A successful transition to clean energy hinges on meeting the world's growing energy demand while reducing greenhouse gas emissions. Achieving this will require significant growth in electricity generation from clean and carbon-free energy sources. Several energy providers have already begun the transition from traditional carbon-based fuels to cleaner alternatives, such as hydrogen and hydrogen enriched natural gas. However, there are still many technical challenges that must be addressed when applying these fuels in gas turbines. The application of hydrogen or hydrogen/natural gas blends to advanced class gas turbines, which have higher operating pressures and temperatures has raised concerns about the potential for leakages or fuel sequencing operations where flammable mixtures of fuel and air could auto-ignite. Public information on the auto-ignition of hydrogen in air at atmospheric pressure is well documented. Such data shows the auto-ignition temperature of hydrogen is roughly 100 °C lower than that of methane. Studies also show that as pressure increases, methane's auto-ignition temperature decreases. However, there was insufficient information in the published literature to characterize the influence of pressure on auto-ignition for hydrogen fuel applications. This study describes the test methodology used to evaluate conditions where auto-ignition occurs for various fuel-air mixtures operating at pressures between 1-30 atmospheres and equivalence ratios between 0.2-1.6. Testing was completed with hydrogen, natural gas and blends at various equivalence ratios using a heated volume with multiple reactant delivery methods. Testing was performed for natural gas to validate the test and data collection methods cited in prior published literature. Results indicate that at atmospheric pressures, an increase in hydrogen concentration results in a reduced auto-ignition temperature. However, at 30 atmospheres, the auto-ignition temperature increased with higher hydrogen concentrations. iv Variations of auto-ignition delay times were also observed during the testing and are compared to modeling predictions, providing insight into auto-ignition characteristics.

Detonation is a high energetic mode of pressure gain combustion that exploits total pressure rise to augment high flow momentum and thermodynamic cycle efficiencies. Detonation is initiated through the Deflagration-to-Detonation Transition (DDT). This process occurs when a deflagrated flame is accelerated through turbulence induction, producing shock-flame interactions that generate violent explosions and a supersonic detonation wave. There is a broad desire to unravel the physical mechanisms of turbulence induced DDT. For the implementation of efficient detonation methods in propulsion and energy applications, it is crucial to understand optimum turbulence conditions for detonation initiation. The study examines the role of turbulence-flame interactions on flame acceleration using a fluidic jet to generate turbulence within the reactant flow field. The investigation aims to classify the turbulent flame dynamics and temporal evolution of the flame stages throughout the turbulent flame regimes. The flame-flow interactions are experimentally studied using a detonation facility and high-speed imaging techniques, including Particle Image Velocimetry (PIV) and Schlieren flow visualization. Flow field measurements enable local turbulence characterization and analysis of flame acceleration mechanisms that result from the jet's high level of turbulent transport. The influence of initial flame turbulence on the turbulent interaction is revealed, resulting in higher turbulence generation and overall flame acceleration. Turbulent intensities are classified, revealing a dynamic fluctuation of flame structure between the thin reaction zone and the broken reaction regime throughout the interaction.

The propagation of a shock and flame from a detonation wave injected orthogonally into a combustible, supersonic flow was observed. The detonation wave was generated through the use of a miniaturized detonation tube, henceforth referred to as a predetonator. Conditions within the test section, including stagnation pressure and equivalence ratio, were varied between cases. Through the use of high-speed schlieren, shadowgraph, and broadband OH chemiluminescence imaging, the leading shock and reaction were recorded as they moved through the test section. Variation of stagnation pressure affected the propagation of the leading shock. Higher stagnation pressures caused greater deflection of the shock wave and jet issued by the predetonator. It was seen that at sufficiently high equivalence ratios, the shock and reaction were able to travel upstream from the test section into the diverging section of the converging-diverging nozzle. Shortly after the shock entered the nozzle, evidence of the initiation of shock induced combustion was observed. Stagnation pressure variation in the range tested had little effect on the ability to initiate a reaction. Multiple behaviors of the shock-induced-combustion were observed, dependent upon the equivalence ratio of the flow through the test section. Behaviors include sustained reaction on the edges of the flow, sustained reaction in the core of the flow, and periodic, non-sustained reaction.

Most combustion machines feature internal flows with very high energy densities. If a small fraction of the total energy contained in the flow is diverted into oscillations, large mechanical or thermal loads on the structure can be the result, which are potentially devastating if not predicted correctly. This is particularly the case for lightweight high performing devices like rockets. The problem is commonly known as "Combustion Instability". Several mechanisms have been identified in the past that link the flow field to the acoustics inside a combustion chamber and thereby drive or dampen oscillations, one of them being vortex shedding. The interaction between the highly sheared flow behind an obstacle and longitudinal acoustic oscillations inside a solid rocket booster is investigated both analytically and experimentally.The analytical approach is developed based on modeling of the second order acoustic energy. The energy model is applied to the specific flow conditions just downstream of a single baffle protruding into the flow. The mean flow profile is assumed to be of the form of a hyperbolic tangent, the unsteady acoustic velocities are assumed to be sinusoidally oscillating. Solutions for the unsteady rotational velocities and the unsteady vorticity are derived. The resulting flow field is utilized in stability calculations for a simplified two-dimensional axial-symmetric geometry. This yields to linear growth rates of the (longitudinal) oscillation modes. The growth rates are functions of the chamber geometry, the mean flow properties and the properties of the shear layer created by the flow restriction.A cold flow experiment is designed, tested and performed in order to validate the analytical findings. Flow is injected radially into a tube with acoustic closed-closed end conditions. A single baffle is installed in the tube, the axial position of the baffle is varied as well as its inner diameter. Frequency spectra of pressure oscillations are recorded. The experimental data is then compared qualitatively to the analytical growth rates. Those longitudinal Normal Modes, which feature the highest theoretical growth rates, are expected to be most prominent in the experimental data. This behavior is clearly observable.

The focus of this study is the prediction of trajectories of solid particles injected into either a cylindrically- shaped solid rocket motor (SRM) or a bidirectional vortex chamber (BV). The Lagrangian particle trajectory is assumed to be governed by drag, virtual mass, Magnus, Saffman lift, and gravity forces in a Stokes flow regime. For the conditions in a solid rocket motor, it is determined that either the drag or gravity forces will dominate depending on whether the sidewall injection velocity is high (drag) or low (gravity). Using a one-way coupling paradigm in a solid rocket motor, the effects of particle size, sidewall injection velocity, and particle-to-gas density ratio are examined. The particle size and sidewall injection velocity are found to have a greater impact on particle trajectories than the density ratio. Similarly, for conditions associated with a bidirectional vortex engine, it is determined that the drag force dominates. Using a one-way particle tracking Lagrangian model, the effects of particle size, geometric inlet parameter, particle-to-gas density ratio, and initial particle velocity are examined. All but the initial particle velocity are found to have a significant impact on particle trajectories. The proposed models can assist in reducing slag retention and identifying fuel injection configurations that will ensure proper confinement of combusting droplets to the inner vortex in solid rocket motors and bidirectional vortex engines, respectively.

Rocket-Based Combined Cycle (RBCC) engines combine the best performance characteristics of air-breathing systems such as ramjets and scramjets with rockets with the goal of increasing payload/structure and propellant performance and thus making LEO more readily accessible. The idea of using RBCC engines for Single-Stage-To-Orbit (SSTO) trans-atmospheric acceleration is not new, but has been known for decades. Unfortunately, the availability of detailed models of RBCC engines is scarce. This thesis addresses the issue through the construction of an analytical performance model of an ejector rocket in a dual combustion propulsion system (ERIDANUS) RBCC engine. This performance model along with an atmospheric model, created using MATLAB was designed to be a preliminary `proof-of-concept' which provides details on the performance and behavior of an RBCC engine in the context of use during trans-atmospheric acceleration, and also to investigate the possibility of improving propellant performance above that of conventional rocket powered systems. ERIDANUS behaves as a thrust augmented rocket in low speed flight, as a ramjet in supersonic flight, a scramjet in hypersonic flight, and as a pure rocket near orbital speeds and altitudes. A simulation of the ERIDANUS RBCC engine's flight through the atmosphere in the presence of changing atmospheric conditions was performed. The performance code solves one-dimensional compressible flow equations while using the stream thrust control volume method at each station component (e.g. diffuser, burner, and nozzle) in all modes of operation to analyze the performance of the ERIDANUS RBCC engine. Plots of the performance metrics of interest including specific impulse, specific thrust, thrust specific fuel consumption, and overall efficiency were produced. These plots are used as a gage to measure the behavior of the ERIDANUS propulsion system as it accelerates towards LEO. A mission averaged specific impulse of 1080 seconds was calculated from the ERIDANUS code, reducing the required propellant mass to 65% of the gross lift off weight (GLOW), thus increasing the mass available for the payload and structure to 35% of the GLOW. Validation of the ERIDANUS RBCC concept was performed by comparing it with other known RBCC propulsion models. Good correlation exists between the ERIDANUS model and the other models. This indicates that the ERIDANUS RBCC is a viable candidate propulsion system for a one-stage trans-atmospheric accelerator.

Because compressibility effects arise in real applications of propellers and turbines, the Actuator Disk Theory or Froude’s Momentum Theory was established for compressible, subsonic flow using the three laws of conservation and isentropic thermodynamics. The compressible Actuator Disk Theory was established for the unducted (bare) and ducted cases in which the disk was treated as the only assembly within the flow stream in the bare case and enclosed by a duct having a constant cross-sectional area equal to the disk area in the ducted case. The primary motivation of the current thesis was to predict the ideal performance of a small ram-air turbine (microRAT), operating at high subsonic Mach numbers, that would power an autonomous Boundary Layer Data System during test flights. The compressible-flow governing equations were applied to a propeller and a turbine for both the bare and ducted cases. The solutions to the resulting system of coupled, non-linear, algebraic equations were obtained using an iterative approach. The results showed that the power extraction efficiency and the total drag coefficient of the bare turbine are slightly higher for compressible flow than for incompressible flow. As the free-stream Mach increases, the Betz limit of the compressible bare turbine slightly increases from the incompressible value of 0.593 and occurs at a velocity ratio between the far downstream and the free-stream that is lower than the incompressible value of 0.333. From incompressible to a free-stream Mach number of 0.8, the Betz limit increases by 0.021 while its corresponding velocity ratio decreases by 0.036. The Betz limit and its corresponding velocity ratio for the ducted turbine are not affected by the free-stream Mach and are the same for both incompressible and compressible flow. The total drag coefficient of the ducted turbine is also the same regardless of the free-stream Mach number and the compressibility of the flow; but, the individual contributions of the turbine drag and the lip thrust to the total drag differs between compressible and incompressible flow and between varying free-stream Mach numbers. It was concluded that overall compressibility has little influence on the ideal performance of an actuator disk.

Lean flame blowout is investigated experimentally within a high-speed combustor to analyze the temporal extinction dynamics of turbulent premixed bluff body stabilized flames. The lean blowout process is induced through fuel flow reduction and captured temporally using simultaneous high-speed particle imaging velocimetry (PIV) and CH* chemiluminescence. The evolution of the flame structure, flow field, and the resulting strain rate along the flame are analyzed throughout extinction to distinguish the physical mechanisms of blowout. Flame-vortex dynamics are found to be the main driving mechanism of flame extinction; namely, a reduction of flame-generated vorticity coupled with an increase of downstream shear layer vorticity. The vorticity dynamics are linked to hydrodynamic instabilities that vary as a function of the decreasing equivalence ratio. Frequency analysis is performed to characterize the dynamical changes of the hydrodynamic instability modes during flame extinction. Additionally, various bluff body inflow velocity regimes are investigated to further characterize the extinction instability modes. Both equivalence ratio and flow-driven instabilities are captured through a universal definition of the Strouhal number for the reacting bluff body flow. Finally, a Karlovitz number-based criterion is developed to consistently predict the onset of global extinction for different inflow velocity regimes.

Performance characterization was undertaken for an air augmented rocket mixing duct with annular cavity configurations intended to produce thrust augmentation. Three mixing duct geometries and a fully annular cavity at the exit of the nozzle were tested to enable thrust comparisons. The rocket engine used liquid ethanol and gaseous oxygen, and was instrumented with sensors to output total thrust, mixing duct thrust, combustion chamber pressure, and propellant differential pressures across Venturi flow measurement tubes. The rocket engine was tested to thrust maximum, with three different mixing ducts, three major combustion pressure sets, and a nozzle exit plane annular cavity (a grooved ring). The combustion pressures tested were , , and allowing for a nozzle pressure ratio range of relative to ambient pressure. The mixture ratio was fuel rich throughout all tests. The engine operated very consistently throughout all the tests performed; however, pressure losses in the feed system prevented higher combustion pressures from being tested. Three mixing ducts of the same outer diameter were tested. The short and diverging ducts were the same length and the long duct was long. The short and long ducts created positive mixing duct thrust and the diverging duct created negative mixing duct thrust. The long duct case did show better performance than the no duct case when the total thrust was divided by combustion pressure and nozzle throat area. The long duct always created several times more mixing duct thrust than either the short or diverging ducts, but none of the mixing ducts created positive overall thrust augmentation in the over expanded cases tested. The mixing duct thrusts ranged between and . As the combustion pressures were increased, getting closer the nozzle’s optimal expansion, the mixing duct thrusts started converging indicating a difference between nozzle operation at over expanded and under expanded. The annular cavity had a noticeable effect on the thrust of the engine and the appearance of the plume. The total thrust of the system was decreased by a maximum of and the plume was more sharply defined when the annular cavity was attached. Better mixing between the primary (engine exhaust) flow and the secondary (ambient air) flow was promoted by the annular cavity because it increased the shear layer’s turbulence and the increased turbulence reduced thrust. The greater mixing also allowed for secondary combustion which made the plumes more sharply defined. The annular cavity was also seen to enhance the mixing duct thrusts for all three mixing ducts.

A wind tunnel investigation of an axisymmetric bluff body, known as a Goldschmied propulsor, was completed. This model conceptually combines boundary layer control and boundary layer ingestion into a single complementary system that is intended to use energy to reduce the axial force on the body by eliminating separation and increasing the pressure recovery aft of the body’s maximum thickness. The goal of the current project was to design, fabricate, and fully document the performance of a wind tunnel model incorporating the Goldschmied propulsor concept and complete an examination of its aerodynamic performance. The investigation took place at California Polytechnic State University, San Luis Obispo in the Aerospace Engineering Department’s subsonic 3ft by 4ft wind tunnel. The model is 38.5 inches in length and 13.5 inches in diameter with a discrete suction slot at 85% of the body length and an embedded propulsor that provides the suction flow, expelling it out of the model’s aft end. The experiment included measurements of surface pressure, total axial force, suction mass flow rate, fan thrust, fan torque, fan speed, and input fan power. The size of the suction slot and amount of input fan power were the main test variables in the 54 data point test matrix that was completed at a length Reynolds number of 1.34 million and a tunnel speed of 66 ft/s (20 m/s). The model was able to achieve fully attached flow on the aftbody with as little as 100W of input power and a net positive (forward) axial force coefficient of 0.12 with as little as 200W of input power. The model was also able to achieve a peak axial pressure force coefficient of 0.005 in the forward direction with an input power of 500W and a slot gap of 1.6% of the body length. A slightly lower axial pressure force coefficient of 0.0045 was achieved with only 200W of input power and a slot gap of 0.7% of the body length. The peak axial pressure force for most tested slot gaps occurred at about 200W of input power, and a slot gap of 0.7% of the body length resulted in the best overall performance for most input power settings. Two different suction slot configurations, a simple gap and a cusp, were tested, and no significant performance differences were seen between them. The pressure coefficient data showed similar trends as test data from 1956 of a similar model at higher Reynolds number, but it did not show complete agreement. Despite these positive aspects of the investigation, a simple power based comparison between the collected data and a conventional non-integrated propulsor does not show a performance improvement for the Goldschmied propulsor.