A seal assembly to seal a gas turbine hot gas path flow at an interface of a combustor liner and a downstream component, such as a stage one turbine nozzle, in a gas turbine. The seal assembly including a piston ring seal housing, defining a cavity, and a piston ring disposed within the cavity. The piston ring disposed circumferentially about the combustor liner. The piston ring is responsive to a regulated pressure to secure sealing engagement of the piston ring and outer surface of the combustor liner. The seal assembly includes at least one of one or more sectional through-slots, bumps or channel features to provide for a flow therethrough of a high-pressure (P.sub.high) bypass airflow exiting a compressor to the cavity. The high-pressure (P.sub.high) bypass airflow exerting a radial force on the piston ring.

A flexible seal for sealing between two adjacent gas turbine components includes a forward end, an aft end axially separated from the forward end, and an intermediate portion between the forward end and the aft end. The intermediate portion defines a continuous curve in the circumferential direction, such that the aft end is circumferentially offset from the forward end. In other cases, the forward and aft ends are axially, radially, and circumferentially offset from one another. A method of sealing using the flexible seal includes inserting, in an axial direction, the aft end of the flexible seal into a recess defined by respective seal slots of two adjacent gas turbine components; and pushing the flexible seal in an axial direction through the recess until the forward end is disposed within the recess.

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 a fence disposed between the forward and aft liner segments, wherein the fence extends in the circumferential direction, and wherein the fence extends into the combustion chamber along the radial direction.

Method for manufacturing a CMC, i.e. ceramic matrix composite material, part provided with at least one cutout, as well as to such a CMC part provided with at least one cutout, the method comprising the following steps: providing (E1) a fibrous reinforcement (10), forming (E2′) a cavity in a portion of the fibrous reinforcement (10), injecting (E3) a slip comprising at least a ceramic powder and a solvent, the slip being injected so as to impregnate the fibrous reinforcement (10′) and to fill the cavity of the fibrous reinforcement (10′), drying (E4) the obtained assembly, carrying out a densification (E6) by infiltration of a liquid densification material and solidification of said densification material, machining (E7) at least one cutout in the obtained blank (30) within the volume corresponding to the cavity of the fibrous reinforcement (10).

The invention relates to a combustion chamber for a gas turbine engine, in which two openings extend through a metal wall (58) and a wall made of a refractory material (16). A device (76) guides a spark plug (48) in said two orifices, the guiding device comprising a metal flange (80) and a floating ring (82), the flange being shrunk onto or welded to the metal wall (58).

A combustor arrangement includes a fuel injector, and a seal arranged around the fuel injector and having an upstream end and a downstream end. The seal includes an annular body at least partially abutting the fuel injector. The annular body includes an inner surface facing the fuel injector and an outer surface radially spaced apart from the inner surface relative to the central axis. The annular body further includes a plurality of slots disposed on the inner surface and circumferentially spaced apart from each other relative to the central axis. Each slot axially extends at least partially from the downstream end to the upstream end. Each slot is disposed in fluid contact with the fuel injector. The seal also includes a flange radially extending from the outer surface of the annular body at the upstream end.

A combustor assembly for a gas turbine engine defining a radial direction and a circumferential direction includes a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and a seal member including a body, a flange, and a radial element, the body defining a body surface extending along the radial direction and positioned adjacent the interface surface of the at least one liner, the flange extending forward from the body, and the radial element coupled to the flange and extending into the radial opening defined by the at least one liner.

A combustor includes an inner liner and an outer liner defining a combustion chamber. The inner liner includes an inner mesh structure, a plurality of hot side planks mounted to a hot side of the inner mesh structure, and a plurality of cold side planks mounted to a cold side of the inner mesh structure. The outer liner includes an outer mesh structure, a plurality of hot side planks mounted to a hot side of the outer mesh structure, and a plurality of cold side planks mounted to a cold side of the outer mesh structure. The combustor further includes a plurality of clips configured to couple the plurality of hot side planks and the plurality of cold side planks to a plurality of structural elements of the inner mesh structure and the outer mesh structure.

A combustor for a gas turbine engine includes a forward liner segment and an aft liner segment positioned downstream from the forward liner segment relative to a direction of flow through the combustor. The forward and aft liner segments at least partially define a combustion chamber. Furthermore, the combustor includes a dilution slot frame positioned between the forward and aft liner segments along a longitudinal centerline of the gas turbine engine. Moreover, the dilution slot frame defines a plurality of dilution slots spaced apart from each other along a circumferential direction of the gas turbine engine such that the plurality of dilution slots provides an annular ring of dilution air to the combustion chamber.

Combustor assemblies are provided. For example, a combustor assembly includes a combustor liner defining a combustion chamber and an annular combustor dome positioned at a forward end of the combustor liner that defines a plurality of dome apertures. The combustor assembly further includes an annular heat shield positioned between the combustor dome and the combustion chamber, a plurality of adapters positioned forward of the heat shield, and a plurality of collars. The heat shield defines a plurality of heat shield apertures that are aligned with the dome apertures. One adapter is attached to the combustor dome at each dome aperture, and the adapters are. One collar extends through each heat shield aperture to couple the heat shield to the combustor dome. Further, ceramic matrix composite (CMC) heat shields are provided that may include an annular body defining a plurality of heat shield apertures, as well as inner and outer wings.