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The purpose of this dissertation is to develop a methodology for designing active control laws in a computational aeroelasticity environment. The methodology involves employing a systems identification technique to develop an explicit state-space model for control law design from the output of a computational aeroelasticity code. The particular computational aeroelasticity code employed in this dissertation solves the transonic small disturbance equation using a time-accurate, finite-difference scheme. Linear structural dynamics equations are integrated simultaneously with the computational fluid dynamics equations to determine the time responses of the structural outputs. These structural outputs are employed as the input to a modern systems identification technique that determines the Markov parameters of an "equivalent linear system". The eigensystem realization algorithm is then employed to develop an explicit state-space model of the equivalent linear system. Although there are many control law design techniques available, the standard Linear Quadratic Guassian technique is employed in this dissertation. The computational aeroelasticity code is modified to accept control laws and perform closed-loop simulations. Flutter control of a rectangular wing model is chosen to demonstrate the methodology. Various cases are used to illustrate the usefulness of the methodology as the nonlinearity of the computational fluid dynamics system is increased through increased angle-of-attack changes. / Ph. D.

With advanced subsonic transports and military aircraft operating in the transonic regime, it is becoming important to determine the effects of the coupling between aerodynamic loads and elastic forces. Since aeroelastic effects can contribute significantly to the design of these aircraft, there is a strong need in the aerospace industry to predict these aero-structure interactions computationally. To perform static aeroelastic analysis in the transonic regime, high fidelity computational fluid dynamics (CFD) analysis tools must be used in conjunction with high fidelity computational structural dynamics (CSD)analysis tools due to the nonlinear behavior of the aerodynamics in the transonic regime. There is also a need to be able to use a wide variety of CFD and CSD tools to predict these aeroelastic effects in the transonic regime. Because source codes are not always available, it is necessary to couple the CFD and CSD codes without alteration of the source codes. In this study, an aeroelastic coupling procedure is developed which will perform static aeroelastic analysis using any CFD and CSD code with little code integration. The aeroelastic coupling procedure is demonstrated on an F/A-18 Stabilator using NASTD (an in-house McDonnell Douglas CFD code)and NASTRAN. In addition, the Aeroelastic Research Wing (ARW-2) is used for demonstration of the aeroelastic coupling procedure by using ENSAERO (NASA Ames Research Center CFD code) and a finite element wing-box code (developed as a part of this research). The results obtained from the present study are compared with those available from an experimental study conducted at NASA Langley Research Center and a study conducted at NASA Ames Research Center using ENSAERO and modal superposition. The results compare well with experimental data. In addition, parallel computing power is used to investigate parallel static aeroelastic analysis because obtaining an aeroelastic solution using CFD/CSD methods is computationally intensive. A parallel finite element wing-box code is developed and coupled with an existing parallel Euler code to perform static aeroelastic analysis. A typical wing-body configuration is used to investigate the applicability of parallel computing to this analysis. Performance of the parallel aeroelastic analysis is shown to be poor; however with advances being made in the arena of parallel computing, there is definitely a need to continue research in this area. / Ph. D.

A method for predicting the unsteady, subsonic, aeroservoelastic response of a wing has been developed. The air, wing, and control surface are considered to be a single dynamical system. All equations are solved simultaneously in the time domain by a predictor-corrector method. The scheme allows nonlinear aerodynamic and structural models to be used and subcritical, critical, and supercritical aeroelastic behavior may be modeled without restrictions to small disturbances or periodic motions. A vortex-lattice method is used to model the aerodynamics. This method accounts for nonlinear effects associated with high angles of attack, unsteady behavior, and deformations of the wing. The vortex-lattice method is valid as long as separation or vortex bursting does not occur. Two structural models have been employed: a linear model and a nonlinear model which accounts for large curvature. Both models consider the flexural-torsional motion of an inextensional wing. / Master of Science

A method for predicting unsteady, subsonic aeroelastic responses has been developed. The technique accounts for aerodynamic nonlinearities associated with angles of attack, vortex-dominated flow, static deformations, and unsteady behavior. The angle of attack is limited only by the occurrence of stall or vortex bursting near the wing. The fluid and the wing together are treated as a single dynamical system, and the equations of motion for the structure and flowfield are integrated simultaneously and interactively in the time domain. The method employs an iterative scheme based on a predictor-corrector technique. The aerodynamic loads are computed by the general unsteady vortex-lattice method and are determined simultaneously with the motion of the wing. Because the unsteady vortex-lattice method predicts the wake as part of the solution, the history of the motion is taken into account; hysteresis is predicted. Two models are used to demonstrate the technique: a rigid wing on an elastic support experiencing plunge and pitch about the elastic axis, and an elastic wing rigidly supported at the root chord experiencing spanwise bending and twisting. The method can be readily extended to account for structural nonlinearities and/or substitute aerodynamic load models. The time domain solution coupled with the unsteady vortex-lattice method provides the capability of graphically depicting wing and wake motion. / Ph. D.

<p>High performance fighter aircraft such as the F-16 experience aeroelastic Limit Cycle Oscillations (LCO) when they carry certain combinations of under-wing stores. This `store-induced LCO' causes serious problems including airframe fatigue, pilot discomfort and loss of operational effectiveness. The usual response has been to restrict the stores carriage envelope based on flight test experience, and accept the accompanying reduction in mission performance.</p><p>Although several nonlinear mechanisms - structural as well as aerodynamic, have been proposed to explain the LCO phenomenon, their roles are not well understood. Consequently, existing models are unable to predict accurately AND reliably the most critical LCO properties, namely onset speed and response level. On the other hand, the more accurate Computational Fluid Dynamics (CFD) based time marching methodology yields results at much greater expense and time. Clearly, there is a critical need to establish methods that are more rapid while providing accurate predictions more in line with flight test results than at present. Such a capability will also aid in future aircraft design and usage.</p><p>This work was undertaken to develop a better understanding of nonlinear aeroelastic phenomena, and their relation to classical flutter and divergence, with a particular focus on store-induced LCO in high performance fighter aircraft. The following systems were studied: (1) a `simple' wing with a flexible and nonlinear root attachment, (2) a `generic' wing with a flexible and nonlinear wing-store attachment and (3) the F-16 aircraft, again with nonlinear wing-store attachments.</p><p>While structural nonlinearity was present in all cases, steady flow aerodynamic nonlinearity was also included in the F-16 case by the use of a Computational Fluid Dynamics model based on the Reynolds Averaged Navier Stokes (RANS) equations. However, dynamic linearization of the CFD model was done for the present computations. The computationally efficient Harmonic Balance (HB) nonlinear solution technique was a key component of this work, with time marching simulations and closed form solutions being used selectively to confirm the findings of the HB solutions. The simple wing and the generic wing were both modeled as linear beam-rods whose displacements were represented using the primitive modes method. The wing aerodynamic model was linear (quasi-steady for the simple wing and based on the Vortex Lattice Method for the generic wing), and the store aerodynamics were omitted.</p><p>The presence of a cubic restoring force (of hardening or softening type, in stiffness or in damping) at the root of the simple wing led to several interesting results and insights. Next, various nonlinear mechanisms including cubic restoring force, freeplay and friction were introduced at the wing-store attachment of the generic wing and these led to a still greater variety in behavior. General relationships were established between the type of nonlinearity and the nature of the resulting response, and they proved very useful for tailoring the F-16 study and interpreting its results.</p><p>The Air Force Seek Eagle Office/Air Force Research Laboratory provided a modal structural model of an LCO-prone store configuration of the F-16 aircraft with stores included. In order to investigate a range of stores attachment configurations, the analysis required modification of the stiffness and damping of the wing-store attachment. Since the Finite Element model of the wing and store structure was not available, the modification was achieved by subtracting the store and adding it back with the necessary changes to the store or attachment using a dynamic decoupling/coupling technique. The modified models were subjected to flutter/LCO analysis using the Duke Harmonic Balance CFD RANS solver, and the resulting flutter boundaries were used in combination with the HB method to derive LCO responses due to the wing-store attachment nonlinearity.</p><p>Comparisons were made between the simulation results and the F-16 flight test LCO data. While multiple sources of nonlinearity are probably responsible for the wide range of observed LCO behavior, it was concluded that cubic softening stiffness and positive cubic damping were the more likely structural mechanisms causing LCO, in addition to nonlinear aerodynamics.</p> / Dissertation

Performance of inflatable wings was investigated through laboratory, wind tunnel and flight-testing. Three airfoils were investigated, an inflatable-rigidazable wing, an inflatable polyurethane wing and a fabric wing restraint with a polyurethane bladder. The inflatable wings developed and used within this research had a unique outer airfoil profile. The airfoil surface consisted of a series of chord-wise \bumps.andamp;quot; The effect of the bumps or \surface perturbationsandamp;quot; on the performance of the wings was of concern and was investigated through smoke-wire flow visualization. Aerodynamic measurements and predictions were made to determine the performance of the wings at varying chord based Reynolds Numbers and angles of attack. The inflatable baffes were found to introduce turbulence into the free-stream boundary layer, which delayed separation and improved performance. Another area of concern was aeroelasticity. The wings contain no solid structural members and thus rely exclusively on inflation pressure for stiffness. Inflation pressure was varied below the design pressure in order to examine the effect on wingtip twist and bending. This lead to investigations into wing deformation due to aerodynamic loading and an investigation of wing flutter. Photogrammetry and laser displacement sensors were used to determine the wing deflections. The inflatable wings exhibited wash-in deformation behavior. Alternately, as the wings do not contain structural members, the relationship between stiffness and inflation pressure was exploited to actively manipulate wing through wing warping. Several warping techniques were developed and employed within this re-search. The goal was to actively influence the shape of the inflatable wings to affect the flight dynamics of the vehicle employing them. Researchers have developed inflatable beam theory and models to analyze torsion and bending of inflatable beams and other inflatable structures. This research was used to model the inflatable wings to predict the performance of the inflatable wings during flight. Design elements of inflatable wings incorporated on the UAVs used within this research are also discussed. Finally, damage resistance of the inflatable wings is shown from results of flight tests.

Flutter é uma instabilidade aeroelástica dinâmica que envolve a interação de forças aerodinâmicas, elásticas e inerciais. Esta instabilidade pode ocorrer em superfícies de aeronaves, como asas, que apresentarão um comportamento oscilatório auto-sustentado e possíveis problemas estruturais se o mesmo não for suprimido. Um dos tipos clássicos de flutter envolve o acoplamento dos modos de vibrar de flexão e torção. Este tipo de flutter binário é conhecido como flutter flexo-torcional. Um dispositivo flexível é desenvolvido para testes de flutter flexo-torcional com asas rígidas em túnel de vento. O procedimento de projeto deste dispositivo flexível é baseado em simulações realizadas com um modelo em elementos finitos cujos resultados são testados em simulações realizadas com um modelo aeroelástico formulado para simular o comportamento aeroelástico do sistema experimental. Então, para se verificar os resultados analíticos, uma análise modal experimental é realizada e as freqüências e a forma dos modos são identificadas utilizando-se o Eigensystem Realization Algorithm. Depois disso, alguns testes em túnel de vento são realizados para a verificação da obtenção do flutter, para a caracterização do flutter e para a identificação do flutter. O desenvolvimento deste sistema experimental permite o estudo e a aplicação de leis de controle para a supressão ativa do flutter, que é o objetivo principal deste trabalho. Um controlador ativo para supressão de flutter através de realimentação de estados é projetado a partir do modelo Aeroelástico previamente desenvolvido. Este controlador é inicialmente testado em simulação e, então, são realizados experimentos em túnel de vento. O objetivo é suprimir o flutter e manter a estabilidade do sistema em malha fechada. O modelo para testes no túnel de vento é uma asa rígida retangular com perfil NACA 0012 com uma superfície de controle no bordo de fuga utilizada como atuador. / 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, which will present a self-sustained oscillatory behaviour and possible structural problems if not suppressed. One of the classical types of flutter involves the coupling of bending and torsion vibration modes. This binary type of flutter is known as flexural-torsional flutter. A flexible mount system is developed for flexural-torsional flutter tests with rigid wings in wind tunnels. The design procedure of this mount system is based in simulations performed with a finite element model which results are tested in simulations performed with an Aeroelastic model formulated to simulate the aeroelastic behaviour of the experimental system. Then, to verify the analytical results, an experimental modal analysis is performed and mode shapes and frequencies are identified using the Eigensystem Realization Algorithm. After this, some wind tunnel tests are performed to verify flutter achievement, for flutter characterization and for flutter identification. The development of this experimental system allows the study and application of active control laws for active flutter suppression, which is the main goal of this work. A state feedback controller for active flutter suppression is designed using the aeroelastic model previously developed. This controller is initially tested in simulations and, then, wind-tunnel experiments are performed. The goal is to suppress flutter and to maintain the stability of the closed loop system. The wind tunnel model is a rigid rectangular wing with a NACA 0012 airfoil section with a trailing edge control surface used as actuator.

Flutter é uma instabilidade aeroelástica dinâmica que envolve a interação de forças aerodinâmicas, elásticas e inerciais. Esta instabilidade pode ocorrer em superfícies de aeronaves, como asas, que apresentarão um comportamento oscilatório auto-sustentado e possíveis problemas estruturais se o mesmo não for suprimido. Um dos tipos clássicos de flutter envolve o acoplamento dos modos de vibrar de flexão e torção. Este tipo de flutter binário é conhecido como flutter flexo-torcional. Um dispositivo flexível é desenvolvido para testes de flutter flexo-torcional com asas rígidas em túnel de vento. O procedimento de projeto deste dispositivo flexível é baseado em simulações realizadas com um modelo em elementos finitos cujos resultados são testados em simulações realizadas com um modelo aeroelástico formulado para simular o comportamento aeroelástico do sistema experimental. Então, para se verificar os resultados analíticos, uma análise modal experimental é realizada e as freqüências e a forma dos modos são identificadas utilizando-se o Eigensystem Realization Algorithm. Depois disso, alguns testes em túnel de vento são realizados para a verificação da obtenção do flutter, para a caracterização do flutter e para a identificação do flutter. O desenvolvimento deste sistema experimental permite o estudo e a aplicação de leis de controle para a supressão ativa do flutter, que é o objetivo principal deste trabalho. Um controlador ativo para supressão de flutter através de realimentação de estados é projetado a partir do modelo Aeroelástico previamente desenvolvido. Este controlador é inicialmente testado em simulação e, então, são realizados experimentos em túnel de vento. O objetivo é suprimir o flutter e manter a estabilidade do sistema em malha fechada. O modelo para testes no túnel de vento é uma asa rígida retangular com perfil NACA 0012 com uma superfície de controle no bordo de fuga utilizada como atuador. / 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, which will present a self-sustained oscillatory behaviour and possible structural problems if not suppressed. One of the classical types of flutter involves the coupling of bending and torsion vibration modes. This binary type of flutter is known as flexural-torsional flutter. A flexible mount system is developed for flexural-torsional flutter tests with rigid wings in wind tunnels. The design procedure of this mount system is based in simulations performed with a finite element model which results are tested in simulations performed with an Aeroelastic model formulated to simulate the aeroelastic behaviour of the experimental system. Then, to verify the analytical results, an experimental modal analysis is performed and mode shapes and frequencies are identified using the Eigensystem Realization Algorithm. After this, some wind tunnel tests are performed to verify flutter achievement, for flutter characterization and for flutter identification. The development of this experimental system allows the study and application of active control laws for active flutter suppression, which is the main goal of this work. A state feedback controller for active flutter suppression is designed using the aeroelastic model previously developed. This controller is initially tested in simulations and, then, wind-tunnel experiments are performed. The goal is to suppress flutter and to maintain the stability of the closed loop system. The wind tunnel model is a rigid rectangular wing with a NACA 0012 airfoil section with a trailing edge control surface used as actuator.