An experimental investigation was conducted on premixed cavity stabilized flames in a high-speed ramjet engine, while varying the mean pressure gradients. The ramjet cavity was designed with a backward-facing step and an aft ramp for flame stabilization in regimes with high Reynolds numbers. To study the effects of mean pressure gradients on the engine performance, the ramjet engine underwent variations in wall geometry to create converging, diverging, and nominal configurations. The reacting flow fields and flame dynamics were captured using high-speed, simultaneous particle image velocimetry (PIV) and chemiluminescence imaging diagnostics. The study found that imposing a larger favorable pressure gradient led to a reduction in the recirculation zone and altered shear layer dynamics, resulting in an increased drag on the cavity. Additionally, a stronger favorable pressure gradient excited a shear layer instability mode with a Strouhal number of St = 0.1. The results from proper orthogonal decomposition (POD) analysis indicate that the instability mode comprised large-scale oscillations occupying the entire cavity flow region, indicating that the excited oscillations were due to a global vortex shedding instability under non-reacting conditions. The findings demonstrate that the mean pressure gradient can significantly influence the performance and stability of the ramjet cavity flame, which is crucial for designing high-speed air-breathing propulsion systems such as dual-mode scramjets.

NASA is financing a mission to study the surface of Titan, Saturn's largest moon, to investigate the terrain, chemical compositions, and potential for existent life. This mission is an exciting advancement from Mars Ingenuity, because the quadrotor lander, Dragonfly, will be the first of its kind with four coaxial rotors. Upon entry into Titan, and at near surface atmospheric conditions, Dragonfly will exit a parachute supported backshell that has shed its protective heatshield. The aerodynamics of this state are significant in comprehending the dynamics of the overall system for successful deployment into powered flight. The studies presented here examine the aerodynamic trends of the Dragonfly lander-backshell combination during Entry, Descent, and Landing (EDL), using computational fluid dynamics (CFD). More specifically, these investigations will focus on the Preparation for Powered Flight (PPF) sub-phase within EDL. The PPF mission phase begins immediately following heatshield separation, where the jettisoning of the heatshield generates an induced rotation on the lander-backshell system. This demands a "despin" capability to regain control authority prior to release of Dragonfly directly into powered flight. Preliminary evaluations of descent on Titan uncovered a suction force interaction between the rotor and lander body which opposed the current method of rotor control used for despin. The research proposes design modifications to regain control authority such as reassigning rotor control designations and altering the rotor cant. The computational model was benchmarked by comparing CFD results to experimental aerodynamic load measurements for similar backshells, bluff bodies, and rotor-body interactions. The model was adapted for Dragonfly to evaluate different descent configurations to gain a comprehensive understanding of the complex flow dynamics which is crucial in formulating strategies aimed at ensuring positive control authority of the system.

Calcium-magnesium-aluminosilicates (CMAS), such as sand or volcanic ash, are ingested by aircraft jet engines during operation. CMAS then becomes molten while traveling through the combustor of the engine before depositing onto turbine blades within the turbine section of the engine. The molten CMAS melt infiltrates and interacts with the high temperature ceramic coated turbine blades. This infiltration increases coating stiffness and promotes coating phase destabilization, encouraging micro-crack formation and increasing the risk of spallation. Thermomechanical effects from CMAS infiltration were mapped over time with confocal Raman spectroscopy. The residual stresses within infiltrated 7YSZ EB-PVD coatings were captured with microscale resolution. The results show an interplay between both the thermomechanical and thermochemical effects influencing the residual stress state of the coating. Thermomechanical mechanisms have a prominent role on the residual stress early on in a coating's CMAS exposure and after 1 h of infiltration, inducing tensile stresses within the coating up to 100 MPa on tetragonal ZrO2 Raman bands. Chemical mechanisms impart a greater influence on a much slower scale and after 10 h of infiltration, inducing compressive stresses within the coating up to 100 MPa. A monoclinic phase volume fraction of about 35% was observed to be a transitional point for thermochemical mechanisms overtaking thermomechanical mechanisms in dominating the residual stress of the coating. These results elucidate, in a non-destructive and non-invasive manner, changes within a coating's residual stress as a result of CMAS exposure and the subsequent CMAS infiltration over varying annealing times. The results aid in the efforts to monitor coating degradation during maintenance and towards implementing CMAS-mitigation strategies in not only 7YSZ EB-PVD coatings, but also as a reference for more novel coating compositions under development.

Detonations are a supersonic mode of combustion witnessed in a variety of applications, from next-generation propulsion devices to catastrophic explosions and the formation of supernovas. Detonations are typically initiated through the deflagration to detonation transition (DDT), a detailed process where a subsonic flame undergoes rapid acceleration increasing compressibility until a hotspot forms on the flame front inciting a detonation wave to form. Due to the complex nature of the phenomena, DDT is commonly investigated in three stages – (i) preconditioning, (ii) detonation onset, and (iii) wave propagation and stability. The research presented explores each of these stages individually, with a focus on preconditioning, to further resolve the governing mechanisms needed to initiate and sustain a detonation. More specifically, this work seeks to investigate the flow field and flame characteristics in reactions with increasing compressibility. Additionally, the research examines detonation onset and wave propagation to attain an all-encompassing concept of the DDT process. The work uses simultaneous high-speed diagnostics, consisting of particle image velocimetry (PIV), OH* chemiluminescence, schlieren and pressure measurements, to experimentally examine the preconditioning stage. Through the comprehensive suite of diagnostics, this research deduces the role of turbulence in detonation onset to an ongoing cycle of flame generated compression that amplifies until the hotspot ignites.

Blisk components are an integral part of jet engines and power generation systems. Composed of a single monolithic structure, integral blisks combine the blade-and-disk portion of primary components such as the fan, compressor stage, or turbine stage. This design modification increases the aerodynamic performance while decreasing the cost and eliminating time of assembly. However, these benefits come with the loss of blade-disk connection interfaces, which greatly reduces the structure's damping and its ability to attenuate unwanted structural vibrations. Blisk dynamics are impacted by mistuning, which is a disruption of the periodic symmetry of otherwise identical bladed sections. This disruption degrades the overall performance of components. This dissertation advances theoretical and experimental studies into mistuning, both in terms of characterization and vibration localization. A theoretical analysis of wave propagation in periodic structures shows that vibration localization arises from two separate generating mechanisms: an isolated defect or random variations throughout the structure. Through two novel metrics—the band flatness factor and the localization amplification criterion—we establish a unique approach to identifying localized modes on cyclic structures. We present a novel framework that distinguishes between the response of the tuned and mistuned system. A characteristic single-degree-of-freedom response of the tuned system to engine-order forcing forms the basis of our mistuning characterization framework. The actual mistuning (or tuned system) characterization is performed through computations of a modified modal assurance criterion, termed here the modal mistuning criterion. The modal mistuning criterion uses frequency response function data obtained from engine order-forcing profiles and tuned system modes derived from a model to compute the modal contribution to a mistuned response. Finally, this mistuning characterization metric is validated through experimental tests on multiple academic blisk components.

Shock tube experiments provide insightful data on combustion, emissions, and ablation characterization for a wide variety of defense and energy sector research topics. Accurate characterization of shock tube facilities is essential to verifying the accuracy of research data. Non-ideal effects, such as boundary layer growth, introduce uncertainty into pressure and temperature conditions, the primary independent variables of interest for shock tube research results. The effect of boundary layer growth on shock tube experiments was investigated by conducting simulations for University of Central Florida's two geometrically different shock tube using StanShock, a quasi-one-dimensional, reacting, compressible flow solver. The characteristic quantities considered for non-ideal effects and their impact on experiments is the post-reflected-shock pressure rise, dp*/dt, and the incident shock wave attenuation, which are calculated from simulated pressure data and developed into correlations for shock tube characterization and experiment planning.

Rotor-powered drones continue to grow in popularity in private and government sectors. The use of these drones in challenging environments and in high stakes applications calls for a certain level of robustness and redundancy. Often, these drones are equipped with sets of paired coaxial rotors, which not only improve the performance of the vehicle, but also ensure that a failure of one motor does not constitute the failure of the whole vehicle. Some applications such as extraterrestrial exploration, which use these coaxial rotors, can benefit from a wing shaped rotor arm to reduce drag and increase lift, extending mission lifetime. This work explores the design of one such coaxial rotor-wing system, using computational fluid dynamics to assess the system performance in a pair of flight conditions. Various parameters of the wing design are adjusted to ascertain the optimal configuration to satisfy various performance criteria.

The characterization of breakup mechanisms in a Liquid jet in Crossflow (LJIC) is of great importance to the propulsion industry. The current study focuses on analyzing and understanding these breakup mechanisms as it pertains to the operability conditions of airbreathing engines. These breakup mechanisms are studied by extracting and identifying spatial patterns and temporal dynamics of their coherent structures that define different modes of a breakup. These coherent structures associated to transverse jets are highly intermittent and cannot be classified by traditional global modal techniques. The primary objective of this study was to identify intermittent coherent structures associated with the four primary column breakup regimes: enhanced capillary breakup, bag breakup, multimode breakup, and shear breakup. The approach used in this study utilizes Proper orthogonal decomposition in conjunction with a novel technique known as the Multi-Resolution Dynamic Mode Decomposition (MrDMD). The applied methodology identifies coherent structures of the liquid surface and is an extension of the currently used Dynamic Mode Decomposition (DMD). The key benefit of MrDMD is it parses nonlinear dynamical systems into multi-resolution time-scaled components to capture intermittent mechanisms. The current methodology to extract MrDMD modes relies purely based on amplitude inspection. However, amplitude inspection becomes redundant for time-resolved snapshots of a highly turbulent system and is therefore not sustainable. In this study, a new method of modal extraction is utilized, this methodology is centered around using the repetition of frequency as detection for relevant modes rather than the magnitude of amplitude. This method will be referred to as Robust MrDMD. Robust MrDMD is applied to time-resolved series of column region snapshots for the four spray regimes. Relative frequencies of each breakup regime are extracted and identified. The frequencies for the characterized fuel jet injection dynamics are linked to critical non-dimensional parameters known as Strouhal number. The results produced a viable new breakup regime map associated with the strouhal number and weber number and its momentum flux ratio. Results conclusively show that all the breakup regimes contain a coherent structure associated with St=0.06-0.08, which dictates ligament breakup. It is also shown that coherent structures associated with small scale shear breakup of St 025-0.28 are shared between all the multimodal the shear breakup cases. These coherent structures are classified with an approximate time scale and found to correlate with a specified Strouhal number directly.

Conventional methods of classifying droplet breakup are evaluated in the context of unique variation in environmental and droplet fluid conditions. Most characterization is developed for subsonic speeds and Newtonian fluids, so this study extends understanding on how these forms change to a span of applications outside these conditions. Presented examples include the impact effects on hypersonic vehicles travelling through precipitation, where even smallest of rain drops at such speeds can cause damage. Before the droplet even reaches the vehicle, it interacts with the detached bow shock that leads it. Another example of exceptional recent concern is risk of viral transmission by breakup function within human saliva in sneezes, coughs and speaking. Such biofluid behavior is complicated by viscous and elastic properties, subject to molecular composition that varies person to person by function of their age, gender, and medical conditions. Both phenomena are difficult to image on the scale of internal droplet fluid flow and droplets of aerosolizing diameters. Thus, this study uses a multi-stage model that couples full scale simulations to simulations of a droplet scale. This multi-scale modelling approach develops a low cost computational method for system evaluation. The hypersonic impact model explores droplet breakup physics that resolve shock transmission through the droplet, with analysis of breakup driving factors of evaporation and cavitation. Similar studies are examined for the viscoelastic breakup of ejected saliva. The results indicate neither example can use conventional methods to characterize the droplet breakup seen. Droplets interacting with a shock experience internal fluid dynamics that present before the expected breakup form. Droplets of viscoelastic nature do not reach the expected breakup form, instead snapping back to prior shape. The results indicate that further experimental and simulation work is needed to address these unique conditions.

The following will outline the methodology and results of validating a coupled Method of Characteristics (MOC) and Direct Simulation Monte Carlo (DSMC) method. This research focused specifically on modeling plume impingement, induced by Reaction Control System (RCS) thrusters that flew on the National Aeronautics and Space Administration's (NASA's) space shuttle Discovery. For each simulation, the continuum portion of the RCS thruster was simulated using MOC for solving hyperbolic Partial Differential Equations (PDEs) and computed with the NASA code, Reacting and Multi-phase Program (RAMP). The solution was then implemented as a starting condition into the NASA DSMC code, Direct Simulation and Monte Carlo Analysis Code (DAC). Typically, DSMC models rely on code-to-code validation for fidelity. The significance of this research is in its ability to validate its models against empirical data. Prior to computing solutions for these simulations, the mesh size and structure were optimized and many variants of DSMC input parameters were iterated on in order to acquire a reliable, mesh-independent, fully optimized numerical solution. This research will discuss the mathematical formulation of MOC for nozzle flow and DSMC for rarefied gases. Additionally, it will provide an explanation of how to implement these mathematical concepts into the two solvers: RAMP and DAC. Ultimately, this research will demonstrate that the overall process illustrated produces results in good agreement with empirical data. As a consequence, the methodology presented is granted an increased level of confidence and will greatly contribute to the aerospace industry and its effort in understanding and predicting rarefied flow fields.