Abstract: It is desirable for a gas turbine system to operate in a wide range operating conditions. However, under certain conditions there exist dynamic boundaries that limit a combustor from reaching its designated condition. Perturbation devices formed of electromagnetic plates can be incorporated into fuel nozzles of the combustor to influence the dynamics so that the range of operating conditions can be widened. The perturbation devices vibrate according to the perturbation signals provided from a dynamics controller. The vibration characteristics of the perturbation devices can be controlled by controlling the attributes of the perturbation signals. The vibrations influence the dynamics of fluid—fuel, oxidant, or both—flowing past the perturbation devices within the fuel nozzles.

Abstract: The present application provides an optical probe system for use with a combustion flame in a combustion chamber. The optical probe system may include a number of optical probes fixedly attached about the combustion chamber and positioned such that the optical probes collect light generated by the combustion flame in a field of view of each of the optical probes. One or more components external to the combustion chamber may produce and analyze signals indicative of the light generated by the combustion flame in the field of view of each of the optical probes.

Abstract: A method for assembling a gas turbine engine combustor is provided. The method includes providing a heat shield defined by a perimeter. The perimeter includes a radially inner edge, a radially outer edge, an axially inner edge, an axially outer edge, and an opening that extends from an upstream side of the heat shield to a downstream side of the heat shield. The method further includes coupling the heat shield to a domeplate such that the perimeter of the heat shield is positioned a distance downstream from an edge of the heat shield defining the opening. The method additionally includes coupling at least one fuel injector to the domeplate such that a portion of the fuel injector extends through the heat shield opening.

Abstract: A method for operating a pulse detonation engine, wherein the method includes channeling air flow from a pulse detonation combustor into a flow mixer having an inlet portion, an outlet portion, and a body portion extending therebetween. The method also includes channeling ambient air past the flow mixer and mixing the air flow discharged from the pulse detonation combustor with the ambient air flow such that a combined flow is generated from the flow mixer that has less flow variations than the air flow discharged from the pulse detonation combustor.

Abstract: In the present invention, at least one detonation initiator is positioned downstream of a main combustion chamber, with the initiator oriented such that it projects a detonation initiation wave forward into the main combustion chamber. The main combustion chamber contains a wave reflection surface. The detonation initiation wave is directed into the main combustion chamber and at the wave reflector surface and is used to initiate, or assist in the initiation of, a fuel and gas mixture in the main combustion chamber. The fuel and gas mixture is detonated, creating a high temperature and high pressure wave that is directed out of the main combustion chamber.

Abstract: The present invention is a pulse detonation combustion system, having a plurality of detonation initiation devices coupled to a main combustion chamber, where each of the detonation initiation devices is operating out-of-phase with each other. Each of the detonation initiation devices assists in the initiation of a detonation in the main combustion chamber, out-of-phase from each other such that the operational frequency of the pulse detonation combustion system is related to the number of detonation initiation devices multiplied by the operational frequency of a single detonation initiation device.

Abstract: A fuel nozzle device suitable for use in a gas turbine engine or the like is provided. The fuel nozzle device includes a fuel line and a plurality of gas orifices disposed at a downstream end of the fuel line, the plurality of gas orifices operable for injecting fuel into an air stream. The acoustic resistance of each of the plurality of gas orifices is chosen to match the acoustic impedance of the fuel line such that the maximum acoustic energy may be transferred between the fuel nozzle device and the combustor, thus enhancing the ability of the fuel nozzle device to control the combustion dynamics of the gas turbine engine system. A fuel injection resonator assembly suitable for use in a gas turbine engine or the like is also provided. The fuel injection resonator assembly includes a plurality of orifices separated by a variable length tube.

Abstract: A fuel nozzle device suitable for use in a gas turbine engine or the like is provided. The fuel nozzle device includes a fuel line and a plurality of gas orifices disposed at a downstream end of the fuel line, the plurality of gas orifices operable for injecting fuel into an air stream. The acoustic resistance of each of the plurality of gas orifices is chosen to match the acoustic impedance of the fuel line such that the maximum acoustic energy may be transferred between the fuel nozzle device and the combustor, thus enhancing the ability of the fuel nozzle device to control the combustion dynamics of the gas turbine engine system. A fuel injection resonator assembly suitable for use in a gas turbine engine or the like is also provided. The fuel injection resonator assembly includes a plurality of orifices separated by a variable length tube.