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.

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This dissertation presents both analytical and numerical approaches to optimizing thrust and thrust efficiency of harmonically deforming thin airfoils and membrane wings. A frequency domain unsteady aerodynamic theory for deformable thin airfoils, with Chebychev polynomials as the basis functions is presented. Stroke-averaged thrust and thrust efficiency expressions are presented in a quadratic matrix form. The motion and deformation of the airfoil is optimized for maximum thrust and efficiency. Pareto fronts are generated showing optimum deformation conditions (magnitude and phase) for various reduced frequencies and constraints. It is shown that prescribing the airfoil to deform in a linear combination of basis functions with optimal magnitude and phase results in a larger thrust as compared to rigid plunging, especially at low reduced frequencies. It is further shown that the problem can be constrained significantly such that thrust is due entirely to pressure with no leading edge suction, and associated leading edge separation. The complete aeroelastic system for a membrane wing is also optimized. The aerodynamic theory for deformable thin airfoils is used as the forcing in a membrane vibration problem. Due to the nature of the two dimensional theory, the membrane vibration problem is reduced to two dimensions via the Galerkin method and nondimensionalized such that the only terms are nondimesional tension, mass ratio and reduced frequency. The maximum thrust for the membrane wing is calculated by optimizing the tension in the membrane so that the the aeroelastic deformation due to wing motion leads to optimal thrust and/or efficiency. A function which describes the optimal variation of spanwise tension along the chord is calculated. It is shown that one can always find a range of membrane tension for which the flexible membrane wings performs better than the rigid wing. These results can be used in preliminary flapping wing MAV design. / Ph. D.

Complex nonlinear aspects of aeroelastic phenomena include unsteady nonlinear aerodynamic loads, structural nonlinearities, as well as nonlinear couplings between the flow and the structural response. Nonlinearities in aerodynamic loads originate from unsteady shocks and/or flow separation. Structural nonlinearities are geometric, or a result of free play. Nonlinear fluid structure couplings result from nonlinear resonance between the aerodynamic load and structural modes. Under different conditions, one or a combination of these aspects could yield flutter or Limit Cycle Oscillations (LCO). The overall goal of this work is to develop the capabilities to quantify the role that these different nonlinear mechanisms could play in observed flutter and LCO. The realization of such a goal would help in providing a benchmark for the detection of nonlinear aeroelastic instabilities and possibly effective means for obtaining improved performance and reduced uncertainties through operation beyond conventional boundaries that are based on linear analysis. Additionally, this effort will provide a benchmark for the validation of computational methodologies. In this thesis, wavelet-based higher order spectra are applied to identify different nonlinear aeroelastic phenomena as encountered in two experiments. First, the analysis is applied to a set of experiments involving a flexible semispan model (FSM) of a High Speed Civil Transport (HSCT) wing configuration conducted by Silva et al. (Experimental Steady and Unsteady Aerodynamic and Flutter Results for HSCT Semispan Models; AIAA/ASME/ASCE/AHS/ASC 41st Structures, Structural Dynamics, and Materials Conference, 2000). The interest is in the identification of nonlinear aeroelastic phenomena associated with a high dynamic response region which was measured over a large range of dynamic pressures around Mach number 0.98. At the top of this region is a ``hard'' flutter point that resulted in the loss of the model. The results show that ``hard'' flutter is related to intermittent nonlinear coupling between the shock motion and large amplitude structural motions. Second, the analysis is applied to identify nonlinear aspects of LCO encountered during test flights of an F-16 aircraft. The results show quadratic and cubic couplings in the acceleration signals of the under-wing launchers and high quadratic coupling levels between flaperon motions and wing oscillations. The implications of applying these techniques in the capacity of a ``flutterometer'' are also discussed. / Ph. D.

Multidisciplinary interactions are expected to play a significant role in the design of future high-performance aircraft (Blended-Wing Body, Truss-Braced wing, High Speed Civil transport, High-Altitude Long Endurance aircraft and future military aircraft). Also, the availability of supercomputers has made it now possible to employ high-fidelity models (Computational Fluid Dynamics for fluids and detailed finite element models for structures) at the preliminary design stage. A necessary step at that stage is to calculate the wing angle-of-attack at which the wing will generate the desired lift for the specific flight maneuver. Determination of this angle, a simple affair when the wing is rigid and the flow regime linear, becomes difficult when the wing is flexible and the flow regime non-linear. To solve this inherently nonlinear problem, a Newton's method type algorithm is developed to simultaneously calculate the deflection and the angle of attack. The present algorithm requires the sensitivity of the aerodynamic pressure with respect to each of the generalized displacement coordinates needed to represent the structural displacement. This sensitivity data is easy to determine analytically when the flow regime is linear. The present algorithm uses a finite difference method to obtain these sensitivities and thus requires only the pressure data and the surface geometry from the aerodynamic model. This makes it ideally suited for nonlinear aerodynamics for which it is difficult to obtain the sensitivity analytically. The present algorithm requires the CFD code to be run for each of the generalized coordinates. Therefore, to reduce the number of generalized coordinates considerably, we employ the modal superposition approach to represent the structural displacements. Results available for the Aeroelastic Research Wing (ARW) are used to evaluate the performance of the modal superposition approach. Calculations are made at a fixed angle of attack and the results are compared to both the experimental results obtained at NASA Langley Research Center, and computational results obtained by the researchers at NASA Ames Research Center. Two CFD codes are used to demonstrate the modular nature of this research. Similarly, two separate Finite Element codes are used to generate the structural data, demonstrating that the algorithm is not dependent on using specific codes. The developed algorithm is tested for a wing, used for in-house aeroelasticity research at Boeing (previously McDonnell Douglas) Long Beach. The trim angle of attack is calculated for a range of desired lift values. In addition to the Newton's method algorithm, a non derivative method (NDM) based on fixed point iteration, typical of fixed angle of attack calculations in aeroelasticity, is employed. The NDM, which has been extended to be able to calculate trim angle of attack, is used for one of the cases. The Newton's method calculation converges in fewer iterations, but requires more CPU time than the NDM method. The NDM, however, results in a slightly different value of the trim angle of attack. It should be noted that NDM will converge in a larger number of iterations as the dynamic pressure increases. For one value of the desired lift, both viscous and inviscid results were generated. The use of the inviscid flow model while not resulting in a markedly different value for the trim angle of attack, does result in a noticeable difference both in the wing deflection and the span loading when compared to the viscous results. A crude (coarse-grain) parallel methodology was used in some of the calculations in this research. Although the codes were not parallelized, the use of modal superposition made it possible to compute the sensitivity terms on different processors of an IBM SP/2. This resulted in a decrease in wall clock time for these calculations. However, even with the parallel methodology, the CPU times involved may be prohibitive (approximately 5 days per Newton iteration) to any practical application of this method for wing analysis and design. Future work must concentrate on reducing these CPU times. Two possibilities: (i) The use of alternative basis vectors to further reduce the number of basis vectors used to represent the structural displacement, and (ii) The use of more efficient methods for obtaining the flow field sensitivities. The former will reduce the number of CFD analyses required the latter the CPU time per CFD analysis. NOTE: (03/2007) An updated copy of this ETD was added after there were patron reports of problems with the file. / Ph. D.

A new numerical methodology to examine fluid-structure interaction of a wing/pylon/store system has been developed. The aeroelastic equation of motion of the complete system is solved iteratively in the time domain using a two-entity numerical code comprised of ABAQUS/Standard and the Unsteady-Vortex-Lattice Method. Both codes communicate through an iterative handshake procedure during which displacements and air loads are updated. For each increment in time the force/displacement equilibrium is found in this manner. The wing, pylon, and store data considered in this analysis are based on an F16 configuration that was identified to induce flutter in flight at subsonic speeds. The wing structure is modeled as an elastic plate and pylon and store are rigid bodies. The store body is connected to the pylon through an elastic joint exercising pitch and yaw degrees of freedom. Vortex-Lattice theory featuring closed ring-vortices and continuous vortex shedding to form the wakes is employed to model the aerodynamics of wing, store, and pylon. The methodology was validated against published data demonstrating excellent agreement with documented key phenomena of fluid-structure iteration. The model correctly predicts the effects of the pylon induced lateral flow disruption as well as wing-tip-vortex effects. It can identify the presence of aerodynamic interference between the store, pylon, and wing wakes and examine its significance with respect to the pressure and lift forces on the participating bodies. An elementary flutter study was undertaken to examine the dynamic characteristics of a stiff production pylon at near-critical airspeeds versus those of a soft-in-pitch pylon. The simulation reproduced the stabilizing effect of the stiffness reduction in the pitch motion. This idea is based on the concept of the decoupler pylon, introduced by Reed and Foughner in 1978 and flight tested in the early 1980's. NOTE: (3/07) An updated copy of this ETD was added after there were patron reports of problems with the file. / Ph. D.

Based on a refined analytical anisotropic thin-walled beam model, aeroelastic instability, dynamic aeroelastic response, active/passive aeroelastic control of advanced aircraft wings modeled as thin-walled beams are systematically addressed. The refined thin-walled beam model is based on an existing framework of the thin-walled beam model and a couple of non-classical effects that are usually also important are incorporated and the model herein developed is validated against the available experimental, Finite Element Anaylsis (FEA), Dynamic Finite Element (DFE), and other analytical predictions. The concept of indicial functions is used to develop unsteady aerodynamic model, which broadly encompasses the cases of incompressible, compressible subsonic, compressible supersonic and hypersonic flows. State-space conversion of the indicial function based unsteady aerodynamic model is also developed. Based on the piezoelectric material technology, a worst case control strategy based on the minimax theory towards the control of aeroelastic systems is further developed. Shunt damping within the aeroelastic tailoring environment is also investigated. The major part of this dissertation is organized in the form of self-contained chapters, each of which corresponds to a paper that has been or will be submitted to a journal for publication. In order to fullfil the requirement of having a continuous presentation of the topics, each chapter starts with the purely structural models and is gradually integrated with the involved interactive field disciplines. / Ph. D.

There is a need for a nonlinear flutter analysis method capable of predicting limit cycle oscillation in aeroelastic systems. A review is conducted of analysis methods and experiments that have attempted to better understand and model limit cycle oscillation (LCO). The recently developed method of nonlinear normal modes (NNM) is investigated for LCO calculation. Nonlinear normal modes were used to analyze a spring-mass-damper system with nonlinear damping and stiffness to demonstrate the ability and limitations of the method to identify limit cycle oscillation. The nonlinear normal modes method was then applied to an aeroelastic model of a pitch-plunge airfoil with nonlinear pitch stiffness and quasi-steady aerodynamics. The asymptotic coefficient solution method successfully captured LCO at a low relative velocity. LCO was also successfully modeled for the same airfoil with an unsteady aerodynamics model with the use of a first order formulation of NNM. A linear beam model of the Goland wing with a nonlinear aerodynamic model was also studied. LCO was successfully modeled using various numbers of assumed modes for the beam. The concept of modal truncation was shown to extend to NNM. The modal coefficients were shown to identify the importance of each mode to the solution and give insight into the physical nature of the motion. The quasi-steady airfoil model was used to conduct a study on the effect of the nonlinear normal mode's master coordinate. The pitch degree of freedom, plunge degree of freedom, both linear structural mode shapes with apparent mass, and the linear flutter mode were all used as master coordinates. The master coordinates were found to have a significant influence on the accuracy of the solution and the linear flutter mode was identified as the preferred option. Galerkin and collocation coefficient solution methods were used to improve the results of the asymptotic solution method. The Galerkin method reduced the error of the solution if the correct region of integration was selected, but had very high computational cost. The collocation method improved the accuracy of the solution significantly. The computational time was low and a simple convergent iteration method was found. Thus, the collocation method was found to be the preferred method of solving for the modal coefficients. / Ph. D.

Several procedures for the simultaneous aerodynamic-structural design optimization of aircraft wings are investigated. These procedures include efficient methods for optimization and sensitivity calculations that are applied to two specific design examples. The first is a subsonic transport aircraft with a composite forwardswept wing. The aerodynamic modeling for this case is provided by vortex-lattice theory and the structural model initially utilizes finite-element analyses. Even with efficient sensitivity methods, the approximate optimization problem still requires a large computational effort. To reduce this cost, a variable-complexity model for the structural analyses is introduced. First, an algebraic equation model for wing weight is used in the optimization procedure to obtain an aerodynamic design that approximately accounts for the effects of wing geometry on wing weight. Then this design is refined by simultaneous aerodynamic-structural optimization based on the finite-element analysis. The net effect of this dual structural model is a substantial reduction in optimization costs. The second example is the wing design of a supersonic High-Speed Civil Transport (HSCT). For this case, the simple wing-weight equations for structures are retained. For the aerodynamics, a variable-complexity model was introduced with the complex models provided by volumetric wave drag analysis and panel methods. In addition, simple algebraic models for wave and drag due to lift provide inexpensive approximations during most of the optimization cycles. With the minimization of the costly complex sensitivity calculations, a reduction in optimization costs is realized. / Ph. D.