Abstract: A method for manufacturing a turbine engine vane a root connected to a blade extending in a longitudinal direction includes the steps of providing a root; and providing mold with a first cavity and a second cavity that together define a recess in which the vane is formed. The recess includes a first space in which the blade is formed and a second space in which the root (1), c) is formed. The method further includes the steps of providing aluminum strips; positioning a fibrous reinforcement; arranging the vane root in the second space; and injecting a foam comprising aluminum or injecting an aluminum alloy into the first space of the recess of the mold such that the foam impregnates the fibrous reinforcement.

Abstract: A method and system control the thrust of an aircraft turbomachine having a high bypass ratio by direct action on a variable-pitch system. The variable-pitch system varies the pitch of the vanes of a stator of a low-pressure compressor for the open-loop control of the thrust of the turbomachine. The method also provides closed-loop control of the pitch of the blades of a propeller based on a rotational speed of the propeller.

Abstract: Disclosed is a cart (2) for transporting an aircraft engine, comprising a base frame (4) equipped with wheels (6) for riding on a floor; at least two engine arms (14, 16) extending horizontally and movable vertically relative to the base frame, structured and designed for supporting the aircraft engine; and at least two adapter arms (20) extending horizontally at a higher level than the at least two engine arms, structured and designed for supporting an adapter coupled at the top of the aircraft engine. Disclosed is also a method of transporting an aircraft engine to a test cell.

Abstract: A transonic compressor of a turbomachine comprising exactly two or exactly three annular rows of rotor blades and respectively exactly one or exactly two annular rows of intermediate stator vanes interposed between two respective rows of rotor blades, wherein the flow velocity relative to the rotor downstream of the intermediate stator vanes is less than or equal to a Mach number of 0.9 over a radial portion of the blades extending over the radially inner 40% of the blades, and less than or equal to a Mach number of 1 over a radial portion extending over the radially inner 80% of the blades, and less than or equal to a Mach number of 1.05 over a radial portion extending over the radially outer 20% of the blades.

Abstract: An internal gear pump comprising: a ring with internal teeth delimiting a cavity and a gear disposed in the cavity, the gear comprising external teeth cooperating with the internal teeth to drive the ring in rotation, remarkable in that it comprises a pinion disposed in the cavity and driving the gear in rotation. The pinion cooperates with the external toothing of the gear. In an alternative embodiment, the gear comprises an internal toothing which cooperates with the pinion.

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 turbine engine heat exchanger for exchanging heat between a first fluid and a second fluid includes a reference axis, a network of tubular meshes having a plurality of meshes each of which is formed, successively in a reference direction, of at least two curvilinear branches, called anterior branches, of a junction where the two anterior branches meet, and of at least two curvilinear branches, called posterior branches, diverging from the junction, wherein the first and second fluid have a respective general direction of flow, and the general direction of flow of the first fluid is parallel to the general direction of flow of the second fluid. The present disclosure also concerns a turbine engine comprising the heat exchanger and a manufacturing method for manufacturing the heat exchanger.

Abstract: A turbine engine heat exchanger for exchanging heat between a first fluid and a second fluid includes a reference axis, a network of tubular meshes having a plurality of meshes each of which is formed, successively in a reference direction, of at least two curvilinear branches, called anterior branches, of a junction where the two anterior branches meet, and of at least two curvilinear branches, called posterior branches, diverging from the junction, wherein the body of the heat exchanger is of cylindrical shape. The present disclosure also concerns a turbine engine comprising the heat exchanger and a manufacturing method for manufacturing the heat exchanger.

Abstract: A matrix for exchanging heat between a first fluid and a second fluid, in particular for an air-oil application in a turbine engine, includes an envelope defining a flow path of the first fluid and a network extending into the flow path and in which the second fluid flows. Along the axis defined by the curvature of the matrix, the dimensions of the envelope vary circumferentially (T(A)) and radially (R(A)). The matrix may be used with a heat exchanger.

Abstract: A heat exchanger includes a plurality of fins arranged as a network and delimiting corridors, and an envelope having an internal wall and an external wall, the internal and external walls delimiting between them a channel for a flow of a first fluid in a main direction, the network of fins being arranged in the channel and connected to the internal and external walls, at least one passage for a flow of a second fluid being embedded in at least one of the internal and external walls, the channel being, in the main direction, divergent and then convergent.

Abstract: The disclosed technology concerns an aircraft turbomachine part comprising a part body drilled with at least one cavity open to the outside and at least one conduit joining the cavity on the one hand and leading to the outside on the other hand. Each cavity receives a pressure sensor, and the conduit corresponds to the cavity guides the cables connected to the sensor to the outside of the part body. The part is an aircraft turbomachine vane.

Abstract: An assembly for the compressor stator of a turbomachine. The assembly comprises: a shroud, in various instances an inner shroud, that is axially divided into two parts; a pocket formed in the shroud; a bearing located in the pocket; and an orientable vane pivotably mounted in the bearing about a pivot axis. The shroud comprises an axial interface separating the parts that is axially offset from the pivot axis of the orientable vane. The invention also provides a process for assembling the assembly.

Abstract: Turbine engine assembly comprising: an external casing (28) of a low-pressure compressor (4), an annular wall (30) and an oil tank (46) with a circular chamber (48) around an axis (14) of the compressor. The wall (30) comprises an inner surface (38) delimiting an primary guide path for the flow of the compressor, and an external surface (40) radially opposite the inner surface and delimiting the internal chamber (48) of the tank (46).

Abstract: The invention relates to a method for handling a rectifier of a turbojet of an aircraft, the rectifier having an axis defining the asymmetry thereof, said method comprising a step of placing the rectifier on the rollers of a supporting structure, the structure and the rollers thereof being arranged such that the axis of the rectifier is inclined at a non-null acute angle in relation to the horizontal, and a step of controlling, maintaining, assembling, handling, storing, deburring and/or cleaning the rectifier, during which the rectifier is pivoted about the axis thereof. The invention also relates to a trolley for handling a rectifier for an axial turbojet, said trolley comprising at least two of the lower rollers that have axes inclined in relation to the horizontal of said angle and at least one upper roller, the axis of which is inclined in relation to the vertical of said angle.

Abstract: Heat exchange devices for aircraft turbine engines include a heat exchanger and an inlet scoop comprising an air intake configured for supplying the heat exchanger. The air intake of the inlet scoop is divided into several mouthpieces, each defining an air flux supplying the heat exchanger.

Abstract: A depolluted test bench (I) for a turbomachine includes a channeled air duct with an air inlet and an air outlet to allow a flow of air between the air inlet and the air outlet. The channeled air duct includes a test chamber) for testing a turbomachine, the test chamber being located between the air inlet and the air outlet. A depolluting system is positioned in the air duct to depollute an air flow generated at least in part by a test of said turbomachine in the test bench. Control means control the depolluting system to allow the depolluting system to be switched on or off when a test of the turbomachine is started or stopped.

Abstract: A tooling holds a blade during friction welding to a rotor element. The tooling includes a fixed jaw having a central part; two arms separated by a distance d for receiving at least one portion of the blade; and a fixed jaw body with bearing surfaces and a bore for receiving at least one portion of the blade. The bearing surfaces are configured to come into contact with the blade. A movable jaw includes a movable bearing surface that comes into contact with the blade, and pressure means that moves the movable jaw towards the fixed jaw body to press the blade.

Abstract: The invention relates to a tank (32) for liquid, in particular of a turbine engine such as an aircraft turbojet engine, the tank comprising an external wall (36) with an inner surface (48); an inner chamber intended for containing the liquid of the tank (32) and defined by the inner surface of the external wall; and a system (42) for measuring the level of liquid. The system comprises: a submersible portion matching the inner surface of the wall (36); a clearance (52) separating the submersible portion from the inner surface; and a liquid detector (54) configured to emit a signal towards the submersible portion through the clearance, said signal being configured to be modified in the event that liquid is present in the clearance (52). The invention likewise relates to a method for measuring a level of liquid, and an associated computer program.

Abstract: 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.

Abstract: A method of testing and/or controlling the operation of an axial turbomachine through which passes a gas stream, includes the following actions: measurement of operating parameters of the turbomachine, said parameters including pressure in the gas stream at different axial positions, and calculation of operating conditions of the turbomachine from the measured parameters and the Laplace coefficient ? of the gas passing through the turbomachine, wherein the measurement of parameters includes a measurement of the temperature of the gas stream, and the calculation of operating conditions includes a determination of the Laplace coefficient ? on the basis of the measurement of the temperature of the gas stream.