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; and 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, wherein the forward and aft liner segments are coupled together at a moveable interface.

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 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 micro-turbine core fabricated as a single part using 3D additive manufacturing (AM) to simultaneously form sequential layers of at least two static components from any of the following static components: central bearing support structure, outer casing, combustor complete, nozzle guide vanes (NGVs), diffuser, diffuser outer casing, fuel manifold, fuel injector(s), igniter mounting boss, oil manifold, oil distribution lines, or turbine outer casing. The single part does not require fastening hardware, welding, and/or bonding processes to create the single part.

In this method for mounting a combustor component, a component hanging step for attaching a string material to a hanging jig mounted on a combustor component and hanging the combustor component with the string material together with the hanging jig, and a component mounting step for mounting a mounting flange of the combustor component at a combustor mounting position of a gas turbine casing are executed. The hanging jig has a hanging tool having a portion which becomes a suspension point contacted by the string material when the string material is attached and hanging the combustor component. The suspension point is positioned on the distal end side of the center of gravity of the combustor component in the combustor axial line direction when the hanging jig has been mounted to the mounting flange.

In this method for mounting a combustor component, a component hanging step for attaching a string material to a hanging jig mounted on a combustor component and hanging the combustor component with the string material together with the hanging jig, and a component mounting step for mounting a mounting flange of the combustor component at a combustor mounting position of a gas turbine casing are executed. The hanging jig has a hanging tool having a portion which becomes a suspension point contacted by the string material when the string material is attached and hanging the combustor component. The suspension point is positioned on the distal end side of the center of gravity of the combustor component in the combustor axial line direction when the hanging jig has been mounted to the mounting flange.

A turbofan gas turbine engine configured to reduce hotspots within combustors. The engine includes an axis and a combustor that is circumferentially disposed about the axis. The combustor includes an annular combustor liner that includes a front portion and a rear portion. The annular combustor liner is joined to an annular combustor dome via front portion and defines a chamber and a nozzle is mounted within the annular combustor dome and is configured to inject fuel into a plurality of swirlers. At least one or more dilution openings is circumferentially distributed around the liner such that a region is fluidly connected through the annular combustor liner to the chamber. Each one of the pluralities of dilution openings includes an opening and a radial support wall that is positioned aft of the opening such that the radial support wall extends into the chamber.

A combustor liner for a gas turbine engine includes at least one liner segment that has an external wall dimensioned to bound a combustion chamber. The external wall extends between leading and trailing edges in an axial direction and extends between opposed mate faces in a circumferential direction. A cooling circuit is defined by the external wall. A plurality of heat transfer features are distributed in the cooling circuit to define first and second prioritized flow regions on opposed sides of a first restricted flow region.

A combustor liner for a gas turbine engine includes at least one liner segment that has an external wall dimensioned to bound a combustion chamber. The external wall extends between leading and trailing edges in an axial direction and extends between opposed mate faces in a circumferential direction. A cooling circuit is defined by the external wall. A plurality of heat transfer features are distributed in the cooling circuit to define first and second prioritized flow regions on opposed sides of a first restricted flow region.

A turbofan gas turbine engine configured to reduce hotspots within combustors. The engine includes an axis and a combustor that is circumferentially disposed about the axis. The combustor includes an annular combustor liner that includes a front portion and a rear portion. The annular combustor liner is joined to an annular combustor dome via front portion and defines a chamber and a nozzle is mounted within the annular combustor dome and is configured to inject fuel into a plurality of swirlers. At least one or more dilution openings is circumferentially distributed around the liner such that a region is fluidly connected through the annular combustor liner to the chamber. Each one of the pluralities of dilution openings includes an opening and a radial support wall that is positioned aft of the opening such that the radial support wall extends into the chamber.