The interaction between an aircraft's structural dynamics, unsteady aerodynamics and flight control system is known as aeroservoelasticity. The problem can occur because the control system sensors are of sufficient bandwidth to sense the structural vibrations as well as the rigid-body motion of the aircraft. This sensed structural vibration can result in further excitation of the structure through both aerodynamic and inertial excitation, leading to a potential closed-loop instability. At present, such an unstable interaction is prevented by the inclusion of notch filters within the feedback path which have a detrimental effect on the aircraft's rigid-body performance. The current clearance procedure is restricted by a poor understanding of the array of complex issues involved. The aim of the project was to develop a clearer understanding of the interactions between system components leading to a reduction in the clearance requirements. Work has concentrated on the effects of system nonlinearities and on the digital nature of modem control systems. A major source of nonlinearities within the control system are the servo-hydraulic actuators. Through detailed actuator modelling confirmed by rig testing of actual hardware, these nonlinearities are analysed and a method for predicting the response of the actuators in the presence of two input signals proposed. As a result, it is demonstrated that an unstable structural oscillation would cause a limit-cycle oscillation as opposed to an unbounded response. Through nonlinear system theory the criteria for the existence of such limit-cycles are obtained, enabling them to be predicted and therefore prevented. Consideration of the true nonlinear nature of the aeroservoelastic system has enabled an alternative design and clearance procedure to be proposed which reduces the attenuation requirements of the structural-mode filters whilst ensuring satisfactory aircraft performance even in the presence of modelling errors. This design procedure is demonstrated on both a model of the aircraft system and a simple test system enabling verification of the nonlinear analysis and comparison between the current and proposed alternative procedures. As a result, it is demonstrated that consideration of the true nonlinear nature of the aeroservoelastic interaction has the potential for allowing a significant reduction in structural filter attenuation requirements. Consequently, a reduction in the phase lag due to the filters is possible resulting in an improvement in closed-loop system performance whilst ensuring the safe operation of the aircraft.

O controle de flutter é um problema cuja solução é almejada ao longo de décadas e que ainda se apresenta como um desafio considerável. Os desafios residem basicamente no projeto de dispositivos de atuação eficientes e na síntese das leis de controle. A lógica difusa se mostra como uma técnica promissora e efetiva de controle. Neste trabalho é proposto um controlador difuso para o controle de flutter em uma asa com aerofólio do tipo NACA 0012, tendo como superfície de controle um flap. Este modelo físico é acoplado a um dispositivo elástico de sustentação da asa - DESA, que possibilita os movimentos de arfagem e deslocamento vertical. Para o desenvolvimento do modelo matemático são utilizadas as Equações de Lagrange e o Princípio do Trabalho Virtual. A determinação dos parâmetros estruturais do conjunto ASA/DESA é realizada a partir de um modelo em elementos finitos e de uma análise modal experimental. Os modos de vibrar são determinados através do ERA - Eigensystem Realization Algorithm. O estudo da atuação do controlador difuso é realizado através de simulação computacional e de análise experimental. Dois modelos difusos são utilizados na construção do controlador, o de Mamdani e o de Takagi-Sugeno-Kang. Os resultados obtidos mostram que o controlador difuso, para os dois modelos, é bastante eficiente no controle do flutter. / Flutter control is a problem for which solution has been longed for decades and still is a considerable challenge. The challenges reside basically in the development of devices with efficient performance and in the synthesis of the control laws. Fuzzy logic appears as a promising and effective technique of control. In this work, a fuzzy controller is proposed for flutter control of a wing with NACA 0012 airfoil section, having as control surface a trailing edge flap. This physical model is coupled to the elastic support device of the wing - DESA, which allows plunge and pitch displacements. For the development of the mathematical modeI, the Lagrange Equations and the Principle of the Virtual Work are used. The determination of the structural parameters of the Wing/DESA device is accomplished starting from a finite element model and from an experimental modal analysis. The vibrating modes are obtained using the Eigensystem Realization AIgorithm - ERA. The study of the fuzzy controllers performance is accomplished through simulation and experimental analysis. Two fuzzy models are used for the controller: the Mamdani and the Takagi- Sugeno-Kang. The results show that both fuzzy controllers are quite efficient in the control of flutter.

O controle de flutter é um problema cuja solução é almejada ao longo de décadas e que ainda se apresenta como um desafio considerável. Os desafios residem basicamente no projeto de dispositivos de atuação eficientes e na síntese das leis de controle. A lógica difusa se mostra como uma técnica promissora e efetiva de controle. Neste trabalho é proposto um controlador difuso para o controle de flutter em uma asa com aerofólio do tipo NACA 0012, tendo como superfície de controle um flap. Este modelo físico é acoplado a um dispositivo elástico de sustentação da asa - DESA, que possibilita os movimentos de arfagem e deslocamento vertical. Para o desenvolvimento do modelo matemático são utilizadas as Equações de Lagrange e o Princípio do Trabalho Virtual. A determinação dos parâmetros estruturais do conjunto ASA/DESA é realizada a partir de um modelo em elementos finitos e de uma análise modal experimental. Os modos de vibrar são determinados através do ERA - Eigensystem Realization Algorithm. O estudo da atuação do controlador difuso é realizado através de simulação computacional e de análise experimental. Dois modelos difusos são utilizados na construção do controlador, o de Mamdani e o de Takagi-Sugeno-Kang. Os resultados obtidos mostram que o controlador difuso, para os dois modelos, é bastante eficiente no controle do flutter. / Flutter control is a problem for which solution has been longed for decades and still is a considerable challenge. The challenges reside basically in the development of devices with efficient performance and in the synthesis of the control laws. Fuzzy logic appears as a promising and effective technique of control. In this work, a fuzzy controller is proposed for flutter control of a wing with NACA 0012 airfoil section, having as control surface a trailing edge flap. This physical model is coupled to the elastic support device of the wing - DESA, which allows plunge and pitch displacements. For the development of the mathematical modeI, the Lagrange Equations and the Principle of the Virtual Work are used. The determination of the structural parameters of the Wing/DESA device is accomplished starting from a finite element model and from an experimental modal analysis. The vibrating modes are obtained using the Eigensystem Realization AIgorithm - ERA. The study of the fuzzy controllers performance is accomplished through simulation and experimental analysis. Two fuzzy models are used for the controller: the Mamdani and the Takagi- Sugeno-Kang. The results show that both fuzzy controllers are quite efficient in the control of flutter.

M. Tech. Electronic Engineering. / Discusses the main objective of this research to develop a robust image-based algorithm to detect and determine the orientation of small helipad using shape descriptors and associated pre-processing techniques.

A multidisciplinary design optimization framework is developed that integrates control system design with aerostructural design for a highly-deformable wing. The objective of this framework is to surpass the existing aircraft endurance limits through the use of an active load alleviation system designed concurrently with the rest of the aircraft. The novelty of this work is two fold. First, a unified dynamics framework is developed to represent the full six-degree-of-freedom rigid-body along with the structural dynamics. It allows for an integrated control design to account for both manoeuvrability (flying quality) and aeroelasticity criteria simultaneously. Secondly, by synthesizing the aircraft control system along with the structural sizing and aerodynamic shape design, the final design has the potential to exploit synergies among the three disciplines and yield higher performing aircraft. A co-rotational structural framework featuring Euler--Bernoulli beam elements is developed to capture the wing's nonlinear deformations under the effect of aerodynamic and inertial loadings. In this work, a three-dimensional aerodynamic panel code, capable of calculating both steady and unsteady loadings is used. Two different control methods, a model predictive controller (MPC) and a 2-DOF mixed-norm robust controller, are considered in this work to control a highly flexible aircraft. Both control techniques offer unique advantages that make them promising for controlling a highly flexible aircraft. The control system works towards executing time-dependent manoeuvres along with performing gust/manoeuvre load alleviation. The developed framework is investigated for demonstration in two design cases: one in which the control system simply worked towards achieving or maintaining a target altitude, and another where the control system is also performing load alleviation. The use of the active load alleviation system results in a significant improvement in the aircraft performance relative to the optimum result without load alleviation. The results show that the inclusion of control system discipline along with other disciplines at early stages of aircraft design improves aircraft performance. It is also shown that structural stresses due to gust excitations can be better controlled by the use of active structural control systems which can improve the fatigue life of the structure.

A multidisciplinary design optimization framework is developed that integrates control system design with aerostructural design for a highly-deformable wing. The objective of this framework is to surpass the existing aircraft endurance limits through the use of an active load alleviation system designed concurrently with the rest of the aircraft. The novelty of this work is two fold. First, a unified dynamics framework is developed to represent the full six-degree-of-freedom rigid-body along with the structural dynamics. It allows for an integrated control design to account for both manoeuvrability (flying quality) and aeroelasticity criteria simultaneously. Secondly, by synthesizing the aircraft control system along with the structural sizing and aerodynamic shape design, the final design has the potential to exploit synergies among the three disciplines and yield higher performing aircraft. A co-rotational structural framework featuring Euler--Bernoulli beam elements is developed to capture the wing's nonlinear deformations under the effect of aerodynamic and inertial loadings. In this work, a three-dimensional aerodynamic panel code, capable of calculating both steady and unsteady loadings is used. Two different control methods, a model predictive controller (MPC) and a 2-DOF mixed-norm robust controller, are considered in this work to control a highly flexible aircraft. Both control techniques offer unique advantages that make them promising for controlling a highly flexible aircraft. The control system works towards executing time-dependent manoeuvres along with performing gust/manoeuvre load alleviation. The developed framework is investigated for demonstration in two design cases: one in which the control system simply worked towards achieving or maintaining a target altitude, and another where the control system is also performing load alleviation. The use of the active load alleviation system results in a significant improvement in the aircraft performance relative to the optimum result without load alleviation. The results show that the inclusion of control system discipline along with other disciplines at early stages of aircraft design improves aircraft performance. It is also shown that structural stresses due to gust excitations can be better controlled by the use of active structural control systems which can improve the fatigue life of the structure.

Aircraft structures have varying stiffness levels making them flexible. Consequently, this elastic property becomes increasingly important at high speeds affecting the flight dynamics of the aircraft. In high speed aircraft such as the F/A-18, elastic structural properties must be accounted for to ensure confidence in predicted flight dynamics in order to avoid adverse aeroelastic phenomena throughout flight. Data from the F/A-18 Active Aeroelastic Wing (AAW) program was used to create aeroservoelastic (ASE) models at varying flight conditions. The discretized ASE models were integrated into the NASA Dryden F/A-18 simulator in parallel with the traditional 6-DOF (degrees-of-freedom) flight dynamics calculations to ensure minimal disruption to the existing operating framework of the simulator. An interpolation scheme was used to construct ASE models within the known flight condition models. Data was processed through the state-space ASE models to compute the elastic effects during flight. Total flight dynamics from the simulation were analyzed and showed expected behavior for the combined elastic and rigid-body components in flight.

Neste trabalho é desenvolvido um sistema de controle ativo para supressão de flutter de uma asa utilizando-se sensores de pressão em pontos estratégicos de sua superfície. O flutter é um fenômeno aeroelástico que caracteriza um acoplamento instável entre estrutura flexível e escoamento aerodinâmico não estacionário. Quando a modificação da estrutura ou da aerodinâmica da asa não é viável, o uso de sistemas de controle passa a ser uma boa opção. Para o desenvolvimento do sistema de controle proposto, é primeiramente desenvolvido um modelo numérico de asa flexível. Com esse modelo numérico e a pressão na superfície da asa medida em certos pontos e realimentada ao sistema controlador, são determinadas correções no ângulo de uma superfície de controle no bordo de fuga. A tentativa de se utilizar um sistema de controle bem simples, com o uso de um único sensor de pressão, mostra a viabilidade de se implementar um sistema deste tipo em aeronaves reais. Esse sistema pode tornar-se uma alternativa aos desenvolvidos até então com o uso de acelerômetros, além de ser útil em sistemas onde se procura prever o estol e observar o comportamento da distribuição de pressão sobre a asa em vôo. / In this work, a wing flutter suppression active control system using pressure sensors in strategic points is developed. Flutter is an aeroelastic phenomenon characterized by an unstable coupling of a flexible structure and a non-stationary aerodynamic flow. When changes of the wing structure or of the aerodynamics are not viable, the use of automatic control systems becomes a good option. For the developing of the suggested control system, a numeric model of a finite flexible wing is firstly done. With this model and the pressure over the wing surface read in certain points and fedback to the control system, changes of the control surface angle on the trailing edge are determined. The attempt to use a simple control system, with a unique pressure sensor shows the viability of implanting this kind of system in real aircrafts. This system may become an alternative to those developed until now, using accelerometers. Yet, it could be useful, in systems where it is necessary to predict stall and observe the pressure load behavior over the wing in flight.

The work in this dissertation comprises aeroservoelastic simulation development, two modal filter design case studies and theoretical improvement of the modal filter. The modal filter is made robust to sensor bias. Studies have shown that the states estimated by the modal filter can be integrated into active structural control. The integration of modal filters into aircraft structural control systems is explored. Modal filters require distributed sensing to achieve accurate modal coordinate estimates. Distributed sensing technology has progressed to the point, where it is being tested on aircraft such as Ikhana and the upcoming X-56A. Previously, the modal filter was criticized for requiring too many sensors. It was never assessed for its potential benefits in aircraft control. Therefore it is of practical interest to reinvestigate the modal filter. The first case study shows that under conditions of sensor normality, the modal filter is a Gaussian efficient estimator in an aeroservoelastic environment. This is a fundamental experiment considering the fact that the modal filter has never been tested in the airflow. To perform this case study a linear aeroservoelastic code capable of modeling distributed sensing is developed and experimentally validated. From this code, a computational wing model is fitted with distributed sensing. A modal filtering design methodology is developed and applied. With distributed sensing and modal filtering feedback control is achieved. This is also compared and contrasted with a controller using state-of-the-art accelerometers. In addition, new methods of active shape control are introduced for warping an aeroelastic structure utilizing the modal filter and control surfaces. The next case study takes place in a realistic setting for an aircraft. Flexible aircraft bring challenges to the active control community. Increased gust loads, possibility of flutter, and off-design drag may detrimentally affect performance and safety. Aeroservoelastic tailoring, gust load alleviation (GLA) and active flutter suppression (AFS) may be required on future flexible air vehicles. It is found that modal filters can theoretically support these systems. The aircraft case study identifies additional steps required in the modal filtering design methodology. Distributed sensing, the modal filter and modal reference shape control are demonstrated on the X-56A flutter-unstable simulation model. It is shown that control of deformations at potentially millions of points on an aircraft vehicle can be achieved through control of a few modal coordinates. Finally modal filter robustness is theoretically improved and computationally verified. State-of-the-art modal filters have high bias sensitivity. In fact, this is so critical that state-of-the-art modal filters may never be certified for aircraft implementation. This is especially true within a flight critical control system. The solution to this problem is found through derivation of the robust modal filter. The filter combines good properties of concentration algorithms with robust re-descending M-estimation. A new trim criterion specific to the strain based modal sensing system is derived making the filter robust to asymmetric or leverage point outliers. Robust starts are introduced to improve convergence of the modal estimation system to the globally optimal solution in the presence of 100s of biased fiber optic sensors.

Neste trabalho é desenvolvido um sistema de controle ativo para supressão de flutter de uma asa utilizando-se sensores de pressão em pontos estratégicos de sua superfície. O flutter é um fenômeno aeroelástico que caracteriza um acoplamento instável entre estrutura flexível e escoamento aerodinâmico não estacionário. Quando a modificação da estrutura ou da aerodinâmica da asa não é viável, o uso de sistemas de controle passa a ser uma boa opção. Para o desenvolvimento do sistema de controle proposto, é primeiramente desenvolvido um modelo numérico de asa flexível. Com esse modelo numérico e a pressão na superfície da asa medida em certos pontos e realimentada ao sistema controlador, são determinadas correções no ângulo de uma superfície de controle no bordo de fuga. A tentativa de se utilizar um sistema de controle bem simples, com o uso de um único sensor de pressão, mostra a viabilidade de se implementar um sistema deste tipo em aeronaves reais. Esse sistema pode tornar-se uma alternativa aos desenvolvidos até então com o uso de acelerômetros, além de ser útil em sistemas onde se procura prever o estol e observar o comportamento da distribuição de pressão sobre a asa em vôo. / In this work, a wing flutter suppression active control system using pressure sensors in strategic points is developed. Flutter is an aeroelastic phenomenon characterized by an unstable coupling of a flexible structure and a non-stationary aerodynamic flow. When changes of the wing structure or of the aerodynamics are not viable, the use of automatic control systems becomes a good option. For the developing of the suggested control system, a numeric model of a finite flexible wing is firstly done. With this model and the pressure over the wing surface read in certain points and fedback to the control system, changes of the control surface angle on the trailing edge are determined. The attempt to use a simple control system, with a unique pressure sensor shows the viability of implanting this kind of system in real aircrafts. This system may become an alternative to those developed until now, using accelerometers. Yet, it could be useful, in systems where it is necessary to predict stall and observe the pressure load behavior over the wing in flight.