Dissertations / Theses on the topic 'Aerospace engineering – Simulation methods'

Hybrid Large Eddy Simulation/Reynolds-Averaged Navier-Stokes (LES/RANS) simulations of several high-speed flows are presented in this work. The solver blends a Menter BSL two-equation model for the RANS part of the closure with a Smagorisnky sub-grid model for the LES component. The solver uses a flow-dependent blending function based on wall distance and a modeled form of the Taylor micro-scale to transition from RANS to LES. Turbulent fluctuations are initiated and are sustained in the inflow region using a recycling/rescaling technique. A new multi-wall recycling/rescaling technique is described and tested. A spanwise-shifting method is introduced that is intended to alleviate unphysical streamwise streaks of high- and low-momentum fluid that appear in the time-averaged solution due to the recycling procedure. Simulations of sonic injection of air, helium and ethylene into a Mach 2 cross-flow of air are performed. Also, simulations of Mach 5 flow in a subscale inlet/isolator configuration with and without back-pressuring are performed. Finally, a Mach 3.9 flow through a square duct is used as an initial test case for the new multi-wall recycling and rescaling method as well as a multi-wall shifting procedure. A discussion of the methods, implementation and results of these simulations is included.

It has been said that engineers create that which never was. The university experience is a key component in preparing engineers who support the creation of products and systems that improve the world we live in. The way in which engineers have been trained in universities has changed throughout history in America, moving from an apprentice-like approach to the still-used engineer scientist. Some in industry and academia feel that this model of engineer preparation needs to change in order to better address the complexities of engineering in the 21st century, and help fill a perceived gap between academic preparation and 21st century industrial necessity. A new model for student preparation centering on engineering design called the Live Simulation Based Learning (LSBL) approach is proposed based upon the theories of situated learning, game-based learning, epistemic frames, and accidental competencies. This dissertation discusses the results of a study of the application of LSBL in a two term capstone design class in aerospace engineering aircraft design at Virginia Tech. It includes LSBLâ s impact on student professional and technical skills in relation to aerospace engineering design practice. Results indicate that the participants found the LSBL experience to be more engaging than the traditional lecture approach and does help students respond and think more like aerospace engineering practicing professionals and thus begin to address the â gapâ between academia and industry.
Ph. D.

The ongoing drive for lighter and more efficient structural components by the commercial engineering industry has resulted in the rapid adoption of the finite element method (FE) for design analysis. Satisfied with the success of finite elements in reducing prototyping costs and overall production times, the industry has begun to look at other areas where the finite element method can save time, and in particular, improve designs. First, the mathematical methods of optimisation, on which the methods of structural design improvement are based, are presented. This includes the methods of: topology, influence functions, basis vectors, geometric splines and direct sensitivity methods. Each method is demonstrated with the solution of a sample structural improvement problem for various objectives (frequency, stress and weight reduction, for example). The practical application of the individual methods has been tested by solving three structural engineering problems sourced from the automotive engineering industry: the redesign of two different front suspension control arms, and the cost-reduction of an automatic brake tubing system. All three problems were solved successfully, resulting in improved designs. Each method has been evaluated with respect the practical application, popularity of the method and also any problems using the method. The solutions presented in each section were all solved using the FE design improvement software ReSHAPE from Advea Engineering Pty. Ltd.

The flow produced by a rotating propeller is inherently unsteady and three-dimensional. Conventional design of propellers uses blade-element theory but becomes inaccurate in capturing three-dimensional vortical and compressible effects at the tips, as well as the effect on downstream bodies. A propeller is always attached to a fixed component that affects its performance, thus the need arises to couple a fixed domain to a rotating domain in an unsteady aerodynamic simulation. A finite element formulation for the simulation of propellers is presented in terms of the Reynolds-averaged Navier-Stokes equations for unsteady, three-dimensional, viscous, compressible flows. The first step consists of preparing a mesh containing two separate domains interfacing at a virtual surface. Then, simulation is run to obtain an initial solution. This step highlights the live/dead interfacing scheme between the fixed and rotating domains without mesh movement. Finally the unsteady simulation performs interpolations at each time step with node movement until a periodic steady state is reached. Mesh movement can be treated by either an ALE formulation or a rotating frame of reference correction. Two test cases are used to validate the code: a two-dimensional pitching airfoil in transonic flow and a 3-bladed 5868-9 propeller with a liquid cooled nacelle.

Flight mechanical simulators play an important role in the design steps during development of a new aircraft. To be able to simulate and evaluate flight mechanical characteristics during development it is important to minimize development time and cost while keeping flight safety high during early flights. The aim of the project presented in this thesis is to develop a method for validating a flight mechanical simulator against flight test data from dynamic maneuvering. An important part in this thesis is about how deviations in the result data can be found and analyzed, for example deviations between aircraft individuals or store configurations. The work presented here results in a good model for comparison of a big amount of data where it is easy to backtrace where the deviation occurs.
Flygmekaniska simulatorer är av stor betydelse under utvecklingen av ett nytt stridsflygplan. Möjligheten att simulera och utvärdera under tidens gång har stor betydelse både ur tid- och kostnadsbesparings perspektiv men även ur flygsäkerhetsperspektiv när det är dags för första flygning. Syftet med det här projektet är att utveckla en metod för jämförelse mellan simulering och flygprov för att validera hur bra den flygmekaniska simulatorn kan förutspå flygplansbeteende. En viktig del i projektet syftar till hur skillnader i resultaten kan hittas och analyseras, till exempel skillnader mellan olika flygplansindivider eller lastkonfigurationer. Arbetet presenterat här har resulterat i en modell som är bra för jämförelse av en stor mängd data där det är enkelt att spåra var skillnaderna har uppstått.

Since its invention at the University of Stuttgart, Germany in the mid-1960, scientists have been trying to understand and explain the mechanism of the plasma interaction inside the magnetoplasmadynamics (MPD) thruster. Because this thruster creates a larger level of efficiency than combustion thrusters, this MPD thruster is the primary cadidate thruster for a long duration (planetary) spacecraft. However, the complexity of this thruster make it difficult to fully understand the plasma interaction in an MPD thruster while operating the device. That is, there is a great deal of physics involved: the fluid dynamics, the electromagnetics, the plasma dynamics, and the thermodynamics. All of these physics must be included when an MPD thruster operates.

In recent years, a computer simulation helped scientists to simulate the experiments by programing the physics theories and comparing the simulation results with the experimental data. Many MPD thruster simulations have been conducted: E. Niewood et al.[5], C. K. J. Hulston et al.[6], K. D. Goodfellow[3], J Rossignol et al.[7]. All of these MPD computer simulations helped the scientists to see how quickly the system responds to the new design parameters.

For this work, a 1D MPD thruster simulation was developed to find the voltage drop between the cathode and the plasma regions. Also, the properties such as thermal conductivity, electrical conductivity and heat capacity are temperature and pressure dependent. These two conductivity and heat capacity are usually definded as constant values in many other models. However, this 1D and 2D cylindrical symmetry MPD thruster simulations include both temperature and pressure effects to the electrical, thermal conductivities and heat capacity values interpolated from W. F. Ahtye [4]. Eventhough, the pressure effect is also significant; however, in this study the pressure at 66 Pa was set as a baseline.

The 1D MPD thruster simulation includes the sheath region, which is the interface between the plasma and the cathode regions. This sheath model [3] has been fully combined in the 1D simulation. That is, the sheath model calculates the heat flux and the sheath voltage by giving the temperature and the current density. This sheath model must be included in the simulation, as the sheath region is treated differently from the main plasma region.

For our 2D cylindrical symmetry simulation, the dimensions of the cathode, the anode, the total current, the pressure, the type of gases, the work function can be changed in the input process as needed for particular interested. Also, the sheath model is still included and fully integrated in this 2D cylindrical symmetry simulation at the cathode surface grids. In addition, the focus of the 2D cylindrical symmetry simulation is to connect the properties on the plasma and the cathode regions on the cathode surface until the MPD thruster reach steady state and estimate the plasma arc attachement edge, electroarc edge, on the cathode surface. Finally, we can understand more about the behavior of an MPD thruster under many different conditions of 2D cylindrical symmetry MPD thruster simulations.

The resurgence of airships has created a need for dynamics models and simulation capabilities of these lighter-than-air vehicles. The focus of this thesis is a theoretical framework that integrates the flight dynamics, structural dynamics, aerostatics and aerodynamics of flexible airships. The study begins with a dynamics model based on a rigid-body assumption. A comprehensive computation of aerodynamic effects is presented, where the aerodynamic forces and moments are categorized into various terms based on different physical effects. A series of prediction approaches for different aerodynamic effects are unified and applied to airships. The numerical results of aerodynamic derivatives and the simulated responses to control surface deflection inputs are verified by comparing to existing wind-tunnel and flight test data. With the validated aerodynamics and rigid-body modeling, the equations of motion of an elastic airship are derived by the Lagrangian formulation. The airship is modeled as a free-free Euler-Bernoulli beam and the bending deformations are represented by shape functions chosen as the free-free normal modes. In order to capture the coupling between the aerodynamic forces and the structural elasticity, local velocity on the deformed vehicle is used in the computation of aerodynamic forces. Finally, with the inertial, gravity, aerostatic and control forces incorporated, the dynamics model of a flexible airship is represented by a single set of nonlinear ordinary differential equations. The proposed model is implemented as a dynamics simulation program to analyze the dynamics characteristics of the Skyship-500 airship. Simulation results are presented to demonstrate the influence of structural deformation on the aerodynamic forces and the dynamics behavior of the airship. The nonlinear equations of motion are linearized numerically for the purpose of frequency domain analysis and for aeroelastic stability analysis. The results from the latter for
L'intérêt renouvelé envers les dirigeables a créé un besoin de modèles dynamique et de simulations de ces véhicules plus légers que l'air. Cette thèse traite d'un cadre théorique qui intègre la dynamique de vol, la dynamique structurale, l'aérostatique et l'aérodynamique des dirigeables flexibles. La recherche débute par une étude d'un modèle dynamique fondé sur l'hypothèse d'un corps rigide. Une approche de calcul d'aérodynamique complète est présentée, où les forces et les moments aérodynamiques sont classés par catégories basées sur différents effets physiques. Une série d'approches de prédiction des différents effets aérodynamiques est unifiée et appliqué aux dirigeables. Les résultats numériques des dérivés aérodynamiques et des réponses simulées à des commandes spécifiés sont comparés à des résultats d'essais provenant d'autre œuvres. Une fois l'aérodynamique et le modèle de corps rigide validés, les équations de mouvement d'un dirigeable élastique sont dérivées avec une formulation Lagrangienne. Le dirigeable est modélisé comme poutre Euler-Bernoulli et les déformations sont représentées par des fonctions de forme choisies. Afin de prendre en considération la dépendance entre les forces aérodynamiques et l'élasticité structurale, la vitesse locale sur le véhicule déformé est employée dans le calcul des forces aérodynamiques. En conclusion, avec les forces d'inertie, de gravité, d'aérodynamique et de commande incorporées, le modèle dynamique d'un dirigeable flexible est exprimé sous la forme d'un ensemble d'équations différentielles ordinaires non-linéaires. Le modèle proposé est mis en pratique sous forme de simulation dynamique afin d'analyser les caractéristiques dynamiques du dirigeable Skyship-500. Des résultats de simulation sont présentés pour démontrer l'influence de la déformation structurale sur les forces aérodynamiques et le comportement dynamique du di

The use of near Earth space has increased dramatically in the past few decades, and operational satellites are an integral part of modern society. The increased presence in space has led to an increase in the amount of orbital debris, which poses a growing threat to current and future space missions. Characterization of the debris environment is crucial to our continued use of high value orbit regimes such as the geosynchronous (GEO) belt. Objects in GEO pose unique challenges, by virtue of being densely spaced and tracked by a limited number of sensors in short observation windows. This research examines the use of a new class of multitarget filters to approach the problem of orbit determination for the large number of objects present. The filters make use of a recently developed mathematical toolbox derived from point process theory known as Finite Set Statistics (FISST). Details of implementing FISST-derived filters are discussed, and a qualitative and quantitative comparison between FISST and traditional multitarget estimators demonstrates the suitability of the new methods for space object estimation. Specific challenges in the areas of sensor allocation and initial orbit determination are addressed in the framework. The sensor allocation scheme makes use of information gain functionals as formulated for FISST to efficiently collect measurements on the full multitarget system. Results from a simulated network of three ground stations tracking a large catalog of geosynchronous objects demonstrate improved performance as compared to simpler, non-information theoretic tasking schemes. Further studies incorporate an initial orbit determination technique to initiate new tracks in the multitarget filter. Together with a sensor allocation scheme designed to search for new targets and maintain knowledge of the existing catalog, the method comprises a solution to the search-detect-track problem. Simulation results for a single sensor case show that the problem can be solved for multiple objects with no a priori information, even in the presence of missed detections and false measurements. Collectively, this research seeks to advance the capabilities of FISST-derived filters for use in the estimation of geosynchronous space objects; additional directions for future research are presented in the conclusion.

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.

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.

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.

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.

In-flight variable stability aircraft and in-flight simulation are described. The uses of these vehicles and the associated requirements are described. Several forms of control architecture are identified for use in a T-6A in-flight simulator. A non-linear model of the T-6A Texan II is developed for use in MATLAB/SIMULINK®. This model is used to design a feedforward response-feedback controller, based on simplified dynamic inversion. This controller is shown to exercise precise control over the T-6A host aircraft dynamics. This architecture is then used to demonstrate simulation of the A-4 and the F-15 by the T-6A. In addition to proving simplified dynamic inversion for in-flight simulation, it is shown that the same configuration is useful in handling qualities training of military test pilots.

Reynolds averaged Navier Stokes (RANS) solvers have become the workhorse for simulating turbulent flows for most practical purposes. While the incompressible turbulence models used with RANS equations have improved considerably in their predictive capability, significant breakthrough has not been achieved for their compressible counterparts. With the advancement in computing power, high resolution direct numerical simulation (DNS) of low Reynolds number turbulent flows has become feasible. DNS of simple turbulent flows provides a detailed database which can be used for developing and testing turbulence models. In this work, we perform DNS of compressible homogeneous turbulence—decaying isotropic turbulence and homogeneous shear flow—for a range of initial turbulent Mach numbers, (M_{t 0} = 0.05–0.4) using the more natural initial conditions. Simulations were performed on grids with 128³ and 256³ points. Compressibility effects on the evolution of turbulent kinetic energy were studied. We found negligible compressibility effects for decaying isotropic turbulence, while homogeneous shear flow demonstrated compressibility effects in the growth rate of turbulent kinetic energy. Compressibility corrections to turbulence models in the form of the ratio &epsis;_d/&epsis; _s, have been tested with the results from the simulations. For decaying isotropic turbulence a [special characters omitted] scaling was found to be better than [special characters omitted] while for homogeneous shear flow it was the opposite. The small value of the ratio &epsis;_d/&epsis;_s in decaying isotropic turbulence makes the [special characters omitted] scaling less relevant. Based on the DNS results of homogeneous shear flow, a new correction parameterized by the gradient Mach number, M_g, is proposed. The parameter C_μ, which is assumed constant for incompressible two equation eddy viscosity models, is computed explicitly from the DNS data. An M_g, dependence of the parameter, C_μ, is proposed.

A six-degree-of-freedom piecewise simulation of aircraft motion is developed in SIMULINK. Using a mathematical model of fixed-wing aircraft, the simulation is used to observe the longitudinal and lateral-directional motions of the aircraft following a pilot input. The mathematical model is in state-space form and uses aircraft stability and control derivatives calculated from the aircraft geometric and aerodynamic characteristics. The simulation takes into account the changed speed and altitude due to pilot input and demonstrates the non-linearity of the aircraft motion due to change in speed and altitude. The results from the simulation are compared with the known results to validate the mathematical model used. The simulation is carried out for a number of airspeed and altitude combinations to examine the effect of changing speed and altitude on the aircraft dynamic response.

The objective of the present study was to investigate the mechanisms of sound generation by subsonic jets. Large eddy simulations were performed along with bandpass filtering of the flow and sound in order to gain further insight into the pole of coherent structures in subsonic jet noise generation.A sixth-order compact scheme was used for spatial discretization of the fully compressible Navier-Stokes equations. Time integration was performed through the use of the standard fourth-order, explicit Runge-Kutta scheme. An implicit low dispersion, low dissipation Runge-Kutta (ILDDRK) method was developed and implemented for simulations involving sources of stiffness such as flows near solid boundaries, or combustion. A surface integral acoustic analogy formulation, called Formulation 1C, was developed for farfield sound pressure calculations. Formulation 1C was derived based on the convective wave equation in order to take into account the presence of a mean flow. The formulation was derived to be easy to implement as a numerical post-processing tool for CFD codes.Sound radiation from an unheated, Mach 0.9 jet at Reynolds number 400, 000 was considered. The effect of mesh size on the accuracy of the nearfield flow and farfield sound results was studied. It was observed that insufficient grid resolution in the shear layer results in unphysical laminar vortex pairing, and increased sound pressure levels inthe farfield. Careful examination of the bandpass filtered pressure field suggested that there are two mechanisms of sound radiation in unheated subsonic jets that can occur in all scales of turbulence. The first mechanism is the stretching and the distortion of coherent vortical structures, especially close to the termination of the potential core. As eddies are bent or stretched, a portion of their kinetic energy is radiated. This mechanism is quadrupolar in nature, and is responsible for strong sound radiation at aft angles. The second sound generation mechanism appears to be associated with the transverse vibration of the shear-layer interface within the ambient quiescent flow, and has dipolar characteristics. This mechanism is believed to be responsible for sound radiation along the sideline directions.Jet noise suppression through the use of microjets was studied. The microjet injection induced secondary instabilities in the shear layer which triggered the transition to turbulence, and suppressed laminar vortex pairing. This in turn resultedin a reduction of OASPL at almost all observer locations. In all cases, the bandpass filtering of the nearfield flow and the associated sound provides revealing details of the sound radiation process. The results suggest that circumferential modes are significant and need to be included in future wavepacket models for jet noise prediction.Numerical simulations of sound radiation from nonpremixed flames were also performed. The simulations featured the solution of the fully compressible Navier-Stokes equations. Therefore, sound generation and radiation were directly captured in the simulations. A thickened flamelet model was proposed for nonpremixed flames. The model yields artificially thickened flames which can be better resolved on the computational grid, while retaining the physically currect values of the total heat released into the flow. Combustion noise has monopolar characteristics for low frequencies. For high frequencies, the sound field is no longer omni-directional. Major sources of sound appear to be located in the jet shear layer within one potential core length from the jet nozzle.
L'objectif de cette étude est d'obtenir la meilleure compréhension des mécanismesde géneration de bruit par des jet subsoniques. Cette étude est basée sur simulations aux grandes échelles de jets réactifs et sans réactifs. Des calculs numériques employant des schéme compacts de sixiéme ordre. L'integration temporelle fut éxéciteé à l'aide de schéme Runge-Kutta de de quatrième ordre. Des schéme à faible dispersion et dissipation numérique. Un formulation intégrale basée sur les analogies acoustiques fut développées pour la prédiction du champ acoustique lointain pour les sources et observateure en mouvement dans un fluide avec vitesse uniforme. La formulation fut implémentée à l'aide d'algorithmes facilitant l'implémentation pour le traitement de données d'écoulement en haute performance utilisant des outils de simiulation à grande échelle. Les champs sonore produit par un jet turbulent non-réactif avec nombre de Mach de 0.9, et un nombre de Reynolds ReD = 400, 000 fut étudié. L'effect de la taille du maillage sur la précision de l'écoulement en champs proche et e champs sonore loin de source fut analysé. La sous-résolution de la couche de cisaillement à la sortie du jet méne à l'apparition de structures cohérentes et forte radiation qui ne sort pas physiquement réalistes. Deux mécanismes principaux de génération sonore par jets subsoniques furent identifiés. Le premier mécanisme est l'étirement et la distorsion de structures tourbillonnaires cohérentes, en particulier prés de la fin du coere potentiel. Ce mécanisme est quadripolaire, et émet principalement vers l'arriére du jet dans la direction de l'écoulement. Le seconde mécanisme semble être constitué de vibration transversale de la couche de cisaillement en réponse à la présemce de structures cohérentes dans la jet. Semblable à la radiation d'une plaque à bonds finis, la contribution de ce méchanisme est dipolaire et domine la champs sonore dans la direction transversale, perpendiculaire au jet. L'utilisation de plusieurs microjet fut investiguée pour la réduction du bruit. L'injection à l'aide de microjets précipite la transition à la turbulence, favorisent le mélange et la destrcutction de structures cohérentes de grande échelle. Un filtrage en bandes de étroites fut effectué. Ce traitement des données numériquepermet de visualiser les relations complexes entre l'écoulement et les onds sonores émises. Les résultats démontrent l'importance de modes circumférenciele, ce qui a des conséquenecs pour les modiles dits de paquets d'onde pour la preédiction du bruit du jet. Des simulation numériques d'écoulement et champs sonore d'une flame sans prémélange furent aussi éxécutées. Les simulations incluent encore une fois l'écoulement et le champ sonore associé, obtenus directement des équations de Navier Stokes compressibles. Un modèle flammelette épaissie fut proposé que donne flammes épaissies artificiellement qui peuvent être mieux résolus sur le maillage. Le bruit de combustion a des caractéristiques monopolaires aux basses fréquences. Principales sources de bruit semblent être situé dans la couche de cisaillement.

A comprehensive survey of available numerical methods and models was performed on the open source computational fluid dynamics solver OpenFOAM version 2.0 for incompressible turbulent bluff body flows. Numerical methods are illuminated using source code for side-by-side comparison. For validation, the accuracy of flow predictions over a sphere in the subcritical regime and delta wing with sharp leading edge is assessed. Solutions show mostly good agreement with experimental data and data obtained from commercial software. A demonstration of the numerical implementation of a dynamic hybrid RANS/LES framework is also presented, including results from test studies.

Two explicit time-integration schemes based on a finite-volume approach for the solution of the Euler equations are developed and used in the study of compressible flows. The starting point is a comparison of the performance of three widely used methods (i.e., Jameson's, MacCormack's and Godunov's) in several rather difficult test problems, characterized by the existence of flow discontinuities or strong nonlinearities. This indicates that the best solutions for such flows are obtained when the numerical method is closely related to the physical behaviour of the fluid, as is the case with Godunov's method, in contrast with the other two methods, which need a special treatment of the discontinuities, and are very prone to numerically induced oscillations. Therefore, a first scheme, which improves the way Jameson's method computes the flux-node variables in that it treats in a more realistic manner the physics of signal propagation in both subsonic and supersonic flow, is developed. The numerical experiments with this scheme suggest that it converges more rapidly and does not need the dissipation terms, thus leading to computer efficiency and a gain in accuracy. The second method is a linear hybrid, in conservative form, between MacCormack's and Godunov's methods, which is shown to keep the best features of both the methods: second order accuracy in smooth regions of the flow and lack of oscillations near discontinuities, where it behaves locally like a first-order monotone scheme.

Composite sandwich plates are widely used in aerospace, automobile and shipbuilding industries. Composite sandwich plates have many different types of failure modes. A comparative study of composite sandwich plates with different finite element modeling approaches for predicting buckling and wrinkling failure response is described in this thesis. The research considers composite sandwich plates with isotropic and anisotropic face-sheets with a thick core. Finite element solutions are obtained using Abaqus/CAE 2016 software by conventional shell element models and conventional shell/solid element models. This study investigates results obtained using finite element methods and compares them to experimental and analytical solutions for overall buckling and face-sheet wrinkling. Results of the study indicate that finite element methods provide an accurate and effective modeling approach for predicting both overall buckling and wrinkling response.

Furthermore, the study also explored buckling response of composite sandwich panels with different core thickness and face-sheet fiber angle orientation. The study found that the shell/solid element model provides an appropriate and effective modeling method to predict both overall buckling and local wrinkling behavior in composite sandwich plates.

The main objective of this thesis is to implement a suitable open-source, density-based solver to simulate the high-speed compressible flows coming out of an aerospike rocket nozzle at different altitudes. The stability of the solver is demonstrated through an assessment of the calculated thrust and the degree of accuracy in the resolution of shocks and other critical compressibility effects. The method that will be used to design the contour has been devised by Angelino, this approach assumes a series of centred, isentropic expansion waves originating at the throat section of the aerospike nozzle and is based on the Prandtl-Meyer function. A recently developed coupled density-based solver on the foam-extend release of OpenFOAM, dbnsFoam (Jasak et al., 2014) has been chosen to simulate the flow field of interest. The solver is based on the HLLC Approximate Riemann Solver algorithm for the solution of the Euler Equations. Therefore, the solver is inviscid, but it represents a perfect starting point, especially to determine a proper set of boundary conditions. Additionally, there is also a viscous version of the solver, called dbnsTurbFoam, that will be useful for future works when viscosity and a model for turbulence will be introduced in the calculations. The HLLC Approximate Riemann Solver managed to obtain a solution for the presented cases. It turned out to be fairly precise in terms of accuracy of the produced thrust and in representing the characteristic waves and other flow features sought in the simulation. Nevertheless, the solver is slow and prone to instabilities when the shocks present in the flow are too strong. In fact, simulations with ambient pressures lower than 5000 Pa didn’t result in a converged solution.

McDaniel, Keith S. Three Dimensional Simulation of Time-DependentScramjet Isolator / Combustor Flowfields Implemented onParallel Architectures, ( Under the directions of Dr. J. R. Edwards). The development of a parallel Navier-Stokes solver for computing time-dependent,three-dimensional reacting flowfields within scramjet (supersonic combusting ramjet)engines is presented in this work. The algorithm combines low-diffusion upwinding methods, timeaccurate implicit integration techniques, and domain decomposition strategies to yield an effectiveapproach for large-scale simulations. The algorithm is mapped to a distributed memoryIBM SP-2 architecture and a shared memory Compaq ES-40 architecture using the MPI-1 message-passingstandard. Two and three-dimensional simulations of time-dependent hydrogen fuel injection into a modelscramjet isolator / combustor configuration at two equivalence ratios are performed. Thesesimulations are used to gain knowledge of engine operability, inlet performance, isolatorperformance, fuel air mixing, flame holding, mode transition, and engine unstart.Results for an injection at a ratio of 0.29 show qualitative agreement withexperiment for the two-dimensional case, but revealed a slow progression towardengine unstart for the three-dimensional case. Injection at an equivalence ratio of 0.61resulted in engine unstart for both two-dimensional and three-dimensional cases.Engine unstart for the three-dimensional case occurs as a response to the formation and growthof large pockets of reversed flow along the combustor side wall. These structuresdevelop at an incipient pressure above 154 kPa and result in significant blockage of the core flow,additional compression, and chemical reaction within the boundary layer. All of these factors promotea much more rapid unstart as compared with the two-dimensional case.

Research is presented on the development and testing of a new procedure for the time dependent, spatially varying numerical simulation of wall bounded turbulent flows. The Flow Simulation Methodology (FSM), as it is now known, was originally proposed by Speziale (1996a) for the purpose of computing complex, non-equilibrium flows which are currently beyond the reach of Smagorinsky based Large-Eddy Simulations (LES). The new method represents a hybrid approach that combines favorable aspects of Reynolds stress modeling [used for Reynolds Averaged Navier-Stokes (BANS) calculations] with the underlying principles of LES. For instance, Reynolds stress models developed for non-equilibrium, anisotropic, and/or rotational flows can be utilized in the unsteady manner of LES, i.e. where the flow field is decomposed into resolved-scale (calculated) and subgrid-scale (modeled) components, thereby reducing computational requirements. The key to the FSM is a contribution function which provides a degree of local turbulence modeling that is dependent upon the ratio of the numerical resolution to the Kolmogorov length-scale, an estimate for the smallest scales of turbulent motion. With this approach, a calculation resolved to the level of a Direct Numerical Simulation (DNS) can proceed continuously to a Reynolds Averaged Navier-Stokes calculation as the numerical resolution is decreased and/or the Reynolds number is increased. In between these two limits, an "untraditional" LES is recovered. The method is untraditional because it replaces the commonly employed Smagorinsky subgrid-scale model, which is known to have considerable limitations, with a more capable Reynolds stress model. A detailed evaluation of the Flow Simulation Methodology is made for the test case of a transitional and turbulent flat plate boundary layer with zero pressure gradient. The relatively simple geometry is chosen because the technical issues associated with combining elements of RANS calculations and LES must be established and the FSM itself must be validated before more complex flows can be attempted. The Reynolds stresses needed for the new method are computed using the two-equation Algebraic Stress Model (ASM) of Gatski & Speziale (1993) developed for non-equilibrium turbulent flows. Results of FSM calculations are compared with results obtained from coarse grid DNS, traditional LES based on the Smagorinsky subgrid-scale model, and RANS, all of which are implemented using an identical core computer code. This approach is extremely valuable to the evaluation of the FSM since a common code allows for certain behaviors to be more easily attributed to the turbulence models as opposed to numerical effects. Further validation is achieved through comparisons of FSM results with various direct numerical simulations and experiments available in the literature.

Modern and near-future Solar Electric Propulsion capabilities enable many new missions that were inconceivable using chemical propulsion systems. Many of these involve highly complex trajectories that are very challenging to design. New tools are needed that effectively utilize the rapidly growing parallel processing capabilities of modern computers. This research improves Gauss-Lobatto collocation methods, which are known to perform very well for low-thrust trajectory optimization, by formulating them as massively parallel processes. The parallelized elements of the problem formulation execute up to 11 times faster, depending on what force model is used and when evaluated by themselves. When accounting for the operations of the nonlinear programming solver, this translates to up to 3.7 times faster performance for solving a complete trajectory optimization problem, again depending on the force model that is used. The remaining barriers to further performance improvements, and the conditions upon which these depend, are clearly identified.

The implemented methods are combined into an optimization tool named Maverick. More general improvements to the formulation of the Gauss-Lobatto collocation methods are also developed and included in Maverick, which permit a more flexible use of these optimization schemes and enable them to find more complex solutions. One example of this is Maverick's ability to autonomously introduce gravity assists into trajectories, which greatly increases the utility and convergence radius of these methods.

In order to demonstrate the benefit of this work, three applications are studied. The first are transfers between halo-like orbits in the Earth-Moon system, which shows this is likely an unattractive region for missions like the New Worlds Observer. The second application investigates stabilization maneuvers in lunar distant retrograde orbits. This work demonstrates the feasibility of these stabilization transfers for a variety of sample return missions, such as the upcoming Asteroid Redirect Mission. The final application discussed is a series of multi-body low-thrust transfers from the Earth to the Moon that efficiently utilize highly variable dynamics to reduce propellant consumption, which is relevant for a variety of future mission concepts. These are computed for a wide range of flight times, showing that reductions up to 45% of the transfer time can be achieved with a propellant consumption as little as 0.5% of the total spacecraft mass. Up to 90% of the flight time can be eliminated for a propellant cost of 4% of the total spacecraft mass, or up to 83% for a propellant cost of less than 2%. The developed algorithm seamlessly transitions its solutions from full low-thrust, low-energy trajectories to the 'pure' low-thrust trajectories that define the shortest transfer trajectories, validating its robust performance. Beyond these quantifiable results, these examples illustrate the complexity of the solutions that can be identified with these improved implementations of Gauss-Lobatto collocation methods, with many instances where the optimization method autonomously introduces powered gravity assists, an unusual capability that has the potential for useful application to many other trajectory optimization problems.

Flight-testing represents an important step of an airplane development. Every new or modified aircraft configuration is tested. Through set of tests, the quality and reliability of aircrafts are guaranteed. At Dassault Aviation, the flight test Directorate responsible for carrying out flight tests is located at Istres, in South of France. One major test before flying is the Ground Vibration Test (GVT). The aim of this test is to measure Eigen frequencies and mode shapes of the structure. Those results are after compared to the ones given by the finite element model in order to verify it or update it. Determining the structural behavior has a significant importance for aircraft safety, for instance, it helps to determine the aircraft’s flutter boundary. Under unsteady aerodynamic loads, the aircraft structure can be dynamically unstable, meaning that the amplitude of oscillations increases with time. This phenomenon called flutter can highly damage the airplane or can even lead to the complete destruction of its structure. GVTs are performed by exciting the aircraft with oscillatory forces on designated parts. Then, hundreds of accelerometers are used to measure the vibrations. Knowing exactly the input excitation and how the structure answers, it is possible to calculate the relations (so-called transfer functions) between the applied oscillatory forces and the acceleration measurements. This measurement method based on accelerometers is highly accurate; however, one accelerometer only provides one punctual measurement. In order to have an infinite number of measurement points and thus, a better understanding of the mode shapes, new measurement methods should be investigated. As part of Ground Vibration Testing, this paper presents an investigation on new innovative measurement methods that could improve and complement the current methods based on accelerometers. This report is structured in three parts; the first part gathers and presents some innovative measurement methods and the two following parts focus on experimentations of two measurement methods using rapids cameras and 3D laser scanners.
Flygtestning är ett viktigt steg av att utveckla ett flygplan. Genom en uppsättning tester blir varje ny eller modifierad flygkonfiguration certifierad. Certifiering garanterar säkerhet, kvalitet och pålitlighet. Avdelningen för flygtestning på Dassault Aviation är ansvarig för genomförandet av flygtester I Istres, södra Frankrike. Ett viktigt test före flygning är Ground Vibration Test (GVT). Målet med detta test är att mäta strukturens egenfrekvenser och lägesformer. Resultaten jämförs sedan med dem som ges av den finita elementmodellen för att kunna verifiera eller uppdatera dem. Att fastställa det strukturella beteendet är av stor vikt för flygplanets säkerhet; till exempel hjälper det att stipulera planets flygstabilitet och fladdergränsen. Under ostadiga aerodynamiska belastningar kan flygplanets struktur vara dynamiskt instabil, vilket innebär att svängningens amplitude ökar med med tiden. Resonansfenomenet som kallas för fladder kan allvarligt skada eller till och med leda till total förstörelse av flygplanets struktur. GVT: er utförs genom att utsätta flygplanet för svängande krafter på utsedda delar av planet. Sedan används hundratals accelerometrar för att mäta deformationerna. Genom att veta exakt inmatningsexitering och hur strukturen svarar, är det möjligt att beräkna förhållandena (de så kallade överföringsfunktioner) mellan de applicerade oscillerande krafterna och accelerationsåtgärderna. Denna mätmetod baserad på accelerometrar är mycket exakt; installationen tar emellertid mycket tid och skapar en viktig överbelastning på grund av ett stort antal accelerometrar och kablar som används. Som en del av markvibrationstest presenterar detta dokument en utredning om nya innovativa mätmetoder som kan förbättra, komplettera eller till och med ersätta de nuvarande metoderna baserade på accelerometrar. Denna rapport är strukturerad i tre delar; en kort konstnärligt samling där det presenteras några innovativa mätmetoder och sedan två delar med fokus på experiment av två mätmetoder med hjälp av rapids kameror och 3D-laser. Arbetet är ett första steg i en lång forskning som säkert på några år kommer att avföra accelerometrar och ersätta dem med nya metoder som är mycket bekvämare.

This work has as an objective to structure the content of a course on Attitude determination methods, part of an Aerospace Engineering Master program. A selection of books, papers, theses, web sites and films was reviewed to identify the most relevant topics within the areas of Static and Dynamic Attitude Determination and the ways to present them in a educational context. Theory is presented in a simplified way and examples were gathered to illustrate the theoretical part.  Finally, a discussion is carried out on the main learning goals and challenges, required time for instruction and exercises and suggestion for a grading system.​
Detta arbete har som mål att strukturera innehållet i en kurs om Attitydbestämningsmetoder inom flyg- och rymdteknikmastersprogram. Ett urval av böcker, artiklar, avhandlingar, webbsidor och filmer granskades för att identifiera de mest relevanta ämnena inom statisk och dynamisk attitydbestämning och de olika sätten att presentera dem i ett utbildningssammanhang. Teorin presenteras på ett förenklat sätt och några exemplar visas för att illustrera den teoretiska delen. Avslutningsvis, diskuteras de huvudsakliga lärandemålen, nödvändig handledning och övningstid, samt betygsättning.

This work uses computational fluid dynamics to study the flowfield around a hypersonic missile with two lateral jets to provide control in place of control surfaces. The jets exhaust an H₂-O₂ mixture at Mach number of 2.9 with a jet pressure ratio of roughly 10,500. The jets are staggered axially and circumferentially in such a way to produce pitch and yaw. The flowfield of such a jet configuration is characterized at several angles of attack and the corresponding force coefficients and amplification factors are provided. The freestream air and H₂-O₂ plume is treated as inert for the majority of the calculations. Special cases are treated with finite rate chemical kinetics and compared to the inert flowfield to ascertain the effects that chemical reactions have on the force coefficients. It was found that the flowfield was only slightly altered from the familiar one jet flowfield when the second jet is active. The flow topology and vortex structures tend to shift towards the second jet but the overall structure remains the same. The normal force amplification factors are close to unity over the range of angle of attack due to the thrust being so high with the two jet configuration having a lower amplification factor compared to firing a single jet. Treating the flowfield as chemically reacting did not affect the force values much: the difference being 0.3% for an angle of attack of 0°.

Spacecraft project management calls for division of project lifetime into phases, with specific goals to be fulfilled at the end of each phase. During first few phases a Preliminary Design Review (PDR) has to be conducted, after which top-level hardware design is not to be changed. This thesis describes a process of creating and demonstrates a software framework supporting teams building small satellites - typically CubeSat student projects - during initial phases of conceptual design, mission planning, and selection and sizing of hardware components. The scope of the thesis covers review of available tools for satellite mission and control system design, then it proposes a self-made MATLAB/Simulink toolbox - Spacecraft Control Architecture Rapid Simulator (SCARS) Toolbox, as a open source tool with gentle learning curve and ease of reverse engineering approach. In further parts of the thesis examples of usage are provided, and conclusions and descriptions of problems are presented. In the end, this thesis should not only serve as a description of SCARS toolbox, but also as an insight into the task of building a small satellite simulation.

The purpose of this research has been to extend a previously developed one-equation model for transitional/turbulent flows (AIAA Journal, Vol. 39, No. 9) for use in the simulation of transitional/turbulent flows over three-dimensional bodies in conventional hypersonic tunnels. This is done computationally through the combination of the Spalart-Allmaras one-equation turbulence model and an eddy viscosity-transport equation based on that proposed by Xiao, Edwards, and Hassan for high disturbance environment (HIDE) induced transition. The blending of these two pieces of the model is achieved through the use of an intermittency function based on the work of Dhawan and Narasimha. The test case used in this research is an elliptic cone of aspect ratio 2:1 in a Mach 8 environment with Reynolds numbers between the range of 1.98x10⁶/ft and 6.09x10⁵/ft. Two separate methods are used to find the boundary layer edge flow properties under the resulting conical shock. The first of these methods uses fluid values extracted from the surface of the cone after an inviscid calculation. The second searches for the boundary layer edge by locating the largest momentum flux under the shock. The second of the two approaches is found to be the most successful in replicating transitional flow heat flux data measured experimentally by Kimmel, Poggie, and Schwoerk. Over the range of Reynolds numbers examined, the model reasonably predicts the location and extent of the transitional region, but does not effectively predict fluid properties within the transitional region.

The following document describes an example mission, which originated from a real life concept of an imaging satellite in a Sun Synchronous Orbit (SSO)around Earth. This report takes the reader through the thermal analysis and evaluation of space equipment performed with Airbus Space & Defence’s Systema/Thermica tool, at Space Structures GmbH. It details the full process of designing a thermal control system for a space project. The project started from a CAD file which was converted into a Geometric Mathematical Model (GMM) inside Thermica. This process requires an extensive knowledge of not only the software, but also the technical background behind what happens to a satellite in such an extreme environment. This thesis addresses this by showing a step-by-step approach of a full thermal evaluation, starting with the required theoretical background of the thermal environment and the different passive and active thermal design techniques. The next step involves gathering the required input information for the software; such as defining the conductance values between the components and calculating the per node power dissipation for each component considering each operational mode. The final step includes the designing, simulation, iteration and presentation of the temperature results across the spacecraft thermal model. The results of the initial simulation showed that some sensitive components were not within the specified temperature requirements, and therefore both radiators and heaters were sized and introduced to the model. After the third iteration of thermal control, the sensitive components’ temperatures were observed to be within the allowable margins of an ECSS Phase A study. This thesis can serve as a guide and complete document for future missions which plan the design of a Thermal Control System of a satellite in orbit around Earth.

Determination of aerodynamic coeffcients and stability derivatives is necessary in defning a model of the Orion CEV dynamics. This involves reducing experimental data, which can include acceleration, angular rate, or orientation data. This sort of extraction of dynamics from experimental data is often performed on data gathered from experiments conducted on uninstrumented models at indoor ballistics ranges. The US Army Research Laboratory (ARL) has developed a high-g survivable stand-alone instrumentation package that can transmit in-flight measurements of acceleration, angular rate, and local magnetic field. This telemetry module (TM) was installed in a scale model of the Orion CEV, which was red from a 175mm cannon at the ARL range. The instrumentation package was upgraded to include pressure transducers to measure forebody pressures. A minimum variance with a priori method was formulated to solve for both the "local" flight parameters of Mach number, angle of attack, and sideslip angle at each timestamp and the "global" parameters of scale factor and bias for each pressure transducer. Results using both simulated and experimental data indicates that these parameters may be estimated and used to compute stability coeffcients. Low pressure differentials between symmetrically-opposed pressure transducers, however, increased uncertainty in the parameter estimates. Validation of this method of data generation and analysis supports a low-cost method of vehicle testing.

This thesis reports on research into the causes of local optima when optimization algorithms are applied to aerospace structural design. A thorough understanding of local optima will enable the engineers to select the algorithm for optimization or to guide the optimization to ensure either global optima or near optimal solutions are achieved. Therefore, a comprehensive literature review has been conducted and several illustrative examples have been identified to help fully understand the cause and importance of local optima. The first application involved the design of the internal structure of a simplified wing spoiler. MSC.NASTRAN was used to optimize each discretized location of an additional rib with the aid of a Patran Command Language (PCL) algorithm. The objective function of minimum weight was approximated as a multimodal function in a 2D smooth curve where the local and global optima were identified. The theory of continuous rectangular plates was used to explain the phenomena. The second problem considered buckling of a wing rib. A PCL code was written to obtain the rib buckling factors as the position of the center of a square cutout was varied within a constrained area. The rib linear buckling factor versus the centre position O(X, Y) of the square cutout was plotted in a 3D surface contour plot. Load path theory and relevant plate buckling theories were used to explain the local and global maxima identified. The final example considered the maximization of the buckling load of a simply supported composite laminated plate under in-plane loading. A conventional Genetic Algorithm was used to examine the local and global optima of the critical buckling load factor. Many local and global optima were identified and explained and many near-optimal solutions were found in a single run. A significant understanding of local optima in aerospace structural design with the optimal utilization of available software and the appropriate selection of optimization algorithms has been achieved. Further work could either include implementing the proposed global optimization strategies or include implementing rapid methods for identifying multiple local optima.

This thesis developed a linear shaped charge (LSC) separation mechanism capable of severing the interstage skin for first stage separation of the Ares I launch vehicle. The derived LSC design solution was found using available data on Explosive Technologys Jetcord LSC and from National Aeronautics and Space Administration (NASA) Marshall Space Flight Centers (MSFC) desired characteristics. Mechanism components are designed after Minuteman IIIs separation mechanism for first stage separation and NASA MSFCs desired characteristics. Mechanism severance is verified through the use of the numerical method capability smoothed particle hydrodynamics that the hydrocode Autodyn offers. Three simulations are conducted to determine feasibility: the first of only the LSC exploding, to numerically validate the explosion process; the second of the LSC penetrating the target, to numerically validate the penetration process and failure mechanisms; and the last of the entire mechanism, to obtain information about the explosion, penetration, failure, and debris generated.

In this thesis, there are generally three contributions to the Ranking and Selection problem in discrete-event simulation area. Ranking and selection is an important problem when people want to select single or multiple best designs from alternative pool. There are two different types in discrete-event simulation: terminating simulation and steady-state simulation. For steady-state simulation, there is an initial trend before the data output enters into the steady-state, if we cannot start the simulation from steady state. We need to remove the initial trend before we use the data to estimate the steady-state mean. Our first contribution regards the application to eliminate the initial trend/initialization bias. In this thesis, we present a novel solution to remove the initial trend motivated by offline change detection method. The method is designed to monitor the cumulative absolute bias from the estimated steady-state mean. Experiments are conducted to compare our procedure with other existing methods. Our method is shown to be at least no worse than those methods and in some cases much better. After removing the initialization bias, we can apply a ranking and selection procedure for the data outputs from steady-state simulation. There are two main approaches to ranking and selection problem. One is subset selection and the other one is indifference zone selection. Also by employing directed graph, some single-best ranking and selection methods can be extended to solve multi-best selection problem. Our method is designed to solve multi-best ranking and selection. And in Chapter 3, one procedure for ranking and selection in terminating simulation is extended based full sequential idea. It means we compare the sample means among all systems in contention at each stage. Also, we add a technique to do pre-selection of the superior systems at the same time of eliminating inferior systems. This can accelerate the speed of obtaining the number of best systems we want. Experiments are conducted to demonstrate the pre-selection technique can save observation significantly compared with the procedure without it. Also compared with existing methods, our procedure can save significant number of observations. We also explore the effect of common random number. By using it in the simulation process, more observations can be saved. The third contribution of this thesis is to extend the procedure in Chapter 3 for steady-state simulation. Asymptotic variance is employed in this case. We justify our procedure in asymptotic point of view. And by doing extensive experiments, we demonstrate that our procedure can work in most cases when sample size is finite

Thesis: Ph. D., Massachusetts Institute of Technology, Department of Chemical Engineering, 2019
Cataloged from PDF version of thesis.
Includes bibliographical references.
This thesis describes the systematic development of methods to perform large scale dynamic simulations of hydrodynamically interacting colloidal particles undergoing Brownian motion. Approximations to the hydrodynamic interactions between particles are built from the periodic fundamental solution for flow at zero Reynolds number and are methodically improved by introducing the multipole expansion and constraints on particle dynamics. Ewald sum splitting, which decomposes the sum of slowly decaying interactions into two rapidly decaying sums evaluated indepently in real space and Fourier space, is used to accelerate the calculation and serves as the basis for a new technique to sample the Brownian displacements that is orders of magnitude faster than prior approaches. The simulation method is first developed using the ubiquitous Rotne-Prager approximation for the hydrodynamic interactions.
Extension of the Rotne-Prager approximation is achieved via the multipole expansion, which introduces the notion of induced force moments whose value is determined from the solution of constraint problems (for example, rigid particles cannot deform in flow), and methods for handling these multipole-based constraints are illustrated. The multipole expansion converges slowly when particles are nearly touching, a problem which is functionally solved for dynamic simulations by including divergent lubrication interactions, in the style of Stokesian Dynamics. The lubrication interactions effectively introduce an additional constraint on the relative motion of closely separated particle pairs. This constraint is combined with the multipole constraints by developing a general method to handle nearly arbitrary dynamic constraints using saddle point matrices. Finally, the methods developed herein are applied to study sedimentation in suspensions of attractive colloidal particles.
The simulation results are used to develop a predictive model for the hindered/promoted settling function that describes the mean sedimentation rate as a function of particle concentration and attraction strength.
"The research in this thesis was supported by the MIT Energy Initiative Shell Seed Fund and NSF Career Award CBET-1 554398"
by Andrew M. Fiore.
Ph. D.
Ph.D. Massachusetts Institute of Technology, Department of Chemical Engineering

Aerated-liquid atomization, which is produced by the introduction of gas directly into a liquid flow immediately upstream of the injector exit orifice to generate a two-phase flow, has been shown to produce well-atomized sprays in a quiescent environment with only a small amount of aerating gas at relatively low injection pressures. A time-derivative preconditioning method using the Low-Diffusion Flux-Splitting Scheme (LDFSS) has been extended to a ?mixture? model of two-phase flow and applied to simulate the structure of internal two-phase flow for aerated-liquid injectors, with each phase governed by its own equation of state. The Continuum Surface Force (CSF) model of Brackbill, et al. is adapted to model compressible fluid flow influenced by interfacial surface tension. A sub-iterative time integration method based on a planar Gauss-Seidel partitioning of the system matrix is used with implicit source terms as a means of solving the three-dimensional, time-dependent form of the governing equations. The calculations are parallelized using domain-decomposition and Message-Passing Interface (MPI) methods, and are optimized for operation on the 720 processor IBM SP-2 at the North Carolina Supercomputing Center (NCSC). Simulation results for 2-D aerated-liquid injector flowfields at gas-to-liquid (GLR) mass ratios of 0.08% and 2.45% are discussed. In accord with experimental visualization data, the results for GLR = 0.08% indicate a combination of slugging and core-annular two-phase flow in the injector. Results at GLR = 2.45% indicate that a core-annular flow mode dominates, again in agreement with experimental results. The effects of the choice of reference velocity and the level of surface tension on the injector flowfield solutions are also examined.

Thesis (MScEng)--Stellenbosch University, 2011.
ENGLISH ABSTRACT: The aerospace industry actively pursues innovation, especially in materials and their use in new applications, to improve their aircraft as well as their competitive position. Ti-6Al-4V has been available now for more than 50 years. Yet, in the new generation of aircraft using structural composites, a dramatic increase in the proportion of Ti-6Al-4V will be seen along with emerging application in automotive and chemical industries. This material possesses superior material properties compared to conventional materials such as steel and aluminium, although it is at the expense of machinability. Researchers are therefore actively searching for improved cutting technologies to improve production rates for Ti-6Al-4V. At higher cutting speeds than the industry norm of 60 - 90 m/min, machining becomes a challenge, resulting in low productivity on titanium parts. The limiting factor in the machining of Ti-6Al-4V is high tool temperatures of the order of 1000oC, caused by its resistance to absorb heat and good mechanical strength at elevated temperatures. The result is extreme temperatures that are concentrated on the cutting edge of the tool. The challenge to improve the tool life is therefore focused on removing heat from the insert. Liquid nitrogen was identified as a good candidate as coolant with the additional advantage of being environmentally friendly. The research presented investigates the use of a gravity feed enclosed liquid nitrogen cooling system to improve the tool life of the cutting inserts. The liquid nitrogen is contained on the insert rake face by means of a tool cap. To improve the effectiveness of the cooling method, a polycrystalline diamond (PCD) insert was used. This insert has a considerably higher thermal conductivity that aids in cooling the cutting edge. Tungsten carbide inserts are used for benchmark testing. The round tungsten carbide inserts with conventional cooling performed exceptionally well for machining titanium compared to square inserts, yielding exceptional tool life improvements while significantly increasing the material removal rate. Positive results were recorded with the liquid nitrogen cooling system when used with the polycrystalline diamond cutting insert. A number of far reaching performance issues are identified relating to the design of the tool cap that hindered clear scientific outputs. From a research perspective, the project makes a contribution to the knowledge base in this field. Additionally a new approach in cooling was investigated, resulting in clear indications of design changes required.
AFRIKAANSE OPSOMMING: Die lugvaart industrie streef aktief innovasie na, veral op die gebied van materiale en hul gebruike, om hul vliegtuie en kompeterende posisie in die mark te verbeter. Ti-6Al-4V is al vir meer as 50 jaar beskikbaar. ‘n Drastiese verhoging in die aanvraag na Ti-6Al-4V deur die lugvaart, motor en chemiese industrieë word verwag wanneer die volgende geslag vliegtuie wat koolstofvesel as strukturele materiaal begin gebruik, in produksie gaan. Die materiaal het beter materiaaleienskappe as konvensionele materiale soos staal en aluminium, maar dit kom egter teen die prys van masjieneerbaarheid. Ti-6Al-4V se masjienering bo die industrie norm van 60 – 90m/min is ‘n groot uitdaging. Navorsers soek daarom deurentyd na verbeterde sny tegnologieë om die produksie tempo van Ti-6Al-4V te verbeter. Die beperkende faktor vir Ti-6Al-4V masjienering is die temperatuur wat genereer word. Die weerstand van die materiaal om hitte te absorbeer en sy goeie meganiese eienskappe veroorsaak dat temperature in die beitel 1000oC bereik. Hierdie temperature word egter op die snykant van die beitel gekonsentreer. Die uitdaging is dus om hierdie temperature in die beitel te beheer. Vloeibare stikstof is geïdentifiseer as ‘n goeie kandidaat vir verkoeling met die bykomende voordeel dat dit omgewingsvriendelik is. Die navorsing wat hier uiteengesit word, ondersoek die gebruik van ‘n geslote kamer beitelverkoelingstelsel wat deur gravitasie met vloeibare stikstof voorsien word om die beitel leeftyd te verbeter. Die oppervlak van die beitel word in hierdie konsep direk aan die vloeibare stikstof blootgestel. Om die effektiwiteit van die stelsel te verbeter word van PCD beitels gebruik gemaak. Die beitel se verbeterde hittegeleidingsvermoë help om die beitel se snykant koel te hou. Tungstenkarbied beitels word gebruik om ‘n standaard te stel vir eksperimentele analise. Die ronde tungstenkarbied beitels en konvensionele verkoeling het verstommend goed presteer vir Ti-6A-4V masjienering in vergelyking met vierkantige beitels. Die materiaalverwyderingstempo is aansienlik verhoog sonder om die beitel se leeftyd in te boet. Positiewe resultate is waargeneem met die vloeibare stikstof sisteem saam met die PCD beitels. ‘n Aantal verreikende uitdagings is geïdentifiseer wat suiwer wetenskaplike afleidings bemoeilik. Hierdie probleme kan almal aan die ontwerp van die toerusting toegeskryf word. Die werk lewer egter steeds ‘n bydrae tot die kennis in die veld. ‘n Bykomende benadering vir verkoeling is ondersoek wat duidelik ontwerp-veranderings aandui.

In March 2019 the number of artificial objects bigger than 1 mm in orbit around the Earth is estimated to be more than 170 millions. Only a small fraction of them (0.03%) is catalogued. An impact of an operational satellite with one of these debris can damage the satellite and undermine its mission. So it is important to catalogue as many objects as possible in order to reduce the risk of a collisions. This is done by using the software tool Space Object Observations and Kalman Filtering (SPOOK), developed in Airbus Defence and Space in Friedrichshafen. The goal of this Master Thesis was to create newfunctionalities to this tool and improve the existing ones. In particular three main goals have been accomplished: • a new model for the lighting ratio has been built to take into account the occultation of the Sun due to a covering body (for example the Earth or the Moon) and itsinfluence on the solar radiation pressure, necessary to have a good model for orbit propagation; • a tracklet building algorithm has been built to distinguish different tracklets (consecutive observations of the same object along its orbit) as a starting point for the association of different measurements belonging to the same object at distant epochs, necessary to update a catalogue of space objects; • a model to take into account the process noise has been improved giving some suggestion on how to tune the different parameters for different kinds of orbit.
I mars 2019 uppgick antalet konstgjorda föremål större än 1 mm i omloppsbana runt jorden till över 170 miljoner. Av dessa är endast en mycket liten andel (0.03%) katalogiserade. En kollision mellan en operativ satellit och ett annat föremål i bana kan helt eller delvis förstöra satelliten. För att reducera risken för kollisioner är det därför viktigt att katalogisera så många föremål som möjligt. Detta görs genom att använda programvaran "Space Object Observations and Kalman Filtering" (SPOOK), som utvecklats av företaget Airbus Defence and Space i Friedrichshafen, Tyskland. Målet med detta examensarbete var att skapa nya funktioner i programvara samt att förbättra de befintliga funktionerna. Tre huvudmål har uppnåtts: • En ny modell för ljusförhållandet har skapats för att ta hänsyn till ocklutationen av solen på grund av en täckande kropp (till exempel jorden eller månen) och dess påverkan på solstrålningstrycket på rymdfarkosten, vilket är nödvändigt för att ha en bra modell för hur omloppsbanan fortplantas • En algoritm för s.k. tracklets, flera observationer av samma föremål längs dess omlopp, har skapats i syfte att skilja mellan olika tracklets som utgångspunkt för bestämma vilka mätningar som tillhör samma föremål vid avlägsna epoker. Detta är nödvändigt för att korrekt kunna uppdatera katalogen över rymdföremål. • Modellen för att ta hänsyn till processbruset har förbättrats och förslag ges om hur man ställer in olika parametrar för olika slags omloppsbanor.

Although it has been used for many years, an existing computer code, developed to simulate the laminar opposed-jet diffusion flame (LOJDF), was found not to be written in an user-friendly fashion. This was especially true for the portion dealing with calculation of thermochemical properties. The purpose of this research was to replace the appropriate portions of the existing program by the corresponding portions of the CHEMKIN package. CHEMKIN has become a recognized standard in inputing chemical kinetics data into program, since the inputing is almost format free and easy to manipulate. A series of test cases show that the updated code is now better structured, user-friendly, and ready to use. The previous LOJDF model, in addition, is modified by adding source terms for species generation in the governing equations. The source-contained LOJDF model has proven to be useful in evaluating the numerical relation of the fate of an impurity to its location and strength.

The design of NextGen and current-day cockpit displays are critical for efficient pilot performance and situation awareness on the flight deck. Before deployment of a design into the cockpit the costs and benefits that a display design imposes on performance and situation awareness should be considered. In this thesis, a design tool was developed to support the design of NextGen displays for situation awareness and performance. This design tool is a library of pilot performance estimates. Through literature reviews and meta-analyses of empirical data, the library was developed to provide display designers 1) qualitative distinctions of display properties that either support or limit full situation awareness, and 2) quantitative performance time estimates until situation awareness as a function of various display formats. A systematic method was also developed for future augmentation of the library.

The analysis of the reliability of CREO simulate will be done step by step. First of all, the precisionof the simulation has to be measured thus a comparison with simple theoretical computations willbe done. Then, the scope of the capacity of the software will be analyzed and if elements deemednecessary to thermal simulations are missing, back up solutions are to be found. Also, the influenceof the meshing will be studied and measured to ensure that the software guarantee convergence evenin the hand of persons unfamiliar with simulations.In parallel, one will experiment with realistic hardware that could be used to compare reality withthe simulations. Those experiments will be handmade using regular materials from the company.

Personal transportation has a large and increasing impact on people, society, and the environment globally. Computational energy-use simulation is becoming a key tool for automotive research and development in designing efficient, sustainable, and consumer acceptable personal transportation systems. Historically, research in personal transportation system design has not been held to the same standards as other scientific fields in that classical experimental design concepts have not been followed in practice. Instead, transportation researchers have built their analyses around available automotive simulation tools, but conventional automotive simulation tools are not well-equipped to answer system-level questions regarding transportation system design, environmental impacts, and policy analysis.

The proposed work in this dissertation aims to provide a means for applying more relevant simulation and analysis tools to these system-level research questions. First, I describe the objectives and requirements of vehicle energy-use simulation and design research, and the tools that have been used to execute this research. Next this dissertation develops a toolset for constructing system-level design studies with structured investigations and defensible hypothesis testing. The roles of experimental design, optimization, concept of operations, decision support, and uncertainty are defined for the application of automotive energy simulation and system design studies.

The results of this work are a suite of computational design and analysis tools that can serve to hold automotive research to the same standard as other scientific fields while providing the tools necessary to complete defensible and objective design studies.

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Electrical Engineering and Computer Science, 1992.
Includes bibliographical references (leaves 51-53).
by Jennifer Anne Lloyd.
M.S.

Thesis (M.S.)--University of Akron, Dept. of Civil Engineering, 2007.
"May, 2007." Title from electronic thesis title page (viewed 4/28/2009) Advisor, Wieslaw K. Binienda; Committee members, Craig C. Menzemer, Ala Abbas; Department Chair, Wieslaw K. Binienda; Dean of the College, George K. Haritos; Dean of the Graduate School, George R. Newkome. Includes bibliographical references.