As industry transitions to a net-zero carbon future, turbulent premixed combustion will remain an integral process for power generating gas turbines and are also desired for aviation engines due to their ability to minimize pollutant emissions. However, accurately predicting the behavior of a turbulent reacting flow field remains a challenge. To better understand the dynamics of premixed reacting flows, this study experimentally investigates the evolution of turbulence in a high-speed bluff-body combustor. The combustor is operated across a range of equivalence ratios from 0.7-1 to quantify the role of heat release and flame scales on the evolution of turbulence as the flow evolves from reactants to products. High-speed particle image velocimetry and CH* chemiluminescence imaging systems are simultaneously employed to quantify turbulent flame and flow dynamics. The results demonstrate that the flame augments turbulence fluctuations as the flow evolves from reactants to products for all cases. However, turbulence fluctuations increase monotonically with the heat of combustion and corresponding turbulent flame speed. Nondimensional spatial profiles of turbulence are used to develop a correlation to predict the increase in turbulent fluctuations in an extended progress variable space. A Reynolds Averaged Navier Stokes (RANS) decomposition is also explored to better characterize the effects of heat release on turbulence evolution dynamics. The correlations and RANS decomposition can guide modeling capabilities to better predict confined turbulent reacting flows and accelerate design strategies for premixed turbines with carbon-free fuels.

The Rotating Detonation Combustor (RDC) has, in recent years, been a subject of great interest in the pressure-gain combustion research community. Researchers have been investigating the RDC as a potential improvement over current combustors in today's turbomachinery-based power generation systems. With the theoretical efficiency boost that detonations provide over deflagrating combustors, RDCs have the promise to be a next step in fuel/cost savings in the power generation industry. One mode of research to push an RDCs capabilities is the potential use of combustible carbon/hydrocarbon solid particles in addition to liquid or gaseous fuels. These particles are a source of high energy density and, once added, can reduce the amount of liquid/gaseous fuel needed while operating at the same fuel-to-air ratio. These organic particles are derived from grown sources making them cost-effective and sustainable, in contrast to mined or drilled fossil fuels. Carbon black, peanut flour, and powdered sugar were seeded into a 6-inch diameter RDC operating on a gaseous hydrogen-air mixture. This was done at the leanest hydrogen-air ratio possible where the combustor, operating on just gaseous fuel, would still experience stable detonation waves. Solid fuel was then seeded in place of the gaseous fuel at varying ratios to study its effects on the ability of the combustor to continue to experience detonations. In general, while stable detonations were achieved when solid fuel began replacing the gaseous fuel, the greater the concentration of solid particles compared to gaseous fuel, the greater the likelihood of irregular detonation modes. These modes were observed using a high-speed camera: taking back-end imaging observations to measure characteristics of present detonation waves, including wave number, speed, stability, and phase angle. CTAPs were also added along the length of the outer body of the RDC to take pressure measurements during operation.

Temperature measurements can be difficult to obtain across many different harsh environments such as engine combustion chambers, engine exhaust temperatures, and explosion fireballs. While there are alternate methods to measure fluid temperature such as laser measurements, acoustic measurements, and camera imaging techniques, these methods can often be expensive, difficult to implement, and not able to see within the environment. Thermocouples are popular sensors because they are cheap and easy to implement across a wide range of applications and can measure temperature in areas where other methods cannot reach or see. However, while these sensors are very popular and versatile, they do have some disadvantages, mainly, the response time. When the testing environment becomes harsh, the thermocouple size increases so that the sensor can survive. Unfortunately, when the thermocouple size increases, so does the time that it takes to sense the gas temperature. For this research, the environment will mimic an explosive environment with very fast temperature rise times that will require quick sensor response. This will not be achievable with a single thermocouple; so, multiple thermocouples will be used. This research focuses on evaluating past multi-thermocouple reconstruction techniques to determine which available method is the most accurate and feasible to implement. Of the methods researched, this work has found that a frequency domain method proposed by Forney and Fralick provides temperature estimates as accurate as 0.5% off the average steady state temperature with an average percent error of 5%.

Fluids-related issues in the Aerospace industry are often multiphase in scope. Numerical modeling, such as computational fluid dynamics, is used to describe these problems, as they are difficult or impossible to describe analytically. This research uses computational fluid dynamics to describe multiphase problems related to melting-solidification and particle impingement. Firstly, a numerical model was established that uses the Volume-of-Fluid method to resolve a melting/solidifying particle. This model was verified against experiments and simplified analytical models, and a mesh independence study was done to ensure the results were independent of the mesh sizing. Next, the model was applied to two separate but related problems. The Artemis program has renewed interest in lunar dust mitigation. It is proposed that lunar regolith partially melts and becomes "sticky" when coming into contact with a jet flame, like a landing rocket. The method above was applied to a lunar regolith particle to show how these "sticky" particles can adhere to surfaces. The direct resolution methodology was also applied to a melted sand particle impinging and infiltrating a yttria-stabilized zirconia thermal barrier coating, as seen in engine turbines. Sand can infiltrate the thermal barrier coating and decrease its effectiveness. The infiltration from a single particle was compared to the infiltration from a stream of melted sand. These three efforts showcase the usefulness of directly resolving small particles using the Volume-of-Fluid method.

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.

No description available.

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.

No description available.