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The formation of NOx in hydrogen-fuelled pulse detonation engines (PDE) is investigated numerically. The computations are based on the axisymmetric Euler equations and a detailed combustion model consisting of 12 species and 27 reactions. A multi-level, dynamically adaptive grid is utilized, in order to resolve the structure of the detonation front. Computed NO concentrations are in good agreement with experimental measurements obtained at two operating frequencies and two equivalence ratios. Additional computations examine the effects of equivalence ratio and residence time on NOx formation at ambient conditions. The results indicate that NOx formation in PDEs is very high for near stoichiometric mixtures. NOx reduction requires use of lean or rich mixtures and the shortest possible detonation tube. NOx emissions for very lean or very rich mixtures are, however, fairly insensitive to residence time.

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• E-mail your question via the Internet to help@ sti.nasa.gov• Fax your question to the NASA STI Help Desk at 301-621-0134• Telephone the NASA STI Help Desk at 301-621-0390• Prepared under Contract NNC05BAZZB and NNC08JF17T AcknowledgmentsThe authors wish to thank Lynn Arrington, Rick Gardins, Gerald Hurd, Kim Kearns, Anthony Ogorzaly, Jason Wendell, and Joseph Zoeckler in applying their considerable technical skills in preparing and operating the test facility for the experimental tests described in this paper. AbstractThe LOX/methane propellant combination is being considered for the Lunar Surface Access Module ascent main engine propulsion system. The proposed switch from the hypergolic propellants used in the Apollo lunar ascent engine to LOX/methane propellants requires the development of igniters capable of highly reliable performance in a lunar surface environment. An ignition test program was conducted that used an in-house designed LOX/methane spark torch igniter. The testing occurred in Cell 21 of the Research Combustion Laboratory to utilize its altitude capability to simulate a space vacuum environment. Approximately 750 ignition test were performed to evaluate the effects of methane purity, igniter body temperature, spark energy level and frequency, mixture ratio, flowrate, and igniter geometry on the ability to obtain successful ignitions. Ignitions were obtained down to an igniter body temperature of approximately 260 R with a 10 torr backpressure. The data obtained is also being used to anchor a CFD based igniter model. Nomenclature IntroductionThe LOX/methane propellant combination is being considered for the Lunar Surface Access Module ascent main engine propulsion system. The LOX/methane propellant combination is being considered for a variety of reasons. This propellant combination would replace toxic propellants that may pose an operational challenge near a lunar base. Use of this propellant combination for lunar ascent would provide operational experience relevant to its use on other planetary bodies where it may be produced from indigenous materials. The proposed switch from the hypergolic propellants used in the Apollo lunar ascent engine to LOX/methane propellants requires the development of igniters capable of highly reliable performance in a lunar surface environment. There has been a flurry of recent activities (ref. 1) to develop such igniters in industry, academia, and at government laboratories. An ignition test program was conducted at the NASA Glenn Research Center (GRC) that used an in-house designed LOX/methane spark torch igniter and the vacuum test facilities in Cell 21 of the Research Combustion Laboratory (RCL) at GRC. Igniter HardwareThe igniter was a three piece modular design consisting of a head end, a chamber section, and a fuel coolant sleeve as shown in figures 1, 2, and 3. At the top of the igniter head end, are the propellant inlets to which the valve offsets were attached. The valve offsets were small tubes attached to the valves at one end and threaded into the...