Abstract: The invention relates to a turbine engine combustion module (10), in particular for an aircraft turbine engine, designed to carry out constant-volume combustion, comprising: at least two combustion chambers (12A, 12B) arranged about an axis, each chamber (12A, 12B, 12C) comprising a compressed gas intake port (16) and a burnt gas exhaust port (18); and an ignition means that triggers combustion in the combustion chambers (12A, 12B, 12C). The module (10) comprises at least one duct (80) which establishes a communication between a first combustion chamber (12A) and at least one second combustion chamber (12B) in order to inject burnt gases from the first combustion chamber (12A) into the second combustion chamber (12B) so as to trigger combustion in the second combustion chamber (12B).

Abstract: A discharge system of a separated twin-flow turbojet for an aircraft, supported by a suspension mast, is disclosed. The system includes a main nozzle delimited by an annular cowl with a slot of annular shape defining upstream and downstream portions of the cowl and which is traversed by the suspension mast. The downstream portion of the cowl of the main nozzle includes a first part extending downstream from the upstream portion of the cowl to a trailing edge of the main nozzle, on either side of the suspension mast along two predefined angular sectors; and a second part formed from an internal contour of the slot and having a trailing edge with a diameter smaller than that of the trailing edge associated with the first part of the downstream portion of the cowl. Connecting walls laterally connect the first and second parts of the downstream portion.

Abstract: A device for circulation of air for a turbomachine, the device including an inlet provided to receive incident airflow, air circulation ducts connected to the inlet, a central part including side walls, peripheral walls arranged at a distance from the central part, so as to form the air circulation ducts, each duct being arranged between one of the side walls of the central part and one of the peripheral walls, each peripheral wall including an upstream end portion delimiting with one of the side walls an inlet orifice of one of the ducts and a downstream end portion delimiting with one of the side walls an outlet orifice of one of the ducts, at least one member mobile in translation relative to the peripheral walls, the at least one mobile member being formed in the central part and arranged so that its displacement varies a section of each of the inlet orifices and a section of at least one of the outlet orifices.

Abstract: A turbine engine convergent-divergent nozzle including an annular central element and an annular cap arranged coaxially around the central element to co-operate with the central element to define an annular flow channel for a gas stream from the engine. Between a throat section and an ejection section of the nozzle, the central element and the cap present respective internal profiles in longitudinal section that are modeled by curves having respective radii of curvature that are identical in absolute value.

Abstract: Heat-exchange and noise-reduction panel for a propulsion assembly, in particular for an aircraft, the panel comprising: a perforated plate comprising a plurality of through-openings; a cellular structure comprising longitudinally oriented structural walls covered by said perforated plate and comprising, between said walls, cavities that define Helmholtz resonators, said through-openings forming necks of said resonators; and means for the circulation of fluid, for example oil, at said perforated plate, wherein said fluid circulation means comprise channels that are formed at least in part in thickened ends of said walls on the same side as said perforated plate, and/or at least in part in regions of the perforated plate situated in the longitudinal extension of said thickened ends.

Abstract: A forward fan rotor is disclosed with a hub of axis of rotation (X) and a cone mounted on the hub of the fan. The cone comprises an air bleed orifice which opens into an air duct of which a forward end portion passes through the fan rotor, said forward end portion comprising mechanical air entrainment means. The air bleed orifice has an annular shape and in that the cone is divided by said orifice into a front vertex portion and a rear frustoconical portion. A turbomachine forward axial spool equipped with such a fan rotor is also disclosed.

Abstract: A turbomachine including a nozzle extending along an axial direction and configured to convey a hot gas stream; and a mounting pylon for mounting the turbomachme, in which the pylon extends radially from a base arranged in the vicinity of the nozzle, is provided. The turbomachine includes a pylon cooler device having at least one auxiliary gas duct having a gas inlet configured to take cold gas from the outside of the nozzle, and a gas outlet opening out in the vicinity of the base of the pylon.

Abstract: In an assembly including a bypass turbine engine and an attachment system attaching the turbine engine to an aircraft structure, the attachment system includes a mast, at least two rear attachments connecting the mast to two rear attachment points of the turbine engine, and a rear covering covering the rear attachments and providing a streamlined fairing. The rear covering includes two side walls, each including a first portion extending along and in proximity of the mast and extending rearwards, and a second portion projecting sideways relative to the first portion towards its outside, to cover the rear attachments. Each second portion is substantially a bulge and is connected to the first portion along a curved connection line extending between front and rear points situated substantially level with an inner longitudinal edge of the first portion to minimize the impact on a flow of bypass stream leaving the bypass passage.

Abstract: According to the invention, the method for defining the shape of a convergent-divergent nozzle of longitudinal axis L-L comprising a convergent part connected, at a throat, to a divergent part the downstream free end of which comprises scallops delimiting noise-reducing chevrons distributed in the circumferential direction, comprises defining beforehand the divergent part of the nozzle using a divergent first portion, a divergent second portion and a convergent portion, attributing an initial value to dimensional parameters defining the chevrons and the three portions, calculating performance criteria and checking that the calculated performance criteria satisfy predefined performance conditions.

Abstract: A turbomachine (10) comprising a nozzle (14) extending along an axial direction (A), said nozzle (14) being configured to convey a hot gas stream (GC), and a mounting pylon (12) for mounting the turbomachine, said pylon (12) extending radially from a base (12a) arranged in the vicinity of the nozzle (14). The turbomachine (10) includes a pylon cooler device (18) having at least one auxiliary gas duct (20) having a gas inlet (20a) configured to take cold gas (GF) from the outside of the nozzle (14), and a gas outlet (20b) opening out in the vicinity of the base (12a) of the pylon (12).

Abstract: A turbine engine convergent-divergent nozzle including an annular central element and an annular cap arranged coaxially around the central element to co-operate with the central element to define an annular flow channel for a gas stream from the engine. Between a throat section and an ejection section of the nozzle, the central element and the cap present respective internal profiles in longitudinal section that are modeled by curves having respective radii of curvature that are identical in absolute value.

Abstract: In an assembly including a bypass turbine engine and an attachment system attaching the turbine engine to an aircraft structure, the attachment system includes a mast, at least two rear attachments connecting the mast to two rear attachment points of the turbine engine, and a rear covering covering the rear attachments and providing a streamlined fairing. The rear covering includes two side walls, each including a first portion extending along and in proximity of the mast and extending rearwards, and a second portion projecting sideways relative to the first portion towards its outside, to cover the rear attachments. Each second portion is substantially a bulge and is connected to the first portion along a curved connection line extending between front and rear points situated substantially level with an inner longitudinal edge of the first portion to minimize the impact on a flow of bypass stream leaving the bypass passage.

Abstract: A discharge system of a separated twin-flow turbojet for an aircraft, supported by a suspension mast, is disclosed. The system includes a main nozzle delimited by an annular cowl with a slot of annular shape defining upstream and downstream portions of the cowl and which is traversed by the suspension mast. The downstream portion of the cowl of the main nozzle includes a first part extending downstream from the upstream portion of the cowl to a trailing edge of the main nozzle, on either side of the suspension mast along two predefined angular sectors; and a second part formed from an internal contour of the slot and having a trailing edge with a diameter smaller than that of the trailing edge associated with the first part of the downstream portion of the cowl. Connecting walls laterally connect the first and second parts of the downstream portion.

Abstract: A core cowling for a bypass turbojet engine consisting of an internal wall (7) bathed in the main airflow (1) of said turbojet engine and an outer wall (6) bathed in the secondary airflow (2) thereof, the two walls between them forming a cavity (5) which receives the ventilation airflow originating from the hot parts of the turbojet engine to be injected into the secondary airflow through said outer wall, wherein said outer wall is pierced with multiple perforations (8) which are uniformly distributed on the circumference of said barrel and spread longitudinally along said outer wall in the direction of flow of the secondary airflow (2), so as to connect said cavity that receives the ventilation airflow to said secondary flow along at least two parallel circles of perforations.