Abstract: The present invention relates to a thrust orienting nozzle, shaped in such a way as to divide a principal propulsion gas flow coming from at least one gas generator into a first flow and a second flow for an ejection in a first half-nozzle and in a second half-nozzle and comprising at least one of following two piloting means: a means of dividing the principal flow into each of the two half-nozzles, and a means of orienting the thrust vector produced by each of the two half-nozzles. The invention applies in particular to the yaw control of an aircraft without a vertical tail unit.

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: The present invention relates to a propulsion gas exhaust assembly, in an aircraft propelled by hot gases produced along the axis of the latter by a gas generator, comprising a duct and a nozzle, wherein said duct comprises a first cylindrical duct element (21) receiving the gases and communicating downstream with two second duct elements (22, 23), the directions of which are divergent in a first plane, each of the two duct elements communicating downstream with a third duct element (25, 27) which emerges in an axial gas-ejection half-nozzle (24, 26).

Abstract: The present invention relates to a propulsion gas exhaust assembly in an aircraft propelled by hot gases produced along the axis of the latter by a gas generator, comprising a duct (21) and a nozzle (24). It is noteworthy by the fact that said duct forms a vertical elbow.

Abstract: The present invention relates to a thrust orienting nozzle, shaped in such a way as to divide a principal propulsion gas flow coming from at least one gas generator into a first flow and a second flow for an ejection in a first half-nozzle and in a second half-nozzle and comprising at least one of following two piloting means: a means of dividing the principal flow into each of the two half-nozzles, and a means of orienting the thrust vector produced by each of the two half-nozzles. The invention applies in particular to the yaw control of an aircraft without a vertical tail unit.

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 divergence measurement antenna for single-pulse radars comprises at least two radiant panels. With their beams having the same center of phase, they are oriented differently. A monotonic function of the angular divergence of a signal received by the antenna is obtained from the ratio of the power of the signal received by the first panel to its power received by the second panel. An even-parity function of the signal received is obtained by taking the sum of these two signals. Application to precise measurements of angular divergence.