Powering spacecraft with electric propulsion is becoming more common, especially in CubeSat-class satellites. On account of the risk of spacecraft interactions, it is important to have robust analysis and modeling tools of electric propulsion engines, particularly of the plasma plume. The Navier-Stokes equations used in classic continuum computational fluid dynamics do not apply to the rarefied plasma, and therefore another method must be used to model the flow. A good solution is to use the DSMC method, which uses a combination of particle modeling and statistical methods for modeling the simulated molecules. A DSMC simulation known as SINATRA has been developed with the goal to model electric propulsion plumes. SINATRA uses an octree mesh, is written in C++, and is designed to be expanded by further research. SINATRA has been initially validated through several tests and comparisons to theoretical data and other DSMC models. This thesis examines expanding the functionality of SINATRA to simulate charged particles and make SINATRA a DSMC-PIC hybrid. The electric potential is calculated through a 7-point 3D stencil on the mesh nodes and solved with a Gauss-Seidel solver. It is validated through test cases of charged particles to demonstrate the accuracy and capabilities of the model. An ambipolar diffusion test case is compared to a neutral diffusion case and the electric field is shown to stabilize the diffusion rate. A steady state flow test case shows the simulation is able to stabilize and solve the electric potential for a plume-like scenario. It includes additional features to simplify further research including a comprehensive user manual, industry-standard version control, text file inputs, GUI control, and simple parallelism of the simulation. Compilation and execution are standardized to be simple and platform independent to allow longevity of the code base. Finally, the execution bottlenecks of linking particles to cells and particle moving were removed to reduce the simulation time by 95%.

A reusable oxidizer-cooled annular aerospike nozzle was designed for testing on a labscale PMMA-N20[1] hybrid rocket motor at Cal Poly-SLO.[2] The detailed design was based on the results of previous research involving cold-flow testing of annular aerospike nozzles and hot-flow testing of oxidizer-cooled converging-diverging nozzles. In the design, nitrous oxide is routed to the aerospike through a tube that runs up the middle of the combustion chamber. The solid fuel is arranged in an annular configuration, with a solid cylinder of fuel in the center of the combustion chamber and a hollow cylinder of fuel lining the circumference of the combustion chamber. The center fuel grain insulates the coolant from the heat of the combustion chamber. The two-phase mixture of nitrous oxide then is routed through channels that cool the copper surface of the aerospike. The outer copper shell is brazed to a stainless steel core that provides structural rigidity. The gaseous N2O flows from the end of aerospike to provide base bleed, compensating for the necessary truncation of the spike. Sequential and fully-coupled thermal-mechanical finite element models developed in Abaqus CAE were used to analyze the design of the cooled aerospike. The stress and temperature distributions in the aerospike were predicted for a 10-sec burn time of the hybrid rocket motor. [1] PMMA stands for polymethyl methacrylate, a thermoplastic commonly known by the brand name Plexiglas®. N2O is the molecular formula for nitrous oxide. [2] California Polytechnic State University, San Luis Obispo

Since 2000, a relatively new electric thruster concept has been in research, development, and production at Thales Electron Devices in Germany. This High Efficiency Multistage Plasma Thruster, or HEMPT, has promising lifetime capabilities due to its plasma confinement system. However, the permanent magnet system that offers this and other benefits also creates a hollow plume, where ions are accelerated at angles rather than up the thruster centerline, causing a dip in ion current along the centerline. A laboratory model, built at JPL, was run at Cal Poly to characterize this plume shape and implement a shield to restore a conical shape to the plume. A similar solution was used on a different type of thruster, a cylindrical hall thruster, at Princeton with excellent results. A shield was designed to shunt the magnetic field outside the thruster, where the Princeton experiments have identified a radial magnetic field as the cause for this hollow plume. The thruster was run with and without the shield, taking measurements of the ion current in the plume using a linear probe drive. The shield fixed the plume shape, increasing centerline current by 48%, however it also had detrimental effects on thruster performance, causing a decrease in thrust, specific impulse, and cut the total efficiency in half. The shield design was reexamined and a new design has been suggested for future testing of the HEMPT to restore performance while still fixing the plume shape.

The circulation control system onboard Cal Poly's Advanced Model for Extreme Lift and Improved Aeroacoustics was a critical component of a highly complex wind tunnel model produced in order to fulfill the requirements of a NASA Research Announcement awarded to David Marshall of the Aerospace Engineering Department. The model was based on a next generation, 150 passenger, regional, cruise efficient, short take-off and landing concept aircraft that achieved high lift through circulation control wings and over-the-wing mounted engines. The wind tunnel model was 10-ft in span, used turbine propulsion simulators, and had a functioning circulation control system driven from tunnel supplied high pressure air. Wind tunnel test results will be compiled into an open-source database intended for validation of predictive tools whose purpose is to advance the state- of-the-art in predictive capabilities for the next generation aircraft configurations. The model's circulation control system produced highly directional, nonuniform flow, and required significant modification in order to generate flow suitable for representation in predictive software. The effort and methods used to generate uniform flow along the circulation control slots is detailed herein. Additionally the results of the system characterization are presented and include a thorough analysis of the slot height, the wing symmetry, and total pressure at the circulation control jet exit. These datasets are intended to aid in making adjustments to the simulation such that it accurately reflects the condition at which the model was tested. Many flow visualization results from the wind tunnel test are also presented to serve as a medium of comparison for results from predictive tools. Oil flow visualization was conducted at many test conditions and provides insight to AMELIA's surface flow in blown and unblown regions. Of particular interest were streamlines at the wingblend, which exhibited some outboard turning, and streamlines on the lower surface where the leading edge stagnation point was investigated. Smoke flow visualization was also utilized to explore the flowfield. The deflection of a individual streamline, under the influence of a changing discharge coefficient as investigated along with the discharge coefficients effect on the extended flowfield. Collectively, the images depict the massive augmentation of the flowfield caused by the presence of the circulation control wing.

The FanWing propulsion system is a novel propulsion system which aerodynamically behaves as a hybrid between a helicopter and a fixed wing aircraft, and if the knowledge base with regards to this novel concept can be fully explored, there could be a new class of aircraft developed. In the current research, only 2D CFD studies have been done for the FanWing, hence the 3D lift characteristics of the FanWing have been unknown thus far, at least in the theoretical domain. Therefore, it was proposed to develop a modified Prandtl's Lifting Line Theory numerical solution and a CFD solution, comparing the results of each. A new variable was introduced into the classical Lifting Line Theory solution, αi,FW, to account for the additional lift produced by the FanWing as opposed to a traditional airfoil. This variable, αi,FW, is a function of the wing angle and the velocities taken at three-quarter chord length on the FanWing. The introduction of this variable was informed by other papers which superimposed velocities when developing Lifting Line Theory for unconventional airfoil planforms. After introducing a correction factor, the numerical model aligned with the 3D CFD results where LLT assumptions were valid. For the 3D simulation, it was observed that the lift per unit span rapidly increases from quarter span to wingtip, which is different from traditional wing planforms. This study provides a valuable first step towards documenting the 3D lift characteristics of the novel FanWing propulsion system.

Atomization of fuel is essential in controlling combustion inside a direct injection engine. Controlling combustion helps in reducing emissions and boosting efficiency. Cavitation is one of the factors that significantly affect the nature of spray in a combustion chamber. Typical fuel injector nozzles are small and operate at a very high pressure, which limit the study of internal nozzle behavior. The time and length scales further limit the experimental study of a fuel injector nozzle. Simulating cavitation in a fuel injector will help in understanding the phenomenon and will assist in further development. The construction of any simulation of cavitating injector nozzles begins with the fundamental assumptions of which phenomena will be included and which will be neglected. To date, there has been no consensus about whether it is acceptable to assume that small, high-speed cavitating nozzles are in thermal or inertial equilibrium. This diversity of opinions leads to a variety of modeling approaches. If one assumes that the nozzle is in thermal equilibrium, then there is presumably no significant delay in bubble growth or collapse due to heat transfer. Heat transfer is infinitely fast and inertial effects limit phase change. The assumption of inertial equilibrium means that the two phases have negligible slip velocity. Alternatively, on the sub-grid scale level, one may also consider the possibility of small bubbles whose size responds to changes in pressure. Schmidt et al. developed a two dimensional transient homogeneous equilibrium model which was intended for simulating a small, high speed nozzle flows. The HEM uses the assumption of thermal equilibrium to simulate cavitation. It assumes the two-phase flow inside a nozzle in homogeneous mixture of vapor and liquid. This work presents the simulation of high-speed nozzle, using the HEM for cavitation, in a multidimensional and parallel framework. The model is extended to simulate the non-linear effects of the pure phase in the flow and the numerical approach is modified to achieve stable result in multidimensional framework. Two-dimensional validations have been presented with simulation of a venturi nozzle, a sharp nozzle and a throttle from Winklhofer et al. Three-dimensional validations have been presented with simulation of ‘spray A’ and ‘spray H’ injectors from the Engine Combustion Network. The simulated results show that equilibrium assumptions are sufficient to predict the mass flow rate and cavitation incidence in small, high-speed nozzle flows.

Scramjet engines are key for sustained hypersonic flights. Analytic models play a critical role in the preliminary design of a scramjet engine configuration. The objective of this research is to develop and validate a quasi-1D model for the scramjet engine encompassing inlet, isolator and combustor, to evaluate the impact of flight conditions and design parameters on the engine functionality. The model is developed assuming isentropic flow in the inlet with a single turn; modified Fanno-flow equations in the isolator that account for the area change of the core flow; and the combustor is modeled using Heiser-Pratt equations accounting for the fuel mixing efficiency. The isolator and combustor models are validated against experimental results. The model accounts for twelve parameters allowing for a decent range of possible configurations. Finally, the model was applied to five sets of parametric studies to evaluate the effect of multiple parameters on the engine functionality.

Combustion emissions from aviation operations contribute significantly to climate change and air pollution. Accordingly, there is increasing interest in advancing battery-powered propulsion for aviation applications to reduce emissions. As batteries continue to improve, it is essential to recognize breakthroughs in battery specific energy in the context of air transport vehicles. Most electric aircraft designs and programs have focused on small aircraft because of restrictive battery performance. This work presents a feasibility assessment for an all-electric airliner based on an Airbus A220-100 with turbofan engines replaced by electric motors and propellers. The analysis compares the performance characteristics of the electric airliner to the A220-100 and establishes several configurations with varying battery pack-specific energy. The short-term electric airliner could replace conventional aircraft on very short, high-density missions. In contrast, the long-term electric airliner requires significant battery technology improvements that are not currently foreseen. The alternative long-term electric airliner could complete half of the A220-100’s missions, but the necessary specific energy value is also not anticipated shortly. All-electric airliners would significantly impact manufacturing, operations, costs, and emissions but are commercially infeasible with current battery technology. Additional development of more advanced battery technology is required to increase the specific energy of battery packs, enhance battery safety and reliability, and develop lighter high-power electric motors.

The need and demand for propulsion devices on nanosatellites has grown over the last decade due to interest in expanding nanosatellite mission abilities, such as attitude control, station-keeping, and collision avoidance. One potential micro-propulsion device suitable for nanosatellites is an electrothermal plasma thruster called Pocket Rocket. Pocket Rocket is a low-power, low-cost propulsion platform specifically designed for use in nanosatellites such as CubeSats. Due to difficulties associated with integrating propulsion devices onto spacecraft such as volume constraints and heat dissipation limitations, a characterization of the heat generation and heat transfer properties of Pocket Rocket is necessary. Several heat-transfer models of Pocket Rocket were considered as a part of this analysis to determine viability and complexity of the analysis, including a lumped thermal model, a finite-element model written in MATLAB, and a finite-volume model constructed using ANSYS Fluent and environmental conditions to closely reflect the experimental environment, both steady-state and transient. Results were validated experimentally. A Pocket Rocket thruster was manufactured for this purpose, and data regressed against model predictions to compare the validity of predicted models. Thermal models compared favorably to experimental measurements, accurately predicting the temperatures obtained at the surface of the thruster within 10 Kelvin after 1.5 hours of operation as well as the temporally-dependent temperature increases during the duration of operation within a standard error of ±6 Kelvin. Mission and integration viability is found to be favorable and within the realm of practicality for use of Pocket Rocket on nanosatellites.

An automated computer simulation of the two-dimensional planar Cal Poly Supersonic Ejector test rig is developed. The purpose of the simulation is to identify the operating conditions which produce the saturated, Fabri choke and Fabri block aerodynamic flow patterns. The effect of primary to secondary stagnation pressure ratio on the efficiency of the ejector operation is measured using the entrainment ratio which is the secondary to primary mass flow ratio. The primary flow of the ejector is supersonic and the secondary (entrained) stream enters the ejector at various velocities at or below Mach 1. The primary and secondary streams are both composed of air. The primary plume boundary and properties are solved using the Method of Characteristics. The properties within the secondary stream are found using isentropic relations along with stagnation conditions and the shape of the primary plume. The solutions of the primary and secondary streams iterate on a pressure distribution of the secondary stream until a converged solution is attained. Viscous forces and thermo-chemical reactions are not considered. For the given geometry the saturated flow pattern is found to occur below stagnation pressure ratios of 74. The secondary flow of the ejector becomes blocked by the primary plume above pressure ratios of 230. The Fabri choke case exists between pressure ratios of 74 and 230, achieving optimal operation at the transition from saturated to Fabri choked flow, near the pressure ratio of 74. The case of optimal expansion yields an entrainment ratio of 0.17. The entrainment ratio results of the Cal Poly Supersonic Ejector simulation have an average error of 3.67% relative to experimental data. The accuracy of this inviscid simulation suggests ejector operation in this regime is governed by pressure gradient rather than viscous effects.