A numerical simulation for evaluating methods of predicting and controlling the response of an elastic wing in an airstream is discussed. The technique employed interactively and simultaneously solves for the response in the time domain by considering the air, wing, and controller as elements of a single dynamical system. The method is very modular, allowing independent modifications to the aerodynamic, structural, or control subsystems and it is not restricted to periodic motions or simple geometries. To illustrate the technique, we use a High Altitude, Long Endurance aircraft wing. The wing is modeled structurally as a linear Euler-Bernoulli beam that includes dynamic coupling between the bending and torsional oscillations. The governing equations of motion are derived and extended to allow for rigid-body motions of the wing. The exact solution to the unforced linear problem is discussed as well as a Galerkin and finite-element approximations. The finite-element discretization is developed and used for the simulations. A general, nonlinear, unsteady vortex-lattice method, which is capable of simulating arbitrary subsonic maneuvers of the wing and accounts for the history of the motion, is employed to model the flow around the wing and provide the aerodynamic loads. Two methods of incorporating gusts in the aerodynamic model are also discussed. Control of the wing is effected via a distributed torque actuator embedded in the wing and two strategies for actuating the wing are described: a classical linear proportional integral strategy and a novel nonlinear feedback strategy based on the phenomenon of saturation that may exist in nonlinear systems with two-to-one internal resonances. Both control strategies can suppress the flutter oscillations of the wing, but the nonlinear controller must be actively tuned to be effective; gust control proved to be more difficult. / Master of Science

The nonlinear interactions between aerodynamic forces and wing structures are numerically investigated as integrated dynamic systems, including structural models, aerodynamics, and control systems, in the time domain. An elastic beam model coupled with rigid-body rotation is developed for the wing structure, and the natural frequencies and mode shapes are found by the finite-element method. A general unsteady vortex-lattice method is used to provide aerodynamic forces. This method is verified by comparing the numerical solutions with the experimental results for several cases; and thereafter applied to several applications such as the inboard-wing/twin-fuselage configuration, and formation flights. The original thought that the twin fuselage could achieve two-dimensional flow on the wing by eliminating free wing tips appears to be incorrect. The numerical results show that there can be a lift increase when two or more wings fly together, compared to when they fly alone. Flutter analysis is carried out for a High-Altitude-Long-Endurance aircraft wing cantilevered from the wall of the wind tunnel, a full-span wing mounted on a free-to-roll sting at its mid-span without and with a center mass (fuselage). Numerical solutions show that the rigidity added by the wall results in a higher flutter speed for the wall-mounted semi-model than that for the full-span model. In addition, a predictive control technique based on neural networks is investigated to suppress flutter oscillations. The controller uses a neural network model to predict future plant responses to potential control signals. A search algorithm is used to select the best control input that optimizes future plant performance. The control force is assumed to be given by an actuator that can apply a distributed torque along the spanwise direction of the wing. The solutions with the wing-tip twist or the wing-tip deflection as the plant output show that the flutter oscillations are successfully suppressed with the neural network predictive control scheme. / Ph. D.

The ability of fish to deform their bodies in steady swimming action is gaining from robotic designers. While bound by the same physical laws, fish have evolved to move in ways that often outperform artificial systems in critical measures such as efficiency, agility, and stealth through thousands of years of natural selection. As we expand our presence in the ocean with deep-sea exploration or offshore drilling for petroleum and natural gas, the demand for prolonging underwater operations is growing significantly. Therefore, it is critical for robotic designers to understand the physics of pisciform (fish-like) locomotion and learn how to effectively implement the propulsive mechanisms into their designs to create the next generation of aquatic robots. Aiming to assist this process, this thesis presents an experimental apparatus called Towed Biolocomotion Emulator (TBE), which is capable of imitating the undulating action of different fish species in steady swimming and can be quickly adapted to different configurations with modular modules. Using the TBE device, an experiment is performed to test its hydrodynamic performance and evaluate the effectiveness of the bio-inspired locomotion implemented on such a mechanical system. The analysis of hydrodynamic data collected from the experiment shows that there exists a small range of kinematic parameters where the undulating motion of the device produces the optimal performance. This result confirms the benefits of utilizing pisciform locomotion for small-scale underwater vehicles. In addition, this thesis also proposes a reduced-order flow model using the unsteady vortex lattice method (UVLM) to predict the hydrodynamic performance of such a system. The proposed model is then validated with the experimental data collected earlier. The tool developed can be employed to quickly explore the possible design space early in the conceptual design stage for such a bio-mimetic vehicle. / Master of Science / It is no surprise that through thousands of years of natural evolution, marine species possess incredible ability to navigate through water. As we expand our presence in the sea, more and more tasks require underwater operations such as ocean exploration, oil-rig maintenance, etc. Yet, most of the underwater robotic vehicles still utilize propellers as the primary propulsive mechanism. In many cases, the bio-inspired propulsion system that mimics the swimming action of fish offers many advantages in agility, maneuverability, and stealth. With the rising interest in the field, the works presented in this thesis aim to expand our understanding of how to implement the bio-inspired propulsive mechanism to robotic design. To achieve this, a mechanical device is designed to mimic the swimming action of different fish species. Then, an experiment is performed to subject the device to different fish-like motions and test their effectiveness. In addition, a reduced-ordered model is also introduced as an alternative method to predict the hydrodynamic performance of this propulsive mechanism. The works presented in this thesis help to expand the toolbox available for the engineer to design the next generation of the underwater robotic vehicle.

This thesis describes the development of a nonlinear flight dynamics model for a small, fixed-wing unmanned aerial vehicle (UAV). Models developed for UAVs can be used for many applications including risk analysis, controls system design and flight simulators. Several challenges exist for system identification of small, low-cost aircraft including an increased sensitivity to atmospheric disturbances and decreased data quality from a cost-appropriate instrumentation system. These challenges result in difficulties in development of the model structure and parameter estimation. The small size may also limit the scope of flight test experiments and the consequent information content of the data from which the model is developed. Methods are presented to improve the accuracy of system identification which include data selection, data conditioning, incorporation of information from computational aerodynamics and synthesis of information from different flight test maneuvers. The final parameter estimation and uncertainty analysis was developed from the time domain formulation of the output-error method using the fully nonlinear aircraft equations of motion and a nonlinear aerodynamic model structure. The methods discussed increased the accuracy of parameter estimates and lowered the uncertainty in estimates compared to standard procedures for parameter estimation from flight test data. The significant contributions of this thesis are a detailed explanation of the entire system identification process tailored to the needs of a small UAV and incorporation of unique procedures to enhance identification results. This work may be used as a guide and list of recommendations for future system identification efforts of small, low-cost, minimally instrumented, fixed-wing UAVs. / MS / This thesis describes identification of a series of equations to model the flight motion of a small unmanned airplane. Model development for small unmanned aerial vehicles (UAVs) is a challenging process because they are significantly affected by small amounts of wind and they usually contain inexpensive, lower quality sensors. This results in lower quality data measured from flying a small UAV, which is subsequently used in the process to develop a model for the aircraft. In this work, techniques are discussed to improve estimation of model parameters and increase confidence in the validity of the final model. The significant contributions of this thesis are a comprehensive explanation of the model development process specific to a small UAV and implementation of unique procedures to enhance the resulting model. This work as a whole may be used as a guide and list of recommendations for future model development efforts of small, low-cost, unmanned aircraft.

Flutter is one of the primary aeroelastic phenomena that must be considered in aircraft design. Flutter is a self-sustaining structural vibration in which energy is extracted from the air flow and transferred to the structure. The amplitude of the vibration grows exponentially until structural failure occurs. Flutter stability requirements often influence the design of an aircraft, making accurate flutter prediction capabilities an essential part of the design process. Advances in computational fluid dynamics and computational power make it possible to solve the fluid flow and structural dynamics simultaneously, providing highly accurate solutions especially in the transonic flow regime. This procedure is, however, too time-consuming to be used in the design optimisation process. As a result panel codes, e.g., the doublet lattice method, and modal-based structural analysis methods are still being used extensively and continually improved. One application that is lagging in terms of accuracy and simplicity (from the user’s perspective) is the flutter analysis of T-tails. The flutter analysis of a T-tail usually involves the calculation of additional aerodynamic loads, apart from the loads calculated by the standard unsteady aerodynamic codes for conventional empennages. The popular implementations of the doublet lattice method do not calculate loads due to the in-plane motion (i.e., lateral or longitudinal motion) of the horizontal stabiliser or the in-plane loads on the stabiliser. In addition, these loads are dependent on the steady-state load distribution on the stabiliser, which is ignored in the doublet lattice method. The objective of the study was to extend the doublet lattice method to calculate the additional aerodynamic loads that are crucial for T-tail flutter analysis along with the customary unsteady air loads for conventional configurations. This was achieved by employing the Kutta-Joukowski theorem in the calculation of unsteady air loads on lifting surface panels. Calculating the additional unsteady air loads for T-tails within the doublet lattice method significantly reduces the human effort required for T-tail flutter analysis as well as the opportunities for introducing errors into the analysis. During the course of the study it became apparent that it was necessary to consider the quadratic mode shape components in addition to the linear mode shape components. Otherwise the unsteady loads due to the rotation (“tilting”) of the steady-state load on the stabiliser, one of the additional aerodynamic loads that are crucial for T-tail flutter analysis, would give rise to spurious generalised forces. In order to reduce the additional burden of determining the quadratic mode shape components, methods for calculating quadratic mode shape components using linear finite element analysis or estimating them from the linear mode shape components were developed. Wind tunnel tests were performed to validate the proposed computational method. A T-tail flutter model which incorporated a mechanism for changing the incidence angle of the horizontal stabiliser, and consequently the steady-state load distribution on the horizontal stabiliser, was used. The flutter speed of this model as a function of the horizontal stabiliser incidence was determined experimentally and compared to predictions. Satisfactory correlation was found between predicted and experimentally determined flutter speeds. / Thesis (M.Ing. (Chemical Engineering))--North-West University, Potchefstroom Campus, 2012

abstract: This study identifies the influence that leading-edge shape has on the aerodynamic characteristics of a wing using surface far-field and near-field analysis. It examines if a wake survey is the appropriate means for measuring profile drag and induced drag. The paper unveils the differences between sharp leading-edge and blunt leading-edge wings with the tools of pressure loop, chordwise pressure distribution, span load plots and with wake integral computations. The analysis was performed using Computational Fluid Dynamics (CFD), vortex lattice potential flow code (VORLAX), and a few wind-tunnels runs to acquire data for comparison. This study found that sharp leading-edge wings have less leading-edge suction and higher drag than blunt leading-edge wings. The blunt leading-edge wings have less drag because the normal vector of the surface in the front section of the airfoil develops forces at opposed skin friction. The shape of the leading edge, in conjunction with the effect of viscosity, slightly alter the span load; both the magnitude of the lift and the transverse distribution. Another goal in this study is to verify the veracity of wake survey theory; the two different leading-edge shapes reveals the shortcoming of Mclean’s equation which is only applicable to blunt leading-edge wings. / Dissertation/Thesis / Masters Thesis Aerospace Engineering 2019

A computational approach to analyze the hydrodynamic performance of a hydrofoil or a marine propeller undergoing unsteady motion has been developed. In order to simulate heave and pitch motion of a hydrofoil, an unsteady boundary element method based modeling is performed. The wake of the hydrofoil is modeled by a continuous dipole sheet and determined in time by applying a force-free condition on its surface. An explicit vortex core model is adapted in this model to capture the rolling up shape and to avoid instability due to roll-up deformation of the wake. The numerical results of the developed model are compared with analytical results and those from the commercial Reynolds-Averaged Navier-Stokes solver (ANSYS/FLUENT). The results show close level of agreement with each other. The problem of flow around a marine propeller performing surge, roll and heave motion in an unbounded fluid is formulated and solved using both a vortex-lattice method and a boundary element method. A fully unsteady wake alignment algorithm is implemented into the vortex-lattice method in order to satisfy the force-free condition on the propeller wake surface. Finally, a comparative study of transient propeller forces on a propeller blade obtained from BEM and VLM (with or without fully aligned wake) is carried out and results are presented. In some cases, results from the presented methods are compared with those from RANS or other numerical methods available in the literature. / text

Flutter is one of the primary aeroelastic phenomena that must be considered in aircraft design. Flutter is a self-sustaining structural vibration in which energy is extracted from the air flow and transferred to the structure. The amplitude of the vibration grows exponentially until structural failure occurs. Flutter stability requirements often influence the design of an aircraft, making accurate flutter prediction capabilities an essential part of the design process. Advances in computational fluid dynamics and computational power make it possible to solve the fluid flow and structural dynamics simultaneously, providing highly accurate solutions especially in the transonic flow regime. This procedure is, however, too time-consuming to be used in the design optimisation process. As a result panel codes, e.g., the doublet lattice method, and modal-based structural analysis methods are still being used extensively and continually improved. One application that is lagging in terms of accuracy and simplicity (from the user’s perspective) is the flutter analysis of T-tails. The flutter analysis of a T-tail usually involves the calculation of additional aerodynamic loads, apart from the loads calculated by the standard unsteady aerodynamic codes for conventional empennages. The popular implementations of the doublet lattice method do not calculate loads due to the in-plane motion (i.e., lateral or longitudinal motion) of the horizontal stabiliser or the in-plane loads on the stabiliser. In addition, these loads are dependent on the steady-state load distribution on the stabiliser, which is ignored in the doublet lattice method. The objective of the study was to extend the doublet lattice method to calculate the additional aerodynamic loads that are crucial for T-tail flutter analysis along with the customary unsteady air loads for conventional configurations. This was achieved by employing the Kutta-Joukowski theorem in the calculation of unsteady air loads on lifting surface panels. Calculating the additional unsteady air loads for T-tails within the doublet lattice method significantly reduces the human effort required for T-tail flutter analysis as well as the opportunities for introducing errors into the analysis. During the course of the study it became apparent that it was necessary to consider the quadratic mode shape components in addition to the linear mode shape components. Otherwise the unsteady loads due to the rotation (“tilting”) of the steady-state load on the stabiliser, one of the additional aerodynamic loads that are crucial for T-tail flutter analysis, would give rise to spurious generalised forces. In order to reduce the additional burden of determining the quadratic mode shape components, methods for calculating quadratic mode shape components using linear finite element analysis or estimating them from the linear mode shape components were developed. Wind tunnel tests were performed to validate the proposed computational method. A T-tail flutter model which incorporated a mechanism for changing the incidence angle of the horizontal stabiliser, and consequently the steady-state load distribution on the horizontal stabiliser, was used. The flutter speed of this model as a function of the horizontal stabiliser incidence was determined experimentally and compared to predictions. Satisfactory correlation was found between predicted and experimentally determined flutter speeds. / Thesis (M.Ing. (Chemical Engineering))--North-West University, Potchefstroom Campus, 2012

A hazardous and not fully understood aeroelastic phenomenon is the transonic dip,the decrease in flutter dynamic pressure that occurs for most aircraft configurationsin transonic flows. The difficulty of predicting this phenomenon forces aircraft manufacturersto run long and costly flight test campaigns to demonstrate flutter-free behaviourof their aircraft at transonic Mach numbers.In this project, subsonic and transonic flutter calculations for the KTH-NASA genericfighter research model have been performed and compared to existing experimentalflutter data from wind tunnel tests performed at NASA Langley in 2016. For the fluttercalculations, industry-standard linear panel methods have been used together with afinite element model from NASTRAN.Further, an alternative approach for more accurate transonic flutter predictions usingthe full-potential solver Phi has been investigated. To predict flutter using this newmethodology a simplified structural model has been used together with aerodynamicmeshes of the main wing. The purpose of the approach was to see if it was possibleto find a method that was more accurate than panel methods in the transonic regimewhilst still being suitable for use during iterative design processes.The results of this project demonstrated that industry-standard linear panel methodssignificantly over-predict the flutter boundary in the transonic regime. It was alsoseen that the flutter predictions using Phi showed potential, being close to the linearresults for the same configuration as tested in Phi. For improved transonic accuracy inPhi, an improved transonic flow finite element formulation could possibly help .Another challenge with Phi is the requirement of an explicit wake from all liftingsurfaces in the aerodynamic mesh. Therefore, a method for meshing external storeswith blunt trailing edges needs to be developed. One concept suggested in this projectis to model external stores in "2.5D", representing external stores using airfoils withsharp trailing edges. / Ett farligt och inte helt utrett aeroelastiskt fenomen är den transoniska dippen, minskningeni dynamiska trycket vid fladder som inträffar för de flesta flygplan i transoniskaflöden. Svårigheten i att prediktera detta fenomen tvingar flygplanstillverkare attbedriva tidskrävande och kostsam flygprovsverksamhet för att demonstrera att derasflygplan ej uppvisar fladderbeteende i transonik inom det tilltänkta användningsområdet.I detta projekt har fladderberäkningar genomförts i både underljud och transonikför en generisk stridsflygplansmodell i skala 1:4 ämnad för forskning, byggd som ettsamarbete mellan KTH och NASA. Beräkningarna har också jämförts med fladderresultatfrån vindtunnelprov genomförda vid NASA Langley under sommaren 2016. Förfladderberäkningarna har industri-standarden linjära panelmetoder används tillsammansmed en befintlig finit element modell för användning i NASTRAN.Vidare har ett alternativt tillvägagångssätt för att förbättra precisionen i transoniskafladderresultat genom att använda potentiallösaren Phi undersökts. En förenkladstrukturmodell har använts tillsammans med aerodynamiska nät av huvudvingen föratt prediktera fladder. Syftet med denna metodik var att undersöka om det var möjligtatt hitta en metod som i transoniska flöden var mer exakt än panelmetoder men somfortfarande kunde användas i iterativa design processer.Resultaten från detta projekt visade att linjära panelmetoder, som de som används iindustrin, är signifikant icke-konservativa gällande fladdergränsen i transonik. Resultatenfrån Phi visade potential genom att vara nära de linjära resultaten som räknadesfram med hjälp av panelmetoder för samma konfiguration som i Phi. För ökad transonisknoggrannhet i Phi kan möjligen en förbättrad transonisk element-formuleringhjälpa.En annan utmaning med Phi är kravet på en explicit vak från alla bärande ytor idet aerodynamiska nätet. Därför behöver det utvecklas en metodik för nätgenereringav yttre laster med trubbiga bakkanter. Ett koncept som föreslås i denna rapport är attmodellera yttre laster i "2.5D", där alla yttre laster beskrivs genom att använda vingprofilermed skarpa bakkanter.

This dissertation presents shape and structural optimization studies on flapping wings for micro air vehicles. The design space of the optimization includes the wing planform and the structural properties that are relevant to the wing model being analyzed. The planform design is parameterized using a novel technique called modified Zimmerman, which extends the concept of Zimmerman planforms to include four ellipses rather than two. Three wing types are considered: rigid, plate-like deformable, and membrane. The rigid wing requires no structural design variables. The structural design variables for the plate-like wing are the thickness distribution polynomial coefficients. The structural variables for the membrane wing control the in-plane distributed forces which modulate the structural deformation of the wing. The rigid wing optimization is performed using the modified Zimmerman method to describe the wing. A quasi-steady aerodynamics model is used to calculate the thrust and input power required during the flapping cycle. An assumed inflow model is derived based on lifting-line theory and is used to better approximate the effects of the induced drag on the wing. A multi-objective optimization approach is used since more than one aspect is considered in flapping wing design. The the epsilon-constraint approach is used to calculate the Pareto optimal solutions that maximize the cycle-average thrust while minimizing the peak input power and the wing mass. An aeroelastic model is derived to calculate the aerodynamic performance and the structural response of the deformable wings. A linearized unsteady vortex lattice method is tightly coupled to a linear finite element model. The model is cost effective and the steady-state solution is solved by inverting a matrix. The aeroelastic model is used to maximize the thrust produced over one flapping cycle while minimizing the input power. / Ph. D.