<p dir="ltr">Rotating detonation engines (RDEs) are a promising technology by which to increase the efficiency of propulsion and power generation systems. Self-sustained, rotating detonation waves within the combustion chamber provide a means for combustion to occur at elevated local pressures, theoretically resulting in hotter temperature product gas than a constant pressure combustion process could provide at equivalent operating conditions. Despite theoretical advantages of RDEs, the thermodynamic benefit has yet to be achieved in experimental applications. Additionally, much of the experimental work to date has been conducted at mean operating pressures lower than industrial applications will require, especially for rocket or gas turbine combustion environments. The sensitivity of these devices to operating pressure has made clear the importance of chemical reaction rates on the successful operation of these combustors. This work addresses critical risks associated with implementing this technology at flight-relevant conditions by advancing the understanding of deflagrative loss mechanisms on delivered performance and by investigating the coupling between chemical kinetic timescales and operating modes produced by the combustor.</p><p dir="ltr">A novel pressure measurement technique was developed in which the stagnation pressure of exhausting gas produced by the RDC is measured through quantification of the under-expanded exhaust plume divergence angle at megahertz-rates. Time-averaged stagnation pressure measurements obtained with this technique are shown to be within 1.5% of the equivalent available pressure (EAP) measured. Time-resolved stagnation pressure measurements produced by this technique provide a means to quantify the detonation cycle pressure ratio. It was shown that increasing the total mass flow rate through the combustor, therefore increasing the mean operating pressure, results in a decrease in both detonation wave velocities and detonation cycle stagnation pressure ratios.</p><p dir="ltr">Numerical modeling of detonations was conducted to understand the coupling of stagnation pressure ratios and wave speeds to deflagrative modes of combustion within rotating detonation combustors. Using the experimental measurements, it is shown that significant amounts of propellant combusts as a result of deflagration prior to (i.e., preburning) and after (i.e., afterburning) the detonation wave. Increasing the RDC operating pressure by 4x is shown to increase the amount of preburned propellant by 4.5x. Relevant chemical kinetic reaction rates of the conditions tested are modeled to increase by 4.5x as well, indicating that the increase in reactant preburning is the result of faster chemical kinetic timescales associated with higher pressure combustion. Results from this testing suggest an operating pressure upper limit for this combustor exists around 20 bar. At these conditions, chemical kinetic rates would be fast enough that deflagration would be the primary mode of combustion and the detonation would not exist. It is suggested that different injector or combustor designs might be able to extend operating limits, however it is unclear if there is a chemical kinetic limit at which no design would be able to overcome.</p><p dir="ltr">Despite significant amounts of deflagrative combustion within the RDC, the vacuum specific impulse produced by the RDC was shown to be between 95.0% and 98.5% of what an ideal deflagrative combustor could produce for most conditions. Given conventional rocket combustors typically operate at specific impulse efficiencies in the range of 90%-99%, it is noted that the RDC tested in this work has demonstrated, at the very least, equal performance to the current state of the art for rocket propulsion combustors while utilizing an effective combustor length (L*) of only 63 mm (2.5 inches). A detailed RDC performance model was developed which considered losses associated with deflagration (both preburning and afterburning) and incomplete combustion. Using measurements obtained from the experiment it is determined that incomplete combustion contributes a larger performance loss than the deflagration which occurs within the combustor.</p><p dir="ltr">A total of 17 parametric studies were conducted experimentally to evaluate the response of the RDC specifically to changes in the propellant chemical reaction timescales. Detonation wave arrival times ranged between 10 microseconds and 178 microseconds as a result of testing at ranges of operating pressures, equivalence ratios, and utilizing nine unique propellant combinations. It was shown that the wave arrival time is primarily a function of chemical kinetic timescales and injector mixing processes. A model using the injector momentum ratio and modeled deflagrative preheat times is shown to be able to closely predict experimentally obtained detonation wave arrival times.</p>

Recent improvements to concentrating solar dish systems lead to further focus on hybridization systems for small-scale power generation applications. Variability of the solar load creates new requirements for combustion systems. This thesis presents a CFD simulation of the air flow inside a new combustor design for the combination of an impinging air solar receiver and a MGT. The system consists of a LPP tubular combustor with radial main swirler and central pilot burner. Focus is made on the pressure loss at the downstream impinging cooling wall for appropriate flow distribution between reacting and bypass air. Heat transfer is not studied due to lack of time. A fully-hexahedral multi-zones mesh of the system without fuel injection has been generated with Ansys ICEM software, making use of its O-grid capabilities. A realizable k-epsilon model is used for turbulence modelling. Several impinging hole’s diameters are studied to find the right balance between the two streams. Streamlines are also observed to confirm the location of recirculation zones and recommend design improvements. / Nya förbättringar av koncentrerade solskålssystem leder till ytterligare fokus på hybridsystem för småskaliga applikationer för elproduktion. Ojämn solstrålning skapar nya krav på förbränningssystem. Detta examensarbete presenterar en CFD-simulering av luftflödet i en ny förbränningsdesign för en kombination av en solfångare med forcerad konvektionskylning och en mikrogasturbin (MGT). Systemet består av en LPP-rörbrännare med radiellt virvelsystem och central pilotbrännare. Studien fokuserar på tryckförlusten vid slaghålsväggen, som används för kylning vid förbränning, och lämplig flödesfördelning mellan reagerande- och förbigående flöde. Värmeöverföring studeras inte på grund av tidsbrist. Ett helt sexkantigt nät med flera zoner i systemet utan bränsleinsprutning har genererats med Ansys ICEM-programvara som använder O-nätfunktioner. En realiserbar k-epsilon-modell används för turbulensmodellering. Flera slaghålsdiametrar studeras för att hitta rätt balans mellan de två strömmarna. Även strömlinjer observeras för att bekräfta placeringen av återcirkulationszoner och kunna rekommendera förbättringar av designen.

Combustion designers have considered back-side impingement cooling as the solution for modern DLE combustors. The idea is to provide more cooling to the deserved local hot spots and reserve unnecessary coolant air from local cold spots. Therefore, if accurate heat load distribution on the liners can be obtained, then an intelligent cooling system can be designed to focus more on the localized hot spots. The goal of this study is to determine the heat transfer and pressure distribution inside a typical can-annular gas turbine combustor. This is one of the first efforts in the public domain to investigate the convective heat load to combustor liner due to swirling flow generated by swirler nozzles. An experimental combustor test model was designed and fitted with a swirler nozzle provided by Solar Turbines Inc. Heat transfer and pressure distribution measurements were carried out along the combustor wall to determine the thermo-fluid dynamic effects inside a combustor. The temperature and heat transfer profile along the length of the combustor liner were determined and a heat transfer peak region was established. Constant-heat-flux boundary condition was established using two identical surface heaters, and the Infrared Thermal Imaging system was used to capture the real-time steady-state temperature distribution at the combustor liner wall. Analysis on the flow characteristics was also performed to compare the pressure distributions with the heat transfer results. The experiment was conducted at two different Reynolds numbers (Re 50,000 and Re 80,000), to investigate the effect of Reynolds Number on the heat transfer peak locations and pressure distributions. The results reveal that the heat transfer peak regions at both the Reynolds numbers occur at approximately the same location. The results from this study on a broader scale will help in understanding and predicting swirling flow effects on the local convective heat load to the combustor liner, thereby enabling the combustion engineer to design more effective cooling systems to improve combustor durability and performance. / Master of Science

A comparison study between axial and radial swirler performance in a gas turbine can combustor was conducted by investigating the correlation between combustor flow field geometry and convective heat transfer at cold flow conditions for Reynolds numbers of 50,000 and 80,000. Flow velocities were measured using Particle Image Velocimetry (PIV) along the center axial plane and radial cross sections of the flow. It was observed that both swirlers produced a strong rotating flow with a reverse flow core. The axial swirler induced larger recirculation zones at both the backside wall and the central area as the flow exits the swirler, and created a much more uniform rotational velocity distribution. The radial swirler however, produced greater rotational velocity as well as a thicker and higher velocity reverse flow core. Wall heat transfer and temperature measurements were also taken. Peak heat transfer regions directly correspond to the location of the flow as it exits each swirler and impinges on the combustor liner wall. Convective heat transfer was also measured along the liner wall of a gas turbine annular combustor fitted with radial swirlers for Reynolds numbers 210000, 420000, and 840000. The impingement location of the flow exiting from the radial swirler resulted in peak heat transfer regions along the concave wall of the annular combustor. The convex side showed peak heat transfer regions above and below the impingement area. This behavior is due to the recirculation zones caused by the interaction between the swirlers inside the annulus. / Master of Science

RANS models in CFD are used to predict the liner wall heat transfer characteristics of a gas turbine annular combustor with radial swirlers, over a Reynolds number range from 50,000 to 840,000. A three dimensional hybrid mesh of around twenty five million cells is created for a periodic section of an annular combustor with a single radial swirler. Different turbulence models are tested and it is found that the RNG k-e model with swirl correction gives the best comparisons with experiments. The Swirl number is shown to be an important factor in the behavior of the resulting flow field. The swirl flow entering the combustor expands and impinges on the combustor walls, resulting in a peak in heat transfer coefficient. The peak Nusselt number is found to be quite insensitive to the Reynolds number only increasing from 1850 at Re=50,000 to 2200 at Re=840,000, indicating a strong dependence on the Swirl number which remains constant at 0.8 on entry to the combustor. Thus the peak augmentation ratio calculated with respect to a turbulent pipe flow decreases with Reynolds number. As the Reynolds number increases from 50,000 to 840,000, not only does the peak augmentation ratio decrease but it also diffuses out, such that at Re=840,000, the augmentation profiles at the combustor walls are quite uniform once the swirl flow impinges on the walls. It is surmised with some evidence that as the Reynolds number increases, a high tangential velocity persists in the vicinity of the combustor walls downstream of impingement, maintaining a near constant value of the heat transfer coefficient. The computed and experimental heat transfer augmentation ratios at low Reynolds numbers are within 30-40% of each other. / Master of Science