Future Navy Capabilities indicate the need for a supersonic cruise missile. Thus the need exists for a low cost, light-weight, and efficient means of supersonic propulsion. NPS has been developing the Pulse Detonation Engine, which in theory has a thermodynamic efficiency greater than 50% as compared to 35% for state of the art constant-pressure cycles currently in use in gas turbines/ramjets/scramjets. Nonetheless, there are two major problems in the development of this engine. These are the increase of the propulsive efficiency by removing the oxygen-assisted initiator currently in use, and the reduction of internal total pressure losses caused by the highly constrictive internal flow-path geometry currently required to promote the deflagration to detonation transition (DDT). The aforementioned problems have been addressed and a viable design proposed through the implementation of a novel Transient Plasma Ignition system and a split-flow path engine geometry as described in this work. Future work will concentrate on the development of a performance measurement test rig to experimentally assess the designs presented herein. / US Navy (USN) author.

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Approved for public release, distribution unlimited / An existing freejet facility was upgraded and its range of operation extended into the high subsonic regime for operation as a test rig for the development of a combined-cycle, turbo-ramjet engine. A combustor was designed, developed, and tested as the afterburner for the turbo-ramjet engine. At subsonic speeds with the afterburner running, an increase in thrust of 40% was measured over the baseline turbojet running at 80% spool speed. A Computational Fluid Dynamics model of the flow through the shrouded turbojet engine was developed and successfully used to assist in predicting the bypass ratio of the engine at different Mach numbers. Numerous recommendations were made to improve the operation of the test rig, to improve the performance of the turbo-ramjet engine, and refine the numerical models. These recommended improvements will extend the present capabilities to design and analyze small combined cycle engines which have an application in unmanned aerial vehicles and missiles. / Lieutenant, United States Navy

To provide quality benchmark data (that can be used in numerical simulation comparisons) and to examine the effects of combustion on a typical ramjet engine flowfield, a water-cooled, stainless steel dump combustor model was developed. A two-component Laser Doppler Anemometer (LOA) was used to measure the mean and turbulent velocities in the axial and tangential directions and provide a comparison between combusting and isothermal flows. However, before any LOA measurements could be taken, the combustor had to be configured to run in a suitably stable mode. Stability was identified by the pressure spectra obtained under various running conditions using piezoelectric pressure transducers wired to a spectrum analyzer. Operational parameters such as fuel composition, fuel injection location, acoustic configuration, and equivalence ratio were varied until instabilities were minimized. The optimal configuration ran with upstream fuel injection (premixed mode) at the duct center line and an orifice plate installed immediately upstream of the fuel injectors, with propane as the fuel. Once stability was achieved, LOA data was taken. The results showed some significant differences between the reacting and nonreacting flows. The most significant effect was the difference between the inherent recirculation regions for each case. Combustion decreased the length of the region by approximately 50 percent, while increasing the maximum negative velocities. This made for a more compact, but stronger, recirculation region. Since the recirculation region acts as the main flame holder and is a major source of turbulence, the changes in this region significantly altered the dump combustor flowfield. / Master of Science

Supersonic inlet starting by mass spillage is studied theoretically and numerically in the present thesis. A quasi-one-dimensional, quasi-steady theory is developed for the analysis of flow inside a perforated inlet. The theory results in closed-form relations applicable to flow starting by the mass spillage technique in supersonic and hypersonic inlets. / The theory involves three parameters to incorporate the multi-dimensional nature of mass spillage through a wall perforation. Mass spillage through an individual slot is studied to determine these parameters; analytical expressions for these parameters are derived for both subsonic and supersonic flow conditions. In the case of mass spillage from supersonic flows, the relations are exact. However, due to the complexity of flow field, the theory is an approximation for subsonic flows. Therefore, a correction factor is introduced which is determined from an empirical relation obtained from numerical simulations. / A methodology is also proposed to determine perforation size and distribution to achieve flow starting for a given inlet at a desired free-stream Mach number. The problem of shock stability inside a perforated inlet designed with the proposed method is also discussed. / The method is demonstrated for some test cases. Time-realistic CFD simulations and experimental results in the literature confirm the accuracy of the theory and the reliability of the proposed design methodology.

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Thesis -- University of Adelaide, 1966. / [Typescript].

Basic experimental studies of the injection of liquid fuel into a two dimensional flowfield designed to represent a sudden-expansion "dump" combustor were performed under cold-flow conditions. Test conditions were as follows: 0.6 entrance Mach number, 25 PSIA total pressure, and nominally 75°F stagnation temperature. Two step heights were investigated, 1.0 in. and 0.5 in., corresponding to area ratios of 1.33 and 1.17. The investigation included Pitot and static pressure distributions, spark and streak shadowgraphs, surface flow visualization, direct photographs and videotape recordings. The backlighted streak and spark shadowgraphs were used to obtain jet penetration and break-up information. Oil drop surface flow studies showed details of the flow in the recirculation region behind the step. The injectant for these cold flow studies was selected as water, which was injected transversely to the air flow 1.0 in. and 0.5 in. upstream of the step at various flow rates. It was found that both the location of the injection port relative to the step and the step height had no measurable effect on jet penetration and break-up. Injectant accumulation on the combustor wall in the base-flow region was found to be substantial under some conditions, and the amount of accumulation was shown to be a strong function of initial liquid jet penetration height. / M.S.