<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>