyzed. The results exhibit an improvement of around 2% with respect to the original airfoil for every flight condition treated.

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 instant asteroid starship is a reference plan for a manned interstellar mission, developed in the wake of the 100YSS program sponsored by DARPA and NASA. This plan consists of a full conceptual design that considers most of the requirements for a manned interstellar mission, including systems and subsystems, strategies, and motivation. For resource utilization, safety considerations, and immediate start of activities, a captured asteroid is engineered as a space settlement. In this article, basic aspects of this plan are summarized; the development is part of a larger effort, sponsored by the group Star Voyager, which should culminate in a reference starship design.

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.

This paper discusses the theory of thermo-conductivity due steady-state heat flow of semi-infinite homogeneous solids containing a perfectly rigid circular inclusion located on the plane where the two half-solids are joint. The half-spaces have different thermo-physical characteristics and a steady heat flow is applied to the solids. The effects of selected spatial heat sources are also investigated. Mathematical model and general solution is presented along with selected particular cases and simulations, followed by pertinent discussions and conclusions.

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.

The fouling effect of seawater samples collected from three New Hampshire beaches on a 90/10 Cu/Ni commercial heat exchanger tube was investigated. A fouling monitoring device, designed on the basis of fouling thermal resistance, was used for this experimental study. The filtered seawater samples were circulated through the closed loop experimental setup for durations of seven and 14 days and the fouling thermal resistance was measured continuously. Analytical microscopy was performed on the tube surface before and after the experiments to see the effect of seawater fouling on the tube surface. The results show different fouling behavior for the seawater samples. This different behavior is confirmed by the different composition of the samples. Fouling monitoring experiments reveal a higher fouling thermal resistance for one of the seawater samples, Hampton seawater, contrary to the results of SEM analysis which show a lower crystallization rate for Hampton sample. Water decomposition analysis shows the lowest sodium content for Hampton seawater compared to the other samples. Accordingly, corrosion of the tube surface occurs with a higher rate for Hampton seawater due to the presence of chlorine ions and a lower concentration of sodium. The high fouling resistance of Hampton seawater can be explained as the result of several simultaneous fouling mechanisms, corrosion and crystallization indicating composite fouling behavior. The results of the current study are critical for the industries which use seawater as the cooling water source.

In this study, aeroelastic analysis of a plate subjected to the external supersonic airflow is carried out. A 3-D rectangular plate element of variable thickness based on absolute nodal coordinate formulation (ANCF) has been developed for the structural model. In the approach to the problem, 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. Linearized first-order potential (piston) theory is coupled with the structural model to account for pressure loading. 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. All the formulations and the computations are built up in a FORTRAN 90 computer program after it was confirmed by Mathematica®, ver. 5. The results of free vibration analysis and flutter are compared with the available references and reasonable good agreement has been found. However, some results indicate that the known problem of locking (ANCF with uniform thickness) still persist in the current developed formulation.

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.

Functionally Graded Materials (FGM) have attracted significant interest as heat-shielding materials for space vehicle, skin sub-structures, gas turbine blades technologies, and many other high-temperature industrial applications. This paper reviews the state-of-the-art in linear and nonlinear aero-thermo-elasticity of FGM panels with emphasis on the authors’ contributions to the topic. An overview of the pertinent literature discussing the linear and nonlinear behavior of flat and curved panels when exposed to high temperature supersonic flow fields is presented first. The effect of material property dependency on temperature is also discussed. The study addresses divergence and flutter and methodologies used to determine these aero-thermo-elastic instabilities. In particular, critical and post-critical behaviors for panels in presence of thermal loads are addressed, along with a series of divergence, flutter, and post-flutter results obtained with linear/nonlinear dynamics approaches. Regular and chaotic...

In this paper, a class of aeroelastic systems with an unmodeled nonlinearity and external disturbance is considered. By using leading- and trailing-edge control surface actuations, a full-state feedforward/feedback controller is designed to suppress the aeroelastic vibrations of a nonlinear wing section subject to external disturbance. The fullstate feedback control yields a uniformly ultimately bounded result for two-axis vibration suppression. With the restriction that only pitching and plunging displacements are measurable while their rates are not, a high-gain observer is used to modify the full-state feedback control design to an output feedback design. Simulation results demonstrate the efficacy of the multi-input multi-output control toward suppressing aeroelastic vibration and limit cycle oscillations occurring in pre and postflutter velocity regimes when the system is subjected to a variety of external disturbance signals. Comparisons are drawn with a previously designed adaptive multi-input multi-output controller.

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.

The aeroelastic modeling and flutter characteristics of a shear deformable wing/stores configuration under pull-up angular velocity is investigated. An isotropic non-uniform wing, which structural model incorporates flexibility in transverse shear and warping effects, is considered. The aeroelastic governing equations and boundary conditions are determined via Hamilton’s variational principle. In order to exactly consider the span wise location and properties of the attached stores the generalised function theory is used. The partial differential equations are transformed into a set of eigenvalue equations through the extended Galerkin’s approach. Numerical simulation highlighting the effects of the pull-up angular velocity and store parameters and configurations, such as mass ratio and their attachment locations, on the flutter speed are presented. The results of flutter analyses are validated with the published results and good agreement is observed. Furthermore, the procedure for an optimal deployment of stores is obtained for the case of the wing with four stores.

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.

This paper deals with the thermally induced dynamic response analysis of a composite blade rotating at a constant speed. Fiber-reinforced composite thin-walled beam are used in a variety of aerospace applications, including helicopter rotor blades, tilt rotor aircraft, turbine engines, compressor blades, only to name a few. The composite beam is modeled as a tapered thin-walled beam which is subjected to a temperature field. A number of non-classical features such as transverse shear, secondary warping, anisotropy of constituent materials, and rotary inertias have been included in the model. The structure of the blade consists of a host graphite epoxy laminate with spanwise distributed transversely isotropic (PZT-4) sensors and actuators. The discussed results reveal that the thermal environment has a detrimental effect on the blade dynamic response. The active controller is implemented via the combined displacement and velocity feedback control methodology, which can alleviate the deleterious effect asso...

Damage detection and structural health monitoring techniques based on vibration data have seen increased attention in recent years. Among the different vibration-based methods, the ones based on random vibrations are of particularly interest, especially when they do not require measurement of the input(s). In this work, several frequency and time domain signal processing techniques are explored in their respective abilities to detect damage in a bolted composite structure. First, a joint loosening model is developed and used to simulate the dynamic response to a stationary Gaussian excitation. Informed by the model, two signal processing techniques are used to assess the connection strength. The first method relies on basic statistical properties of the measured strains and their time derivatives, while the second is based on the signal power in different frequency bands. Both approaches are then used to assess progressive bolt loosening on an experimental composite-to-metal joint. All strain response data were obtained using a fiber optic strain sensing system. Results are presented in the form of Receiver Operating Characteristic (ROC) curves, showing both Type-I and Type-II errors associated with the proposed detection schemes.

Higher-order spectral analysis techniques are often used to identify nonlinearities in complex dynamical systems. More specifically, the auto- and cross-bispectrum have proven to be useful tools in testing for the presence of quadratic nonlinearities based on knowledge of a system's input and output. In this paper, analytical expressions for the auto- and cross-bispectrum are developed using a Volterra functional approach under the assumption of a zero-mean, stationary Gaussian input; proper simplifications are presented when the whiteness of the input signal is also imposed. These formulae show the contributions of the bispectrum in terms of the system frequency response function and elementary physical properties of the system. Simulations based on a stochastic numerical integration technique accompany the analytical solutions for a mechanical mass–spring–damper system possessing quadratic damping and stiffness coefficients and subjected to Gaussian white noise excitation. Subsequent estimates of the bispectrum based on the simulated signals show excellent agreement with theory. These results show how modes may interact nonlinearly producing intermodulation components at the sum and/or difference frequency of the fundamental modes of oscillation. The presence and extent of nonlinear interactions between frequency components are identified. Advantages of using higher-order spectra techniques will be revealed and pertinent conclusions will be outlined.

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.

Higher-order spectra appear often in the analysis and identification of nonlinear systems. The auto-bispectral density is one example of a higher-order spectrum and may be used in the analysis of stationary structural response data to detect the presence of certain types of structural nonlinearities. In this work a closed-form expression for the auto-bispectral density, derived previously by the authors, is used to find the bispectral frequency most sensitive to the nonlinearity. The properties of nonlinearity detectors based on estimates of the magnitude of the auto-bispectral density at this frequency are then explored. Estimates of the auto-bispectral density are obtained using the direct method based on the discrete Fourier transform. The bias associated with this estimator is derived here and combined with previously derived expressions for the estimator variance to give both Type-I and Type-II errors for the detector. Detector performance is quantified using a receiver operating characteristic (ROC) curve illustrating the trade-off between false positives (Type-I error) and power of detection (1.0-Type-II error). Theoretically derived ROC curves are compared to those obtained via numerical simulation and show excellent agreement. Results are presented for different levels of nonlinearity in both the stiffness and damping terms for a spring–mass system. Possible consequences are discussed with regard to the detection of damage-induced nonlinearities in structures.

In the current work, the study of the aeroelastic behaviour of a wing with external store(s), such as a missile or drop fuel tank, is presented. The aeroelastic governing equations derived for a cantilever wing with coupled bending and torsion modes account for structural and aerodynamic non-linearities. Coupling terms retained in the aeroelastic governing equations are due to: (a) the non-linear beam theory, (b) the aerodynamic non-linearities of a quasi-steady model with stall, and (c) the non-linear kinematics terms of the store(s). As presented in the paper, this aircraft configuration can induce pathologies, such as store(s)-induced limit cycle oscillations (or si-LCOs), very different from the one of a clean wing configuration, and from the one obtained from the linearized form of the aeroelastic governing equations. Time domain simulation, phase portrait, and bifurcation analyses are performed for various velocities, initial conditions, and store(s)-sensitive parameters — such as store mass, number, location along the wing — to examine the dynamic aeroelastic instabilities of the system (e.g. the onset of flutter and LCO). Numerical studies indicate the presence of regions of subcritical Hopf-bifurcation, corresponding to an unstable LCO, for velocities below the linear flutter velocity.

Via the use of leading- and trailing-edge control surface actuation, an adaptive output feedback controller is designed for suppressing aeroelastic vibrations on a nonlinear wing section. Although a single flap under adaptive control can suppress vibrations, the response rate is limited by the system zero dynamics. Under the restriction that only pitching and plunging variables are available for measurement but their rates are not, the proposed algorithm addresses the problem of designing a singularity-free adaptive output feedback controller when the control inputs are coupled via an input gain matrix for which the parameters are uncertain. The stability result achieved is global asymptotic tracking. Simulation results demonstrate the efficacy of the multi-input/multi-output control toward suppressing flutter and limit-cycle oscillations, as well as reducing the vibrational level in the subcritical flight-speed range. Pertinent conclusions are outlined.

The flutter and post flutter of a two-dimensional double-wedge lifting surface with combined freeplay and cubic stiffness nonlinearities in both plunging and pitching degrees-of-freedom operating in supersonic/hypersonic flight speed regimes have been analyzed. In addition to the structural nonlinearities, the third-order piston theory aerodynamics is used to evaluate the unsteady non-linear aerodynamic force and moment. Such model accounts for stiffness and damping contributions produced by the aerodynamic loads. Responses involving limit cycle oscillation and chaotic motion are observed over a large number of parameters that characterizes the aeroelastic system. Results of the present study show that the freeplay in the pitching degree-of-freedom and soft/hard cubic stiffness in the pitching and plunging degrees-of-freedom have significant effects on the LCOs exhibited by the aeroelastic system in the supersonic/hypersonic flight speed regimes. The simulations also show that the aeroelastic system behavior is greatly affected by physical structural parameters, such as the radius of gyration and the frequency ratio, especially in post-flutter regimes, when accounting for all system nonlinearities. It has been shown that at high Mach numbers the non-linear aerodynamic stiffness yields detrimental effects from the aeroelastic point of view, while the damping one do not.

This paper addresses stability analysis, control design, and simulation of high-altitude, long-endurance unmanned aerial vehicles. These aircraft are highly flexible and have very low structural frequencies. Hence, the separation between the elastic and rigid body motions no longer exists, and an approach that unifies these two motions must be used. The available data of the aircraft considered include the geometry of the aircraft, aerodynamic data, and mass, flexural rigidity, and torsional rigidity distributions. The wing of the aircraft is modeled as a flexible beam undergoing bending and torsion, whereas the remaining members are assumed to be rigid. The equations of motion are obtained by means of the Lagrangian equations in quasi coordinates. A perturbation approach separates the problem into nominal dynamics and perturbation dynamics. The equations for nominal dynamics are used to design desired maneuvers and to determine the corresponding structural deformations. The equations for perturbation dynamics are used to address stability of the aircraft on the desired flight paths, to design feedback controls to maintain stable flights, and to simulate the motion of the aircraft.

By using a combined analytical-computational methodology, a unified modelling of aerodynamic indicial functions covering the incompressible, subsonic compressible, transonic, and supersonic flight speed regimes is presented. The procedure is carried out in conjunction with a computational fluid dynamic analysis. For a plunging-pitching airfoil, selected unsteady aerodynamic load expressions have been supplied, and appropriate procedures enabling one to obtain these loads via the indicial function approach have been presented. While a single indicial function is needed to describe the aerodynamic loads in the incompressible flight speed regime, for cases where the compressibility effects play a dominant role, four indicial functions are needed. Having in view the usefulness of indicial functions towards determination of unsteady aerodynamics loads in both time and frequency domains, and implicitly for aeroelastic response and flutter predictions, the advantages of their implementation and use appear evident. Comparisons and validations of the aerodynamic model against numerical, analytical, and experimental results are presented, and pertinent conclusions are drawn.

The performance of a robust control strategy applied to a two-dimensional supersonic flap-wing aeroelastic system impacted by a pressure pulse in the subcritical supersonic flight speed regime is analyzed. The piston theory aerodynamic model modified to account for the flap deflection has been considered. The study is carried-out via implementation of the Linear Quadratic Gaussian (LQG) control methodology with sliding mode observer. Its performance toward suppressing flutter, reducing the vibrational level amplitudes, as well as toward eliminating the observation spillover due to the unmeasured states is revealed through numerical simulations, and pertinent conclusions are outlined.

Parks containing manufactured homes and trailers exist in the hurricane-prone areas of the southern United States. Homes in these parks may suffer heavy damage from high-speed winds. This paper’s objective is to present a methodology and to provide suggestions to the engineering community and protection agencies such as FEMA and the U.S. Army Corps of Engineers (USACE) about a method to lessen damage to groups of manufactured homes. To support the development of this methodology with analysis, an unsteady, Reynolds-averaged, Navier-Stokes computational tool was employed to model a manufactured-home park to evaluate how well a wind barrier installed along its upwind periphery could establish skimming flow over all the buildings in the park. A manufactured-home park that had existed in Punta Gorda, Florida, was studied because satellite pictures before and after it was heavily damaged by Hurricane Charley on August 13, 2004, were available. Computations were performed using a computer program and th...

The limit cycle oscillations (LCOs) exhibited by long-span suspension bridges in post-flutter condition are investigated. A parametric dynamic model of prestressed long-span suspension bridges is coupled with a nonlinear quasi-steady aerodynamic formulation to obtain the governing aeroelastic partial differential equations adopted herewith. By employing the Faedo–Galerkin method, the aeroelastic nonlinear equations are reduced to their state-space ordinary differential form. Convergence analysis for the reduction process is first carried out and time-domain simulations are performed to investigate LCOs while continuation tools are employed to path follow the post-critical LCOs. A supercritical Hopf bifurcation behavior, confirmed by a stable LCO, is found past the critical flutter condition. The analysis shows that the LCO amplitude increases with the wind speed up to a secondary critical speed where it terminates with a fold bifurcation. The stability of the LCOs within the range bracketed by the Hopf and fold bifurcations is evaluated by performing parametric analyses regarding the main design parameters that can be affected by uncertainties, primarily the structural damping and the initial wind angle of attack.

In this paper, we formulate the stacking sequence design of laminated composite flat panels for maximum supersonic flutter speed and maximum thermal buckling capacity. The design is constrained to have stable limit cycle oscillation in the vicinity of the flutter boundary. For the purpose of comparison, we consider both linear and nonlinear panel flutter analyses. Because of manufacturing constraints, the stacking sequence of the laminate is selected from a discrete set of orientation angles. The plate is modeled using von Karman plate equations and assumed to be subjected to uniform temperature differential. The aerodynamic load on the plate is computed using third-order piston theory. The plate equations of motion are derived using Hamilton's principle and discretized using the Rayleigh-Ritz method. The nonlinear panel flutter analysis is performed using an approximate perturbation approach. The Pareto fronts describing the trade-off between thermal buckling and flutter are generated for different panel aspect ratios. It is observed that optimal designs obtained for both linear and nonlinear flutter analyses are similar, with only the predicted flutter speeds different This suggests that, for the purpose of preliminary stacking sequence design, linear flutter analysis is sufficient.

In this paper we present, the design and modeling of the novel nonlinear limiter control feedback control plant [Myneni et al., 1999; Corron et al., 2000; Corron & Pethel, 2002], applied for the first time here in an aeroelastic system, and actuated as a jet reaction torquer control of a wing with potentially chaotic dynamics. This study will provide a better understanding of the nonlinear dynamics of the open/closed-loop aeroelasticity of flexible wings with either steady or unsteady aerodynamic loads. The limiter control can be applied to either the plunging or pitching characteristic of the wing or to both of them. We show that the control can effectively suppress Limit Cycle Oscillations (LCO) and chaos well beyond the nominal flutter speed. This could lead to a practical implementation of the control mechanism on actual and future generation aircraft wings via implementation of a combination of propulsive/jet type forces, micro surface effectors and fluidic devices. Analysis of this control produced favorable results in the suppression of LCO amplitude and increased flutter boundaries for plunging and pitching motion. The limiting control has asymptotically zero power, and is simply implemented, making it a feasible solution to the problem of the chaotic dynamics of the oscillating airfoil.

This paper concerns the flutter, post-flutter and adaptive control of a non-linear 2-D wing-flap system operating in supersonic/hypersonic flight speed regimes. An output feedback control law is implemented and its performance toward suppressing flutter and limit cycle oscillations (LCOs) as well as reducing the vibrational level in the subcritical flight speed range is demonstrated. This control law is applicable to minimum phase systems and we provide conditions for stability of the zero dynamics. The control objective is to design a control strategy to stabilize the pitch angle while adaptively compensating for uncertainties in all the aeroelastic model parameters. It is shown that all the states of the closed-loop system are asymptotically stable.

Adaptive control of a nonlinear two-dimensional wing-flap system operating in an incompressible flowfield is studied. An output feedback control law is implemented, and its performance toward suppressing flutter and limit cycle oscillations, as well as reducing the vibrational level in the subcritical flight speed range, is demonstrated. The control law proposed here is applicable to minimum phase systems, and conditions for stability of the zero dynamics are provided. The control objective is to design a control strategy to drive the pitch angle to a setpoint while adaptively compensating for uncertainties in all of the aeroelastic model parameters. It is shown that all of the states of the closed-loop system are asymptotically stable. Furthermore, an extension is presented to include flap actuator dynamics. Simulations have been presented to validate the efficacy of the proposed strategy. Pertinent conclusions have been outlined.

The performance of a robust control strategy applied to a three-degrees-of-freedom (3-DOF) flap-wing aeroelastic system impacted by a pressure pulse in the subcritical flight speed regime is investigated. The goal of its implementation is to suppress flutter instability and reduce the vibrational level in the subcritical flight speed range. To this end, the linear quadratic Gaussian (LQG) control methodology in conjunction with a sliding mode observer (SMO) are used. Comparisons with the counterpart results obtained via implementation of LQG controller with conventional Kalman filter (KF) are also provided and pertinent conclusions are outlined.

Active aeroelastic control is a recently emerging technology aimed at providing solutions to a large class of problems involving the aeronautical/aerospace flight vehicle structures that are prone to instability and catastrophic failures, and to oscillations that can yield structural failure by fatigue. In order to prevent such damaging phenomena to occur, the linear/nonlinear aeroelastic control technology should be applied. Its goals are among others: (i) to alleviate and even suppress the vibrations appearing in the subcritical flight speed range, (ii) to enlarge the flight envelope by increasing the flutter speed, and (iii) to enhance the post-flutter behavior by converting the unstable limit cycle oscillation to a stable one. A short review of the available control techniques and capabilities is presented first. Attention is focused on the open/closed-loop of 2D and 3D lifting surfaces as well as on panels exposed to supersonic flowfields. A number of concepts involving various control methodologies, such as proportional, velocity, linear quadratic regulator, modified bang-bang, sliding mode observer, time-delay control, fuzzy, etc., as well as results obtained with such controls are presented. Emphasis is placed on theoretical and numerical results obtained with the various control strategies that are considered in a comparative way. Finally, conclusions and directions for further work are presented.

Two basic issues related to the open/closed-loop aeroelasticity of 2-D lifting surfaces in an incompressible flow field are considered. These concern the subcritical aeroelastic response to external time-dependent excitations, and the flutter instability of actively controlled airfoils involving a time-delayed feedback control. Results and comparisons regarding the flutter instability obtained via the first Volterra kernel in conjunction with a frequency eigenvalue analysis are presented. In the same context, the implications on the instability boundary and aeroelastic response of the presence of time-delays in the feedback control are investigated and pertinent conclusions are supplied.

Wind tunnel experiments are performed to investigate the aerodynamic characteristics of unconventional airships docking in-flight. The aim is to provide wind tunnel data for multiple feeders docking to a cruiser and to identify key issues in docking approach strategies. The conceptual design of a docking mechanism and its integration into a baseline configuration is also addressed. Force and moment data are collected on a 1:7.5x10 -4 scale model for the cruiser configurations and at various cruiser-feeder vertical separation distances to simulate docking. Results indicate that the there is a trade-off between drag performance and the implementation of the docking mechanism. Drag decrease is measured when the modified cruiser is docked with the feeders. Best docking practices are also identified from the pitching moment coefficients showing that the cruiser-feeder can be trimmed during docking.