An arrangement for an aircraft turbine engine combustion chamber including an injection system and a fuel injector is provided. The injection system includes an injector nozzle guide, the inner surface of which delimits an opening for centering the nozzle, which includes an outer casing. The arrangement further includes a sealing device between the inner surface of the guide and the outer casing. The sealing device includes a first part accommodated in a groove of the outer casing, the groove being delimited, in part, by a downstream delimiting surface, the first part having a first sealing surface and bearing axially against the downstream delimiting surface; and a second part having a second sealing surface bearing radially against the inner surface of the guide.

A turbomachine and a method of operating the turbomachine are disclosed. The turbomachine includes a stator, a rotor including a rotor bearing face, and a face seal assembly including a first segmented seal ring and a second segmented seal ring. The first segmented seal ring includes a plurality of joints and a first flat-contact surface and the second segmented seal ring includes a plurality of segment ends and a second flat-contact surface. One of the first and second segmented seal rings includes a seal bearing face. The second segmented seal ring is coupled to the first segmented seal ring such that the second flat-contact surface is in contact with the first flat-contact surface. The plurality of segment ends is circumferentially offset from the plurality of joints. The first segmented seal ring is slidably coupled to the stator and defines a face seal clearance between the rotor and seal bearing faces.

An epicyclic gear train for a turbine engine includes a gutter with an annular channel. A rotating structure includes a ring gear that has an aperture that is axially aligned with the annular channel. Axially spaced apart walls extend radially outward relative to the rotating structure to define a passageway. The passageway is arranged radially between and axially aligned with the aperture and the annular channel. The walls are configured to inhibit an axial flow of an oil passing from the aperture toward the annular channel.

A turbine rotor blade that includes a tip shroud attached to the outboard tip of the airfoil. The tip shroud may include an axially and circumferentially extending planar component in which an inboard surface opposes an outboard surface, and a shroud edge that connects the inboard surface to the outboard surface and defines an outboard profile of the tip shroud. The tip shroud may include a seal rail protruding from the outboard surface of the tip shroud and a cutter tooth disposed on the seal rail. The cutter tooth may be formed as a circumferential section of the seal rail that is axially thickened. The seal rail may further include a leakage gap formed therethrough that is configured to increase a leakage level during operation.

When the compressed air is not applied, an inclined surface, which is positioned at an upstream side with respect to the supply direction of the compressed air, of the lip part is in contact with the inclined surface of the inner ring. When the compressed air is applied, a contact area between the inclined surface, which is positioned at the upstream side with respect to the supply direction of the compressed air, of the lip part and the inclined surface of the inner ring becomes smaller, as compared to when the compressed air is not applied.

A seal segment has an arcuate body having: a first end; a second end circumferentially opposite the first end; a first face; a second face axially opposite the first face; an inner diameter (ID) face; and an outer diameter (OD) face. The seal segment is shaped to interfit with a plurality of identical seal segments first end to second end to form a seal surrounding a central longitudinal axis. The first face has: a circumferential channel closer to the ID face than the OD face; and a plurality of channels extending from the circumferential channel to the OD face. The ID face has: a circumferential channel closer to the first face than the second face; and a plurality of channels extending from the circumferential channel to the second face. The ID face circumferential channel has an open end.

A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

A seal assembly for a gas turbine engine having a seal formed of a carbon material; and a seal seat positioned for rotation relative to the seal, wherein the seal and the seal seat each have a sealing surface which together define a sliding seal, and further having a carbon film on the sealing surface of the seal seat.

A rotating machine includes a rotating body rotatably supported in a casing; a rotor blade fixed to an outer peripheral portion of the rotating body; a stator blade arranged on a downstream side in a fluid flow direction with respect to the rotor blade and fixed to an inner peripheral portion of the casing; a sealing device arranged between the inner peripheral portion and a front end of the rotor blade; a swirling flow generation chamber provided in the casing on the downstream side from the sealing device along a circumferential direction of the rotating body; and guiding members provided in the swirling flow generation chamber along a radial direction of the rotating body and in the circumferential direction at predetermined intervals. The swirling flow generation chamber has a wall surface located on the downstream side from an edge of the stator blade on an upstream side.

Sealing arrangements and turbomachines are provided. A sealing arrangement includes a transition duct having an upstream end and a downstream end. The transition duct includes an aft frame that circumferentially surrounds the downstream end of the transition duct. A first stage nozzle is spaced apart from the aft frame and defines a gap therebetween. A sealing assembly is coupled to the aft frame. The sealing assembly includes a flexible sealing element that extends from the aft frame, across the gap, to the first stage nozzle. The flexible sealing element is forced into sealing engagement with the first stage nozzle by pressure from a compressed working fluid. The sealing assembly further includes a heat shield disposed between the flexible sealing element and the aft frame. The heat shield terminates within the gap.