Trailing-edge Control Surface Research Articles

Efficient roll control using distributed control surfaces and aeroelastic effects

The potential of using multiple leading and trailing edge control surfaces and aeroelastic effects for efficient roll manoeuvring is investigated. Numerical optimization in combination with a simulation model including aeroelastic dynamics is used to design a controller for roll angle tracking. The controller distributes the control power to the individual surfaces such that it minimizes the control effort yet fulfilling roll performance requirements in a wide airspeed envelope. The controller is implemented and experimentally validated using an elastic wind-tunnel model equipped with 16 individual control surfaces. Good correlation between simulations and experiments is obtained although some deviations are observed and discussed. Finally, the choice of the most efficient control surface layout is investigated by evaluating control laws which utilize a subset of the available control surfaces.

A flutter suppression active controller

Flutter is a dynamic aeroelastic instability that involves the interaction of aerodynamic, elastic, and inertial forces. This instability may occur in aircraft surfaces, like wings and tails, leading them to a divergent oscillatory motion. A classical two-degrees-of-freedom flutter is an interaction of bending and torsional modes of vibration of a structure. A flexible mount system has been developed for flutter tests with rigid wings in wind tunnels. This flexible mount has to provide a well-defined two-degrees-of-freedom system on which rigid wings encounter flutter. Active control schemes for flutter suppression can be tested using this experimental set-up. An aeroelastic model is formulated to simulate the aeroelastic behaviour of this system. The equations of motion are developed using Lagrange' equations and the Principle of Virtual Work, resulting in a state space representation. This is a convenient way to determine pitch and plunge time responses for several initial conditions and velocities. Using this model, a state feedback controller of the form u = – Kx is determined. The wind tunnel model is a rectangular wing with a NACA 0012 airfoil section and a trailing edge control surface as actuator. The main goal is to suppress flutter and to maintain the stability of the closed loop system.

Adaptive and neural control of a wing section using leading- and trailing-edge surfaces

This paper treats the question of control of nonlinear aeroelastic responses of a prototypical wing section with structural nonlinearity using leading- and trailing-edge control surfaces. It is assumed that all the aerodynamic, structural and inertia parameters are unknown to the designer. As such the limitation of a recent control design reported in the literature, which requires complete knowledge of aerodynamic derivatives and inertia parameters, is removed. An adaptive controller and a neural control system are designed for the trajectory control of the plunge displacement and pitch angle. For the derivation of the adaptive control law, a linearly parameterized model is used but the neural controller is designed by treating the stiffening-type structural nonlinearity as an unstructured function (not parameterizable). It is shown that the adaptive and neural controllers accomplish trajectory control in the closed-loop system. Simulation results are presented which show that these controllers are effective in regulating the nonlinear responses to the origin in the state space in spite of large model uncertainties. Moreover unlike the model with a single trailing-edge surface, two control surfaces provide flexibility in shaping both the plunge and pitch responses.

Limit-cycle stability reversal via singular perturbation and wing-flap flutter

A three-degree-of-freedom aeroelastic typical section with a trailing-edge control surface is theoretically modelled, including nonlinear springs for both the nonlinear description of the torsional stiffness and of the hinge elastic moment. Furthermore, augmented states for linear unsteady aerodynamic of 2-D incompressible potential flow, have been considered in the model. First, the system response is determined by numerically integrating the governing equations using a standard Runge–Kutta algorithm in conjunction with a ‘shooting method’. The numerical analysis has revealed the presence of stable and unstable limit cycles, along with stability reversal in the neighborhood of a Hopf bifurcation. Consequently, the equations of motion are analysed by a singular perturbation technique based on the normal-form method. This method, originally introduced by strictly applying a resonance condition, is herein extended by applying a near-resonance condition in order to improve the semi-analytical description of the stability reversal behavior. Therefore, amplitudes and frequencies of limit cycles depending on the flow speed V are obtained from the normal-form equations, and the terms which are essentially responsible for the nonlinear system behavior are identified.

Transient response study on rolling effectiveness of multiple control surfaces

A study of the effectiveness of trailing- and leading-edge control surfaces has been made for a rolling wing-fuselage model. An experimental model and wind tunnel test are used to assess the theoretical results. The theoretical model includes the inherently nonlinear dry friction damping moment that is present between the spindle support and the experimental aeroelastic wing model. The roll trim equation of motion and the appropriate aeroelastic equations are solved for different combinations of leading- and trailing-edge control surface rotations using a reduced-order aerodynamic model based upon the fluid eigenmodes of three-dimensional vortex lattice aerodynamic theory. The focus is on the transient response of the system. The present paper provides new insights into the transient dynamic behavior and design of an adaptive aeroelastic wing using trailing- and leading-edge control surfaces.

Overview of the DARPA Smart Wing Project

The recently completed DARPA/AFRL/NASA Smart Wing program, performed by a team led by Northrop Grumman Corporation (NGC), addressed the development and demonstration of smart materials based concepts to improve the aerodynamic and aeroelastic performance of military aircraft. This paper presents an overview of the program. The program was divided into two phases. Under Phase 1 (January 1995 to February 1999), the NGC-led team developed adaptive wing structures with integrated actuation mechanisms to replace standard hinged control surfaces and provide variable, optimal aerodynamic shapes for a variety of flight regimes. An important limitation of the Phase 1 effort, the low bandwidth achievable in Shape Memory Alloy (SMA)-based actuation, was addressed in Phase 2 (January 1997 to November 2001). Under Phase 2, a 30-percent scale full span wind tunnel model of an NGC Uninhabited Combat Air Vehicle (UCAV) design was developed. For the Phase 2 first wind tunnel test, completed in March 2000, SMA-actuated, hingeless, smoothly contoured, flexible leading and trailing edge control surfaces were incorporated on one wing of the model and conventional trailing edge control surfaces actuated using electric motors on the other. This test provided baseline steady-state data at Mach numbers ranging from 0.3 to 0.8. The test also demonstrated the benefits of the smart leading edge control surface to compensate for loss of aileron effectiveness with increasing dynamic pressure. The second test, performed in May 2001, demonstrated a hingeless, smoothly contoured, structurally compliant, trailing edge control surface actuated using piezoelectric motors. Spanwise and chordwise shape control was demonstrated for the smart trailing edge control surface at deflection rates of up to 80°/s. Performance improvements in terms of increased rolling and pitching moments and lower control surface deflections were quantified. The work performed under this program has demonstrated the feasibility of developing smart control surface designs to provide optimal aerodynamic performance at a wide range of flight conditions.

Development of High-rate, Adaptive Trailing Edge Control Surface for the Smart Wing Phase 2 Wind Tunnel Model

The DARPA/AFRL/NASA Smart Wing program, led by Northrop Grumman Corporation (NGC) under the DARPA Smart Materials and Structures initiative, addressed the development of smart technologies and demonstration of relevant concepts to improve the aerodynamic performance of military aircraft. In Phase 2, Test 2 of the program, the main objective was to demonstrate high-rate actuation of hingeless, spanwise, and chordwise deformable control surfaces using smart materials-based actuators on a 30% scale, full span wind tunnel model of a proposed NGC uninhabited combat air vehicle (UCAV). A minimum actuation rate of 25° flap deflection in 0.33 s, producing a slew rate of 75°/s, was desired. This slew rate is representative of many operational military aircrafts with hinged control surfaces. Numerous trade studies were performed on a variety of smart materials and flexible structure configurations before arriving at the final trailing edge structure design that consisted of a flexcore center and elastomeric outer skin actuated by high-power ultrasonic motors using an eccentric motion. The trailing edge control surface fitted onto the wind tunnel model comprised 10 eccentric-driven segments connected together by a continuous outer skin and a flexible hinge pin at the trailing edge tip. This pinned configuration allowed the segments partial freedom to rotate about each other, but constrained any lateral motion thus giving a smooth trailing edge shape for nonuniform spanwise deflections. To control the 10 segments of the trailing edge, a VME-based control system with high speed, simultaneously sampled A/D and D/A boards and a dedicated DSP board was developed. This paper describes the analysis and design of the flex structure, ultrasonic motor selection and performance, element and coupon tests to verify analysis, control system development, model integration, and results from the wind tunnel test.

Aerodynamic Performance of the Smart Wing Control Effectors

This paper presents and discusses the wind tunnel results obtained during Phase 1 and Phase 2 of the Smart Wing program. A series of four tests was conducted, two in each phase. The purpose of the tests was to evaluate the performance of smart material based control effectors in representative aerodynamic environments. In Phase 1, wing twist using a Shape Memory Alloy (SMA) torque tube and smoothly contoured trailing edge surfaces enabled by SMA wires were evaluated on a semi-span, 16% scale model of a typical fighter wing. The focus of Phase 2 was to address the design and demonstration of a high-frequency, large-deflection, and smoothly contoured trailing edge control surface capable of a spanwise variation in deflection. The actuation system for this control surface was based on piezoelectric (PZT) ultrasonic motors. A smoothly contoured leading edge control surface was also built and tested using SMA wires. These designs were evaluated on a 30%, full-span model of a representative Unmanned Combat Air Vehicle (UCAV). In each wind tunnel entry comparisons were made to the performance of conventional control surfaces. Successful results included: improved aileron effectiveness at high dynamic pressures; demonstrated improvements in lateral and longitudinal control effectiveness with smoothly contoured control surfaces over conventional hinged control surfaces; continuous spanwise shape control; and, large deflections at rates over 80°/s, which is well within the desired deflection aileron rates for control of fighter aircraft.

Control of a Nonlinear Wing Section Using Leading- and Trailing-Edge Surfaces

The control of nonlinear aeroelastic response of a wing section with a continuous stiffening-type structural nonlinearity is examined through analytical and experimental studies. Motivated by the limited effectiveness of using a single, trailing-edge control surface for the suppression of limit-cycle oscillations of a typical wing section, improvements in the control of limit-cycle oscillations are investigated through the use of multiple control surfaces, namely, an additional leading-edge control surface. The control methodology consists of a feedback linearization approach that transforms the system equations of motion via Lie algebraic methods and a model reference adaptive control strategy that augments the closed-loop system to account for inexact cancellation of nonlinear terms due to modeling uncertainty. Specifically, uncertainty in the nonlinear pitch stiffness is examined. It is shown through simulations and experiments that globally stabilizing control may be achieved by using two control surfaces.

Aeroelastic Wing with Leading- and Trailing-Edge Control Surfaces

It is well known that the effectiveness of a trailing-edge control surface can be substantially diminished due to the elastic twist of an airfoil or wing. This aeroelastic phenomenon is known as control surface reversal when the lift or roll rate vanishes at a sufficiently large ratio of flow dynamic pressure to wing stiffness. However, a leading-edge control surface can be used to counteract control surface reversal, and indeed, in principle, a leading-edge control surface may entirely cancel the tendency of the trailing-edge control surface to undergo reversal. Moreover, analysis shows that by using a simple control strategy one can use a combination of leading- and trailing-edge control surface rotations to maintain lift and roll effectiveness and minimize control surface rotations. The beneficial effects of leading-edge control surfaces on control surface reversal are known to practioners. However, the present simple model makes these especially transparent and suggests an advantageous strategy using a combination of leading- and trailing-edge control surfaces.

Experimental and Theoretical Study on Rolling Effectiveness of Multiple Control Surfaces

It is well known that the effectiveness of a trailing-edge control surface can be substantially diminished due to the elastic twist of an airfoil or rolling wing. This aeroelastic phenomenon is known as control surface reversal when the lift or rolling rate vanishes at a sufe ciently large ratio of e ow dynamic pressure to airfoil or wing stiffness. However, a leading-edge control surface can be used to counteract control surface reversal, and, indeed, in principle, a leading-edge control surface can entirely cancel the tendency of the trailing-edge control surface to undergo reversal. Moreover, analysis shows that by using a simple adaptive control strategy, one can use a combination positive and negative control surface rotations to maximize lift and rolling effectiveness or minimize control surface rotations. In the present work, a theoretical ‐experimental study of the effectiveness of trailingand leading-edge control surfaces has been made for a rolling wing ‐fuselage model. An experimental model and wind-tunnel test are used to assess the theoretical results. The theoretical model includes the inherently nonlinear dry friction damping moment between the spindle support and the experimental aeroelastic wing model for the rollingdegreeoffreedom.Athree-dimensionalvortexlatticeaerodynamictheoryisemployed.Newinsightsintothe behavior and design of an adaptiveaeroelastic wing using trailing- and leading-edge control surfacesare provided. Nomenclature c = wing chord length K® = torsional stiffness km, kn = numbers of vortex elements on the wing in x and y directions, respectively kmm = total number of vortices on both the wing and wake in the x direction L = total lift on the wing Llocal = local lift along the span

Aerodynamic and Aeroelastic Characteristics of Wings with Conformal Control Surfaces for Morphing Aircraft

Investigations are conducted on lifting surfaces with conventional and conformal trailing-edge control surfaces. The Sohngen inversion formula is used with the thin-airfoil integral equation to determine the aerodynamic pressure for various control surface chord-to-airfoil chord ratios. Comparisons to a conventional control surface show increases in lift and pitching moment of the airfoil with a conformal control surface. Aerodynamic pressure distributions acting on a wing with control surfaces are determined with the vortex lattice technique. Predicted aerodynamic pressures and roll moments are compared to available wind-tunnel data and provide a more general understanding of theaerodynamicbehavior observed there. Roll performance of a rectangular wing is determined for various control surface chord-to-wing chord ratios. It is found that the maximum roll rate is greater for a wing with a conformal control surface, but has a lower reversal dynamic pressure than the wing with a conventional control surface. The aerodynamic and aeroelastic results obtained from this investigation provide some insight for wings designed with conformal control surfaces.

Sliding Mode Control Application for Two-Dimensional Active Flutter Suppression.

This paper applies a sliding mode controller for suppressing two-dimensional incompressible flow flutter problems in which aerodynamic uncertainties and structural nonlinearity are incorporated. To consider the aerodynamic uncertainties, the steady aerodynamic theory is regarded for the controller design, although the quasi-steady aerodynamic theory is employed in numerical simulation. The structural nonlinearity is included about pitch motion to make a limit cycle oscillation appearing in flutter phenomenon. The control inputs are leading and trailing edge control surfaces. At a velocity of 20m/s, the sliding mode controller is designed to suppress the initial response within 1 second while the uncontrolled wing shows limit cycle oscillation. Furthermore, the sliding mode controller shows robust characteristics in a wide range of the flow velocity. These results indicate that the sliding mode controller is effective for suppressing the flutter problems that have aerodynamic uncertainties and structural nonlinearity.

スライディングモード制御による遷音速フラッタのアクティブ制振

This paper describes two-dimensional active flutter suppression to cope with the transonic dip using the sliding mode control. The airfoil model has plunge and pitch degrees of freedom with leading and trailing edge control surfaces. The aerodynamic forces acting on the airfoil, lift and pitching moment, are calculated by solving Euler's equations using computational fluid dynamics. At a specific altitude, flutter occurs between Mach number of 0.7 and 0.88, which corresponds to the transonic dip. The sliding mode control makes the airfoil to be stable all through the Mach number including the transonic dip. The sliding mode controller gives wider flutter margin than a linear quadratic regulator. These characteristics indicate that the sliding mode control is useful for active flutter suppression in the transonic flight.

Active Control of Flutter in Compressible Flow and Its Aeroelastic Scaling

A new two-pronged approach suitable for the general aeroelastic and aeroservoelastic scaling of any aeroelastic cone guration is presented. The method produces aeroelastic scaling laws for general cone gurations, and it is particularly useful for situations involving active controls and smart-materials-based actuation. This approach is illustrated by applying it to a two-dimensional wing section in compressible e ow, combined with a trailing-edge control surface. Augmentedaerodynamicstatesarereconstructedusing a Kalman e lter, and linearoptimal control isused to design a full-statefeedback e uttersuppression controller. Constraints on actuator dee ection and rateare alsoconsidered.Inasecondexample,e uttersuppressionfora typicalcrosssectionwithaconventionaltrailing-edge control surface is compared with that obtained with piezoelectric actuation utilizing bend/twist coupling. Nomenclature [A] = system matrix [AH] = Hamiltonian system matrix a = nondimensional offset between elastic axis (EA) and midchord [B] = open-loop control matrix in state space b = airfoil semichord [C] = observation matrix, y(t) = [C]{x(t)} C(k) = Theodorsen’ s lift dee ciency function Ch, Cl, Cm = hinge moment, lift, and pitch moment coefe cients per unit span CP, ¯ CP = power coefe cient and its average value

Computational Fluid Dynamics and Doublet-Lattice Calculation of Unsteady Control Surface Aerodynamics

Accurate prediction of control surface aerodynamics has been a challenge since the dawn of aviation. Whereas this has been an important problem for many years, recent increases in the use of control surfaces for active control (load alleviation and e utter suppression ) have increased the importance of accurate steady and unsteady control surface aerodynamics. Because of the strong ine uence of viscosity on the pressures on a trailing-edge control surface, the aerodynamic theories based on the linear potential equation have had only marginal success in predicting control surface aerodynamics, and in practice, large corrections (based on wind-tunnel data ) are often required for acceptable accuracy. Recent advances in computing technology and unsteady aerodynamic codes based on the Navier ‐Stokes equation areallowing moreaccurate analysesto be performed. Unsteady aerodynamic calculations due to control surface oscillations are made using a linear potential code (N5K) and a Navier‐Stokes code (CFL3D.AE-BA version 4.1 ). The Navier‐Stokes calculations are performed in the time domain using an exponential pulse technique and are transformed to the frequency domain using Fourier transform. For low reduced-frequency cases, theNavier ‐Stokes calculations arecompared to the doublet-lattice method and to experiment, and the advantages of the nonlinear analysis are clearly demonstrated. Correlation between Navier ‐Stokes and doublet-lattice results is then studied for higher reduced frequencies.

Fundamental Aeroservoelastic Study Combining Unsteady Computational Fluid Mechanics with Adaptive Control

A computational two-dimensional aeroservoelastic study in the time domain is described. The model, which is based on exact inviscid aerodynamics, captures the large amplitude motions and the associated strong shock dynamics in the transonic regime. The aeroservoelastic system consists of a two-degree-of-freedom airfoil with a trailing-edge control surface. By the use of first-order actuator dynamics, a digital adaptive controller is applied to provide active flutter suppression. Comparisons between time responses of the open-loop and closed-loop systems show the ability of the trailing-edge control surface to suppress nonlinear transonic aeroelastic phenomena. A relation between actuator dynamics, sampling time step, and limits on the flap deflection angle to guarantee the effectiveness of the adaptive controller is illustrated by the results.

Benchmark Active Controls Technology Model Aerodynamic Data

The Benchmark Active Controls Technology (BACT) model is a part of the Benchmark Models Program (BMP). The BMP is a NASA Langley Research Center program that includes a series of models which were used to study different aeroelastic phenomena and to validate computational fluid dynamics codes. The primary objective of BACT testing was to obtain steady and unsteady loads, accelerations, and aerodynamic pressures due to control surface activity in order to calibrate unsteady CFD codes and active control design tools. Three wind-tunnel tests in the Transonic Dynamics Tunnel (TDT) have been completed. The first and parts of the second and third tests focused on collecting open-loop data to define the model''s aeroservoelastic characteristics, including the flutter boundary across the Mach range. It is this data that is being presented in this paper. An extensive database of over 3000 data sets was obtained. This database includes steady and unsteady control surface effectiveness data, including pressure distributions, control surface hinge moments, and overall model loads due to deflections of a trailing edge control surface and upper and lower surface spoilers.

Computational Test Cases for the Benchmark Active Controls Model

Unsteady aerodynamic and aeroelastic computational test cases have been selected from the data set for the Benchmark Active Controls Technology model. Data sets for trailing-edge control surface oscillations and uppersurfacespoileroscillationsforarangeofMachnumbers, angleofattack, andstaticcontroldee ectionsareincluded. Cases also include static angles of attack, static trailing-edge control, and static upper-surface spoiler dee ections for a range of conditions near those for the oscillation cases. The aeroelastic test cases include three types of e utter instability and dynamic response measurements for oscillations of the controls on the e exible mount.

Stability and Control of a Structurally Nonlinear Aeroelastic System

The authors examine the stability properties of a class of nonlinear controls derived via feedback linearization techniques for a structurally nonlinear prototypical two-dimensional wing section. In the case in which the wing section has a single trailing-edge control surface, the stability of partial feedback linearization to achieve plunge primary control is studied. It is shown for this case that the zero dynamics associated with the closed-loop system response are locally asymptotically stable for a range of e ow speeds and elastic axis locations. However, there exist locations of the elastic axis and speeds of the subsonic/incompressible e ow for which this simple feedback strategy exhibits a wide range of bifurcation phenomena. Both Hopf and pitchfork bifurcations evolve parametrically in terms of the e ow speed and elastic axis location. In the case in which the wing section has two control surfaces, the global stability of adaptive control techniques derived from full feedback linearization is studied. In comparison with partial or full feedback linearization techniques, the adaptive control strategies presented do not require explicit knowledge of the form of the structural nonlinearity.