In some aspects, a gas turbine combustor assembly is arranged around a longitudinal axis. The gas turbine combustor comprises a first fuel/air mixer assembly, the mixer assembly comprising a first fuel injector and a plurality of first mixer elements, each mixer element defining an air flow passage therethrough having an outlet in a first plane. A second fuel/air mixer assembly comprises a second fuel injector and a plurality of second mixer elements, and each second mixer element defines an air flow passage therethrough having an outlet in a second plane, longitudinally offset from the first plane.

A combustor having a main chamber and a trapped vortex cavity. The main chamber includes an outer liner and an inner liner. The trapped vortex cavity extends from at least one of the outer liner or the inner liner. A plurality of mixing assemblies operably injects a fuel-air mixture into the trapped vortex cavity to produce combustion gases. The trapped vortex cavity injects the combustion gases into the main chamber. A steam system is in fluid communication with the main chamber. The steam system operably injecting steam into the main chamber such that the steam flows downstream of the trapped vortex cavity.

A method for fabricating a main body of a trapped vortex fuel injector having a main body defining a fuel circuit. The method includes determining three-dimensional information of the main body including the fuel circuit where the fuel circuit is fully circumscribed within the main body and extends between an annular portion and a semi-annular portion of the main body and where the three-dimensional information of the main body further includes a plurality of fuel injection ports which provide for fluid communication between the fuel circuit and a trapped vortex pre-mix zone. The method further includes converting the three-dimensional information into a plurality of slices that define a cross-sectional layer of the main body and successively forming each layer of the main body by fusing a metallic powder using laser energy or electron beam energy.

A combustor assembly including a combustor liner defining therein a combustion chamber for the downstream flow of a main fluid. At least two annular trapped vortex cavities are located on the combustor liner and staged axially spaced apart. A cavity opening is located at a radially inner end of each of the at least two annular trapped vortex cavities spaced apart from a radially outer wall and extending between an aft wall and a forward wall of each cavity. A plurality of injectors are configured tangentially relative to the circular radially outer wall to provide for an injection of air and fuel to form an annular rotating trapped vortex of a fuel and air mixture within a respective annular trapped vortex cavity. The annular rotating trapped vortex of the fuel and air mixture at the cavity openings is substantially perpendicular to the downstream flow of the main fluid. A gas turbine engine including the combustor assembly is disclosed.

Disclosed herein is an apparatus. The apparatus comprises an injector coupled to a head portion of a combustion chamber, the injector comprising a plurality of injector elements distributed away from an inner annulus and in an outer annulus. A geometry of combustion chamber comprises a body portion, an optional shoulder portion, and a throat portion. An inner wall of combustion chamber converges radially inward towards the throat. The plurality of injector elements in combination with the geometry of the combustion chamber are configured to confine a predetermined percentage of mass flow associated with combustion to a predetermined outer annulus of the chamber.

This assembly comprises a wall body in which is formed a through-hole passing through in a thickness direction, and a component fixed to one surface of the wall body so as to cover the through-hole, a welded part welded so as to fix the component to the one surface being accommodated within the through-hole.

A combustor of a turbine engine includes a first combustion zone operable to combust a first fuel and air mixture, a first fuel inlet for providing a first fuel, a first air inlet for providing first zone air, the first fuel and the first zone air combining to form the first fuel and air mixture in the first combustion zone, a second combustion zone operable depending on turbine engine operating parameters, for combusting a second fuel and air mixture, a second fuel inlet providing a second fuel, that operates when the second combustion zone is operating and does not operate when the second combustion zone is not operating, and a second air inlet providing second zone air, the second fuel and the second zone air combining to form the second fuel and air mixture in the second combustion zone, wherein the first fuel and the second fuel are disparate fuels.

This assembly comprises a wall body in which is formed a through-hole passing through in a thickness direction, and a component fixed to one surface of the wall body so as to cover the through-hole, a welded part welded so as to fix the component to the one surface being accommodated within the through-hole.

A gas turbine engine includes a compressor configured to receive inlet air at a compressor inlet and generate compressed air at a compressor exit, a combustor positioned fluidically and physically downstream of the compressor, a turbine positioned fluidically and physically downstream of the combustor, and a shaft mechanically connecting the turbine and the compressor. The combustor is fluidically connected to the compressor to receive a first portion of the compressed air as combustor primary inlet air and also includes a toroidal recirculation zone configured to receive and combust fuel in a rich combustion zone, an ignitor positioned to ignite an air/fuel mixture in the rich combustion zone, a rapid quench zone downstream of the toroidal recirculation zone, a lean combustion zone downstream of the rapid quench zone, and a cooling air flow path configured to direct a second portion of the compressed air around an outer combustor liner.

A system includes at least one component of a turbine section. The at least one component includes an inert gas port formed into a wall of the at least one component. The inert gas port is configured to inject an inert gas into a chamber of the turbine section in a downstream direction. The at least one component also includes a fuel port formed into the wall upstream of the inert gas part. The fuel port is configured to inject a fuel in the downstream direction toward the injected inert gas. The inert gas causes the injected fuel to be lifted away from the surface of the wall and propelled into the hot gas flow path, where the fuel can be ignited by the combustion gases to increase gas turbine engine efficiency and performance.