General Electric Company

Ge Infrastructure Sensing, Inc.

Hejie Li

Michael Anthony Benjamin

Brandon Wayne Miller

Shawn David Wehe

Joel Meier Haynes

Stefan Joseph Cafaro

Eric Richard Westervelt

Yufeng Huang

Robert Jon McQuiston

Anthony Kowal

John McKinley Poole

Xiaoyong Liu

Gene Smith Berkowitz

Shai Birmaher

Michael Robert Millhaem

Anquan Wang

Gerardo Antonio Salazar Lois

Gerardo Perez

Andrew Campbell

Alejandro Federico Medrano Perez

Francisco Ernesto De La Pena

Christopher R Bosco

Clayton Stuart Cooper

Lawrence Bernard Kool

Hiranya Kumar Nath

Guanghua Wang

Samhita Dasgupta

Craig Alan Gonyou

Nirm Velumylum Nirmalan

Nicholas Ryan Overman

Mark Alan Rosenzweig

Karthick Vilapakkam Gourishankar

Brian Lewis Devendorf

Keith Robert McManus

Edward Randall Furlong

Arundhati Sengupta

Inenhe Mohammed Khalid

Daniel Roy Kiracofe

Narayanan Janakiraman

Keith Franklin Beatty

Sanjay Kumar Sondhi

Mathew Paul Thariyan

Jason Scott Dangel

Diana Julieta Méndez Maya

Habeeb Kunnummal Manat

Michael J Gambuzza

Robert Arthur Clark

Daniel M Stearns

Stephen Thomas Walls

Rodrick Gordon Brown

K'Ehleyr James

Jorge Camacho

Joshua Friedman

Jorge Omar Mañón Cantú

Kehleyr James

Keith F Beatty

A combustor for a gas turbine engine includes a dome-deflector structure and a fuel nozzle-swirler assembly having a mounting wall with a plurality of purge orifices extending therethrough. A circumferential purge cavity is defined between a fuel nozzle-swirler assembly housing, a fuel nozzle-swirler assembly opening of the dome-deflector structure, and the mounting wall. The purge orifices provide a purge airflow to the circumferential purge cavity. In a first state, when a radial height of the circumferential purge cavity is constant, the dome-deflector structure overlaps a portion of the purge orifices to block a portion of each purge orifice, and, in a second state when the fuel nozzle-swirler assembly is radially shifted, the dome-deflector structure increases blockage of at least one purge orifice on a second side of the circumferential purge cavity, and reduces blockage of at least one purge orifice on a first side of the circumferential purge cavity.

Combustor for a gas turbine engine

A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system can be controlled to adjust one or more parameters of the fuel based on data received from the sensors.

Methods of control for management of hot fuel

A combustor of a turbine engine combusts fuel and air. The combustor includes a dome and a heat shield fastened to the dome. The heat shield is unconstrained relative to dome, to permit relative thermal expansion and contraction in a radial direction. A heat shield inner portion is generally annular in shape, is arranged about a dome eyelet longitudinal centerline axis, is smaller in diameter than a dome inner portion, and is assembled inside of the dome inner portion. A plurality of pinned assemblies fastens the heat shield to dome. Each pinned assembly includes a fastener with a pin end, the pin end assembled in a clearance fit into a hole in the heat shield inner portion. The fastener is fixed to the dome inner portion, the pin end sliding in the hole in the heat shield during relative thermal expansion and contraction between the dome and the heat shield.

Combustor including a heat shield for a turbine engine

A combustor of a turbine engine has a combustion chamber for combusting fuel and air, generating heat. A fuel nozzle-mixer assembly mixes and injects fuel and air into the combustion chamber for combustion. The assembly has a heat shield at an aft end for shielding the assembly from the heat. A centerbody outer shell defines a cooling air flow cavity. A cooling air flow gap is defined between the centerbody outer shell and the heat shield. The centerbody outer shell has cooling air exit holes between the cooling air flow cavity and the cooling air flow gap, the air passing through the cooling air exit holes and into the cooling air flow gap. The cooling air exits the cooling air flow gap at a cooling air exit angle, the cooling air exit angle having a non-zero axially aft component, displacing the combustion downstream, to protect the assembly.

Turbine engine combustor including a heat shield

A gas turbine engine including a combustor, a plurality of fuel nozzles, and at least one fuel manifold. The combustor includes a combustion chamber. The plurality of fuel injects fuel into the combustion chamber of the combustor. The gas turbine engine may include a first fuel circuit and a second fuel circuit. The first fuel circuit includes a first fuel manifold fluidly connected to at least one fuel nozzle of the plurality of fuel nozzles to distribute the fuel to the at least one fuel nozzle at a first temperature. The second fuel circuit includes a second fuel manifold fluidly connected to at least one fuel nozzle of the plurality of fuel nozzles to distribute the fuel to the at least one fuel nozzle at a second. The second temperature is less than the first temperature.

Multi-temperature fuel injectors for a gas turbine engine

A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system adjusts parameters of the fuel based on data received from the sensors.

A method is provided for operating a gas turbine engine. The method includes: determining data indicative of an operation of a cooling system of the gas turbine engine during a shutdown of the gas turbine engine, following the shutdown of the gas turbine engine, or both; and modifying a startup schedule of the gas turbine engine in response to the determined data indicative of the operation of the cooling system of the gas turbine engine.

Gas turbine engine cooling system control

A fuel temperature control system and a method of controlling a fuel temperature for a turbine engine including supplying an aviation fuel to a fuel nozzle fluidly coupled to a combustion chamber. A fuel temperature sensor for determining at least one input parameter to define an inlet fuel temperature of the aviation fuel in the fuel nozzle. A controller for receiving the at least one input parameter and for calculating a calculated flow number in the fuel nozzle. The controller capable of comparing the calculated flow number and a reference flow number associated with a threshold during a steady state condition to determine if the aviation fuel is boiling inside the fuel nozzle.

Fuel temperature control system

A fuel system including a fuel metering unit, a heat exchanger, a plurality of fuel nozzles, and a coke filtration system. The fuel metering unit is configured to meter a flow of a hydrocarbon fuel. The heat exchanger is fluidly connected to the fuel metering unit. The heat exchanger is configured to heat the hydrocarbon fuel to improve engine performance or efficiency. The plurality of fuel nozzles is fluidly connected to the heat exchanger downstream of the heat exchanger to receive the hydrocarbon fuel heated by the heat exchanger. The coke filtration system includes at least one filter to remove coke particles in the heated fuel. The coke filtration system is located upstream of the plurality of fuel nozzles and directly downstream of the heat exchanger.

Fuel system having a coke filtration system

A component and a system for mitigating coke formation during delivery of a hydrocarbon fluid. The component includes a contact surface configured to be in contact with the hydrocarbon fluid. Tuning the zeta potential of the contact surface allows selective attraction and/or repulsion of coke-catalyzing materials, metal ions, heteroatomic hydrocarbons, and/or coke precursors present in the hydrocarbon fluid. A method of mitigating coke formation during delivery of a hydrocarbon fluid includes tuning a zeta potential of the contact surface of the component and injecting or circulating the hydrocarbon fluid through the system such that the contact surface selectively attracts and/or repels coke-catalyzing materials, metal ions, heteroatomic hydrocarbons, and/or coke precursors present in the hydrocarbon fluid.

Tuning the zeta potential of surfaces for coke mitigation in fuel and oil systems

A gas turbine engine includes a core cowl and a core contained within the core cowl. The core includes a compressor in fluid communication with a downstream combustor and a downstream turbine, the compressor including a compressor bleed port, wherein an undercowl space is defined between the core cowl and the core. The gas turbine engine further includes a cooling duct disposed at least partially in the undercowl space and having an inlet and an outlet, wherein the cooling duct is in fluid communication with a source of cooling air and is further in fluid communication with the compressor bleed port; a valve assembly including at least one valve disposed in the cooling duct; and a cooling blower disposed within the engine and operable to move an air flow from the inlet of the cooling duct towards the outlet of the cooling duct and through the compressor bleed port.

A gas turbine engine includes a liner positioned within a compressor section or a turbine section of the gas turbine engine and at least partially defining a core air flowpath through the gas turbine engine. The gas turbine engine also includes a casing at least partially enclosing the liner. Additionally, the gas turbine engine includes a pressure sensor assembly having a body, an extension member, and a pressure sensor. The pressure sensor is positioned at least partially within the body and the body is positioned at least partially on an outer side of the casing, the extension member extending from the body through a casing opening in the casing and towards a liner opening in the liner. The extension member defines a continuous sense cavity exposing the pressure sensor to the core air flowpath.

Pressure sensor assembly for a turbine engine

Systems and methods for indirect monitoring of combustor dynamics in a gas turbine engine include collecting vibration data acquired by a vibration sensor, which is mounted proximate to an operational component positioned relative to a combustor of a gas turbine engine. The vibration data can be transformed into a frequency domain representation at periodic intervals. The relative signal strength of the vibration data can be determined over the one or more identified frequency bands. The relative signal strength can be adjusted with hysteresis at each of the one or more identified frequency bands. Occurrence of a combustor dynamics event at one or more specific resonant frequencies can be determined based at least in part on evaluation of the adjusted relative signal strength relative to one or more event detection threshold levels.

Indirect monitoring of aircraft combustor dynamics

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Schenectady, NY, United States of America

Hejie Li