Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

Abstract: A low-pressure compressor for an axial-flow turbomachine, such as a jet engine, includes an annular row of vanes and a between-vanes passage with a connecting surface that links the pressure surface of a first vane to the suction surface of a second vane of the row. The connecting surface includes a main protuberance which includes a first elevation and a second elevation that are spaced apart from one another.

Abstract: A turbine engine compressor blade includes a leading edge, a trailing edge, a suction surface, and a pressure surface. In addition, the blade includes at least one irregularity in the form of a projecting protuberance of the suction surface or the pressure surface or in the form of a recess nested in the suction surface or the pressure surface. The irregularity may have a direction of longest dimension substantially parallel to the leading edge or substantially axial.

Abstract: In an axial turbomachine low-pressure compressor, the compressor, or booster, includes a first vane with an intrados surface, a second vane with an extrados surface, an inter-vane passage with a guiding surface, which connects the intrados surface to the extrados surface delimiting an annular vein, and which has a bulge and a recess forming a so-called 3D contouring geometry, or non-axisymmetric surface. The guiding surface includes an axially symmetrical or substantially flat zone, which axially extends through the inter-vane passage so as to separate the recess from the bulge.

Abstract: 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.

Abstract: A system for measuring turbulence of a flow of a turbine engine, notably of a turbine engine compressor includes a first housing with a first pressure sensor and a first inlet, a second housing with a second pressure sensor and a second inlet inclined relative to the first inlet, and a temperature sensor. The system is configured to calculate at least two orientation components of the velocity of the flow on the basis of the pressure sensors and the temperature sensor. The inlets are disposed at the vane foot, on the leading edge at the level of an internal shell.

Abstract: 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.

Abstract: A compressor for a turbine engine includes an annular row of blades including a first blade with an extrados surface and a second blade with an intrados surface; and a connecting surface joining the intrados surface to the extrados surface, the connecting surface having a bulge of volume VB and a recess of volume VC, the bulge and recess forming a contouring 3D. The volume of the bulge VB is greater than four times the volume VC of the recess, the connecting surface includes a flat or axisymmetric band circumferentially separating the bulge from the recess, and the bottom of the recess being formed against the intrados surface.

Abstract: A turbine engine compressor blade includes a leading edge, a trailing edge, a suction surface, and a pressure surface. In addition, the blade includes at least one irregularity in the form of a projecting protuberance of the suction surface or the pressure surface or in the form of a recess nested in the suction surface or the pressure surface. The irregularity may have a direction of longest dimension substantially parallel to the leading edge or substantially axial.

Abstract: An aircraft turbojet low-pressure compressor includes a sealing device for a turbine engine. The device includes an internal stator shroud connected to an external casing via an annular row of stator vanes and a rotor with two brush seals arranged inside the shroud and respectively upstream and downstream. The internal shroud includes a frustoconial internal surface, which is associated with each brush seal. Each frustoconical surface lies flush with and is inclined with respect to the bristles of the brush seal, so as to seal with the said brush. A method for assembling a sealing device is also proposed.

Abstract: The present application relates to an axial turbine engine with a turbine, a low-pressure compressor, and a fan. The turbine drives the compressor and the fan via a plurality of epicyclic reducing gears. The rotor of the compressor includes upstream a disc with a row of vanes and downstream a drum with a plurality of rows of vanes. The disc and the drum are connected to different parts of the epicyclic reducing gears, so as to be counter-rotating. The turbine engine includes a row of stator vanes arranged between two rows of rotor blades of the drum of the compressor, so as to straighten the flow and improve the efficiency. The turbine engine includes at least two epicyclic reducing gears, so that the vane rows of the drum turn faster than that downstream, which itself turns faster than the fan.

Abstract: The invention concerns a support casing (24) for a rotor (12) of a turbine engine such as a ducted fan turbojet engine used for propulsion of an aircraft. The casing (24) comprises: an outer annular wall (38) with an inner annular surface (44); an inner hub (40) able to support the rotor (12) of the axial turbine engine and comprising an outer annular surface (42); an annular passage (46) between the annular wall (38) and the inner hub (40); an annular row of arms (48) passing radially through the annular passage (46). Each arm (48) of the casing (24) comprises an orifice (50) arranged in the annular passage (46) radially at the level of one of said annular surfaces (42; 44). Inserts are fitted to the orifices (50) to control the flow passing through.

Abstract: The present application proposes an axial turbomachine compressor comprising a rotor with at least one annular row of rotor blades, a stator casing surrounding the row of rotor blades, the casing including a device for generating counter-vortexes. During operation of the compressor, the movement of the blades creates leakage vortexes at the blade tip. The generating device in turn injects counter-vortexes rotating in the opposite direction to the leakage vortexes in order to counter the leakage vortexes. This improves the surge margin of the compressor. The present application also provides a method for controlling the stability of a turbomachine compressor by counter-vortex injection.

Abstract: A system for controlling a variable orientation vane of a turbomachine compressor, for example a low-pressure compressor of a turbojet engine. Such a vane is also known as a variable stator vane. The system comprises a support, an orientable vane that is movable in rotation relative to the support and that comprises a lever for controlling the orientation of the orientable vane (26), and a magnetic field source that defines an air gap with the lever. When the source is powered electrically, it forms an electromagnet attracting the lever by induction such that the orientable vane changes orientation.

Abstract: In an axial turbomachine low-pressure compressor, the compressor, or booster, includes a first vane with an intrados surface, a second vane with an extrados surface, an inter-vane passage with a guiding surface, which connects the intrados surface to the extrados surface delimiting an annular vein, and which has a bulge and a recess forming a so-called 3D contouring geometry, or non-axisymmetric surface. The guiding surface includes an axially symmetrical or substantially flat zone, which axially extends through the inter-vane passage so as to separate the recess from the bulge.

Abstract: A low-pressure compressor for an axial-flow turbomachine, such as a jet engine, includes an annular row of vanes and a between-vanes passage with a connecting surface that links the pressure surface of a first vane to the suction surface of a second vane of the row. The connecting surface includes a main protuberance which includes a first elevation and a second elevation that are spaced apart from one another.

Abstract: A compressor for a turbine engine includes an annular row of blades including a first blade with an extrados surface and a second blade with an intrados surface; and a connecting surface joining the intrados surface to the extrados surface, the connecting surface having a bulge of volume VB and a recess of volume VC, the bulge and recess forming a contouring 3D. The volume of the bulge VB is greater than four times the volume VC of the recess, the connecting surface includes a flat or axisymmetric band circumferentially separating the bulge from the recess, and the bottom of the recess being formed against the intrados surface.

Abstract: A rotor, in particular a drum of a low-pressure compressor of a turbojet aero engine, is disclosed. The rotor includes an outer annular wall delimiting a primary annular flow of the turbine engine, sealing devices with two rubbing strips or annular ribs formed on the wall. The rubbing strips cooperate by abrasion with inner shrouds. In addition, the annular wall includes rows of intake orifices for leakages which are arranged between each pair of rubbing strips in order to aspirate the recirculation leakages there. A plenum for leakages is formed inside the rotor by means of a composite partitioning, then evacuates the parasitic flow downstream of the turbine engine via the central shaft.