Airbus Operations Patente

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 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 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.

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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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.

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.

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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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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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.

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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).

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