Airbus Operations Patente

This slat cove filling device (24) for a slat (18) of an aircraft (10) comprises a sheltering panel (46) extending along a spanwise direction (Y) and comprising a proximal edge (50), configured for being fastened to a slat body (22) of the slat (18), and a distal edge (52), configured for being slidable relative to the slat body (22),the sheltering panel (46) comprising a plurality of guiding passages, each guiding passage being configured for cooperating with a guiding element of the slat body (22) for guiding the distal edge (52) along a sliding direction (S) relative to the slat body (22).

A fuel mixing assembly comprising, a fuel vessel, such as a pipe 15 for conveying flow of a first fuel; and a second-fuel conduit 17 for conveying flow of, for example an additive, into the fuel vessel 15. The fuel mixing assembly further comprises a vortex generator 21, the vortex generator 21 being configured to generate vortices to mix the first fuel and the additive. The additive may be released from an outlet in the vortex generator 21.

An aircraft maintenance system, method, and computer-readable storage medium are provided. The aircraft maintenance system comprises a processor and storage comprising a digital twin of at least part of an aircraft and a machine learning classifier. The aircraft maintenance system is configured to obtain aircraft operational data, update the digital twin based on the aircraft operational data, identify an altered operation of the aircraft based on the aircraft operational data, and determine a set of conditions responsible for the altered operation using the machine learning classifier to evaluate the digital twin. Output data indicative of the set of conditions responsible for the altered operation is generated.

A composite part for a vehicle is provided. The composite part has a first surface (22a) and an opposing second surface (22b), wherein at least part of the first surface forms an outer aerodynamic surface of a vehicle. A region of the part has a curvature in two orthogonal directions. The part has an undulation (40) defining a channel (60) in one of the first surface or the second surface, and a corresponding ridge (50) in the other of the first surface and the second surface.

Disclosed is an enclosure device configured to enclose a concave surface portion of an aerodynamic component of an aircraft wing. Disclosed also is, an aerodynamic component comprising an enclosure device, and an aircraft wing and aircraft comprising one or more enclosure devices. Also disclosed is a method of counteracting unwanted airflow circulation in a concave surface portion of an aerodynamic component of an aircraft wing.

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 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 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.

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 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).

Es werden ein Robotersystem (1) mit wenigstens einem Roboter (2) und wenigstens zwei Roboterarmen (5) sowie ein Verfahren zur Ausführung einer Arbeitsaufgabe mit einem Robotersystem (1) mit wenigstens einem Roboter (2) und wenigstens zwei Roboterarmen (5) vorgeschlagen, die dazu eingerichtet sind, mit wenigstens einem Endeffektor (6) eine Arbeitsaufgabe auszuführen, wobei ein adaptives Sockelmodul (4) an einem Untersatz (3) des wenigstens einen Roboters (2) dazu eingerichtet ist, die wenigstens zwei Roboterarme (5) an ihrer jeweiligen Basis (9a) zu halten und wenigstens zwei verschiedene an die jeweilige Arbeitsaufgabe angepasste Betriebszustände (A, B, C, D) anzunehmen.

Disclosed are a panel element, in particular for use as an aircraft interior part, comprising at least one plate component having a front side and a rear side, and a stiffening structure forming enforced areas of the plate component, wherein the stiffening structure comprises at least an edge protection frame surrounding outer edges of the at least one plate component, and local stiffening elements extending over adjacent surface areas of the front side and the rear side, wherein the stiffening structure is visible in a final state of the panel element on both sides, and an aircraft interior part comprising at least one panel element.

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.

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.

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.

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).

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).

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).

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 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.

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.

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.

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.

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 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.

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.

The method for assembling parts comprises the following steps: providing a first and second parts (1, 2) to be assembled; applying a shim and sealant material (3) on the first part (1) to be assembled; positioning the first part (1) on the second part (2) to be assembled, with the shim and sealant material (3) placed between the first and second parts (1, 2); detecting a contour of the first or second part (1, 2) at least partially covered by the other of the first or second part (1, 2); drilling one or more holes in the first and second parts (1, 2), based on the detected contour; and placing a fastener (6) in the or each drilled hole.It permits to provide a simplified assembling method that reduces these costs.

Disclosed is a free-space optical communications (FSO) terminal comprising: a first arm addressing a first port of a beamsplitter and comprising a first component configured to transmit a diverging first transmitted beam and/or utilise a first received beam; a second arm addressing a second port of the beamsplitter and comprising a second component configured to transmit a diverging second transmitted beam and/or utilise a second received beam; and a third arm addressing the third port of the beamsplitter and comprising a pointing unit configured to steer the first and/or second transmitted beam to a target and/or the first and/or second received beam received from a target. The third arm comprises an optical element to collimate transmitted beams to provide a collimated transmitted beam to the pointing unit, and focus received beams so as to provide a converging beam to the first/second arm. A method of beam steering is also disclosed.

Disclosed is a free-space optical communications (FSO) terminal. The FSO terminal comprises a first arm comprising a combined transmit-receive system configured to produce a data beam, encoding data, for transmission to a target and receive a received data beam from the target, a second arm comprising a component configured to transmit a further transmitted beam or utilise a further received beam, a pointing unit configured to steer the data beam to the target and/or the received data beam received from the target, and an optical fibre comprising a single-mode portion configured to carry the transmitted data beam and received data beam and couple the first arm to the pointing unit, and a multi-mode portion configured to carry the further transmitted beam or the further received beam and couple the second arm to the pointing unit. An FSO system and vehicle comprising an FSO terminal are also disclosed.

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Zusammenfassung wird geladen …

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