Abstract: A bridge positioned between two parts, such as a nozzle support pipe envelope and a heat shield liner, includes a central part with inversed curvature thus having two contact zones with the nozzle support pipe envelope. Furthermore, the bridge includes two lugs, one of which can be mounted to slide in the heat shield liner.

Abstract: The after-burner device receives a “hot” central primary flow from the turbine of the turbojet and a “cold” peripheral secondary flow, and it has an after-burner ignition zone situated in the secondary flow that reaches the after-burner device. A fraction of the primary flow is brought into the after-burner ignition zone in order to raise the temperature in said zone to a value that is higher than that of the secondary flow so as to encourage after-burner ignition.

Abstract: A device for feeding air and fuel to a burner ring in an after-burner chamber, including a plurality of flame-holder arms extending radially about the axis of the chamber from an outer casing, and a burner ring made up of ring sectors mounted substantially end to end on the flame-holder arms, the ring sectors containing fuel feed rail and air manifold boxes.

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: The invention relates to a burner ring for turbofan jet engines. The ring includes an upstream annular envelope which forms a channel which is open axially in the downstream direction, and a manifold for injecting fuel into the channel. The ring includes multiple sectors connected together. Each sector has a fuel inlet which is connected to the fuel injection manifold. Part of the upstream annular envelope is in the core flow. Each sector has a connector which receives the fuel inlet and a ventilation duct which extends through the channel and upstream of the injection manifold. Each sector is fitted with an inlet for a bypass air, which air is then emitted by the ventilation duct to cool the injection manifold. A sector of downstream annular envelope is arranged downstream of the injection manifold to protect this manifold.

Abstract: A turbojet having an after-burner chamber presenting a specific injector opening into the burner ring to provide secure ignition. The after-burner chamber has an annular fuel injection duct, an ignition spark plug, and an ignition injector installed facing the spark plug and connected to specific fuel feed means.

Abstract: The bridge piece makes it possible to limit mechanical stresses in the heat shield liner and the nozzle support pipe envelope. According to the invention, it comprises a central part with inversed curvature (3) thus having two contact zones (4) with the nozzle support pipe envelope (2). Furthermore, one of the two lugs (5B) is mounted sliding in the heat shield liner (1). Application to turboprops with cooling bypass air flow.

Abstract: A device for fastening a fluid duct such as a fuel duct in an orifice of a turbojet casing, the device comprising a ring screwed into the orifice of the casing and having an annular rim coming to bear against a washer at the end of the duct, and a nut screwed onto the end of the duct so as to clamp the annular rim of the ring against the washer of the duct in order to fasten the duct to the casing.

Abstract: The invention concerns a double flow turbo-jet engine, including a re-heating channel of the primary flow and an ejection nozzle, the secondary flow emerging at least partially in the primary flow upstream of the re-heating channel, the re-heating channel including a burner ring, including a flame holder gutter in the form of a ring portion whereof the upstream section is situated in the secondary flow, receiving a fuel manihold and a protective tubular screen of the manihold. The invention is characterised in that the gutter includes an upstream recess linked with the secondary flow, and the screen includes, on its upstream section, at least one ventilation orifice of the space situated between the screen and the manihold.

Abstract: The turbojet of the invention includes a combustion chamber, a channel for heating the gas stream, where the heating channel includes at least one device for injection of fuel into the gas stream, which includes an open chamber, with a U-shaped section, having at least one wall within which extend fuel-injection means, which inject the fuel in at least one direction. It is characterised by the fact that a cooling jacket is provided in the chamber, alongside the wall forming the base of its U-section, and the fuel injection device includes protection means interposed between the fuel-injection means and the wall, in a fuel-injection direction. Thus, as a result of the invention, the fuel injection device, which is bathed in a very hot environment, is protected against thermal shocks due to the projection of colder fuel onto its walls, and its life expectancy is therefore increased.

Abstract: The invention relates to ventilating the confluence sheet of an afterburner chamber of an aviation turbomachine, including, upstream from the afterburner chamber: a diffuser defined by a confluence sheet disposed inside a casing, said casing and said confluence sheet defining between them an annular channel for a cold secondary flow, upstream fuel injectors being disposed at the inlet to said diffuser and flame-catchers being disposed downstream from said upstream injectors, said confluence sheet presenting, between the radial plane containing said upstream injectors and the radial plane situated at the rear ends of said flame-catchers two bends so as to flare downstream in order to slow down the primary flow F1 downstream from said upstream injectors.

Abstract: The invention relates to ventilating the confluence sheet of an afterburner chamber of an aviation turbomachine, including, upstream from the afterburner chamber: a diffuser defined by a confluence sheet disposed inside a casing, said casing and said confluence sheet defining between them an annular channel for a cold secondary flow, upstream fuel injectors being disposed at the inlet to said diffuser and flame-catchers being disposed downstream from said upstream injectors, said confluence sheet presenting, between the radial plane containing said upstream injectors and the radial plane situated at the rear ends of said flame-catchers two bends so as to flare downstream in order to slow down the primary flow F1 downstream from said upstream injectors.

Abstract: The after-burner device receives a “hot” central primary flow from the turbine of the turbojet and a “cold” peripheral secondary flow, and it has an after-burner ignition zone situated in the secondary flow that reaches the after-burner device. A fraction of the primary flow is brought into the after-burner ignition zone in order to raise the temperature in said zone to a value that is higher than that of the secondary flow so as to encourage after-burner ignition.

Abstract: A turbomachine comprising, downstream from the turbine, a post-combustion chamber defined by a thermal protection lining inside a casing defining an annular channel for a secondary flow, an annular diaphragm secured at the downstream end of said channel, said nozzle comprising a plurality of flaps hinged to the upstream end of said casing, each flap fitted on its inside face with a thermal protection plate to define a passage for cooling air delivered by said diaphragm, wherein the cooling air is provided by an annular duct defined by a first flexible annular gasket in sliding contact against the downstream inside face of the casing and against the upstream inside faces of the flaps under the pressure of the secondary flow, and defined on the inside by a second flexible annular gasket whose upstream end is fixed to the radially inner zone of the diaphragm, and whose downstream end is in sliding contact against the upstream inside face of the protection plates.

Abstract: A device for feeding air and fuel to a burner ring in an after-burner chamber comprises a plurality of flame-holder arms extending radially about the axis of the chamber from an outer casing, a burner ring made up of ring sectors mounted substantially end to end on the flame-holder arms, and means for feeding air and fuel to the burner ring sectors, said means extending between the burner ring and the outer casing of the after-burner chamber inside the flame-holder arms.

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: The invention relates to an afterburner ring 19 for turbofan jet engines. The ring comprises an upstream annular envelope which forms a channel which is open axially in the downstream direction, and a manifold 4 for injecting fuel into the channel. The ring has sectors 20 which each comprise a sector 1 of the upstream annular envelope. Each sector 1 has a fuel inlet 35 which is connected to the injection manifold 4. Part of the upstream annular envelope is in the core flow. Each sector 20 has a connecting means which receives the fuel inlet and a ventilation duct 2 which extends through the channel along the sector 1, at a point upstream of the injection manifold 4. Each sector 1 is fitted with an inlet for bypass air, which air is then emitted by the ventilation duct 2 to cool the injection manifold 4. A sector 5 of downstream annular envelope is arranged downstream of the injection manifold 4 to protect it.

Abstract: The invention relates to an aviation turbomachine comprising, downstream from the turbine, a post-combustion chamber extended by at least one nozzle, said chamber being defined radially by a thermal protection lining disposed inside a casing, said casing and said lining together defining an annular channel in which, in operation, there flows a secondary flow, an annular diaphragm secured to said casing being disposed at the downstream end of said channel, said nozzle comprising a plurality of flaps hinged to the upstream end of said casing, each flap being fitted on its inside face with a thermal protection plate co-operating with said flap to define a passage for being fed with cooling air delivered by said diaphragm, wherein the feed of cooling air to said passages is provided by an annular duct defined on the outside by a first flexible annular gasket retained, in operation, pressed in sliding contact against the downstream inside face of the casing and against the upstream inside faces of the flaps under