Abstract: A vibration damper between outer platforms of adjacent composite-material blades of a turbine engine rotor wheel, the damper including a peg having a substantially spherical portion at each of its two ends, and two hollow inserts for being received in tangential cavities that are provided facing each other in outer platforms of two adjacent composite-material blades of a turbine engine rotor wheel, each spherical portion of the peg coming to bear with point contact against the inside of one of the inserts.

Abstract: A turbine engine, including a final centrifugal compressor stage associated with a diffuser for supplying air to a combustion chamber, and a ventilation element for ventilating a high-pressure turbine wheel, is disclosed. The ventilation element includes an injection element for injecting air onto the wheel and a take-up element for taking up a flow for cooling the impeller of the compressor, where the take-up element has a labyrinth seal mounted at the outlet of the injection element, and the air outlet orifices are installed between the injection element and the labyrinth seal and leading upstream of the turbine wheel.

Abstract: A turbomachine rotor wheel having composite material blades is disclosed. The wheel includes a slideway disk having a circumferential groove at its outer periphery, the groove having an opening that is narrower than its bottom, with a plurality of composite material blades each having a blade root engaged in the groove of the slideway disk and presenting a longitudinal section adapted to the longitudinal section of the slideway disk in order to be held radially therein by co-operating shapes, and an orientation disk fastened coaxially around the slideway disk and including, in its outer periphery, a plurality of axial slots each of which is open at one end. Each slot presents a cross-section that is adapted to the cross-section of a corresponding blade root so as to hold it angularly by co-operating shapes.

Abstract: A gate valve for a system controlling cooling in a turbomachine is disclosed. The gate valve includes a gate mounted so that it can pivot about an axis between a position in which it shuts off an air passage orifice and a position in which it opens the orifice, and a rotating device for rotating the gate about the axis. The rotating device includes two superposed coaxial rotary members collaborating with one another via cam surfaces designed such that rotating the first rotary member from the shut-off position causes a translational movement of the second rotary member and of the gate along the axis of rotation followed by a rotation of this member and of the gate about this axis.

Abstract: This system for guiding a drone during the approach phase to a platform, particularly a naval platform, with a view to landing the same, is characterized in that the platform is equipped with a glide slope indicator installation emitting an array of optical guide beams over an angular sector predetermined from the horizontal, and in that the drone is equipped with a beam acquisition camera connected to image analysis means and to computing means of orders for commanding automatic piloting means of the drone to cause it to follow the guide beams.

Abstract: A method of fabricating a composite-material turbomachine blade having platform incorporated therein includes separately making a blade platform element that comprises rigidified fiber reinforcement and that has a shape close to that of a platform of the blade to be fabricated, with at least one opening being formed through the platform element. The opening has the shape of the profile of the airfoil of the blade to be fabricated. Three-dimensional weaving is used to make a flexible fiber blank as a single piece that includes blade airfoil-and-root preform portions. The platform element is engaged, via its opening, on the flexible fiber blank while deforming it so as to assemble the platform element with the fiber blank. The fiber blank is shaped together with the platform element assembled thereon to obtain a blade preform of a shape that is close to the shape of the blade to be fabricated. After shaping, the blade preform is consolidated in its shape in a shaper.

Abstract: The invention relates to an aircraft support pylon (1), designed to attach a turbomachine housed in a nacelle (7) to an aircraft fuselage (8). The pylon comprises a cooler (21), located substantially in the vicinity of a leading edge (12) of the pylon, and able to be heated by a hot fluid, said hot fluid intended to be cooled by heat exchange with cold air outside said cooler. The pylon also comprises at least one air evacuation duct (22), at the exit from the cooler (21) and extending longitudinally in an internal volume of the pylon to a trailing edge (11) of the pylon.

Abstract: An aircraft includes a heat exchanger between, on the one hand, a first fluid that flows into the narrow pipes (22) of the exchanger and that is able to contain solid elements that can melt, and, on the other hand, a second fluid to be cooled. The heat exchanger includes a device for preheating the first fluid upstream from the pipes (22) of the exchanger, a number of pipes arranged in at least one plane that is secant to the direction of flow of the first fluid, whereby the spacing between two adjacent pipes is suitable for retaining the solid elements of large dimensions and for allowing them to pass only when they have melted at least partially.

Abstract: The invention relates to a door (20) for an aircraft engine nacelle (6), comprising an opening leaf (22), locking means (32, 36) with which the opening leaf may be locked in the closed position, as well as means (28) for setting the opening leaf (22) into motion with which the latter may be displaced from the closed position to the open position. According to the invention, said locking means have a mechanical fuse (50) intended to break under the effect of a force exerted by the opening leaf (22) and resulting from an opening gas pressure being applied on the opening leaf, with a value greater than or equal to a predetermined value and causing displacement of the opening leaf from the closed position to the open position after breakage of the mechanical fuse (50).

Abstract: The present invention relates to a turbomachine rotor balancing system comprising a balancing flange (52) provided with through-passageways (78), the system further comprising balance weights (54) each mounted fixedly on the flange (52) by means of a first and a second screw/nut assembly (56a, 56b) having a first and a second screw (84, 84) passing through a first and a second passageway (78, 78). According to the invention, said first screw is pressing against a first tangential end side (116) of the first through-passageway, and at a distance from a second tangential end side, and the second screw is pressing against a second tangential end side (118) of the second through-passageway, and at a distance from the first tangential end side.

Abstract: A stationary vane assembly of a turbine engine including an inner casing and a high-pressure turbine guide vane assembly including at least two angular sectors including blades, the angular sectors being fixed on the casing, each angular sector including a platform and a fastening tab. The fastening tab protrudes from the platform radially inwardly. The casing includes a first annular groove receiving the fastening tabs, radial retaining pins for the angular sectors being introduced into bores made in the casing and the fastening tabs, the axial retaining pins being axially stopped by a ring mounted in a second annular groove of the casing.

Abstract: A turbomachine rotor balancing system including a balancing flange provided with through-passageways, and balance weights mounted fixedly on the balancing flange by at least one screw/nut assembly passing through one of the passageways is disclosed. For each of the passageways, the flange has on its first face a first recess passing via the radial summit of the passageway. This first recess being deprived of contact with the screw head pressing on this first face. The flange has on its second face a second recess passing via the radial summit. The second recess being deprived of contact with the balance weight pressing on this second face.

Abstract: A stationary vane assembly for a turbine engine including an inner casing and at least two angular sectors forming a high-pressure guide vane assembly, each angular sector including a fastening tab for fastening the angular sector on the casing. The stationary vane assembly including a ring of abradable material fixed on the casing, on the side of a mobile vane assembly using a plurality of screw-nut assemblies, the stationary vane assembly also including radial retaining pins for the angular sectors introduced into first bores formed in the fastening tabs and in the mounting clip, the bores having axes substantially parallel to the longitudinal axis, the mounting clip being fixed on the casing by the same screw-nut assemblies as those fixing the ring of abradable material support on the casing.

Abstract: The invention relates to method characterised in that said method comprises steps of acquiring movements, calculating the mean position, calculating position predictions, and calculating minimum grid (5) movement speeds, and a step of acquiring the position of the drone (4) so that, if the drone cannot follow the movements of the grid and the movements of the grid are small, i.e. smaller than the radius of the latter, it is possible to apply a landing strategy by monitoring the mean position of the grid, and if the movements of the grid are large, i.e. larger than the radius of the grid, it is possible to apply a landing strategy by positioning at the minimum speeds of the grid, and if the drone (4) can follow the movements of the grid (5) and the movements of the grid are small, i.e. smaller than the radius of the grid, it is possible to apply a landing strategy according to the mean position of the grid, and if the movements of the grid are large, i.e.

Abstract: The invention relates to an aircraft propeller (1) comprising a turbomachine (8) housed in a nacelle (10) and a cooler (45) capable of being traversed by a hot fluid, which is to be cooled by thermal exchange with cold air external to the cooler. The propeller (1) comprises an air stream (13) (13b) capable of directing pressurized air towards an air duct (20) realized between an outer wall (6) and an inner wall (60) of the nacelle (10). The cooler comprises volumetric cooling means (14) located in the air duct (20) and surface cooling means (15), unrelated to said volumetric cooling means, located at an outside wall (6, 101) of the aircraft.

Abstract: The invention relates to an aircraft propeller (1) comprising a turbomachine (8) housed in a nacelle (10) and a cooler (14) capable of being traversed by a hot fluid, which is to be cooled by thermal exchange with cold air external to the cooler. The propeller (1) comprises an air vein (13) (13b) capable of directing pressurized air towards an air duct (20) realized between an outer wall (6) and an inner wall (60) of the nacelle (10). The cooler (14) comprises a first cooling means, called first surface cooling means (145), on a first surface (141), arranged at the outer wall (6) of the propeller nacelle (10) and a second surface cooling means (146), on a second surface (142) arranged at a wall (23) of the air duct (20). The invention also concerns an aircraft equipped with such a propeller.

Abstract: Turbine engine, including a final centrifugal compressor stage associated with a diffuser for supplying air to a combustion chamber, and ventilation means for ventilating a high-pressure turbine wheel, including injection means for injecting air onto the wheel and take-up means for taking up a flow for cooling the impeller of the compressor, wherein these take-up means comprise a labyrinth seal mounted at the outlet of the injection means and the air outlet orifices installed between the injection means and said labyrinth seal and leading upstream of the turbine wheel.

Abstract: A method for bonding a first part made of a composite to a second part to a predetermined dimension is disclosed. The method includes placing an insert into the first part, a first surface of the insert protruding beyond the predetermined dimension; machining the insert until the first surface is at the predetermined dimension; placing the second part on the insert; and fastening the second part via the insert. The use of an insert makes it possible to produce, with high precision, a reproducible assembly, which can still be dismantled.

Abstract: A device attaching an afterburner to a turbojet casing. The device comprises a transverse plate for attachment to the casing and at least two side brackets for supporting two legs of the afterburner. The brackets and the two legs extend longitudinally. The afterburner includes a flame-arrestor arm with the two legs.

Abstract: An exhaust cone of composite material. The exhaust cone is mounted to extend an adapter ring and it is held by an elastically deformable element that acts axially.