This dissertation is the first comprehensive attempt to address a new engineering problem: control of a vehicle maneuvering in a full unsteady flow field. The approach to the solution is focused in three main areas: modeling of a vehicle in full unsteady flow, control of a vehicle in full unsteady flow, and synthesizing the fluid loads for use in control of a vehicle maneuvering in a full unsteady flow field. To model a vehicle maneuvering in a full unsteady flow field this dissertation develops the Coupled Fluid Vehicle (CFV) model in which the fluid, which is a sum of a finite number of spatially dependent velocity fields whose contributions vary with time, is coupled to the vehicle rigid-body equations of motion. To control a vehicle maneuvering in a full unsteady flow field this dissertation develops the Fluid Compensation Control (FCC) strategy which gives the designer an opportunity to include the fluid states, in addition to the vehicle states, in the control law and an opportunity to balance reducing the fluid dynamic load through compensation and reducing the state error through regulation. To synthesize the fluid loads this dissertation has attempted to forward current work on the prediction of fluid loads from stagnation and separation point measurements using the Kutta principle, which says that the velocity around a vehicle is a smoothly varying function and that it is determined up to a multiplicative constant by its nodes (stagnation, separation, and reattachment points/lines), and by conducting an experiment to attempt to determine the correlation of the fluidic loads from the orientation and separation lines on a 3-dimensional bluff body.

One of the pillars of aviation safety is assuring sound engineering practices through airworthiness certification. As Unmanned Aircraft Systems (UAS) grow in popularity, the need for airworthiness standards and verification methods tailored for UAS becomes critical. While airworthiness practices for large UAS may be similar to manned aircraft, it is clear that small UAS require a paradigm shift from the airworthiness practices of manned aircraft. Although small in comparison to manned aircraft these aircraft are not merely remote controlled toys. Small UAS may be complex aircraft flying in the National Airspace System (NAS) over populated areas for extended durations and beyond line of sight of the operators. A comprehensive systems engineering framework for certifying small UAS at the system level is needed. This work presents a point based tool that evaluates small UAS by rewarding good engineering practices in design, analysis, and testing. The airworthiness requirements scale with vehicle size and operational area, while allowing flexibility for new technologies and unique configurations.

Two approaches for the autonomous control of aerospace vehicle formations are developed. The development of the approaches relies on fundamental work in the areas of distributed control; specifically modal, robust, optimal, and decentralized control. The algorithms are shown to satisfy five separation principles that simplify design and enable the algorithms to be implemented reliably. The autonomous controllers uniformly dampen the modes of the formation (global control) using a decentralized approach and a nearest-neighbor approach. A numerical example illustrates robust formation changes from 9-vehicle (3 x 3) grids to V-type formations.

A distributed actuation array was installed on a blended-wing-body uninhabited aerial vehicle and tested in the 12-foot subsonic wind tunnel at NASA?s Langley Research Center. From the results of these tests, a discussion is given of the baseline aircraft, its conventional control surfaces, and the distributed array. Each effector in the distributed array was tested individually as well as pre-determined configurations incorporating all 12 effectors on each wing. From the tests on the individual effectors, a method was created that allows for the prediction of the control authorities for any configuration of the array. The six pre-determined shapes served as bases for comparison to determine the accuracy of the prediction scheme. Additionally, the shapes were compared to the conventional control surfaces to determine if a distributed array could completely replace those control surfaces.

This study investigates the application of Fiber Bragg Grating (FBG) optical sensors interrogation techniques over several time scales to monitor damage in composite structures due to low velocity impacts events. Optical fiber sensors are embedded into carbon fiber/epoxy resin woven composites using a single-step cure process. The composite specimens are subjected to multiple low energy impacts until failure. Impact events are characterized by acceleration and position sensors integral to the impactor head. The embedded FBG sensors are interrogated using three different interrogation techniques. Low speed, full spectrum measurements are recorded using a tunable laser source. High speed, peak wavelength detection data is taken using a commercial peak wavelength interrogation system. Finally, high speed full spectrum measurements are recorded using new instrumentation developed at Brigham Young University. By qualitatively examining the responses of these three techniques and comparing the FBG data with impact characterization data, a more complete picture of the composite health is available.

The use of forward-swept wings has aerodynamic benefits at high angles of attack and in supersonic regimes. These consist of reduction in wave drag, profile drag, and increased high angle of attack handling qualities. These increased benefits are often offset due to an increase in structural components, to overcome flutter and wing tip divergence due to high loading of the wing tips at high angles of attack. The use of composite materials and aeroelastic tailoring of the structures eliminates these instabilities without a significant increase in weight. This work presents the design of an aeroelastic wing structure for a highly forward-swept wing, and the verification of the aerodynamic and structural finite element analysis through experimental testing.

This research focuses on evaluating empirical methods and implementing a prototype Transitional Airblast Model (TRAM) for facilitating communication between one-dimensional and two-dimensional airblast models. An overview of detonation phenomena is presented, especially concerning detonation waves and accompanying airblast effects. Two existing airblast models are discussed that were designed to predict the effects of a detonation in two separate types of geometries, one-dimensional and two-dimensional. The functionality and behavior of each airblast model will be scrutinized giving particular insight into their performance in applications with both one-dimensional and two-dimensional components. The strengths and deficiencies of the different airblast models will offer motivation for the development of the TRAM prototype. The TRAM prototype consists of two separate methodologies, one for translating one-dimensional airblast propagation to two dimensions, and another for translating two-dimensional airblast propagation to one dimension. The selection of those two methodologies will be presented, along with results of detonation scenarios using both existing airblast models as well as the TRAM prototype. The TRAM prototype performed well for both types of detonation scenarios and is recommended for further development.

The purpose of this research is to model and understand the deployment dynamics of space systems with long and short tethers. This research is divided into two parts; in the first part, a model for short and medium length tether systems is developed and simulated by solving equations of motion. A detailed parametric study is conducted after identifying important parameters affecting the deployment and studying the effect of each parameter on the deployment performance. Certain tools are developed to assist mission planners in predicting the deployment performance of a space tether system for a given set of parameters. The second part of the research is motivated by Space Elevator (SE) dynamics. SE is a futuristic and highly challenging technology based on the idea of connecting Earth and Space by an approximately 100,000 km long tether. The tether used for the SE would be deployed from Geostationary Orbit(GEO). With this motivation, the short tether analysis from the previous section is extended to the analysis of long tethers. A model for long tether deployment is developed and governing equations of motions are formulated. Critical parameters are identified and problems involved in SE deployment are investigated. Tether mass is initially included in the model, but it is found that that the mass of the tether has very little effect on the overall qualitative dynamics of the system. Hence, for further analysis, a massless tether model is adapted. Upon simulating the system, it is found that long tethers can be highly unstable during deployment and can crash onto the Earth. However, a considerable fraction of the tether can be deployed successfully without any external control mechanism before the instability manifests itself. Hence, alternate SE designs with shorter tether deployment requirements may be a possibility.

The formation of NO and NO2 are sensitive indicators of both temperature and residence time. In this work, the NOx emission index is measured in an unsteady counterflow diffusion flame for methane, propane and ethylene, as a function of average strain rate and amplitude and frequency of imposed sinusoidal oscillation. The flames studied vary from non-sooting to high soot loading, and from low average strain rate to near extinction. Due to the relatively long time scales associated with NOx formation, the effect of unsteadiness on emission index is weaker than on either temperature or soot volume fraction. Time average global measurements were taken using a California Analytical Instruments Model 400 HCLD NO/NOx analyzer. Results are compared with unsteady calculations using a modified OPPDIF code included in the Chemkin package.

Serpentine inlet ducts are utilized in many aircraft where the inlet capture area is located off the thrust line or there is a desire to conceal the engine compressor face. Due to the curvature that characterizes a compact serpentine duct, issues with flow distortion and total pressure loss at the engine face arise leading to reduction in propulsion system performance. Computational analysis has shown that flow control implementing micro-fluidic vortex generators significantly reduces the losses. Previous work at North Carolina State University has demonstrated the benefits of a fluidic flow control of this type in a highly compact serpentine inlet duct through the design and experimental static testing of a propulsion system for an uninhabited aerial vehicle. With the implementation of flow control, engine face distortion was reduced and propulsion system performance was increased. This work continues the investigation of the effectiveness of the fluidic flow control by examining the performance of the system during dynamic situations through high speed taxi testing of an uninhabited aerial vehicle implemented with this technology. Additionally, the collected data was used to compare calculated takeoff parameters to values calculated using standard takeoff analysis.