VULCAN (Viscous Upwind aLgorithm for Complex flow ANalysis) is a cell centered, finite volume code used to solve high speed flows related to hypersonic vehicles. Two algorithms are presented for expanding the range of applications of the current Navier-Stokes solver implemented in VULCAN. The first addition is a highly implicit approach that uses subiterations to enhance block to block connectivity between adjacent subdomains. The addition of this scheme allows more efficient solution of viscous flows on highly-stretched meshes. The second algorithm addresses the shortcomings associated with density-based schemes by the addition of a time-derivative preconditioning strategy. High speed, compressible flows are typically solved with density based schemes, which show a high level of degradation in accuracy and convergence at low Mach numbers (M < 0.1). With the addition of preconditioning and associated modifications to the numerical discretization scheme, the eigenvalues will scale with the local velocity, and the above problems will be eliminated. With these additions, VULCAN now has improved convergence behavior for multi-block, highly-stretched meshes and also can accurately solve the Navier-Stokes equations for very low Mach numbers.

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.

With increasing interest in the use of adaptive lifting surfaces for improved aircraft performance, it is necessary to study the stability and control characteristics of an aircraft with an adaptive wing. This research builds on recent development of an automated cruise flap for adapting a wing shape to achieve low drag over a large lift range. Such an automated cruise flap system was shown to have unusual lift and pitching moment curves, which prompted the need for studying the effect on aircraft stability and control. In this thesis, the static stability considerations are used to show that when an automated cruise flap is used on an airfoil to continuously adjust the flap to the optimum angle, there is a need for an accompanying controller that achieves the desired lift coefficient by adjusting the airfoil angle of attack. Likewise, when an automated flap is used on an aircraft wing, there is a need for an accompanying airspeed controller to adjust the elevator. The thesis presents Simulink models for analyzing the dynamic behavior of an aircraft with an automated flap. Two schemes were studied for the flap and elevator controllers. The most desirable results were obtained when the flap and elevator controllers were coupled so that trim changes due to the flap are immediately compensated by adjustment of the elevator angle. The results of the simulation show that the aircraft does not exhibit any undesirable behavior with the automated cruise flap. The study, therefore, provides the confidence needed to implement the system on an uninhabited aerial vehicle.

The objective of the thesis is to analyze the dynamics and controls of a solar sail operating on a Solar Storm Warning Mission. The Solar Storm Warning Mission was chosen for analysis because it has a high priority mission to provide data on solar storm activity. Solar storms can cause blackouts in power grids and disable satellites that are currently orbiting the Earth. A payload extends from the sailcraft by four tethers attached near the ends of the booms to provide stability. The tether length that produced a stable sailcraft was equal to or greater than 84.31 meters. A parametric study was completed to examine the effects of the roll, pitch, and yaw disturbances on the sailcraft dynamics. The roll and pitch coupling (in the equations of motion) had a large effect on the sailcraft using short tether lengths, but had a much smaller effect at the longer tether lengths of 84.31 meters or more. Thrusters were placed at the ends of the sail booms to provide dampening to the sailcraft. Full-state feedback (pole placement) provided settling times and number of maneuvers for each specific thruster/tether length combination. If the disturbances (roll, pitch, and yaw) expected to impact the sailcraft are known, an ideal tether length can be found as a function of the settling time and number of maneuvers. For example, the 125 mN thruster requires a tether length of approximately 250 meters to provide the most maneuvers for any disturbance of less than twenty degrees.

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.

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 benefits of highly compact serpentine inlet ducts extend from reductions in overall aircraft weight to higher survivability, as well as allow the aircraft designer greater flexibility in propulsion system integration. Unfortunately, due to the extreme wall curvature, these ducts result in significant flow distortion and total pressure losses at the engine face. It has been shown that active flow control in the form of micro-fluidic vortex generators significantly helps to reduce these losses. To date, these systems have only been tested in a laboratory setting in which items such as flow control air supply, system and subsystem size, weight, and location are not major factors. Subscale unmanned aerial vehicles provide a real world test bed to help overcome these constraints at a lower cost and lower risk as compared to full scale aircraft testing. This work presents the design, integration, testing, and analysis of an unmanned aerial vehicle?s propulsion system that implements fluidic flow control inside a highly compact serpentine inlet duct in order to reduce engine face distortion and increase propulsion system performance.

This work presents the background, flight testing, and resulting change in the aerodynamic and handling qualities of a C-130A Hercules modified with wing tip tanks. The data collected during the baseline and modified flight tests of this aircraft demonstrated the potential aerodynamic benefits of a tip tank design that incorporates a greater aspect ratio and end plating effects. Wing mounted pressure belts measured a 24% increase in local Cl near the tip tanks. This local increase in lift contributed to a 38% increase in CL max for the airplane. The pressure and dynamic data was gathered using a LIFT (Linux In Flight Testing) system, and it laid the foundation for finding the longitudinal and lateral directional stability coefficients of the airplane. Then using MATLAB® and the System IDentification Programs for AirCraft (SIDPAC) to reduce this data, it was possible to generate aerodynamic, lateral, and longitudinal parameters that clearly proved the overall benefits of the design change. The demonstrated lift benefits of these uniquely designed tip tanks for the C-130A cargo transport proved that by capitalizing on the benefits of a combination tip tank and end plate design it is possible to generate increased lift without adversely affecting the stability and dynamic parameters of the aircraft.

Peterson, Joseph Scott, Design Study of an Autonomous Unmanned Air Vehicle Controller. (Under the direction of Dr. Charles Hall) The purpose of this document is to both quantitatively and qualitatively compare three varying approaches to control of unmanned air vehicles: dynamic inversion, classical gain scheduling, and robust gain scheduling through H? synthesis. The quantitative comparisons include robust performance and robust stability measures of the aircraft and controller linearized about a trim operating point. A second study will look at the time response of the system with varying perturbation to the nominal plant dynamics. The qualitative analysis looks at the complexity of the controller and time required to implement, this included comparison of iterative solving methods for system gains both in the frequency and time domains. A full nonlinear form of the dynamic inversion controller was implemented with full non-linear simulation in Simulink and shows comparable results to the linearized studies. The results have shown that a dual loop controller with a dynamic inversion inner loop and H? outer loop has the largest robust performance over the flight envelope but one of the more advanced forms of control for implementation.

A reconstruction-based discontinuous Galerkin (RDG) method is presented for the solution of the compressible Euler equations on arbitrary grids. By taking advantage of handily available and yet invaluable information, namely the derivatives, in the context of the discontinuous Galerkin methods, a polynomial solution of one degree higher is reconstructed using a least-squares method. The stencils used in the reconstruction involve only the von Neumann neighborhood (face-neighboring cells) and are compact and consistent with the underlying DG method. The resulting RDG method can be regarded as an improvement of a recovery-based DG method, in the sense that it shares the same nice features, such as high accuracy and efficiency, and yet overcomes some of its shortcomings such as a lack of flexibility, compactness, and robustness. The developed RDG method is used to compute a variety of flow problems on arbitrary meshes to demonstrate its accuracy, efficiency, robustness, and versatility. The numerical results indicate that this RDG method is third-order accurate at a cost slightly higher than its underlying second-order DG method, at the same time providing a better performance than the third order DG method, in terms of both computing costs and storage requirements.