The motivation for this research is the need to improve performance of the autonomous flight of a tailsitter UAV. Tailsitter aircraft combine the hovering and vertical take-off and landing capability of a rotorcraft with the long endurance flight capability of a fixed-wing aircraft. The particular aircraft used in this research is the V-Bat, a tailsitter UAV with a conventional wing and the propeller and control surfaces located within a ducted-fan tail assembly. This research focuses on identifying the models and parameters of the V-Bat in hover and level flight as a basis for the design of the control systems for hover, level, and transition modes of flight.Models and parameters were identified from experimental data. Wind-tunnel tests, bench tests, and flight tests were performed in a variety of flight conditions. Wind tunnel tests yielded force and moment coefficients over the full flight envelope of the V-Bat. Models and parameters for longitudinal, lateral, and hover flight are presented. Bench tests were conducted to enhance understanding about the ducted-fan propulsion system and the effectiveness of the control surfaces. The thrust characteristics of the ducted fan were measured. Control derivatives were derived from force and moment measurements. Flight tests were completed to obtain dynamic models of the V-Bat in hover flight. Using frequency-domain system identification methods, frequency-response and transfer function models of roll, pitch, and yaw responses to aileron, elevator, and rudder control input were derived.The results obtained from these experimental tests were used to identify models and parameters of the V-Bat aircraft, giving insight into its behavior and enhancing the control analysis and simulation capabilities for this aircraft, thus providing the increased levels of understanding needed for autonomous flight.

Cavitation is detrimental to the performance of ships and submarines, causing noise, erosion, and vibration. This study seeks to understand cavitation inception and delay on a typical ducted propulsor by utilizing the SimCenter's unstructured simulation and design system: U2NCLE. Specifically, three fundamental questions are addressed: 1. What are the macroscale flow physics causing cavitation inception? 2. How does cavitation inception scale with Reynolds number? 3. How can tip-leakage vortex cavitation inception be suppressed? To study the physics of cavitation inception, a ducted propulso simulation is developed and extensively validated with experimental results. The numerical method is shown to agree very well with experimental measurements made in the vortex core. It was discovered that the interaction of the leakage and trailing edge vortices cause the pressure to drop to a local minimum, providing ideal conditions for inception to occur. However, experimental observation shows that inception does not occur at the minimum pressure location, but rather at the point where the two vortices completely coalesce. At the point of coalescence, the simulation reveals that the streamwise core velocity decelerates, causing the air nuclei to stretch and burst. A Reynolds number scaling analysis is performed for the minimum pressure and maximum velocity in the vortex core. First, the numerical method is validated on a flate plate at various Reynolds numbers to assess the ability of typical turbulence models to predict Reynolds numbers ranging from one million to one billion. This scaling analysis methodology is then applied to the propulsor simulation, revealing that the minimum pressure in the vortex core is much less dependent on Reynolds number than was previously hypothesized. Lastly, to investigate means of delaying cavitation inception, the propulsor is parameterized and studied using design optimization theory. Concepts of vortex alleviation evident in nature are used to suggest suitable parameterizations. Also, dimension reduction is used to reduced the number of design variables. Finally, the concepts are implemented, evaluated, and shown to completely decouple the two vortices causing cavitation inception. Moreover, the minimum pressure in the vortex core is significantly increased.

The present work was conducted with the intent of developing a high-fidelity numerical model of a unique combustion flow problem combining multi-phase fuel injection with substantial momentum and temperature into a highly complex turbulent flow. This important problem is very different from typical and more widely known liquid fuel combustion problems and is found in practice in pulverized coal combustors and ducted rocket ramjets. As the ducted rocket engine cycle is only now finding widespread use, it has received little research attention and was selected as a representative problem for this research. Prior to this work, a method was lacking domestically and internationally to effectively model the ducted rocket engine cycle with confidence. In the ducted rocket a solid fuel gas generator is used to deliver a fuel-rich multi-phase mixture to the combustion chamber. When a valve is used to vary the fuel generator pressure, and thereby the delivered fuel flowrate, the engine is known as a Variable Flow Ducted Rocket (VFDR). The Aerojet MARC-R282 ramjet engine represents the worlds first VFDR flown, and the first in operational use. Although performance requirements were met, improvements are sought in the understanding of the ramjet combustion process with a future aim of reducing the visible exhaust and correcting uneven combustor heating patterns. For this reason the MARC-R282 combustor was selected as the baseline geometry for the present research, serving to provide a documented baseline case for numerical modeling and also being a good candidate to benefit from an improved understanding of the combustion process. In order to proceed with the present research, experiments were first carried out to characterize the gas generator particulate exhaust in terms of composition and particle size. Equilibrium thermochemistry was used to supplement these data to develop a gas phase combustion model. The gas phase reactions and resulting particle definition were modeling using the FLUENT Computational Fluid Dynamics (CFD) code for the baseline GQM-163A Supersonic Sea Skimming Missile (SSST) operating conditions. These results were compared to direct-connect ramjet ground tests in order to validate the analysis tool. Data were developed to understand the gas and solid phase fuel exhaust characteristics at the propellant surface, exiting the gas generator injector, and following secondary combustion with air. Particles were collected and analyzed from the fuel generator exhaust. While exhibiting some variation with time in the firing, they were roughly an average of 20 microns in diameter, in line with prior experience with pulverized coal combustion experiments. A computational model was developed based on coal combustion parameters using FLUENT. However, despite considerable effort, the CFD analysis was not able to predict effective burning of the carbon particles to the degree seen in testing. In addition, using equilibrium thermochemistry as a basis for determining the carbon particle content in the fuel exhaust, the CFD analysis resulted in trends in performance opposite to the test results. These facts led to a hypothesis that there was actually a significant fraction of small particles or much less carbon produced than equilibrium thermochemistry would predict. A parametric analysis was performed replacing the 20 micron soot particles with fine fraction particles, representing a fraction of the predicted equilibrium carbon soot being still in the gas phase as higher molecular weight hydrocarbons, or in the form of sub-micron particles. When almost all particles were replaced with fine fraction particles, the model was able to correctly predict absolute values of combustion efficiency as well as trends for different injector geometries. The presence of particles was apparent from the visible exhaust and collection data, however they were found not to play a significant role in the combustion process for this fuel and engine configuration. The robustness of the computational model was also evaluated by examining the effects of turbulence model, order of discretization, and grid size. Comparable trends and results were seen for all cases examined. With the successful development of this modeling tool and an improved understanding of the combustion process, future work is enabled to develop improved combustor flow management and fuel injection schemes to improve existing designs and develop new configurations. This research has served to advance the field of combustion modeling by providing: 1) a solid ducted rocket combustion modeling tool considering solid and gas phase combustion, 2) a correlation between primary combustion theory and particulate exhaust sampling, 3) low length/diameter ratio ducted rocket combustor modeling, and 4) combustor CFD coupled with solid particle tracking and combustion models. / Ph. D.

Unmanned aerial vehicles (UAVs) are increasingly being used for both military and commercial applications to replace more costly and dangerous manned operations. Vehicles with vertical take-off and landing (VTOL) and hovering capabilities are of interest for functions such as surveillance and inspection where the ability to hold the position of the vehicle is desired. Ducted fan vehicles are of particular interest because of their high efficiency per unit diameter when compared to the more commonly seen multirotor vehicles. This makes ducted fan UAVs very well suited for size-constrained missions such as indoor inspection or urban reconnaissance. However, the advantages of ducted fans come at the cost of complex nonlinear dynamics which present challenging modeling and control problems. This thesis provides a detailed discussion of the instrumentation, modeling, and control of a ducted fan UAV. The dynamic model of the UAV is computed from a simplified parametric model. Unknown parameters of the model are found from system identification based on flight data. Synthesis of a linear state feedback controller based on this model is discussed, and it is demonstrated in hardware that this controller can effectively stabilize the vehicle. / Master of Science

Aviation industry has a crucial impact on society on the grounds that it offers wider social and economic benefits. The demand of transportation is increasing and it is expected that the worldwide fleet of aircraft and rotorcraft will increase accordingly. This growth will introduce an increased environmental impact which can be controlled with the introduction and the implementation of new and greener technologies which can provide both a reduced carbon foot-print and increased efficiency. Therefore, the simulation of new designs with tools that can capture the flow physics accurately is crucial, on the grounds that an accurate simulation could provide novel designs and new ways in order to design from scratch new vehicles as well as providing a better appreciation of the physics that are involved. This work has a central aim to propose a methodology which combines CFD simulations and the method of performance mapping. It focuses on the application of a ducted tail rotor which can offer significant performance benefits compared to a conventional tail rotor. The developed methodology was tested against the results of an in-house rotorcraft comprehensive code and provided a reasonable qualitative correlation. In principle, this methodology can work for all helicopter flight phases such as hover, climb, cruise, descend but due to the complexity of the investigations, together with the lack of experimental data that can be used to refine the CFD model, only the hover and forward flight were considered. Although CFD studies of a ducted tail rotor currently exist in the literature (though scarce), this work can be considered, to the best knowledge of the author as a first attempt in investigating the performance of the configuration, from low to high forward flight speed, by combining CFD and performance mapping.

abstract: For a conventional quadcopter system with 4 planar rotors, flight times vary between 10 to 20 minutes depending on the weight of the quadcopter and the size of the battery used. In order to increase the flight time, either the weight of the quadcopter should be reduced or the battery size should be increased. Another way is to increase the efficiency of the propellers. Previous research shows that ducting a propeller can cause an increase of up to 94 % in the thrust produced by the rotor-duct system. This research focused on developing and testing a quadcopter having a centrally ducted rotor which produces 60 % of the total system thrust and 3 other peripheral rotors. This quadcopter will provide longer flight times while having the same maneuvering flexibility in planar movements. / Dissertation/Thesis / Experimental flight test for ductless quadcopter configuration / Masters Thesis Mechanical Engineering 2019

This thesis extends the use of neural-network-based model reference adaptive control to systems that occur as cascades. In general, these systems are not feedback linearizable. The approach taken is that of approximate feedback linearization of upper subsystems whilst treating the lower-subsystem states as virtual actuators. Similarly, lower-subsystems are also feedback linearized. Typically, approximate inverses are used for linearization purposes. Model error arising from the use of an approximate inverse is minimized using a neural-network as an adaptive element. Incorrect adaptation due to (virtual) actuator saturation and dynamics is avoided using the Pseudocontrol Hedging method. Using linear approximate inverses and linear reference models generally result in large desired pseudocontrol for large external commands. Even if the provided external command is feasible (null-controllable), there is no guarantee that the reference model trajectory is feasible. In order to mitigate this, nonlinear reference models based on nested-saturation methods are used to constrain the evolution of the reference model and thus the plant states. The method presented in this thesis lends itself to the inner-outer loop control of air vehicles, where the inner-loop controls attitude dynamics and the outer-loop controls the translational dynamics of the vehicle. The outer-loop treats the closed loop attitude dynamics as an actuator. Adaptation to uncertainty in the attitude, as well as the translational dynamics, is introduced, thus minimizing the effects of model error in all six degrees of freedom and leading to more accurate position tracking. A pole-placement approach is used to choose compensator gains for the tracking error dynamics. This alleviates timescale separation requirements, allowing the outer loop bandwidth to be closer to that of the inner loop, thus increasing position tracking performance. A poor model of the attitude dynamics and a basic kinematics model is shown to be sufficient for accurate position tracking. In particular, the inner-outer loop method was used to control an unmanned helicopter and has subsequently been applied to a ducted-fan, a fixed-wing aircraft that transitions in and out of hover, and a full-scale rotorcraft. Experimental flight test results are also provided for a subset of these vehicles.

Measuring the farfield noise levels of full-scale rotor systems is not trivial and can be costly. Researchers prefer to perform small-scale experiments in the laboratory so that they can extrapolate the model scaled results to the larger scale. Typically Inverse Square Law (ISL) is used to extrapolate the sound pressure levels (SPL), obtained from model-scale experiments at relatively small distances to predict noise at much larger distances for larger scale systems. The assumption underlying this extrapolation is that the source itself can be treated as a point sound source. At what distance from a rotor system it can be treated as a point source has never been established. Likewise, many theoretical models of shielding by hard surfaces assume the source to be a point monopole source. If one is interested in shielding the noise of a rotor system by interposing a hard surface between the rotor and the observer, can the rotor system really be considered to be a monopole? If rotating noise sources are under consideration what is the effect of configuration and design parameters? Exploring the validity of point source assumption alluded to above for a rotor for farfield acoustic measurements with and without shielding form the backbone of the present work.

Aviation industry has a crucial impact on society on the grounds that it offers wider social and economic benefits. The demand of transportation is increasing and it is expected that the worldwide fleet of aircraft and rotorcraft will increase accordingly. This growth will introduce an increased environmental impact which can be controlled with the introduction and the implementation of new and greener technologies which can provide both a reduced carbon foot-print and increased efficiency. Therefore, the simulation of new designs with tools that can capture the flow physics accurately is crucial, on the grounds that an accurate simulation could provide novel designs and new ways in order to design from scratch new vehicles as well as providing a better appreciation of the physics that are involved. This work has a central aim to propose a methodology which combines CFD simulations and the method of performance mapping. It focuses on the application of a ducted tail rotor which can offer significant performance benefits compared to a conventional tail rotor. The developed methodology was tested against the results of an in-house rotorcraft comprehensive code and provided a reasonable qualitative correlation. In principle, this methodology can work for all helicopter flight phases such as hover, climb, cruise, descend but due to the complexity of the investigations, together with the lack of experimental data that can be used to refine the CFD model, only the hover and forward flight were considered. Although CFD studies of a ducted tail rotor currently exist in the literature (though scarce), this work can be considered, to the best knowledge of the author as a first attempt in investigating the performance of the configuration, from low to high forward flight speed, by combining CFD and performance mapping.

The technology of UAV (Unmanned Aerial Vehicle) has developed since its conception many years ago. UAVs have several features such as, computerised and autonomous control without the need for an on-board pilot. Therefore, there is no risk of loss of life and they are easier to maintain than manned aircraft. In addition, UAVs have an extended range/endurance capability, sometimes for several days. This makes UAVs attractive for missions that are typically