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: The present application relates to an axial turbomachine gaseous flow-guiding element, such as a compressor inner shroud or vane. The element includes a plasma generator including: a layer of dielectric material with a guiding surface in contact with the gaseous flow, a first electrode placed in the guiding surface, and a second electrode electrically isolated from the first electrode by means of the dielectric layer. The plasma generator drives the gaseous flow along the guiding surface from the first electrode to the second electrode and includes a third electrode covered by the dielectric layer and electrically connected to the second electrode, so as to participate in the generation of the plasma in combination with the first electrode and the second electrode, the second electrode being closer to the guiding surface than the third electrode.

Abstract: The invention relates to an axial-flow turbomachine blading comprising a semi-circular row of blades extending radially. Each blade comprises a airfoil and lateral branches or lateral legs. Each branch has an end joined to the airfoil and an end opposite the airfoil. The blading comprises a first shroud segment connected to at least two opposite ends of two adjacent branches of two adjacent blades and/or of the same blade so as to connect the airfoils to the shroud segment via the lateral branches of the airfoils. The blading also comprises a second shroud segment concentric with the first shroud segment. The segments are connected to one another via the airfoils and the branches. The blading forms a rigid casing and is provided by additive manufacturing with a titanium-based powder.

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 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: An assembly for a turbojet, the assembly comprising: an outer shroud and an inner shroud that are concentric, the inner shroud being segmented and comprising circumferential clearances between the segments thereof; an annular row of stator vanes connecting the inner shroud to the outer shroud; a drive with a reduction ratio, which drive is intended to be coupled to a fan; and a circuit for cooling and for lubricating the drive, which 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 system for active control of radial gap around an annular row of rotor blades of a turbine engine, notably rotor blades of a low-pressure compressor of an aircraft turbojet engine. The system comprises an annular row of rotor blades; an outer casing around the annular row of rotor blades; a radial gap between the rotor blades and the outer casing; an oil circuit which is suitable for recovering the calories from a reduction gear box such as a planetary gear train which drives the fan. The oil circuit includes an expansion module which is configured to be expanded by the calories recovered from the oil. The expansion module is placed inside the outer casing so as to modulate its diameter around the rotor blades.

Abstract: A blade of a compressor or of an axial-flow turbomachine. The blade comprises a joining airfoil, or main airfoil, extended by branches, or auxiliary branch airfoils. The branches are adjacent and extend the joining airfoil radially. The branches form a row over the circumference and comprise faces facing one another. The branches comprise joining edges connected to the joining airfoil and are coincident or joined so as to form a circumferential material continuation between the adjacent branches. The leading edges and the trailing edges of the branches are a continuation of the leading edges and the trailing edges of the joining airfoil. The branch airfoils offer more anchoring points for a blade. The overlap of the branch airfoils strengthens the edges of the joining airfoil, which can thus be formed thinner. The blade comprises a fixing platform connected to the branch airfoils.

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: 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: The invention relates to a blade of an axial turbomachine comprising an airfoil which extends radially and a first set of branches or divisions, which radially extend a radial end of the airfoil, and a second set of branches which radially extend the other of the radial ends of the airfoil and which are offset over the circumference of the turbomachine. The branches extend axially along the entire length of the airfoil. The sets have different numbers of branches. The branches further occupy the channel of the turbomachine and further counteract debris in the event of intake.

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 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.

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 relates to a blade of an axial turbomachine comprising an airfoil which extends radially and a first set of branches or divisions, which radially extend a radial end of the airfoil, and a second set of branches which radially extend the other of the radial ends of the airfoil and which are offset over the circumference of the turbomachine. The branches extend axially along the entire length of the airfoil. The sets have different numbers of branches. The branches further occupy the channel of the turbomachine and further counteract debris in the event of intake.

Abstract: The invention relates to a blade of a compressor or of an axial-flow turbomachine. The blade comprises a joining airfoil, or main airfoil, extended by branches, or auxiliary branch airfoils. The branches are adjacent and extend the joining airfoil radially. The branches form a row over the circumference and comprise faces facing one another. The branches comprise joining edges connected to the joining airfoil and are coincident or joined so as to form a circumferential material continuation between the adjacent branches. The leading edges and the trailing edges of the branches are a continuation of the leading edges and the trailing edges of the joining airfoil. The branch airfoils offer more anchoring points for a blade. The overlap of the branch airfoils strengthens the edges of the joining airfoil, which can thus be formed thinner. The blade comprises a fixing platform connected to the branch airfoils.

Abstract: The invention relates to an axial-flow turbomachine blading comprising a semi-circular row of blades extending radially. Each blade comprises a airfoil and lateral branches or lateral legs. Each branch has an end joined to the airfoil and an end opposite the airfoil. The blading comprises a first shroud segment connected to at least two opposite ends of two adjacent branches of two adjacent blades and/or of the same blade so as to connect the airfoils to the shroud segment via the lateral branches of the airfoils. The blading also comprises a second shroud segment concentric with the first shroud segment. The segments are connected to one another via the airfoils and the branches. The blading forms a rigid casing and is provided by additive manufacturing with a titanium-based powder.

Abstract: The present application relates to an axial turbomachine gaseous flow-guiding element, such as a compressor inner shroud or vane. The element includes a plasma generator including: a layer of dielectric material with a guiding surface in contact with the gaseous flow, a first electrode placed in the guiding surface, and a second electrode electrically isolated from the first electrode by means of the dielectric layer. The plasma generator drives the gaseous flow along the guiding surface from the first electrode to the second electrode and includes a third electrode covered by the dielectric layer and electrically connected to the second electrode, so as to participate in the generation of the plasma in combination with the first electrode and the second electrode, the second electrode being closer to the guiding surface than the third electrode.

Abstract: The present invention relates to a pressure seal (7) of a turbomachine stator, said seal (7) comprising a first abradable surface opposite a rotor portion of the turbomachine and a second surface in contact with an inner shroud (3) of the stator, said seal (7) comprising a plurality of component units (10), each component unit (10) having, on its first abradable surface, a circumferential step (9) creating an obstacle in the circumferential direction of the inner shroud (3).

Abstract: The present invention relates to a pressure seal (7) of a turbomachine stator, said seal (7) comprising a first abradable surface opposite a rotor portion of the turbomachine and a second surface in contact with an inner shroud (3) of the stator, said seal (7) comprising a plurality of component units (10), each component unit (10) having, on its first abradable surface, a circumferential step (9) creating an obstacle in the circumferential direction of the inner shroud (3).