General Electric Company

Keith Robert McManus

Anthony John Dean

Kapil Kumar Singh

Lewis Berkley Davis, Jr

Venkat Eswarlu Tangirala

Joel Meier Haynes

Sravan Kumar Dheeraj Kapilavai

Mathew Paul Thariyan

Suhui Li

Narendra Digamber Joshi

Gregory Allen Boardman

Timothy James Held

Manampathy Gangadharan Giridharan

Adam Rasheed

Fei Han

Jun Xu

Hejie Li

Michael Louis Vermeersch

Andrew Maxwell Peter

Mark Anthony Mueller

Mark David Durbin

Jeffrey Scott Goldmeer

Edip Sevincer

John Thomas Herbon

Simon Ralph Sanderson

Shawn David Wehe

Owen Graham

Sara Rocci Denis

Venkatraman Ananthakrishman Iyer

Liangyu Wang

Zekai Hong

Herbert Kopacek

Michael Breit

Zhongtao Dai

Venkatraman Ananthakrishnanlyer

Lewis Berkley Davis Jr

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.

Premixer for low emissions gas turbine combustor

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 NOemissions.

Fuel-air premixer for a gas turbine

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.

Combustor for rotating detonation engine and method of operating same

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.

Air fuel premixer for low emissions gas turbine combustor

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.

System and method for reducing pressure oscillations within a gas turbine engine

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.

Advanced laser ignition systems for gas turbines including aircraft engines

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.

System and method for real-time measurement of equivalence ratio of gas fuel mixture

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.

Method and apparatus for assembling a gas turbine engine

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.

Optical interrogation sensors for combustion control

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.

Three-dimensional optical sensor and system for combustion sensing and control

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.

Methods and apparatus for operating a pulse detonation engine

An injection assembly for use with a combustor is provided. The injection assembly includes an effusion plate that has a plurality of plate openings and a plate sleeve having a sidewall portion that includes a forward edge. The forward edge is coupled to the effusion plate such that the effusion plate is oriented obliquely with respect to a centerline extending through the combustor. The injection assembly also includes a plurality of ring extensions where each of the ring extensions is coupled to one of the plurality of plate openings. Each ring extension extends rearwardly into the plate sleeve.

Injection assembly for a combustor

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.

Multiple detonation initiator for frequency multiplied pulsed detonation combustion

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. The area ratio of the plurality of orifices may be adjusted using an automated valve system or the like to modify and/or control the relative flow resistance of the plurality of orifices. The resulting fuel injection resonator assembly acts as a tunable acoustic waveguide operable for delivering fuel to the combustor.

Growing community of inventors

Clifton Park, NY, United States of America

Keith Robert McManus