Abstract: A premixer for a gas turbine combustor includes a centerbody, a swirler assembly, and a mixing duct. The swirler assembly includes an inner swirler with vanes that rotate air in a first direction and an outer swirler with vanes that rotate air in an opposite direction. The inner swirler vanes and the outer swirler vanes are separated by an annular splitter. The outer swirler vanes define an outlet plane, and the inner swirler vanes each have a trailing edge that is disposed at an acute angle relative to the outlet plane. In one aspect, the inner swirler is axially offset from the outer swirler. The mixing duct may also define fuel passages that deliver fuel to fuel outlets on the downstream end of the mixing duct. The premixer is designed for operation on gaseous fuel or liquid fuel.

Abstract: A premixer for a gas turbine combustor includes a centerbody, a swirler assembly, and a mixing duct. The swirler assembly includes an inner swirler with vanes that rotate air in a first direction and an outer swirler with vanes that rotate air in an opposite direction. The inner swirler vanes and the outer swirler vanes are separated by an annular splitter. The outer swirler vanes define an outlet plane, and the inner swirler vanes each have a trailing edge that is disposed at an acute angle relative to the outlet plane. In one aspect, the inner swirler is axially offset from the outer swirler. The mixing duct may also define fuel passages that deliver fuel to fuel outlets on the downstream end of the mixing duct. The premixer is designed for operation on gaseous fuel or liquid fuel.

Abstract: The present disclosure relates to a fuel-air premixer for a turbine system. The fuel-air premixer includes a swirler and a centerbody. The swirler is configured to direct a flow of air through the premixer, and the centerbody is configured to inject fuel into the flow of air. Additionally, the centerbody includes an airfoil shape that reduces and/or substantially eliminates recirculation pockets to prevent autoignition and/or flame holding in a combustion chamber. Accordingly, the turbine system may produce fewer NOx emissions.

Abstract: A combustor is configured to operate in a rotating detonation mode and a deflagration mode. The combustor includes a housing and at least one initiator. The housing defines at least one combustion chamber and is configured for a deflagration process to occur within the at least one combustion chamber during operation in the deflagration mode and a rotating detonation process to occur within the at least one combustion chamber during operation in the rotating detonation mode. The at least one initiator is configured to initiate the rotating detonation process within the at least one combustion chamber during operation in the rotating detonation mode and to initiate the deflagration process within the at least one combustion chamber during operation in the deflagration mode.

Abstract: A combustor is configured to operate in a rotating detonation mode and a deflagration mode. The combustor includes a housing and at least one initiator. The housing defines at least one combustion chamber and is configured for a deflagration process to occur within the at least one combustion chamber during operation in the deflagration mode and a rotating detonation process to occur within the at least one combustion chamber during operation in the rotating detonation mode. The at least one initiator is configured to initiate the rotating detonation process within the at least one combustion chamber during operation in the rotating detonation mode and to initiate the deflagration process within the at least one combustion chamber during operation in the deflagration mode.

Abstract: A turbine engine includes at least one compressor configured to increase pressure of a fluid flow and a primary combustor coupled in flow communication with the at least one compressor. The primary combustor is configured to receive pressurized fluid flow from the at least one compressor. The primary combustor includes a housing defining at least one combustion chamber. The primary combustor is configured for a rotating detonation process to occur within said at least one combustion chamber. The turbine engine also includes at least one supplemental combustor coupled in flow communication with the primary combustor. The at least one supplemental combustor is configured to receive combustion products and perform a combustion operation. The turbine engine further includes a turbine assembly coupled in flow communication with the at least one supplemental combustor. The turbine assembly is configured to receive combustion products from the at least one supplemental combustor.

Abstract: The present disclosure relates to a fuel-air premixer for a turbine system. The fuel-air premixer includes a swirler and a centerbody. The swirler is configured to direct a flow of air through the premixer, and the centerbody is configured to inject fuel into the flow of air. Additionally, the centerbody includes an airfoil shape that reduces and/or substantially eliminates recirculation pockets to prevent autoignition and/or flame holding in a combustion chamber. Accordingly, the turbine system may produce fewer NOx emissions.

Abstract: A system for premixing fuel and air prior to combustion in a gas turbine engine includes a mixing duct, a centerbody fuel injector located along a central axis of the mixing duct an outer annular swirler located adjacent an upstream end of the mixing duct for swirling air flowing therethrough in a first swirl direction and an inner annular swirler located adjacent of the mixing duct upstream end for swirling air flowing therethrough in a second swirl direction. The system includes a hub separating the inner and outer annular swirlers to permit independent rotation of an air stream therethrough and multiple hollow paths located radially outward around the centerbody fuel injector and at a radially inward side of the inner annular swirler for allowing a flow of sweeping air over the surface of the centerbody fuel injector.

Abstract: In one embodiment, a system for reducing pressure oscillations within a gas turbine engine includes at least one fuel injector configured to inject fuel into a combustor. The system also includes a valve fluidly coupled to the at least one fuel injector. The system further includes a controller communicatively coupled to the valve. The controller is configured to cycle the valve between an open position and a closed position at a first frequency and a first duty cycle while a magnitude of pressure oscillations within the combustor is less than a threshold value, to cycle the valve between the open position and the closed position at a second frequency and a second duty cycle while the magnitude of the pressure oscillations within the combustor is greater than or equal to the threshold value, and to adjust the second frequency based on a measured frequency of the pressure oscillations.

Abstract: A system for premixing fuel and air prior to combustion in a gas turbine engine includes a mixing duct, a centerbody fuel injector located along a central axis of the mixing duct, an outer annular swirler located adjacent an upstream end of the mixing duct for swirling air flowing therethrough in a first swirl direction and an inner annular swirler located adjacent of the mixing duct upstream end for swirling air flowing therethrough in a second swirl direction. The system includes a hub separating said inner and outer annular swirlers to permit independent rotation of an air stream therethrough and multiple hollow paths located radially outward around the centerbody fuel injector and at a radially inward side of the inner annular swirler for allowing a flow of sweeping air over the surface of the centerbody fuel injector.

Abstract: A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system comprises at least one laser light source configured to generate a laser beam and an optical beam guidance component. The optical beam guidance component is configured to transmit the laser beam to irradiate on an oxygenated fuel mixture supplied into the combustion chamber at a region of highest ignitability to generate a combustor flame in a flame region. The system further includes an integrated control diagnostic component configured to detect at least a portion of a light emission and operable to control one or more combustion parameters based in part on the detected light emission. The system further includes additional enhanced ignition control configurations. A method for igniting a fuel mixture in an internal combustion engine is also presented.

Abstract: In one embodiment, a system for reducing pressure oscillations within a gas turbine engine includes at least one fuel injector configured to inject fuel into a combustor. The system also includes a valve fluidly coupled to the at least one fuel injector. The system further includes a controller communicatively coupled to the valve. The controller is configured to cycle the valve between an open position and a closed position at a first frequency and a first duty cycle while a magnitude of pressure oscillations within the combustor is less than a threshold value, to cycle the valve between the open position and the closed position at a second frequency and a second duty cycle while the magnitude of the pressure oscillations within the combustor is greater than or equal to the threshold value, and to adjust the second frequency based on a measured frequency of the pressure oscillations.

Abstract: A real-time monitoring of an equivalence ratio of a gas-fuel mixture of a gas turbine engine is provided. The system includes multiple optical probes arranged on a plurality of fuel nozzles for transmitting laser beams directly through a gas-fuel mixture or indirectly by reflecting the laser beams from a surface of a centerbody or burner tube of the fuel nozzle. The system also includes one or more detectors to measure the transmitted laser beams from the multiple optical probes. Further, the system includes a data acquisition subsystem for acquiring and processing signals from the one or more detectors to determine the equivalence ratio of the gas-fuel mixture of the nozzle.

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 combustion nozzle includes at least one passage having a mixing section and an exit section. The mixing section includes an air inlet, and a fuel inlet. The mixing section has a first length and a first diameter. The exit section has a second length different from the first length, and a second diameter different from the first diameter.

Abstract: Certain embodiments of the invention may include systems and methods for providing optical interrogation sensors for combustion control. According to an example embodiment of the invention, a method for controlling combustion parameters associated with a gas turbine combustor is provided. The method can include providing an optical path through the gas turbine combustor, propagating light along the optical path, measuring absorption of the light within the gas turbine combustor, and controlling at least one of the combustion parameters based at least in part on the measured absorption.

Abstract: A real-time monitoring of an equivalence ratio of a gas-fuel mixture of a gas turbine engine is provided. The system includes multiple optical probes arranged on a plurality of fuel nozzles for transmitting laser beams directly through a gas-fuel mixture or indirectly by reflecting the laser beams from a surface of a centerbody or burner tube of the fuel nozzle. The system also includes one or more detectors to measure the transmitted laser beams from the multiple optical probes. Further, the system includes a data acquisition subsystem for acquiring and processing signals from the one or more detectors to determine the equivalence ratio of the gas-fuel mixture of the nozzle.

Abstract: A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system comprises at least one laser light source configured to generate a laser beam and an optical beam guidance component. The optical beam guidance component is configured to transmit the laser beam to irradiate on an oxygenated fuel mixture supplied into the combustion chamber at a region of highest ignitability to generate a combustor flame in a flame region. The system further includes an integrated control diagnostic component configured to detect at least a portion of a light emission and operable to control one or more combustion parameters based in part on the detected light emission. The system further includes additional enhanced ignition control configurations. A method for igniting a fuel mixture in an internal combustion engine is also presented.

Abstract: A system includes an optical sensor that optically measures and spatially resolves in three dimensions at least one chemical species within a flame produced by a device and a component that correlates the three dimensionally measured at least one chemical species to at least one parameter of the device.

Abstract: Certain embodiments of the invention may include systems and methods for providing optical sensors for combustion control. According to an example embodiment of the invention, a method for controlling combustion parameters associated with a gas turbine combustor is provided. The method can include providing at least one optical path adjacent to a flame region in the combustor, detecting at least a portion of the light emission from the flame region within the at least one optical path, and controlling at least one of the combustion parameters based in part on the detected light emission.