Hybrid motors that employ non-toxic, non-explosive components with a liquid oxidizer and a solid hydrocarbon fuel grain have inherently safe operating characteristics. The inherent safety of hybrid rocket motors offers the potential to greatly reduce overall operating costs. Another key advantage of hybrid rocket motors is the potential for in-flight shutdown, restart, and throttle by controlling the pressure drop between the oxidizer tank and the injector. This research designed, developed, and ground tested a closed-loop throttle controller for a hybrid rocket motor using nitrous oxide and hydroxyl-terminated polybutadiene as propellants. The research simultaneously developed closed-loop throttle algorithms and lab scale motor hardware to evaluate the fidelity of the throttle simulations and algorithms. Initial open-loop motor tests were performed to better classify system parameters and to validate motor performance values. Deep-throttle open-loop tests evaluated limits of stable thrust that can be achieved on the test hardware. Open-loop tests demonstrated the ability to throttle the motor to less than 10% of maximum thrust with little reduction in effective specific impulse and acoustical stability. Following the open-loop development, closed-loop, hardware-in-the-loop tests were performed. The closed-loop controller successfully tracked prescribed step and ramp command profiles with a high degree of fidelity. Steady-state accuracy was greatly improved over uncontrolled thrust.

A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor’s performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

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The two body problem and the rocket equation r̈ + ∊ α ṙ + k/r3r = 0 have been expressed in numerous ways. However, the combination of the rocket equation with the two-body problem has not been studied to any degree of depth due to the intractability of the resulting non-linear, non-homogeneous equations. The goal is to use perturbation techniques to approximate solutions to the combined two-body and rocket equations.

A numerical investigation on the combustion mechanics of a hybrid rocket engine is performed through unsteady Reynolds-averaged Navier-Stokes simulation. The hybrid rocket engine model is based on an experimental laboratory scale engine design operating on GOX and HDPE as a propellant. A simple convection heat flux model is used to determine the heat transfer to the fuel wall. The project is done with the goal of finding the fuel regression rate in mind, as it is an essential parameter for determining engine performance. The results show early results of the fluid- and thermodynamics occurring in the combustion chamber. Propellant mixing is shown to not be optimal as a significant part of the exit flow consists of high concentrations of oxidizer that has not reacted with the fuel. The flame temperature is shown to be relatively high inside the combustion chamber. It is concluded from the simulations that the model needs further improvement in order to accurately compute the flow as well as the heat transfer to the fuel. To determine the regression rate, radiation should be implemented into the heat transfer model.

Microphone arrays are useful for measuring acoustic energy quantities (e.g. acoustic intensity) in the near-field of a full-scale solid rocket motor. Proper characterization of a rocket plume as a noise source will allow for more accurate predictions in engineering models that design for protection of structures, payloads and personnel near the rockets. Acoustic intensity and energy density quantities were measured in three rocket noise fields and have shown that the apparent source region of the rocket becomes smaller and moves upstream as frequency increases. Theoretical results accounting for some scattering and finite-difference errors arising in these types of energy-based measurements have been previously discussed by other authors. This thesis includes results from laboratory experiments which confirm some of this previous theoretical work as well as gives insight into the physical limitation of specific microphone array designs. Also, calibrations for both magnitude and directional response of the microphones are demonstrated. Of particular interest is the efficacy of phase calibration of array microphones for the low-frequency regime below 200 Hz.

The application of emission spectroscopy to monitor combustion products of solid rocket propellant combustion can potentially yield valuable data about reactions occurring within the volatile environment of a strand burner. This information can be applied in the solid rocket propellant industry. The current study details the implementation of a compact spectrometer and fiber optic cable to investigate the visible emission generated from three variations of solid propellants. The grating was blazed for a wavelength range from 200 to 800 nm, and the spectrometer system provides time resolutions on the order of 1 millisecond. One propellant formula contained a fine aluminum powder, acting as a fuel, mixed with ammonium perchlorate (AP), an oxidizer. The powders were held together with Hydroxyl-Terminated-Polybutadiene (HTPB), a hydrocarbon polymer that is solidified using a curative after all components are homogeneously mixed. The other two propellants did not contain aluminum, but rather relied on the HTPB as a fuel source. The propellants without aluminum differed in that one contained a bimodal mix of AP. Utilizing smaller particle sizes within solid propellants yields greater surface area contact between oxidizer and fuel, which ultimately promotes faster burning. Each propellant was combusted in a controlled, non-reactive environment at a range of pressures between 250 and 2000 psi. The data allow for accurate burning rate calculations as well as an opportunity to analyze the combustion region through the emission spectroscopy diagnostic. It is shown that the new diagnostic identifies the differences between the aluminized and non-aluminized propellants through the appearance of aluminum oxide emission bands. Anomalies during a burn are also verified through the optical emission spectral data collected.

Control of vortex-driven instabilities was demonstrated via a scaled-down, cold-flow simulation that modeled closed-end acoustics. When vortex shedding frequencies couple with the natural acoustic modes of a choked chamber, potentially damaging low-frequency instabilities may arise. Although passive solutions can be effective, an active control solution is preferable. An experiment was performed to demonstrate an active control scheme for the reduction of vortex-driven oscillations. A non-reacting experiment using a primary flow of air, where both the duct exit and inlet are choked, simulated the closed-end acoustics. Two plates, separated by 1.27 cm, produced the vortex shedding phenomenon at the chamber's first longitudinal mode. Two active control schemes, closed-loop and open-loop, were studied via a cold-flow simulation for validating the effects of reducing vortex shedding instabilities in the system. Actuation for both control schemes was produced by using a secondary injection method. The actuation system consisted of pulsing compressed air from a modifed, 2-stroke model airplane engine, controlled and powered by a DC motor. The use of open-loop only active control was not highly effective in reducing the amplitude of the first longitudinal acoustic mode, near 93 Hz, when the secondary injection was pulsed at the same modal frequency. This was due to the uncontrolled phasing of the secondary injection system. A Pulse Width Modulated (PWM) signal was added to the open-loop control scheme to correct for improper phasing of the secondary injection flow relative to the primary flow. This addition allowed the motor speed to be intermittently increased to a higher RPM before returning to the desired open-loop control state. This proved to be effective in reducing the pressure disturbance by approximately 46%. A closed-loop control scheme was then test for its effectiveness in controlling the phase of the secondary injection. Feedback of the system's state was determined by placing a dynamic pressure transducer near the chamber exit. Closed-loop active control, using the designed secondary injection system, was proven as an effective means of reducing the problematic instabilities. A 50% reduction in the FFT RMS amplitude was realized by utilizing a Proportional-Derivative controller to modify the phase of the secondary injection.

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.