Increasing energy demands, and the subsequent need for cleaner energy conversion to combat climate change, creates a challenge that requires both short- and long-term solutions. To that end, new energy conversion cycles such as the Allam-Fetvedt cycle uses the combustion products (CO2) as the working fluid to increase efficiency and reduce emissions. There are several challenges regarding the implementation of these cycles, namely the extreme combustor conditions required (approximately 300 bar). The new High Pressure, Extended Range Shock Tube for Advanced Research (HiPER-STAR) was designed, built, and characterized to study combustion at these conditions to aid in the development of these sCO2 systems, among other extreme environments such as rocket chamber conditions. Further, development of chemical kinetics models used to predict combustion in these conditions typically assume reactions only in the homogeneous bulk gas region, while in these systems there are stagnation regions where hot gases are in contact with a heated wall for extended durations. Heterogeneous reactions are historically difficult to study, as typically there are coupled gas dynamic and transport-related complications that affect the reactions. A shock tube is an ideal location to mitigate and decouple these effects. The current work explores reactive and non-reactive end wall effects at high pressure, an area of interest for implementation by industry and resultantly where better efficiency can be achieved. Further designs have been completed and fabrication is underway to improve the capabilities of the facility to better decouple thermal wall effects and catalytic surface effects, as well as improve other combustion diagnostic capabilities of the facility.

Planar laser-induced fluorescence (PLIF) is a spectroscopic diagnostic method used widely in combustion research. In this study, imaging with formaldehyde as the tracer species was used in the diagnosis of jet engine performance at the UCF Propulsion and Energy Research Laboratory (PERL). PLIF imaging was first conducted on a laboratory Bunsen burner in order to validate the technique, identify the individual correction components, and demonstrate the results are consistent with other turbulent freejet formaldehyde PLIF literature. Once validated, PLIF imaging was then used to examine the concentration of formaldehyde in the reacting jet of a high pressure axially staged combustor. The results were processed to convert from recorded fluorescence to quantitative concentration profiles. This allowed for simultaneous visualization of the flame structure and the spatial distribution of formaldehyde vapor concentration in the reacting jet in crossflow for different equivalence ratios. Additionally, our concentration distributions in the instantaneous cross-sectional images showed regions of higher formaldehyde fluorescence near the preheat zone, and moderate formaldehyde fluorescence in the region preceding the preheat zone. Recommendations were made for improvements to the procedures used in this study for future work. Preliminary work was also done for the future integration of hydroxyl (OH) PLIF to be used simultaneously with formaldehyde PLIF for even more in-depth performance analysis.

Liquid fuels are desirable in aerospace applications due to their higher energy density when compared to gaseous fuels. With the advent of detonation-based engines, it is necessary to characterize and analyze how liquid fuel interacts with detonation waves as well as shocks to ignite. While liquid fuel sprays have been proven to successfully aid and sustain detonations, the physical mechanism by which the individual liquid droplets accomplish this is yet to be understood. Such knowledge allows for more predictable detonation properties, which in turn can let detonation-based engines be sustained more easily. This research seeks to quantify and characterize interactions of liquid fuels with detonations and shocks, analyzing the breakup mechanism as well as the ignition of select fuels. Such effects will be characterized for several different mixture compositions as well as shock and detonation speeds. Primary analysis techniques include shadowgraph, Schlieren, and chemiluminescence imaging. Data on pressure will also be taken with pressure transducers to confirm shock and detonation properties. This research will further the progress of liquid fuel detonation-based engines by enabling more predictable and sustainable detonations.

Although concerns about carbon dioxide (CO2) emissions and their impact on climate change has led to an increase in renewable energy electricity generation, natural gas power plants remain the dominant source of electricity generation in the United States. Until the capacity of renewable energy sources can meet growing electricity demand, natural gas power generation will likely remain an important source of electricity generation. Supercritical CO2 (sCO2) power generation cycles offer an alternative to traditional gas turbines by reusing and sequestering CO2 from the combustion process to prevent its release into the atmosphere. This study seeks to understand natural gas ignition in highly CO2 diluted mixtures at conditions relevant to sCO2 cycles, as well as rocket engines, using a high-pressure shock tube facility. Experiments were performed using a natural gas mixture (C1-C4 alkanes) with and without CO2 dilution for varying equivalence ratios at pressure up to 213 atm for a temperature range of 1016 K to 1286 K. Experiments were also performed using a second natural gas mixture (C1-C2), as well as using methane for a baseline comparison with similar studies. Ignition delay times were measured using OH radical emission measurements and compared to model predictions. Laser absorption spectroscopy measurements at a wavelength of 3.39 µm were used as a qualitative indicator of methane depletion during ignition. It was found that misinterpretation of OH radical emission measurements for experiments with significant reflected-shock bifurcation suggests better agreement with model predictions than observed using a combination of emission and absorption measurements. A comparison of chemical kinetics models shows inconsistent agreement with methane ignition measurements at 200 bar with CO2 and argon as the primary bath gases. Reaction pathway analysis was conducted to investigate the predicted effects of CO2 on chemical kinetics for these conditions. In addition, model predictions did not capture the effect of CO2 on natural gas ignition at 200 bar for the mixtures and temperature ranges studied. More data is needed to support the refinement of chemical kinetics mechanisms to better model methane and natural gas ignition in CO2 at 200 bar.

This thesis aims to perform a detailed analysis on a 5th Order Polynomial Nozzle, verifying its effectiveness in improving the thrust performance of a Rotating Detonation Rocket Engine. Rotating detonation engines are a promising engine type that uses detonations as a means of combustion rather than traditional conflagration. Through this method, these engines can produce significant amounts of energy while burning less fuel in the process. However, exhaust flow instabilities and swirl limit the engine's potential for use as a means of propulsion. The 5th Order Polynomial Nozzle was previously demonstrated to reduce and control this swirl; however, analysis was limited to side and back-end imaging. Using a recently built thrust stand, direct performance measurements were made with the nozzle being testing in several configurations. Discussed will be the data collected from the thrust stand, side-imaging to confirm flow behaviors similar to previous tests, and future work that is being done to analyze the exhaust flow.

Non-conventional bio-derived fuels have been evaluated for use in hybrid rocket motors. Tests were conducted at combustion pressures in the range of 100 – 220 psig and thrust levels of 40 – 170 newtons. Beeswax was tested with oxygen as the oxidizer and showed a regression rate at least three times as high as traditional hybrid propellant combinations such as hydroxyl-terminated polybutadiene (HTPB) and liquid oxygen (LOX). This provides the promise of a high thrust hybrid rocket motor using a simple, single port geometry and overcomes the main weakness of traditional hybrid rocket motor propellants, which are low regression rates. Beeswax was also tested with nitrous oxide as an oxidizer, but further testing is needed to attain high enough combustion chamber pressures to achieve stable combustion. Experimental evaluation of the specific impulse for beeswax and oxygen was moderately successful for lab scale testing, but needs further refinement. Analytical studies were performed to evaluate the theoretical performance of non-conventional hybrid rocket motors. This analysis indicates beeswax, lard, a mixture of paraffin and lard, and combinations of beeswax and aluminum should all perform better than traditional hybrid rocket propellants considered when burned with oxygen. For a combustion chamber pressure of 500.38 psig, beeswax and oxygen yielded a maximum specific impulse of 327 s. The high specific impulse combined with a high regression rate combine to make beeswax and oxygen a potentially high performing hybrid rocket motor propellant for launch vehicles, suborbital rockets, or orbital kick motors.

Applied research and testing has been conducted at the Cal Poly San Luis Obispo High-pressure Blow-Down facility to study the affects of throttling in a thrust-vectored aerospike nozzle. This study supports the ongoing research at Cal Poly to effectively thrust vector a hybrid rocket motor. Such thrust vectoring is achieved by small secondary ports in the nozzle body that are perpendicular to the main nozzle. The testing conducted included characterizing and comparing the performance of a straight aerospike nozzle to that of a thrust-vectored aerospike nozzle. Throttling effects on the aerospike nozzle in an unvectored and in a vectored configuration were also investigated. The interaction between throttling and thrust vectoring of an aerospike nozzle is the focus of this thesis research. This research shows that large-throat/high-thrust operation of an aerospike nozzle provides little thrust vector generation. Conversely, small-throat/low-thrust operation provides ample thrust vector generation. These results have implications in the effectiveness of thrust vectoring an aerospike nozzle with secondary ports. Rockets having an aerospike nozzle with throttling capabilities will be subject to the minimum and maximum turn angles for a given throttle position. As such, certain vehicle maneuvers might not be obtainable at certain throttle operations. Conversely, at lower throttling conditions, higher turn angles will be achievable.

Picosatellites are defined as satellites with a mass between 0.1 and 1kg (Miniaturized satellite). Picosatellites are typically designed to work together or function in formations (Miniaturized satellite). A specific type of Picosatellite known as CubeSats were introduced in 1999 and since then have increased in popularity so that there are now over 80 CubeSat programs around the world. CubeSats are defined as cubic units 10cm on each side and no more than 1kg in mass. CubeSats are required to conform to the CubeSat Standard created by California Polytechnic State University and Stanford University and be compatible with Cal Poly’s P-POD deployment system (Toorian, 2005). Some CubeSat uses include earth imaging, communications projects and various scientific experiments. CubeSats currently require attitude control and in the future, may require, maintaining a specific orbit, or changing orbit. With this ability many new activities may be possible for CubeSats. These activities could include rendezvous, vehicle inspection, formation flying and de-orbiting. For these activities to be possible, a high performance propulsion system is required. The goal of this thesis is to design and test an affordable, safe, and effective micro-propulsion system for CubeSats.

With the advent of smaller satellites, along with the need for less than 0.1 μN precision attitude control for interferometry and imaging missions, finer micro- to sub-micro- thrusters have become an area of high interest. As thrusters are developed and ground-tested, it is necessary to evaluate their thrust performance on-orbit. On-orbit measurements offer actual thrust performance in mission conditions, free from ground facility vibrations and miniaturization restraints, and allow a thruster system to achieve a NASA Technology Readiness Level (TRL) of 7-8. A review is conducted of existing and proposed ground and on-orbit thrust measurement techniques. Experimental gaps and complementary methods are examined along with the current thrust resolution limits. A novel fusion technique combining attitude determination, torsional balance, and filtering techniques is proposed to increase resolution beyond current on-orbit minimums, 4μN, via a dedicated sub-μN on-orbit thrust measurement mission. A simulated case study in the application of this measurement technique to a theoretical Casimir-thruster-equipped, 10-7-10-13 N, smallsat mission is explored. A detailed error analysis is conducted, and the technique is found to be analytically viable for greater than or equal to 10-7 N on a 1U nanosat equipped with sun sensor and three-axis gyroscope, as well as physically viable at a TRL 7-9 level. Recommended next steps are modification of the post-processing technique to decrease gyroscope noise and mass restrictions or exploration of suggested alternate methods, including orbit estimation, direct force sensing, and formation flying.

The objective of this project is to present a Lithium Ion battery voltage model derived using systems engineering principles. This paper will describe the details of the model and the implementation of the model in practical use in a power system. Additionally, the model code is described and results of the model output are compared to battery cell test data. Finally, recommendations for increased model fidelity and capability are summarized. The modeling theory has been previously documented in the literature but detailed implementation and application of the modeling theory is shown. The detailed battery cell test voltage profiles are proprietary; as such this project will not include axis values, often used in presentation of proprietary data in the public domain. The objective of this presentation is still achieved, as the modeling implementation and results are clearly demonstrated.