Abstract: Sail wing aircraft which includes a wing (6) and at least one propulsion engine (8). It includes an upper beam (22) which is firmly fixed at its front end to a first frame (12) located on an air inlet (14) of the propulsion engine and which is in addition firmly fixed at its median part to a second frame (16) located to the rear of the first frame. The sail wing aircraft includes in addition a pylon (26) for attachment of the engine onto the fuselage, where the engine is fixed to the pylon (26).

Abstract: A fluid conduit coupling for a turbofan engine wherein the engine has a gas path defined by inner and outer duct walls and the exterior surface of the outer duct wall is readily accessible for visual inspection. In one example, a first fluid line extends across the gas path between the inner and outer duct walls and a second fluid line is connected to the first fluid line at the level of the outer duct wall. The first fluid line may include a female coupling member and the second fluid member may include a mating male coupling member. An annular sealing member is associated with the first and second couplings and a leak path is defined between the annular sealing member and the exterior surface of the outer duct wall, whereby fluid may be detected by visual inspection of the outer duct wall if a leak should occur at the annular sealing member.

Abstract: The present invention relates to a nacelle for a turbojet, comprising an air inlet structure capable of channelling a flow of air to a turbojet fan and a middle structure (5) comprising a casing (9) designed to encircle said fan, with the air inlet structure attached to it, said structure having a peripheral inside surface at least partly equipped with a sound-attenuating structure (13) extending without geometrical interruption around at least part of the casing, said structure being characterized in that a gap (14) is provided between the sound-attenuating structure and the casing.

Abstract: A suspension of a multi-flow turbojet engine provided with an intermediate casing and an exhaust casing from a pylon that can be attached to the structure of an aircraft is disclosed. The suspension includes a forward attachment device between the hub of the intermediate casing and the pylon, a rear attachment device between the exhaust casing and the pylon, and a connection device rigidly connecting the intermediate casing to the pylon. The rear attachment device includes an actuator for compensating for the variations in diameter of the exhaust casing so as to keep the axis of the exhaust casing coaxial with the axis of the intermediate casing through the various phases of flight of the aircraft.

Abstract: A mounting system for a gas turbine engine includes a mounting linkage assembly and a tangential link positioned generally transverse to the mounting linkage assembly. The mounting linkage assembly reacts at least a thrust load. The tangential link reacts at least a vertical load, a side load, and a torque load of the gas turbine engine.

Abstract: A securing mast for securing an engine to the wing of an aircraft, wherein the device includes two fasteners, a first end of each fastener being secured to the engine and the second end of each fastener being secured to the wing, wherein the fasteners include hinged junction means capable of modifying the geometry of the aforementioned fasteners in order to modify a distance between the engine and the wing.

Abstract: A supporting structure for a gas turbine engine includes at least one annular member, and a plurality of circumferentially spaced elements that extend in a radial direction of the annular member and that are connected to the annular member. At least one of the circumferentially spaced elements has an airfoil shape in cross section. The annular member includes at least two substantially flat panel segments that are arranged side by side in a circumferential direction of the annular member. The panel segments are connected to each other in a connection region that extends along facing end parts of the two panel segments. The airfoil shaped element is connected to the annular member at the connection region. A mean camber line of the airfoil shaped element at least along a portion of its length is inclined in relation to an axial direction of the annular member, and the connection region at least along a portion of its length is inclined in relation to the axial direction of the annular member.

Abstract: An aircraft assembly including a wing element and an attachment pylon of a turboengine, including a rigid structure forming a caisson defined externally by a first longeron, a second longeron, and two lateral panels. The assembly also includes an attachment mechanism of the structure on the wing element, fitted with two front attachments, each including two first fittings solid with the associated lateral panel. For each front attachment, the two first fittings are arranged respectively on either side of their associated lateral panel.

Abstract: An attachment device for an aircraft engine that includes a rigid structure and an attachment mechanism of the engine to the rigid structure. The attachment mechanism includes a device reacting to thrust forces generated by the engine, the reacting device including two link rods mechanically connected to a balance beam at a rear end by a first mechanical connection, a main fitting fixed to the rigid structure to which the balance beam is mechanically connected, and two additional fittings fixed to the rigid structure to which the link rods are mechanically connected with play. The mechanical connections between the balance beam and the main fitting and between the link rods and the additional fittings each include a pin orthogonal to a mid-plane of transmission of the forces.

Abstract: A rear lower aerodynamic fairing for an attachment device for an aircraft engine, including two side panels assembled together with transverse stiffening inner ribs spaced apart from each other along a longitudinal direction of the fairing, and a heat protection floor to be closely followed by a primary flow of the engine. The floor includes plural floor segments distributed along the direction and non-rigidly bound to each other, and the fairing further includes an assembling mechanism assembling the floor segments, offsetting the latter of the ribs to which are connected the assembling mechanism.

Abstract: A mounting system for a gas turbine engine includes a thrust ring and a linkage assembly. The linkage assembly is at least partially received by the thrust ring. The linkage assembly reacts at least a side load and a thrust load communicated from the thrust ring.

Abstract: A gas turbine engine has a casing component, such as a mid turbine frame system, which includes an outer case surrounding a spoke or strut casing. Load transfer spokes of the spoke casing are detachably connected to the outer case and radially extend between the outer case and an inner case of the spoke casing. A bearing housing is supported by the inner case of the spoke casing. An arrangement for radially centring the bearing within the outer case is provided, the arrangement including at least three radial locators mounted to the outer case.

Abstract: Provided is a seal structure, according to the present invention, for sealing opposite surfaces of flanges between adjacent tail ducts, which can be prevented from being worn or aged deteriorated due to a thermal deformation in a high temperature atmosphere or vibration of a gas turbine combustor, and which can maintain a satisfactory sealing function for a long time.

Abstract: An aircraft engine assembly including an engine and a mount device for the engine. The mount device can be positioned between a wing of the aircraft and the engine, and includes a rigid structure and a mechanism to mount the engine on the rigid structure, the mounting mechanism including a first attachment and a second attachment. The first attachment includes two side semi-attachments each fixed to a fan case of the engine, and the second attachment is fixed onto a forward part of a central case of the engine.

Abstract: A turbomachine including an annular combustion chamber is disclosed. The combustion chamber includes a radially inner wall and a radially outer wall exhibiting symmetry of revolution. An annular chamber end wall equipped with a fuel injection device connects upstream ends of the inner and outer walls. A suspension device suspends the upstream end of the combustion chamber from an outer casing.

Abstract: There is provided a gas turbine capable of improving the durability of a main shaft by absorbing the deformation of a casing due to heat and pressure and suppressing stresses acting on the main shaft. A link plate 31 of a movable supporting leg 30 that supports a casing 26 of a gas turbine 20 on the turbine 24 side is installed beforehand in an inclined manner at an angle ?1, whereby the displacement X of a lower casing 26L in the height direction; which occurs between a pin 31a in a top end portion of the link plate 31 and a top surface of a lower casing 26L when the operation status of the gas turbine 20 shifts from a stop condition to an operating condition is compensated for. This prevents a main shaft 25 from changing in the height direction in the part of the movable supporting leg 30 and suppresses bending stresses acting on the main shaft 25.

Abstract: An aircraft engine assembly including an attachment mechanism of a turbojet onto a rigid structure of a mounting strut, including a first, second, and third forward engine attachment for taking up thrust loads brought onto the fan frame, and arranged such that the third attachment passes through a first diametrical plane of the turbojet, the first and second attachments being disposed on one side and another of the first plane. The first and second forward engine attachments are respectively brought to the fan frame at two points, situated beyond the second diametrical plane of the turbojet, orthogonal to the first diametrical plane, with respect to the third attachment.

Abstract: A propulsion unit comprises a turbojet engine and a thermal exchanger, situated above the turbojet engine. The turbojet engine comprises a regulation device to regulate an air flow-rate in a cooling air lead-in conduit conveying a stream of cooling air into a thermal exchanger, The regulation device regulates the air flow-rate upstream with respect to the cooling air lead-in conduit.

Abstract: A power system includes a gas turbine engine. The gas turbine engine may include a first portion, a second portion, and a longitudinal axis. The power system may also include a first moveable support engaged to the first portion of the gas turbine engine. The first moveable support may be operable to at least partially constrain movement of the first portion of the gas turbine engine during operation of the gas turbine engine. Additionally, the first moveable support may be operable to at least partially support the first portion of the gas turbine engine as the first portion of the gas turbine engine is moved away from the longitudinal axis independently of the second portion of the gas turbine engine.

Abstract: An aircraft engine assembly including an annular structure surrounding a central casing and mechanically connected to the central casing by the intermediary of a mounting mechanism. To each of the first and second engine mounts is associated a reinforcing structure forming a shear plane, fixedly connected: in an area of the annular structure at a first anchoring point; in an area of the fan casing at a second anchoring point; and in an area of a structural arm or of the intermediate casing at a third anchoring point. The structure extends along a radial imaginary plane passing through an anchoring point of the mount on the fan casing.

Abstract: A fastener for a turbine engine includes a bolt and a ceramic matrix composite (CMC) cap mechanically secured to the bolt. CMC components are attached to an exhaust nozzle case with the fasteners. The fastener includes a bolt made of metal and having a head. A CMC cap is located on the head of the bolt and secured to the bolt by a mechanical device. The mechanical device allows the CMC cap to have a different thermal expansion than the bolt. The CMC cap also acts as a thermal shield for the bolt, aiding in the reduction of a heat signature of the turbine engine.

Abstract: An annulus filler assembly for a rotor of a turbomachine, the assembly comprising: an annulus lid (56) having a radially outwardly facing surface (58) for forming an inner wall of a flow annulus of the rotor and a radially inwardly facing surface; and a frame for supporting the annulus lid (40), the frame being mountable to a disc of the rotor such that the annulus lid (56) is spaced away from the disc, wherein the frame (40) comprises a connection portion (46) which, in use, passes through an aperture (60) in the annulus lid (56) from the radially inwardly facing surface towards the radially outwardly facing surface (58) such that at least a portion of the connection portion (46) is visible from the radially outwardly facing surface (58); the assembly further comprising a locking element (68) which locks the connection portion (46) to the annulus lid (56) via the visible portion of the connection portion (46).

Abstract: A rigid structure of a suspension pylon for an aircraft engine, which is in a shape of a box closed by first and second box side panels. The structure includes inter-rib spaces, each space delimited by two directly consecutive box transverse ribs. For each inter-rib space forming part of a group including at least three arbitrary and directly consecutive inter-rib spaces, a single access opening is provided on the side panels, the openings arranged alternately on the first and second side panels.

Abstract: In an annular torsional rigid static component for an aircraft engine, the component includes at least one arrangement for mounting the engine to an aircraft. The mounting arrangement comprises a body formed in a composite material.

Abstract: A coupling assembly for an engine core of a turbomachine is provided. The coupling assembly comprises a duct having a substantially circular shape and surrounding the engine core, and a plurality of coupling struts, each coupling strut extending only in a substantially vertical direction relative to the direction of gravity when the engine core is at a rest position with zero pitch and zero roll, each of the plurality of coupling struts coupling the engine core to the duct.

Abstract: A cam-type apparatus is included in a support link between outer and inner cases of gas turbine engines for centering the cases one to another. The cam-type apparatus is lockable to allow locking an adjusted position of the cam-type apparatus.

Abstract: A gas turbine engine case having a working fluid flow, includes inner and outer case portions defining an annular duct for directing the working fluid flow, and a plurality of struts positioned within the annular duct and extending between the inner and outer case portions. The struts are welded to the inner and outer case portions with a first weld along a peripheral line of the respective struts and with second weld, of fillet type, in selected locations for additionally connecting a portion of each strut to the respective inner and outer case portions.

Abstract: A gas turbine engine having a mid turbine frame comprising an annular outer case providing a portion of an engine casing; an interturbine duct (ITD) disposed within the outer case, the ITD including outer and inner rings radially spaced apart one from another and being interconnected by a plurality of radially extending and circumferentially spaced hollow strut fairings, the inner and outer rings co-operating to provide a portion of a hot gas path through the engine; a tube for delivering or discharging a lubricant fluid to or from a bearing housing, the tube extending radially through one of the hollow struts; and an insulation structure radially extending through one said hollow strut fairing, the insulation structure surrounding the tube and being spaced apart from the tube and from a hot internal surface of the one hollow strut fairing for shielding the tube from heat radiated from the hot internal surface of the one hollow strut fairing and for preventing the lubricant fluid from contacting the hot inter

Abstract: A mid-turbine frame is connected to at least one mount of a gas turbine engine for transferring a first load from a first bearing and a second load from a second bearing to the mount. The mid-turbine frame includes a first load structure, a second load structure, and a plurality of struts. The first load structure combines the first load and the second load into a combined load. The second load structure transfers the combined load to the mount. The struts are connected between the first load structure and the second load structure and transfers the combined load from the first load structure to the second load structure.

Abstract: Provided is a seal structure, according to the present invention, for sealing opposite surfaces of flanges between adjacent tail ducts, which can be prevented from being worn or aged deteriorated due to a thermal deformation in a high temperature atmosphere or vibration of a gas turbine combustor, and which can maintain a satisfactory sealing function for a long time.

Abstract: A mid-turbine frame connected to at least one mount of a gas turbine engine transfers a first load from a first bearing and a second load from a second bearing to the mount. The mid-turbine frame includes a load transfer unit, a torque box rotatably positioned within the load transfer unit, and a plurality of struts. The load transfer unit has a first locking element and combines the first load and the second load into a combined load. The torque box has a second locking element that is engagable with the first locking element of the load transfer unit. The plurality of struts are connected between the torque box and the mount, and transfer the combined load from the torque box to the mount. The first locking element and the second locking element are at least one of a rib or a groove.

Abstract: A mounting element for transferring loads between a gas turbine engine and an airframe is provided, the mounting element presenting a cylindrical hole for receiving a fastening element, such as a pin or a bolt. The mounting element includes a collapsible element surrounding the hole, and is adapted to operate in a first mode, and to operate, upon a collapse of the collapsible element, in a second mode, in which the stiffness of the mounting element is lower than in the first mode. In a second aspect of the invention the energy absorption capacity of the mounting element is higher in the second mode than in the first mode.

Abstract: A system for moving an auxiliary component of a gas turbine engine is disclosed. The system may have a mobile support platform configured to support the auxiliary component during operation of the gas turbine engine. The system may further have at least one guide assembly operably connected to the mobile support platform and configured to guide movement of the mobile support platform. The system may also have a drive assembly operably connected to the mobile support platform and configured to impart movement of the mobile support platform between an elevated position and a lowered position.

Abstract: A casing for a gas turbine includes a construction providing improved structural efficiency. Improved load paths and means for transmitting loads in the engine case are disclosed.

Abstract: A gas turbine engine mid turbine frame having an annular interturbine duct and vane ring assembly includes a duct having outer and inner duct walls of sheet metal interconnected by radial hollow struts of sheet metal and a vane ring is connected to the duct to provide the assembly. The interturbine duct and vane ring assembly may be provided within a mid turbine frame in a manner which is independent of a bearing load path through the mid turbine frame.

Abstract: This invention relates to a force transmission system (50) between a turboprop and an attachment pylon, comprising: a first and a second force transmission device (52a, 52b) each comprising a first hollow cylinder (54) with an axis (56) that will be made parallel to a first force transmission direction, said first cylinder defining a first space (58) full of a fluid (60) and closed off in a sealed manner at its two opposite ends, by a first damping layer (62) that will bear in contact with a bearing element (66) of the turboprop, and by a second damping layer (64) that will bear in contact with a bearing element (68) of the attachment device, said first and second layers possibly being compressed by the pressure of said fluid; and first means (70) of fluid communication between the two spaces (58, 58).

Abstract: A mid turbine frame of a gas turbine engine includes an outer case which supports a spoke casing co-axially positioned therein. The spoke casing has load transfer spokes extending radially from an inner case and secured to the outer case. A load transfer device is provided to transfer load from the spokes to the outer case in addition to load transfer through a first group of fasteners securing the spokes to the outer case, thereby forming a secondary load transfer path from the spokes. The load transfer device includes an opening of the outer case into which at least some of the spokes are inserted.

Abstract: An attachment device for an aircraft engine that includes a rigid structure and an attachment mechanism of the engine to the rigid structure. The attachment mechanism includes a device reacting to thrust forces generated by the engine, the reacting device including two link rods mechanically connected to a balance beam at a rear end by a first mechanical connection, a main fitting fixed to the rigid structure to which the balance beam is mechanically connected, and two additional fittings fixed to the rigid structure to which the link rods are mechanically connected with play. The mechanical connections between the balance beam and the main fitting and between the link rods and the additional fittings each include a pin orthogonal to a mid-plane of transmission of the forces.

Abstract: An inlet assembly for an aircraft engine nacelle includes an inner barrel, an outer barrel radially spaced from the inner barrel, and a rear support for supporting the outer barrel relative to the inner barrel. The rear support includes at least one radially extending stiffener and at least one energy absorber defining a zone of deformation adjacent the inner barrel. The rear support is deformable in the zone of deformation in response to forces applied during a fan blade-out event to prevent fracture of the rear support.

Abstract: A hinged device for at least one door (46.1, 46.2) of an aircraft nacelle (40), whereby the door is able to block an opening (44) that is made in the fairing of the nacelle, the device including a base to which is connected at least one door, able to occupy a first retracted state in which the base is located at least in part inside the nacelle and a second extended state in which the base is located outside the fairing of the nacelle so as to be able to move, in a direction that is essentially parallel to the longitudinal axis of the nacelle, outside of the nacelle by entraining the at least one door (46.1, 46.2) to release the opening (44) at least in part.

Abstract: A structural turbine engine casing is disclosed. The casing includes a central hub, at least one intermediate shroud, and an external ring, which are concentric, the external ring and the intermediate shroud being connected via links, the intermediate shroud and the central hub being connected by a series of radially fastened hollow arms. The series of arms includes at least one main arm and at least one secondary arm, each main arm being arranged in the continuation of a link and having a mass which is greater than that of a secondary arm arranged at a distance from the links.

Abstract: System for attaching an engine to the structure of an aircraft, such as a sail wing aircraft. The system includes an upper beam which is firmly fixed at its front end to a first frame mounted on the aircraft structure. The upper beam is in addition firmly fixed at its median part to a second frame located to the rear of the first frame. The engine is hooked beneath the upper beam. The system is orientated in relation to a longitudinal axis X, a transverse axis Y and a vertical axis Z. Two front engine attachment points are arranged symmetrically on either side of the vertical median plane of the engine. The front attachment points include a shaft (44, 46) which is inclined relative to the Y direction and to the Z direction so that they bear the forces along the longitudinal direction and forces which are inclined relative to the Z direction and Y direction.

Abstract: The nacelle of the invention includes an air-inlet front section, a middle section for encasing the fan of the turbofan (7) and a rear section including means for connection to a pylon to be connected to the rigid structure or an airplane. The rear section includes a structural framework (18) having at least one aerodynamic smoothing and acoustic panel (21). The acoustic panel (21) is secured on the structural framework (18) by floating or elastic fixation means so that said acoustic panel (21) can be deformed in a direction (31) essentially centrifugal and radial relative to the turbofan (7) in the presence of overpressure air (3) in the engine compartment.

Abstract: The invention relates to an engine assembly for an aircraft comprising a turbojet, a engine mount pylon, and a plurality of engine mounts inserted between the engine mount pylon and the turbojet. According to the invention, the plurality of engine mounts comprises a engine mount (6a) designed so as to uniquely resist forces applied along a direction (Y) transverse to the turbojet, this engine mount (6a) comprising an intermediate fitting (46) assembled onto a first fitting (40) fixed to the pylon through two swivel axes (48) oriented parallel along a vertical direction (Z), and a pin (56) oriented along a longitudinal direction (X) and fixed to the intermediate fitting, the pin being installed on a second fitting (58) fixed to the turbojet, with clearance along the longitudinal direction (X).

Abstract: A mounting device for an aircraft engine includes a rigid structure and a device for mounting the engine on the rigid structure, the mounting device including a thrust load recovery device with two lateral thrust load recovery rods each having a rear end mounted on a crossbar of the recovery device with a first mechanical connection. At the rear end of each of the two rods, there is also provided a second mechanical connection with play between the rear end and a support fitting integral with the rigid structure. The first mechanical connection is implemented using a pin provided on the crossbar, passing through an orifice provided in the rear end of the rod.

Abstract: The present invention relates, on the one hand, to a nacelle (1) for a double flow turboreactor (2) comprising a front air intake section (5), a median section (6) intended for surrounding a fan (3) of the turboreactor and a rear section (7), said rear section having an internal structure (7b) intended for serving as a housing to a rear portion of the turboreactor, characterised in that the internal structure possesses attachment means suitable for enabling the nacelle to be fastened to a pylon (12) intended to be connected to a fixed structure (13) of an aeroplane on at least one portion of said internal structure, and on the other hand, to a propellant assembly and to an aircraft provided with such a nacelle.

Abstract: Turboshaft engine including two subassemblies defining between them an annular chamber housing a seal. The two subassemblies are assembled under axial stress thereby defining an annular chamber housing the seal and an interposed part is inserted between the butting surfaces of the two parts of the annular chamber.

Abstract: A mounting system for a gas turbine engine includes a thrust ring and a linkage assembly. The linkage assembly is at least partially received by the thrust ring. The linkage assembly reacts at least a side load and a thrust load communicated from the thrust ring.

Abstract: The invention concerns a fixing system for at least one component of a pod of a turbojet on an engine strut to which the pod is attached and comprising at least one coupling zone provided by the component capable of co-operating with at least one complimentary coupling zone provided by the strut, and fixing means for providing with complementary fixing means a rigid and detachable connection between said component and the strut. The invention is wherein the fixing means and the complementary fixing means are capable of slightly pivoting relative to each other, as well as relative to a corresponding pod.

Abstract: A gas turbine engine assembly includes a gas turbine engine mounted on a three-point mounting frame. The three-point mounting frame includes an engine support structure having first and second beams in a V-configuration, and a secondary support structure that is positioned about the engine support structure and has a rectangular footprint. The frame may include a Y-configuration that includes the V-configuration, and may be implemented in a machine mounting system, such as in a marine vessel.