A composite vane includes both a vane body extending in a longitudinal direction between an inner end and an outer end and also at least one fastener base connected to the inner end or to the outer end of the vane body. The vane includes metal reinforcement associated with an organic matrix. The metal reinforcement defines integrally as a single part both a longitudinal core extending between the inner and outer ends of the vane body and also the fastener base, the longitudinal core having the organic matrix overmolded thereon, the organic matrix defining the outside shape of the vane.

An acoustic liner can comprise a support layer including a set of partitioned cavities defining a set of cells with open faces, as well as a first and second facing sheet operably coupled to the support layer that overlie and close the open faces. A set of aerogel fillings can be provided within at least some of the set of cells.

A nosecone for a turbine engine includes a nosecone body and a nosecone mount. The nosecone body extends along an axis between a tip end and a base end. The nosecone body is configured from or otherwise includes thermoplastic material. The nosecone body includes a shell and an arrangement of ribs, which structurally support at least a portion of the shell. A thickness of the arrangement of ribs is greater than or substantially equal to approximately one half of a thickness of the shell. The nosecone mount is adapted to connect the nosecone body to a component of the turbine engine.

A low-pressure compressor for a turbine engine, such as an aircraft turbojet engine includes a rotor with two rows of rotor blades between which two annular ribs are positioned; and one annular row of stator blades between the rotor blades. An internal shroud is connected to the stator blades. The internal shroud includes abradable material collaborating with the annular ribs, and annular teeth made of an abradable material and which extend radially towards the rotor, so as to provide sealing. The system may be used in a method for manufacturing a bypass turbojet engine compressor.

An assembly for a turbojet, wherein the assembly includes an outer shroud and an inner shroud that are concentric, wherein the inner shroud is segmented and includes circumferential clearances between the segments thereof. The assembly additionally includes an annular row of stator vanes connecting the inner shroud to the outer shroud, a drive with a reduction ratio that is intended to be coupled to a fan, and a circuit for cooling and for lubricating the drive. The circuit is configured to heat up at least the outer shroud during the operation of the turbine engine such as to circumferentially reduce the circumferential clearances between the segments.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

A bearing element for at least one adjustable guide vane of a turbomachine, in particular an aircraft engine, is provided. The bearing element comprising a porous matrix made of carbon and/or graphite and at least one metallic phase or a metal salt that is at least partially arranged inside the pores of the matrix, so that what results is a metal-infiltrated or a metal-salt-infiltrated material.

A composite vane includes both a vane body extending in a longitudinal direction between an inner end and an outer end and also at least one fastener base connected to the inner end or to the outer end of the vane body. The vane includes metal reinforcement associated with an organic matrix. The metal reinforcement defines integrally as a single part both a longitudinal core extending between the inner and outer ends of the vane body and also the fastener base, the longitudinal core having the organic matrix overmolded thereon, the organic matrix defining the outside shape of the vane.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan having a plurality of fan blades, and a nacelle surrounding the plurality of fan blades, a compressor section including a low pressure compressor and a high pressure compressor, the low pressure compressor including a plurality of stages, and the high pressure compressor including 6 or more stages. A turbine section includes a fan drive turbine that drives the fan section through a gear arrangement, and including a second turbine that drives the high pressure compressor. A power ratio is provided by the combination of a first power input of the low pressure compressor and a second power input of the high pressure compressor, the power ratio defined by the second power input divided by the first power input, and the power ratio is less than or equal to 1.0

A fiber preform for a turbine engine blade or vane obtained by single-piece three-dimensional weaving. The preform includes a first longitudinal segment suitable for forming a root; a second longitudinal segment extending the first longitudinal segment and suitable for forming an airfoil portion; and a first transverse segment extending transversely from the junction between the first and second longitudinal segments and suitable for forming a first platform.