A turbomachine fuel injector system includes an air distributor configured to be mounted to and contained within a casing and to provide air to mix with a fuel from one or more fuel distribution systems. The system includes a fuel manifold configured to be mounted indirectly to the casing such that the fuel manifold system is contained within the casing but is independent such that fuel manifold does not touch or directly mount to either an interior of the casing or touch the air distributor within the casing to prevent or reduce thermal transfer from the air distributor to the fuel manifold.

A combustor adapted for use in a gas turbine engine a combustor shell, a heat shield, and a heat shield retainer. The combustor shell is made from metallic materials and is formed to define an internal cavity. The heat shield is formed from ceramic matrix composite materials and is coupled to the dome panel. The heat shield retainer is configured to retain the heat shield to the combustor shell.

A combustor for a gas turbine engine includes a combustor shell, a heat shield and a burner seal. The combustor shell includes metallic materials and is formed to define an interior combustion space. The heat shield includes ceramic matrix composite materials and is configured to shield a portion of the combustor shell from the interior combustion space. The burner seal includes ceramic matrix composite materials and is configured to extend through apertures formed in the combustor shell and the heat shield.

An annular combustor is disclosed in which free-vortices are generated which: enhance fuel air mixing, recirculate the air, provide cooling for the combustor walls, and provide low emissions and a substantially uniform exit temperature profile. The combustor is provided fuel and air through a plurality of fuel injectors which atomizes the fuel and promote individual vortices emanating from each injector. The combustor includes an annular prechamber, an annular main chamber, and an annular dilution zone. A first swirler is provided on the inner side of the annular prechamber and a second swirler on the outer side of the annular prechamber, the swirlers causing flow to rotate in mutually opposite directions to intensify the vortices imposed by the injectors. The vortices cause a pressure depression to promote some backflow that enhance mixing to: promote complete combustion, produce low emissions, and provide cooling.

Combustion sections and sealing systems for fuel ignition assemblies of gas turbine engine combustion sections are provided. For example, a sealing system comprises a ferrule positioned on an outer surface of a ceramic matrix composite (CMC) combustor liner an aperture defined in the CMC liner; a sleeve positioned within an adapter of the fuel ignition assembly such that an inner end portion of the sleeve is in contact with the ferrule, the sleeve having an end wall that forms an inner boundary of a cavity defined by the sleeve; and a biasing member positioned within the cavity. The biasing member extends between a bushing and the end wall of the sleeve. The biasing member continuously urges the sleeve into contact with the ferrule to seal the aperture against fluid leakage therethrough. The exemplary sealing system may be part of a fuel ignition assembly of a gas turbine engine combustion section.

A combustor adapted for use in a gas turbine engine includes a metallic combustor shell forming an interior space, a heat shield, and a liner arranged in the interior space of the metallic case. The liner defines a combustion chamber in which fuel is burned during operation of a gas turbine engine.

A liner panel for a combustor of a gas turbine engine includes a nominal wall thickness and a thickened wall thickness in the region of a hot spot.

A combustor assembly for a gas turbine engine includes a dome having a forward surface and an inner surface. The forward surface and the inner surface of the dome at least partially define a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end of the liner is positioned within the slot of the dome. The forward end of the liner includes an axial interface surface and a radial interface surface. The axial interface surface defines a radial gap with the inner surface of the dome and the radial interface surface defines an axial gap with the forward surface of the dome. At least one of the radial gap or the axial gap is less than about 0.150 inches during operating conditions of the combustor assembly to prevent an undesirable airflow.

A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an arcuate surface section therebetween in the axial profile between the forward section and the aft section.

A fuel injector for a gas turbine engine of an aircraft having a fuel nozzle including a fuel swirler and/or an outer air swirler. The fuel swirler may include a manifold for receiving fuel from a fuel conduit, and a plurality of fuel passages to direct fuel from the manifold to discharge orifices that direct fuel with swirling flow. The fuel swirler may be configured to provide uniform spray while minimizing recirculation zones; reduce residence time as fuel enters the manifold; minimize flow disruptions, boundary layer growth, and/or pressure drop as fuel flows through the fuel passages; reduces coking internally of the nozzle; reduces thermal stresses; and is simple and low-cost to manufacture.