This work seeks to analyze the transient characteristics of a high-speed inlet with a variable-geometry, rotating cowl. The inlet analyzed is a mixed compression inlet with a compression ramp, sidewalls and a rotating cowl. The analysis is conducted at nominally Mach 4.0 wind tunnel conditions. Advanced Computational Fluid Dynamics techniques such as transient solutions to the Unsteady Reynolds-averaged Navier-Stokes equations and relative mesh motion are used to predict and investigate the unstart and restart processes of the inlet as well as the associated hysteresis. Good agreement in the quasi-steady limit with a traditional analysis approach was obtained. However, the new model allows for more detailed, time-accurate information regarding the fully transient features of the unstart, restart, and hysteresis to be obtained that could not be captured by the traditional, quasi-steady analysis. It is found that the development of separated flow regions at the shock impingement points as well as in the corner regions play a principal role in the unstart process of the inlet. Also, the hysteresis that exists when the inlet progresses from the unstarted to restarted condition is captured by the time-accurate computations. In this case, the hysteresis manifests itself as a requirement of a much smaller cowl angle to restart the inlet than was required to unstart it. This process is shown to be driven primarily by the viscous, separated flow that sets up ahead of the inlet when it is unstarted. In addition, the effect of cowl rotation rate is assessed and is generally found to be small; however, definite trends are observed. Finally, a rigorous assessment of the computational errors and uncertainties of the Variable-Cowl Model indicated that Computation Fluid Dynamics is a valid tool for analyzing the transient response of a high-speed inlet in the presence of unstart, restart and hysteresis phenomena. The current work thus extends the state of knowledge of inlet unstart and restart to include transient computations of contraction ratio unstart/restart in a variable-geometry inlet. / Doctor of Philosophy / Flight at high speeds requires efficient engine operation and performance. As the vehicle traverses through its flight profile, the engine will undergo changes in operating conditions. At high speeds, these changes can lead to significant performance loss and can be detrimental to the vehicle. It is, therefore, important to develop tools for predicting characteristics of the engine and its response to disturbances. Computational Fluid Dynamics is a common method of computing the fluid flow through the engine. However, traditionally, CFD has been applied to predict the static performance of an engine. This work seeks to advance the state of the art by applying CFD to predict the transient response of the engine to changes in operating conditions brought about by a variable geometry inlet with rotating components.

Colorectal cancer is the third most common form of cancer, and is the number two cancer-related death in the United States. Receiving regular colonoscopies can reduce the average person's risk of dying from colon cancer by 90%. However, only 54% if adults over the age of 50 get regular colonoscopies. This low percentage can be attributed to the exam's poor availability, severe discomfort, high cost, and the risk of procedural complications. The Endoscope Propulsion System, or EPS, will assist in the colonoscopy procedure. This device will enable a lesser skilled physician to effectively perform the colonoscopy, thus increasing the procedure's availability. In addition to requiring less skill, the assistive nature of the EPS will also decrease the chance of complications due to colon perforation. The EPS will greatly reduce the discomfort cause by the colonoscope, which will eliminate the need for anesthesia and recovery, therefore greatly reducing the cost of the procedure. The Endoscope Propulsion System design described in this paper is an update to the device outlined in Dr. M. Jonathan Bern's patent application (20060270901). The criteria and requirements of the design are discussed along with the final design and analysis. Finally, a prototype was built to ensure the validity of the proposed invention. / Master of Science

A lithium vaporizer for a high-power magnetoplasmadynamic (MPD) thruster is modeled using a parallel approach. A one-dimensional, thermal-resistive network is developed and used to calculate the required vaporizer length and power as a function of various parameters. After comparing results predicted by this network model with preheat power data for a 200 kW MPD thruster, we investigate performance over a parameter space of interest for the Advanced Lithium-Fed, Applied-field, Lorentz Force Accelerator (ALFA2) thruster. Heater power sensitivity to cathode tube emissivity, mass flow rate, and vapor superheat are presented. The cold-start heater power for 80 mg/s is found to range from 3.38 to 3.60 kW, corresponding to a vaporizer (axial) length of 18 to 26 cm. The strongest drivers of vaporizer performance are cathode tube emissivity and a conduction heat sink through the mounting flange. Also, for the baseline ALFA2 case, it is shown that increasing the vapor superheat from 100 K to 300 K has the effect of lowering the vaporizer thermal efficiency from 57% to 49%. Also, a finite-volume computational fluid dynamic (CFD) is implemented in FLUENT 6.2 which includes conjugated heat transfer to the solid components of the cathode. This model uses a single-fluid mixture model to simulate the effects of the two-phase vaporizer flow with source terms that model the vaporization. This model provides a solution of higher fidelity by including three-dimensional fluid dynamics such as thermal and momentum boundary layers, as well as calculating a higher resolution temperature distribution throughout the cathode assembly. Results from this model are presented for three mass flow rates of interest (40 mg/s, 80 mg/s, and 120 mg/s). Using a fixed power and length taken from the conceptual ALFA2 design, the dryout point ranges from 12.3-17.6 cm from the base of the cathode assembly for 40 mg/s and 80 mg/s, respectively. For the 120 mg/s case, the two-phase flow never reaches dryout. Finally, results two modeling approaches are compare favorably, with a maximum disagreement of 13.0 percent in prediction of the dryout point and 4.2 percent in predictions of the exit temperature.

This study seeks to understand the interplay of vorticity and elasto-dynamics that forms the basis for a fluttering flag and flapping wing propulsion, and factors that distinguish one from the other. The fluid dynamics is assumed two dimensional and incompressible, and comprises potential and viscous flow simulations. The elastic solid is one dimensional and governed by the Bernoulli-Euler flexure model. The fluid and elastic solid models are coupled using a predictor-corrector algorithm. Flutter of a flag or foil is associated with drag and we show that the pressure on the foil is predominantly circulatory in origin. The circulatory pressure generated on the foil depends primarily on the slope and curvature. The wake vorticity exhibits a wide range of behavior starting from a Kelvin-Helmholtz type instability to a von Kármán wake. Potential flow simulations do not capture the wake accurately both at high and low mass ratios. This is reflected in the flutter boundary and pressure over the foil when compared with viscous flow simulations. Thrust due to heaving of a flexible foil shows maxima at a set of discrete frequencies that coincide with the frequencies at which the flapping velocity of the foil tip is a maximum. The propulsive efficiency shows maxima at a set of discrete frequencies that are close but distinct from the thrust maxima set of frequencies. These discrete frequencies are close to the natural frequencies of vibration of a cantilevered foil vibrating in vacuum. At low frequencies thrust is a consequence of a strong leading edge vortex developed over the foil and it remains attached to the foil as it is convected due to the favorable pressure gradient presented by the time and spatially varying shape of the foil. At moderate and high frequencies of oscillation the pressure, and consequently the thrust, generated by the foil is non-circulatory in origin and they are high where the accelerations of the foil are high. At high frequencies the leading edge vortex is weak. Except in the low frequency range, potential flow simulations qualitatively compares well with viscous flow predictions. We show that thrust and drag on a flexible foil oscillating in a flow is caused by the phase difference between the slope of the foil and the fluid pressure on it. Propulsive efficiency though is governed by the phase difference between foil velocity and fluid pressure and inertia forces. Thus, the interplay of vorticity and elasto-dynamics determine the behavior of a flutter and propulsion of a flexible foil in a fluid flow.

This paper is an investigation into a quantitative selection process of either a mechanical or electrical system architecture for the transmission of propulsion power in naval combatant vessels. A database of historical naval ship characteristics was statistically analyzed to determine if there were any predominant ship parameters that could be used to predict whether a ship should be designed with a mechanical power transmission system or an electric one. A Principal Component Analysis was performed to determine the minimum number of dimensions required to define the relationship between the propulsion transmission architecture and the independent variables. Combining the results of the statistical analysis and the PCA, neural networks were trained and tested to separately predict the transmission architecture or the installed electrical generation capacity of a given class of naval combatant.

As conventional chemical rockets reach the outer limits of their abilities, significant research is going into alternative thruster technologies, some of which decouple the maximum thrust and efficiency from the propellant's internal chemical energy by supplying energy to the propellant as needed. Of particular interest and potential is the electrically powered thruster, which promises very high specific thrust using relatively inexpensive and stable propellant gasses. Some such thrusters, specifically ion thrusters, have achieved significant popularity for various applications. However, there exist other classes of electrical thrusters which promise even higher levels of efficiency and performance. This thesis will focus on one such thruster type - the magnetoplasmadynamic thruster - which uses an ionized propellant flow and large currents to accelerate the propellant gas by electrical and magnetic force interactions. The necessary background will be presented in order to understand and characterize the operation of such devices, and a theoretical model will be developed in order to estimate the levels of performance which can be expected. Simulations will be performed and analyzed in order to better understand the principles on which these devices are designed. Finally, a thruster package will be designed and built in order to test the performance of the device and accuracy of the model. This will include a high-current power supply, ignition circuit, gas delivery system, and nozzle. Finally, the measured performance of this thruster package will be measured and compared to the theoretical predictions in order to validate the models constructed for this type of thruster.

The rectangular pulsed plasma thruster (PPT) is an electromagnetic thruster that ablates Teflon propellant to produce thrust in a discharge that lasts 5-20 microseconds. In order to integrate PPTs onto spacecraft, it is necessary to investigate possible thruster plume-spacecraft interactions. The PPT plume consists of neutral and charged particles from the ablation of the Teflon fuel bar as well as electrode materials. In this thesis a novel application of quadruple Langmuir probes is implemented in the PPT plume to obtain electron temperature, electron density, and ion speed ratio measurements (ion speed divided by most probable thermal speed). The pulsed plasma thruster used is a NASA Glenn laboratory model based on the LES 8/9 series of PPTs, and is similar in design to the Earth Observing-1 satellite PPT. At the 20 J discharge energy level, the thruster ablates 26.6 mg of Teflon, creating an impulse bit of 256 mN-s with a specific impulse of 986 s. The quadruple probes were operated in the so-called current mode, eliminating the need to make voltage measurements. The current collection to the parallel to the flow electrodes is based on Laframboise's theory for probe to Debye length ratios between 5 and 100, and on the thin-sheath theory for ratios above 100. The ion current to the perpendicular probe is based on a model by Kanal and is a function of the ion speed ratio, the applied non-dimensional potential and the collection area. A formal error analysis is performed using the complete set of nonlinear current collection equations. The quadruple Langmuir probes were mounted on a computer controlled motion system that allowed movement in the radial direction, and the thruster was mounted on a motion system that allowed angular variation. Measurements were taken at 10, 15 and 20 cm form the Teflon fuel bar face, at angles up to 40 degrees off of the centerline axis at discharge energy levels of 5, 20, and 40 J. All data points are based on an average of four PPT pulses. Data analysis shows the temporal and spatial variation in the plume. Electron temperatures show two peaks during the length of the pulse, a trend most evident during the 20 J and 40 J discharge energies at 10 cm from the surface of the Teflon fuel bar. The electron temperatures after the initial high temperature peak are below 2 eV. Electron densities are highest near the thruster exit plane. At 10 cm from the Teflon surface, maximum electron densities are 1.04e20 ± 2.8e19 m-3, 9.8e20 ± 2.3e20 m-3, and 1.38e21 ± 4.05e20 m-3 for the 5 J, 20 J and 40 J discharge energy, respectively. The electrons densities decrease to 2.8x1019 ± 8.9e18 m-3, 1.2e20 ± 4.2e19 m-3, and 4.5e20 ± 1.2e20 m-3 at 20 cm for the 5 J, 20 J, and 40 J cases, respectively. Electron temperature and density decrease with increasing angle away from the centerline, and with increasing downstream distance. The plume is more symmetric in the parallel plane than in the perpendicular plane. Ion speed ratios are lowest near the thruster exit, increase with increasing downstream distance, but do not show any consistent angular variation. Peak speed ratios at a radial distance of 10 cm are 5.9±3.6, 5.3±0.39, and 4.8±0.41 for the 5 J, 20 J and 40 J discharge energies, respectively. The ratios increase to 6.05±5.9, 7.5±1.6, and 6.09±0.72 at a radial distance of 20 cm. Estimates of ion velocities show peak values between 36 km/s to 40 km/s, 26 km/s to 30 km/s, and 26 km/s to 36 km/s for the % J, 20 J, and 40 J discharge energies, respectively.

Beyond Atlas is a Swedish private company with the goal of exploring the solar system with cheap and reliable spacecraft. Part of their maiden mission, Hugin, aims to demonstrate navigation, propulsion, and communication technology on a 3U CubeSat. This thesis aims to investigate the feasibility of using the Enpulsion NANO electric propulsion (EP) system for deep-space applications and how to best demonstrate its capabilities in low-Earth orbit. Literature reviews of scientific papers and software simulations were conducted to gain an understanding of the underlying processes involved in EP in-orbit operations. Analyses were made on orbital maneuvers, momentum unloading, power and thermal restrictions. The results suggest that the EP system's capabilities is mainly limited by the saturation time of the reaction wheels restricting longer duration orbital maneuvers. Orbital maneuvers for demonstrating the capabilities are proposed based on the limitations imposed on the EP system by the rest of the spacecraft. On the basis of the results of this research, it can be concluded that the Enpulsion NANO thruster's operational range can be utilized both as a low thrust efficient main drive and as a high thrust maneuvering thruster for deep-space applications but is limited by the high power consumption and low thrust-to-power ratio.

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