Trailing-edge Control Surface Research Articles

Roll Maneuverability of Transonic High-Aspect-Ratio-Wing Aircraft with Flared Folding Wingtips

This paper investigates the roll maneuvers of a very flexible, transonic high-aspect-ratio-wing aircraft with flared folding wingtips, representing a potential future commercial transport configuration. The computational study leverages a fully coupled nonlinear aeroelastic-flight dynamics framework for simulating roll maneuvers commanded by deflecting control surfaces inboard or outboard of the flared folding wingtip hinge. Releasing the flared folding wingtips during flight makes the aircraft roll faster only when deflecting outboard-of-hinge leading-edge control surfaces. However, the achieved roll angles and rates are much lower than the ones obtained by deflecting inboard-of-hinge trailing-edge control surfaces, whose roll control effectiveness is not impacted or even degrades when releasing the flared folding wingtips. These trends are explained by analyzing the role of sweep-induced washout effects in the very flexible, transonic high-aspect-ratio wing, revealing physical mechanisms not identified in previous studies on straight or swept wings with moderate flexibility. Higher wing out-of-plane bending stiffness increases the effectiveness of flared folding wingtips in enhancing roll maneuvers at low dynamic pressure but plays a slight role at high dynamic pressure. This study highlights the need to consider the impact of washout effects for different flight conditions and levels of wing flexibility when assessing the ability of flared folding wingtips to enhance the roll maneuverability of transonic high-aspect-ratio-wing aircraft.

Exploring the Impact of Rapidly Actuated Control Surfaces on Drone Aerodynamics

This study investigates the use of rapidly actuated leading-edge and trailing-edge control surfaces to improve the control authority of small fixed-wing drones. Static and dynamic characteristics were investigated and presented in two separate papers. In this paper, the focus is on the dynamic effects observed from rapidly actuated 30% chord leading- or trailing-edge hinged control surfaces affixed to two flat-plate airfoils. Forces were resolved from surface pressure measurements and are augmented by PIV measurements, smoke flow visualization and analyses. The static study revealed that trailing-edge control surfaces exhibited higher effectiveness in producing time-averaged CL compared to leading-edge control surfaces. However, leading-edge control surfaces exhibit significantly less fluctuation in pressure and lift coefficients at fixed angles of attack and control surface deflections, indicating better stability. Unsteady aerodynamic effects of the airfoil at α=0∘ and “ramp” deflections of trailing- and leading-edge control surfaces from 0∘ to 40∘ with variations in actuation rates showed that CL peaks are approximately three to four times greater than static values for the case of the leading-edge control surface. This has significant implications for fixed-wing drone maneuverability and countering the effects of atmospheric turbulence.

An Optimisation of a Chordwise Slot to Enhance Lateral Flow Control on a UCAV

This research aims to optimise a chordwise slot so that lateral flow control of a flying wing configuration can be enhanced. This was achieved by maximising the airflow rate over the trailing edge control surfaces of the wing. A higher rate of airflow over the trailing edge control surfaces will increase the operationality of control surfaces, resulting in enhanced lateral control of the air vehicle at medium to high angles of attack. Four variables describing the chordwise slot are identified for numerical optimisation. They are: location, width, length, and angle of trajectory of chordwise slot relative to freestream. The results of the airflow rate for the wing with an optimised chordwise slot are compared with a baseline clean configuration. The flying wing configuration with a chordwise slot has shown a higher mass flow rate over the control surfaces of the wing in comparison to the baseline clean configuration, demonstrating that an optimised chordwise slot can be implemented on a flying wing configuration as it successfully controls the lateral flow by maximising the flow rate over the trailing edge control surfaces.

An aeroelastic beam finite element for time domain preliminary aeroelastic analysis

An aeroelastic beam finite element based on Euler-Bernoulli beam theory and two-dimensional unsteady aerodynamics is presented. Based on the flexural-torsional equations of motion of the wing the finite element formulation is provided by means of their weak form, taking into account the presence of trailing edge control surfaces that are considered flexible. Each aeroelastic beam presents transverse and rotation displacements due to bending and torsion of the wing box, torsion of the control surface, and the aerodynamic lag state. The developed finite element model is validated with literature and numerical flutter results of wings with and without control surfaces.

Enhancing Flight Envelope for a Nonlinear Aeroelastic Wing-Section Using Adaptive Fuzzy Sliding Mode Control Law

This paper presents an adaptive fuzzy sliding mode controller (AFSMC) for the control of a nonlinear aeroelastic wing section. The proposed controller estimates both the model dynamics of the system and the switching control of the sliding mode. Furthermore, the designed controller’s strategy is to enhance quality and ensure flight safety by providing additional damping to the already stable system and increasing the flight speed beyond the open-loop flutter speed. The aeroelastic system describes the plunge and pitch motions of the aircraft wing-section equipped with leading and trailing edge control surfaces. Additionally, the selected two-degree-of-freedom model includes structural stiffness nonlinearities and quasi-steady aerodynamics. The simulation results show the effectiveness of the proposed controller toward suppressing flutter and LCOs and reducing the vibrational level in the subcritical flight speed range. Additionally, the designed control strategy effectively controls and derives the state trajectories to the origin despite uncertainties and gust loads.

Nonlinear State Dependent Riccati Equation Controller for 3-DOF Airfoil with Cubic Structural Nonlinearity

In this paper, Limit Cycle Oscillations (LCOs) of a Two-Dimensional (2-D) airfoil with cubic structural nonlinearity in the plunge Degree of Freedom (DOF) are investigated. The aerodynamic loads are modelled using the unsteady Theodorsen’s theory. Wagner function and Jones’ approximation are used to transform the unsteady aerodynamic loads from frequency domain into time domain. The aeroelastic differential equations are solved using the routine ODE45 in MATLAB to get the system response. Both subcritical and supercritical LCOs are observed in the 3-DOF airfoil with structural nonlinearity in the plunge DOF. Such LCOs are undesirable phenomena and should be suppressed within the flight envelope. A nonlinear state feedback controller is designed to minimize the amplitude of LCOs presented in the considered nonlinear aeroelastic system. The State-Dependent Riccati Equation (SDRE) approach in combination with Kalman filter technique is used to design a controller for the 2-D airfoil with trailing edge control surface (flap). The forces and moments produced by this flap’s action are used to stabilize the system and suppress the existence of LCOs. The efficiency of the proposed nonlinear controller by the SDRE and Kalman filter in suppressing the existence of subcritical and supercritical LCOs are verified by the numerical simulation results.

Computational Assessment of Transonic Airfoil-Gust Aeroelastic Response

This study assesses the aeroelastic response that could arise from a periodic vertical gust of different frequencies encountering an airfoil which consists of a trailing-edge control surface (typically used for transonic aileron-buzz studies) and an airfoil undergoing pitching and plunging aeroelastic motion in transonic flow. The flutter phenomena and limit cycle oscillation correspond to two different types of self-induced instabilities, which happen at a Mach number, freestream velocity, and angle of attack combination for which the airfoil or control surface attains a stable oscillation. Gust acts as an external force to an aerostructural system. The airfoil shows a beating pattern under a periodic gust when the gust frequency coincides with the flutter frequency. However, this pattern evident at flutter velocity is absent when the periodic gust encounters an aeroelastic system at limit cycle oscillation velocity. The present work also shows the impact of periodic gust front on the shock/boundary-layer interaction modifying the coefficient of pressure and coefficient of skin friction distribution and leading to a multiple-frequency and multiple-amplitude aeroelastic motion. Fast Fourier transform is applied to the structural response to explore the frequency content corresponding to the structural response and dominant fluid modes for different flow Mach number regimes under the influence of the periodic gust. These frequency characteristics of the structural responses and flowfield in inviscid and viscous flowfield are distinct for different test cases and reveal the underlying physics of several multiple-frequency and multiple-amplitude aeroelastic motions, frequency coalescence, generation of a new frequency, and so on, as shown in this study.

Feedback control of limit cycle oscillations and transonic buzz, using the nonlinear transonic small disturbance aerodynamics

In this paper, a systematic method to suppress transonic buzz with feedback is presented. A trailing edge control surface in the form of part-span flap was used only to modify and control the unsteady aerodynamic loading on the wing. The flap rotation was used to provide feedback, which consisted of a weighted linear combination of the amplitudes of the principal modes of the structure, referred to as the control law. A linear, optimal feedback control law, that is synthesized systematically based on pseudo-spectral time domain analysis, may be used in principle, to assess its capacity to actively suppress the buzz in the transonic flow domain by using a servo-controlled control surface to modify the unsteady, nonlinear aerodynamic loads on the wing. Thus, it is essential that a set of feasible control laws are first constructed. In this paper, this is done by applying the doublet-lattice method. Restrictions, such as near-zero structural damping in the flap mode, were imposed on the aeroelastic model to facilitate the occurrence of transonic buzz. The feasible set of control laws were then assessed using the nonlinear transonic small disturbance theory and an optimum control is selected to suppress the buzz. The essential differences of the behaviour of the closed-loop system in nonlinear transonic flow, when compared to the applications of linear optimal control in linear potential flow, are presented and discussed.

Nonlinear aeroelastic analysis of a multi-element airfoil with free play using continuation method

In this study, the nonlinear aeroelastic response of a NACA0012 airfoil with leading-edge and trailing-edge control surfaces is studied numerically using a modified filtered impulse function method and continuation method. The source of nonlinearity comes from the free play behavior of both control surfaces which leads to the nonlinear properties of the structural stiffness matrix. The aerodynamic responses of these aeroelastic systems for varying oscillation frequencies are assumed to be linear for small deflections. These responses are developed using a reduced-order model by using the proposed modified filtered impulse function method on inviscid rotational flows computed using CFD methods for the Euler Equations. Key factors for maintaining accuracy using the proposed filtered impulse function are discussed with examples. The linearized aeroelastic equations are solved using the continuation method with the adaptive step size for three different situations that could arise in practice. The analysis process is conducted in the modes tracking and parameter variation separately. This study shows that compared to the airfoil with a single trailing-edge control surface in free play, the presence of the leading-edge flap motion could decrease the flutter speed and dominate the flutter mode in the absence of free play behavior. However, only slight differences could be observed from the limit cycle oscillations (LCO) analysis for which the structural nonlinearity is dominant owing to free play.

Data-driven active flutter control of airfoil with input constraints based on adaptive dynamic programming method

The stable flight region can be extended by adding control flap at the wing trailing edge and combined with active control technology. We studied the active flutter control by considering the input constraints. By designing the data-driven optimal controller, the limit cycle oscillations of a typical two-dimensional airfoil wing can be suppressed with single trailing edge control surface. The traditional control methods always need a precise mathematical model of the system, which put high requirements on system modeling. In this study, a novel data-driven optimal controller is proposed by using the input–output data and without depending on the nonlinear system dynamic model. This model-free approach avoids the effects of modeling errors and system uncertainty. When the data-driven controller is applied, the limit cycle oscillation phenomenon of the airfoil wing is eliminated within several seconds. It can be seen from the numerical simulation result that the data-driven adaptive dynamic programming control method possess superiority and feasibility.

Characterization of typical aeroelastic sections under combined structural concentrated nonlinearities

Structural nonlinearities are usually present in aeroelastic systems. The analysis of this system commonly comprises a study involving only one type of nonlinearity, influencing a particular motion of the airfoil. However, practical applications of aeroelastic systems can be affected by different types of structural nonlinearities. It becomes essential to study the stability of the aeroelastic system under these conditions to assess more real operational flight procedures. In this context, this article presents an investigation of a typical aeroelastic section response with trailing edge control surface subjected to combinations of concentrated structural nonlinearities. Different nonlinear scenarios involving cubic hardening stiffness in pitching and free play, free play with preload, and slip dry friction in the trailing edge control surface motion are analyzed. The mathematical model is based on linear unsteady aerodynamics coupled to a three-dof typical aeroelastic section. Hopf bifurcations diagrams are obtained from direct time integration of the equation of motion. The post-flutter limit cycle oscillations are investigated, revealing supercritical and subcritical bifurcations. A complete parametric study of the nonlinear parameters is carried out, thereby allowing a sensitivity analysis of each nonlinear scenario. The results show that aeroelastic tailoring considering the mild post-flutter behavior can be achieved through an appropriate choice of combined nonlinear effects. Moreover, combined nonlinearities can mitigate the undesired subcritical aeroelastic responses caused by free play.

Generalized composite noncertainty-equivalence adaptive control of a prototypical wing section with torsional nonlinearity

The paper presents a generalized composite noncertainty-equivalence adaptive control system for the control of a prototypical aeroelastic wing section using a single trailing-edge control surface. The plunge–pitch (two-degree-of-freedom) dynamics of this aeroelastic system include torsional pitch-axis nonlinearity. The open-loop system exhibits limit cycle oscillations beyond a critical free-stream velocity. It is assumed that parameters of the model are not known. The objective is to suppress the oscillatory responses of the system. Based on the immersion and invariance approach, a generalized composite noncertainty-equivalence adaptive (NCEA) control system for regulation of the pitch angle is designed. The control system consists of a control module and a composite parameter identifier—designed independently. The composite integral parameter estimation law is based on (1) the immersion and invariance (I&I) theory, (2) gradient-based adaptation algorithm, and (3) classical certainty-equivalence adaptive (CEA) update rule. Besides the composite integral component, the full parameter estimate also includes a judiciously chosen nonlinear algebraic function. This composite identifier inherits stronger stability properties. Using the Lyapunov analysis, asymptotic suppression of the limit cycle oscillations and the boundedness of system trajectories are established. Interestingly, in the closed-loop system including the composite update rule, there exist two attractive manifolds to which the system’s trajectories converge. Simulation results are presented which show the suppression of the oscillatory plunge displacement and pitch angle responses despite uncertainties in the model parameters. Furthermore, the performance and stability properties of this composite NCEA control system—including the gradient-based adaptation and the update rule of the CEA system—are better than the simple NCEA system.

The interaction between active aeroelastic control and structural tailoring in aeroservoelastic wing design

This paper presents an analysis of the interaction and trade-off between active aeroelastic control and passive structural tailoring on a free-flying fully flexible aircraft model. Both technologies are included in the preliminary design of a typical transport aircraft configuration with a conventional control surface layout containing trailing edge control surfaces and spoilers. The passive structural tailoring is facilitated by exploiting the anisotropic properties of composite materials to steer the static and dynamic aeroelastic behaviour. Active aeroelastic control is implemented by scheduled control surface deflections redistributing the aerodynamic loads during manoeuvres to achieve manoeuvre load alleviation and a feed-forward control law for gust load alleviation. The panel-based aerodynamic modelling of spoiler deflections is improved by a correction of the spatial distribution of the boundary condition derived from higher fidelity simulation data. The optimisation of active control laws requires the consideration of constraints of the actuation system, namely rate and deflection saturation, in a nonlinear manner. The interaction of manoeuvre load alleviation, gust load alleviation and passive structural tailoring is investigated on the basis of results of different aeroservoelastic optimisations. Therefore the primary wing structure is simultaneously optimised with the individual technologies being activated or deactivated, resulting in eight different wing structures. The results of the individual and combined optimisations reveal significant design differences. The potentials of the different technologies can only be optimally exploited by simultaneous optimisation. The paper concludes with a study of the sensitivity of the major findings with respect to the knockdown factor for failure applied to the material properties. A substantial shift of effectiveness from active aeroelastic control to passive structural tailoring is observed with increased allowables resulting in more flexible and hence less stiff wing designs.

Flutter and Limit Cycle Oscillations of Cantilevered Plate in Supersonic Flow

Research interest is growing for theoretical models of highly deflected structures in aeroelastic settings. Presented here is a model of a cantilevered plate subjected to axial supersonic flow to determine the flutter boundary and postflutter characteristics of a system such as a trailing edge control surface. The structural model is a nonlinear inextensible beam model with inertia and stiffness geometric nonlinearities, while the aerodynamic model used is both first-order linear and third-order nonlinear Piston Theory with a new geometric modification to account for large deflections of the cantilevered configuration. Comparisons are made between linear and nonlinear structural models as well as linear and nonlinear Piston Theory, with and without this new geometric modification. It is shown that the model is highly sensitive to the inclusion of each nonlinear aerodynamic or structural component and that the new geometric modification to Piston Theory leads to stable limit cycles which otherwise may be unstable. Finally, the use of Piston Theory for these unconventionally large deflections is validated by comparing pressures on the structure to those computed by the Euler equations acting on the structure’s prescribed motion.

Computational Evaluation of Control Surfaces Aerodynamics for a Mid-Range Commercial Aircraft

Computational fluid dynamics is employed to predict the aerodynamic properties of the prototypical trailing-edge control surfaces for a small, regional transport, commercial aircraft. The virtual experiments are performed at operational flight conditions, by resolving the mean turbulent flow field around a realistic model of the whole aircraft. The Reynolds-averaged Navier–Stokes approach is used, where the governing equations are solved with a finite volume-based numerical method. The effectiveness of the flight control system, during a hypothetical conceptual pre-design phase, is studied by conducting simulations at different angles of deflection, and examining the variation of the aerodynamic loading coefficients. The proposed computational modeling approach is verified to have good practical potential, also compared with reference industrial data provided by the Leonardo Aircraft Company.

Experimental Characterization of Nonlinear Damping in the F-16 Wing Structure

Earlier work did not find any static stiffness nonlinearities in the F-16 wing structure, but suggested that dynamic nonlinearities may be present. Computational studies have demonstrated that small amounts of nonlinear structural damping are sufficient to change a diverging flutter oscillation into a limited amplitude oscillation. Experimental damping measurements of a fixture-mounted F-16 wing were made to determine whether damping nonlinearities are significant and to quantify the nature of any structural damping present in the built-up wing structure. Three fundamental component configurations of the wing were examined: wing box only, wing (i.e., wing box with leading- and trailing-edge control surfaces), and wing with wingtip launcher. A ground vibration test approach was used with a stepped-sine excitation technique over a range of excitation forces. Damping and frequency were extracted from measured frequency response functions of the stepped-sine test data using a modified single degree-of-freedom modal analysis technique. Nonlinear damping with respect to bending mode displacement was observed for all test configurations. The overall magnitude of the damping increased when control surfaces were installed on the wing box. The presence of a wingtip launcher decreased the damping at higher displacements.

Low Reynolds number aerodynamics of leading-edge and trailing-edge hinged control surfaces: Part I statics

The use of a leading-edge hinged control surface, as opposed to a traditional trailing-edge hinged control surface, is investigated in the context of improving control authority of fixed-wing Micro Air Vehicles when subject to gusts. Static and dynamics forces are considered in two separate papers (Part I and II). In Part I we document the static lift characteristics of two flat-plate airfoils, having either a 30% chord leading- or trailing-edge hinged control surface, resolved from surface pressure measurements captured at angles of attack ranging from 0∘ to 30∘ and control surface deflections from 0∘ to ±45∘. The lift response of the airfoil with the leading-edge control surface was complex with a reversal in lift noted beyond specific control surface deflections (for e.g., at α=0∘, CL reversed beyond ψ=±20∘). The use of a trailing-edge control surface was found to be more effective than the leading-edge control surface. Analysis of the CP and CL fluctuations at fixed α and control surface deflections revealed the fluctuations were significantly less with leading-edge control surfaces than trailing-edge control surfaces. Furthermore, a simple potential flow model exhibited good correlation of CL against experimental results at small angles of attack and control surface deflections.

Flutter Control of a Two-dimensional Airfoil based on Adaptive Dynamic Programming

The severe aeroelastic vibrations can lead to structural damage and threaten the flight safety. It is an important topic to develop active controllers for aeroelastic stabilization and flutter suppression. The traditional control methods need the clearly established mathematic model, whereas establishing an accurate airfoil model with structural and aerodynamic nonlinearities is a difficult task. So a novel data-driven optimal design scheme, adaptive dynamic programming (ADP), is proposed in this paper for a nonlinear airfoil system with one trailing-edge control surface. The proposed ADP algorithm iteratively updates the control policy by using the state and input data without knowing the system dynamical model. Calculation shows that the control policy is convergent after several iterations. Finally, the superiority and feasibility of the control method are verified by simulation.

Multiple-Input Multiple-Output Experimental Aeroelastic Control Using a Receptance-Based Method

This paper presents the first experimental study of multiple-input multiple-output (MIMO) active vibration suppression by pole placement using the receptance method. The research is based on a purpose-built modular flexible wing equipped with leading- and trailing-edge control surfaces and two displacement sensors for measuring its position. The MIMO controller has the advantage of being designed entirely on frequency response functions, which are measured between the control surface position (control inputs) and the structural displacements (outputs), and include the actuator dynamics. There is no requirement to evaluate or to know the structural , , and matrices or the aerodynamic loads; and the formulation eliminates the need for a state observer. The controller is first implemented numerically and then experimentally on the aeroelastic system. Both frequencies and damping are assigned (together or independently) for the first two vibration modes. The research includes a procedure for assessing the control effort required and demonstrates an effective means of increasing the flutter margin while the effort is minimized.

Transonic flutter suppression for a three-dimensional elastic wing via active disturbance rejection control

The paper presents an error-based active control law for the transonic flutter suppression of a three dimensional elastic wing in a wide range of Mach numbers and dynamic pressures. The extended state observer is the key issue for the presented controller and provides the observed output signal and the observed total disturbance signal at the same time. The observed output signal obtained via the extended state observer enables one to design a feedback control law for adjusting the output errors, while the observed total disturbance signal linearizes the controlled system and guarantees the anti-interference of the proposed controller. To demonstrate the performance of the proposed controller system for transonic flutter suppression, the transonic aeroservoelastic behaviors of a three-dimensional elastic wing with both leading-edge and trailing-edge control surfaces are illustrated as the test cases. During the numerical simulations, only the tailing-edge control surface is used as the control input and the leading-edge control surface is fixed. The numerical results show that the active control law can effectively suppress the transonic flutter at a wide range of Mach numbers and dynamic pressures. Moreover, the controller system has robustness with respect to certain measurement noises and time delay.