Aeroelastic Analysis Research Articles

Stochastic aeroelastic analysis of laminated composite plate with variable fiber spacing

In this paper, the dynamic and aeroelastic analysis of variable fiber spacing composite (VFSC) laminated plates are carried out. The effects and benefits of changing the fiber distribution pattern on natural frequency and mode shape are explored taking various boundary conditions. Taking cantilever boundary condition flutter characteristics of VFSC plate with different fiber distribution patterns are compared. Further stochasticity of flutter velocity due to randomness in material properties that could arise due to complex manufacturing and fabrication process of VFSC laminates is investigated. The perturbation technique is implemented to perform the stochastic as well as reliability analysis. Various parametric studies are conducted to check the accuracy and efficiency of perturbation technique, by comparing the results with that of Monte Carlo simulation and the first-order reliability method.

Sloshing reduced-order models for aeroelastic analyses of innovative aircraft configurations

The present paper introduces the sloshing effects on aeroelastic stability and response of flying wing configurations. Using reduced order model based on data provided by computational fluid dynamics data, the description of the slosh dynamics is introduced into an integrated modelling that accounts for both rigid and elastic behaviour of flexible aircraft. The fully unsteady aerodynamics, modelled in the frequency domain via doublet lattice method, is recast in time-domain state-space form through the rational function approximation. The reference aircraft is the Body Freedom Flutter equipped with a single tank of cylindrical geometry, partially filled with water, and located underneath the center of mass. The results of the stability and response analyses are compared with the frozen fluid case with the final aim of showing how the fluid movement affects the complete aircraft behaviour.

Coupled Aeroelastic Study of a Flexible Micro Air Vehicle-Scale Flapping Wing in Hovering Flight

Instantaneous force and structural deformation experiments were performed on a flexible, structurally characterized, low-aspect-ratio representative flapping wing. A six-component force balance was used to measure aerodynamic-force time history over a flap cycle. A VICON motion capture setup tracked the wing deformations while flapping. Measured aerodynamic forces and wing deformations were compared against coupled computational fluid dynamics/computational structural dynamics (CFD/CSD) aeroelastic analysis results. The CFD solver is an unsteady Reynolds-averaged Navier–Stokes solver. The CSD solver is a general-purpose multibody dynamics solver capable of modeling geometrically nonlinear beam and shell elements. The CFD/CSD results were able to capture the overall trend in aerodynamic forces and wing deformations. Although predicted and measured variations in lift were similar, the drag-force magnitudes tended to be underpredicted by the coupled aeroelastic solver. The coupled CFD/CSD solver was used to evaluate the influence of wing flexibility on wing performance. Results showed that, in particular instances, decreasing wing stiffness increased the time-averaged aerodynamic lift and lift-to-power ratio. Decreased wing stiffness led to reduced leading-edge vortex circulation strength. This suggests that the increased wing performance is mainly due to a redirection of the resultant force vector allowed by the increased wing compliance.

Structural damping models for passive aeroelastic control

Aeroelastic qualification requirements are typically met by sizing aircraft to achieve adequate stability margins and keep peak gust responses below specified thresholds. A possible alternative approach is delaying flutter and alleviating gust response by embedding dissipative materials into structural components. This approach requires accurate damping models applicable to the analysis of complex configurations. In this paper, the effect of damping models in the evaluation of flutter boundaries and gust response of an aeroelastic test-bed is studied. In particular, three damping models completely different in frequency are considered to model the presence of skin patches for passive aeroelastic control: viscous damping, hysteretic damping and Biot damping models. In order to make them comparable the three models are tuned in order to dissipate the same amount of power at the flutter frequency by means of the introduction and definition of a generalized loss factor. The damping models are compared by evaluating their effect on flutter suppression and gust load alleviation. Finally, the sensitiveness of the stability and response aeroelastic analyses to the considered damping models are outlined and the obtained results are compared to provide modeling recommendations for passive flutter suppression and gust alleviation studies. The proposed methodology is applicable to any linear material model for which the related complex stiffness can be expressed in the frequency domain meeting the Hilbert restriction.

A Parametric Study on the Aeroelasticity of Flared Hinge Folding Wingtips

This paper presents a parametric study on the aeroelasticity of cantilever wings equipped with Flared Hinge Folding Wingtips (FHFWTs). The finite element method is utilized to develop a computational, low-fidelity aeroelastic model. The wing structure is modelled using Euler–Bernoulli beam elements, and unsteady Theodorsen’s aerodynamic strip Theory is used for aerodynamic load predictions. The PK method is used to estimate the aeroelastic boundaries. The model is validated using three rectangular, cantilever wings whose properties are available in literature. Then, a rectangular, cantilever wing is used to study the effect of folding wingtips on the aeroelastic response and stability boundaries. Two scenarios are considered for the aeroelastic analysis. In the first scenario, the baseline, rectangular wing is split into inboard and outboard segments connected by a flared hinge that allows the outboard segment to fold. In the second scenario, a folding wingtip is added to the baseline wing. For both scenarios, the influence of fold angle, hinge-line angle (flare angle), hinge stiffness, tip mass and geometry are assessed. In addition, the load alleviation capability of FHFWT is evaluated when the wing encounters discrete (1-cosine) gusts. Finally, the hinge is assumed to exhibit cubic nonlinear behavior in torsion, and the effect of nonlinearity on the aeroelastic response is assessed and analyzed for three different cases.

Aeromechanics and Aeroelastic Stability of Coaxial Rotors

Coaxial rotor aeroelasticity is complex due to the counter-rotating wake system, rotor lift offset, periodic blade passage loads, unsteady rotor wake interactions, reduced rotor speed, and stiff hingeless blades. In this study, the aeroelastic stability of a coaxial rotor is examined in hover and forward flight. The rotor wake is modeled with the viscous vortex particle method, a grid-free approach for calculating vortex interactions over long distances. The spanwise blade aerodynamic loading is calculated using a computational-fluid-dynamics-based reduced-order model in attached flow, and the ONERA dynamic stall model in separated flow. Two propulsive trim procedures are developed: one with the propulsor not operating, and the other with the vehicle at level attitude. An aeroelastic stability analysis based on Floquet theory is applied to the periodic system. A novel graphical method is developed to identify coupling between blade modes of the two rotors. The effects of lift offset and advance ratio on the hub loads, inflow distribution, and aeroelastic stability are examined to provide an improved physical understanding of the aeroelastic interactions. Results indicate that the blade passage effect is caused by the bound-circulation-induced inflow. The first and second lag modes are the least stable modes in hover and forward flight.

Reduced-Order Nonlinear Damping Model: Formulation and Application to Postflutter Aeroelastic Behavior

Recent studies of postflutter limit-cycle oscillations (LCOs) have suggested the presence/effects of a nonlinear structural damping in addition to other potential sources of nonlinearity. Such a nonlinearity occurs, for example, for structures with linear viscoelastic properties when their responses are in the nonlinear geometric regime. The present effort focuses on such situations; first on developing a structural reduced-order model (ROM) which could be used in aeroelastic analyses. Adopting a linear Kelvin–Voigt constitutive model in the undeformed configuration, the ROM governing equations are obtained and found to be of a generalized Van der Pol–Duffing form. A nonintrusive identification approach is next developed to determine the parameters of these governing equations from a structural finite element model constructed in a commercial software. Finally, the effects of this nonlinear damping on postflutter response are analyzed on the Goland wing assuming a linear aerodynamic model. It is found that the nonlinearity in the damping can stabilize the unstable aerodynamics and lead to finite amplitude limit-cycle oscillations, even when the stiffness-related nonlinear geometric effects and aerodynamic nonlinearities are neglected. The dependence of the LCO amplitude and frequency on the parameters of the Kelvin–Voigt model is analyzed to provide insight into this nonlinearity.

Numerical study on the trajectory of a long-range flexible rocket with large slenderness ratio

With the requirements of the speed, maneuverability and agility, the structure with large slenderness ratio is widely employed in the design of the modern rocket. The interaction among unsteady aerodynamic load, structural elastic response, as well as flight motion will be more significant. This paper focuses on the aeroelastic effects in the trajectory simulation of a rotary rocket with a large slenderness ratio. A calculation method based on computational fluid dynamics / computational structural dynamics / computational flight mechanics (CFD/CSD/CFM) coupling system is proposed, in which an in-house CFD/CSD strong-coupled program is connected with the six-degree-of-freedom flight mechanics' equations. The CFD/CSD coupling subsystem as well as CFD/CFM coupling subsystem are validated by aeroelastic dynamic stability analysis of AGARD wing and trajectory simulation of a spinning projectile, respectively. Then, the coupling algorithm is performed on a long-range trajectory simulation of an uncontrolled rotary flexible rocket. The unsteady aerodynamic loads, structural deformation, flight motions, as well as the varying thrust direction are obtained and analyzed. Through the Monte Carlo Experiment, the landing points dispersion of the elastic rocket is also calculated and compared with the rigid case. The results show that the aeroelastic effect can decrease the range and reduce the landing accuracy of the rocket. The proposed method can provide technical support for the correction of the firing table and the design of the control system of the rocket.

The stochastic aeroelastic response analysis of helicopter rotors using deep and shallow machine learning

This paper addresses the influence of manufacturing variability of a helicopter rotor blade on its aeroelastic responses. An aeroelastic analysis using finite elements in spatial and temporal domains is used to compute the helicopter rotor frequencies, vibratory hub loads, power required and stability in forward flight. The novelty of the work lies in the application of advanced data-driven machine learning (ML) techniques, such as convolution neural networks (CNN), multi-layer perceptron (MLP), random forests, support vector machines and adaptive Gaussian process (GP) for capturing the nonlinear responses of these complex spatio-temporal models to develop an efficient physics-informed ML framework for stochastic rotor analysis. Thus, the work is of practical significance as (i) it accounts for manufacturing uncertainties, (ii) accurately quantifies their effects on nonlinear response of rotor blade and (iii) makes the computationally expensive simulations viable by the use of ML. A rigorous performance assessment of the aforementioned approaches is presented by demonstrating validation on the training dataset and prediction on the test dataset. The contribution of the study lies in the following findings: (i) The uncertainty in composite material and geometric properties can lead to significant variations in the rotor aeroelastic responses and thereby highlighting that the consideration of manufacturing variability in analyzing helicopter rotors is crucial for assessing their behaviour in real-life scenarios. (ii) Precisely, the substantial effect of uncertainty has been observed on the six vibratory hub loads and the damping with the highest impact on the yawing hub moment. Therefore, sufficient factor of safety should be considered in the design to alleviate the effects of perturbation in the simulation results. (iii) Although advanced ML techniques are harder to train, the optimal model configuration is capable of approximating the nonlinear response trends accurately. GP and CNN followed by MLP achieved satisfactory performance. Excellent accuracy achieved by the above ML techniques demonstrates their potential for application in the optimization of rotors under uncertainty.

Reliability-Based Design of an Aircraft Wing Using a Fuzzy-Based Metaheuristic

The purpose of this paper is to design aircraft wing using reliability-based design optimization concerned to fuzzy uncertainty variables. A possibilistic safety index-based design optimization (PSIBDO) with fuzzy uncertainties is proposed to overcome difficult tasks from the original probabilistic problem. The design problem is to minimize mass of a composite aircraft wing subject to aeroelastic and structural constraints through consideration of the material properties are the uncertainties. The design variables include aircraft wing structure dimensions. The reliability-based design approach is needed to alleviate such a problem. Due to the complexity of the aircraft wing structures design and aeroelastic analysis, nonprobability-based design is an alternative choice to increase computational efficiency in the design process. The optimum results show the efficiency of our proposed approach.

Geometrically nonlinear aeroelastic characteristics of highly flexible wing fabricated by additive manufacturing

Additive manufacturing technology has a potential to improve manufacturing costs and may help to achieve high-performance aerospace structures. One of the application candidates would be a wind tunnel wing model. A wing tunnel model requires sophisticated designs and precise fabrications for accurate experiments, which frequently increase manufacturing cost. In this paper, manufacturing accuracy and aeroelastic characteristics of an additively manufactured wing model are evaluated numerically and experimentally. The feasibility of such wings to use in wind tunnel tests is also demonstrated. In addition, a geometrically nonlinear aeroelastic analysis model, which have been developed in the previous study, is validated for static calculations by comparing with results of the wind tunnel test for the additively manufactured highly flexible wing model.

Aeroelastic stability analysis of a flexible panel subjected to an oblique shock based on an analytical model

Aeroelastic stability analysis of a flexible panel subjected to an oblique shock based on an analytical model

Efficient nonlinear aeroelastic analysis of a morphing wing via parameterized fictitious mode method

Efficient nonlinear aeroelastic analysis of a morphing wing via parameterized fictitious mode method

Numerical and Experimental Study of a Flexible Trailing Edge Driving by Pneumatic Muscle Actuators

A static aeroelastic analysis of the flexible trailing edge is conducted to calculate the deformed shape, aerodynamic coefficients and corresponding driving pressure. A physical flexible trailing edge model is manufactured using a honeycomb structure, which is measured based on binocular vision. The quadratic response surface method is adopted to establish the pneumatic artificial muscle actuator model. The wire-pulley transmission model is built to identify the existence of equivalent forces and produce the equivalent forces as the substitute of actuation force. A finite element model of the flexible trailing edge is established, which is validated by the test data. A nonlinear relationship is found between the driving pressure and deflection angle. The pressure needed to bear the structural stiffness is found to be much larger than that of the aerodynamic load. With the increase in pressure, the magnitude of the lift coefficient increases less. However, the magnitude of the drag coefficient increases more with the increase in pressure under 0.2 MPa. When the driving pressure exceeds 0.2 MPa, the relationship between them is nearly linear.

Kinematic characteristics of longitudinal double folding wings

ABSTRACTA folding wing is a tactical missile launching device that needs to be miniaturised to facilitate storage, transportation, and launching; save missile and transportation space; and improve the combat capability of weapon systems. This study investigates the aeroelastic characteristics of the secondary longitudinal folding wing during the unfolding process. First, the Lagrange equation is used to establish the structural dynamics model of the folding wing, the kinematics characteristics during the deformation process are analysed, and the unfolding movement of the folding wing is obtained using the dynamic equations in the process. Then, the generalised unsteady aerodynamic force is calculated using the dipole grid method, and the multi-body dynamics equation of the folding wing is obtained. The initial angular velocity required for the deployment of the folding wing is analysed through structural model simulation, and the influence of the initial angular velocity on the opening process is studied. Finally, aeroelastic flutter analysis is performed on the folding wing, and the physical model of the folding wing verified experimentally. Results show that the type of aeroelastic response is sensitive to the initial conditions and the way the folding wing opens.

Time-Domain Aeroelasticity Analysis by a Tightly Coupled Fluid-Structure Interaction Methodology

A tightly coupled fluid-structure interaction (FSI) methodology is developed for aeroelasticity analysis in the time domain. The preconditioned Navier–Stokes equations for all Mach numbers are employed and the structural equations are tightly coupled with the fluid equations by discretizing their time derivative term in the same pseudo time-stepping method. A modified mesh deformation method based on reduced control points radial basis functions (RBF) is utilized, and a RBF based mapping algorithm is introduced for data exchange on the interaction interface. To evaluate the methodology, the flutter boundary and the limit cycle oscillation of Isogai wing and the flutter boundary of AGARD 445.6 wing are analyzed and validated.

Delamination effects on the unsteady aero-elastic behavior of composite wing by modal analysis

This article deals with the investigation of delamination effects on the flutter behavior of a composite wing subjected to unsteady aerodynamic loads by providing an integrated procedure using exact methods. For this purpose, an undamaged wing structure is modeled as a cantilevered Bernoulli–Euler beam with coupled bending–torsion and delaminated wing structure is modeled as three interconnected intact beams. The modal analysis of the intact and delaminated beams is performed and mode shapes are obtained based on the development of the dynamic stiffness matrix, a computer code and an iterative method. The fundamental modes of the intact and delaminated beam derived from the free vibration analysis are used in Galerkin’s method and the flutter and divergence speeds are obtained with aerodynamic forces applied. In fact, modified shape functions were developed to reflect the delamination effect in the aeroelastic analysis. Also, for the unsteady aeroelastic analysis, an iterative algorithm is developed. It was found that the delamination location and its length can have a positive or a negative influence on the flutter behavior of the wing which depends on the fiber angle.

Effect of Fuel Weight Distribution on the Aeroelastic Instability of Swept Wings

In this work, a study has been conducted to observe the influence of fuel weight distribution on the flutter characteristics of a high aspect ratio swept wing. In this paper, B777-200 wing model having a 34º sweptback angle is used as a baseline and two other models with the swept angles of 0º (straight wing) and 30º (forward swept) are considered in this study. Aeroelastic analysis is performed ranging from sea level up to 35,000 ft and the influence of in-flight fuel management in several flight altitudes is also investigated. There are four different fuel distribution model are investigated and it was found that some fuel distribution configuration are more critical which may lead to flutter. Moreover it was found that the flight altitude significantly affects the aeroelastic instability.

Frequency-domain analyses on the aeroelastic characteristics of thrust-vectored system on airship

This paper investigates the aeroelastic performance of a thrust-vectored system on aerostat vehicle, which is characterized by capsule-beam-rotor or capsule-wing-rotor configuration. The distinct inherent and environmental factors, like rotor tilt, wing sweep, air condition etc., are taken into account. The variation principle is utilized to establish the governing equations, where the aerodynamic loads of wing and rotor which highly couple with the structural deformations are modeled with accordant accuracies. The physical model is applied, which is especially appropriate for the aeroelastic analyses of tilt-rotor systems in low-speed flight condition. The Rayleigh-Ritz method is employed to derive the frequency-domain modal solutions. The case studies consider three contrastive models to investigate the aeroelastic mechanisms by means of parametric analyses. The results show that, except for the installation geometries of rotor, all the distinct parameters cause implications on the amplitudes and trends of modal frequencies and damping rates; the rotor rotation interacts with the structural deformations severely, leading to the occurrence of complicated coupling phenomenon; in addition, the relevant parameters also show optimizable potentials to postpone the aeroelastic instabilities.

Aeroelastic stability analysis and failure damage evaluation of wind turbine blades under variable conditions

ABSTRACT Failure during the start and brake processes of a 1.5 MW horizontal axis wind turbine blade was examined using a measured wind field with fluid-structure coupling. The maximum blade displacement and stress with angular acceleration during the start process were 7.14% and 16.27%, respectively, larger than those without acceleration, whereas during braking they were 37.71% and 26.96% larger, respectively, without acceleration. Further, angular acceleration had a greater impact on the second-order frequency of vibration than the first-order, significantly impacting the braking process. The maximum dynamic stress and amplitude of the blades during starting were 4.35 MPa and 0.56 m, and during braking were 35.86 MPa and 3.13 m, respectively. Based on drone images of the blades, several failure modes were identified. This research will provide key inspection area guidance for employing drones to inspect blades to improve inspection efficiency. A large amount of damage images are collected and classified into damage levels, and an image processing system are established, which can quickly identify damage characteristics of the blade and export the inspection report quickly and accurately. Through a large number of drone detection experiments, this technology has a wide range of application prospects and extremely high efficiency.