Abstract: A pilot fuel nozzle assembly includes a fuel nozzle, a swirler, and a vented pilot venturi. The vented pilot venturi has an annular wall with an oxidizer flow passage therein. An expansion flow surface portion of the venturi has a larger diameter at an outlet than at a throat of the venturi. A plurality of venturi oxidizer outlet ports extend through the expansion flow surface to the oxidizer flow passage within the annular wall to provide a flow of oxidizer through the venturi wall into a mixing cavity of the venturi and at an outlet end of the venturi. The oxidizer outlet ports are circumferentially spaced about a circumference of the expansion flow surface, and may be arranged in a plurality of rows. The oxidizer outlet ports may be angled with respect to the expansion flow surface and may angled circumferentially in a co-swirl direction with the swirler.

Abstract: A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.

Abstract: A turbine engine includes a compression section, combustion section, and turbine section in serial flow arrangement and defining an engine centerline. The combustion section has a combustor including a combustor liner at least partially defining a combustion chamber exhausting at a combustion chamber outlet, with the combustor liner having an inner liner and an outer liner. A liquid fuel injector configured to inject a first mixture comprised of liquid fuel and air or atomized liquid fuel and air into the combustion chamber and a gaseous fuel injector is configured to inject a second mixture comprised of gaseous fuel and air or atomized fuel and air into the combustion chamber.

Abstract: A combustion section of a turbine engine having a primary combustion chamber and a secondary combustion chamber. The primary combustion chamber is defined, at least in part, by a combustor liner and a dome wall. A set of primary fuel nozzles have outlets located at the dome wall, wherein the outlets of the set of primary fuel nozzles are fluidly coupled with the primary combustion chamber. At least one opening located in an outer liner axially aft of the dome wall fluidly couples the secondary combustion chamber and the primary combustion chamber. A set of secondary fuel nozzles are fluidly coupled with the secondary combustion chamber.

Abstract: A turbine engine has a compressor section, a combustion section, and a turbine section in serial flow arrangement. The combustion section has a combustion liner and dome wall collectively forming at least a portion of a combustion chamber. The dome wall has a fuel nozzle opening. The combustion section has a fuel nozzle assembly extending through the fuel nozzle opening. The fuel nozzle assembly has a gaseous fuel supply, a compressed air supply, a mixer, a fuel nozzle and a first swirler.

Abstract: An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. Increasing efficiency and meeting emission needs can be met with the use of alternative fuels, which combust at higher temperatures or higher speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.

Abstract: A gas turbine engine, comprising: a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a fuel nozzle assembly, comprising: a liquid fuel supply to supply a liquid fuel; a hydrogen fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle body defining a fuel nozzle assembly centerline, the fuel nozzle body fluidly coupled with the liquid fuel supply and the hydrogen fuel supply to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber.

Abstract: A turbine engine can combust fuel to drive a turbine, which drives the engine. A fuel nozzle can supply fuel for combustion or ignition of the fuel. The fuel nozzle can include a central body defining a first fuel passage terminating at a first orifice. An annular first wall circumferentially surrounds the central body and is spaced therefrom to define a first airflow passage therebetween. A set of first swirler vanes is circumferentially spaced about the central body and disposed within the first airflow passage. The fuel nozzle can supply a mixture of fuel and air for combustion.

Abstract: A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; a wall coupled to the combustor liner; a first fuel supply to supply a first fuel; a gaseous fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle assembly coupled to the wall and fluidly coupled to the first fuel supply and the gaseous hydrogen fuel supply, the fuel nozzle assembly comprising: a main mixer; and a fuel nozzle comprising an outer wall defining a pilot channel and an outer wall fuel orifice to emit at least one of the first fuel or the gaseous fuel radially outward into the main mixer; and a secondary mixer disposed in the pilot channel.

Abstract: A fuel injector for a gas turbine engine includes a compression section, combustion section, and turbine section in serial flow arrangement. The fuel injector arranged in the combustion section and includes an inner nozzle defining a fuel injector axis and an outer nozzle in annular arrangement about the inner nozzle. The outer nozzle at least partially defines a mixing region. An inner air passage and an outer air passage exhaust to the mixing region. A fuel body positions between the inner air passage and the outer air passage and has an outer fuel passage exhausting to the mixing region.

Abstract: A gas turbine engine, comprising: a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a fuel nozzle assembly, comprising: a liquid fuel supply to supply a liquid fuel; a hydrogen fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle body fluidly coupled with the liquid fuel supply and the hydrogen fuel supply to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber.

Abstract: A turbine engine and method of operation with a combustor as shown and described. The turbine engine including a combustor comprising an annular array of rich cups and an annular array of lean cups. A method for controlling nitrogen oxides including injecting a fuel/air mixture from the lean cups and the rich cups.

Abstract: A combustor comprising a dome wall, an annular liner, a combustion chamber, a set of fuel cups, and a set of dilution passages for each fuel cup of the set of fuel cups. The set of fuel cups circumferentially spaced along the dome wall relative to the combustor centerline. The set of dilution passages terminating in a plurality of slots spaced about the corresponding fuel cup in the set of fuel cups.

Abstract: A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; a wall coupled to the combustor liner; a first fuel supply to supply a first fuel; a gaseous fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle assembly coupled to the wall and fluidly coupled to the first fuel supply and the gaseous hydrogen fuel supply, the fuel nozzle assembly comprising: a main mixer; and a fuel nozzle comprising an outer wall defining a pilot channel and an outer wall fuel orifice to emit at least one of the first fuel or the gaseous fuel radially outward into the main mixer; and a secondary mixer disposed in the pilot channel.

Abstract: A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.

Abstract: A method of operating a gas turbine engine includes controlling operating of the gas turbine engine through a first portion of a total operating period utilizing first data of a first control table to control an operational aspect related to an engine operating condition of an estimated performance margin table, selecting one of a second control table through an Nth control table for operating the gas turbine engine through a second portion of the total operating period of the gas turbine engine, and controlling operating of the gas turbine engine through the second portion of the total operating period utilizing the second through N data of the respective control table selected in the selecting to control the operational aspect related to the engine operating condition of the estimated performance margin table.

Abstract: An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion, which can supply a primary fuel supply and a secondary fuel supply. Increasing efficiency and reducing emission require the use of alternative fuels, which combust at higher temperatures or burn at faster burn speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.

Abstract: A turbine engine has a compressor section, a combustion section, and a turbine section in serial flow arrangement. The combustion section has a combustor liner and dome wall collectively forming at least a portion of a combustion chamber. The dome wall has a fuel nozzle opening. The combustion section has a fuel nozzle assembly extending through the fuel nozzle opening. The fuel nozzle assembly has a first body, a second body, a first swirler and a second swirler.

Abstract: A turbine engine includes a compression section, combustion section, and turbine section is serial flow arrangement. A supply of gaseous fuel is mixed with a supply of inert gas to form a mixture of fuel and inert gas. The mixture of fuel and inert gas is provided to or intermixed within a fuel injector for combustion within a combustor provided within the combustor section to drive the turbine section.

Abstract: A turbine engine with a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. A combustion section for the turbine engine, having a primary combustor liner including an inner liner and an outer liner annular about an engine centerline. A dome wall extending between the inner liner and the outer liner. A set of primary dome inlets located in the dome wall and circumferentially arranged about the engine centerline. A set of secondary combustors fluidly coupled to a primary combustion chamber, the set of secondary combustors including a first mini combustor and a second mini combustor.