The present disclosure is directed to a fuel injector assembly and a gas turbine engine including the fuel injector assembly. The fuel injector assembly includes a centerbody extended along a lengthwise direction. The centerbody defines a first fuel nozzle. An annular shroud defining a second fuel nozzle surrounds the centerbody and is extended along the lengthwise direction. A passage is defined through the annular shroud and extended generally along the lengthwise direction. The passage defines an exit opening disposed at a downstream end adjacent to the combustion chamber and in fluid communication therewith. The annular shroud defines a fuel inlet opening disposed at an upstream end of the passage. The annular shroud further defines an air inlet opening in fluid communication with the passage. The air inlet opening is disposed between the fuel inlet opening and the exit opening.

A combustion chamber comprises an upstream end wall structure and inner and outer annular wall structures. The upstream end wall structure comprises an upstream wall and a plurality of circumferentially arranged heat shields secured to the upstream wall. The upstream wall has a plurality of circumferentially spaced fuel injector apertures. Each heat shield has radially outer and radially inner ends and a fuel injector aperture aligned with a corresponding fuel injector aperture in the upstream wall. The radially outer and inner ends of each heat shield have outer and inner rails spacing the heat shield from the upstream wall. The radially outer and inner ends of each heat shield have first and second pluralities of circumferentially spaced apertures extending there-through and through the associated outer and inner rails to direct coolant over the surface of the outer and inner annular wall structures to form respective films of coolant.

The gas turbine engine includes a fluid system fluidly connecting at least two components of the gas turbine engine, and a tunable resonator in fluid flow communication with the fluid system. The tunable resonator has a resonating volume that varies as a function of a volume of an inflatable member located inside the tunable resonator. The inflatable member having a means for varying the volume of the inflatable member, to thereby tune the resonating volume to a selected frequency of pressure fluctuations or acoustic waves within the fluid system.

A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber; and an intermediate member disposed at least partially between the forward and aft liner segments and extending in the circumferential direction.

A combustor may comprise an outer combustor panel and an inner combustor panel radially inward of the outer combustor panel. A bulkhead panel may extend radially between the outer combustor panel and the inner combustor panel. An outer spacer may be located between an outer flange of the bulkhead panel and the outer combustor panel. An inner spacer may be located between an inner flange of the bulkhead panel and the inner combustor panel.

The present invention generally relates to a gas turbine and more in particular it is related to a damper assembly for a combustion chamber of a gas turbine. According to preferred embodiments, the present solution provides a damper assembly including protrusions on a wall of the neck. These protrusions result in a side wall reactance to the acoustic field that has the effect of decreasing the effective speed of sound in the neck. The decrease of the effective speed of sound in the neck is equivalent to an increase of the effective neck length.

An axial fuel staging (AFS) system for a combustor is provided. The AFS system includes a secondary fuel injector, an injector housing surrounding the secondary fuel injector, and a conduit assembly in flow communication with the secondary fuel injector. The conduit assembly includes a first conduit for conveying fuel to the secondary fuel injector, and a second conduit circumscribing the first conduit such that a channel is defined between the first conduit and the second conduit. The first conduit and the second conduit terminate within the injector housing.

A combustor assembly for a gas turbine engine includes a liner and a combustor dome. The combustor dome and liner together at least in part define a combustion chamber. The combustor dome includes a transition portion and a forward wall. The transition portion extends from the forward wall towards the liner. Additionally the transition portion may define an angle relative to the forward wall and/or may define a radius of curvature between the forward wall of the combustor dome and a flat of the transition portion of the combustor dome for increasing an aerodynamic efficiency of the combustor dome.

A combustor wall for a turbine engine with an axial centerline comprising a combustor support shell comprising a plurality of impingement apertures; a combustor heat shield comprising a plurality of effusion apertures fluidly coupled with the plurality of impingement apertures; and at least one shaped pad formed in said combustor heat shield, said at least one shaped pad extending through a cutout in said combustor support shell.

A combustor of a gas turbine engine including a combustor shell having an interior surface, a first panel mounted to the interior surface at a first position and a second panel mounted to the interior surface at a second position. The first panel has a first end, a first combustion chamber surface parallel with the interior surface, a first rail extending from the first combustion chamber surface toward the interior surface of the combustor shell, and a first extension extending axially from the first rail to the end of the first panel. The second panel has a second end, a second combustion chamber surface, and a second rail extending from the second combustion chamber surface toward the interior surface of the combustor shell. The first end and the second end are proximal to each other and define a circumferentially extending gap there between.