Abstract: The present invention relates to an abradable coating for a turbine engine part, characterized by the fact that said coating includes a layer made of an abradable material, the surface asperities of which are filled with thermally bonded ceramic grains, forming a smooth, free surface having a low roughness.

Abstract: There is provided a seat break over device. The seat break over device features linkage members that connect the seat back to the quadrant arm. In the static condition, the linkage member acts as a rigid member, constraining rotation of the seat back relative to the quadrant arm. Upon impact of a certain inertial force, the linkage member interacts with a weighted member. The weighted member may include a suspended weight or mass that pivots about a horizontal axis. During a gravity load of a certain amount (e.g., a 16 g event), inertial loads act on the weighted member/hammer, which transfers a force onto the linkage member. This force moves the linkage member into an unstable mode, allowing forward rotation of the seat back with respect to the quadrant arm.

Abstract: The invention concerns a method for generating auxiliary power in an aircraft, comprising the step consisting of: starting up an auxiliary power unit (6) of the aircraft by supplying compressed air to the auxiliary power unit (6) from a supercharger (7), and transferring non-propulsive energy from the auxiliary power unit (6) to the aircraft. The invention also concerns a system (5) for generating auxiliary power in an aircraft and an aircraft implementing such a method.

Abstract: The invention relates to a system for controlling variable pitch blades for a turbine engine, comprising an annular row of variable pitch blades extending about an axis (A) and each comprising a blade connected at the radially outer end thereof to a pivot (20) that defines a substantially radial axis of rotation of the blade and which is connected by a lever (34) to control means (40a, 40b) extending about said axis. The invention is characterized in that said control means include first links (40a) supported by said pivots and second links (40b) extending between said first links, said first and second links extending substantially along a same circumference of said axis and being connected to one another and to actuation means (56).

Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: A nacelle for a propulsion unit includes at least one scoop opening to the external air, and a canal supplying external air from the scoop to two ducts which separate downstream of the canal, forming a dividing lip, and at least one exchanger to cool a stream of fluid entering the exchanger. The exchanger is arranged in the region of and in contact with the separation lip. The external air circulates from the scoop towards the two ducts making it possible to cool the stream of fluid entering the exchanger.

Abstract: When cold and in the non-coated state, a turbine blade includes an aerodynamic profile that is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: A method for modeling an offset portion of a blade of a non-ducted propeller is provided. The method includes: parametrizing a C1-class curve representing a deformation of the blade characterizing the offset, according to a position along a section at a given height in the blade, the curve intersecting consecutively through a first bend control point, a central control point, and a second bend control point, the first and second bend control points defining the extent of the blade section, the parametrization being implemented according to a first deformation parameter defining the abscissa of the central control point, a second parameter of deformation defining the ordinate of the second bend point, and a third deformation parameter defining the angle of the tangent to the curve at the second bend control point; optimizing one of the deformation parameters; and plotting the values of the optimized parameters on an interface.

Abstract: A device for cooling a rolling bearing (112), in particular for a turbine engine. The device includes an annular bearing support (124) having a substantially cylindrical wall (134) of which an inner peripheral face (136) is configured to surround an outer ring (122) of the rolling bearing. The wall is able to dispense cooling fluid via a fluidic network for circulating the fluid and for supplying annular cooling grooves (144) of the bearing with fluid.

Abstract: A turbomachine turbine blade includes a suction face, a pressure face, a leading edge and a trailing edge as well as a squealer at the tip thereof. The squealer tip includes at least one internal rib that extends from a point of attachment of the rim of the squealer on the suction-face side to a point of attachment of the rim of the squealer on the pressure-face side, and includes a portion for absorbing the load of leakage flows which extends from the suction face and a deflector-forming portion which extends the load absorber portion with an inflection and guides the leakage flows toward the pressure face.

Abstract: A method of regenerating a platinum bath by flow reaction, the method comprising the successive steps of: drawing off fluid from the platinum bath by means of a draw-off flow; complexing platinum by mixing together the draw-off flow and a regeneration solution flow containing platinum, mixing taking place in an intensified reactor; and feeding the platinum bath with the mixture resulting from the platinum complexing step, by means of a regenerated bath flow; all of these steps being performed as a continuous flow.

Abstract: A control assembly for a stage of variable-pitch vanes for a turbine engine generally includes an actuating ring having an annular body, a vane guidance means for connecting to the cylindrical pivots of the variable pitch vanes, a rotating means for rotating the actuating ring about a casing of the turbine engine, and a centering means for cooperating with the casing to center the actuating ring. The centering means may have at least one skid bearing against the casing. At least one of the skid or the rod may be configured to cooperate by radial sliding with an inner wall of a housing. The skid may be configured to be mobile during operation by sliding between a radially internal position and a radially external position.

Abstract: The invention relates to a guide vane located in a fan air flow (15) in a aircraft twin-spool turbomachine, the vane being made with an extrados body (32a) and an intrados body (32b) between which there is a thermal conduction matrix (80). Furthermore, the attachment means between the two spools (32a, 32b) are arranged outside the aerodynamic part (32) of the vane.

Abstract: The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream at the radially inner ends of the leading edges of the fan blades, divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.

Abstract: Communication system, used in a drone system, based on a communication system of the TDMA type in which the time is divided into a plurality of temporal breakdown hierarchy levels. All the hierarchy levels being suitable for complying with a plurality of constraints related to the drone system comprising a first constraint meaning that each communication between a drone and a station controlling the drone system must be compatible with a conversation audio data transmission, a second constraint meaning that the control station must be capable of controlling a predefined number of drones simultaneously and a third constraint meaning that, the drone system using transmissions by bursts, each burst must be transmitted over a stationary channel.

Abstract: A fuel supply system for a turbomachine. The supply system includes a starter circuit, at least one first hydraulic resistance and a purge circuit. The purge circuit includes a duct including an opening at the exterior of the supply system. The purge circuit is configured to make purge air flow between a starter injector and the opening through the first hydraulic resistance. The supply system includes a device for measuring a value representative of pressure, which devices are configured to measure a value representative of pressure between the starter injector and the first hydraulic resistance when purge air flows between the starter injector and the first hydraulic resistance.

Abstract: An electrical power supply device in an aircraft, the device including an enclosure, the enclosure containing at least: a fuel cell; a dihydrogen generator; a gas feed circuit connecting the dihydrogen generator to the anode of the fuel cell; an oxygen feed device feeding the cathode of the fuel cell; and a cooling circuit of the fuel cell associated with at least one heat exchanger; wherein the enclosure is mounted on an actuator system, the actuator system being configured to move the enclosure from a first position in which the enclosure is housed inside an outer wall of the aircraft, to a second position in which the enclosure projects from the outer wall.

Abstract: When in a cold state and in a non-coated state, the aerodynamic profile for an arm of a structural casing of a turbine having a hub and a shroud is substantially identical to a nominal profile determined by specific Cartesian coordinates X,Y, Zadim. The coordinate Zadim is the quotient D/H where D is a distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: A guide vane for a dual flow aircraft turbine engine includes an aerodynamic part. The aerodynamic part includes an inner duct for lubricant cooling extending in a main direction and being partly bounded by a pressure side wall and a suction side wall of the vane. The aerodynamic part is embodied as a single piece, and includes heat transfer fins arranged in the duct expanding substantially parallel to a direction of the duct. The fins are spaced from each other depending on the direction of the duct and a transversal direction of the vane, so that at least some of the fins are arranged substantially staggered.

Abstract: A turbine ring assembly includes ring sectors forming a turbine ring, and a ring support structure having two annular flanges, each ring sector having a portion forming an annular base with an inner face defining the inside face of the turbine ring and an outer face from which there project at least two tabs, the tabs being retained between the two annular flanges. Each tab of the ring sectors includes a projecting portion on its face situated facing one of the two annular flanges, this projecting portion co-operating with a housing present in the annular flange. Each tab of the ring sectors includes an opening in which there is received a portion of a retention element secured to the annular flange situated facing the tab. The retention element is made of a material having a thermal expansion coefficient that is greater than that of the material of the ring sectors.