Computations of transonic viscous flows are very challenging. The major difficulty comes from the discontinuity in the solution across a shock wave, causing undesired oscillations in the solution. In this work we focus on minimizing the oscillations by the use of a limiter to control the amount of diffusivity. This limiter provides the right amount of viscosity to capture a sharp shock and an accurate solution in high gradient regions. The limiter employs changes in pressure and entropy and has been implemented into the Streamline Upwind Finite Element Method. A mesh adaptation strategy has been employed to further enhance the accuracy of the solution. Results of simulations over RAE 2822 airfoil and ONERA M6 wing indicate significant improvements to the solution with this implementation.

This thesis examines a new method of detecting the presence of structural cracks in wing-like structures at an incipient stage. This method is based on the analysis of the dynamics of damaged structures with bonded piezoelectric strips executing flexural oscillations. Such oscillations can be generated by mechanical loads, piezoelectric actuators, or unsteady aerodynamic loads in certain flight conditions of the aircraft. The proposed method of crack detection uses pairs of piezoelectric strip sensors bonded on the opposite sides of the structure and is based on the fact that the presence of a crack causes a difference between the strains measured by the two sensors of a given pair. The structural analysis presented in this thesis uses a nonlinear model for the cracks and a finite element formulation for the piezoelectric strips coupled with the structure. A 3D panel method is used to determine the unsteady aerodynamic loads acting on the oscillating wing. This study includes the dynamic analysis in time domain of cracked wing-like structures undergoing forced flexural vibrations in a range of frequencies generated by a pair of piezoelectric actuators, as well as the analysis of the oscillating wings with piezoelectric strips subjected to unsteady aerodynamic loads. The numerical simulations have shown that the presence of a crack in wing-like structures can be efficiently detected at an early stage by monitoring the response of the piezoelectric sensor pairs. / Cette thèse étudie une nouvelle méthode de détection de la présence de fissures structurelles à un stade précoce dans une structure de type aile. Cette méthode est basée sur l'analyse des oscillations en flexion des structures endommagées munies de bandes piézoélectriques collées. Ces oscillations peuvent être générées par des charges mécaniques, des actionneurs piézoélectriques, ou des charges aérodynamiques instationnaires dans certaines conditions de vol de l'avion. La méthode de détection des fissures proposée utilise des paires de capteurs piézoélectriques collés sur les côtés opposés de la structure et est basée sur le fait que la présence d'une fissure entraîne une différence entre les déformations mesurées par les deux capteurs d'une paire donnée. L'analyse structurale présentée dans cette thèse utilise un modèle non linéaire pour les fissures et une formulation par éléments finis pour les bandes piézoélectriques couplées avec la structure. Une méthode de panneau tridimensionnelle est utilisée pour déterminer les charges aérodynamiques instationnaires agissant sur l'aile oscillante. Cette étude comprend l'analyse dynamique dans le domaine temporel de structure de type aile fissurée subissant des vibrations en flexion forcées dans une gamme de fréquences générées par une paire d'actionneurs piézoélectriques, ainsi que l'analyse des ailes oscillantes équipées de bandes piézoélectriques soumises à des charges aérodynamiques instationnaires. Les simulations numériques ont montré que la présence d'une fissure dans ces structures peut être efficacement détectée à un stade précoce en surveillant la réponse des capteurs piézoélectriques.

This thesis deals with mathematical programming problems of flight mechanics of a jet aircraft. The aircraft is assumed to be in quasi-steady flight. Different flight phases are considered here; they cover rectilinear level flight, climbing flight, gliding flight, and curvilinear level flight. In this thesis, the problems studied include feasibility and optimization. These problems are solved via the modified quasilinearization algorithm for mathematical programming problems. Results for the different flight phases are presented here. They illustrate the performances of a jet aircraft in rectilinear level flight, climbing flight, gliding flight, and curvilinear level flight.

<p> One of the main research focus areas of the WVU Flight Control Systems Laboratory (FCSL) is the increase of flight safety through the implementation of fault tolerant control laws. For some fault tolerant flight control approaches with adaptive control laws, the availability of accurate post failure aircraft models improves performance. While look-up tables of aircraft models can be created for failure conditions, they may fail to account for all possible failure scenarios. Thus, a real-time parameter identification program eliminates the need to have predefined models for all potential failure scenarios. The goal of this research was to identify the dimensional stability and control derivatives of the WVU Phastball UAV in flight using a frequency domain based real-time parameter identification (PID) approach.</p><p> The data necessary for this project was gathered using the WVU Phastball UAV, a radio-controlled aircraft designed and built by the FCSL for fault tolerant control research. Maneuvers designed to excite the natural dynamics of the aircraft were implemented by the pilot or onboard computer during the steady state portions of flights. The data from these maneuvers was used for this project.</p><p> The project was divided into three main parts: 1) off-line time domain PID, 2) off-line frequency domain PID, and 3) an onboard frequency domain PID. The off-line parameter estimation programs, in both frequency domain and time domain, utilized the well known Maximum Likelihood Estimator with Newton-Raphson minimization with starting values estimated from a Least-Squares Estimate of the non-dimensional stability and control derivatives. For the frequency domain approach, both the states and inputs were first converted to the frequency domain using a Fourier integral over the frequency range in which the rigid body aircraft dynamics are found. The final phase of the project was a real-time parameter estimation program to estimate the dimensional stability and control derivatives onboard the Phastball aircraft. A frequency domain formulation of the least-squares estimation process was used because of its low computational and memory requirements and robustness to measurement noise and sensor information dropouts. Most of the onboard parameter estimates obtained converge to the values determined using the off-line parameter estimation programs (though a few typically show a bias) within four to six seconds for longitudinal estimates and four to eight seconds for the later estimates. For the experiments conducted, the real-time parameter estimates did not diverge after the conclusion of the maneuver.</p>

<p> The main objective of this thesis is to study the flow characteristics of a rotary cylinder with a symmetric end plate. We performed the simulations for different velocity ratios (0, 5, 10, and 15), aspect ratios (9.053 and 18) at high Reynolds numbers (1.15 x 10⁴ &le; Re &ge; 6.44 x 10⁵). We then studied the wake structure, the vortices formed in the wake region, the effect of vortex formation on the aerodynamic forces such as lift and drag. We performed computational fluid dynamics (CFD) simulations using a CFD solver, STAR-CCM+ from CD-Adapco. The results show that when the circular cylinder is stationary, the vortex shedding frequency is high; the upper and lower vortices show an asymmetrical process with the wake centerline. A significant vortex pairing can also be seen. But, with the rotation and increase in velocity ratio, the strength of vortex shedding decreases and after velocity ratio 5 the periodic vortex shedding is suppressed. The structure of the wake also modifies depending on the direction of the rotation. When aspect ratio of the circular cylinder is increased, the lift force generated on the cylinder surface is decreased. When an end plate is introduced in the region close to the stationary circular cylinder, it creates interference in the vortex formation and hence, the instabilities in the fluid flow due to vortices are decreased. The geometry of the stationary circular cylinder with an end plate seems to behave similar to a symmetric airfoil at zero angle of attack. Hence, aerodynamic forces generated on the geometry are constant. When the circular cylinder with the end plate is given a constant rotation, then the vortex formation is suppressed, the wake moves further downstream due to the end plate, the lift force generated on the surface increases and a significant decrease in drag force is also observed.</p>

<p> The objective of the present study is to investigate the characteristics of a flow around a rotating circular cylinder with and without an end plate near a wall boundary. The different cases which are taken into consideration in the current investigations were with gap ratios of 0.1d, 0.5d, 1.0d, 1.5d and 2.0d. A symmetric end plate is attached behind the rotating circular cylinder at a distance of 0.1d from the cylinder and a gap ratio of 1.5d. We performed Computational Fluid Dynamics (CFD) simulation of the flow around a rotating circular cylinder near a plane wall boundary using a CFD solver, STAR-CCM+. Free-stream velocity is kept constant at 5 m/s and the Reynolds number calculated is 3.24X10⁴. We then studied the flow characteristics such as lift and drag generated on the circular cylinder with and without an end plate and the wake structure. We observed that the vortex suppression is increased when the gap ratio is reduced, i.e., when the circular cylinder is nearer to the plane wall boundary. As the gap ratio increases the drag force generated decreases and the lift force increases considerably. In the case of rotating circular cylinder with an end plate, the wake area has moved upwards and the lift generated has increased manifold.</p>

<p> The objective of this study was to determine the deflected wing shape and the out-of-plane loads of a large-scale carbon-composite wing of an ultralight aerial vehicle using Fiber Bragg Grating (FBG) technology. The composite wing was instrumented with an optical fiber on its top and bottom surfaces positioned over the main spar, resulting in approximately 780 strain sensors bonded to the wings. The strain data from the FBGs was compared to that obtained from four conventional strain gages, and was used to obtain the out-of-plane loads as well as the wing shape at various load levels using NASA-developed real-time load and displacement algorithms. The composite wing measured 5.5 meters and was fabricated from laminated carbon uniaxial and biaxial prepreg fabric with varying laminate ply patterns and wall thickness dimensions. A three-tier whiffletree system was used to load the wing in a manner consistent with an in-flight loading condition.</p>

<p> Although multiphase gas and liquid phenomena occurs widely in engineering problems, many aspects of multiphase interaction like within droplet dynamics are still not quantified. This study aims to qualify the Lattice Boltzmann (LBM) Interparticle Potential multiphase computational method in order to build a foundation for future multiphase research. This study consists of two overall sections. </p><p> The first section in Chapter 2 focuses on understanding the LBM method and Interparticle Potential model. It outlines the LBM method and how it relates to macroscopic fluid dynamics. The standard form of LBM is obtained. The perturbation solution obtaining the Navier-Stokes equations from the LBM equation is presented. Finally, the Interparticle Potential model is incorporated into the numerical LBM method. </p><p> The second section in Chapter 3 presents the verification and validation cases to confirm the behavior of the single-phase and multiphase LBM models. Experimental and analytical results are used briefly to compare with numerical results when possible using Poiseuille channel flow and flow over a cylinder. While presenting the numerical results, practical considerations like converting LBM scale variables to physical scale variables are considered. Multiphase results are verified using Laplaces law and artificial behaviors of the model are explored. </p><p> In this study, a better understanding of the LBM method and Interparticle Potential model is gained. This allows the numerical method to be used for comparison with experimental results in the future and provides a better understanding of multiphase physics overall.</p>

The three-dimensional rendezvous between two spacecraft is considered: a target spacecraft on a circular orbit around the Earth and a chaser spacecraft initially on some elliptical orbit yet to be determined. The chaser spacecraft has variable mass, limited thrust, and its trajectory is governed by three controls, one determining the thrust magnitude and two determining the thrust direction. We seek the time history of the controls in such a way that the propellant mass required to execute the rendezvous maneuver is minimized. Two cases are considered: (i) time-to-rendezvous free and (ii) time-to-rendezvous given, respectively equivalent to (i) free angular travel and (ii) fixed angular travel for the target spacecraft. The above problem has been studied by several authors under the assumption that the initial separation coordinates and the initial separation velocities are given, hence known initial conditions for the chaser spacecraft. In this paper, it is assumed that both the initial separation coordinates and initial separation velocities are free except for the requirement that the initial chaser-to-target distance is given so as to prevent the occurrence of trivial solutions. Two approaches are employed: optimal control formulation (Part A) and mathematical programming formulation (Part B). In Part A, analyses are performed with the multiple-subarc sequential gradient-restoration algorithm for optimal control problems. They show that the fuel-optimal trajectory is zero-bang, namely it is characterized by two subarcs: a long coasting zero-thrust subarc followed by a short powered max-thrust braking subarc. While the thrust direction of the powered subarc is continuously variable for the optimal trajectory, its replacement with a constant (yet optimized) thrust direction produces a very efficient guidance trajectory. Indeed, for all values of the initial distance, the fuel required by the guidance trajectory is within less than one percent of the fuel required by the optimal trajectory. For the guidance trajectory, because of the replacement of the variable thrust direction of the powered subarc with a constant thrust direction, the optimal control problem degenerates into a mathematical programming problem with a relatively small number of degrees of freedom, more precisely: three for case (i) time-to-rendezvous free and two for case (ii) time-to-rendezvous given. In particular, we consider the rendezvous between the Space Shuttle (chaser) and the International Space Station (target). Once a given initial distance SS-to-ISS is preselected, the present work supplies not only the best initial conditions for the rendezvous trajectory, but simultaneously the corresponding final conditions for the ascent trajectory. In Part B, an analytical solution of the Clohessy-Wiltshire equations is presented (i) neglecting the change of the spacecraft mass due to the fuel consumption and (ii) and assuming that the thrust is finite, that is, the trajectory includes powered subarcs flown with max thrust and coasting subarc flown with zero thrust. Then, employing the found analytical solution, we study the rendezvous problem under the assumption that the initial separation coordinates and initial separation velocities are free except for the requirement that the initial chaser-to-target distance is given. The main contribution of Part B is the development of analytical solutions for the powered subarcs, an important extension of the analytical solutions already available for the coasting subarcs. One consequence is that the entire optimal trajectory can be described analytically. Another consequence is that the optimal control problems degenerate into mathematical programming problems. A further consequence is that, vis-a-vis the optimal control formulation, the mathematical programming formulation reduces the CPU time by a factor of order 1000. Key words. Space trajectories, rendezvous, optimization, guidance, optimal control, calculus of variations, Mayer problems, Bolza problems, transformation techniques, multiple-subarc sequential gradient-restoration algorithm.

This thesis uses linear covariance analysis to model a landmark tracking camera for lunar navigation on manned missions during a loss of communication scenario. The research provides evidence that this method satisfies Crew Exploration Vehicle requirements for autonomous navigation for returning astronauts safely to Earth from lunar orbit. This study broadens NASA's existing research efforts by creating a 6-degree-of-freedom linear covariance analysis tool to simulate the navigation errors generated in lunar orbit in the absence of ground updates. This methodology is capable of generating results which approximate those of a Monte Carlo study in a fraction of the time. Evidence is shown that landmark tracking can substantially reduce position and velocity errors while actively tracking a realistic set of lunar features.