A ducting arrangement (10) in a combustion stage downstream of a main combustion stage of a gas turbine engine is provided. A duct (18) is fluidly coupled to receive a cross-flow of combustion gases from the main combustion stage. Duct (18) includes a duct segment (23) with an expanding cross-sectional area (24) where one or more injector assemblies (26) are disposed. Injector assembly (26) includes one or more reactant-guiding structures (27) arranged to deliver a flow of reactants into the downstream combustion stage to be mixed with the cross-flow of combustion gases. Disclosed injector assemblies are arranged in expanding cross-sectional area (24) to reduce total pressure loss while providing an effective level of mixing of the injected reactants with the passing cross-flow. Respective duct components or the entire ducting arrangement may be formed as a unitized structure, such as a single piece using a rapid manufacturing technology, such as 3D Printing/Additive Manufacturing (AM) technologies.

A combustor may comprise an outer wall defining, at least, a portion of a combustion chamber. A dilution chute may extend from an interior surface of the outer wall. A support structure may extend between the dilution chute and the interior surface of the outer wall. A spall plate may extend from the interior surface of the outer wall. The spall plate may be located between the support structure and an outlet of the combustion chamber.

A turbine combustor section has a flow conditioner including a plurality of conduits arranged to convey pressurized air to an air chamber for entrance into a plurality of fuel nozzles. Each conduit includes an inlet configured to receive the pressurized air from an annular passage and an outlet configured to deliver the pressurized air to the air chamber. A cross-sectional area of each conduit varies between the inlet and the outlet so as to reduce a pressure drop across the flow conditioner.

A grommet for a combustor of a gas turbine engine is disclosed. The grommet includes a lower platform and a raised platform. The lower platform includes a plurality of top slots located in the top surface and a plurality of bottom slots located in the bottom surface. Each top slot and bottom slot may extend radially along the lower platform. The plurality of top slots may be spaced equidistantly from one another along the circumference of the lower platform. The plurality of bottom slots may be spaced equidistantly from one another along the circumference of the lower platform.

The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a liner defining a combustion chamber therewithin and a pressure plenum surrounding the liner. The liner defines an opening and includes a walled chute disposed at least partially through the opening. A plurality of flow openings is defined through the walled chute.

A fuel nozzle for a gas turbine engine that includes: an elongated centerbody; an elongated peripheral wall formed about the centerbody so to define a primary flow annulus; a primary fuel and air supply in fluid communication with the primary flow annulus; and a pilot nozzle. The pilot nozzle, formed in the centerbody, may have axially elongated air and mixing tubes extending between inlets and outlets defined, respectively, through upstream and downstream faces of the pilot nozzle. A secondary air supply may be communicate with the inlets of the air and mixing tubes. A fuel port may be positioned in the mixing tubes for connecting each to a secondary fuel supply. An uninterrupted sidewall sealing structure in each of the air tubes may segregate an air flow therethrough from the secondary fuel supply. The air and mixing tubes may be configured as canted tubes so to induce a downstream swirling flow.

A mixer having a housing, a duct within the housing, a first and a second injector arranged to inject a fluid at a centre zone of the duct, a third and a fourth injector arranged to inject the fluid at a wall zone of the duct. The first/third injectors are at a distance D1=v/2f.sub.1 or odd integer multiples of it from the second/fourth injectors in the absence of an acoustic node between them, or at a distance D1=λ.sub.conv=v/f.sub.1 or full wave length integer multiples of it in the presence of an acoustic node between them. Advantageously f.sub.1 is greater than f.sub.2.

A cooling system (10) for a fuel system in a turbine engine (14) that is usable to cool a fuel nozzle (16) is disclosed. The cooling system (10) may include one or more cooling system housings (18) positioned around the fuel nozzle (16), such that the cooling system housing (18) forms a cooling chamber (20) defined at least partially by an inner surface (22) of the cooling system housing (18) and an outer surface (24) of the fuel nozzle (16). The fuel nozzle (16) may extend into a combustor chamber (26) formed at least in part by a combustor housing (32). The fuel nozzle (16) may include one or more fuel exhaust orifices (28) with an opening (30) in an outer surface (24) of the fuel nozzle (16) and configured to exhaust fluids unrestricted by the housing (18) forming the cooling system cooling chamber (20). The cooling system (10) may provide cooling fluids to cool the fuel nozzle (16) within the cooling system cooling chamber (20) regardless of whether the fuel nozzle (16) is in use.

A multi-tube combustor and a gas turbine having the same are provided. The multi-tube combustor includes a nozzle casing configured to receive compressed air from a compressor section and to receive fuel from outside, a liner coupled to the nozzle casing and defining a combustion chamber in which a mixture of fuel ejected from the nozzle casing and the compressed air is combusted, a transition piece connected to the liner and configured to supply the combustion gas generated in the combustion chamber defined by the liner to a turbine section, a plurality of fuel nozzles disposed in a nozzle tube provided inside the nozzle casing, each fuel nozzle having a cavity, a plurality of compressed air supply tubes connected to the plurality of fuel nozzles and configured to supply the compressed air to the plurality of fuel nozzles, a center body having a first end connected to a nozzle base of the nozzle casing, extended to enter the cavity of the fuel nozzle, and having a second end having a fuel ejection hole through which the fuel is ejected, and an on/off valve provided on the plurality of compressed air supply tubes to open and close the compressed air supply tubes. The fuel and the compressed air are mixed inside the plurality of fuel nozzles, and the plurality of fuel nozzles are divided into a plurality of fuel nozzle groups, and a mixture of the fuel and the compressed air is ejected from one or more selected fuel nozzle groups of the plurality of fuel nozzle groups according to a combustion load condition or during a ramp-up process.

An assembly for a gas turbine having a combustion chamber, a swirler, a combustion zone in the combustion chamber, and an air feed. In a transition region from the air feed to the swirler that is flowed through by the air, a plenum is formed and the assembly adjoining the plenum has the swirler, the combustion chamber, and a cover closing the combustion chamber. The assembly has an air conduction channel designed to conduct part of the air flow flowing into the assembly into the combustion chamber, so that the air flow leading through the air feed is divided into a main flow leading through the swirler into the combustion zone and a bypass flow leading past the combustion zone.