Exhaust frequency mitigation apparatus, exhaust diffusers, and turbomachines are provided. An exhaust frequency mitigation apparatus includes a main body extending along an axial centerline from a base to a tip. the base defines a first diameter and the tip defining a second diameter. the main body converges from the first diameter to the second diameter with respect to the axial centerline. The base is configured to extend from an inner shell of a turbomachine exhaust diffuser. The exhaust frequency mitigation apparatus further includes at least one rib extending from a root coupled to the main body to a free end.

The invention relates to an exhaust gas turbocharger having a fluid dynamic bearing having a rotor (10) and a counter-bearing part (50) assigned to the rotor (10), wherein a rotor bearing surface of the rotor (10) and a counterface of the counter-bearing part (50) face each other, to form a fluid dynamic bearing, wherein the rotor bearing surface and/or the counterface form(s) a continuous bearing contour when cut longitudinally and through the axis of rotation (R) in sectional view, which bearing contour(s) are formed of at least two contour sections (44.1 to 44.3; 53.1 to 53.3) to generate fluid dynamic load capacities in both the radial and the axial directions, wherein the bearing surface of the rotor (10) is formed by a rotor part (40), which is connected to a rotor shaft (11) and is secured on the rotor shaft (11), and wherein the rotor part (40) is supported relative to the rotor shaft (11) in the area of a support section (14) of the rotor shaft (11). In order to be able to provide such an exhaust gas turbocharger with a compact and efficient bearing arrangement having a fluid dynamic bearing, wherein at the same time the fluid dynamic bearing can be easily mounted using few parts, provision is made according to the invention that the support section (14) and at least one of the contour sections (53.1 to 53.3) of the counter-bearing part (50) at least sectionally overlap in the direction of the axis of rotation (R).

A turbine comprises a casing, an outer metal shroud, an inner metal shroud and an annular distributor having a plurality of CMC ring sectors, each sector comprising a mast, an inner platform, an outer platform and at least one blade having a hollow profile that defines an inner housing, the inner and outer platforms each having an opening communicating with said inner housing, and the mast passing through said openings and the inner housing and being secured to said casing and connected to said annular sector. Each blade comprises at least one first radial shoulder projecting axially towards the inside of the blade, and each mast comprises at least one second shoulder projecting axially towards the outside of the mast configured to radially cooperate with a first shoulder and radially press the blade against the mast.

A triple-flow turbomachine for an aircraft, including a power transmission module including a torque input connected to a turbine shaft, a first torque output of a gearbox connected to a main shaft for rotatably driving a main fan propeller, and a second torque output of a planet gear connected to a secondary shaft for rotatably driving a secondary fan propeller. The planet gear is independent of the gearbox and arranged downstream of the gearbox.

A nozzle assembly for a gas turbine engine and methods for assembling a nozzle assembly are provided. In one example aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced from the outer wall. The outer wall defines a plurality of mounting openings spaced circumferentially from one another. The inner wall defines a plurality of mounting openings spaced circumferentially from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The nozzle assembly includes vanes that are inserted through the mounting openings of the outer wall in a radially inward direction and vanes that are inserted through the mounting openings of the inner wall in a radially outward direction in an alternating manner.

Disclosed is a method of repairing a rear cone segment erosion coating including removing a portion of an existing fluoroelastomer coating from a rear cone segment to form a repair area; forming a cured replacement erosion coating corresponding to the repair area, wherein the cured replacement erosion coating has a layer of fluoroelastomer and a layer of fibrous reinforcement; preparing the repair area and the cured replacement erosion coating for adhesion; locating the prepared cured replacement erosion coating on the prepared repair area; and adhering the prepared cured replacement erosion coating to the prepared repair area. Also disclosed is a cured replacement erosion coating and a repaired inlet nose cone.

Assembly for axial turbomachine, in particular for an aircraft turbojet, the assembly comprising: an annular casing with an internal surface (40); an annular row of stator baffles (26) with at least one stator baffle (26) comprising an airfoil (50) which extends radially from a fixing platform (34), the fixing platform (34) being fixed to the casing and having a polygonal outline; characterised in that it further comprises a gasket (80) comprising a frame, the outline of which conforms to the polygonal outline of the fixing platform (34), the frame being in radial contact with the fixing platform (34) and the casing in order to ensure a seal.

An ejector comprises a primary nozzle having an annular wall forming part of an outer boundary of an exhaust portion of a primary flow path of a gas turbine engine. The annular wall has a downstream end defining a plurality of circumferentially distributed lobes. The ejector further comprises a secondary nozzle having an annular wall disposed about the primary nozzle, the primary nozzle and the secondary nozzle defining a secondary flow passage therebetween for channeling a secondary flow. The secondary nozzle defines a mixing zone downstream of an exit of the primary nozzle. A flow guide ring is mounted to the primary nozzle lobes. The ring has an aerodynamic surface extending from a leading edge to a trailing edge respectively disposed upstream and downstream of the exit of the primary nozzle. The aerodynamic surface of the ring is oriented to guide the high velocity primary flow into the mixing zone.

An engine assembly for use with a gas turbine engine includes a first component, a second component, and a retention locking plug. The first component is formed to define a passage that extends into the first component. The second component is received in the passage defined in the first component. The retention locking plug extends into the first component and the second component to couple the second component with the first component.

A gas turbine engine with a lubrication system includes a ball bearing assembly and a rotor circumscribing a rotational axis and journaled within the ball bearing assembly. The gas turbine engine also includes a lubrication system located radially outward from a rotational axis and radially outward and adjacent to the ball bearing assembly, which includes a lubrication channel having an inlet and an outlet and a dispersion cone adjacent to the outlet of the lubrication channel.