Airbus Operations Patente

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).

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 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 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.

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.

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').

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".

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.

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>)

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 present disclosure relates to a spacer (100) for covering a gap between a sidewall (220) arranged in a longitudinal direction and a monument wall (200) arranged in a transverse direction in an aircraft. The spacer comprises an upper mount (111), and a lower mount (113), wherein the lower mount (113) restricts a relative movement between the monument and the spacer in the longitudinal direction, and allows a relative movement in the transverse direction.

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.

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.

The present invention relates to a oxide oxidation unit (600) for converting at least one reductant to thermal energy under the production of exhausts (G) and further exhausts (H) from a first supply flow (24) comprising an oxidant and a second supply flow (26) comprising the reductant, respectively; the solid oxide oxidation unit (600) comprising a duct wall (14) configured for separating the first supply flow (24) from the second supply flow and at least sectionwise comprising an electrolyte layer allowing for a transfer of ions of the oxidant to the reductant; wherein the duct wall (14) comprises an electrically conducting material (113) allowing for a transfer of electrons from the reductant to the oxidant for enabling a full oxidation of the reductant contained in the second supply flow (26). Furthermore, the present invention relates to a fuel cell arrangement (200) comprising at least one solid oxide oxidation unit (600) and at least one fuel cell setup (100) comprising a carrier structure (10) comprising a duct wall; wherein a primary power coating layer (28) is applied on a surface of the duct wall for being arranged between the first supply flow and the second supply flow; and wherein the primary power coating layer (28) is configured for generating electrical energy from the first supply flow and the second supply flow; to an energy supply system (300), comprising at least one solid oxide oxidation unit (600) and/or at least one fuel cell arrangement (200), and to an apparatus, in particular a vehicle, such as an aircraft (400), comprising same.

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The present disclosure relates to a water supply system (100) for a monument (11, 12, 13, 14) in an aircraft (1) and comprises a pressureless water buffer tank (110) configured to store a predetermined amount of potable water, a potable water consumer (120) connected to the pressureless water buffer tank via a water supply pipe (115), a grey water unit (130) configured to collect and store grey water from the water consumer, and a toilet (135). A pressurisation unit (137) is fluidly connected to the grey water unit and is configured to pressurise grey water from the grey water unit and to provide pressurised grey water to the toilet (135). The pressureless water buffer tank is installed in the monument above the potable water consumer, and the water supply pipe is configured to guide water from the pressureless water buffer tank to the potable water consumer only by gravity.

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Disclosed is a vehicle propulsion system (1000) comprising a propulsion device (201), a turbine (210) mechanically coupled to the propulsion device (201) and configured to drive the propulsion device (201), a fuel cell (110) having an exhaust outlet (111), and an exhaust duct (220) connecting the exhaust outlet (111) of the fuel cell (110) with the turbine (210). The turbine (210) is driven by exhaust gas of the fuel cell (110) conducted through the exhaust duct (220). Furthermore, a vehicle (1) having such vehicle propulsion system (1000) is also disclosed.

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A laminate structure comprising: a fibre laminate impregnated with a laminate matrix material; and a veil of carbon nanotubes impregnated with a veil matrix material. The laminate matrix material and the veil matrix material doped with carbon particles. The veil provides lightning strike protection. The structure is manufactured by co-curing the laminate matrix material and the veil matrix material to bond the veil of carbon nanotubes to the fibre laminate.

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Es wird eine Isolierungsanordnung (8) zum thermischen und akustischen Isolieren einer Wandung (12) eines Flugzeugs (42), aufweisend: eine erste, luftdichte Materialschicht (10), eine davon beabstandete zweite, luftdichte Materialschicht (14), mindestens ein zwischen der ersten Materialschicht (10) und der zweiten Materialschicht (14) angeordnetes, luftdurchlässiges Isolierelement (16), mindestens einen Lufteinlass (18), und mindestens einen Luftauslass (20), wobei das Isolierelement (16) dazu ausgebildet ist, zumindest stellenweise einen Abstand zwischen der ersten Materialschicht (10) und der zweiten Materialschicht (14) auf einen vorbestimmten Wert zu begrenzen, und wobei der mindestens eine Lufteinlass (18) und der mindestens eine Luftauslass (20) an voneinander beabstandeten Bereichen des Isolierelements (16) angeordnet sind und dazu ausgebildet sind, Luft durch den mindestens einen Lufteinlass (18) in das Isolierelement (16) einzuleiten und aus dem mindestens einen Luftauslass (20) wieder auszuleiten.

An aircraft wing structure comprises first and second unitary shell structures 7, 8, preferably formed from a fibre-reinforced composite material. Each shell structure 7, 8 comprises a partial front spar, a partial rear spar and a wing skin. The partial front spars of each shell structure and the partial rear spars of each shell structure are arranged to overlap to form front and rear spars respectively. A wing box structure 6 constructed according to the invention is lightweight, strong and durable, with a lower overall part count than has been conventionally achievable. Aircraft systems equipment may be attached to the shell structures 7, 8 prior to, or after, assembly into a wing box.

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The invention provides an aircraft structure dynamic loading testing device, including a nozzle 141, 142, an attachment mechanism comprising a first portion 124 attached to the nozzle such that a nozzle force direction 141c is constrained at a set orientation and a second portion 125 for attaching to the aircraft structure 110 and configured to be constrained at a given orientation with respect to the aircraft structure, and a length of conduit 121, 123, wherein, in use, fluid is supplied along the length of conduit to the nozzle and ejected from the nozzle such that the nozzle exerts a dynamic force on the aircraft structure in the direction of the nozzle force direction. The invention also provides a dynamic loading testing device, a kit of parts and methods of dynamic load testing of a structure.

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Disclosed is a drag reduction system (13) for an aircraft (1), comprising an aircraft component (3) comprising a skin panel (17) having an inner surface (19) and an outer surface (23) for contact with an ambient flow (25), wherein the outer surface (23) comprises an upstream area (27) for laminar flow and a downstream area (29) for turbulent flow and/or transitional flow, wherein the skin panel (17) comprises a plurality of micro pores (39) for blowing air from inside (21) the aircraft component (3) into the ambient flow (25). The object to provide a drag reduction system that can further reduce the drag of the aircraft component, is achieved in that the micro pores (39) are arranged in the downstream area (29) of the outer surface (23) of the skin panel (17).