A combustion system includes an electrically actuated flame location control mechanism.

A combustor assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a combustion chamber, and a fuel injector assembly in communication with the combustion chamber that has a swirler body situated about a nozzle to define an injector passage that converges to a throat. The throat is defined at a distance from the combustion chamber. The nozzle includes a primary fuel injector along a first fuel injector axis and at least one secondary plain jet fuel injector axially forward of the primary fuel injector.

A combustion system includes an electrically actuated flame location control mechanism.

An airblast fuel injector for a gas turbine engine fuel spray nozzle has, in order from radially inner to outer, a coaxial arrangement of an inner air swirler passage, an annular fuel passage, an annular outer air swirler passage, and an annular shroud air swirler passage. The injector further has an annular shroud having an inner surface profile. Relative to the overall axial direction of flow through the injector the shroud inner surface profile has a convergent section followed by a divergent section, the transition of which forming a first inwardly directed annular nose. The injector further has an annular wall having an outer surface profile, and having an inner surface profile. Relative to the overall axial direction of flow through the injector the wall outer surface profile has a convergent section followed by an outwardly turning section which faces across the shroud air passage to the first nose.

A fuel delivery system for a combustor of a gas turbine engine including a primary fuel tank configured to store a primary fuel, a secondary fuel tank configured to store a secondary fuel, a swirler configured to produce a main flame within a combustion chamber of the combustor, and a fuel nozzle configured to produce a pilot flame within the combustion chamber of the combustor. The fuel nozzle includes a nozzle outlet that is located proximate to an end of the swirler or at the end of the swirler, the end of the swirler being located at an inlet of the combustor. The fuel delivery system also includes a primary fuel line fluidly connecting the primary fuel tank to the fuel nozzle and a secondary fuel line fluidly connecting the secondary fuel tank to the swirler.

A dual-fuel fuel nozzle includes a centerbody that at least partially defines an air plenum therein. A tip body is disposed at a downstream end of the centerbody. The tip body includes an upstream side that is axially spaced from a downstream side. The downstream side defines a circular slot and a circular recess that is positioned radially inwardly from the circular slot. The circular recess includes a floor and a side wall, The floor defines an insert opening and the side wall defines an annular groove that includes an undercut surface. The tip body further defines a plurality of orifices annularly arranged around the insert opening radially inward from the side wall of the circular recess. Each orifice includes an inlet that is in fluid communication with the air plenum and an outlet that is oriented towards the undercut surface.

A combustion system includes an electrically actuated flame location control mechanism.

An air and fuel injection system for a turbine engine combustion chamber dome, including at least two fuel injection devices including a central injector and a peripheral annular injector arranged around the central injector, and an inner annular air inlet channel into which the central injector opens up to enable mixing between the fuel from the central injector and the air inlet into the inner annular channel, and at least one outer annular air inlet channel to enable enrichment of this mix with air and to stabilize the pilot combustion zone, wherein the inner annular channel and at least one outer annular channel are separated by an intermediate annular wall extending around the central injector and having a convergent inner profile.

A fuel nozzle apparatus for a gas turbine engine includes: a first pilot fuel injector disposed on a centerline axis of the fuel nozzle which defines a direction of air flow through the fuel nozzle, the first pilot fuel injector being of a pressure atomizing type; an annular second pilot fuel injector at least partially surrounding the first pilot fuel injector, the second pilot fuel injector being of an air blast type and having a fuel outlet disposed axially downstream and radially outboard of the first pilot fuel injector; an annular venturi surrounding the first and second pilot fuel injectors, the venturi including a throat of minimum diameter; an array of inner swirl vanes extending between the first pilot fuel injector and the second pilot fuel injector; and an array of outer swirl vanes extending between the second pilot fuel injector and the venturi.

The present application provides a fuel nozzle for a gas turbine engine using a primary fuel and a secondary fuel. The fuel nozzle may include a number of primary fuel injection ports for the primary fuel, a water passage, a number of secondary fuel injection ports, and a secondary fuel evaporator system for atomizing the secondary fuel.