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The first two steps in the development of an integrated multidisciplinary design optimization procedure capable of analyzing the nonlinear fluid flow about geometrically complex aeroelastic configurations have been accomplished in the present work. For the first step, a three-dimensional unstructured grid approach to aerodynamic shape sensitivity analysis and design optimization has been developed. The advantage of unstructured grids, when compared with a structured-grid approach, is their inherent ability to discretize irregularly shaped domains with greater efficiency and less effort. Hence, this approach is ideally suited fro geometrically complex configurations of practical interest. In this work the time-dependent, nonlinear Euler equations are solved using an upwind, cell-centered, finite-volume scheme. The discrete, linearized systems which result from this scheme are solved iteratively by a preconditioned conjugate-gradient-like algorithm known as GMRES for the two-dimensional cases and a Gauss-Seidel algorithm for the three-dimensional; at steady-state, similar procedures are used to solve the accompanying linear aerodynamic sensitivity equations in incremental iterative form. As shown, this particular form of the sensitivity equation makes large-scale gradient-based aerodynamic optimization possible by taking advantage of memory efficient methods to construct exact Jacobian matrix-vector products. Various surface parameterization techniques have been employed in the current study to control the shape of the design surface. Once this surface has been deformed, the interior volume of the unstructured grid is adapted by considering the mesh as a system of interconnected tension springs. Grid sensitivities are obtained by differentiating the surface parameterization and the grid adaptation algorithms with ADIFOR, an advanced automatic-differentiation software tool. To demonstrate the ability of this procedure to analyze and design complex configurations of practical interest, the sensitivity analysis and shape optimization has been performed for several two- and three-dimensional cases. In two-dimensions, an initially symmetric NACA-0012 airfoil and a high-lift multi-element airfoil were examined. For the three-dimensional configurations, an initially rectangular wing with uniform NACA-0012 cross-sections was optimized; in additions, a complete Boeing 747-200 aircraft was studied. Furthermore, the current study also examines the effect of inconsistency in the order of spatial accuracy between the nonlinear fluid and linear shape sensitivity equations. The second step was to develop a computationally efficient, high-fidelity, integrated static aeroelastic analysis procedure. To accomplish this, a structural analysis code was coupled with the aforementioned unstructured grid aerodynamic analysis solver. The use of an unstructured grid scheme for the aerodynamic analysis enhances the interactions compatibility with the wing structure. The structural analysis utilizes finite elements to model the wing so that accurate structural deflections may be obtained. In the current work, parameters have been introduced to control the interaction of the computational fluid dynamics and structural analyses; these control parameters permit extremely efficient static aeroelastic computations. To demonstrate and evaluate this procedure, static aeroelastic analysis results for a flexible wing in low subsonic, high subsonic (subcritical), transonic (supercritical), and supersonic flow conditions are presented. / Ph. D.

The aviation industry is growing at a steady rate but presently, the industry is highly dependent on fossil fuel. As the world is running out of fossil fuels and the wide-spread acceptance of climate change due to carbon emissions, both the governments and industry are spending a significant amount of resources on research to reduce the weight and hence the fuel consumption of commercial aircraft. A commercial fixed-wing aircraft wing consists of spars which are beams running in span-wise direction, carrying the flight loads and ribs which are panels with holes attached to the spars to preserve the outer airfoil shape of the wing. Kapania et al. at Virginia Tech proposed the concept of reducing the weight of aircraft wing using unconventional design of the internal structure consisting of curvilinear spars and ribs (known as SpaRibs) for enhanced performance. A research code, EBF3GLWingOpt, was developed by the Kapania Group. at Virginia Tech to find the best configuration of SpaRibs in terms of weight saving for given flight conditions. However, this software had a number of limitations and it can only create and analyze limited number of SpaRibs configurations. In this work, the limitations of the EBF3GLWingOpt code has been identified and new algorithms have been developed to make is robust and analyze larger number of SpaRibs configurations. The code also has the capability to create cut-outs in the SpaRibs for passage of fuel pipes and wirings. This new version of the code can be used to find best SpaRibs configuration for multiple objectives such as reduction of weight and increase flutter velocity. The code is developed in Python language and it has parallel computational capabilities. The wing is modeled using commercial FEA software, MSC.PATRAN and analyzed using MSC.NASTRAN which are from within EBF3GLWingOpt. Using this code a significant weight reduction for a transport aircraft wing has been achieved. / PHD

A identificação de mecanismos modais é uma tarefa que requer um grande esforço ao se considerar geometrias complexas. O uso de materiais inteligentes como tecnologia nesse tipo de identificação vem sendo bastante difundido, principalmente o uso de sensores piezelétricos, como o piezo-fiber composite (PFC). Esse tipo de aplicação pode se tornar uma ferramenta bastante prática no estudo de instabilidades aeroelásticas, em especial o mecanismo modal de flutter. A proposta desse trabalho é criar uma metodologia de análise de flutter simulando o desempenho de materiais piezelétricos, aderidos em laminados compósitos, como sensores modais. Inicialmente, é realizada uma análise aeroelástica da estrutura para se identificar o mecanismo e os modos dominantes para o surgimento do flutter. Em seguida, os modos identificados são detectados pelos sensores com uma determinada potência de sinal. A sensibilidade desse sinal é avaliada de acordo com a posição e configuração do laminado embebido no sensor. Para realizar essa simulação, um modelo de asa é gerado e suas frequências naturais e modos são determinados pelo método dos elementos finitos (MEF). Com esses dados, é possível caracterizar o modelo nas equações de movimento aeroelásticas. O carregamento aerodinâmico dessas equações é obtido utilizando o método dos anéis de vórtice, do inglês: vortex lattice method (VLM). A simulação é realizada em cada velocidade de fluxo e a resposta dos sensores piezelétricos é obtida no domínio do tempo e domínio da freqüência para se analisar a potência do sinal. Foi realizada uma prévia análise de um modelo de asa representado por uma placa e as configurações de maior potência de sinal são identificadas. A posição dos sensores se demonstrou mais sensível do que a configuração do laminado e a utilização de apenas um sensor foi suficiente para identificação do mecanismo modal, o que pode tornar essa tecnologia viável em ensaios de flutter em estruturas de material compósito. / For complex aeronautical structures, modal mechanism identification requires a great deal of effort. The use of smart materials has been developed in this application, mainly the sensor application with piezo-fiber composites (PFC). It can become a useful tool in aeroelastic instabilities studies, especially on flutter modal mechanism. This work intends to develop a methodology of flutter analysis evaluating the piezoelectric materials performance, using composites impregnation effects, and working as a modal sensor. First, one aeroelastic analysis is done to identify the flutter mechanism and its dominant modes. Then, it modes is detected by sensors with some specific power of electric signal, whose sensitivity is evaluated according with position and embeeded laminate configuration. This simulation uses a plate model representing a wing, whose natural frequencies and modes are determined by finite element method (FEM). So, given this data, is possible to define the wing model using an equation of motion, whose aerodynamic load is obtained by vortex lattice method (VLM). That equation is solved step by step, for each airspeed considered, then, the PFC response is obtained both in the frequency and time domain. The analysis was done using a metric that qualifies the best configuration according with the power of signal. The sensor position was more significant than the laminate configuration; however, the use of only one sensor is sufficient to identify the modal mechanism, which becomes this technology feasible in flutter test of composite structures.

A identificação de mecanismos modais é uma tarefa que requer um grande esforço ao se considerar geometrias complexas. O uso de materiais inteligentes como tecnologia nesse tipo de identificação vem sendo bastante difundido, principalmente o uso de sensores piezelétricos, como o piezo-fiber composite (PFC). Esse tipo de aplicação pode se tornar uma ferramenta bastante prática no estudo de instabilidades aeroelásticas, em especial o mecanismo modal de flutter. A proposta desse trabalho é criar uma metodologia de análise de flutter simulando o desempenho de materiais piezelétricos, aderidos em laminados compósitos, como sensores modais. Inicialmente, é realizada uma análise aeroelástica da estrutura para se identificar o mecanismo e os modos dominantes para o surgimento do flutter. Em seguida, os modos identificados são detectados pelos sensores com uma determinada potência de sinal. A sensibilidade desse sinal é avaliada de acordo com a posição e configuração do laminado embebido no sensor. Para realizar essa simulação, um modelo de asa é gerado e suas frequências naturais e modos são determinados pelo método dos elementos finitos (MEF). Com esses dados, é possível caracterizar o modelo nas equações de movimento aeroelásticas. O carregamento aerodinâmico dessas equações é obtido utilizando o método dos anéis de vórtice, do inglês: vortex lattice method (VLM). A simulação é realizada em cada velocidade de fluxo e a resposta dos sensores piezelétricos é obtida no domínio do tempo e domínio da freqüência para se analisar a potência do sinal. Foi realizada uma prévia análise de um modelo de asa representado por uma placa e as configurações de maior potência de sinal são identificadas. A posição dos sensores se demonstrou mais sensível do que a configuração do laminado e a utilização de apenas um sensor foi suficiente para identificação do mecanismo modal, o que pode tornar essa tecnologia viável em ensaios de flutter em estruturas de material compósito. / For complex aeronautical structures, modal mechanism identification requires a great deal of effort. The use of smart materials has been developed in this application, mainly the sensor application with piezo-fiber composites (PFC). It can become a useful tool in aeroelastic instabilities studies, especially on flutter modal mechanism. This work intends to develop a methodology of flutter analysis evaluating the piezoelectric materials performance, using composites impregnation effects, and working as a modal sensor. First, one aeroelastic analysis is done to identify the flutter mechanism and its dominant modes. Then, it modes is detected by sensors with some specific power of electric signal, whose sensitivity is evaluated according with position and embeeded laminate configuration. This simulation uses a plate model representing a wing, whose natural frequencies and modes are determined by finite element method (FEM). So, given this data, is possible to define the wing model using an equation of motion, whose aerodynamic load is obtained by vortex lattice method (VLM). That equation is solved step by step, for each airspeed considered, then, the PFC response is obtained both in the frequency and time domain. The analysis was done using a metric that qualifies the best configuration according with the power of signal. The sensor position was more significant than the laminate configuration; however, the use of only one sensor is sufficient to identify the modal mechanism, which becomes this technology feasible in flutter test of composite structures.

Incorporating uncertainties in the aeroelastic analysis increases the confidence levels of computational predictions and reduces the need for validation with experimental or flight test data. Helicopter rotor blades, which play a dominant role in the overall vehicle performance, are routinely made of composites. The material properties of composites are uncertain because of the variations in manufacturing process and other effects while in service, maintenance and storage. Though nominal values are listed, they are seldom accurate. In this thesis, the effect of uncertainty in composite material properties on the computational predictions of cross-sectional properties, natural frequencies, blade tip deflections, vibratory loads and aeroelastic stability of a four-bladed composite helicopter rotor is studied. The effect of material uncertainty is studied with the composite rotor blades modeled as components of soft-inplane as well as stiff-inplane hingeless helicopter rotors. Aeroelastic analysis based on finite elements in space and time is used to evaluate the helicopter rotor blade response in hover and forward flight. Uncertainty analysis is performed with direct Monte Carlo simulations based on a sufficient number of random samples of material properties. It is found that the cross-sectional stiffness parameters and natural frequencies of rotor blades show considerable scatter from their baseline predictions. The uncertainty impact on the rotating natural frequencies depends on the level of centrifugal stiffening of each mode. The propagation of material uncertainty into aeroelastic response causes large deviations from the baseline predictions. The magnitudes of 4/rev vibratory loads show deviations of 10 to 600 percent from their baseline predictions. The aeroelastic stability in hover and forward flight conditions also show considerable uncertainty in the predictions. In addition to the effects of material uncertainty, various factors influencing the propagation of material uncertainty are studied with the first-order based reliability methods. The numerical results have shown the need to consider the uncertainties in the helicopter aeroelastic analysis for reliable computational predictions. Uncertainty quantification using direct Monte Carlo simulation is accurate but computationally expensive. The application of response surface methodologies to reduce the computational cost of uncertainty analysis is studied. Response surface approximations of aeroelastic outputs are developed in terms of the composite material properties. Monte Carlo simulations are then performed using these computationally less expensive response surface models. The results of this study show that the metamodeling techniques can effectively reduce the computational cost of uncertainty analysis of composite rotor blades. In the last part of the thesis, an aeroelastic optimization method to minimize the vibration level is developed with due consideration to material uncertainty. Second-order polynomial response surfaces are used to approximate the objective function which smooths out the local minima or numerical noise in the design space. The aeroelastic optimization is carried out with the nominal values of composite material properties and the performance of final design is found to be optimum even for the perturbed values of material properties.

In this thesis, computational aerodynamics, structural and aeroelastic analyses of the composite tactical unmanned air vehicle which is designed and manufactured in the Department of Aerospace Engineering are performed. Verification of the structural integrity of the air vehicle is shown at the minimum maneuvering and the dive speeds at the static limit loads which are calculated by the computational aerodynamics analysis of the full aircraft model. In the current work, aerodynamic loads are re-calculated for more accurately determined dive speed angle of attack in an effort to match the overall vertical pressure load more closely to the half of the aircraft weight at the positive load factor. Finite element models of the fuselage, wing and the vertical-horizontal tail plane are prepared including the filament wound boom connecting the wing and the tail plane. Structural analyses of the composite wing, vertical and horizontal tail plane are performed under the limit aerodynamic loads calculated at the corner points of the V-N diagram using the structural finite element model of the wing-tail plane combination only. Global finite element analysis of the wing-tail plane combination showed that composite and isotropic materials of the wing-tail plane combination have positive margins of safety. Woven carbon and E-glass fabric that was procured to be used for the serial production version of the airplane are characterized for the tensile properties by the tests. Comprehensive aeroelastic stability analyses of the airplane are conducted by adding one sub-structure at a time to the aeroelastic model. Specifically, aeroelastic models which are used are the wing only, wing-tail plane combination, complete air vehicle with and without wing control surfaces. With such a study it is intended to address the effect each sub-structure adds to the aeroelastic model on the critical aeroelastic stability modes and speeds, and to see how sensitive the aeroelastic stability modes and speeds are to model fidelity. Detailed structural and aeroelastic analyses showed that the airplane has sufficient structural integrity under the action of static limit loads, and no aeroelastic instability is expected to occur within the flight envelope of the airplane.

Helicopter aeroelastic analysis is highly complex and multidisciplinary in nature; the flexibility of main rotor blades is coupled with aerodynamics, dynamics and control systems. A key component of an aeroelastic analysis is the vehicle trim procedure. Trim requires calculation of the main rotor and tail rotor controls and the vehicle attitude which cause the six steady forces and moments about the helicopter center of gravity to be zero. Trim simulates steady level flight of the helicopter. The trim equations are six nonlinear equations which depend on blade response and aerodynamic forcing through finite element analysis. Simulating the behavior of the helicopter in flight requires the solution of this system of nonlinear algebraic equations with unknowns being pilot controls and vehicle attitude angles. The nonlinear solution procedure is prone to slow convergence and occasional divergence causing problems in optimization and stochastic simulation studies. In this thesis, an attempt is made to efficiently solve the nonlinear equations involved in helicopter trim. Typically, nonlinear equations in mathematical physics and engineering are solved by linearizing the equations and forming various iterative procedures, then executing the numerical simulation. Helicopter aeroelasticity involves the solution of systems of nonlinear equations in a computationally expensive environment. The Newton method is typically used for the solution of these equations. Due to the expensive nature of each aeroelastic analysis iteration, Jacobian calculation at each iteration for the Newton method is not feasible for the trim problems. Thus, the Jacobian is calculated only once about the initial trim estimate and held constant thereafter. However, Jacobian modifications and updates can improve the performance of the Newton method. A comparative study is done in this thesis by incorporating different Jacobian update methods and selecting appropriate damping schemes for solving the nonlinear equations in helicopter trim. A modified Newton method with varying damping factor, Broyden rank-1 update and BFGS rank-2 update are explored using the Jacobian calculated at the initial guess. An efficient and robust approach for solving the strongly coupled nonlinear equations in helicopter trim based on the modified Newton method is developed. An appropriate initial estimate of the trim state is needed for successful helicopter trim. Typically, a guess from a simpler physical model such as a rigid blade analysis is used. However, it is interesting to study the impact of other starting points on the helicopter trim problem. In this work, an attempt is made to determine the control inputs that can have considerable effect on the convergence of trim solution in the aeroelastic analysis of helicopter rotors by investigating the basin of attraction of the nonlinear equations (set of initial guess points from which the nonlinear equations converge). It is illustrated that the three main rotor pitch controls of collective pitch, longitudinal cyclic pitch and lateral cyclic pitch have significant contribution to the convergence of the trim solution. Trajectories of the Newton iterates are shown and some ideas for accelerating the convergence of trim solution in the aeroelastic analysis of helicopter are proposed.

Submitted by Kilvya Braga (kilvyabraga@hotmail.com) on 2018-04-27T11:35:02Z No. of bitstreams: 1 DIEGO PAES DE ANDRADE PEÑA - DISSERTAÇÃO (PPGEM) 2016.pdf: 5093848 bytes, checksum: c6e79e54502ec5a0ff9ff5b410ffd362 (MD5) / Made available in DSpace on 2018-04-27T11:35:02Z (GMT). No. of bitstreams: 1 DIEGO PAES DE ANDRADE PEÑA - DISSERTAÇÃO (PPGEM) 2016.pdf: 5093848 bytes, checksum: c6e79e54502ec5a0ff9ff5b410ffd362 (MD5) Previous issue date: 2016-09-02 / CNPq / A aeroelasticidade é o campo da ciência que estuda a correlação entre as forças aerodinâmicas, elásticas e de inércia. Tal ciência é de grande importância no campo aeronáutico uma vez que as estruturas alares são flexíveis, devem suportar os esforços aerodinâmicos e serem rígidas o suficiente para garantir que esteja livre de todos os problemas aeroelásticos característicos (divergência, eficiência de controle, flutter e buffeting) dentro da faixa operacional de velocidades desenvolvida pela aeronave. Realizou-se uma análise modal da estrutura a fim de se conhecer os modos naturais de vibração e as respectivas frequências naturais. Para tal, utilizou-se o ANSYS Structural e o método dos elementos finitos, além de um estudo de malha para verificar a convergência dos resultados. Estudou-se também a influência da posição do lastro na ponta da placa plana, que causa a diminuição da segunda frequência natural. Além disso, realizou-se uma análise bidimensional de um volume de controle do tipo C-Grid, uma vez que o tamanho do volume de controle em uma análise aerodinâmica computacional é um fator extremamente importante. Com um volume de controle grande, tem se mais elementos na malha, caso o mesmo seja pequeno, as condições de contorno juntamente com os tamanhos dos elementos podem interferir nos resultados dos campos de velocidade e pressão em torno da estrutura. Nesse contexto, utilizou-se do software ANSYS Fluent para a simulação aerodinâmica da placa plana inclinada e obtenção dos coeficientes aerodinâmicos de sustentação e arrasto CL e CD. Os resultados foram comparados com resultados experimentais em túnel de vento de Goudeseune (SELIG; ROBERT; WILLIAMSON, 2011). Através do cálculo do Grid Convergence Index (GCI) e da comparação dos resultados numéricos com os dados experimentais constatou-se a convergência e conseguiu-se determinar um tamanho de volume de controle com erro baixo e aceitável. A análise fluido-estrutura acoplada de duas vias foi realizada com o ANSYS Structural para analisar a dinâmica estrutural através do método dos elementos finitos e o ANSYS CFX para resolver o campo do escoamento mediante método dos volumes finitos. Obtiveram-se o comportamento oscilatório da estrutura, além do coeficiente de amortecimento e tensões de von Mises. Analisando o comportamento transiente da dinâmica estrutural mediante um fluxo aerodinâmico constante (velocidade fixa). As simulações representaram bem o fenômeno, já que com o aumento da velocidade, o escoamento induz maior amortecimento à estrutura quando comparado com baixas velocidades. / The aeroelasticity is the field of science that studies the relationship between the aerodynamic elastic and inertia forces. Such knowledge is of great importance in the aviation field since the wing structures are flexible, must withstand the aerodynamic loads and be rigid enough to ensure that it is free from all aeroelastic problems like divergence, control efficiency, flutter and buffeting within the operating speed range. We carried out a modal analysis of the structure in order to know the natural vibration modes and natural frequencies. To this end, we used the ANSYS Structural with finite element method, a mesh study to verify the convergence of the results. It is also studied the influence of the slender body position of the tip of the flat plate, which causes the decrease of the second natural frequency. Furthermore, there was a twodimensional analysis of a volume control type C-Grid, since the control volume aerodynamic size in a computational analysis is an extremely important factor. A large volume of control has more elements in the mesh if it is small, the boundary conditions together with the sizes of elements may affect the results of the velocity field and pressure around the structure. In this context, we used the ANSYS FLUENT for the aerodynamic simulation of the inclined flat plate, and obtaining the aerodynamic support, and drag coefficients CL and CD. The results were compared with experimental results of Goudeseune wind tunnel (SELIG; ROBERT; WILLIAMSON, 2011). By calculating the Grid Convergence Index (GCI) and comparing the numerical results with experimental data found the convergence and managed to determine a control volume size with low and acceptable error. The fluid-structure coupled two-way analysis was performed using ANSYS Structural to analyze the structural dynamics through the finite element method and ANSYS CFX to resolve the flow field by the finite volume method. It was possible to obtain the oscillatory behavior of the structure, besides the damping coefficient and von Mises stresses. Analyzing the transient behavior of structural dynamics by a constant aerodynamic flow (fixed speed), the simulations represented the phenomenon as well, since with the increase in speed, the flow induces cushioning structure as compared to low speed

The unsteady aerodynamic and aeroelastic analysis of flapping flight under various kinematics and flow parameters is presented in this dissertation. The main motivation for this study arises from the challenges facing the development of micro air vehicles. Micro air vehicles by requirement are compact with dimensions less than 15-20 cm and flight speeds of around 10-15 m/s. These vehicles operate in low Reynolds number range of 10,000 to 100,000. At these low Reynolds numbers, the aerodynamic efficiency of conventional fixed airfoils significantly deteriorates. On the other hand, flapping flight employed by birds and insects whose flight regime coincides with that of micro air vehicles offers a viable alternate solution. For the analysis of flapping flight, a boundary fitted moving grid algorithm is implemented in a flow solver, GenIDLEST. The dynamic movement of the grid is achieved using a combination of spring analogy and trans-finite interpolation on displacements. The additional conservation equation of space required for moving grid is satisfied. The solver is validated with well known flow problems such as forced oscillation of a cylinder, a heaving airfoil, a moving indentation channel, and a hovering fruitfly. The performance of flapping flight is analyzed using Large Eddy Simulation (LES) for a wide range of Reynolds numbers and under various kinematic parameters. A spiral Leading Edge Vortex (LEV) forms during the downstroke due to the high angle of attack, which results in high force production. A strong spanwise flow of the order of the flapping velocity is observed along the core of the LEV. In addition, the formation of a negative spanwise flow is observed due to the tip vortex, which slows down the removal of vorticity from the LEV. This leads to the instability of the LEV at around mid-downstroke. Analysis with different rotation kinematics shows that a continuous rotation results in better propulsive efficiency as it generates thrust during the entire flapping cycle. Analysis with different angles of attack shows that a moderate angle of attack which results in complete shedding of the LEV offers high propulsive efficiency. The analysis of flapping flight at Reynolds numbers ranging from 100 to 100,000 shows that higher lift and thrust values are obtained for Re?100. The critical reasons are that at higher Reynolds numbers, the LEV is closer to the surface and as it sheds and convects it covers most of the upper surface. However, the Reynolds number has no or little effect on the lift and thrust as identical values are obtained for Re=10,000 and 100,000. The analysis with different tip shapes shows that tip shapes do not have a significant effect on the performance. Introduction of stroke deviation to kinematics leads to drop in average lift as wing interacts with the LEV shed during the downstroke. A linear elastic membrane model with applied aerodynamic load is developed for aeroelastic analysis. Analysis with different wing stiffnesses shows that the membrane wing outperforms the rigid wing in terms of lift, thrust and propulsive efficiency. The main reason for the increase in force production is attributed to the gliding of the LEV along the camber, which results in a high pressure difference across the surface. In addition, a high stiffness along the spanwise direction and low stiffness along the chordwise direction results in a uniform camber and high lift and thrust production. / Ph. D.