Abstract: Use is made of an organic compound of formula CHnI4-n, n being between one and three, in order to absorb ionizing radiation, preferably when its energy is greater than approximately 10 eV and more preferably still when it concerns X-rays or gamma rays. This material can be used in particular in devices for protecting from ionizing radiation and as absorbent labeling agent for radiography.

Abstract: The invention proposes a long life nozzle flap for aircraft turbojets. This flap is remarkable in that it is composed of a hollow tapered body (10) flattened in the transverse direction along straight geometric generating lines (11), the body (10) forming a thin wall (12) with an approximately constant thickness E, this wall (12) comprising a continuous inner surface with a radius of curvature equal to at least 2×E, the wall (12) being made of a refractory composite material composed of reinforcing fibers (18) embedded in a matrix also made from a refractory material, the reinforcing fibers (18) being arranged in a plurality of continuous layers of fibers (18) stacked on each other, the fibers (18a) of a layer crossing the fibers (18b) of any adjacent layer.

Abstract: A stationary ring assembly forming a rotor shroud for a gas turbine, the stationary ring assembly comprising a plurality of segments having adjacent side faces that are placed end to end with sealing means interposed therebetween, the sealing means comprising at least one axial sealing strip and at least one radial sealing strip respectively received in axial slots and in radial slots formed facing one another in the adjacent side faces of the segments, at least one end of each radial slot opening out into the corresponding axial slot, each axial slot of a segment presenting a depth that is greater than the depth of the corresponding radial slot, and the axial sealing strip presenting a width that is greater than the width of the radial strip.

Abstract: A turbine blade of a turbomachine, having at least a bottom platform connected to the base of the blade by a bottom connection zone, and a cooling circuit consisting of at least one cooling cavity, of a plurality of evacuation slots arranged along the trailing edge of the blade, said blade having a bottom evacuation slot that is disposed near the blade base, the bottom evacuation slot including an end wall provided with an opening that opens into the cavity, a setback wall, a bottom wall disposed beside the blade base, a bottom edge formed between the setback wall and the bottom wall, and a bottom shoulder formed between the bottom wall and the bottom connection zone, both the bottom edge of the bottom evacuation slot and the bottom shoulder of the bottom evacuation slot having respective right sections of substantially rounded shape, thereby avoiding any protruding angles between the opening of said slot and the bottom connection zone.

Abstract: In a layout with two concentric shaft lines (15, 16), one of which projects beyond the other, a seal support sleeve (38) placed between them has a conical shape and bears on a spacer (40) of the external shaft line, so that an assembly nut (19) of the forward end is accessible from the back passing through the space (43) between the two shaft lines (15, 16), so that the nut can be unscrewed and the external shaft line can be disassembled. The invention may be applicable to front bearings of gas turbine high pressure shaft lines and makes it possible to disassemble the high pressure body from behind, without needing to disassemble equipment such as a fan on the front of the machine. This does not hinder the installation of a movement control gear (20).

Abstract: Method of manufacturing a hollow blade (1) for a turbomachine in which the blade (1) is made from a preform (14) obtained from external primary parts (28, 30). In order to reduce costs, the root part (2) of the blade is formed from spare material (34) entirely located on only one of the two external primary parts of the preform (14).

Abstract: Method to introduce a deliberate mismatch into a turbomachine bladed wheel so as to reduce vibration amplitudes of the wheel in forced response, characterised by the fact that it consists of determining an optimum value of the standard deviation for the mismatch as a function of operating conditions of the wheel inside the turbomachine, with respect to the maximum vibration amplitude response required on the wheel, and of at least partly placing blades with different natural frequencies on the said wheel such that the standard deviation of the frequency distribution of all blades is equal to at least the said mismatch value, the said mismatch value being determined statistically.

Abstract: The present invention concerns a turbofan jet engine comprising an outer fan duct, an inner fan duct, fluid-conveying ancillaries located outside the outer fan duct, fluid-conveying ancillaries located inside the inner fan duct, characterized by the fact that at least one removable modular arm forming a connection for ancillaries is arranged between the outer fan duct and the inner fan duct. With the invention, the connection of ancillaries outside the outer fan duct to ancillaries inside the inner fan duct is achieved simply by means of a removable arm that is easy to mount and dismount.

Abstract: A cooling device for a stationary ring surrounding a gas-turbine hot-gas flow path, said ring being made up of a plurality of ring segments that are fastened by upstream and downstream fastener systems onto a plurality of spacer segments forming a support spacer surrounding said ring so as to co-operate therewith to define at least one annular impact cavity into which at least one air supply orifice opens out, each ring segment having walls that are pierced with a plurality of air exhaust holes opening out both into the impact cavity, and into the hot-gas flow path, the device further comprising means for channeling the air coming from leakage through the seals of the fastener systems, and for directing it towards at least one of the axial ends of the ring segments so as to cool said end(s).

Abstract: This invention concerns a turbofan jet engine comprising an outer casing supported by arms, an inner casing facing the casing and ancillaries. An ancillaries distribution support is formed along the axis of an arm, on the inner casing, comprising a first sole plate for reception of ancillaries on the outer side of the inner casing. The invention reduces the longitudinal dimension necessary for the ancillaries to pass through.

Abstract: A device for lubricating a component in an assembly of parts comprises a lubricant distribution chamber arranged in the vicinity of the component and formed by an annular setback in a ring mounted on a part of the assembly, said setback being closed automatically when the ring is mounted by means of a corresponding surface of the part into which there open out channels or ducts formed in the part and feeding the setback with lubricant and/or connecting the setback to means for taking lubricant to the component.

Abstract: A device for assembling together annular flanges by means of fastening bolts, in particular in a low-pressure turbine stage of the turbojet, one of the flanges has an element projecting from its face that faces the other flange and that is received in a notch in said other flange when the relative angular positioning of the two flanges is correct.

Abstract: Device for fixing a burner ring onto flameholder arms (14) in an afterburner combustion chamber of a turbojet engine, the burner ring being formed of ring sectors (12) placed more or less end to end and the ends of which are housed and guided between two circumferentially-directed parallel plates (34, 36) of the flameholder arms (14) and are immobilized between these plates (34, 36) by locking pieces (52) which are themselves held between the plates (34, 36) by immobilizing means (54, 56) engaged in aligned orifices (58, 60, 64) in the plates (34, 36) and in the locking pieces (52).

Abstract: A turbomachine including a wheel with hollow blades that are cooled internally by a forced flow of cooling air. According to an important characteristic of the invention, a hook-annulus (70) is interposed between the outer edge of the labyrinth-disk (52) and the rotor disk (36), the hook-annulus including a retaining shoulder (72) constituting an outer annular hook in which the outer edge of the labyrinth-disk is engaged, and a bearing shoulder (74) constituting an inner hook engaged in a groove around the circumference of the rotor disk.

Abstract: The present invention relates to a forging die including on one face, a half-imprint of a part to be forged such as a turbine engine blade. The die is characterized by the fact that it includes on said face, a means forming a marking along two directions relatively to which the position of said imprint is defined, said means consisting of two pads protruding relatively to said face and each including two notches defining said two directions. This means preferably consists of two pads machined on the die at the same time as the imprint, and each including two notches in the form of a cross. The invention thus allows visual checking of the positioning of the imprints and of the dies on the platens of the press.

Abstract: A turbomachine comprising at least a first rotary assembly with a first shaft, a stator, and bearings secured to the stator and suitable for supporting said shaft, one of said bearings being capable of giving way when an unbalance occurs in the rotary assembly, at least one segment of said shaft being covered in a covering which comprises at least one piece of fabric impregnated with resin and which is for coming into contact with a surrounding portion of the turbomachine when said unbalance occurs in order to protect the first shaft and ensure that there is no risk of said first shaft rupturing. The turbomachine is intended to be used mainly in the field of aviation, and more particularly as an airplane turbojet.

Abstract: The turbojet according to the invention comprises a fixed structure, a fan rotor fixed to a drive shaft (5) supported by a first bearing (6) and a second bearing (7), mounted on the said structure fixed through a bearing support part (11, 19), the first bearing (6) being mounted on the fixed structure of the turbojet through a device (13) enabling it to be decoupled (13) from the fixed structure, characterised by the fact that the second bearing (7) is mounted on the bearing support part (19) through a link acting as a ball joint (23, 24) and in that the turbojet also comprises means enabling axial displacements of the second bearing (7) with respect to the fixed structure of the turbojet if the first bearing (6) should be decoupled. The centring function of the second bearing (7) after decoupling of the first bearing (6) is thus provided without introducing excessive stresses in the bearing (7).

Abstract: The invention relates to a hybrid cycle HP compressor (104) for a turbine engine, comprising in particular a first rotor wheel (16), a first rectifier (118) and a second rotor wheel (20), as well as a separation slot (22) fitted between the first rectifier and the second rotor wheel in order to supply the turbine engine with a secondary air flux. According to the invention, the first rectifier is a double grid rectifier comprising a grid of vanes upstream (30) as well as a grid of vanes downstream (34), these upstream and downstream vane grids being associated respectively with variable setting systems upstream and downstream (38,40) controlled in autonomous manner.

Abstract: The invention relates to aerodynamically matching the rear fan of a turbojet having two fans and a low-pressure compressor at the front of the intermediate casing, said fans being driven by two independent shafts. Said compressor is disposed between the blades of the two fans and includes at least one ring of rotor blades around the periphery of a wheel driven by the drive shaft for the front fan, and at least two grids of stator vanes disposed on either side of the ring of rotor blades and inside a grid carrier ring. A stationary outer grid and a stator of variable pitch connect the ring to the fan casing.

Abstract: The invention relates to a turbojet section comprising an external casing having a surface that extends radially inwards, an internal casing having an essentially axial wall that extends toward said surface, and a seal located between said wall and said surface to prevent communication between the high and low pressure regions situated on either side of said seal, in which said seal is made in the form of an annular sheet-metal ring comprising an essentially cylindrical first part fixed in a leaktight manner to one face of the axial wall and a second part continuing from said first part and situated in the space separating said axial wall from said radial surface, said second part exhibiting, in section on a radial plane containing the axis of the turbojet, a V-shaped profile and having an end portion in leaktight sliding contact with said radial surface.