Abstract: In a turbomachine comprising a metal material shell containing in a gas flow direction F: a fuel injection assembly; a composite material combustion chamber; and a metal material sectorized nozzle forming the inlet stage with fixed blades of a high pressure turbine, provision is made for the combustion chamber to be held by a sectorized flexible sleeve of metal material having one end fixed to the combustion chamber by first fixing means and a flange-forming opposite end fixed to the shell by second fixing means. The first fixing means also serve to connect the combustion chamber to the sectorized nozzle.

Abstract: A device for controlling fuel pressure and delivery as necessary for the operation of the servovalves of an aircraft engine is disposed downstream of a fuel injector feed unit of the engine. The device comprises a chamber maintained at the servovalve feed pressure by the movement of a closure member to vary the opening of a port of the chamber. The closure member moves under the action of opposing forces comprising inter alia a force in the opening direction proportional to the servovalve feed pressure. An equilibrium position of the closure member is a function of pressure resulting from the engine speed. An injector delivery control loop is therefore isolated from a servovalve pressure and delivery control loop.

Abstract: A variable cycle propulsion system for a supersonic airplane comprises at least one engine capable of generating thrust for flight at supersonic speeds together with at least one auxiliary propulsion assembly that is separate from the engine and that is capable of generating additional thrust for takeoff, landing, and flight at subsonic speeds. The auxiliary propulsion assembly does not have a gas generator and means are provided for transmitting a fraction of the mechanical power produced by the engine to the auxiliary propulsion assembly in order to enable it to generate the additional thrust for takeoff, landing, and subsonic cruising flight. Means are provided for decoupling the mechanical transmission means for supersonic cruising flight.

Abstract: An emergency device for relighting a windmilling turbojet, the jet comprising a fan driven by a low-pressure turbine via a first shaft and a compressor driven by a high-pressure turbine via a second shaft disposed coaxially around the first shaft, a differential interconnecting said first and second shafts while compensating for their different speeds of rotation in normal operation or the turbojet, and a braking system connected to the differential so as to enable it to be slowed down or blocked when the turbojet shuts down, thereby enabling the first shaft to entrain the second shaft so that it reaches a speed that favors relighting of the turbojet.

Abstract: A variable cycle propulsion system for a supersonic airplane, the system comprising at least one engine having means for producing exhaust gas and a gas exhaust nozzle for generating thrust for supersonic flight speeds, and at least one separate auxiliary propulsion assembly dissociated from said engine, having no gas generator, and capable of generating thrust for takeoff, landing, and subsonic flight speeds. The system further comprises gas flow tapping means movable between a position in which they tap off at least a fraction of the exhaust gas produced by said engine and feed it to said propulsion assembly to enable it to generate thrust for takeoff, landing, and subsonic cruising flight, and a position in which the gas produced by the engine is directed solely to the engine nozzle for supersonic cruising flight.

Abstract: A combustion chamber for a gas turbine made up of outer and inner side walls, the combustion chamber being received in a casing so as to define an annular space between the combustion chamber and the casing in which there flows air for combustion, for dilution, and for cooling the combustion chamber, the side walls of the combustion chamber being pierced by a plurality of holes in which bushings of substantially elliptical right section are fixed to define air injection passages for injecting air into the combustion chamber, each bushing having a peripheral wall in which at least one additional orifice is formed opening out into the combustion chamber in the immediate vicinity of the side wall of the combustion chamber in which said bushing is fixed so that the air passing through said orifice flows substantially along said peripheral wall, the peripheral wall of each bushing having at least one groove opening out into the annular space and into which the or each orifice opens out so as to be fed with air and

Abstract: A variable cycle propulsion system for a supersonic airplane, the system comprising at least one engine having at least one compressor and capable of generating thrust for supersonic flight speeds, and at least one auxiliary propulsion assembly separate from said engine, and capable of generating additional thrust for takeoff, landing, and subsonic cruising speeds, at least one pipe connecting the engine to the auxiliary propulsion assembly to convey thereto compressed air produced by the compressor so as to enable the auxiliary propulsion assembly to generate additional thrust for takeoff, landing, and subsonic cruising flight, there being at least one valve enabling the pipe to be closed off for supersonic cruising flight.

Abstract: An annular combustion chamber has outer and inner axially-extending side walls of composite material and an end wall of metal material held in position on the outer and inner side walls by fixing means. Provision is made for the fixing means to pass through annular cavities that are designed to receive cylindrical end portions of the outer and inner side walls, and that are created between peripheral edges of the end wall and facing portions folded downstream, a determined amount of clearance J being provided between the peripheral edges and the facing faces of the outer and inner side walls in such a manner as to allow expansion to take place freely, in operation, in a radial direction between said end wall and said side walls.

Abstract: A turbomachine has inner and outer annular shells of metal material containing, in a gas flow direction F, a fuel injector assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine. Provision is made for the combustion chamber to be held in position between the inner and outer metal annular shells by a plurality of flexible metal tabs having first ends interconnected by a metal ring fixed securely to each of the annular shells by first fixing means, and second ends fixed by second fixing means on a ring of composite material fixed securely to said composite material combustion chamber, the flexibility of said metal fixing tongues allowing expansion to take place freely in a radial direction at high temperatures between said composite material combustion chamber and said metal annular shells.

Abstract: A turbomachine has inner and outer annular shells of metal material containing, in a gas flow direction F, a fuel injector assembly, a composite material annular combustion chamber, and an nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine. Provision is made for the combustion chamber to be held in position between the inner and outer metal annular shells by a plurality of flexible metal tabs having first ends interconnected by a flange-forming metal ring fixed securely to each of the annular shells by first fixing means, and second ends fixed by second fixing means firstly to the composite material combustion chamber and secondly to one end of a composite material wall whose other end forms a bearing plane for a sealing element secured to the nozzle and providing sealing for the gas stream between the combustion chamber and the nozzle.

Abstract: A turbomachine comprises an annular shell of metal material containing in a gas flow direction F a fuel injection assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine, said nozzle being supported by the annular shell and being fixed thereto by first releasable fixing means, and provision being made for the combustion chamber to be mounted in floating manner inside the annular shell and held in position solely by the nozzle to which it is fixed in resilient manner by second releasable fixing means.

Abstract: A device for controlling fuel pressure and delivery as necessary for the operation of the servovalves of an aircraft engine is disposed downstream of a fuel injector feed unit of the engine. The device comprises a chamber maintained at the servovalve feed pressure by the movement of a closure member to vary the opening of a port of the chamber. The closure member moves under the action of opposing forces comprising inter alia a force in the opening direction proportional to the servovalve feed pressure. An equilibrium position of the closure member is a function of pressure resulting from the engine speed. An injector delivery control loop is therefore isolated from a servovalve pressure and delivery control loop.

Abstract: In a turbomachine comprising inner and outer annular shells of metal material containing, in a gas flow direction F: a fuel injection assembly; an annular combustion chamber of composite material having a plurality of air feed orifices in a primary zone and/or in a dilution zone; and an annular nozzle of metal material forming the inlet stage with fixed blades for a high pressure turbine, provision is made for the combustion chamber to be held in position between the inner and outer annular metal shells by means of a plurality of flexible metal tongues having first ends interconnected by respective metal rings fixed securely to said inner and outer annular shells by first fixing means, and second ends fixed securely to the composite material combustion chamber at it air feed orifices by second fixing means.

Abstract: In an annular combustion chamber having inner and outer axially-extending side walls of composite material and an end wall of metal material, provision is made to hold the end wall in position between the side walls of the annular combustion chamber by means of a plurality of flexible metal tongues fixed firstly to said side walls by fixing means and secondly to said end wall by brazing or welding, said end wall further comprising means for providing sealing between said end wall and said side walls.

Abstract: In a turbomachine comprising an annular shell of metal material containing in a gas flow direction F: a fuel injection assembly; an annular combustion chamber of composite material; and an annular nozzle of metal material forming the inlet stage with fixed blades of a high pressure turbine, provision is made for the combustion chamber to be held in position inside the annular metal shell by a plurality of flexible metal tongues each comprising three branches connected together in a star configuration, the ends of two of these three branches being fixed securely to a downstream end of the combustion chamber via respective first and second fixing means, and the end of the third branch being fixed securely to the annular shell via third fixing means.

Abstract: A spacer (36) made to be installed between two permanently connected discs (31 and 32), without a bolted flange, and comprising axial thrust (43, 47) as well as outer radial thrust (45, 48) alignment surfaces (41). It comprises moreover an outer cover (37) to provide sealing and a cooling passageway, and stiffening tongues (38) which accommodate thrust loads.

Abstract: A method for peening the tops of cooled blades which include cooling air passage orifices. Peening is performed by mobilization of a plurality of beads in a sealed chamber containing at least one blade tip. The plurality of beads is mobilized by the active surface of a sonotrode which delimits part of said chamber. The beads have a diameter greater than that of the cooling air passage orifices.

Abstract: A suspension system for an aircraft powerplant attached to an aircraft frame and having a housing. The suspension system includes a spindle mounted in a spindle support rigidly joined to the aircraft frame, a bearing cage affixed to the housing and having a bearing positioned therein, a pair of suspension arms connected to the spindle support, and a pair of lateral linkrods configured to link the housing and the suspension arms. The side surfaces of the bearing cage and the suspension arms define a clearance therebetween concentric to the Y-axis. The bearing cage is configurable along the spindle such that the bearing cage is limited to displacement relative to the longitude of the spindle while maintaining contact with the spindle.

Abstract: A connector (30) in a network of lubrication pipes (31, 32, 33) is provided with a valve (45) that automatically closes off the connector (30) whenever one of the pipes is withdrawn following an incident. Lubricant leaks are thus avoided. The valve plug is cylindrical-conical (46, 47) and it is fitted with a needle (48).

Abstract: A sealing device is disclosed comprising a least one sealing leaf maintained axially in position by a retaining bracket which comprises a first end bearing on the sealing leaf and a bent-back second end engaged in a groove in a structural element. The sealing leaf is preferably retained radially in position by lugs which are bent so as to engage below lugs provided on the retaining bracket. When the sealing device comprises at least two adjacent sealing leaves placed edge to edge, the device also comprises a cover sheet so disposed between the sealing leaves and the retaining bracket as to cover the joint between the two adjacent leaves.