Abstract: An assembly including a sector of fixed blades for turbomachine distributor including at one circumferential end a protuberance and at another circumferential end a contact surface, such that the protuberance of a sector arrives opposite the contact surface of another adjacent sector, a detachable cap including a surface with anti-wear material, configured to be engaged detachably on the protuberance of the sector, such that the anti-wear material is in contact with the contact surface of the other sector.

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: The invention relates to an aircraft undercarriage provided at its bottom end with a rocker arm (2) carrying at least two axles (4, 5), including one in front of a hinge between the rocker arm and the undercarriage and the other behind said hinge. According to the invention, one of the axles is fitted with one or more braked wheels (6) and the other axle is fitted with one or more motor-driven wheels (8).

Abstract: A support case for a turbomachine ring sector includes an annular body around a longitudinal axis of the case, and a downstream edge for attachment of the ring sector. The downstream edge extends from the body radially towards the inside of the case. The downstream edge extends circumferentially along a junction region between the body and a fixed vanes stage. The junction region includes a central portion and two lateral portions located on each side of the central portion, along a circumferential direction of the case with respect to its longitudinal axis. The central portion projects radially outwards from the case with respect to the lateral portions.

Abstract: The invention relates to a lubrication device for a turbine engine, comprising an oil intake pipe (23) provided with a pump (24) for supplying oil and control means (30) located downstream from the supply pump (24), a supply pipe (26) intended for supplying oil to a member to be lubricated and a recirculation pipe (27) connected upstream from the supply pump (24), the control means (30) making it possible to direct all or part of the flow of oil from the intake pipe (23) towards the supply pipe (26) and/or towards the recirculation pipe (27).

Abstract: A compressor shroud for an aircraft turbine engine, the shroud being arranged between two bladed rotary wheels and radially in line with a deflector, the shroud including a sealing device including one or more sealing elements, including a downstream end sealing element on which, projecting downstream, an air entraining and diverting structure is provided, designed to axially divert the discharge air issuing from the end sealing element.

Abstract: The invention relates to a method for guiding an aircraft along a reference path, said method comprising: a) when the altitude of the aircraft is greater than a threshold, estimating (E1) a relative location of the aircraft in relation to a taxiing starting point using a map of the platform and reference points on the ground, b) when the altitude of the aircraft is less than said threshold and before the aircraft is located at the taxiing starting point, estimating (E2) a relative location of the aircraft in relation to the taxiing starting point using data relating to the absolute location of the aircraft and the last relative location estimated at step a), c) when the aircraft is located at the taxiing starting point, guiding (E3) the aircraft on the basis of a location of the aircraft relative to the reference path as estimated using data relating to a set of indicators on the ground.

Abstract: The invention relates to a nacelle (1) for an aircraft turbojet, the nacelle comprising a frustoconical radially outer annular wall (13), a pylon (9) extending radially outwards from said annular wall, the pylon (9) being designed to fasten the nacelle (1) to a fixed portion of the airplane, said nacelle being characterized in that it is provided with at least one projecting zone (22) extending radially outwards from the frustoconical outer wall (13), said projecting zone (22) being situated in the vicinity of a downstream annular edge (14) of said outer wall (13) and in the vicinity of the pylon (9).

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 disclosure provides an iterative method for draping fiber plies to produce a preform for a composite wall with an organic matrix. The method comprises the sequential stacking of woven fiber plies on a template. The plies are covered with reference markings, particularly with tracer threads forming a grid pattern, the positions of the threads being indicated by means of illuminated marks projected by a laser. The reference markings of the plies are moved in turn until they coincide with their corresponding marks, thereby ensuring the correct positioning of the corresponding ply. The stack contains plies with threads orientated at ?45°/+45° and at 0°/90°. The illuminated marks change with each type of ply so as to project dedicated indications onto the plies. The present disclosure also provides a draping installation and a composite wall of a casing of an axial turbine engine, or of a motor vehicle structure.

Abstract: A turbine ring assembly includes a ring support structure and a plurality of CMC ring sectors forming a turbine ring. Each ring sector is K-shaped in radial section, with tabs extending from the outside face of the annular base over end portions of the annular base, the tabs and the end portions of each ring sector being held respectively facing tabs and end portions of ring sectors that are adjacent in the ring. The turbine ring assembly has a plurality of rigid gaskets, each extending axially between adjacent ring sectors, and resilient holder devices exerting a force suitable for holding the gaskets in contact with the end portions or the tabs of two adjacent ring sectors.

Abstract: The invention relates to a fairing (37) for a mixer (28) of a nozzle (26) of a dual-flow turbomachine (20), said mixer being of an overall annular shape and extending along a longitudinal axis (24) of the turbomachine, said mixer comprising an upstream part (29) provided with a flange (31) extending radially toward the outside of the mixer, intended to be fastened to an exhaust casing (21) of the turbomachine, and a downstream part (33) forming a flow mixing area, the fairing (37) being of an overall annular shape and configured to extend around the upstream part (29) of the mixer (28) and to be attached to the flange (31) of said upstream part (29), the fairing being without means for fastening to the downstream part (33) of the mixer (28), in such a way as to connect the fairing (37) to the mixer (28) at a distance from the flow mixing area.

Abstract: The invention relates to tooling (24) for machining an annular groove of a turbine engine casing, wherein said tooling (24) comprises a machining tool (25), a baseplate (33), first means of positioning (28) the machining tool (25) in relation to the baseplate (33) along a first axis (Y) forming a radial axis, second means of positioning (30) the machining tool (25) in relation to the baseplate (33) along a second axis (X) perpendicular to the first axis (Y), wherein said second axis (X) extends along the axis of the groove and of the annular casing and third means of positioning capable of positioning the baseplate (33) axially and radially in relation to the groove of the casing.

Abstract: The invention relates to a suspension pylon comprising: at least one arm (3) for suspending a propulsion unit (4) from the fuselage (2) of an airplane (1), an oil exchange circuit (70) configured to be connected, on the one hand, to an oil pump (5) in the fuselage (2) and on the other hand, to the propulsion unit (4) the arm (3) whereof provides for suspension, said circuit (70) comprising a feed line (72) and a return line (74), which both extend inside said arm (3), an intermediate shaft (6) which extends in said arm (3), said shaft being configured, on the one hand, to be driven by the propulsion unit (4) the arm (3) whereof provides for suspension and, on the other hand, to drive the oil pump (5) in the fuselage (2), wherein the arm (3) has a structure adapted for cooling the oil exchange circuit at the arm (3).

Abstract: A method of regulating a temperature (TM) associated with a heat exchanger assembly (10) of a turbine engine (1), the method comprising, in a single cycle: measuring (500) the temperature (TM) of an air stream at the outlet from a heat exchanger (14); receiving a setpoint temperature (TC) for the air stream at the outlet from the heat exchanger (14); estimating (510) a theoretical temperature (TT) for the air stream at the outlet from the heat exchanger (14) as a function of an estimate of the position of a shutter of a controlled valve (15) bleeding off a cooling air stream for the heat exchanger (14); determining a correction current (ICO) from the difference (?T1) between the measured temperature (TM) and the theoretical temperature (TT); and determining (530) a control current (IC) for the shutter from the difference (?T2) between the measured temperature (TM) and the setpoint temperature (TC) and while taking account of the correction current (ICO) determined during the preceding cycle, the position of t

Abstract: A monitoring system for monitoring a piece of aircraft equipment, the system comprising a master electronic module and a slave electronic module, each fitted with first wireless communication means, the slave electronic module further comprising measurement means for the purpose of taking measurements of a parameter of the aircraft equipment, and power supply means making the slave electronic module independent in terms of energy, the master electronic module further comprising detector means adapted to detect a stage of flight in which the aircraft is to be found, and control means for acting via the first wireless communication means to control the measurement means of the slave electronic module so as to adapt the measurements taken by the measurement means to the detected stage of flight.

Abstract: Described is a breakover mechanism for a passenger seat that includes a frame member with a rear portion attached to two opposing sides of a seat back of the passenger seat wherein the frame member is at least partially disposed under a seat pan of the passenger seat, at least one moving portion attached to a forward portion of the frame member, a carriage portion fixedly attached to the seat pan; and a single point mechanism attached to the carriage portion. Once a threshold loading condition occurs, the single point mechanism changes states to facilitate movement of the moving portion relative to the carriage portion.

Abstract: The invention relates to a bar spacer (10) for a braked aircraft wheel, the bar being for fitting to a rim (1) of the wheel in order to drive the rotor disks of the brake in rotation, the spacer being for interposing between the bar and the rim, the spacer comprising two bearing blocks (11) connected together by a core (15) that is pierced to pass a screw for fastening the bar to the rim. According to the invention, the bearing blocks (11) of the spacer include voids defining a void fraction of at least 30% of the volume of the bearing blocks.

Abstract: The invention relates to a blade (16), in particular a fan blade, for resurfacing an abradable coating (26) on an internal surface of a turbine engine case (18) intended to surround the blade. The comprises a root (20) and an airfoil (24) having a tip opposite the root, said root comprising a housing (32) for the storage or passage of a polymerisable refill resin. The airfoil comprises longitudinal channels (34), the first ends of said channels being connected to the housing and the opposite ends opening at the aforementioned tip.

Abstract: A pitch actuation system for a turbomachine propeller, comprising an actuator with a movable portion configured to be connected to blades of the propeller in order to rotate them relative to the blade-pitch setting axes, wherein the actuator is an electromechanical actuator and comprises: blade-pitch control having a transmission screw rotated about an axis; a nut through which said transmission screw passes; and a member configured to engage with the blades in order to move the blades, wherein the nut is connected to the member via a decoupler configured such that a translational movement of the nut causes a translational movement of the member, but rotational movement of the member does not cause rotational movement of the nut.