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.

No description available.

Accurate flow path modeling of scramjet engines is a key step in the development of an airframe integrated engine for hypersonic vehicles. A scramjet system model architecture is proposed and implemented using three different engine components: the isolator, combustor, and nozzle. For each component a set of intensive properties are iterated to match prescribed conditions, namely the mass flow. These low-fidelity one-dimensional models of hypersonic propulsion systems are used in tandem with Sandia Labs' Dakota optimization toolbox with the goal of accelerating the design and prototyping process. Simulations were created for the various components of the propulsion system and tied together to provide information for the entire flow-path of the engine given an inlet state. The isolator model incorporated methods to compute the intensive properties such as temperature and pressure of the flow path whether a shock-train exists internally as a dual-mode ramjet or if the engine is operating as a pure scramjet with a shock free isolator. A Fanno flow-like model was implemented to determine the friction losses in the isolator and a relation is iterated upon to determine the strength and length of the shock train. Two combustor models were created, the first of which uses equilibrium chemistry to estimate the state of the flow throughout the combustor and nozzle. Going one step further, the second model uses a set of canonical reactors to capture the non-equilibrium effects that may exist in the combustor/nozzle. The equilibrium combustor model was created to provide faster calculations in early iterations, and the reactor model was created to provide more realistic data despite its longer computational time. The full engine model was then compared and validated with experimental data from a scramjet combustor rig. The model is then paired with an optimization toolbox to yield a preliminary engine design for a provided design space, using a finite element analysis to ensure a feasible design. The implemented finite element analysis uses a coarse mesh with simple geometry to reduce computational time while still yielding sufficiently accurate results. The results of the optimization are then available as the starting point for higher fidelity analyses such as 2-D or 3-D computational fluid dynamics. / Master of Science / Ramjets and scramjets are the key to sustained flight at speeds above five times the speed of sound. These propulsion systems pose a challenging simulation environment due to the wide range of flow seen by the system structure. A scramjet simulation model is formulated using a series of combustion models with the goal of accurately modelling the combustion processes throughout the engine. The combustor model is paired with an isolator model and the engine model is compared against previous studies. A structural analysis model is then paired with the engine simulation, and the combined model is used within an optimizer to find an optimum design.

The concept of Magneto-Hydrodynamic (MHD) propulsion can be used to implement a propeller-less propulsion system for marine vehicles. The basic principle behind MHD is to use the (Lorentz) force produced by the interaction of electric and magnetic fields to generate thrust on a conducting fluid in motion. Electrodes are lined up along the walls of the duct which act as the source of the electric field. Seawater acts as the conducting medium for the current when it passes through the duct. This medium is then subjected to a strong magnetic field within the duct, thereby producing an axial force, i.e., an axial thrust. Propulsion systems based on MHD require virtually no mechanical components, therefore a good application would be to design a propulsor which produces very little noise for small underwater vehicles. Results of a preliminary feasibility study on this application are presented in this thesis. An approximate, consistent MHD propulsion theoretical model to assess the performance of a MHD propulsor for small underwater vehicles is introduced and analyzed. The model is generalized from the hydrodynamic point of view to consider inlet and outlet diffusers. The general model is applied systematically varying the main design parameters with respect to a given autonomous underwater vehicle (AUV) size. The results show that larger magnetic fields, longer propulsor lengths and smaller inlet flow speeds are preferred to get the highest propulsion efficiency and thrust. To check the consistency of the theoretical model, experiments are conducted. The results of these experiments show an approximate relation between the theoretical equations and the actual phenomenon. / Master of Science / In recent years, there has been an increase in the usage of small autonomous (unmanned) underwater vehicles (AUV) for various purposes such as exploration, mining and military applications. Most of these AUVs use the conventional system of a motor and propeller to drive the vehicle. This thesis proposes a different method of propulsion, one without any mechanical moving parts such as a rotor or a motor, for certain applications of these AUVs. The proposed system uses the concept of Magneto-Hydrodynamics (MHD) to propel the vehicle using an interaction between the applied magnetic and electric fields inside the propulsion channel. These applied fields produce a force (Lorentz) on the fluid that is present in the channel, thereby creating thrust to propel the vehicle. In the present case, the fluid is the electrically conducting seawater. Since, propulsion systems based on MHD require no mechanical components, they produce very little noise and are ideal for applications that require stealth. A feasibility study on this application is introduced, analyzed and presented in this thesis. Parameters such as applied fields, propeller configurations, and propeller shape and size are varied with respect to a given AUV size, to understand how each variable effects the system. The results show that larger magnetic fields, longer propulsor lengths and smaller inlet flow speeds are preferred to get the highest propulsion efficiency and thrust. To check the consistency of the theoretical model, experiments are conducted. The results of these experiments show an approximate relation between the theoretical equations and the actual phenomenon.

This thesis investigates low-energy xenon-molybdenum (Xe+-Mo) sputtering yields for ion energies of 100 eV and less. Sputtering yield data at these energies are important for ion thruster design and lifetime prediction. The basic principles of sputtering phenomena are discussed. An overview of various popular types of experimental sputtering yield methods is presented with an emphasis on the techniques that have been used to find Xe+-Mo sputtering yields in the past. Sputtering yields in this study are found through both models and experiments. Sputtering yields are calculated using the Sigmund, Bohdansky, Yamamura, and Wilhelm formulas. The computed sputtering yields for these models varied widely at low-energy. TRIM (The TRansport of Ions in Matter), a Monte-Carlo simulation program, was adapted to study sputtering yields, and energy and angular distributions of sputtered atoms. Simulations were run at various combinations of ion energy and ion incidence angle. TRIM did not prove to be an adequate model for low-energy sputtering. Experimental measurements of sputtering were made using both Rutherford backscattering spectrometry (RBS) and mass-loss methods. Sputtering was performed in a small vacuum facility using an ion gun. For the RBS technique, sputtered material was collected on aluminum foil substrates. The area density of the deposited Mo film on the substrates was measured using RBS. These measurements enabled calculation of differential sputtering yields, which were integrated to find the total sputtering yield. Sputtering yield was found by the mass-loss technique by simply comparing the mass of the sample both before and after sputtering using a microbalance. Sputtering yields at 100 eV, 90 eV, 80 eV, 70 eV, and 60 eV were found using the RBS technique. The mass-loss technique was only successful in the 80 eV experiment. The experimental results were unexpected. The measured sputtering yields were significantly higher than those reported by other researchers. Also, sputtering yields were found to increase with decreasing ion energy from 90 eV down to 60 eV. / Master of Science

ABSTRACT Developments in the field of microturbine technology and gas turbine driven aircraft has been progressing without much progress in light aircraft predominantly propelled by piston engines. Because of inhibitive maintenance and overhaul costs of such however, propulsion via a gas turbine engine has been proposed with the potential of eventually replacing current engine configurations. Subsequently, the objective was to conceptually design a replacement gas turbine engine in the 150 kW range. A selection of case studies was used to illustrate the changing technologies to illustrate the technological viability of micro-gas turbines for light aircraft. Advantages and disadvantages of both engine types were discussed and a concise description of gas turbine operations and its components was given. A brief overview of fundamentals as well as the transmission layout was also supplied. Three configurations were isolated, namely the single spool design, a twin spool design featuring a free power turbine and the effect of a fuel conserving recuperator. Calculations were performed using Microsoft Excel, which proved sufficient in effectively calculating complex formulae - even under the necessary iterative feed-back conditions the design process demanded. Eventually, variable-specific design criteria were derived regarding the three engine types. Because fuel consumption still proved inhibitive, the effect of recuperation was investigated which yielded a very competitive engine - should the possibility of recuperator technology exist on time. As a result, one particular recuperated, single spool gas turbine engine was successfully identified. Having met all the design criteria sufficiently, this preliminary prototype design was numerically described and put within context of principal, peripheral working components such as a compatible gearbox layout.