Abstract: The tank for fluid under pressure comprises one or an assembled-together plurality of individual containers or modules made at least in part out of composite material. The or each individual container (20) comprises a cylindrical body (22) of composite material, two end plates (30) closing the axial ends of the cylindrical body, and at least one belt passing around the container substantially in a longitudinal direction and bearing against portions of the outside faces of the end plates.

Abstract: An endoscope with deflected distal viewing, comprises a rigid tube containing observation means, themselves comprising a distal deflector prism having a viewing axis that is inclined relative to the axis of the tube, and ultraviolet light guide means housed in the tube and opening out at their distal end onto a prism for deflecting the illuminating light beam in a direction that is substantially parallel to the viewing axis. The invention applies in particular to inspecting parts by the penetration test technique in the aviation field.

Abstract: The composite turbomachine blade of the present invention comprises a preform made of yarns or fibers woven in three dimensions and a binder maintaining the relative disposition between the yarns of the preform. Said preform is made up of warp yarns and of weft yarns, the direction of the warp yarns forming the longitudinal direction of the preform. In characteristic manner, said preform comprises at least a first portion made using a first weave forming the airfoil of the blade, and a second portion made using a second weave forming the root of the blade, and the first and second portions are united by a transition zone in which the first weave is progressively modified to end up with the second weave, thereby obtaining a reduction at least in the thickness of the blade between the second portion and the first portion. The invention is applicable to making a fan blade for a turbojet.

Abstract: An attachment device for attaching a blade to a turbine rotor disk in a turbomachine, the blade root forming a male portion engaged and held radially in a cavity of complementary shape in the rotor disk, with means for preventing the blade root from moving axially in the cavity, the blade axis being inclined relative to a perpendicular to the axis of rotation of the rotor, in a direction and at an angle that are designed to compensate at least in part for the displacement of the axis of the blade caused by the mechanical and thermal stresses applied to the rotor during operation of the turbomachine.

Abstract: The casing of the invention is of the type supporting series of stationary blades having disposed between them series of blades that are movable in rotation about a longitudinal axis. In characteristic manner, said casing comprises a main element and, at least facing one of the series of moving blades, an assembly comprising a plate made of a honeycomb material and a covering disposed on the face of the plate facing away from the blades so that said cells are open towards the blades, said covering being provided with holes that open out into cells of the plate, thereby forming open cells, said covering closing other cells of the plate, thereby forming closed cells, a cavity being formed between said plate and the inside face of said main element. Preferably, the casing forms the casing of an axial compressor operating at low pressure in a turbojet engine.

Abstract: A system for decoupling a fan of a turbojet, the turbojet including a moving fan shaft mounted on bearings each connected to a stationary structure of the turbojet by bearing supports, the bearing supports being fixed to the stationary structure of the turbojet via mechanical links serving, during normal operation of the turbojet, to transmit the forces generated by rotation of the fan, the mechanical links include explosive charges enabling the links to be broken, the explosive charges being controlled by a full authority digital engine control computer on the basis of means for measuring the mechanical stresses in the turbojet and of a computer model simulating the static, dynamic, and thermodynamic behavior of the turbojet.

Abstract: The invention relates to a rotation-locking device of a sector of rectifier blades (5) mounted in a sector of the turbojet housing. A sector of blades comprises a support base (3). A housing sector comprises two recesses perpendicular to the axis of the housing, each receiving an edge (34-1, 34-2) of the support base. At one and the same axial end of the housing sector, each recess opens out into a seat (30-1, 30-2) defining an axial face (32-1). The device comprises an axial stop (26) adapting itself on each side in the seats (30-1; 30-1) of the two recesses, and once the complementary sector of the housing is installed and during operation of the turbojet, bearing on the axial faces (32-1) under the thrust of the complementary sector of rectifier blades in order to provide rotation locking.

Abstract: A device is dedicated to setting one or more pivoting axial flow synchronizing rings (R), comprising a drive mechanism (4-6) intended to rotate their rotary spindle (3) so as to set them angularly depending on angular position data supplied by a means of control. These means of control comprise firstly means of illumination (8, 10) designed to supply light to an illumination zone (11), secondly an encoder element (12) attached to the spindle of the synchronizing ring (R) in the illumination zone (11) and designed to interact with the illumination light in order to generate optical signals representing the angular position of the synchronizing ring, thirdly means (13a) designed to collect the optical signals and fourthly means of processing (7) designed to determine angular position data based on the optical signals collected.

Abstract: The device for detection of an oil unbalanced mass inside the rotor (2) of a machine includes means of entraining oil (100) present in the rotor (2) in rotation, such that the oil rotation speed is practically the same as the rotation speed of said rotor.

Abstract: The present invention pertains to turbine engine mounting onto an aircraft pylon, comprising a bracket (12) arranged to be attached to the pylon, a lever (15;15?) attached in its central part (150;150?) to the bracket by first connection means (150-130) with a pivot pin (136;136?) and two thrust links (20,22; 20?,22?) each joined aft to the lever by second connection means (200-155, 220-157; 200?-155?; 220?-157?) and including engine attachment means at the forward end, characterized in that the first connection means (150-130) consists of two parallel arms (151,153) spaced apart on central part (150) of the lever, cooperating with an attachment lug (130) integral with the bracket via a pivot pin, and lever (15,15?) comprises two third side connection means on standby (155-135, 157-137; 155?-135?; 157?-137?) with bracket (12,12?). The invention provides for simplified mounting with improved safety.

Abstract: A rolling bearing for a turbomachine supporting a first shaft to rotate relative to a second shaft, and comprising: a plurality of rolling elements engaged between an inner annular ring mounted on the second shaft and an outer annular ring mounted on the first shaft; an oil film formed at an annular contact surface between said outer ring and said first shaft; at least one duct passing through said inner ring and opening out in an inner race for said rolling elements in order to lubricate them; and at least one additional duct passing through said inner ring and opening out outside the inner race in order to feed oil through said outer ring to said oil film under the effect of the centrifugal force generated by said second shaft rotating, said outer ring having an annular shoulder projecting radially towards the inside of said annular space, said shoulder having at least one channel opening out substantially in register with said additional duct and opening out into said oil film.

Abstract: The invention relates to a turbomachine (1) for an aircraft comprising a gas generator (4) designed such that hot gases are ejected from a combustion chamber (28) towards the upstream side of the turbomachine, and also comprising a plurality of hollow mixer struts (40) connected to an output (36) from the gas generator through which hot gases can pass, each mixer strut comprising an output mixer portion (40c) located inside the annular fan duct (8) so as to mix the secondary air and the hot gases ejected by these output mixer portions towards the downstream side of the turbomachine.

Abstract: The invention relates to a ceramic thermal barrier composition of low conductivity and high high-temperature strength. In characteristic manner, the composition comprises a zirconia base, at least one trivalent oxide enabling the zirconia to be stabilized and enabling the thermal conductivity of the zirconia to be reduced in optimum manner, and at least one pentavalent oxide enabling the content of oxygen vacancies to be reduced so as to make it substantially equal to that of a partially stabilized zirconia. Preferably, said trivalent oxide is present with a molar content constituted by a first portion enabling the zirconia to be stabilized partially and a second portion which introduces point defects into the lattice said pentavalent oxide is present with a molar concentration equal to said molar concentration of said second portion of the trivalent oxide. The invention is applicable to making a superalloy machine part having a ceramic coating with said composition, and possibly also a bonding underlayer.

Abstract: An inter-vane platform (3) for a vane support disk (1) of a fan of a turbine engine, comprising a deflecting part (4) includes a lower face (5) provided with at least one first securing foot (6), provided with at least one aperture (7) for the passage of a securing pin (8) intended to connect it to a second securing foot (12) of the support disk (1), between two adjacent vanes.

Abstract: The invention relates to a hollow blade comprising an internal cooling passage (24), an open cavity (30) located at the tip (14) of the blade and bounded by an end wall (26) and a rim (28) and cooling channels (32) that connect the said internal cooling passage (24) to the outer face of the pressure wall (16), the said cooling channels (32) being inclined to the pressure wall (16) in such a way that they emerge on the outer face of the pressure wall (16) near the top (28a) of the said rim. By way of feature, a reinforcement of material (34) is present between the rim (28) and the end wall (26) of the cavity (30) along at least one portion of the pressure wall (16), whereby the said rim (28) is widened at its base adjacent to the said end wall (26) in such a way that the cooling channels (32) emerge near the top (28a) of the rim (28) without reducing the mechanical strength of the tip (14) of the blade.

Abstract: The present invention relates to a convergent-divergent turbojet nozzle (1) comprising driven divergent flaps (4a), follower divergent flaps (4b) interposed between the driven flaps, and means (32) for supplying cooling air to the follower flaps (4b), said follower flaps (4b) having a box structure and having lateral openings for delivering cooling air towards the inner face of said driven flaps (4a), in such a way as to limit the heating up of these flaps when the turbojet is in operation.

Abstract: The casing of the invention is of the type supporting series of stationary blades having disposed between them series of blades that are movable in rotation about a longitudinal axis. In characteristic manner, said casing comprises a main element and, at least facing one of the series of moving blades, an assembly comprising a plate made of a honeycomb material and a covering disposed on the face of the plate facing away from the blades so that said cells are open towards the blades, said covering being provided with holes that open out into cells of the plate, thereby forming open cells, said covering closing other cells of the plate, thereby forming closed cells, a cavity being formed between said plate and the inside face of said main element. Preferably, the casing forms the casing of an axial compressor operating at low pressure in a turbojet engine.

Abstract: The invention relates to the sealing of the cavity from which air is bled off to the cabin, which cavity is delimited, on the one hand, by the external shell (6) of the compressor and an annular structure (7) connected to the shell, and, on the other hand, by the external casing (12) of the diffuser grating and a strut (13) connected to said external casing and to an external engine casing shell (14), this external casing shell (14) being fastened to the annular structure (7) by bolting together flanges, using sealing means provided between the annular structure (7) and the external casing (12) of the diffuser grating.

Abstract: System for injecting an air/fuel mixture into a turbomachine combustion chamber, comprising a hollow tubular structure for the flow of the air/fuel mixture into the combustion chamber fuel injection means placed at an upstream end of the hollow tubular structure air injection means placed downstream of the fuel injection means cold plasma generating means placed downstream of the air injection means so as to generate active species in the flow of the air/fuel mixture and to cause prefragmentation of the molecules of the air/fuel mixture and, means for controlling the said cold plasma generating means depending on the speed of operation of the turbomachine.

Abstract: A device for controlling the clearance between the tips of rotary blades and a stationary ring assembly of a gas turbine, the device comprising a circular control box having at least two annular air circulation strips that are spaced apart from each other in the axial direction, each having a plurality of perforations in order to modify the temperature of the stationary ring assembly by discharging air, an annular air feed channel radially spaced from the air circulation strip, at least one air duct for feeding the feed channel with air, and a plurality of hollow distribution spacers connecting the air feed channel to the air circulation strips in order to feed the strips with air while enabling the air that has been discharged against the stationary ring assembly to flow between the feed channel and the circulation strips in order to be evacuated therefrom.