Airbus Operations Patente

A landing gear assembly for an aircraft, the landing gear assembly comprising: a forward brace having a first end for attachment to an airframe of the aircraft, and a second end; an aft brace having a first end for attachment to the airframe of the aircraft, and a second end, wherein the second end of the aft brace is coupled to the second end of the forward brace; an actuator piston having a first end and a second end, wherein the first end of the actuator piston is pivotally attachable to the airframe of the aircraft; a bellcrank linkage having a first end, a second end, and a pivot point between the first end and the second end, such that a first portion of the bellcrank linkage extends from the first end to the pivot point, and a second portion of the bellcrank linkage extends from the pivot point to the second end, wherein the pivot point of the bellcrank linkage is pivotally attachable to the airframe of the aircraft, and wherein the first end of the bellcrank linkage is pivotally coupled to the second end of the actuator piston; a primary strut having a first end and a second end, wherein the first end of the primary strut is pivotally coupled to the second end of the forward brace and the second end of the aft brace, and wherein one or more wheels of the aircraft are mounted at the second end of the primary strut; and a secondary strut (e.g. a shock absorber) having a first end and a second end, wherein the first end of the secondary strut is pivotally coupled to the second end of the bellcrank linkage, and wherein the second end of the secondary strut is pivotally coupled to a pivot point near to the second end of the primary strut; wherein the landing gear assembly is configured to adopt a deployed configuration when the actuator piston is in an extended configuration, and wherein the landing gear assembly is configured to adopt a retracted configuration when the actuator piston is in a retracted configuration. Also provided is an aircraft (such as a blended wing body aircraft) having such a landing gear assembly.

A method for use in assembling an aerofoil structure. A torsion box has a first mounting feature and a second mounting feature. A fixed leading or trailing edge structure has a third mounting feature and a fourth mounting feature. The first mounting feature and the third mounting feature each have a fastener hole prior to aligning the leading or trailing edge structure to the torsion box. The second mounting feature and the fourth mounting feature are without fastener holes for joining the torsion box to the fixed leading or trailing edge structure prior to aligning the leading or trailing edge structure to the torsion box. The method includes aligning the fixed leading or trailing edge structure to the torsion box so as to align the fastener holes of the first mounting feature and the third mounting feature, drilling through the second mounting feature and the fourth mounting feature to form fastener holes in the second mounting feature and the fourth mounting feature when the fixed leading or trailing edge structure is aligned with the torsion box, and installing a fastener through the fastener holes of the first mounting feature and the third mounting feature, and installing a fastener through the fastener holes of the second mounting feature and the fourth mounting feature, so as to join the fixed leading or trailing edge structure to the torsion box.

The present invention relates to a riblet for a fixed leading edge (FLE) assembly. The aforementioned riblet comprises a first riblet portion having a first aft mounting for attaching the first riblet portion to a wing box assembly and also having a forward mounting opposite the first aft mounting. The aforementioned riblet also comprises a second riblet portion having a second aft mounting for attaching the second riblet portion to the forward mounting of the first riblet portion.

An inerting system (22) has an inerting apparatus (44) which receives an inlet air flow and separates it into nitrogen enriched air (NEA) and oxygen enriched air (OEA). The inerting system also has an oxygen concentration sensor (46) and a controller (48) connected to the oxygen concentration sensor (46). The controller (48) adjusts operation of the inerting system (22) during use, based at least in part on feedback received from the oxygen concentration sensor (46).

A landing gear assembly (30) for an aircraft (10), the landing gear assembly comprising: a forward brace (40) having a first end (41) for attachment to an airframe of the aircraft, and a second end (43); an aft brace (50) having a first end (51) for attachment to the airframe of the aircraft, and a second end (53), wherein the second end (53) of the aft brace is coupled to the second end (43) of the forward brace; an actuator piston (70) having a first end (71) and a second end, wherein the first end (71) of the actuator piston is pivotally coupled to the first end (41) of the forward brace; a bellcrank linkage (60) having a first end (62), a second end (64), and a pivot point (63) between the first end and the second end, such that a first portion of the bellcrank linkage extends from the first end (62) to the pivot point (63), and a second portion of the bellcrank linkage extends from the pivot point (63) to the second end (64), wherein the pivot point of the bellcrank linkage is pivotally coupled to the first end (51) of the aft brace, and wherein the first end (62) of the bellcrank linkage is pivotally coupled to the second end of the actuator piston (70); a primary strut (90) having a first end (91) and a second end (96), wherein the first end of the primary strut is pivotally coupled to the second end (43) of the forward brace (40) and the second end (53) of the aft brace (50), and wherein one or more wheels (100) of the aircraft are mounted at the second end (96) of the primary strut (90); and a secondary strut (80, e.g. a shock absorber) having a first end (81) and a second end (83), wherein the first end of the secondary strut is pivotally coupled to the second end (64) of the bellcrank linkage (60), and wherein the second end (83) of the secondary strut (80) is pivotally coupled to a pivot point near to the second end (96) of the primary strut (90); wherein the landing gear assembly is configured to adopt a deployed configuration when the actuator piston (70) is in a retracted configuration, and wherein the landing gear assembly is configured to adopt a retracted configuration when the actuator piston (70) is in an extended configuration. Also provided is an aircraft (such as a blended wing body aircraft BWB) having such a landing gear assembly.

An aircraft wing comprising: a wing structure supporting a wing cover, the wing cover having an outer aerodynamic surface, which forms part of an aerofoil profile of the wing, and a recess stepped back from the wing aerofoil profile; and a spoiler device comprising a spoiler coupled to the wing structure by a spoiler actuation mechanism for moving the spoiler between a retracted position and a deployed position, the spoiler having a surface which follows a contour of the wing aerofoil profile and where the spoiler is housed entirely in the recess when the spoiler is in the retracted position, and wherein the spoiler surface is extended away from the wing cover when the spoiler is in the deployed position.

A method of monitoring a landing gear shock absorber (4) comprises monitoring a tyre gas pressure of a landing gear tyre; at a first time, when the aircraft is stationary before a flight, recording a first ambient temperature and a first tyre gas pressure; at a second time, when the landing gear is unloaded during the flight, recording a second tyre gas pressure and a shock absorber pressure; calculating a second ambient temperature at the second time; comparing the recorded shock absorber pressure with a predetermined pressure range for the second temperature; and providing an alert if the recorded shock absorber pressure is outside the predetermined pressure range. This allows a health check of the shock absorber (4) to be performed without the need to directly measure the temperature of the shock absorber (4).

A wing assembly 10 for an aircraft, comprising a fixed wing 18 and a wing tip device 12 rotatable about a joint assembly 16. The wing assembly is configurable between a flight configuration for use during flight and a ground configuration for use during ground-based operations. The wing assembly 10 comprises joint-bordering structure located adjacent the joint assembly 16, the joint-bordering structure comprising a rib 34, an upper cover 26 and a lower cover. The joint assembly comprises a moveable panel 32. The panel 32 comprises a first hard stop element 38a, and the joint-bordering structure comprises a second hard stop element 40a. The first hard stop element 38a is integrally formed in the panel 32, and the second hard stop element 40a is integrally formed in one of the rib 34, the upper cover 26 or the lower cover of the joint-bordering structure. The hard stop elements 38a, 40a may be configured to self-align.

A method of measuring deformation of a structure by image correlation, the method comprising: generating a non-periodic tiling with a set of prototiles; generating a non-periodic pattern based on the non-periodic tiling; applying the non-periodic pattern to a structure; generating a reference image of the non-periodic pattern on the structure; generating a deformed image of the non-periodic pattern on the structure; and measuring deformation of the structure by comparing the deformed image with the reference image.

A wing hinge assembly (13) is disclosed. The wing hinge assembly (13) comprises a wing tip fitting (19) pivotally mounted to a fixed wing fitting (15). The wing tip fitting (19) is rotatable about a hinge axis (H) with respect to the fixed wing fitting (15). One of the wing tip fitting (19) and the fixed wing fitting (15) defines a locking channel (26). The other of the wing tip fitting (19) and the fixed wing fitting (15) comprises a locking member (34) which defines a locking member axis. The wing hinge assembly (13) is configurable to a locked configuration in which the locking member (34) is received in the locking channel (26) by moving the locking member (34) in a direction perpendicular to the locking member axis.

An aircraft fuel tank monitoring system, comprising: at least one first sensor for sensing a pressure of fuel within a fuel tank; at least one second sensor for sensing a pressure of an ullage space above the fuel within the fuel tank; a processor for monitoring the change in the pressure of the fuel within the fuel tank over a time period and for monitoring the change in the pressure of the ullage space within the fuel tank over the time period, and to determine a fault if the monitored pressure changes exceed normal operating conditions.

A filling level monitoring device (100) for monitoring a filling level (210) of a fluid container (200) is provided that includes an exciting element (10), a sensing element (21), a signal source (30), a suspension element (70) with a defined stiffness, and a processing unit (40). The exciting element (10) is configured to mount on the fluid container (200) or suspension element (70) to activate a specific suspension mode (25) of vibration. The suspension element (70) attaches rigidly to the fluid container (200) so that both vibrate together. The sensing element (21) detects these vibrations and sends data to the processing unit (40). The signal source (30) generates a multi-frequency input signal (31) for the exciting element (10). The processing unit (40) uses the vibration data to identify the suspension mode (25), determine its modal frequency (80), and calculate the fluid container's (200) current filling level (210) based on this frequency.

The invention provides an aircraft wing assembly with a fixed wing and a wing tip device. The wing tip device is moveable between a flight configuration and a ground configuration. In the ground configuration the span of the aircraft wing assembly is reduced. The fixed wing comprises a fixed wing lug and the wing tip device comprises a wing tip lug. A locking pin is provided. In the flight configuration the fixed wing lug and wing tip device lug are aligned such that the locking pin may be inserted through the fixed wing lug and wing tip device lug into a locking position to lock the wing tip device in the flight configuration. The locking pin is hollow and comprises a heating element.

The invention relates to a latch mechanism 12 for locking an aircraft folding wing. The mechanism comprises at least one lug 13, 14 having an aperture 15a-15f and a latch pin 16 moveable within the aperture between a latched configuration in which the latch pin is received in the aperture, and an unlatched configuration in which the latch pin is withdrawn from the aperture. At least part of one of the pin 16 and the aperture 15a-15f has a hydrophobic coating. A hydrophobic coating may be applied to all apertures in the mechanism and/or to the latch pin. The provision of a hydrophobic coating helps to prevent the formation of moisture and hence ice on the latch pin and/or in the aperture, which ice can inhibit the pin from moving along the aperture into and out of its latching position.

A wing (3) for an aircraft (1), comprising a main wing (5) and a leading edge high lift assembly (6) comprising a slat (7) and a connection assembly (9) movably connecting the slat (7) to the main wing (5), such that the slat (7) is movable between a retracted position (11), a first extended position (13), and a second extended position (15), wherein the connection assembly (9) comprises an elongate slat track (17) that extends along a track longitudinal axis (19) between a first end (21) and a second end (23) and has an intermediate portion (25) between the first and second ends (21, 23), wherein the first end (21) and/or the intermediate portion (25) of the slat track (17) is mounted to the slat (7), wherein the slat track (17) is mounted to the main wing (5) by a roller bearing (27) such that the slat track (17) is movable along the track longitudinal axis (19). The object to provide a wing having a beneficial ratio of aerodynamic performance and noise generated by extension of the slat, is achieved in that the slat track (17) is mounted to the slat (7) rotatably about a slat axis (22), and the connection assembly (9) comprises a slat angle mechanism (10) mounted to the slat (7) and configured to control the angle of the slat (7) about the slat axis (22) when the slat (7) is moved between the retracted position (11), the first extended position (13) and the second extended position (15).

This actuation mechanism (24) is configured for moving a high lift surface (22) relative to a wing body (20) and comprises:- a guiding linkage (26), configured for guiding a displacement of the high lift surface (22) relative to the wing body (20),- a rocker (28), configured to be moveable in rotation relative to the wing body (20);- a rotational actuator (30), comprising a housing (38) and an actuation element (40), the housing (38) being fastened to the rocker (28) and the actuation element (40) being moveable in rotation relative to the housing (38); and- an actuation linkage (32) configured for displacing the rocker (28) and the high lift surface (22) relative to the wing body (20) upon rotation of the actuation element (40) relatively to the housing (38).

Disclosed is a wing (3) for an aircraft (1), comprising a main wing (5) and a leading edge high lift assembly (6) comprising a slat (7), and a connection assembly (9) movably connecting the slat (7) to the main wing (5), such that the slat (7) is movable between a retracted position (11), a first extended position (13), and a second extended position (15), wherein the connection assembly (9) comprises an elongate slat track (17) that extends along a track longitudinal axis (19) between a first end (21) and a second end (23) and has an intermediate portion (25) between the first and second ends (21, 23), wherein the first end (21) and/or the intermediate portion (25) of the slat track (17) is mounted to the slat (7), wherein the slat track (17) is mounted to the main wing (5) by a roller bearing (27) such that the slat track (17) is movable along the track longitudinal axis (19), wherein the roller bearing (27) comprises a plurality of bearing units (34a-c), each of the bearing units (34a-c) being mounted to the main wing (5) and having a bearing surface (22) for engaging an engagement surface (35) provided on the slat track (17), wherein the plurality of bearing units (34a-c) comprises at least a first end roller unit (34a) and an intermediate roller unit (34b), wherein the plurality of bearing units (34a-c) are positioned such that, relative to the track longitudinal axis (19), the intermediate roller unit (34b) is positioned between the first end roller unit (34a) and the second end (23) of the slat track (17). The object to provide a wing having a beneficial ratio of aerodynamic performance and noise generated by extension of the slat, is achieved in that the slat track (17) and the roller bearing (27) are configured such that in the retracted position (11) and in the first extended position (13), the slat (7) rests against the main wing (5), whereas in the second extended position (15), a gap (37) is formed between the trailing edge of the slat (7) and the leading edge of the main wing (5).

The present invention relates to a valve apparatus for a fluid delivery system comprising a valve housing having a fluid flow path, a valve member, a valve actuator operably connected to the valve member and configured to move the valve member between first and second positions, and a thermally-activated locking mechanism which provides the valve apparatus with an additional passive safety feature which will automatically lock and inhibit movement of the valve member upon reaching a pre-determined activation temperature.

The invention relates to a drive assembly (2) for driving a movable flow device (4) of an aircraft. The drive assembly comprises a single actuator (10) comprising a longitudinal transmission element (48), a first support element (12) and a second support element (18). The longitudinal transmission element of the single actuator (10) is configured to be oriented transversely to an extending direction of the movable flow device (4), is configured to be mounted to a fixed support structure (6) of the aircraft, and is configured to act on the first and second support elements (12, 18) such that the movable flow device (4) is movable between extended or retracted positions relative to the fixed support structure (6). The invention also relates to a drive system, a high lift system, an aircraft wing, an aircraft and a method of actuating a movable flow device.

A mould for moulding a sleeve of protective material 25 onto a wiring harness 15. The mould comprises: a mould body 33; an inlet 37; an outlet 39; and a channel 35 extending through the mould body 33 and connecting the inlet 37 to the outlet 39. The mould also comprises a plurality of centralising spacers 41, configured to hold the wiring harness 15 centrally within the channel 35 as the wiring harness 15 extends between the inlet and the outlet. The mould also comprises an injection port and an exit port arranged to convey fluid sleeve material into and out of the channel. The channel 35 follows a path along a helical curve as it extends between the inlet and the outlet of the mould.

A cryogenic liquid storage system comprising: an outer tank (22); an inner tank (24) inside the outer tank and configured to store a cryogenic liquid; an isolation valve (40); a collection chamber (26) outside the outer tank and fluidically connected to the inner tank to enable fluid to flow from the inner tank to the collection chamber via the isolation valve, wherein the isolation valve can be actuated to isolate the collection chamber from the inner tank; and a measurement probe (50) movable between the inner tank and the collection chamber.

The invention relates to a process (9) for setting a phase between the engines (3) of a multi-engine aircraft (A) to minimize cabin noise and vibration, the aircraft (A) comprising a master engine (3a) and at least one slave engine (3b), each engine (3) being coupled to a fan or a propeller, the process (9) comprising the steps of:a)performing a first synchronization step (91) aimed at minimizing vibration; andb) performing a second fine-tuning step (92) aimed at minimizing noise.Such a process makes it possible to reduce, in a simple and reliable manner, the vibrations and noise emitted by a multi-engine aircraft.

The present invention relates to lavatories on board an aircraft. In order to provide lavatories with a facilitated use for persons with reduced mobility, a lavatory (10) of an aircraft cabin comprises at least a first lavatory space (12) with a toilet (14), a housing structure (16) enclosing the lavatory space at least partly and a bracket arrangement (18) for a temporary mount of an upper support (20) for a user sitting on the toilet. The housing structure comprises at least one housing segment (22) of the group of: a first lateral wall segment (24), a second lateral wall segment (26), a front wall segment (28), a rear wall segment (30) and a lavatory ceiling segment (32). The bracket arrangement comprises at least two brackets (34) that are fixedly attached on the at least one housing segment such that the upper support can be temporarily mounted in a support position such that it is arranged above a user's head when sitting on the toilet.

The present invention relates to a frame (10) for an aircraft fuselage. The frame (10) comprises a primary frame segment (12) and a secondary frame segment (14). The primary frame segment (12) provides a primary arc with two primary end sections (18) and comprises a primary inner surface (22) and a primary outer surface (24). The secondary frame segment (14) provides a secondary arc with two secondary end sections (20) and comprises a secondary inner surface (26) and a secondary outer surface (28). The primary frame segment (12) and the secondary frame segment (14) are configured to be couplable to each other such that their end sections (18, 20) partially overlap so that a portion of the secondary inner surface (26) of the secondary end sections (20) of the secondary frame segment (14) abuts on a portion of the primary outer surface (24) of the primary end sections (18) of the primary frame segment (12), thereby forming transition areas from the primary frame segment (12) to the secondary frame segment (14).

The present invention provides Aircraft fuselage (100), in particular of an open fan propeller aircraft (400), comprising a primary structure having frame (105) members extending in a circumferential direction (C) of the fuselage (100) and stringers (111) extending in a longitudinal direction (A) of the fuselage (100) and a ballistic shield (101) covering an inner surface of a longitudinal section (123) of the fuselage, wherein the ballistic shield (101) is inserted between adjacent frame (105) members of the primary structure and extends in a longitudinal direction (A) and a circumferential direction (C) of the fuselage (100) section, a ballistic shield (101) for use on an inner surface (103) of an aircraft fuselage (100), a ballistic shield (101) assembly and aircraft (400)..

An aircraft wing structure (100) comprising a spar (101) and a systems module (200) is disclosed. The systems module (200) is releasably secured to a plurality of brackets (300) on a first side of the spar (101) and the second, opposite side of the spar is configured to define a wall of a fuel tank of an aircraft wing. The systems module (200) is releasably secured to at least one bracket (300) of the plurality of brackets (300) by a fastener (207) comprising a bolt (211) and a barrel nut (209). The barrel nut (209) is configured to engage with a part of the bracket (300) to prevent rotation of the barrel nut (209) when rotating the bolt (211) within the barrel nut (209) to tighten or untighten the fastener (207). Methods of installing and removing the systems module are also described.

The invention relates to an overhead lid (34) for a gas-liquid separator device (18), the separator device (18) comprising a separation area (20) for separating a liquid (12) from a fluid (14) comprising a gas (16) flowing from a separator inlet (22) of the separator device (18) to a separator outlet (24) of the separator device (18), and a reservoir component (26) for collecting the liquid (12) from the separation area (20), the overhead lid (34) being configured for mounting on the gas-liquid separator device (18) and including a cavity area (36) for collecting the liquid (12) from the separation area (20) opposite to the reservoir component (26).

This vortex generator (16) comprises:- a base (22), configured to be connected to an external surface (S) of an aircraft (10),- a baffle (24) being connected to the base (22) and moveable in rotation around a rotation axis (R) relative to the base (22), and- an actuator, configured to move the baffle (24) in rotation around the rotation axis (R) relative to the base (22),the actuator (26) being controllable into:- at least one active configuration, in which the actuator (26) is moving the baffle (24) to a predetermined active position relative to the base (22), and- a passive configuration, in which the actuator (26) is allowing rotation of the baffle (24) relative to the base (22),wherein the baffle (24) is in a feathered position relative to the base (22) when the actuator (26) is in its passive configuration.

This vortex generating system (18) comprises: - a vortex generator (16) comprising + a base (22), + a baffle (24), moveable in rotation around a rotation axis (R) relative to the base (22), and + an actuator (26), configured to move the baffle (24) in rotation around the rotation axis (R) relative to the base (22); and - a control unit (20), connected to the vortex generator (16) and comprising: + an acquisition module (36), configured for acquiring a target active position of the baffle (24), + a control module (34), configured for controlling the actuator (26) into an active configuration, in which the actuator (26) is applying a controlling torque to the baffle (24) depending on the acquired target active position, and + an evaluation module (40), configured for generating an inconsistency evaluation signal as a function of the controlling torque and of the target active position.

A robot arrangement for interacting with an object, comprising: a plurality of robots adjacent an object; the plurality of robots configured to interact with the object sequentially such that: a first robot of the plurality of robots is configured to perform a first portion of a first task on the object and then, whilst the first robot is not interacting with the object, a second robot of the plurality of robots is configured to perform a second portion of the first task on the object, and the plurality of robots are configured to simultaneously interact with the object, before or after completion of the first task, so as to collectively perform a second task on the object.

The invention provides an aircraft wing (100) with a wing tip device (14) movable between a flight configuration for use during flight and a ground configuration for use during ground-based operations. In the ground configuration the wing tip device (14) is moved away from the flight configuration such that the span of the aircraft wing is reduced. A hinge arrangement (16) connects the fixed wing and the wing tip device. The hinge arrangement comprises a plurality of fixed wing interface rings (40), a plurality of wing tip device interface rings (41), and a geared rotary actuator (38, GRA) partially located within the fixed wing interface rings and wing tip device interface rings. The GRA is arranged to drive the wing tip device interface rings to move the wing tip device between the flight configuration and ground configuration.

A fixation assembly (60) and a fixation arrangement (10) for fixing a mounting structure (8) to a support structure (2) of a vehicle (1), in particular a fuselage of an aircraft, and a vehicle (1) are provided, the fixation assembly (60) comprising at least one fixation element (61) adapted for transferring operational loads (K) of the mounting structure (8) into the support structure (2), and at least one redundancy element (62) configured to transfer the operational loads (K) alternatively and/or additionally to the at least one fixation element (61).

A compliant mount for an aircraft, for attaching equipment (such as a latching unit, for example) to a structural member of the aircraft, the compliant mount comprising: a mounting member for attachment to the structural member of the aircraft; a receptacle for supporting the equipment, the receptacle being coupled to the mounting member along a longitudinal axis; and one or more spring elements arranged to resiliently bias the receptacle into alignment with the longitudinal axis and to allow angular movement of the receptacle away from the longitudinal axis. Also provided is an aircraft equipment assembly comprising such a compliant mount, and aircraft equipment supported by the receptacle, and an aircraft having one or more such equipment assemblies. The aircraft may be a folding wingtip aircraft and the equipment assembly may be part of a folding wingtip actuation mechanism.

Disclosed is a computer-implemented method. The computer-implemented method comprises receiving data from an acoustic sensor, the data indicative of a leak in a vacuum-insulated assembly of a cryogenic fluid system. The computer-implemented method also comprises causing an action to be performed based on the data.

Face milling a cover layer (14) of metal sheets (10) with a milling tool (21) configured to rotate around a rotational axis (R), having at least one milling face (31, 32) extending essentially perpendicularly to the rotational axis (R). A fixture assembly (23) fixes the metal sheet (10) and defines a working plane (E) with which at least flat sections (A) of the metal sheet (10) are supposed to be aligned, such that a surface normal (N) of the flat section (A) extends essentially perpendicularly to the working plane (E); wherein the rotational axis (R) is being tilted under at least one work angle (γ) with respect to the surface normal (N) of the flat section (A) during the milling process.

Disclosed are a fuel tank 100, 100", 1001, 1002 of or for an aircraft 1, and an aircraft 1 comprising such fuel tank. The fuel tank is configured to contain liquid hydrogen LH2, and it comprises at least one adjustable anti-sloshing baffle 110, 110', 110". The fuel tank may preferably comprise an adjustment mechanism 130, 130" adapted to adjust the at least one anti-sloshing baffle 110, 110', 110".Further disclosed is a fuel tank arrangement comprising such fuel tank 100, 100", 1001, 1002 with an adjustment mechanism 130, 130", and a control unit 200 for control of the adjustment mechanism 130, 130".

This conduit (22) for a heat exchanger (16) comprises: - an outer tube (40), comprising an outer inlet (40<i>) and an outer outlet (40<o>) and defining an outer passage (40<p>) configured for guiding a working fluid (W) from the outer inlet (40<i>) to the outer outlet (40<o>), and - an inner tube (42), comprising an inner inlet (42<i>) and an inner outlet (42<o>) and defining an inner passage (42<p>) configured for guiding the working fluid (W) from the inner inlet (42<i>) to the inner outlet (42<o>), wherein the inner tube (42) is arranged in the outer passage (42<p>) and defines with the outer tube (40) an annular space (A), the inner outlet (42<o>) opening in the outer passage (40<p>)

The invention relates to an in-situ thermoforming and (co-)consolidation, wherein a pre-deformation of a plastic-based component (2) in order to correspondent to a contour of a component (1) to be joint with can be omitted, as the adaption of the component (2) to the contour of the component (1) to be joint with is done directly before the joining step with the other component in the same tooling (12, 13) by adapting process parameters such as heat and pressure.

A landing gear assembly for an aircraft, the landing gear assembly comprising: a main body on which one or more wheels of the aircraft are mounted, the main body being aftwardly pivotable, about a pintle line, into a landing gear bay of the aircraft; and a bracing structure coupled to the main body, the bracing structure being configurable into an extended configuration to support the main body in a deployed configuration, and into a retracted configuration in which the main body is held in a retracted configuration. The bracing structure includes at least a first component of preferential weakness configured to fail when subjected to a critical load when the main body is in the deployed configuration, to thereby decouple the main body from at least part of the bracing structure and enable the main body to pivot aftwards, into the landing gear bay. Also provided is an aircraft (such as a blended wing body aircraft) having such a landing gear assembly.

This actuation system (22) comprises: - an actuation device (26), comprising a first end (30) configured to be connected to a main portion (18) and a second end (32) configured to be connected to a control surface (20), and - at least one hinging device (28) comprising at least one hinging member (34) and being configured for guiding a movement of the control surface (20) relative to the main portion (18), wherein the actuation device (26) is configured to move its first (30) and second ends (32) in translation along an actuator axis (A-A'), wherein the hinging device (28) defines a plurality of hinging axes (H-H') intersecting to define a pivot axis (P-P'), the pivot axis (P-P') being offset to the hinging device (28), the hinging device (28) being configured for guiding the movement of the control surface (20) in rotation around the pivot axis (P-P').