A 3d parallel CFD code is written to investigate the characteristics of and differences between Large Eddy Simulation (LES) models in the context of simulating a thermal buoyant plume. An efficient multigrid scheme is incorporated to solve the Poisson equation, resulting from the fractional step, projection method used to solve the Low Mach Number (LMN) Navier-Stokes equations. A wide range of LES models are implemented, including a variety of eddy models, structure models, mixed models and dynamic models, for both the momentum stresses and the temperature fluxes. Generalised gradient flux models are adapted from their RANS counterparts, and also tested. A number of characteristics are observed in the LES models relating to the thermal plume simulation in particular and turbulence in general. Effects on transition, dissipation, backscatter, equation balances, intermittency and energy spectra are all considered, as are the impact of the governing equations, the discretisation scheme, and the effect of grid coarsening. Also characteristics to particular models are considered, including the subgrid kinetic energy for the one-equation models, and constant histories for dynamic models. The argument that choice of LES model is unimportant is shown to be incorrect as a general statement, and a recommendation for when the models are best used is given.

The problem of low Mach number (non-Boussin´esq) conjugate laminar natural convection combined with surface radiation in a vertical annulus with a centrally located vertical heat generating rod is studied numerically, taking into account the variable transport properties of the fluid. Such problems arise often in practical applications like spent nuclear fuel casks, cooling of electrical and electronic equipment, convection in ovens, cooling of enclosed vertical bus bars and underground transmission cables. The physical model consists of a vertical heat generating rod, a concentric outer isothermal boundary and adiabatic top and bottom surfaces. The heat generation in the rod drives the natural convection in the annulus. Surface radiation is coupled to natural convection through the solid-fluid interface condition and the adiabatic condition of the top and bottom surfaces. A mathematical formulation is written using the governing equations expressing the conservation of mass, momentum and energy for the fluid as well as the energy balance for the solid heat generating rod. The governing equations are discretized on a staggered mesh and are solved using a pressure-correction algorithm. Steady-state solutions are obtained by time-marching of the time dependent equations. The discretized equations for the dependent variables are solved using the Modified Strongly Implicit Procedure. A global iteration is introduced on the variables at each time step for better coupling. The parameters of the problem are the heat generation and gap width based Grashof number, aspect ratio, radius ratio and the solid-to-fluid thermal conductivity ratio. The coupling of radiation introduces the wall emissivity and the radiation number as the additional parameters and also necessitates the calculation of radiation configuration factors between the elemental surfaces formed by the computational mesh. The radiant heat exchange is calculated using the radiosity matrix method. A parametric study is performed by varying Grashof number from 106 to 1010 , aspect ratio from 1 to 15, radius ratio from 2 to 8, the solid-to-fluid thermal conductivity ratio from 1 to 100, with the Prandtl number 0.7 corresponding to air as the working medium. The characteristic dimension and the outer boundary temperature are fixed. For Radiative calculations, and the emissivity is varied between 0.25 and 0.75. Converged solutions with laminar model could be obtained for high Grashof numbers also as the heat generation based Grashof number is generally two orders of magnitude higher than the temperature difference based Grashof number. Results are presented for the flow and temperature distributions in the form of streamline and isotherm maps. Results are also presented for the variation of various quantities of interest such as the local Nusselt numbers on the inner and outer boundaries, the axial variation of the centerline and interface temperatures, maximum solid, average solid and average interface temperature variations with Grashof number and the average Nusselt number variation for the inner and outer boundaries with Grashof number. The results show that simplification of conjugate problems involving heat generation by the prescription of an isoflux boundary condition on the rod surface is inadequate because a truly isoflux condition cannot be realised on the one hand and because the solid temperature distribution remains unknown with such an approach. The average Nusselt numbers on the inner and outer boundaries show an increasing trend with the Grashof number. For pure natural convection, the Boussin´esq model predicts higher temperatures in the solid and lower average Nusselt numbers on the inner and outer boundaries, compared to the non-Boussin´esq model and the Boussin´esq approximation appears to be adequate roughly upto a Grashof number of 109, beyond which the non-Boussin´esq model is to be invoked. The average pressure in the annulus is found to increase with an increase in the Grashof number. Radiation is found to cause convective drop and homogenize the temperature distribution in the fluid.

In the design of high-hydrogen content gas turbines for power generation, flashback of the turbulent flame by propagation through the low velocity boundary layers in the premixing region is an operationally dangerous event. The high reactivity of hydrogen combined with enhanced flammability lim- its (compared to natural gas) promotes flame propagation along low-speed boundary layers adjoining the combustion walls. This work focuses on the simulation of boundary layer flashback using large-eddy simulations (LES). A canonical channel configuration is studied to assess the capabilities of LES and determine the modeling requirements for boundary layer flashback simulations. To extend this work to complex geometries, a new reactive low-Mach number solver has been written in an unstructured code. / text

A 3D parallel CFD code is written to investigate the characteristics of and differences between Large Eddy Simulation (LES) models in the context of simulating a thermal buoyant plume. An efficient multigrid scheme is incorporated to solve the Poisson equation, resulting from the fractional step, projection method used to solve the Low Mach Number (LMN) Navier-Stokes equations. A wide range of LES models are implemented, including a variety of eddy models, structure models, mixed models and dynamic models, for both the momentum stresses and the temperature fluxes. Generalised gradient flux models are adapted from their RANS counterparts, and also tested. A number of characteristics are observed in the LES models relating to the thermal plume simulation in particular and turbulence in general. Effects on transition, dissipation, backscatter, equation balances, intermittency and energy spectra are all considered, as are the impact of the governing equations, the discretisation scheme, and the effect of grid coarsening. Also characteristics to particular models are considered, including the subgrid kinetic energy for the one-equation models, and constant histories for dynamic models. The argument that choice of LES model is unimportant is shown to be incorrect as a general statement, and a recommendation for when the models are best used is given.

Large heat transfer rates are a major source of possible failure in flight vehicles due to increases in temperature being linked to weakening material properties. Aircraft in high-Mach number flow generate excessive aerodynamic heat that may increase temperatures above limits of structural integrity. Even without reducing speed or changing material, it is possible to mitigate heat transfer by altering vehicle geometry. The purpose of this thesis is to study the extent of heat transfer in gap regions of various sizes by computationally simulating flow over an idealized missile-fin configuration. Maximum levels of heat transfer are analyzed as well as surface distributions that identify key design points. The Department of Defense software package with computational fluid dynamics capabilities, Kestrel, was employed to use the Reynolds-averaged Navier-Stokes equations to simulate turbulent Mach~6 flow over the missile model. Results are compared to data obtained by the Air Force Research Laboratory via wind tunnel tests of the same flow. Experiments and simulations both found an order of magnitude increase in heat transfer when an offset fin was attached, but this heating could be reduced by minimizing the offset distance. Simulated baseline properties agreed very well with experimental measurements and simulations of the gap region more precisely identified the locations of maximum heating.

A tailored integration scheme is developed to treat stiff reaction-diffusion prob- lems. The construction adapts a stiff solver, namely VODE, to treat reaction im- plicitly together with explicit treatment of diffusion. The second-order Runge-Kutta- Chebyshev (RKC) scheme is adjusted to integrate diffusion. Spatial operator is de- scretised by second-order finite differences on a uniform grid. The overall solution is advanced over S fractional stiff integrations, where S corresponds to the number of RKC stages. The behavior of the scheme is analyzed by applying it to three simple problems. The results show that it achieves second-order accuracy, thus, preserving the formal accuracy of the original RKC. The presented development sets the stage for future extensions, particularly, to multidimensional reacting flows with detailed chemistry.

In the field of supersonic jet noise, measurements of full-scale afterburning jet engines are infrequent and provide unique opportunities to better understand jet noise phenomena. This thesis represents a phenomenological jet noise analysis using far-field noise data collected from a T-7A-installed GE F404-103 turbofan engine. One issue with the far-field acoustic data from the T-7A was the effects of ground reflections present in the spectra generated from the measured waveforms. A previously developed ground reflection model was implemented into the data to account for this interference. This work represents the first time this model has been implemented in data collected from a full-scale aircraft. Spectra and spatiospectral maps are used to show that, while imperfect, this model represents a step in the right direction for accounting for ground reflections. From the ground reflection-corrected data, sound power values were calculated at varying engine powers. These values were compared against mechanical power values calculated using provided engine parameters at the corresponding engine conditions. It is shown that the observed increase in sound power with mechanical power at supersonic engine conditions follows classical jet noise theory, while the increase between transonic engine conditions is much greater than predicted by classical jet noise theory. This divergence is currently unexplained. Finally, far-field noise directivity measured from the T-7A is connected back to both physics-based and empirically derived definitions of the convective Mach number, a dimensionless parameter used to describe the velocities of coherent structures in the turbulent mixing layer of a jet. For supersonic jets, where Mach wave radiation is the dominant noise source, the convective Mach number should be useful for predicting peak directivity angles. The evaluated definitions show that the convective Mach number associated with Kelvin-Helmholtz instability waves best predicts the peak directivity of the T-7A.

One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semi-global transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime. This research presents several unique contributions to the literature. First, different classes of injection are compared at the same flow conditions to evaluate their suitability for forebody injection. A novel comparison methodology is presented that allows for a technically defensible means of identifying outperforming concepts. Second, potential wall cooling schemes are identified and simulated in a parametric manner in order to identify promising autoignition mitigation methods. Finally, the presence of instabilities in the shock-induced combustion zone of the flowpath are assessed and the analysis of fundamental physics of blunt-body premixed, shock-induced combustion is accelerated through the reformulation of the Navier Stokes equations into a rapid analysis framework. The usefulness of such a framework for conducting parametric studies is demonstrated.

This document details original numerical studies performed by the author pertaining to the propellant oxidizer, ammonium perchlorate (AP). Detailed kinetic mechanisms have been utilized to model the combustion of the monopropellants AP and ADN, and a two-dimensional diffusion flame model has been developed to examine the flame structure above an AP/HTPB composite propellant. This work was part of an ongoing effort to develop theoretically based, a priori combustion models. The improved numerical model for AP combustion utilizes a “universal” gas-phase kinetic mechanism previously applied to combustion models of HMX, RDX, GAP, GAP/RDX, GAP/HMX, NG, BTTN, TMETN, GAP/BTTN, and GAP/RDX/BTTN. The universal kinetic mechanism has been expanded to include chlorine reactions, thus allowing the numerical modeling of AP. This is seen as a further step in developing a gas-phase kinetic mechanism capable of modeling various practical propellants. The new universal kinetic mechanism consists of 106 species and 611 reactions. Numerical results using this new mechanism provide excellent agreement with AP's burning rate, temperature sensitivity, and final species data. An extensive literature review has been conducted to extract experimental data and qualitative theories concerning ADN combustion. Based on the literature review, the first numerical model has also been developed for ADN that links the condensed and gas phases. The ADN model accurately predicts burning rates, temperature and species profiles, and other combustion characteristics of ADN at pressures below 20 atm. Proposed future work and modifications to the present model are suggested to account for ADN's unstable combustion at pressures between 20 and 100 atm. A two-dimensional model has been developed to study diffusion in composite propellant flames utilizing a vorticity formulation of the transport equations. This formulation allows for a more stable, robust, accurate, and faster solution method compared to the Navier-Stokes formulations of the equations. The model uses a detailed gas-phase kinetic mechanism consisting of 37 species and 127 reactions. Numerical studies have been performed to examine particle size, pressure, and formulation effects on the flame structure above an AP/HTPB propellant. The modeled flame structure was found to be qualitatively similar to the BDP model. Results were consistent with experimental observations. Three different combustion zones, based on particle size and pressure, were predicted: the AP monopropellant limit, the diffusion flame, and a premixed limit. Mechanistic insights are given into AP's unique combustion properties.

ITC/USA 2006 Conference Proceedings / The Forty-Second Annual International Telemetering Conference and Technical Exhibition / October 23-26, 2006 / Town and Country Resort & Convention Center, San Diego, California / Telemetry data are usually collected for analysis at some later time and can be monitored to follow the progress of a test. In the case of an Air Data System the signals from the sensors are sent to a computer that calculates the air data parameters for use on multiple LabView-generated displays, as well as to the Data Acquisition System. The readouts on the multiple displays need to be real-time so they are useful to the flight crew. Equations that control the different air data values are determined by what telemetry data are available and the preference of those doing the test planning. These systems need to display the information in a format useful to the flight crew and be reliable.