Limit cycle oscillations of cantilever rectangular wings designed using topology optimisation

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in Proceeding of AIAA SciTech 2020 Forum (2020, January)

A closed form state-space model for the nonlinear aeroelastic response of thin cantilevered flat plates is derived using a combination of von Karman thin plate theory and a linearized continuous time vortex lattice aerodynamic model. The modal-based model is solved for the amplitude and period of the limit cycles of the flat plates using numerical continuation. The resulting predictions are compared to experimental data obtained from identical flat plates in the wind tunnel. Both conventional and topologically optimised flat rectangular plates are investigated. It is shown that the aeroelastic model predicts the linear flutter conditions and nonlinear response of the plates with reasonable accuracy, although the predicted limit cycle amplitude variation with airspeed is different to the one measured experimentally due to unmodelled physics.

A database of flutter characteristics for simple low and medium aspect ratio wings at low speeds

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in Proceeding of AIAA SciTech 2020 Forum (2020, January)

Flutter is a catastrophic aeroelastic instability that must be prevented for all engineering structures operating in an airflow. Several standard methods exist for predicting the flutter onset airspeed and flutter frequency of lifting surfaces, such as the Doublet Lattice Method and the Vortex Lattice Method. Such techniques can be applied by means of industrial standard aeroelastic software packages, such as Nastran and ZAERO. Nevertheless, researchers are constantly developing new aeroelastic prediction software for various applications and with different capabilities. Such developments must be validated using either standard benchmark solutions or experimental data. While such benchmarks do exist, they are mostly targeted at compressible flows around aircraft wings. The purpose of the present work is to present a new experimental database of flutter airspeeds and frequencies for a significant number of simple wings at low subsonic airspeeds. The wings feature a variety of aspect and taper ratios, focusing in the low and medium aspect ratio range, from 2 to 5.4. This paper presents results from the wind tunnel experiments, including flutter speeds, flutter frequencies, as well as natural frequency and damping ratio variation with airspeed for all wings. Predictions for the flutter characteristics of the wings are obtained from the Vortex Lattice method and compared to the experimental measurements. It is shown that the predictions are very good, except for the wings with Aspect Ratios close to 2, for which errors in flutter speed of up to 28% were obtained. The flutter frequency was predicted accurately for all wings.

Numerical and Experimental Investigation of Tandem Wing Flyers

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in Proceedings of the AIAA SciTech 2019 Forum and Exhibition (2019, January)

The recent focus on micro-UAV systems and bio-inspired drones has generated interest in tandem wing applications. Dragonfly-based configurations are of significance for very low Reynolds numbers; for larger drones, Microraptor-based geometries could prove to be efficient. The present study of tandem wing flyers aims at understanding the basic principles governing the aerodynamic properties of tandem wings in close proximity. The analysis includes both numerical simulations by means of the Unsteady Vortex Lattice Method and wind tunnel experimentation applied to generic rectangular wing geometries. Preliminary conclusions include the facts that increasing the rear wing’s angle of attack results in a bigger increase in lift than increasing the front wing’s angle of attack. The dihedral angles of the two wings also seem to have significant impact on the lift, some configurations leading to an increase in lift coefficient of up to 25%. The insight provided by the results will be used in the future to test and validate different flight configurations for the Microraptor and, hopefully, to shed some light on its preferred in-flight configuration and its flight capabilities.

Computation of Leishman-Beddoes model parameters using unsteady RANS simulations

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in Proceedings of the AIAA SciTech 2019 Forum and Exhibition (2019, January)

The determination of the Leishman-Beddoes (LB) model parameter values from experimental measurements and Computational Fluid Dynamic (CFD) simulations is presented. Two-dimensional unsteady RANS simulations are carried out for experimental test cases of two airfoils oscillating in pitch in the wind tunnel at Reynolds numbers of the order of Re=1.8e4. The RANS results are first compared directly to the experimental measurements and it is shown that the simulations cannot represent some important aspects of the physics of the phenomenon, such as the effect of a laminar separation bubble occurring at low angles of attack and the effect of the resulting leading edge vortex once it has started travelling over the surface of the airfoil. The flowfields computed from the CFD simulations are then used to estimate the values of three parameters for the Leishman-Beddoes model. The aim is to explore the possibility of using the LB model as a reduced order model for CFD simulations. It is shown that the inaccuracies of the CFD simulations lead to inaccurate parameter values, such as an overestimation of the leading edge vortex shedding time. Nevertheless, the resulting LB model can smooth the oscillations in the post-stall load responses predicted by the CFD. It is concluded that higher-fidelity simulations are necessary, involving a boundary transition model or even Large Eddy Simulation schemes.

A Full Potential Static Aeroelastic Solver for Preliminary Aircraft Design

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in Proceedings of the 18th International Forum on Aeroelasticity and Structural Dynamics (IFASD2019) (2019)

There is a consensus in the aerospace research community that future aircraft will be more flexible and their wings will be more highly loaded. While this development is likely to increase aircraft efficiency, it poses several aeroelastic questions. Current aeroelastic tailoring practice for early preliminary aircraft design relies on linear aerodynamic modeling, unable to predict shocks. On the other hand, nonlinear solvers, although they provide a wide range of functionality and are reliable, often consist in monolithic code structures and cannot be efficiently coupled to external structural mechanics codes. They are therefore usually not readily usable for coupled fluid-structure interaction computations. The objective of the present work is to carry out aerodynamic and static aeroelastic computations in the context of preliminary aircraft design. To this end, an open-source, fast and reliable, unstructured finite element, Full Potential solver has been developed. Preliminary results are presented and show a significant improvement over the classical linear potential method and are in good agreement with higher fidelity nonlinear solvers.

Demonstration of a unified and flexible coupling environment for nonlinear fluid-structure interaction

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Conference (2018, September 13)

Higher Fidelity Transonic Aerodynamic Modeling in Preliminary Aircraft Design

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in Proceedings of the 31st Congress of the International Council of the Aeronautical Sciences (2018, September 11)

There is a consensus in the aerospace research community that future aircraft will be more flex- ible and their wings will be more highly loaded. While this development is likely to increase air- craft efficiency, it poses several aeroelastic ques- tions. Current aeroelastic tailoring practice for early preliminary aircraft design relies on linear aerodynamic modeling, which is unable to pre- dict shocks and boundary layers. The objective of this research is to enhance the linear aerodynamic modeling methodology, thus allowing fast and re- liable aerodynamic loads prediction for aeroelas- tic computations. First, the different levels of fi- delity of aerodynamic modeling that can be used in aircraft design are reviewed and compared on benchmark test cases. A Field Panel Method is subsequently developed and implemented. Pre- liminary results are presented and possible future enhancements are detailed.

Inviscid and viscous flow modeling for fast transonic flutter calculations

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in Proceedings of the International Council of Aeronautical Sciences, ICAS 2018 (2018, September 10)

Transonic aeroelastic analysis at the design level relies on linear panel methods, such as the Doublet Lattice approach, usually after application of transonic corrections. The results from these calculations cannot predict shock motion, shock-boundary layer interactions and the effects of such phenomena on flutter behavior, even after corrections are applied, since the latter are generally quasi-steady. This paper proposes a higher-fidelity approach that involves the solutions of the flow equations in order to obtain the unsteady flow response to relatively small amplitude periodic deformations of a structure over a large range of oscillation frequencies. The main idea is to perform a few high-fidelity CFD simulations, such as Euler or RANS simulations, with an imposed structural deformation at selected oscillation frequencies so as to capture the most dominant nonlinear dynamic modes of the flow response. These fluid dynamic modes are then interpolated to estimate the flow response for any other oscillation frequency. The methodology can then be used to obtain a frequency-domain generalized aerodynamic force matrix, and stability analysis can be performed using standard flutter calculation methods such as the p-k method. The present methodology provides a very good estimate of the flutter boundary for the 2D Isogai airfoil validation case, but at much lower computational cost than the traditional higher-fidelity Fluid-Structure Interaction (FSI) simulations.

Unsteady aerodynamics and nonlinear dynamics of freefalling rotating seeds

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in Proceedings of the International Conference on Noise and Vibration Engineering, ISMA 2018 (2018, September)

This work presents a three-dimensional numerical tool that is suitable for studying the aerodynamics and nonlinear dynamics of free-falling rotating seeds like samaras. The proposed simulation framework consists of a modified version of the unsteady vortex-lattice method (enhanced by including a diffusion model and the leading-edge vortex contribution by means of the Polhamus analogy) coupled with a multibody rigid dynamic model for the whole seed. The numerical scheme adopted by the aerodynamic subsystem is based on an explicit low-order integrator (Euler's explicit first-order method). On the other hand, the equations of motion associated with the structural part are integrated in the time domain using a second-order Lie group integrator based on an extension of the classical generalized- method for dynamical systems. Among the main results obtained, it is found that the predicted terminal descending velocity and angular velocity (around the vertical axis) are in close agreement with experimental results reported in the literature.

New aerodynamic modeling methodology for transonic flutter calculations

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Conference (2018, July 04)

Two Methods for Modeling Unsteady Transonic Flows at Low Computational Cost

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Conference (2017, July 06)

Unsteady pressure distributions on a 4:1 rectangular cylinder: comparison of numerical and experimental results using decomposition methods.

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Conference (2017, July 04)

Detached flows around bluff bodies are ubiquitous in civil engineering applications. In this work, the flow around a static 4:1 rectangular cylinder at moderate Reynolds number and at different angles of incidence is studied using both Experimental Fluid Dynamics (EFD) and Computational Fluid Dynamics (CFD). Typically, the integration of EFD and CFD allows a better understanding of the flow of interest by leveraging the complementary of their respective outputs. However, the comparison of computational and experimental results is an important but difficult step of this integration, particularly in the case of local quantities related to unsteady flows. In this work, decomposition methods are used to compare unsteady loads and pressure distributions coming from EFD and CFD. In particular, Proper Orthogonal Decomposition (POD) and Dynamic Mode Decomposition (DMD) are used to extract the dominant structures of the aerodynamic coefficients. The experimental data are obtained from dynamic pressure measurements in wind tunnel while numerical data come from two-dimensional unsteady Reynolds-Averaged Navier-Stokes (uRANS) simulations and tri-dimensional Delayed-Detached Eddy Simulations (DDES). This work shows that the decomposition methods represent a powerful tool enabling the analysis and the quantitative comparison of the main spatial and temporal characteristics of unsteady flows. Moreover, the accuracy of uRANS and DDES results is analyzed in light of the capacity of both CFD techniques to capture the reattachment occurring on the upper part of the rectangular cylinder.

A Discussion on the Advancement of Blade Tip Timing Data Processing

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in Proceedings of the Turbomachinery Technical Conference & Exposition, TURBO EXPO 2017 (2017, June 29)

The Blade Tip Timing method (BTT) is a well-known approach permitting individual blade vibration behavior characterization. The technique is becoming increasingly popular among turbomachinery vibration specialists. Its advantages include its non-intrusive nature and its capability of being used for long-term monitoring, both in on-line and off-line analysis. However, the main drawback of BTT is frequency aliasing. Frequency aliasing effects in tip timing can be reduced by means of the application of different methods from digital signal analysis that can exploit the non-uniform nature of the data sampled by BTT. This non-uniformity is due to the fact that an optimization of the circumferential distribution of BTT probes is usually required in order to improve the data quality for targeted modes of blade vibration and/or orders of excitation. The BTT data analysis methods considered in this study are the non-uniform Fourier transform, the minimum variance spectrum estimator approach, a multi-channel technique using in-between samples interpolation, the Lombe-Scargle periodogram and an iterative variable threshold procedure. These methods will be applied to measured data representing quite a large scope of events occurring during gas-turbine compressor operation, e.g. synchronous engine order resonance crossing, rotating stall, suspected limit-cycle oscillations. Finally, the frequency estimates obtained from all these methods will be summarized.

Limit cycle oscillations of cantilever rectangular flat plates in a wind tunnel

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2017 (2017, June 27)

A closed form state-space model of the nonlinear aeroelastic response of thin cantilevered flat plates is derived using a combination of Von Karman thin plate theory and a linearized continuous time vortex lattice aerodynamic model. The modal-based model is solved for the amplitude and period of the limit cycles of the flat plates using numerical continuation. The resulting predictions are compared to experimental data obtained from identical flat plates in the wind tunnel. It is shown that the aeroelastic model predicts the linear flutter conditions and nonlinear response of the plates with reasonable accuracy, although the predicted limit cycle amplitude variation with airspeed is different to the one measured experimentally due to unmodelled physics.

Dynamic stall onset variation with reduced frequency for three stall mechanisms

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2017 (2017, June 27)

A set of unsteady aerodynamic load measurement is performed on three oscillating airfoils with distinct stall mechanisms: a flat plate, a NACA0012, and a NACA0018. The airfoils are forced to oscillate in pitch around the stall angle of attack with prescribed frequency and amplitude. A criterion proposed by Sheng et al. is used to locate the onset of the flow separation process associated with dynamic stall, and quantify its variation with an equivalent reduced pitch rate. The validity of this criterion is tested for the three airfoils at low Reynolds number, Re = 2 × 10^4. Results are compared with the experimental data obtained by Sheng et al. at higher Reynolds number of Re = 1.5 × 10^6.

Research on Fast Aeroelastic Modeling Methods for the Transonic Regime

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2017 (2017, June 27)

Two methods for modeling unsteady transonic flows at low computational cost are presented as a first step towards a fast and accurate aeroelastic calculation methodology for the preliminary design stage in the transonic flow regime. The first approach corresponds to a quasi-steady approximation based on few steady simulations that is improved through the use of an unsteady filter. The second approach is based on the interpolation of dynamic modes between solutions at different frequencies that are obtained either from Dynamic Mode Decomposition (DMD) of unsteady simulations or directly from Harmonic Balance (HB) simulations. The two methods are illustrated in the case of a pitching airfoil in the transonic regime. Results show that the first method is fast and provides a first approximation of the unsteady dynamics. The computational cost of the second approach is higher, but the method provides better results in predicting aerodynamic forces and shock motion for a large range of reduced frequencies.

Freeplay-induced limit cycle oscillation mitigation using linear and nonlinear tuned vibration absorbers

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in Proceeding of the IFASD 2017 Conference (2017, June 27)

Structural nonlinearities such as freeplay in control surface bearings and actuators or in connections between wings and external payloads sometimes lead to aeroelastic limit cycle oscillations at airspeeds lower than the linear flutter speed of the aircraft. In parallel, numerous studies demonstrated the potential of linear and nonlinear tuned vibration absorbers to increase the flutter speed of linear and continuously hardening aeroelastic systems such as two-degree- of-freedom wings or long span bridges. In this work, the effect of linear and nonlinear tuned vibration absorbers is studied on a wing with pitch plunge and control surface deflection degrees of freedom and with freeplay in pitch. Depending on the tuning of the linear absorber, the linear flutter speed of the system can be increased by 10% or the onset of limit cycle oscillations due to the freeplay can be delayed by 7.7% and their amplitude can be significantly decreased. The addition of cubic hardening forces on the absorber can further decrease the limit cycle amplitude in a limited airspeed range at the cost of an increase in limit cycle amplitude in another airspeed range. Conversely, the addition of a freeplay hardening force on the absorber can decrease the limit cycle amplitude without any detrimental effect.

System Eigenvalue Identification Of Mistuned Bladed Disks Using Least-Squares Complex Frequency-Domain Method

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in Proceeding of the Turbomachinery Technical Conference & Exposition TURBO EXPO 2017 (2017, June 26)

This paper presents the results from a research effort on eigenvalue identification of mistuned bladed rotor systems using the Least-Squares Complex Frequency-Domain (LSCF) modal parameter estimator. The LSCF models the frequency response function (FRF) obtained from a vibration test using a matrix-fraction description and obtains the coefficients of the common denominator polynomial by minimizing the least squares error of the fit between the FRF and the model. System frequency and damping information is obtained from the roots of the denominator; a stabilization diagram is used to separate physical from mathematical poles. The LSCF estimator is known for its good performance when separating closely spaced modes, but few quantitative analyses have focused on the sensitivity of the identification with respect to mode concentration. In this study, the LSCF estimator is applied on both computational and experimental forced responses of an embedded compressor rotor in a three-stage axial research compressor. the LSCF estimator is first applied to computational FRF data obtained from a mistuned first-torsion (1T) forced response prediction using FMM (Fundamental Mistuning Model) and is shown to be able to identify the eigenvalues with high accuracy. Then the first chordwise bending (1CWB) computational FRF data is considered with varied mode concentration by varying the mistuning standard deviation. These cases are analyzed using LSCF and a sensitivity algorithm is developed to evaluate the influence of the mode spacing on eigenvalue identification. Finally, the experimental FRF data from this rotor blisk is analyzed using the LSCF estimator. For the dominant modes, the identified frequency and damping values compare well with the computational values.

Staggered strong coupling between existing fluid and solid solvers through a Python interface for fluid-structure interaction problems

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in Proceedings of the VII International Conference on Coupled Problems in Science and Engineering (2017, June)

Flutter and limit cycle oscillation suppression using linear and nonlinear tuned vibration absorbers

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in Proceedings of the SEM IMAC XXXV (2017, February)

Aircraft are more than ever pushed to their limits for performance reasons. Consequently, they become increasingly nonlinear and they are more prone to undergo aeroelastic limit cycle oscillations. Structural nonlinearities affect aircraft such as the F-16, which can undergo store-induced limit cycle oscillations (LCOs). Furthermore, transonic buzz can lead to LCOs because of moving shock waves in transonic flight conditions on many aircraft. This study presents a numerical investigation of passive LCO suppression on a typical aeroelastic system with pitch and plunge degrees of freedom and a hardening stiffness nonlinearity. The absorber used is made of a piezoelectric patch glued to the plunge springs and connected to a resistor and an inductance forming a RLC circuit. A mechanical tuned mass damper absorber of similar configuration is also considered. The piezoelectric absorber features significant advantages in terms of size, weight and tuning convenience. The results show that both types of absorber increase the linear flutter speed of the system in a similar fashion but, when optimal, they lead to a sub-critical bifurcation while a super-critical bifurcation was observed without absorber. Finally, it is shown that the addition of a properly tuned nonlinear spring (mechanical absorber) or capacitor (piezo- electric absorber) can restore the super-criticality of the bifurcation. The tuning of the nonlinearity is carried out using numerical continuation.

Unsteady lifting line theory using the Wagner function

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in Proceedings of the 55th AIAA Aerospace Sciences Meeting (2017, January)

A method is presented to model the incompressible, attached, unsteady lift and moment acting on a thin three-dimensional wing in the time domain. The model is based on the combination of Wagner theory and lifting line theory trough the unsteady Kutta-Joukowsky theorem. The result is a set of closed form linear ordinary di erential equations that can be solved analytically or using a Runge-Kutta-Fehlberg algorithm. The method is validated against numerical predictions from an unsteady Vortex Lattice method for rectangular and tapered wings undergoing step or oscillatory changes in plunge or pitch. As the aerodynamic loads are written in state space form in the proposed method, they can be easily included in aeroelastic and flight dynamic calculations.

Dynamic interactions of a supercritical aerofoil in the presence of transonic shock buffet

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in Proceedings of the International Conference on Noise and Vibration Engineering, ISMA 2016 (2016, September 19)

Within a narrow transonic flight region, shock-wave/boundary-layer interactions yield large amplitude, self sustained shock oscillations that are detrimental to both platform handling quality and structural integrity. In this study, the aeroelastic interactions between this transonic buffet instability and a spring-suspended supercritical aerofoil are investigated by means of Reynolds-Averaged Navier-Stokes simulations. Single degree-of-freedom pitching simulations are performed for a range of structural to aerodynamic frequency ratios, sectional mass ratios and levels of structural damping. The results show that for a range of pitch eigenfrequencies above the fundamental buffet frequency, sychronisation of the aerodynamic and structural modes occurs. This so called lock-in phenomenon acts as a mechanism for large amplitude Limit Cycle Oscillation in aircraft structures within the transonic flow regime. The sectional mass and the addition of structural damping are both found to have a pronounced effect on the nature of the limit cycles

Assessment of the fluid-structure interaction capabilities for aeronautical applications of the open-source solver SU2.

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in Proceedings of the VII European Congress on Computational Methods in Applied Science and Engineering (2016, September)

Passive flutter suppression using a nonlinear tuned vibration absorber

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Poster (2016, July)

A recent study showed that the addition of a linear tuned vibration absorber could increase the flutter speed of a rigid wing with pitch and flap degrees of freedom by about 35%. However, the absorber was turning the initial super-critical bifurcation into a sub-critical one. This work shows numerically that adding a nonlinear restoring force to the absorber can restore the su- percritical behaviour of the bifurcation and further reduce the post-instability limit cycle amplitude.

Research on Fast Aeroelastic Modeling Methods in the Transonic Regime

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Conference (2016, June 02)

Influence of propeller configuration on propulsion system efficiency of multi-rotor Unmanned Aerial Vehicles

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in Proceedings of the International Conference on Unmanned Aircraft Systems, ICUAS 2016 (2016, June)

Multi-rotor Unmanned Aerial Vehicles make use of multiple propellers, mounted on arms, to produce the required lift. This article investigates the influence on propulsion system efficiency in hover due to the configuration of these propellers. Influence of pusher or puller configuration of the propeller, number of blades, shape and dimensions of the arm, coaxial and overlapping propellers, is presented. A dedicated test bench that allows testing of various experimental setups is designed and built in order to realistically represent multi-rotor arms. Test results show that a two-bladed pusher configuration is most efficient and slenderness of the arm has more influence on efficiency than shape. A coaxial propulsion system approaches the efficiency of a single-prop system at high disk loadings. Finally, interference effects due to overlapping propellers are discussed.

Experimental Passive Flutter Mitigation Using a Linear Tuned Vibrations Absorber

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in Proceedings of the IMAC XXXIV Conference (2016)

The current drive for increased efficiency in aeronautic structures such as aircraft, wind turbine blades and helicopter blades often leads to weight reduction. A consequence of this tendency can be increased flexibility, which in turn can lead to unfavourable aeroelastic phenomena involving large amplitude oscillations and nonlinear effects such as geometric hardening and stall flutter. Vibration mitigation is one of the approaches currently under study for avoiding these phenomena. In the present work, passive vibration mitigation is applied to an experimental aeroelastic system by means of a linear tuned vibration absorber. The aeroelastic apparatus is a pitch and flap wing that features a continuously hardening restoring torque in pitch and a linear one in flap. Extensive analysis of the system with and without absorber at subcritical and supercritical airspeeds showed an improvement in flutter speed around 34%, a suppression of a jump due to stall flutter, and a reduction in LCO amplitude. Mathematical modelling of the experimental system showed that optimal flutter delay can be obtained when two of the system modes flutter simultaneously. However, the absorber quickly loses effectiveness as it is detuned. The wind tunnel measurements showed that the tested absorbers were much slower to lose effectiveness than those of the mathematical predictions.

PIV-based estimation of unsteady loads on a flat plate at high angle of attack using momentum equation approaches

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Conference (2015, November 23)

The direct measurement of loads with force balance can become challenging when the forces are small or when the body is moving. An alternative is the use of Particle Image Velocimetry (PIV) velocity fields to indirectly obtain the aerodynamic coefficients. This can be done by the use of control volume approaches which lead to the integration of velocities, and other fields deriving from them, on a contour surrounding the studied body and its supporting surface. This work exposes and discusses results obtained with two different methods: the direct use of the integral formulation of the Navier-Stokes equations and the so-called Noca's method. The latter is a reformulation of the integral Navier-Stokes equations in order to get rid of the pressure. Results obtained using the two methods are compared and the influence of different parameters is discussed. The methods are applied to PIV data obtained from water channel testing for the flow around a 16:1 plate. Two cases are considered: a static plate at high angle of attack and a large amplitude imposed pitching motion. Two-dimensional PIV velocity fields are used to compute the aerodynamic forces. Direct measurements of dynamic loads are also carried out in order to assess the quality of the indirectly calculated coefficients.

Suppression of simulated self-excited oscillation using smart materials on flexible wing structure

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2015 (2015, July 01)

Suppression of simulated self-excited oscillation due to aeroelastic effects using piezoelectric patches is reported. The focus of the present work is suppressing simulated flutter oscillation using piezoelectric patches bonded to the wing structure. First, a clean wing is exposed to airflow in the wind tunnel where it experienced limited amplitude oscillation. The responses were recorded. This is followed by bonding piezoelectric patches to the identical wing that was tested in the wind tunnel. Two of the patches were used as actuators to simulate self-excited oscillation in a control manner. The selected mode for excitation is 1st bending mode. The other two patches were used as suppressor for active control using negative velocity feedback Single-Input, Single-Output approach. The single input signal for negative feedback is also sensed using piezoelectric patches. The controller manages to suppress the simulated flutter response to a lower oscillation amplitude values. This works demonstrated the used of piezoelectric material as actuator to reproduce the oscillation amplitude during self-excited oscillation and suppress the oscillation at the same time.

Asymmetric limit cycle oscillations in systems with symmetric freeplay

in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2015 (2015, June 30)

In this paper, a simple 2D aeroelastic system with degrees of freedom in pitch, plunge and control surface deflection is investigated with freeplay in the pitch degree of freedom. It is shown that this system features a fixed point at the origin as well as two anti-symmetric fixed points. Asymmetric limit cycles that span only two of the three piecewise linear subdomains of the phase plane can orbit these anti-symmetric fixed points. A single three-domain symmetric cycle and two two-domain cycles appear as a result of a grazing bifurcation occurring at the flutter speed of the underlying linear system. The two-domain cycles can undergo further fold, period doubling and torus bifurcations. They can cause both periodic and aperiodic oscillations, including highly chaotic responses at parameter values where they interact strongly.

Using 2D-PIV measurements to compute unsteady aerodynamic loads on a flat plate at high angle of attack

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in Proceedings of the 16th International Forum on Aeroelasticity and Structural Dynamics, IFASD 2015 (2015, June 30)

This work exposes and discusses results obtained for aerodynamic forces using an indirect calculation based on Particle Image Velocimetry (PIV) measurements. The methodology used is based on the integral formulation of the Navier-Stokes equations and is applied to spatio-temporal data for different flows around a plate with a 16:1 chord-to-thickness ratio at high angle of attack. Experimental data are obtained in a water channel for both a static and a pitching plate. In addition to PIV data, direct measurements of aerodynamic loads are carried out to assess the quality of the indirect calculation. It is demonstrated that this indirect method is able to compute the mean and the temporal evolution of the lift and drag coefficients with a reasonable accuracy. It is also shown that the noise sensitivity of the method can be partly alleviated through the use of Dynamic Mode Decomposition (DMD) as a pre-processing step to smooth the spatio-temporal data.

Unsteady Viscous-Inviscid Coupling Simulations of Separated Laminar Flows Around 2D Airfoils

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2015 (2015, June 29)

An interactive boundary layer model has been developed in 2D in order to solve the unsteady flow around an airfoil. The inviscid problem is solved using a panel method, by the discretization of the airfoil into linear-varying vortex panels. The solution of the boundary layer equations is carried out using a finite volume scheme. Viscous-inviscid coupling is preformed by imposing a permeation velocity on the skin panels of the airfoil and the addition of a free wake at each separation point.

Theoretical and experimental investigation of a structurally and aerodynamically nonlinear pitch and flap wing

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2015 (2015, June 29)

This paper presents and experimental and theoretical investigation of a novel nonlinear aeroelastic system. It consists of a wing with pitch and flap degrees of freedom, suspended from a leaf spring secured in a nonlinear clamp. Both the structural and the aerodynamic forces acting on the wing can be nonlinear, depending on the amplitude of oscillations. Wind tunnel experiments show that the system undergoes a supercritical Hopf bifurcation that leads to small amplitude limit cycle oscillations. At a particular airspeed, the pitch amplitude jumps to a much higher value and dynamic stall starts to occur. Three mathematical models of the system are formulated, one based on linear aerodynamics and two based on the Leishman-Beddoes dynamic stall model. The objective of the modelling is to determine whether the jump in pitch oscillation amplitude is due to dynamic stall. The predictions for amplitude, frequency and mean angle of the limit cycle oscillations are compared to the experimental observations. All three models predict the small amplitude oscillations with satisfactory accuracy. The complete Leishman-Beddoes model predicts the occurrence of a jump in pitch amplitude but the magnitude of this jump is signi cantly overestimated. The other two models completely fail to model the jump. The failure of the Leishman-Beddoes model to predict the correct post-jump oscillation amplitude may be due to the values selected for the model parameters.

Flapping around in a wind tunnel

Conference (2015, March 17)

Aerodynamic investigations of animal flight have usually concentrated on static conditions. These are good enough to analyse gliding flight but not other types of flight, such as cruise (migrations), take-off and landing or manoeuvres. In this work we present wind tunnel experiments on geese flying in a wind tunnel and unsteady aerodynamic simulations of these flights, based on the Vortex Lattice Method.

Internal Resonance and Stall-Flutter in a Pitch-Flap Wing in the Wind-Tunnel

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in Proceedings of the SEM IMAC XXXII (2015, February)

Nonlinear aeroelastic phenomena such as store-induced LCOs, transonic buzz and stall flutter are the burden or modern aircraft: they reduce the performance and can even limit the flight envelope in both civil and military cases. Several nonlinear setups were studied experimentally in the last decades by the scientific community but most of them have pitch and plunge degrees of freedom and feature a rigid wing. In this paper, we study a new nonlinear aeroelastic apparatus that features nonlinear pitch and flap degrees of freedom, coupled with a flexible wing. The model is tested experimentally in the wind tunnel to determine its dynamic behaviour. Preliminary observations demonstrate that the system undergoes a supercritical Hopf bifurcation due to the hardening nonlinearity followed by an amplitude jump that is the consequence of either dynamic stall (i.e. stall flutter) or internal resonance (i.e. interaction between the hardening nonlinearity and higher modes).

Impact of Circularity Defect of Helical Fillets HDPE Bridge Stay Covers: Analysis of Bi-Stability at Critical Reynolds Number by Bifurcation Diagrams

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in Proceedings of the Symposium on the Dynamics and Aerodynamics of Cables - SDAC 2014 (2014, September)

Dry galloping of cables is still considered by the scientific community as an important phenomenon lacking investigation on cable stayed bridge. The helical fillet cylinders developed to overcome the Rain-Wind Induced Vibration (RWID) were reported to be affected by the appearance of vibration in dry conditions. This paper focuses on a method developed to process data issued from wind tunnel tests on original High Density Polyethylene (HDPE) cable covers with helical fillets, in a range of Reynolds numbers from the sub-critical regime to the critical regime. Corresponding values range from Re = 9.6x10^4 to Re = 3.6x10^5 for the present work. The experiment consists of investigating the effect of the helical fillets with natural or controlled deformation of his circularity defect. Previous studies ([1], [3] and [4]) reported that dry galloping is caused by the appearance of a negative pressure bubble on one side of a circular cylinder at the critical Reynolds number range, leading to a rapid drop in the drag coefficient and the appearance of a non negligible lift force. Because the flow oscillates between two states, say TrBL0 and TrBL1, giving way to a fluctuating lift force supposed to cause the vibration of cables. However, [1] reports that the TrBL1 regime disappears in the presence of helical fillets. The aim of the current study was to investigate a possible appearance of the TrBL1 regime when the circularity defect threshold exceeds the filet thickness. In this case, the second purpose was to characterize the bi-stability of this phenomenon and provide new tools for wind engineering processing.

Wind Tunnel Experiments On Bridge Stays Cables Protection Tubes In Dry Galloping Conditions: Processing Method For Bi-Stable Phenomenon

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in Proceedings of the Symposium on the Dynamics and Aerodynamics of Cables - SDAC 2014 (2014, September)

Surface irregularity is one of the parameters that have not been deeply examined on stay cables of cable-stayed bridges. This paper focuses on wind tunnel tests on original High Density Polyethylene cable covers, in a range of Reynolds numbers from the sub-critical regime to the critical one, corresponding to values ranging from Re = 9.6x104 to Re = 3.3x105. The experiment consists in measurement of parietal pressures on cable covers in order to investigate the effect of surface irregularity on the mechanism of dry galloping excitation. Previous studies ([1], [3] and [4]) reported that dry galloping of cables with a circular cross section is linked to the appearance of a negative pressure bubble, on one side of the circular cylinder at the critical Reynolds number range, leading locally to a rapid drop in the drag coefficient and the appearance of a no negligible lift force. But such a steady change in local lift force does not generate an alternate excitation on a whole cable. The question of the spatial correlation of this “one bubble “ regime along a cable in real conditions and the capability of this lift force to be varied with time are still not resolved. The present paper aims at showing one first exploration of the spatial and temporal correlation of the pressure pattern along the HDPE tube with respect to the natural circularity defect. The Proper Orthogonal Decomposition (POD) was used to characterize the bi-stability phenomenon occurring at critical Reynolds number regime.

Advanced Aeroservoelastic Modeling for Horizontal axis Wind Turbines

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in Cunha, A.; Caetano, E.; Riberio, P.; Müller, G. (Eds.) Proceedings of the 9th International Conference on Structural Dynamics, EURODYN 2014 (2014, July)

This paper describes the development of a complete methodology for the unsteady aeroelastic and aeroservoelastic modeling of horizontal axis wind turbines at the design stage. The methodology is based on the implementation of unsteady aerodynamic modeling, advanced control strategies and nonlinear finite element calculations in the S4WT wind turbine design package. The aerodynamic modeling is carried out by means of the unsteady Vortex Lattice Method, including a free wake model. The complete model also includes a description of a doubly fed induction generator and its control system for variable speed operation and enhanced power output. The S4WT software features a non-linear finite element solver with multi-body dynamics capability. The complete methodology is used to perform complete aeroservoelastic simulations of a 2MW wind turbine prototype model. The interaction between the three components of the approach is carefully analyzed and presented here.

Wind Tunnel Testing of a VTOL MAV Propeller in Tilted Operating Mode

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in Proceedings of the 2014 International Conference on Unmanned Aircraft Systems (ICUAS 2014) (2014, May 30)

This paper presents experimental results of the full 3-axis force vector and 3-axis moment vector acting on a propeller, commonly used for a Vertical Take Off and Landing Micro Aerial Vehicle (VTOL MAV). Measurements were carried out in a wind tunnel using a high resolution 6-axis force/moment sensor embedded in a customized test rig at several wind speeds, propeller rotational speeds and angles of the propeller shaft with respect to the air stream. Results show strong moments acting on the propeller in forward flight and unstable conditions in descending flight. Power calculations reveal a decrease in power consumption during slow forward flight and how motor efficiency can be maximized.

Using POD and DMD for comparing CFD and experimental results in unsteady aerodynamics

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Conference (2014)

Dry galloping on bridge cables: roughness and shape effects on an inclined circular cylinder

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in Proceedings of the European-African Conference on Wind Engineering 2013 (2013, July 09)

Dry galloping of stays in cable-stayed bridges is a phenomenon that is still of interest to the scientific community. Over the course of the current three-year research project numerous experiments will be carried out to investigate the effect of turbulence and inclination angle on dry galloping. The novel aspect of this study consists of an in-depth investigation into the effect of non uniformity of the pipe surrounding the stay cables on the dry galloping excitation mechanism. Previous studies reported that dry galloping is caused by the appearance of a negative pressure bubble on one side of the circular cylinder at the critical Reynolds number range, leading to a rapid drop in the drag coefficient and the appearance of a non negligible lift force. This paper presents results from initial experiments in the wind tunnel at CSTB – Nantes, France.

Numerical and Experimental Investigation of Slot-Blowing Air Over a Cylinder

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in Proceedings of the 5th European Conference for Aeronautics and Space Sciences (2013, July)

Tangential slot blowing around a cylinder in the Coanda effect has been proposed as a mechanism for yaw control in a novel coaxial helicopter. The proposed design has been investigated computationally, using the unsteady k-ω SST solver in NUMECA, as well as experimentally in the wind tunnel at the Université de Liège. While the concept in general is promising, it has been shown that there are some potential problems, including pitch-yaw coupling and some unsteady flow conditions. These problems exist under varied circumstances, and, exist at least in part due to the complicated flow-field in this problem, even in 2-dimensions.

Using Proper Orthogonal Decomposition and Dynamic Mode Decomposition Methods for Comparing CFD Results Experimental Measurements

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in Proceedings of the 15th International Forum on Aeroelasticity and Structural Dynamics, IFASD 2013 (2013, June 26)

A method for the quantitative comparison of numerical and/or experimental data of unsteady aerodynamics around static and oscillating bodies is introduced. It is based on Proper Orthogonal Decomposition (POD) and Dynamic Mode Decomposition (DMD) to extract the dominant structures of the unsteady flow. The proposed method is applied to spatio-temporal data for the flow around a 4:1 rectangular cylinder. Exper- imental data are obtained from wind tunnel testing and two dimensional Time-resolved Particle Image Velocimetry (Tr-PIV) measurements, while unsteady Reynolds Averaged Navier-Stokes (uRANS) are used to compute numerical results. It is demonstrated that the two approaches are complementary and represent a powerful tool that enables the analysis and the quantitative comparison of the main spatial (POD) and temporal (DMD) characteristics of unsteady aerodynamic data.

Aeroelastic response of a 2-DOF wing with structural and aerodynamic nonlinearity

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2013 (2013, June 25)

The aeroelastic behaviour of a wing oscillating in the heave and pitch degrees of freedom with continuos nonlinear stiffness is examined by means of wind tunnel experiments. The phenomena of interest are classical flutter and limit cycle oscillation. The focus of the present work is the transformation of classical flutter into LCOs by varying the stiffness from linear to nonlinear stiffness. The interaction of flow separation-related nonlinearity with structural nonlinearities is also of interest. The measured aeroelastic responses are analyzed and the bifurcation behavior of the dynamic system is characterized. The analysis shows that the bifurcation behaviour is dictated neither by the structural nor from the aerodynamic nonlinearity but by a combination of the two.

Simulation of bird wing flapping using the unsteady vortex lattice method

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in Proceedings of the International Forum on Aeroelasticity and Structural Dynamics, IFASD 2013 (2013, June 24)

The flight of barnacle geese at airspeeds representing high-speed migrating flight is investigated using experiments and simulations. The experimental part of the work involved the filming of three barnacle geese flying at different airspeeds in a wind tunnel. The video footage was analysed in order to extract the wing kinematics. Additional information, such as wing geometry and camber was obtained from a 3D scan of a dried wing. An unsteady vortex lattice method was used to simulate the aerodynamics of the measured flapping motion. It was found that the simulation correctly predicted the shape of the wake and the fact that the wing is aerodynamically inactive during the upstroke. The predicted aerodynamic lift and thrust forces were very sensitive to the wing's camber. Future work will attempt to identify the sensitivity to all the wing shape and kinematic parameters. Additional phenomena will also be modelled, including passive twisting of the wingtip, body lift and leading edge suction.

Dynamic Stall and Stall Flutter Simulations for a 2D Airfoil Using Viscous-Inviscid Coupling

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in Proceedings of the 54th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2013, April 10)

An interactive boundary layer model has been developed and coupled with a pitch plunge airfoil in 2d in order to solve the unsteady flow around the airfoil when undergoing light dynamic stall. The inviscid problem is solved by means of a panel method, by the discretization of the airfoil into vortex panels. The boundary layer is solved in a mixed manner, starting the solution in a direct way by imposing the external velocity and continuing it in an inverse way by imposing the displacement thickness. The solution of the boundary layer equations is carried out using a finite volume scheme. Viscous-inviscid coupling is preformed through the imposition of a permeation velocity on the skin panels of the airfoil and the addition of a free wake at each separation point.

Using Proper Orthogonal Decomposition Methods for Comparing CFD Results to Experimental Measurements

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in Proceedings of the 5th Symposium on Integrating CFD and Experiments in Aerodynamics (Integration 2012) (2012, October 03)

This work presents a method for quantitative comparison of numerical results to experimental measurements. It is based on the concept of Proper Orthogonal Decomposition. This technique is selected in order to compare the unsteady aerodynamic flows around static and oscillating bodies obtained from wind tunnel testing and numerical simulations. Two dimensional Time-resolved Particle Image Velocimetry measurements are carried out on the upper surface a 4:1 rectangular cylinder. Simulations are performed using unsteady Reynolds-Averaged Navier-Stokes and an unsteady Discrete Vortex Method. It is demonstrated that the proposed technique is a good preliminary step for comparing the main characteristics of unsteady aerodynamic data.

Damping identification of linear dynamic systems using Common-base Proper Orthogonal Decomposition

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in Proceedings of ISMA2012 (2012, September 19)

This paper presents a novel identification technique of the modal damping of linear systems. It is based on the Proper Orthogonal Decomposition (POD) of the free response of the system and extended to the Common-base POD (CPOD) approach. Different initial conditions are applied to the system and the corresponding free responses are considered simultaneously. The CPOD technique leads to a unique vector basis which is likely to contain more information about the dynamics of the system than a vector basis obtained by the classic POD technique, based on one set of initial conditions only. The ability of the technique to estimate the modal damping is demonstrated on a simulated mass-spring-damper system and an experimental system. Good agreement is shown between the damping estimates of the CPOD technique and the ones of the LSCF technique. The advantage and limitations of the present technique are discussed.

Numerical simulations of torsional flutter oscillations of a bluff body: Energy issues, GraSMech poster session

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Poster (2012, May)

The possibility of harvesting energy from aeroelastic phenomena is assessed through numerical simulations. The unsteady aerodynamics around an aeroelastic structure are modeled using the Discrete Vortex Method (DVM). The Torsional Flutter oscillations of a rectangular cylinder are studied in this work. The phenomenon is characterized by Limit Cycle Oscillations (LCOs) around the pitch axis of the rectangle. The complete bifurcation behaviour is investigated numerically and compared to wind tunnel measurements for validation. The energy issues are investigated in terms of the sensibility of the energy output to variations in the structural damping.

Initial Investigation of Chordwise Flexible Flat Aerofoil

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in Proceedings of the 53rd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2012, April 25)

The size of small Unmanned Air Vehicles (UAVs) makes them susceptible to gusts, hence an evaluation of their aeroelastic performance is a necessary requirement. With advances in materials, new aircraft are ever lighter and more flexible, both in the span-wise and chord-wise direction. An experiment is set up to evaluate the aeroelastic performance of a 2D chord-wise aerofoil structure under gust loading, where the wing of a small Unmanned Air Vehicle (UAV) is considered by coupling a Finite Element (FE) structural model with a Doublet-Lattice Method (DLM) aerodynamic model; a new semi-analytical modelling approach is also described. The flexible aerofoil structure is then optimised by means of a genetic algorithm (GA) framework for the minimum weight, subject to aeroelastic constraints of flutter, divergence and structural displacement.

Computational Considerations for the Prediction of Stall Flutter

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in Proceedings of the 53rd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2012, April 25)

A solver has been developed within the OpenFoam framework to compute large amplitude motion of two-dimensional rigid configurations. The results obtained with this code were successfully validated on rigid airfoils at static and dynamic conditions, as well as correlated with experimental data and numerical solutions from similar unsteady solvers. The results demonstrated that while current computational methods are able to predict the self-sustained oscillations characterizing a pitch-dominated stall flutter, including energy transfer, improvements are needed. The influence of grid, temporal integration, turbulence modeling, and flow equations is examined for the stall flutter starting solution of dynamic stall.

A Cross-Validation Approach to Approximate Basis Function Selection of the Stall Flutter Response of a Rectangular Wing in a Wind Tunnel

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in Proceedings of the 53rd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2012, April 25)

The stall flutter response of a rectangular wing in a low speed wind tunnel is modelled using a nonlinear difference equation description. Static and dynamic tests are used to select a suitable model structure and basis function. Bifurcation criteria such as the Hopf condition and vibration amplitude variation with airspeed were used to ensure the model was representative of experimentally measured stall flutter phenomena. Dynamic test data were used to estimate model parameters and estimate an approximate basis function.

Experiments on a pitch-plunge wing undergoing limit cycle oscillation

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in Proceedings of the 53rd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2012, April 25)

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The aeroelastic behaviour of a wing oscillating in the heave and pitch degrees of freedom is examined by means of wind tunnel experiment. The phenomena of interest are classical flutter and limit cycle oscillation. Classical flutter is normally associated with the exponential growth of the response amplitude. Linear flutter theory only predicts the critical flutter speed. Any excitation or disturbance beyond the critical speed is assumed to cause exponential growth in the response amplitude. In contrast, any limited amplitude oscillations occurring post-fultter suggest the existence of nonlinear properties in the system. Such properties can originate from the aerodynamic forces in the form of flow separation and reattachment. On the structural side, damping and stiffness can also contribute nonlinear properties. Furthermore, these nonlinearities can manifest themselves even at pre-flutter conditions, depending on the values of some governing parameter. The focus of the present work is the transformation of classical flutter into stall flutter as the equilibrium angle of attack of heaving and pitching wing is increased. The interaction of stall-related nonlinearity with structural nonlinearities is also of interest. The measured aeroelastic responses are analyzed and the bifurcation behavior of the dynamic system is characterized. Structural responses as well as flow field visualization through Particle Image Velocimetry show the origin of nonlinearity does not solely come from the manifestation of separation and the shedding of vortices, but from the structural nonlinearity which limits the response amplitude.

Discrete Vortex Simulations of the torsional flutter oscillations of a 4:1 rectangular cylinder

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in Proceedings of the 53rd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2012, April 24)

This work presents aeroelastic simulations of a 2D bluff-body around its pitching degree of freedom. The numerical tool consists in an aerodynamic solver based on the Discrete Vortex Method (DVM), coupled with a linear structural model. The shape of the bluff-body is a sharp edged rectangular cylinder with a side ratio equal to 4. The numerical results are compared to the experimental measurements recently obtained by the authors.8 The validation is carried out in three steps: first the frequency content of the flow-field in the wake of the static body is investigated. Then the simulated unsteady flow-field around the imposed pitching motion of the body is compared to experimental flow visualizations. This comparison is performed using Proper Orthogonal Decomposition (POD). Finally, the simulation of the global aeroelastic behaviour, based on the coupling of the DVM code to the structural model of the pitching degree of freedom is carried out and the results are compared to measured aeroelastic responses.. Very good agreements are found between numerical and experimental results, demonstrating the capabilities of the numerical tool to simulate complex unsteady aerodynamics around an oscillating bluff-body.

Computational Identification of Non-Linear Damping in an Aeroelastic System

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in Proceedings of the 53rd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2012, April 23)

An energy-based method is proposed to identify damping parameters from time histories of responses to sets of single-frequency harmonic excitation. The method is intended to be practically applicable to real structures and is able to identify the value of viscous damping, Coulomb friction and eventually other forms of non-linear damping models in aeroelastic systems. The inputs required are simply the accelerometer signals and the forces applied. It will be shown that if the system is undergoing Limit Cycle Oscillations, no external force is required for the identification process.

Unsteady Navier-Stokes simulation of low-Reynolds stall flutter

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in Proceedings of the 50th AIAA Aerospace Sciences Meeting (2012, January)

A solver has been developed within the OpenFoam framework to compute large amplitude motion of two-dimensional rigid configurations. The results obtained with this code were successfully validated on rigid airfoils at static and dynamic conditions, as well as correlated with experimental data and numerical solutions from similar unsteady solvers. The results demonstrate that current computational methods are, within the constraints imposed by spatial grids, temporal integration and turbulence modeling, capable of capturing the self-sustained oscillations characterizing stall flutter event with reasonable accuracy, including the mechanisms of energy transfer.

Aerodynamic modeling of horizontal axis wind turbines

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in Proceedings of the 13th International Conference on Wind Engineering (2011, July 14)

The paper presents the development of a vortex lattice aerodynamic modeling routine for SAMCEF for Wind Turbine (S4WT). S4WT is a general-purpose commercial code for wind turbine design and analysis. At present, it uses blade element momentum theory (BEM) for the estimation of the aerodynamics loads on HAWT rotor blades. BEM is a simple aerodynamic modeling approach that is currently used by several software packages for wind turbine design (as well as other rotor-based applications). It is based on the assumption that the flow can be treated as quasi-steady and quasi-2D, so that the steady, 2D aerodynamic loads acting on a strip of a rotor blade are used to estimate the instantaneous unsteady, 3D loads acting on a complete blade. This approach ignores the effect of the unsteady wake of the blades on the aerodynamic loads and simplifies the true 3D load distribution over the blades. A higher fidelity calculation of the time varying aerodynamic forces and moments acting on the blades is the main focus of this work. A good compromise between speed and accuracy to calculate these forces is the 3D unsteady vortex lattice method with a freely deforming wake. The vortex lattice results are compared to the BEM results from S4WT. The ultimate aim is to integrate the vortex lattice calculation as a subroutine in S4WT in order to calculate the unsteady aerodynamic forces on the rotor blades during the design process. This new method in S4WT will provide more representative results to the user, which can be very important for designing a more efficient wind turbine.

Influence of a static wing wake on the stall flutter behavior of a flexible wing

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in Proceedings of the 13th International Conference on Wind Engineering, ICWE13 (2011, July 13)

The subject of this paper is the experimental study of the aeroelastic behavior of a wing undergoing stall flutter in the vicinity of second, static wing. While stall flutter has been the subject of several investigations, such work has almost always concentrated on isolated wings. Stall flutter is a phenomenon that is mostly encountered in rotating blades, such as wind turbine or helicopter blades. In such cases, the phenomenon is influenced by the wake of the preceding blade. This paper presents a series of experiments carried out at the Goldstein Laboratory of the University of Manchester, concerning the phenomenon of stall flutter influenced by the proximity of a static wing. The work is an extension of the single wing stall flutter experiments presented by Dimitriadis and Li (2009).

Experimental Analysis of the Bifurcation Behaviour of a Bridge Deck Undergoing Across-Wind Galloping

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in De Roeck, G.; Degrande, G.; Lombaert, G.; Müller, G. (Eds.) Proceedings of the 8th International Conference on Structural Dynamics, EURODYN 2011 (2011, July 05)

The phenomenon of aeroelastic galloping is a very important design consideration for bridges and other slender structures. It has been investigated by a number of researchers but, most frequently, the analysis is limited to quasi-steady aerodynamic and linearized aeroelastic considerations. Such treatment has been shown to be effective for simple cross-sectional shapes, such as rectangles. In this work, an aeroelastic model of a realistic bridge deck cross-section is tested in a low speed wind tunnel. Both static and dynamic tests are carried out and the resulting force and vibration measurements are presented. The static force results are used to set up a quasi-steady mathematical model. The dynamic responses are used to draw a complete bifurcation diagram within a chosen airspeed range and to discuss the stability of the system. It is shown that the experimental system undergoes a subcritical Hopf bifurcation, its phase space including both a stable and an unstable limit cycle. As consequence, throughout the chosen airspeed, the system can either remain stable or undergo limit cycle oscillations. The quasi-steady analysis fails completely in capturing this type of behaviour. The predicted galloping onset speed is too conservative and the predicted oscillation amplitudes too high. The reason for this failure is the fact that the quasi-steady mathematical model is incapable of modelling subcritical Hopf bifurcations.

Double wake vortex lattice modeling of horizontal axis wind turbines

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in Proceedings of the 15th International Forum on Aeroelasticity and Structural Dynamics, IFASD 2011 (2011, June 29)

This paper presents the development of a vortex lattice based aerodynamic calculation for wind turbine blades. Emphasis is placed on the modeling of flow separation using the vortex lattice approach. This modeling is carried out by shedding a second free wake at the leading edge, which travels downstream at the local air velocity. The decision on whether to shed a wake or not is taken by looking at the sectional lift characteristics of the blade’s airfoil. If the local angle of attack exceeds the maximum lift angle, then a leading edge wake element is shed. Results from the method are presented for both attached and separated flow and compared with predictions obtained from the Blade Element Momentum theory. It is shown that the shedding of the leading edge wake can increased significantly the agreement between vortex lattice and Blade Element results in cases where there are significant regions of separated flow. This improvement concerns mostly forces acting normal to the rotor plane; tangential forces depend more strongly on the drag and neither method calculates a full representation of the drag.

Wind tunnel analysis of separated aerodynamics leading to different types of torsional flutter in bluff-bodies.

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in Proceedings of the 15th International Forum on Aeroelasticity and Structural Dynamics, IFASD 2011 (2011, June 29)

A new experimental apparatus has been developed at the wind tunnel laboratory of University of Liège with the objective to study the torsional flutter instability of a rectangular cylinder. Different configurations, based on the position of the pitching axis, the stiffness of the restoring springs and the static equilibrium angle of the model have been tested and analysed. Sub- and supercritical bifurcations behaviour has been identified. The dynamics of the system, as well as the aerodynamic phenomena taking place around the model have been measured. The frequency content of the pitching displacement and the longitudinal component of the velocity in the wake showed the occurrence of two types of aerodynamic instability: large separation from the leading edge and vortex shedding in the Strouhal sense. Furthermore, preliminary Particle Image Velocimetry visualisation confirmed the presence of complex aerodynamic phenomena over the upper surface and in the wake of the model. These measurements allow a better understanding of the sources of aerodynamic excitation leading to torsional flutter.

Interactive boundary layer calculation of separated flows around 2D airfoils

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in Proceedings of the 15th International Forum on Aeroelasticity and Structural Dynamics, IFASD 2011 (2011, June 27)

Stall flutter is an important undesirable phenomenon affecting lifting surfaces such as helicopter and wind turbine blades, but is particularly difficult to model. Existing methodologies are not universally applicable, and have significant limitations concerning applicability, time of computation and accuracy. This paper presents a method that combines interactive boundary layer with point vortex shedding. It is applied to a NACA0012 airfoil, at different angles of attack in the −40◦ < a < 40◦ range, considering a free flow velocity of 20[m/s]. Results are compared to those obtained at the Goldstein Laboratory of the University of Manchester.

Wind tunnel experiments on a flapping drone

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in Proceedings of the 15th International Forum on Aeroelasticity and Structural Dynamics, IFASD 2011 (2011, May 29)

The motivation for this paper is to obtain a better understanding of the unsteady aerodynamic phenomena involved in flapping wing flight. The work is mainly experimental but also makes use of numerical results obtained from a vortex lattice approach. The flapping of 3D wings produces vortical structures which differ from those produced in 2D flow. It is a phenomenon experienced by any 3D lifting surfaces that are undergoing time dependent motion. In avian flight, it is known that different types of kinematics produce different kinds of vortical wake structures, depending on the wing aspect ratio. Understanding the wake structures and their effects can lead to the optimisation of flapping flight through the manipulation of these unsteady flow features. The objective of this work is to investigate the kinematics of 3D wings along with the evolution of the resulting vortex and wake structures while varying the oscillation parameters. The parameters in question are reduced frequency, flapping and pitching kinematics and wing profiles.The experiments make use of an enhanced version of a dynamically scaled mechanical flapping wing, which is modelled on large migrating birds with simplified kinematics. Pure flapping and combined pitching and flapping are tested in the wind tunnel and simulated at 6.0m/s, 9.4m/s and 14.8m/s.The model is forced to oscillate at four different frequencies.

Complete bifurcation behaviour of aeroelastic systems with freeplay

in Proceedings of the 52nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference (2011, April)

Over the last couple of decades, a significant amount of research has been carried out on the aeroelastic behaviour of aeroelastic systems with freeplay. It has been established that such systems can undergo Limit Cycle Oscillations (LCO), both periodic and aperiodic. It has also been shown that several LCOs can occur at the same flight conditions, depending on initial conditions. A lot of the work has been applied to a pitch-plunge airfoil with a control surface and freeplay in the control rotation spring but, even for this simple model, the complete LCO behaviour has not been calculated. In this work, a combined approach using equivalent linearization, a shooting-based numerical continuation scheme and branch following is used to calculate the full bifurcation behaviour of such a system. It is shown that the primary LCO branches depend on the underlying linear systems but that there are two branching points from which secondary periodic solution branches emanate and wrap themselves around the primary branches. Up to 13 different LCOs can coexist at a single flight condition. The system undergoes Hopf, fold, flip and Neimark-Sacker bifurcations and the proposed solution method can identify and all of them.

Limit Cycle Oscillations of a Delta Wing in a Low Speed Wind Tunnel

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in Proceedings of ISMA2010 (2010, September 22)

Limit cycle oscillations involving Delta wings are an important area of research in modern aeroelasticity. Such phenomena can be the result of structural or aerodynamic nonlinearity. In this paper, a flexible half-Delta wing is tested in a low speed wind tunnel in order to investigate its dynamic response. Specifically, an investigation to determine the effects of a steady angle of attack on nonlinear Limit Cycle Oscillations (LCO) of a delta wing-plate model in low subsonic flow has been undertaken. It was found that, at several combinations of airspeed and angle of attack, the wing undergoes limit cycle oscillations. Two types of such oscillations are observed. One of them is low amplitude, low complexity limit cycle oscillations that occur at lower airspeeds; the other is high amplitude, high complexity limit cycle oscillations that occur a higher airspeeds and can appear abruptly. Some of the LCOs are the result of a subcritical Hopf bifurcation occurring at low steady angles of attack. At higher angles, a nontypical bifurcation was observed, whereby LCOs appear, grow with airspeed but then diminish and finally disappear as the airspeed is increased further.

Damping identification in a non-linear aeroelastic structure

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in Proceedings of ISMA2010 (2010, September 20)

An energy-based method is proposed to identify damping parameters from time histories of responses to sets of single-frequency harmonic excitation. The method is intended to be practically applicable to real structures and is able to identify the value of viscous damping, Coulomb friction and eventually other forms of non-linear damping models in aeroelastic systems. The inputs required are simply the accelerometer signals and the forces applied. It will be shown that if the system is undergoing Limit Cycle Oscillations, no external force is required for the identification process.

Bifurcation analysis of a wing undergoing stall flutter oscillations in a wind tunnel

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in Proceedings of ISMA2010 (2010, September 20)

Stall flutter is a Limit Cycle Oscillation (LCO) caused by the periodic separation of the flow around a wing immersed in a uniform fluid flow. The separation could be either partial or complete [1]. The work presented focuses on a wing undergoing stall flutter in the pitch degree of freedom. The phenomenon is analyzed from the structural and aerodynamic response perspectives. The objective of this study is to promote the understanding of stall flutter by characterizing the complete bifurcation behavior of the selected system. The wing section chosen for this study is NACA 0018 profile applied to a rectangular wing. The tests are carried out at different airspeeds and angles of attack. The measured aeroelastic responses are analyzed and the behavior of the dynamic system is characterized by fully describing its bifurcation. Structural accelerations as well as unsteady pressures around the mid-span point of the wing are measured and examined. Furthermore, flow field visualization by means of the Particle Image Velocimetry technique is used to demonstrate aspects of the unsteady flow field, such as the manifestation of separation and vortices.

Development and Evaluation of a VTOL Observation Platform

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in Proceedings of the 27th International Congress of the Aerospace Sciences (2010, September)

An aerodynamic study of the ULB-developed ducted rotor MAV using the results of full-scale wind tunnel tests allowed the determination of the platform’s positive speed envelope, power requirements and endurance characteristics for ISA sea level conditions. In this study, the power consumption appears to be majorly depending on the rotor rotational speed, while an increase in horizontal speed results in a duct operating more as a circular wing, also unveiling a power bucket as is the case with conventional helicopters. A positive influence of the ground proximity on the total thrust has been monitored.

Etude du comportement aéroélastique des structures soumises à des écoulements décrochés

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Conference (2010, March 02)

L'objectif du travail de recherche est le développement d'un modèle numérique simulant les phénomènes aéroélastiques affectant les corps non-profilés, caractérisés par des effets de décrochage/raccrochage du fluide.L'application de ce modèle aux tabliers de point est prévue.

A Novel Concept for Helicopter Rotor Drives

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in Proceedings of the 3rd European Conference for Aero-Space Sciences (2009, July)

The REDT (Rotor à Entraînement Direct par Turbine – Direct Turbine Driven Rotor) is a new concept for helicopter rotor drives developed by Sagita in Belgium. It works without any mechanical link between the engine on one side and the rotor drive on the other side. It uses a fuselage-mounted compressor that powers two contra-rotating rotor-mounted free turbines. These free turbines drive a pair of contrarotating rotors that are fitted with rigid hingeless main rotor blades. This novel rotor drive eliminates the need for either mechanical transmission or a tail rotor. The aim of the REDT concept is to lower the maintenance costs and the accident rate, as well as to extend the flight envelope towards much higher airspeeds compared to classical helicopters.

Analysis of dynamic response of a very flexible Delta wing model in a wind tunnel

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in Proceedings of the 2009 International Forum on Aeroelasticity and Structural Dynamics (2009, July)

Limit cycle oscillations involving Delta wings are an important area of research in modern aeroelasticity. Such phenomena can be the result of geometric nonlin- earity, aerodynamic nonlinearity or under-wing store nonlinearity. In this paper, a flexible half-Delta wing without stores is tested in a low speed wind tunnel in order to investigate its dynamic response. It is found that, at several combinations of airspeed and angle of attack, the wing undergoes high amplitude limit cycle oscillations. Three types of such oscillations are observed. Type 1 oscillations occur only at low angles of attack and are the result of a Hopf-type bifurcation. Type 2 limit cycle oscillations occur at intermedi- ate angles of attack and are the result of an atypical bifurcation. In other words, these oscillations appear as the airspeed is increased but disappear at even higher airspeeds. Type 3 oscillations occur at even higher angles of attack. A bispectrum analysis shows that type 3 limit cycle oscillations feature quadratic phase coupling. No such coupling was measured for type 2 oscillations, leading to the conclusion that the nonlinearity must be of higher order.

Experiments on a 3-D Flapping and Pitching Mechanical Model

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in Proceedings of the 2009 International Forum on Aeroelasticity and Structural Dynamics (2009, June)

The motivation for this paper is to focus on the better understanding of flapping wing flight, including flapping flight involving high pitching amplitudes. The investigation will be primarily experimental. The experiment will utilize a mechanical flapping and pitching wind tunnel model, which is modeled on large migrating birds. The 3-D rectangular wings are forced to flap and pitch sinusoidally in a low speed wind tunnel. The unsteady aerodynamic forces generated are measured using a three-component force balance. The main objective of this work is to investigate the effect that varying the oscillation parameters has on the aerodynamic forces acting on the 3-D wing undergoing flapping and pitching. The parameters in question are reduced frequency, flapping ampli- tude and pitching amplitude. Comparisons are made with aerodynamic force predictions obtained from unsteady vortex lattice calculations.

Using the Discrete Vortex Method to Simulate the Stall Flutter Phenomenon

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in Proceedings of the 2009 International Forum on Aeroelasticity and Structural Dynamics (2009, June)

The Discrete Vortex Method (DVM) has been recently used by several researchers in order to simulate the dynamic stall phenomenon on 2D airfoils oscillating at high angles of attack or undergoing stall flutter oscillations. The aim of the present work is to investigate whether such simulations are representative of the true physics governing such oscillation. A DVM method is described and implemented. A basic validation of the method is performed on the case of an impulsively started static airfoil at a low angle of attack. It is shown that the steady state lift value obtained from the DVM method agrees with experimental measurements. Subsequently, a DVM simulation is performed for the case of a NACA 0012 airfoil undergoing symmetric stall flutter oscillations. The simulation results are compared to experimental results. It is shown that, while the general shape of the simulated lift variation with time agrees with the experiment, there are significant mean value and phase differences between experiment and simulation. Several justifications are suggested and improvements to the DVM simulation proposed.

Wind tunnel investigation of the stall flutter oscillations of a bridge deck

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in Proceedings of the 8th National Congress on Theoretical and Applied Mechanics (2009, May)

Torsional Stall Flutter (TSF) is a one degree of freedom, self-excited phenomenon induced by the periodic separation/re-attachment of the flow around a body. The present paper proposes an experimental investigation of TSF on a bridge deck. A section model free to oscillate in torsion is tested in a low-speed wind tunnel. Static and dynamic tests are performed to enable a complete interpretation of the phenomena involved. For low airspeeds, the Karman vortex shedding excites the structure, resulting in small oscillations at the vortex shedding frequency. For high airspeeds, the deck undergoes Limit Cycle Oscillations (LCO) due to the TSF phenomenon. Nevertheless the Karman vortex shedding excitation is still present and appears in the measured forces signals. Peaks and troughs in the lift coefficient are detected for each pitching period. They are interpreted as the separation/re-attachment of the flow around the deck. The large amplitudes motion of the deck undergoing TSF reduces the amplitude of the aerodynamic lift coefficient.

Transonic Limit Cycle Oscillation Prediction From Simulated In-Flight Data

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in Proceedings of the 2008 International Conference on Noise and Vibration Engineering (2008, September)

This paper investigates the transonic Limit Cycle Oscillation (LCO) phenomenon from a computational/simulation point of view. A CFD/FE coupled solution is applied to a 3D model of a rectangular wing with tip store (referred to as the Goland wing) in transonic flow, in order to study the resulting LCO motions. The aim of the study is to develop a LCO prediction strategy from in-flight data; at this stage of the work the data will be obtained from simulations. The effect of varying the thickness of the Goland wing on the LCO onset flight condition is explored. The work then addresses the presence or absence of nonlinearity prior to the onset of LCOs. This is crucial as the presence of any nonlinearity for this type of wing can act as a telltale sign of impending LCOs. The clearance between the nonlinearity onset flight condition and the LCO onset flight condition is quantified for various values of the wing design parameters. It is shown that this clearance can be used in order to develop a useful ‘oncoming LCO’ warning criterion.

Data Clustering for the Identification of the Bifurcation Behaviour in Non-Linear Aeroelastic Systems using a Coupled Harmonic Balance/Genetic Algorithm Approach

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in Proceedings of the 2008 International Conference on Noise and Vibration Engineering (2008, September)

This paper describes an efficient method for calculating the bifurcation behaviour of an aeroelastic system using a Harmonic Balance expansion coupled with a Genetic Algorithm, combined with a clustering algorithm in order to determine all the solutions at every single flight condition. It will be shown how it is possible to obtain all the bifurcation branches in one step. Two clustering algorithms, K-Means and PAM, together with a number of cluster index techniques, such as Davies-Boulding, Calinski-Harabasz are investigated. The method is applied to an aeroelastic galloping problem as this phenomenon presents a number of co-existing limit cycles at a range of airspeeds.

Numerical Continuation of Aeroelastic Systems: Shooting vs Finite Difference Approach

in Limit Cycle Oscillation and Other Amplitude-Limited Self Excited Vibrations (2008, May)

Nonlinearity in aeroelastic systems has become an important area of research. It is widely recognised that nonlinearities due to the structure, control system or aerodynamic forces can affect aeroelastic systems, such as aircraft, and can cause oscillatory phenomena that cannot be predicted using a purely linear analysis. An important class of methods for predicting such phenomena from the equations of motion of a nonlinear aeroelastic system is numerical continuation. There are several techniques for the application of numerical continuation, each with its advantages and disadvantages. Here, the finite differences and shooting approaches are implemented and compared. Both approaches are applied to a simple pitchplunge nonlinear aeroelastic mathematical system.

Novel Instrumentation for the Measurement of the Unsteady Pressure Distribution around a Wing Undergoing Stall Flutter Oscillations

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in Autumn Conference of the Institute of Acoustics 2007: Advances in Noise and Vibration Engineering (2007, October)

In this work, a wind tunnel model of a wing undergoing stall flutter oscillations is studied. It is a rectangular wing with pitch and plunge degrees of freedom and low pitch stiffness. The objective of this study is to promote the understanding of the stall flutter phenomenon by measuring the unsteady pressure distribution around the wing as well as the wing displacement, during unforced motion in two degrees of freedom. Both steady and unsteady pressures must be measured with sufficient accuracy during two types of tests. In the static tests the wing is to be clamped in position and not allowed to move and the steady pressures are to be measured around the centre-span section at different angles of attack. Thus the stall angle of attack can be identified and the stall mechanism characterized. In the dynamic tests the wing will be allowed to move and the unsteady pressures will be measured and recorded during a number of cycles of the oscillation, at a number of free stream airspeeds.

Approximate Numerical Continuation for Aeroelastic Systems Undergoing Aperiodic Limit Cycle Oscillations

in Proceedings of the 2007 International Forum on Aeroelasticity and Structural Dynamics (2007, June)

This paper presents a modified numerical continuation approach for predicting the bifurcation behaviour of aeroelastic systems undergoing aperiodic limit cycles oscillations. Such oscillations can occur due to a number of nonlinear functions. Here, backlash nonlinearity in the aileron stiffness for a Generic Transport Aircraft is considered. It is shown that classical numerical continuation will fail due to the aperiodic nature of the limit cycles and the inability to perform period scaling and phase fixing. An alternative, approximate numerical continuation method is proposed, based on longer numerical integration sequences and a heuristic method for determining the period of the limit cycle oscillations. The approach is applied successfully to a simulated aeroelastic model of the Generic Transport Aircraft with backlash.

Linear and Non-Linear Transonic Flow Behaviour of the Goland+ wing

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in Proceedings of the 2007 International Forum on Aeroelasticity and Structural Dynamics (2007, June)

This paper is part of a study investigating the prediction of the aeroelastic behavior of aircraft subjected to transonic aerodynamic forces. The main objective of the work is the creation of Reduced Order Models from coupled Computational Fluid Dynamic and Finite Element calculations. The novelty of the approach lies in the identification of different types of Reduced Order Model in different flight regimes. Linear modal models are used in the Mach range range where the full CFD/CSD system is linear and nonlinear modal models in the transonic flight regime where the CFD/CSD system undergoes Limit Cycle Oscillations. Static solutions of the CFD/CSD system are used in order to determine the extent of the nonlinear Mach number range. The model treated in this work is a three-dimensional wing in a transonic flowfield.

Experimental Study of Stall-Induced LCOs of Free Vibrating Wing

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in Proceedings of the 2007 International Forum on Aeroelasticity and Structural Dynamics (2007, June)

Stall flutter is a limit cycle oscillation phenomenon occurring when all or part of the flow over a wing separates and re-attaches periodically. This paper reports on experimental investigations carried out on the subsonic stall flutter of a pitch plunge wing wind tunnel model. The object of the experiments was to investigate the aeroelastic effects of the nonlinearity introduced by boundary layer growth and flow separation and the ensuing stall-induced flutter, on a freely vibrating aeroelastic system. The motion of the wing and the pressure distribution around its surface were measured at various conditions. Three types of motion were observed, sometimes at the same wind tunnel airspeed; decay to equilibrium, symmetric limit cycle oscillations (LCO) and asymmetric LCOs. The aerodynamic mechanisms that give rise to these LCOs are described and analyzed. The effect of torsional wing stiffness on the stall flutter response of the wing is also investigated. Finally, repeating the experiments in a smaller wind tunnel showed that wind tunnel blockage has a significant effect on stall flutter.

Adaptive Aeroelastic Structures for Improved Aircraft Performance

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in Spencer Jr; Tomizuka, M.; Yun, C. B.; Chen, W. M.; Chen, R. W. (Eds.) World Forum on Smart Materials and Smart Structures Technology (2007, May)

There is a growing interest in the development of adaptive aeroelastic structures to allow aeroelastic deflections to be used in a beneficial manner. Part of the 3AS research programme was devoted towards investigating the use of changes in the internal aerospace structure in order to control the static aeroelastic behaviour. Such an approach is desirable and arguably advantageous compared to other possible concepts.

Improved Implementation of the Harmonic Balance Method

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in Proceedings of the 48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (2007, April)

Harmonic Balance (HB) methods have been applied to non-linear aeroelastic problems since the 1980s. As the computational power available to researchers has increased, so has the order of calculated HB solutions. However, the computational cost of a HB solution increases with the square of the order. Additionally, the traditional Newton-Raphson, Broyden, Toeplitz Jacobian and other techniques used for the solution of the non-linear algebraic problem at the heart of the HB methodology rely on a good initial guess for the unknown coefficients. If there are many such coefficients the probability that a good guess will be available is very low and the HB scheme may well fail. In this paper a search procedure using Genetic Algorithms (GA) is introduced to evaluate the coefficients of a harmonic balance solution. It is shown that the GA can provide high quality initial guesses for the HB coefficients. The method is applied to an aeroelastic galloping-type problem.

Flight-Regime Dependent Reduced Order Models of CFD/FE aeroelastic systems in transonic flow

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in Proceedings of the 48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (2007, April)

This paper is part of a study investigating the prediction of the aeroelastic behavior of aircraft subjected to transonic aerodynamic forces. The main objective of the work is the creation of Reduced Order Models from coupled Computational Fluid Dynamic and Finite Element calculations. The novelty of the approach lies in the identification of different types of Reduced Order Model in different flight regimes. Linear modal models are used in the Mach range range where the full CFD/FE system is linear and nonlinear modal models in the transonic flight regime where the CFD/FE system undergoes Limit Cycle Oscillations. Static solutions of the CFD/FE system are used in order to determine the extent of the nonlinear Mach number range. The model treated in this work is a three-dimensional wing in a transonic flowfield.

Drag Minimisation Using Adaptive Aeroelastic Structures

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in Proceedings of the 48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (2007, April)

This paper describes the latest developments in a research program investigating the development of “adaptive internal structures” to enable adaptive aeroelastic control of aerospace structures. Through controlled changes of the second moment of area, orientation or position of the spars, it is possible to control the bending and torsional stiffness characteristics of aircraft wings or tail surfaces. The aeroelastic behaviour can then be controlled as desired. A number of different adaptive internal structure concepts (rotating, moving and split spars) are compared here using a simple rectangular wing structure in order to determine which are the most effective for achieving minimum drag at different points in a representative flight envelope. A genetic algorithm approach is employed to determine the optimal spar orientation for rotating spars concept. It is shown that it is feasible to adjust the structure and trim characteristics of such wing structures in order to achieve minimum drag at all conditions.

Nonlinearity Characterization for Nonlinear Dynamic System Identification Using an Expert Approach

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in Proceedings of the 2006 International Conference on Noise and Vibration Engineering (2006, September)

The identification of nonlinear dynamic systems can be rendered significantly more parsimonious if the nonlinearity present in the system is known. While there are many successful non-parametric nonlinear system identification methods, the resulting models do not describe the nonlinearity in physical terms and are difficult to obtain due to the large number of candidate terms that must be examined. In this paper an expert approach towards the characterization of nonlinearities in a dynamic system is presented. The methodology is based on simulations of dynamic systems with a variety of commonly occurring nonlinear functions. The responses of such systems to various types of excitation are analysed and rules are developed as to what nonlinearity is likely to be present in a system given the dynamic characteristics of measured responses.

Identification of Structural Free-play Non-linearities using the Non-Linear Resonant Decay Method

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in Proceedings of the 2006 International Conference on Noise and Vibration Engineering (2006, September)

Structural non-linearities are becoming of increasing importance in determining the performance of a range of vibrating mechanical structures. As a consequence, the identification of systems with non-linearities is starting to become a necessary part of vibration testing procedures. NL-RDM (Non-Linear Resonant Decay Method) is an approach for the identification of non-linear multi-degree of freedom systems in modal space on a mode by mode basis, using an appropriated sine excitation to isolate modes or groups of modes. However, the application of NL-RDM to a multi-degree of freedom system with a discontinuous free-play non-linearity has not been attempted yet, except to treat it using high order polynomial terms. The difficulty of using NL-RDM when seeking a discontinuous free-play model lies in the choice of the so-called underlying linear model to be used in uncoupling the linear equations. In this paper, a simulated two degree of freedom lumped parameter system with a free-play non-linearity

An experimental investigation of the subsonic stall flutter

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in Proceedings of the 2006 International Conference on Noise and Vibration Engineering (2006, September)

This paper reports on experimental investigations of the subsonic stall flutter of a wing in a wind tunnel. Stall flutter (also know as Dynamic Stall) is a LCO phenomenon occurring when all or part of the flow over a wing separates and re-attaches at least once during a full oscillation period. In order to investigate the aeroelastic effects of the nonlinearity introduced by the boundary layer growth and flow separation and the ensuing stall-induced LCO, a wind tunnel model of a wing undergoing stall flutter is designed, built and tested. The model is a rectangular wing with constant cross-section free to move in the pitch and plunge directions, restrained by torsional and linear springs respectively. The motion of the wing is measured using laser displacement probes. Two types of stall flutter are observed and measured: 1. Non-symmetric stall flutter, where the flow separates over one side of the wing only, and 2. Symmetric stall flutter, where the flow separates over both sides of the wing (deep stall). The bifurcation behaviour of the wing is very complex and both types of LCO can be observed during a single response history.

An Expert System for the Identification of Nonlinear Dynamical Systems

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in Huang, De-Shuang; Li, Kang; Irwin, George William (Eds.) Intelligent Computing (2006, August)

This paper describes an Expert System that can detect and quantify the nonlinearity present in a given dynamical system and, subsequently, determine and apply the most suitable nonlinear system identification method. The internal workings, algorithms and decision making processes of the Expert System are discussed. For demonstration purposes the Expert System is applied to a nonlinear experimental test-rig. The results show that the Expert System is an automatic tool that will detect nonlinearity, choose the best class of model for the system under investigation and perform optimal parameter estimation, so that the resulting identified models are parsimonious and accurate.

Development of An Expert System for the Identification of Nonlinear Vibrating Systems

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in Brennan, M. J.; Liguore, S.; Mace, B. R.; Muggleton, J. M.; Murphy, K. D.; Petersson, B. A. T.; Rizzi, S. A.; Shen, R. (Eds.) Proceedings of the IX International Conference on Recent Advances in Structural Dynamics (2006, July)

The aim of the present work is to attempt to create a logical framework to be used for the identification of nonlinear systems. It is assumed that no single identification method is general enough to work with a significant range of systems. Therefore, this framework is based on the development of an Expert System that will detect and quantify the nonlinearity present in a given dynamical system and, subsequently, determine and apply the most suitable nonlinear system identification method or methods.

Aeroelastic System Identification using Transonic CFD data for a 3D Wing

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in Brennan, M. J.; Liguore, S.; Mace, B. R.; Muggleton, J. M.; Murphy, K. D.; Petersson, B. A. T.; Rizzi, S. A.; Shen, R. (Eds.) Proceedings of the IX International Conference on Recent Advances in Structural Dynamics (2006, July)

This paper is part of a study investigating the prediction of aeroelastic behaviour subjected to non-linear aerodynamic forces. Of interest here is whether the sub-critical vibration behaviour of the aeroelastic model gives any information about the onset of the LCO. It would be useful to be able to use system identification methods to estimate aeroelastic models that characterise the LCO. Such a methodology would be very useful, not only for analysis with coupled CFD/FE models, but also during flight flutter testing. In this paper, the responses to initial inputs on the Goland Wing [9] CFD/FE model at different flight speeds are analysed to determine the extent of the non-linearity below the critical onset of LCO. Analysis is also performed using a linear identification model.

Identification of Non-Linear Dynamic Systems using an Expert Approach

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in Proceedings of the 47th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (2006, May)

An Expert System approach for the identification of non-linear systems is presented. The Expert System is an attempt to bring order into the non-linear system identification process. The final objective is to deliver a parsimonious mathematical model of the dy- namical system under investigation. It works by defining the key stages of the procedure and iterating between them if necessary. It takes advantage of numerous methodologies to accomplish the tasks in each of the stages and uses the results from many of them. The Expert System applies the excitation forces appropriate to each method and analyses the responses. Each stage ends with a set of recommendations that can be used to begin the next stage. The Expert System is applied to a simple non-linear dynamic system. It is shown that the Expert System procedure can automatically detect, locate and quantify the non- linearity using its array of techniques. Furthermore, it can choose a suitable model struc- ture, select appropriate terms and estimate the model parameters using an extensive set of rules. The resulting identified model is validated and shown to be an accurate represen- tation of the experimental system.

Application of Higher-Order Harmonic Balance to Non-Linear Aeroelastic Systems

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in Proceedings of the 47th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (2006, May)

The influence of non-linearities on modern aircrafts is becoming of increasing impor- tance. The ability to accurately characterise LCOs and to predict at which speed they occur is very important. Higher Order Harmonic Balance (HOHB) methods have attracted some interest from the aeroelastic community over the last two decades. Such methods carry the promise of high quality stability prediction and Limit Cycle Oscillation (LCO) amplitude and frequency prediction for non-linear aeroelastic systems. In this paper, a Higher Order Harmonic Balance scheme is devised to extend the effectiveness of the method to systems undergoing secondary Hopf bifurcations. It is shown that the proposed harmonic shifting technique can allow the HOHB method to accurately estimate both branches of limit cycles occurring after the second bifurcation.

Flutter Clearance of a Non-linear aircraft

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in Proceedings of the 2005 International Forum on Aeroelasticity and Structural Dynamics (2005, June)

Flight flutter testing is always carried out under the assumption that aircraft are linear. Recently, this assumption has started to come under question, especially as far as military aircraft are concerned. This paper deals with possible methodologies for flight flutter testing of aircraft that are no longer assumed linear. Simulated flight testing is performed for a simple non-linear aeroelastic system with cubic stiffness. The flutter speeds predicted using some of the classical linear flutter prediction methods as well as a non-linear method are compared. It is shown that, for non-linear system undergoing Hopf Bifurcations, classical linear flutter prediction can predict the flutter envelope with reasonable accuracy. However, fully non-linear system identification and stability analysis can not only predict the flutter point but also determine whether it is a linear or non-linear flutter point (i.e. whether divergent or Limit Cycle Oscillations will ensue). Additionally, the non-linear method can predict the amplitudes of LCOs that will occur post-critically. The application of the nonlinear method was successful for noise free data, but the problem of noise corruption still needs further investigation.

Stability and LCO Amplitude Prediction for Aeroelastic Systems with Aerodynamic and Structural Nonlinearities Using Numerical Continuation

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in RTO-MP-AVT-123 Flow-Induced Unsteady Loads and the Impact on Military Applications (2005, April)

This paper deals with the prediction of stability boundaries and Limit Cycle Oscillation amplitudes for aeroelastic systems with nonlinear unsteady aerodynamic loads and/or nonlinearity in the structure. The Numerical Continuation method is used to accurately predict bifurcation conditions and LCO amplitudes for aeroelastic systems with various types of nonlinearity without the need for extensive CFD calculations. It is shown that it is possible to completely characterise the stability of systems undergoing subcritical and supercritical bifurcations. The method is applied to a pitch-plunge airfoil subjected to transonic aerodynamics and freeplay structural nonlinearity. The results from this analysis are compared to those obtained from full numerical simulation to ensure their accuracy.

Non-Linear Identification Using a Genetic Algorithm Approach for Model Selection

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in Proceedings of the 23rd International Modal Analysis Conference (2005, January)

The Non-Linear Resonant Decay Method is an approach for the identification of non-linear systems with large numbers of degrees of freedom. The identified non-linear model is expressed in linear modal space and comprises the modal model of the underlying linear system with additional terms representing the non-linear behaviour. Potentially, a large number of these non-linear terms will exist but not all of them will be significant. The problem of deciding which and how many terms are required for an accurate identification has previously been addressed using the Forward Selection and Backward Elimination techniques. In this paper, a Genetic Algorithm optimisation is proposed as an alternative to those methods. A simulated 5-DOF lumped parameter non-linear system is used to demonstrate the proposed optimisation. The use of separate data sets for the identification and validation of the modal model is also investigated. It is found that the Genetic Algorithm approach yields significantly better results than the Backward Elimination and Forward Selection algorithms in many cases.

Validating Blade Vibration Amplitudes from Blade Tip-Timing Data Analysis

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in Vibrations in Rotating Machinery (2004, September)

Blade Tip-Timing (BTT) is a method for the measurement of blade vibration in rotating bladed assemblies such as those found in turbomachinery. BTT data needs to be analysed by specially formulated methods, to yield the vibration frequencies and amplitudes. Although such methods have been shown to recover frequencies successfully, they have not been validated experimentally for amplitude recovery. In the present paper an improved experimental procedure for amplitude measurement is described. The paper presents a comparison between the experimental observation of the blade vibration amplitudes and the answers obtained from the BTT data analysis methods, providing experimental evidence of the ability of these methods to extract the correct blade vibration amplitudes from BTT data.

On the solution of the aeroelastic galloping problem

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in Proceedings of the 2004 International Conference on Noise and Vibration Engineering (2004, September)

A global stability analysis of the transverse galloping of a square section beam in a normal steady ow was performed. The analysis was applied to a mathematical model using experimentally measured stationary aerodynamic forces. The system was modelled as an ordinary differential equation with small non-linearity in the velocity term. Three methods are used for the stability analysis: 1. a harmonic balance approach, 2. normal form theory, 3. cell mapping. The resulting stability predictions were compared with each other and with results obtained from numerical integration. It is shown that the hysteretic stability of the non-linear aeroelastic oscillator was captured by all the methods. Additionally, the methods had a varying degree of success in predicting the amplitude of limit cycle oscillations undergone by the aeroelastic oscillator.

Stability and Limit Cycle Oscillation Amplitude Prediction for Multi-DOF Aeroelastic Systems with Piecewise Linear Non-Linearities

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in Proceedings of the 2004 International Conference on Noise and Vibration Engineering (2004, September)

Discontinuous non-linearities such as freeplay and bilinear stiffness are often encountered in aeroelastic systems, sometimes as a result of wear and tear. It is important to predict the effect of such non-linearities on the dynamic behaviour of a system, so that adequate safety guidelines can be drafted. As a consequence, the prediction of the bifurcation behaviour of a system featuring a discontinuous nonlinearity is crucial. Additionally, the post-bifurcation behaviour of the system is also of interest since it may consist of relatively harmless Limit Cycle Oscillations (LCO) of low amplitude or of unexpected catastrophic high amplitude LCOs. In this paper the bifurcation and post-bifurcation behaviour of a simulated Multi-DOF aeroelastic system with bilinear and freeplay nonlinearities are investigated using the Harmonic Balance method and a novel method for the prediction of the bifurcation conditions and LCO amplitudes. The method is based on the fact that the nonlinearities investigated are piecewise linear. The ratios of the real parts of the system eigenvalues in the various ranges of the bilinear spring are used in order to infer LCO amplitude information. By means of a demonstration on a simulated aeroelastic system with piece-wise linear stiffness, it is shown that the proposed approach is successful in yielding the full bifurcation and post-bifurcation behaviour of the system. Comparison of the amplitude predictions obtained from the Harmonic Balance technique and the Piecewise Linearisation proposed approach show that the latter are more consistent and closer to the true amplitudes throughout the airspeed range. The bifurcation analysis is extended to the special case where the inner stiffness of the bilinear spring is equal to zero, i.e. freeplay stiffness. It is shown that the Piecewise Linear analysis fails to capture the bifurcation behaviour for this case, while the Harmonic Balance method still produces some accurate predictions.

Nonlinear System Identification using Interpolated Short Time Fourier Transform

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in Akay, A.; Arnas, O.; Cooper, J. E.; Egrigan, N.; Franco, F.; Stronge, W. (Eds.) Proceedings of 7th Biennial ASME Conference on Engineering Systems Design and Analysis (2004, July)

For the purpose of constructing the backbone of nonlinear systems, the Interpolated Short Time Fourier Transform (ISTFT) is proposed as a means to improve the estimation accuracy of the instantaneous amplitudes and frequencies of response signals. It is shown that the backbone curves estimated by the ISTFT agree with theoretical backbone curves very well. Additionally, the restoring force can be reconstructed to specify the type of nonlinear stiffness. A curve-fitting technique is introduced to estimate the parameters of nonlinear systems on the basis of theoretical backbone curves. It is shown that a number of typical nonlinear stiffness functions such as cubic, bilinear and pre-compressed springs can be identified accurately using this new method.

Subspace Monitoring of Multivariate Dynamic Systems

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in Akay, A.; Arnas, O.; Cooper, J. E.; Egrigan, N.; Franco, F.; Stronge, W. (Eds.) Proceedings of 7th Biennial ASME Conference on Engineering Systems Design and Analysis (2004, July)

In this article, the monitoring of continuous processes using linear dynamic models is presented. It is outlined that dynamic extensions to conventional multivariate statistical process control (MSPC) models may lead to the inclusion of large numbers of variables in the condition monitor. To prevent this, a new dynamic monitoring scheme, based on subspace identification, is introduced, which can (i) determine a set of state variable for describing process dynamics and (ii) produce a reduced set of variables to monitor process performance. This is demonstrated by an application study to a realistic simulation of a chemical process.

Multiple Frequency Analysis Methods for Blade Tip-Timing Data Analysis

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in Akay, A.; Arnas, O.; Cooper, J. E.; Egrican, N.; Franco, F.; Stronge, W. (Eds.) Proceedings of 7th Biennial ASME Conference on Engineering Systems Design and Analysis (2004, July)

Blade Tip-Timing (BTT) is a method for the measurement of blade vibration in rotating bladed assemblies such as those found in turbomachinery. The system aims to replace strain gauge technology. However all current BTT analysis methods fail to recover the correct frequencies when two blade modes are excited simultaneously by a synchronous vibration. In this paper, five new methods that can recover simultaneous frequencies from BTT data are presented. The methods are based on the auto-regressive approach. The approaches make use of data either from a single blade and single revolution or from multiple revolutions. Furthermore, some of the methods are designed to allow for the presence of measurement errors. The techniques are validated on three test cases in which simulated data was used. It is shown that most of the methods produce accurate estimates for the vibration frequency, even in the presence of significant noise levels, provided that a suitable amount of the response waveform is measured. The most consistent estimates are obtained from the methods that make use of data from multiple revolutions.

Identification and model Updating of a non-stationary vibrating system

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in Akay, A.; Arnas, O.; Cooper, J. E.; Egrigan, N.; Franco, F.; Stronge, W. (Eds.) Proceedings of 7th Biennial ASME Conference on Engineering Systems Design and Analysis (2004, July)

Non-stationary systems, which are commonly encountered in many fields of science, are characterized by time-varying features and require time-frequency methods for their analysis. This study considers the problem of identification and model updating of a non-stationary vibrating system. In particular, a number of identification methods and a model updating procedure are evaluated and compared through application to a time-varying “bridge-like” laboratory structure. The identification approaches include Frequency Response Function based parameter estimation techniques, Subspace Identification and Functional Series modelling. All methods are applied to both output-only and input-out-put data. Model updating is based upon a theoretical model of the structure obtained using a Rayleigh-Ritz methodology, which is updated to account for time-dependence and nonlinearity via the identification results. Interesting comparisons, among both identification and model updating results, are performed. The results of the study demonstrate high modelling accuracy, illustrating the effectiveness of model updating techniques in non-stationary vibration modelling.

Stability and Limit Cycle Oscillation Amplitude Prediction for Simple Nonlinear Aeroelastic Systems

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in Proceedings of the 45th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference (2004, April)

The prediction of the bifurcation and post-bifurcation behaviour of nonlinear aeroelastic systems is becoming a major area of research in the aeroelastic community due to the need for improved transonic aeroelastic prediction, the use of non-linear control systems, and new construction techniques that reduce the amount of inherent damping. In this paper, a novel application of the Centre Manifold Theorem is used to accurately predict bifurcation conditions and Limit Cycle Oscillation amplitudes for simple aeroelastic systems with various types of nonlinearity. A simple aeroelastic system with hardening cubic stiffness nonlinearity is considered and is demonstrated to display a wide variety of bifurcation phenomena. These make it dif cult for some of the standard existing methods, such as Normal Form, Cell Mapping and Tangential Linearisation, to quantify the Limit Cycle Oscillation amplitudes through the entire speed range of the system. Then, the proposed approach is introduced and applied to the same system. It is shown that it can accurately predict the limit cycle amplitudes of the system undergoing all types of bifurcation. Finally, the new technique is applied to the same system but with softening cubic stiffness nonlinearity. It is shown that the method can accurately predict both the static and dynamic divergence boundaries and that it can be used to draw a worst-case stability boundary, inside which the solution is always stable.

Limit Cycle Prediction For Subsonic Aeroelastic Systems Using Nonlinear System Identification

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in Ferman, M. A.; Petersson, B. A. T.; Rizzi, S. A.; Wentz, K. (Eds.) Proceedings of the VIII International Conference on Recent Advances in Structural Dynamics (2003, July)

The prediction of aeroelastic instabilities caused by nonlinear unsteady aerodynamic forces acting on aircraft has recently become an important area of research. Emphasis is placed on the capability to predict the occurrence of Limit Cycle Oscillations (LCOs) at both the design and prototype testing stages. In this paper, the prediction of LCOs is attempted for a simulated aeroelastic system subjected to nonlinear subsonic unsteady aerodynamic forces, using system identification. Response data from the simulated system are curve-fitted by means of a series of polynomial basis functions. This approach yields very accurate identified models of the actual system at individual flight conditions. These identified models are extrapolated to a global aeroelastic identified model. Using this model, the flight conditions at which LCOs occur is accurately predicted but the amplitude of the oscillations is underestimated.

Nonlinear Aeroelastic Research at the Dynamics and Aeroelasticity Research Group

Conference (2003, June)

This work is a presentation of the recent research in nonlinearity carried out at the Dynamics and Aeroelasticity Research Group (DARG) of the University of Manchester, School of Engineering

On the use of control surface excitation in flutter testing

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in Proceedings of the 2003 International Forum on Aeroelasticity and Structural Dynamics (2003, June)

Flutter testing is used to demonstrate that the aircraft flight envelope is flutter free. Response measurements from deliberate excitation of the structure are used to identify and track frequency and damping values against velocity. In this paper, the common approach of using a flight control surface to provide the excitation is examined us ing a mathematical model of a wing and control surface whose rotation is restrained by a simple actuator. In particular, it is shown that it is essential to use the demand signal to the actuator as a reference signal for data processing. Use of the actuator force (or strain) or control angle (or actuator displacement) as a reference signal is bad practice because these signals contain response information. It may also be dangerous in that the onset of flutter may not be seen in the test results. Control sur face flutter is of particular concern.

Improvements of some analysis methods for flight flutter test data using unsteady aerodynamic considerations

in Proceedings of the 2003 International Forum on Aeroelasticity and Structural Dynamics (2003, June)

To date the analysis of data from ight utter tests is routinely carried out under the assumption of steady or, at best, quasi-steady aerodynamics. This approach is favored mainly because of its simplicity. In this paper it is shown that, even for some very simple simulated aeroelastic systems, utter predictions that ignore the unsteadiness of aerodynamic loads can be signi cantly inaccurate. Then, two simple system identi cation techniques are proposed that can model both the steady and unsteady aerodynamic loads acting on an aeroelastic system in incompressible ow. These methods can be used in order to obtain a complete description of an aeroelastic system subject to unsteady aerodynamic forces, including its dependence on ight conditions. Finally, the new methods are applied to a simple simulated unsteady aeroelastic model. The resulting utter predictions are compared to those obtained from two standard ight utter test data analysis approaches and are shown to be signi cantly more accurate.

Development of a Multiple Modes Simulator of Rotating Bladed Assemblies for Blade Tip-Timing Data Analysis

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in Van Hal, B.; Sas, P. (Eds.) Proceedings of the 2002 International Conference on Noise and Vibration Engineering (2002, September)

Blade Tip Timing (BTT) is a promising technique for the measurement of blade-tip vibrations in bladed assemblies. After an engine test is performed, the vibration’s Engine Order (EO) and amplitude information is recovered from the collected BTT data using suitable analysis methods. The validation of such analysis methods can be carried out more conveniently on simulated BTT data. In this paper, a BTT data simulator is developed to produce controlled BTT data for bladed assemblies in which multiple modes of vibration are excited. The mathematical model used allows for multiple blade modes and multiple assembly modes to be simulated. Additionally, single or multiple, integral and non-integral EO excitations can be selected and the rotational speed of the assembly can be varied linearly. Results from several test cases are shown, demonstrating that the simulator successfully emulates assembly and blade mode combinations. A number of BTT analysis methods were able to recover with accuracy the correct EOs from the simulated BTT data.

Identification of the nonlinear function in a vibrating nonlinear beam

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in Lees, A. W.; Prells, U.; Norton, J. (Eds.) Proceedings of the Third International Conference on Identification in Engineering Systems (2002, April)

The Constant Level Identification (CLI) method for nonlinear dynamical systems has been shown to identify accurately a wide variety of systems with discrete degrees of freedom. In this paper, the method is applied to a continuous nonlinear system, namely a simulated model of a vibrating beam attached to a nonlinear spring. A Rayleigh-Ritz series representation of the beam is assumed in order to extract the desired number of modal responses. Then, a combination of modal and physical coordinates are used in the identification process. It is shown that the CLI method can accurately capture the shape of the nonlinearity in the system.

Experimental Validation of the Constant Level method for Identification of Nonlinear Multi Degree of Freedom Systems

in Starr, A. G.; Rao, R. B. K. N. (Eds.) Proceedings of the 14th International Congress on Condition Monitoring and Diagnostic Engineering Management (2001, September)

System identification methods for nonlinear dynamical systems could find uses in many applications such as condition monitoring, finite element model validation and stability determination. The effectiveness of existing nonlinear system identification techniques is limited by various factors such as the complexity of the system under investigation and the type of nonlinearities present. In this work, the Constant Level Identification approach, which can identify multi degree-of-freedom systems featuring any type of nonlinear function, including discontinuous functions, is validated experimentally. The method is shown to yield accurate identifications of an experimental dynamical system featuring two types of stiffness nonlinearity. The full equations of motion are accurately extracted, even in the presence of a discontinuous nonlinearity.

Online Flight Flutter Testing

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in Link, M. (Ed.) Proceedings of the International Conference on Structural System Identification (2001, September)

A number of different identification methods are considered as candidate approaches to undertake online flight flutter testing. Rather than flying the test aircraft to a new flight condition and then performing a data analysis at a constant flight condition, it is shown how it is feasible to track changes in frequencies and damping ratios as the flight speed is increased. It is also possible to provide an updated estimate of the predicted flutter speed, giving the flight engineers added confidence in the margin of safety. Both frequency and time domain methods are evaluated on a simulated 3 DOF aeroelastic system. These preliminary results show that such an approach is feasible although further work is required to determine the best testing and analysis approach.

Aeroelastic and Structural Design of an All-Moveable Aircraft Fin

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in Hönlinger, Heinz (Ed.) Proceedings of the CEAS Conference on Multidisciplinary Aircraft Design and Optimization (2001, June)

The Lagrange Multi-Disciplinary Design Optimisation (MDO) code is used for the preliminary study into design aspects of a generic all-movable aircraft fin. It is shown how the control effectiveness at high speeds can be increased through variation of the attachment position and stiffness, whilst still meeting strength and aeroelastic stability constraints.

A time-frequency technique for the stability analysis of impulse responses from nonlinear aeroelastic systems

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in Proceedings of the 2001 International Forum on Aeroelasticity and Structural Dynamics (2001, June)

A time-frequency method is proposed for the analysis of response time histories from nonlinear aeroelastic systems. The approach is based on a time-varying curve-fit of the Short Time Fourier Transform of the impulse response. It is shown that the method can be used in order to obtain a clear picture of the sub-critical stability of a number of aeroelastic systems with a variety of structural and aerodynamic nonlinearities. Additionally, frequency and amplitude information can be obtained for both the linear and nonlinear signatures of the response signals in the sub and post-critical regions. Finally, it is shown that, given certain types of nonlinear functions, subcritical damping trends can be extrapolated to predict bifurcation airspeeds.

Limit Cycle Oscillation Control

Poster (2000, December)

Over the last fifteen years the Aerospace industry has started to show a continuously increasing interest in the investigation of the vibrations of aircraft subjected to nonlinear structural and aerodynamic forces. Here, a control scheme designed to alleviate or even stop some types of nonlinear aircraft vibrations is presented.

Characterization of nonlinear aeroservoelastic behaviour

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in RTO-MP-036 Structural Aspects of Flexible Aircraft Control (1999, October)

The characterisation of the behaviour of nonlinear aeroelastic systems has become a very important re- search topic. Nevertheless, most of the work carried out to date concerns the development of unsteady CFD solutions in the transonic region. Important though this work is, there is also a need for research which aims at understanding the behaviour of non- linear systems, particularly the occurance of Limit Cycle Oscillations (LCOs). The purpose of this pa- per is to study the stability of a simple aeroservoelas- tic system with nonlinearities in the control system. The work considers both structural and control law nonlinearities and assesses the stability of the system response by use of bifurcation diagrams. It is shown that simple feedback systems designed to increase the stability of the linearised system also stabilise the nonlinear system, although their effects can be less pronounced. Additionally, a nonlinear control law designed to limit the control surface pitch response was found to increase the flutter speed considerably by forcing the system to undergo limit cycle oscilla- tions instead of fluttering. Finally, friction was found to affect the damping of the system but not its sta- bility, as long as the amplitude of the frictional force is low enough not to cause stoppages in the motion.

A Comparison of Blade Tip-Timing Data Analysis Methods

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in Leung, A. Y. T.; Wright, J. R.; Sandoz, D. J.; Starr, A. G. (Eds.) Proceedings of the First International Conference on the Integration of Dynamics, Monitoring and Control (1999, September)

The determination of the vibration characteristics of rotating engine blades is very important for fatigue failure considerations. One of the most promising techniques for measuring blade vibrations is blade tip-timing. Three vibration analysis methods were used on data obtained from a simple mathematical blade tip-timing simulator. The results from the methods were compared statistically. One of the methods, the Global Autoregressive Instrumental Variables approach produced satisfactory results at realistic noise levels. However, all of the techniques produced biased results under certain circumstances.

Identification of Nonlinear Aeroelastic Systems

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in Proceedings of the 1998 International Conference on Noise and Vibration Engineering (1998, September)

The successful identification of nonlinear systems could find uses in many aerospace applications including the validation of finite element models and tracking the stability of aircraft during flight flutter testing. The effectiveness of existing nonlinear system identification techniques is limited by various factors such as the complexity of the system under investigation and the type of nonlinearities present. In this work, a new approach is introduced that is able to identify multi degree-of-freedom systems featuring any type of nonlinear function, including discontinuous functions. The method is shown to yield accurate identifications of three different simulated aeroelastic systems containing a wide range of structural nonlinearities.

Analysis and Prediction of Instability Behaviour in Non-Linear Aeroelastic Systems

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in Tagungsband Aeroelastik-Tagung der DGLR (1998, June)

The effect of non-linearities, be they structural, aerodynamic or control, on the aeroelastic behaviour of aircraft is becoming of increasing importance in the Aerospace industry. As well as the prediction of °utter instabilities, there is a need for the prediction and modelling of the onset of Limit Cycle Oscillations (LCO). A thorough understanding of the possible types of aeroelastic phenomena that can occur due to non-linearities in the system is required before attempting to predict the behaviour. In this paper, the onset and suppression of LCOs is investigated by means of an examination of the stability and energy content of limit cycles.

Prediction of maximum loads due to turbulent gusts using nonlinear system identification

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in Proceedings of the 1997 International Forum on Aeroelasticity and Structural Dynamics (1997, June)

Two nonlinear system identification methods, the NARMAX method and a proposed method based on the Restoring Force Surface method, were used in conjunction with two gust load prediction methods, the Matched Filter Based 1-dimensional search and the Deterministic Spectral procedure - or Matched Filter Based multi-dimensional search - in order to demonstrate the feasibility of obtaining maximised gust loads using identified models of nonlinear aeroelastic systems. The procedure was tried on two different mathematical systems, a simple, one-degree-of-freedom system with cubic nonlinearity and a model of a wing with control surface and discontinuous nonlinearities. The maximised gust loads predicted from the identified systems were very close to the ones obtained for the actual systems. Furthermore, it was demonstrated that performing the identification procedure at low excitation levels does not adversely affect the prediction of the maximised loads and critical gust profiles.