An aircraft assembly (25) comprising: a wing structure; and a pipe assembly (30a, 30b) rotatably coupled to the wing structure by a fixture arrangement (25), the pipe assembly extending along a longitudinal direction, the fixture arrangement configured to restrict movement of at least a portion of the pipe assembly in the longitudinal direction relative to the wing structure and allow rotation of the pipe assembly relative to the wing structure.
F16L 39/04 - Joints or fittings for double-walled or multi-channel pipes or pipe assemblies allowing adjustment or movement
F16L 39/00 - Joints or fittings for double-walled or multi-channel pipes or pipe assemblies
F16L 27/04 - Universal joints, i.e. with mechanical connection allowing angular movement or adjustment of the axes of the parts in any direction with partly-spherical engaging surfaces
F16L 3/16 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets with special provision allowing movement of the pipe
An aircraft assembly comprising: a rib (10) of a wing; an aperture (11) in the rib; a pipe assembly (30a, 30b) extending through the aperture, the pipe assembly comprising a first rigid pipe section (30a), a second rigid pipe section (30b) and a flexible pipe section (33) between the first and second rigid pipe sections; and a fixture arrangement (25) that couples one of the pipe sections to the wing rib (10), wherein the flexible pipe section (33) is located at a plane of the rib and configured to allow relative movement between the first rigid pipe section and the second rigid pipe section. [Figure 5]
A multi-material component is provided, comprising two members formed of a nickel-based superalloy and titanium joined together by a transition member. The transition member is formed to be complementary to both the nickel-based superalloy and titanium member. The transition member may comprise a solid layer of tantalum or BAu-4. The transition member may be joined to the two members by virtue of growing the members via additive manufacturing on the transition member. Accordingly the transition member and optionally the whole component may be obtained via additive manufacture. The multi-material component may be suitable for use on aircraft.
A multi-material component is provided, comprising two members formed of a nickel-based superalloy and a titanium-based material joined together by a transition member. The transition member is formed to be complementary to both the nickel-based superalloy and titanium-based material member. The transition member comprises a layer of titanium and Inconel that have been joined together by a low heat welding method, such as explosion welding, linear friction welding or rotary welding. The layers of the transition member may be separated by a diffusion barrier, which may be tantalum. The transition member may be joined to the two members by virtue of growing the members via additive manufacturing on the transition member. The multi-material component may be suitable for use on aircraft.
B23K 20/12 - Non-electric welding by applying impact or other pressure, with or without the application of heat, e.g. cladding or plating the heat being generated by friction; Friction welding
B23K 35/00 - Rods, electrodes, materials, or media, for use in soldering, welding, or cutting
B23K 35/32 - Selection of soldering or welding materials proper with the principal constituent melting at more than 1550°C
Disclosed is a brake component (151) comprising a composite material (200), the composite material comprising a matrix (202) and reinforcements (201) embedded in and supported by the matrix, wherein the matrix is a glass-ceramic matrix. Also disclosed is a braking system, an aircraft braking system, an aircraft and a method of operating an aircraft.
A computer-implemented method of determining a maintenance schedule or maintenance plan for a fuel tank of an aircraft. In a training phase, a predictive model is trained with training data. The training data is associated with fuel tanks of plural aircraft. The training data comprises environmental data indicative of an environmental parameter at locations of the plural aircraft, and test data which is indicative of a level of microbial contamination of fuel in fuel tanks of the plural aircraft. In a prediction phase, input data is received and the predictive model is operated, along with an optimisation algorithm, to determine a maintenance schedule or maintenance plan. One or more microbial contamination maintenance tasks are performed on the fuel tank based on the maintenance schedule or maintenance plan.
An aerostructure component (6) comprises a core (10) sandwiched between two skins (7, 8), each skin comprising at least one ply (e.g. 9a) of a plurality of elongate reinforcing elements in a metal matrix (36) and the core comprising a plurality of hollow spheres (11) in a metal matrix (37). The component (6) is inherently both lightweight and strong, prevents the ingress of moisture, contains embedded thermal insulation and is electrically conductive. Such properties make it suitable for the formation of a wing box cover or a wing box itself.
B64C 1/12 - Construction or attachment of skin panels
B22D 19/02 - Casting in, on, or around, objects which form part of the product for making reinforced articles
B22D 19/14 - Casting in, on, or around, objects which form part of the product the objects being filamentary or particulate in form
B32B 15/01 - Layered products essentially comprising metal all layers being exclusively metallic
B32B 15/02 - Layered products essentially comprising metal in a form other than a sheet, e.g. wire, particles
B32B 15/04 - Layered products essentially comprising metal comprising metal as the main or only constituent of a layer, next to another layer of a specific substance
B64C 1/26 - Attaching the wing or tail units or stabilising surfaces
A multi-material component is provided, comprising two members formed of different materials joined together at a multi-material join. The two members have complementary protrusions that form a multi-material join with a zig-zag interface, such that a tension between them is experienced in at least part as a shear force at the interface between the protrusions. The component further comprises a third member connected to the first member by a lattice, which is capable of elastic deformation so as to substantially isolate the interface between the complementary protrusions from deformation of the third member such as that caused by thermal expansion. Methods of designing and constructing such a component in a single piece via additive manufacturing are provided. The component is particularly suitable for use in pylons connecting jet engines to aircraft wings.
B32B 3/30 - Layered products essentially comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar form; Layered products essentially having particular features of form characterised by a layer with cavities or internal voids characterised by a layer formed with recesses or projections, e.g. grooved, ribbed
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
B64D 27/26 - Aircraft characterised by construction of power-plant mounting
9.
A METHOD OF DETERMINING AN OPERATING CONDITION OF A VALVE OF AN AIRCRAFT SYSTEM
Disclosed is a method of determining an operating condition of a valve of an aircraft system. The method includes obtaining a first time period associated with actuating the valve, and a second time period associated with actuating the valve. The method includes providing an indication of an altered operating condition associated with actuation of the valve based on the first and second time periods.
F02C 7/232 - Fuel valves; Draining valves or systems
F16K 37/00 - Special means in or on valves or other cut-off apparatus for indicating or recording operation thereof, or for enabling an alarm to be given
A first aspect of the present invention provides a load-bearing structure configured to, during operation of the structure, transfer load from a first part of the structure to a second part of the structure via a load path. The component comprises a matrix material, a plurality of longitudinal first reinforcing elements embedded in the matrix material, and a plurality of longitudinal second reinforcing elements embedded in the matrix material. The long axis of each first reinforcing element is substantially aligned with a first direction and the long axis of each second reinforcing element is substantially aligned with a second direction, the second direction being substantially perpendicular to the first direction. The structure has a predefined crack-propagation region configured to control the propagation of a crack in the structure. The crack-propagation region either comprises multiple first reinforcing elements and does not comprise any second reinforcing elements; or comprises multiple second reinforcing elements and does not comprise any first reinforcing elements.
An aircraft wing comprising: a wingbox; a fuel tank; a fuel cell system comprising a fuel cell; a fuel line configured to deliver fuel from the fuel tank to the fuel cell system; a propulsion system carried by the wingbox; and an electrical power line configured to deliver electrical power from the fuel cell system to the propulsion system. The fuel tank and the fuel cell system are located inside the wingbox, and the propulsion system is located outside the wingbox.
A tyre monitor (11) comprises a sensor device (12) arranged to detect an operating parameter of a tyre, such as tyre pressure or temperature. A resiliently flexible housing (14) comprising an elastic material is arranged to substantially surround the sensor device. The tyre monitor (11) is able to move freely when installed in a tyre. The elastic, resiliently flexible housing (14) around the sensor permits the monitor to be easily installed inside a tyre and to bounce, roll and travel inside it as the tyre moves. Thus, the monitor can directly sense the operating parameter within the tyre itself. The monitor may also include a wireless communications device to enable data from the sensor to be shared with crew.
Checking the operating parameters of a tyre, such as tyre pressure, is an important part of tyre maintenance, particularly in aircraft. A tyre monitor (11) comprises a sensor device (12) arranged to detect an operating parameter of a tyre and a protective housing (14) for the device, the device being retained on a ring (15) having spaced end portions (16a, 16b), the ring being resiliently radially biased. The provision of a retaining ring facilitates installation of the tyre monitor within a tyre assembly, so that the sensor device can directly monitor the operating parameter of the tyre. The protective housing protects the sensor device from impact, vibration and temperature extremes.
There is provided a sensor apparatus for measuring a property of a wheel assembly comprising a tire mounted on a wheel. The sensor apparatus comprises at least one sensor configured to acquire measurement data; and a mounting member attached to the at least one sensor. The mounting member is configured to engage with a wheel such as to retain the sensor on the wheel within an enclosed space formed by the wheel and a tire mounted on the wheel. The mounting member is further configured to enable the sensor to move around the circumference of the wheel when the mounting member is engaged with the wheel.
It has been proposed to detect the operating pressure of a tyre by installing a tyre monitor within the tyre, together with a wireless communication interface arranged to transmit readings from the tyre monitor to an external reader. However, such a system needs to be energised by batteries, which have a limited life. The invention provides a tyre monitor (26) comprising a sensor device (12) configured to operate within an enclosed space formed by the wheel and the tyre, and an energy harvester unit (28, 29). The energy harvester is arranged to convert a first type of energy experienced by the harvester unit in use (e.g. mechanical energy, thermal energy) into electrical energy to energise the sensor device (12). The energy harvester may be, for example, a piezoelectric device or a thermoelectric generator. In use, mechanical stresses experienced by the tyre monitor are converted into electrical energy which can be used to energise the sensor as a supplement to energy from the battery (18) within the tyre monitor, thus prolonging its service life.
There is provided a sensor device for use in determining at least one property of a wheel assembly comprising a tire mounted on a wheel. The sensor device is configured to be attachable to an outer circumferential surface of the wheel which faces an inner circumferential surface of the tire. The sensor device is configured to measure a distance between the sensor device and an object remote from the sensor device.
A flow control device (60) on a structure such that strain in the structure is at least partially transferred to the flow control device. The flow control device has at least two states, or shapes, separated by an elastic instability region. The flow control device is arranged to rapidly transition, or snap through, from the first state to the second state when strain in the structure exceeds an activation threshold of the flow control device (60). In an example, a spoiler (12) on an aerofoil has a rest position where it is substantially flush with the low pressure surface and an activated position where it protrudes from the low pressure surface and modifies the airflow over that surface. The spoiler (12) bends to move from the rest position to the activated position when the strain in the aerofoil crosses a threshold. The deployed spoiler (12) reduces the lift on the aerofoil, acting to reduce the lift induced strain of the aerofoil to which the spoiler is attached.
A method of manufacturing an aircraft aerofoil such as a winglet. The aircraft aerofoil comprises: a first cover, a second cover, an aerofoil leading edge, and an aerofoil trailing edge where the first and second cover are joined at a trailing edge interface. The first cover is formed by: a lay-up step in which a plurality of fibre plies are laid up to obtain a preform, wherein the preform has a thickness and a preform trailing edge, and the preform is laid up during the lay-up step with a ramp of decreasing thickness where the thickness of the preform decreases towards the preform trailing edge; and a curing step in which the preform is cured to form the first cover, wherein the ramp of decreasing thickness cures during the curing step to form a contact surface of the first cover. The trailing edge interface is assembled by engaging the contact surface of the first cover with a contact surface of the second cover.
B29C 70/30 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
An aircraft (101) comprising: one or more main engines (104a, 104b); a first fuel in a first fuel tank (111a) configured to be fed towards at least one of the one or more main engines through a first fuel line; and a second fuel different to the first fuel in a second fuel tank (111b) configured to be fed towards at least one of the one or more main engines through a second fuel line; the first fuel being anhydrous ammonia and the second fuel being a fuel other than anhydrous ammonia.
A cover panel for an aircraft, comprising: a first region, a second region for forming an overlap with a second panel, and a ramp, wherein at least a portion of the ramp is between the first and second regions, wherein the ramp includes a tapered region such that a width of the ramp tapers towards a free-edge of the first region, and wherein the free-edge is configured to form a butt-joint with the second panel.
A winglet (10) is disclosed, with first and second covers (20, 22), a front spar (24), a rear spar (26), and a mid spar (28) between the front spar (24) and the rear spar (26). Each spar is joined to the first cover (20) and to the second cover (22). The mid spar (28) has an I- shaped cross-section comprising a mid spar web (29), a first pair of mid spar caps (29a, 30a) joining the mid spar (28) to the first cover (20), and a second pair of mid spar caps (29b, 30b) joining the mid spar (28) to the second cover (22). The mid spar (28) has a length and the mid spar (28) curves along all or part of its length. A winglet cover assembly (100) is also disclosed, with a cover (20) and a spar (28) formed from fibre-reinforced composite material. The spar (28) has an l-shaped cross-section comprising a spar web (29), a first pair of spar caps (29a, 30a) joining the spar web (29) to the cover (20), and a second pair of spar caps (29b, 30b); and the cover (20) is bonded to the first pair of spar caps (29a, 30a).
B64C 23/06 - Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
B64F 5/10 - Manufacturing or assembling aircraft, e.g. jigs therefor
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
Disclosed is an assembly comprising a bearing having an axis of rotation, and a fibre-based sensor for sensing strain or temperature of the bearing. The sensor extends in a direction parallel to the axis of rotation. Also disclosed is an aircraft system comprising a wheel supported on an axle by a first bearing and a second bearing. The system further comprises a first fibre optic sensor for sensing a strain or temperature of the first bearing, a second fibre optic sensor for sensing a strain or temperature of the second bearing, and an interrogator to analyse optical signals from the sensors to determine differences in the strains or temperatures of the first bearing and the second bearing.
G01H 9/00 - Measuring mechanical vibrations or ultrasonic, sonic or infrasonic waves by using radiation-sensitive means, e.g. optical means
G01L 1/24 - Measuring force or stress, in general by measuring variations of optical properties of material when it is stressed, e.g. by photoelastic stress analysis
G01D 5/353 - Mechanical means for transferring the output of a sensing member; Means for converting the output of a sensing member to another variable where the form or nature of the sensing member does not constrain the means for converting; Transducers not specially adapted for a specific variable using optical means, i.e. using infrared, visible or ultraviolet light with attenuation or whole or partial obturation of beams of light the beams of light being detected by photocells influencing the transmission properties of an optical fibre
G01B 11/16 - Measuring arrangements characterised by the use of optical techniques for measuring the deformation in a solid, e.g. optical strain gauge
G01K 11/32 - Measuring temperature based on physical or chemical changes not covered by group , , , or using changes in transmittance, scattering or luminescence in optical fibres
G01D 5/48 - Mechanical means for transferring the output of a sensing member; Means for converting the output of a sensing member to another variable where the form or nature of the sensing member does not constrain the means for converting; Transducers not specially adapted for a specific variable using wave or particle radiation means
G01K 11/3206 - Measuring temperature based on physical or chemical changes not covered by group , , , or using changes in transmittance, scattering or luminescence in optical fibres at discrete locations in the fibre, e.g. using Bragg scattering
23.
AIRCRAFT SYSTEM WITH ELECTRICAL CONNECTION APPARATUS
Disclosed is an electrical connection apparatus for providing an electrical connection between a first structure of an aircraft system and a second structure of an aircraft system, the second structure movable relative to the first structure. The electrical connection apparatus has a female connector, and a male connector at least partially housed within the female connector such that telescoping motion is enabled between the male and female connectors. The male and female connectors have respective electrical contacts that are contactable with one another. The male and female connectors have respective constant radii of curvature along their respective lengths, and the constant radii of curvature share a common centre of curvature.
An aircraft system has a first structure, and a second structure coupled to the first structure and movable between first and second positions relative to the first structure. The aircraft system has an electrical connector for providing an electrical connection running between respective components housed within the first and second structures. The electrical connector has a cable harness housed within the first structure, and a connector body coupled to an end of the cable harness. The connector body extends through an aperture formed in the first structure, and the connector body is coupled to the second structure such that movement of the second structure between the first and second positions relative to the first structure causes the connector body to move through the aperture.
There is provided an aircraft assembly comprising a wing component and an engine mounting pylon component. The wing component comprises three fastening locations arranged in a triangle such that they define a plane. The engine mounting pylon component is connected to the wing component by three tension fasteners, wherein each of the tension fasteners passes through a different one of the fastening locations. The wing component comprises three first spherical surfaces positioned such that the centre of each first spherical surface is at a different one of the fastening locations. The engine mounting pylon component comprises three second spherical surfaces having equal and opposite curvature to the first spherical surfaces. Each of the second spherical surfaces is in contact with a different one of the first spherical surfaces.
B64D 27/26 - Aircraft characterised by construction of power-plant mounting
F16B 43/02 - Washers or equivalent devices; Other devices for supporting bolt-heads or nuts with special provisions for engaging surfaces which are not perpendicular to a bolt axis or do not surround the bolt
An apparatus includes an outlet configured to output hot air, a flow control device movable between a first position and a second position, and a phase change actuator. The phase change actuator is connected to the flow control device and configured to move the flow control device between the first position and the second position at a predetermined temperature where a change of phase of a working substance of the phase change actuator occurs. In the second position, the flow control device may be configured to interact with an airflow and/or the hot air so as to protect a structure downstream of the outlet from overheating. The flow control device may be a vortex generator. In the second position, the vortex generator may be configured to interact with an airflow and/or the hot air from the outlet so as to redirect and/or reduce the temperature of the hot air.
A method of landing an aircraft, the aircraft comprising a first aircraft, the first aircraft comprising a landing system. The landing system is operated to: receive information from a second aircraft via a direct aircraft-to-aircraft communication from the second aircraft to the first aircraft; determine a landing plan of the first aircraft based on the information; and land the first aircraft based on the landing plan of the first aircraft. The use of a direct aircraft-to-aircraft communication (i.e. a communication from the second aircraft to the first aircraft which does not travel via an intermediary such as a land station, air traffic controller or satellite) makes the method reliable because such a communication is inherently secure and difficult to hack. The first aircraft is a member of a virtual fleet of aircraft. The landing system of the first aircraft is operated to: receive landing plans of all other aircraft of the virtual fleet; and input the landing plans of all other aircraft of the virtual fleet into an optimisation algorithm which determines a set of landing plans for the virtual fleet of aircraft which is optimised for the entire virtual fleet, the optimised set of landing plans for the virtual fleet of aircraft including the landing plan of the first aircraft.
An aircraft fuselage comprising a fuselage shell; and a fuel tank outside the fuselage shell, wherein the fuselage shell has a double-lobed cross section and at least part of the fuel tank is located in a channel between the two lobes.
An aircraft comprising: a fuselage having a pressure shell, a fairing, and an unpressurised space between the pressure shell and the fairing. One or more fuel lines extend inside the unpressurised space from the fuel tank to an engine. The fairing may have a window in line with a window of the pressure shell. The fuel tank may be configured to carry a cryogenic fuel such as liquid hydrogen.
An apparatus attached near the tip end of a cantilevered aircraft wing for controlling the loads on the wing. The apparatus comprises a first body having a mass rotatable about a first axis by an actuator device; and a second body adjacent to the first body and having a mass rotatable about a second axis by an actuator device. The first and second bodies each have a centre of mass offset from their respective axes and are configured to counter rotate about their respective axes.
An apparatus and method for adjusting the position of a tool relative to a component using an x-ray backscatter device, including an x-ray backscatter device that is positioned adjacent a first side of a component and configured to emit x-rays towards the component and receive backscattered x-rays. A computer receives information relating to the backscattered x-rays from the x-ray backscatter device. A tool is located adjacent the first side of the component, and is moved to a position adjacent the first side of the component based on a datum determined from the information received by the computer determined datum.
Disclosed is an aircraft control system (100) comprising an input interface (102), an output interface (114) and a processing engine (108) comprising a classifier (110). The classifier (110) applies input data (104) generated by the input interface (102) to the classifier (110) to generate output control data (112). The classifier (110) has a plurality of parameters which represent a control policy for operating the aircraft (800). The output interface (114) generates control outputs to control the aircraft (800) based on the output control data (112). Also disclosed is machine learning system (900) for training the classifier (110). The machine learning system (900) comprises an environment (902), a pathway evaluation engine (904), storage (906), and a training engine (908). The machine learning system (900) generates training data (912) by selecting a pathway representing an operating procedure using the pathway evaluation engine (904). The training engine (908) trains the classifier (110) using the training data (912).
Disclosed is a monitoring system for a speed determination system in an aircraft and a method, which may be implemented by a computer program, of monitoring a speed determination system for an aircraft. The speed determination system comprises a first speed measurement system in form of a GPS system, and a second speed measurement system based on the integration of the output of an accelerometer of the IRS system. The second speed measurement system is associated with a predetermined behaviour characteristic, leading to a constant increase of the derived speed value caused by the integration of a bias value of the accelerometer. The monitoring system comprises a processor arranged to receive speed data provided by the speed determination system and to perform a correspondence determination process comprising processing the received speed data to determine whether a correspondence condition is satisfied. This correspondence condition comprises a correspondence between the received speed data and the predetermined behaviour characteristic of the second speed measurement system. In response to determining that the correspondence condition is satisfied, the monitoring system determines that the first speed measurement system is in an error condition, So that it cannot regularly correct the errors present in the output of the second speed measurement system.
G01P 21/02 - Testing or calibrating of apparatus or devices covered by the other groups of this subclass of speedometers
G01P 7/00 - Measuring speed by integrating acceleration
G01S 19/47 - Determining position by combining measurements of signals from the satellite radio beacon positioning system with a supplementary measurement the supplementary measurement being an inertial measurement, e.g. tightly coupled inertial
G01C 21/16 - Navigation; Navigational instruments not provided for in groups by using measurement of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
The present application relates to a duct stringer assembly (25, 26, 40, 41, 42) with a bulkhead (50). The duct stringer assembly has duct walls providing a duct with a closed cross-section. The duct is adapted to transport fluid. A bulkhead is in the duct, and the bulkhead is adapted to block the flow of fluid along the duct. The bulkhead has a bulkhead body, and a gasket sealing a gap between the bulkhead body and the duct walls.
F16L 55/132 - Means for stopping flow in pipes or hoses by introducing into the pipe a member expandable in situ introduced axially into the pipe or hose the closure device being a plug fixed by radially deforming the packing
35.
SPOILER ACTUATION APPARATUS FOR MOVING AN AIRCRAFT SPOILER
The present invention relates to the provision of a spoiler actuation apparatus for moving an aircraft spoiler. The spoiler is moveable between a stowed configuration and a deployed configuration. The spoiler actuation apparatus comprises a guide member (210), a rack (220) mounted on the guide member and slideable along a longitudinal axis of the guide member, and a gear (240) coupled to the rack, the gear arranged to move the spoiler (130) in response to sliding of the rack. The rack is held at a first position when the spoiler is in the stowed configuration. An actuator moves the rack from the first position to a second position along the longitudinal axis. When the rack is at the second position, the rack is operable to accelerate relative to the guide member away from the second position by an aerodynamic force acting on the spoiler.
The present invention relates to an aircraft (100) comprising an active drag control system such as a Laminar Flow Control (LFC) system having a port LFC apparatus (118) and a starboard LFC apparatus (120). The aircraft comprises a control system (126) that is configured to test how efficiently the LFC system is working by differentially operating the port LFC apparatus (118) and the starboard LFC apparatus (120), for example by deactivating either LFC apparatus (118, 120), and measuring the effect on the direction of flight of the aircraft (100). The control system (126) is also configured to change the direction of the aircraft (100), and trim the aircraft (100), by differentially operating the port LFC apparatus (118) and the starboard LFC apparatus (120)..
An aircraft wing comprising a first structure connected to a second structure such that the first and second structures are relatively moveable between a retracted configuration in which a contact edge of the first structure abuts the second structure, and a deployed configuration in which a gap exists between the contact edge and the second structure. The contact edge is pivotable relative to another part of the first structure. The first structure comprises a lever and the second structure comprises a stop, which are mutually configured such that a first part of the lever contacts the stop in the retracted configuration. Urging the first part and the stop against each other causes a second part of the lever to transmit a rotational force to the contact edge, the direction of the rotational force being such that the contact edge is pressed against the second structure.
A T-shaped assembly may have a central noodle to fill the void between its constituent parts. By placing pins on the surface of the noodle the connection between the parts penetrated and the noodle is enhanced and the strength of the assembly is increased. Such an assembly may be suitable for introducing out-of-plane loads to a composite panel for use as a rib or stiffener in an aircraft skin section. The noodle may be of unitary construction or comprise a plate having pins and a noodle body. The plate or noodle may be formed of Titanium and obtained by additive manufacturing.
B29C 65/56 - Joining of preformed parts; Apparatus therefor using mechanical means
B29C 65/64 - Joining a non-plastics element to a plastics element, e.g. by force
B29C 65/00 - Joining of preformed parts; Apparatus therefor
B29C 70/02 - Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising combinations of reinforcements and fillers incorporated in matrix material, forming one or more layers, with or without non-reinforced or non-filled layers
B29C 70/42 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29K 105/06 - Condition, form or state of moulded material containing reinforcements, fillers or inserts
39.
AIRCRAFT WING WITH TRAILING EDGE FLIGHT CONTROL SURFACE
An aircraft wing (2) having a main wing (3) and a trailing edge flight control surface (4) movable between a retracted position, a first extended position in which the control surface is positioned rearwardly in the chord wise direction relative to its retracted position, and a second extended position in which the control surface is rotated relative to its retracted position. A closure panel (12), mounted to the main wing (3), extends from the main wing to the control surface (4), to provide an air flow surface between the main wing and control surface, both when the control surface is in its retracted position and its first extended position. The closure panel (12) is movable, relative to the control surface (4), to an open configuration in which it opens an airflow passage (40) provided between the control surface (4) and an opposed surface of the aircraft wing when the control surface (4) is in its second extended position.
The invention provides an aircraft wing (102, Figure 2) comprising a leading edge slat (110') and a fixed aerofoil portion (104). The leading edge slat (110') is moveable between a stowed position and deployed position. In the deployed position a trailing edge of the leading edge slat comprises a sealed portion (112) and an unsealed portion (114). The sealed portion (112) forms a seal with the fixed aerofoil portion and the unsealed portion provides an airflow gap between the leading edge flap and the fixed aerofoil portion. The invention provides substantially the same benefits as a sealed leading edge slat whilst improving stall control behaviour in local zones around the unsealed portion.
Disclosed is an aircraft system (10) for an aircraft (1), the aircraft system (10) having a controller (100) that is configured to receive at least one signal during a landing procedure of the aircraft (1) and to determine that the aircraft (1) is at a predetermined stage in the landing procedure, on the basis of the at least one signal. The predetermined stage in the landing procedure is before a command to extend at least one landing gear (600) of the aircraft (1) is issued during the landing procedure. The controller (100) is configured to cause initiation of at least a part of a procedure to interrogate an aircraft braking system (200), on the basis of the determination.
Disclosed is a method of detecting the presence of an external aerial vehicle in the vicinity of an aircraft in flight. The method comprises receiving image data representing a first image captured by a first aircraft-mounted image sensor having a first field of view and processing the image data to determine whether an external aerial vehicle candidate is present in a target space of the first captured image; and receiving image data representing a second image captured by a second aircraft-mounted image sensor having a second field of view, which encompasses the target space, and processing the image data to determine whether the external aerial vehicle candidate is present in the second captured image. Finally, the method comprises generating a signal indicating that an external aerial vehicle is determined to be present in the vicinity of the aircraft if the external aerial vehicle candidate is determined to be present in both the first captured image and in the second captured image.
Disclosed is a method of analysing a surface of an aircraft. The method comprises receiving image data representing a first image of a surface region of an aircraft, the image data captured by a first aircraft-mounted image sensor of a plurality of aircraft-mounted image sensors, the first image sensor having a field of view comprising said surface region. The image data is processed to identify the presence of any unexpected surface features in the first image of the surface region; and responsive to determining the presence of an unexpected surface feature, generating an output indicating the presence of the unexpected surface feature.
A method of removing hydraulic fluid from an aircraft hydraulic system. The hydraulic system comprises a hydraulically actuated mechanism that is actuated by an electrohydraulic servo valve, a hydraulic fluid port through which hydraulic fluid can escape, and a hydraulic fuse with a closed state and an open state between the electrohydraulic servo valve and the hydraulic fluid port. The hydraulic fluid port is opened, and then the activation of the electrohydraulic servo valve is controlled to force hydraulic fluid to escape from the hydraulic system via the hydraulic fluid port, the control being so that the hydraulic fuse does not enter and remain in the closed state.
A method of operating a media scanner to protect a target machine from malware on a removable storage device. The target machine and the removable storage device each comprise a respective data line, and the media scanner comprises a data switch. Data is scanned on the removable storage device with malware detection software of the media scanner via a first data path, the first data path comprising the data line of the removable storage device connected to the data line of the media scanner by the data switch in a first switching state. After the data has been scanned with the malware detection software, the data switch is operated to switch from the first switching state to a second switching state, thereby disconnecting the data line of the removable storage device from the data line of the media scanner and connecting the data line of the removable storage device to the data line of the target machine. Data is transferred from the removable storage device to the target machine via a second data path, the second data path comprising the data line of the removable storage device connected to the data line of the target machine by the data switch in the second switching state. After the device has been removed, the data switch is returned back to its first switching state.
G06F 21/56 - Computer malware detection or handling, e.g. anti-virus arrangements
G06F 21/85 - Protecting input, output or interconnection devices interconnection devices, e.g. bus-connected or in-line devices
G06F 21/81 - Protecting specific internal or peripheral components, in which the protection of a component leads to protection of the entire computer by operating on the power supply, e.g. enabling or disabling power-on, sleep or resume operations
Disclosed is a speed determination system for an aircraft, comprising one or more interfaces arranged to receive first speed data from a first speed measurement system and second speed data from a second speed measurement system. The first speed measurement system is arranged to provide the first speed data based on a first speed measurement based on global positioning system data. The second speed measurement system is arranged to provide the second speed data based on a second speed measurement. The speed determination system comprises a processor arranged to determine whether the data received from the first speed measurement system is reliable. If global positioning data is determined to be reliable, the speed determination system determines a speed of the aircraft from the first speed data and determines one or more correction values for the second speed measurement system; and if global positioning data is determined to be unreliable, the speed determination system determines a speed of the aircraft from the second speed data and the determined one or more correction values.
Disclosed is a heading control system for an aircraft. The heading control system is arranged to maintain a heading of an aircraft by controlling a nose wheel angle of the aircraft. The heading control system comprises an interface arranged to receive a bias signal indicating a bias towards the port or the starboard of the aircraft and one or more processors. The one or more processors are arranged to determine, based on the bias signal, an offset angle defining an offset from a longitudinal axis of the aircraft and to perform a control process to control the nose wheel angle within an angular range based on the offset angle.
Disclosed is test apparatus (100) for testing at least part of an aircraft system. The test apparatus (100) is configured to obtain first data (210) comprising a plurality of input values and process the input values using a classifier (230) configured to model an operation of the at least part of an aircraft system (220). Second data is obtained, comprising a set of output values (240b). The output values (240b) from the second data are processed with output values (240a) generated by the classifier to identify differences between operation of the at least part of an aircraft system (220) and the operation as modelled by the classifier (230).
Disclosed is an aircraft controller (5), aircraft braking system (13) and method for determining braking parameters. The aircraft controller (5) is configured to determine information associated with an aircraft taxiing route associated with an aircraft (1). On the basis of the information, the aircraft controller (5) determines a first braking parameter for a first main landing gear (8) of the aircraft (1) and a second braking parameter for a second main landing gear (9) of the aircraft (1) during a landing event. The first and second braking parameters are determined to result in asymmetrical braking between the first (8) and second (9) main landing gears during the landing event.
A packaged bearing unit (12) includes an outer surface for receiving an aircraft landing gear wheel (10) and an inner surface for receiving an axle (14). The unit may comprise first and second spaced-apart sets (18a), b of tapered roller bearings, held between inner and outer raceways (20, 22). There may be a bearing setting spacer (52) which can be used when setting and pre-loading the bearings. The bearing unit may (12) be a sealed unit which includes a preset amount of lubricant (grease 50) for lubricating the bearings. One or more sensors (56, 58) may be retained in a void (54) between the sets of bearings (18). The bearing unit (12) may remain fixed to the axle (14) when the wheel (and optionally brake packs 28) are removed. The bearing unit (12) may be serviced less frequently than the wheels (10), and may thus have a longer lifetime, possibly comparable with the lifetime of the major components of the landing gear.
F16C 19/38 - Bearings with rolling contact, for exclusively rotary movement with bearing rollers essentially of the same size in one or more circular rows, e.g. needle bearings for both radial and axial load with two or more rows of rollers
There is provided a component life prediction system configured to predict the remaining life of a consumable component installed on an aircraft. The system comprises a processor configured to: receive wear information indicative of a current wear state of the component; receive schedule information indicative of a future flight schedule for the aircraft; and determine the remaining life of the component based on the received wear information and the received schedule information.
The present application relates to an aircraft assembly. The aircraft assembly has a first structural component and a second structural component. A fastener fastens the first component to the second component. The first structural component comprises a body and an insert in the body. The insert has a machined hole through which the fastener extends. The material hardness of the insert is lower than the material hardness of the body.
Disclosed is an aircraft system (10) for an aircraft (1). The aircraft system comprises a controller (100) that is configured to receive at least one signal during a take-off procedure of the aircraft. The signal may comprise information representative of at least one parameter of the aircraft. The controller is configured to determine whether a current or future aircraft climb rate associated with the take-off procedure meets a criterion, on the basis of the at least one signal. The controller is configured to determine at least one remedial action to be taken, such as performance of at least a portion of a procedure to retract a landing gear of the aircraft, when the controller determines that the aircraft climb rate does not meet the criterion.
Retraction of a landing gear assembly (4) on an aircraft (1) is actuated by a hydraulic actuator (12). The actuator (12) includes a piston (28) that travels within a cylinder (26) along a stroke length between a first position corresponding to the landing gear assembly when extended and a second position corresponding to the landing gear assembly when retracted. Movement of the piston (28) along its stroke length is snubbed at one end by a different amount according to the direction of travel, for example by use of an orifice plate (Fig 10; Fig 11) that has a discharge coefficient that is greater in one direction (Fig 13) than in the opposite direction (Fig 12). Asymmetric snubbing is thus provided, which enables the landing gear to retract faster.
The present invention provides an aircraft hydraulic actuation system for retracting an aircraft landing gear. The actuation system comprises a supply line arranged to carry hydraulic fluid pressurised by a pump, a return line arranged to return hydraulic fluid to a reservoir, and a hydraulic actuator 128. In a first mode of operation, a first chamber 130 of the actuator 128 is supplied with pressurised hydraulic fluid from the supply line such that a piston 134 is moved in a first direction so as to move a load such as a landing gear. In a second mode of operation, the first chamber 130 is taken out of fluid communication with the supply line and a second chamber 132 is in fluid communication with the return line, such that the piston 134 is able to be moved under the influence of the load, for example when the landing gear extends under gravity.
A first aspect of the present invention provides an aircraft assembly comprising a wing and an engine mounting pylon. An aft end of the engine mounting pylon is connected to the wing by a spigot and at least one fastener. The aircraft assembly is configured such that, during operation of the aircraft assembly on an aircraft, the spigot transfers only lateral load between the engine mounting pylon and the wing and the at least one fastener transfers only vertical load between the engine mounting pylon and the wing.
Disclosed is a gap filler (1), system and method for at least partially filling a gap (10) between at least two aircraft components (2, 12). The gap filler comprises a first component (3) and a second component (4) which comprises a channel (5). The first component is configured to be locatable to a variable extent within the channel, so that a position of the first component relative to the second component can be selected so that the gap filler at least partially fills the gap.
Disclosed is an avionics unit for an aircraft. The unit comprises a communications interface arranged to interface with one or more other avionics units. The communication interface is arranged to transmit a signal having a first state indicating a fault condition of the avionics unit to the one or more other avionics units if the first state persists for a predetermined time. A processor is arranged to modulate an output at the communication interface to provide a modulated signal to indicate a characteristic of a present state of the avionics unit with respect to the fault condition. The modulated signal comprising the first state and a second state, different from the first state, wherein, in the modulated signal, the first state has a duration less than the predetermined time
Disclosed is an apparatus for monitoring the condition of a pneumatic tyre. The apparatus comprises at least one pressure sensor to measure an internal pressure within the tyre. At least one speed sensor is provided to measure a rotational speed of the tyre. A controller is configured to receive data from the pressure sensor and the speed sensor and output indications related to a condition of the tyre. The controller determines the output indications by tracking the internal pressure with respect to time during use of the tyre, identifying at least one change in internal tyre pressure in a time interval. The time interval is between a first reference time and a second reference time, the reference times corresponding to points at which the speed sensor indicates different rotational speeds of the tyre. The controller further monitors the pressure change as a function of rotational speed of the tyre to provide an indicator of tyre condition. A method of monitoring wear in a pneumatic tyre.
Disclosed is an aircraft brake control system for controlling a plurality of brakeable wheels of a landing gear. Each brakeable wheel comprises a brake actuator and a wheel speed sensor. The system comprises a controller configured to receive aircraft control parameters and provide brake commands to the brake actuator of each wheel. The controller being configured to activate pre-retraction braking in response to an aircraft control parameter indicating that landing gear retraction is required and execute a functional brake test during pre-retraction braking. A method of operating an aircraft braking system is also disclosed.
B60T 8/32 - Arrangements for adjusting wheel-braking force to meet varying vehicular or ground-surface conditions, e.g. limiting or varying distribution of braking force responsive to a speed condition, e.g. acceleration or deceleration
B60T 8/00 - Arrangements for adjusting wheel-braking force to meet varying vehicular or ground-surface conditions, e.g. limiting or varying distribution of braking force
B60T 13/00 - Transmitting braking action from initiating means to ultimate brake actuator with power assistance or drive; Brake systems incorporating such transmitting means, e.g. air-pressure brake systems
An aircraft landing gear support (1) is provided comprising a support member (such as a landing gear rib (2)) and a plurality of pintle supports (5, 7, 8) (typically in the form of lugs) which form a pintle support arrangement for holding a pintle on which a landing gear assembly may be rotatably supported. The lugs are removably attached to the gear rib.
Disclosed is a method of operating a landing gear controller for an aircraft. The method comprises the landing gear controller receiving one or more signals from at least one position sensor. The one or more signals indicate a position of one of a landing gear bay door and a landing gear when the landing gear bay door or the landing gear, respectively, is part way through a range of travel between two limits of travel. The method also comprises the landing gear controller controlling movement of the other of the landing gear bay door and the landing gear on the basis of the one or more signals.
Disclosed is an autonomous inspection system comprising at least one autonomous mobile robot. The robot has a plurality of two-dimensional LiDAR scanners each scanner having a two-dimensional scanning plane. The plurality of scanners are mounted on the autonomous mobile robot with scanning plane orientations which are non-coplanar. A processor comprises an input to receive point data from the plurality of LiDAR arrays and an output to provide inspection data. The processor is configured to compile the point data from the plurality of LiDAR scanners into three-dimensional plot of the surrounding of the autonomous mobile robot and identify the dimensions and profile of articles within the three-dimensional plot. A method of inspection is also disclosed.
A method of copying data from a first device to a second device, each device comprising a wireless communication interfaces having a maximum range, is disclosed. The first device and the second device are separated by a distance greater than the maximum range. The method comprises: transmitting, by the first device, the data to a third device positioned within the maximum range of the first device; receiving the data at the third device and storing the data in a memory of the third device; moving the third device to a physical location within the maximum range of the second device; and transmitting, by the third device, the data to the second device.
A method of configuring a tyre monitoring device is disclosed. The method comprises, at the tyre monitoring device: entering a configuration mode; responsive to entering the configuration mode, transmitting a first message to a second device, the message indicating that the second device should transmit configuration data to the tyre monitoring device; receiving configuration data from the second device in response to the first message; and configuring the tyre monitoring device based on the configuration data.
A method for initialising a tyre monitoring device comprising a wireless communication interface and tyre monitoring devices and systems using the method are described. The method comprises: operating the tyre monitoring device in a first mode, wherein in the first mode the tyre monitoring device is only responsive to an initialisation instruction received over the wireless communication interface from the second device; receiving an initialisation instruction over the wireless communication interface while operating in the first mode; and, responsive to the initialisation instruction, operating the tyre monitoring device in a second mode in which the tyre monitoring device is responsive to instructions received over the wireless communication interface. The first mode draws lower power than the second mode so power draw prior to initialisation, for example while storage awaiting use, is reduced.
A method of determining a status of a tyre monitoring device is discussed. The tyre monitoring device comprises a counter initiated on a first use of the tyre monitoring device on a wheel and the method comprises: determining a current value of the counter; determining a status of the tyre monitoring device based on the current value of the counter; and providing an indication based on the determined status. In some examples, the status is a remaining service lifetime and the indication may be a replacement notification. A method of cross-checking data between tyre monitoring devices is also discussed. The method comprises receiving data from a plurality of tyre monitoring devices associated with respective tyres of a same vehicle; comparing the received data from different ones of the plurality of tyre monitoring devices; determining inconsistent data associated with at least one of the plurality of tyre monitoring devices based on the comparison; and providing an indication of the inconsistent data.
Methods of restricting commands for a tyre monitoring device are discussed, including a method for controlling entry into a configuration mode and a method of restricting operation based on detection of an abnormal command pattern. In one example, a method comprises, at a tyre monitoring device: receiving a configuration request from a first device, the configuration request indicating that the tyre monitoring device is to operate in a configuration mode, the configuration request provided using a first wireless communication protocol; responsive to receipt of the configuration request, initiating communication using a second wireless communication protocol; establishing successful communication using the second wireless protocol; and entering the configuration mode responsive to establishing successful communication using the second wireless communication protocol. In another example, a method of restricting the operation of a tyre monitoring device comprising a wireless communication interface comprises, at the tyre monitoring device: receiving instructions via the wireless communication interface; detecting an abnormal instruction pattern based on the received instructions; and restricting the operation of at least one component of the tyre monitoring device responsive to detecting the abnormal instruction pattern.
A method of checking tyre pressures using a control device having a wireless communication interface and at least one tyre monitoring device is disclosed. Each tyre monitoring device is mounted on a respective wheel of a vehicle and comprises a wireless communication interface and a pressure sensor for sensing an inflation pressure of a tyre on the respective wheel. The method comprises, at the control device: receiving an input and, responsive to the input: identifying a plurality of tyre monitoring devices within a wireless communication range of the control device, wherein the identifying comprises receiving responses from each tyre monitoring device within the wireless communication range, each response including a vehicle identifier associated with a vehicle upon which the tyre monitoring device is installed; selecting a vehicle identifier to be the subject of the request to check a tyre pressure based on the received responses; and sending a request to check tyre pressure to at least one tyre monitoring device associated with the selected vehicle identifier.
The application relates to a pipe support block for an aircraft fluid system. The pipe support block is arranged to support a pipe on a structural member, such as a bracket. The pipe support has a first part and a second part which together form a collar. The collar has a pipe receiving bore and is configured to support a fluid pipe on the structural member. First and second fastener holes extend through both the first part and the second part. The first and second fastener holes are each arranged to receive a fastener to fasten the first and second parts to each other and to the structural member.
F16L 3/10 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets substantially surrounding the pipe, cable or protective tubing divided, i.e. with two members engaging the pipe, cable or protective tubing
A tyre monitoring device and method is described in which an error in a tyre monitoring device can be detected and indicated. A tyre monitoring device comprises an indicator and a wireless communication interface. A method of operating the tyre monitoring device comprises: receiving an instruction to measure a tyre pressure via the wireless communication interface; measuring a tyre pressure; determining an error status of the tyre pressure monitoring device; and responsive to a determination that the error status is not indicative of an error, activating the indicator to provide an indication based on the tyre pressure.
A method, apparatus and system are disclosed in which an indication given by at least one tyre monitoring device is confirmed by input received from another source is disclosed. The method comprises causing, by the control device, at least one tyre monitoring device mounted on a wheel to measure a tyre pressure and to indicate a status using an indicator provided on the tyre monitoring device; receiving, from the at least one tyre monitoring device, data representative of the status of the at least one tyre monitoring device; receiving input representative of the status indicated on the indicator of the at least one tyre monitoring device; comparing the received data representative of the status with the received input representative of the status; and taking action based on the result of the comparing.
The present invention provides a method of operating an aircraft, comprising the steps of, during a non-braking time period (21), taxiing the aircraft by using driving torque (25) provided by a landing gear drive system, and not providing a braking torque (26) from the landing gear brake system, and, during a braking time period (22, 23), providing the brake command device at one or more command levels within a sub-range of a brake command range, and controlling the landing gear drive system, in response to the level of the brake command device, to reduce the driving torque provided.
B60T 8/17 - Using electrical or electronic regulation means to control braking
B60T 8/32 - Arrangements for adjusting wheel-braking force to meet varying vehicular or ground-surface conditions, e.g. limiting or varying distribution of braking force responsive to a speed condition, e.g. acceleration or deceleration
74.
AN ARRANGEMENT FOR AVOIDING CLASHING BETWEEN AN ACTUATION ASSEMBLY AND AN UPPER COVER OF A FOLDING WING TIP
An aircraft wing (10) comprising a fixed wing (1) and a wing tip device (12) rotatable about a hinge (16) at the tip of the fixed wing. The wing further comprises an actuation assembly (11) for rotating the wing tip device about the hinge. The upper cover (26) of the wing tip device comprises an offset region (34) extending inboard of the hinge axis (18), such that when the wing tip device is rotated to the ground configuration, the offset region extends into a space previously-occupied by the actuation assembly, but is outside the space occupied by the actuation assembly in the ground configuration.
The present application relates to a rib mounting assembly for an aircraft. The rib mounting assembly includes a rib post (60) and a seal member (80, 81). The rib post has a rib post foot (63) to mount with a longitudinal spar (31) and a rib post web (62) upstanding from the rib post foot. The seal member has a seal body (82) and a mounting flange (87). The mounting flange is mounted with the rib post foot and the seal body extends from an end of the rib post.
The present application relates to a spark containment cap. The cap forms a sealed cavity around an end of a fastener protruding from a structure. The cap has a cap body defining an air cavity arranged to enclose the end of the fastener. The cap body includes an annular base (210) terminating at a base rim (211) which surrounds an opening into the air cavity. An annular skirt (230) of the cap provides an annular sealing volume extending around the base rim (211) arranged to receive an annular bead of a curable sealing material around the opening into the cavity to provide a seal between the cap body and the structure to seal a volume of gas within the cavity. The cap body comprises a cap upper arranged to receive at least part of the end of the fastener. The cap upper (220) is positionable so that an axis of the cap upper (220) is offset from an axis of the annular skirt (230).
An aircraft assembly comprises a longitudinal spar and an aerofoil cover integrally formed from a composite laminate material to form a spar-cover such that the composite material of the spar extends continuously into the cover through a fold region created between the spar and the cover. The spar and cover are separated by a recess at a longitudinal end of the fold region to define a spar end region and a cover end region, and a reinforcement element extends between the spar end region and the cover end region to couple the spar end region with the cover end region.
A retractable fuselage mounted landing gear assembly is disclosed. The landing gear assembly comprises a main strut, a drag stay and a landing gear main fitting. A first end of the drag stay is attached to the main strut and a second end of the drag stay is connected to the fuselage. A first end of the landing main fitting is attached to the main strut and a second end of the main fitting is connected to the fuselage. When the landing gear is extended, substantially all the landing gear loads are transferred from the landing gear to the fuselage via one or more of the drag stay and the main fitting.
Disclosed is an aircraft assembly comprising: a first aircraft structure, a second aircraft structure, first and second tension bolts arranged to suspend the first aircraft structure below the second aircraft structure, the first and second tension bolts connected to the first aircraft structure such that an end of each of the tension bolts is accessible from respective outer surfaces of the first aircraft structure and first and second fail-safe mechanisms each associated with a respective one of the first and second tension bolts and arranged to suspend the first aircraft structure below the second aircraft structure responsive to failure of the respective tension bolt. Also disclosed is a method of mounting an engine mounting pylon to a wing box of an aircraft wing and an aircraft.
An aircraft wing comprising a fixed wing and a wing tip device rotatable about a hinge at the tip of the fixed wing. The wing further comprises a hinged panel at the boundary between the wing tip device and the fixed wing. In the flight configuration the panel is closed such that the panel is substantially flush with the upper surface of the wing, and the hinged panel is located in the path of a moveable element (for example part of an actuation mechanism, such as a curved rack). In the ground configuration the panel is hinged open to a position outside the path of the moveable element such that clashing of the moveable element with the stationary structure at the tip of the fixed wing is avoided.
An aircraft panel assembly with: a panel; and a plurality of stiffeners on the panel. Each stiffener has an attachment part attached to the panel and a structural part spaced apart from the panel. A rib foot beam crosses the stiffeners at a series of intersections. At each intersection the rib foot beam is located between the panel and the structural part of a respective one of the stiffeners.
B29C 33/50 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles with means for collapsing or disassembling elastic
B29C 43/10 - Isostatic pressing, i.e. using non-rigid pressure-exerting members against rigid parts or dies
B29C 43/36 - Moulds for making articles of definite length, i.e. discrete articles
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B29C 70/84 - Moulding material on preformed parts to be joined
B29D 99/00 - Subject matter not provided for in other groups of this subclass
Disclosed is a landing gear system uplock arrangement (300, 500) comprising: a hook (310) defining a gap (314), wherein the hook is movable between a first position, at which the hook is for retaining a movable element (299, 499) of a landing gear system in the gap, and a second position, at which the hook is for permitting movement of the element into and out from the gap; a lock (320) that is changeable between a locked state, in which the lock prevents movement of the hook from the first position to the second position, and an unlocked state, in which the lock permits movement of the hook from the first position to the second position; and a sensor arrangement (330) for sensing whether the element is present in the gap.
The invention relates to aircraft with wing tip devices, for example, folding wing tips. The folding wing tip may have a flight configuration for use during flight, and a ground configuration for use in ground-based operations. The ground configuration creates a shorter wing span than when the aircraft is in the flight configuration. A hinge arrangement protrudes beyond an outer surface of the fixed wing and wing tip device. In order to reduce the aerodynamic effect of the protruding hinge a fairing may be provided. In the flight configuration the fairing comprises a rear portion with a concave surface, such that the fairing acts as a lift-generating surface when the aircraft is in flight.
The invention relates to aircraft with wing tip devices, for example, folding wing tips. The folding wing tip may have a flight configuration for use during flight, and a ground configuration for use in ground-based operations. The ground configuration creates a shorter wing span than when the aircraft is in the flight configuration. A hinge arrangement protrudes beyond an outer surface of the fixed wing and wing tip device. In order to reduce the aerodynamic effect of the protruding hinge a fairing may be provided. In the flight configuration the fairing comprises a front portion blended into and extending rearwardly from the leading edge of the fixed wing and the wing tip device.
An aircraft wing (102) includes a closed surface wing tip device (104) which includes an element or actuator (330) within the wing tip for deforming / morphing the shape of the wing tip between geometrical configurations having different aerodynamic properties, for example including one with better overall fuel efficiency for a shorter journey and one with overall fuel efficiency better suited for a longer journey. The device (400) includes a lower winglet (406) with an essentially planar portion (410) spaced apart from the main body of the wing by a blended transition region which is shaped such that the curvature of the local dihedral increases in the outboard direction. The device (400) includes an upper aerofoil structure (414) which with the winglet (406) essentially forms the closed surface. There is also disclosed an aircraft wing tip device(101)having a sigmoid shaped (e.g. S-shaped) aerofoil structure (110) blending in with a main wing of the aircraft.
The present invention relates to a method of manufacturing an aerodynamic structure. More particularly, but not exclusively, the invention relates to a method of manufacturing an aircraft wing tip. The method comprises the following step providing an upper cover (28) comprising a plurality of lugs (16, 18, 20), providing a lower cover (32) comprising a plurality of lugs (22, 24, 26), measuring the distance between a lug on the upper cover (28) and corresponding lug on the lower cover (32), selecting, from the plurality of different sized links, a link suitably sized to correspond to the measured distance between the lug on the upper cover and corresponding lug on the lower cover, connecting the selected link to the lug on the upper cover and corresponding lug on the lower cover; and repeating those steps such that each lug on the upper cover is joined to a corresponding lug on the lower cover.
There is provided a fitting (10) for connecting a first aircraft (20) structure to a second aircraft structure (27). The fitting (10) comprises a first substantially planar surface (11), a second substantially planar surface (12), a body (13), and a cavity (14) in the body (13) configured to retain a captive nut. The body (13) connects the first surface (11) to the second surface (12) such that the first surface (11) is at an angle of less than or equal to 90° to the second surface (12). The cavity (14) is configured such that a captive nut retained therein is adjacent the first surface (11) and the second surface (12).
A pipe joint with a socket (31a), and a pipe (30a) fitted into the socket. An annular seal (50a) is compressed between the socket and the pipe. The seal is electrically conductive and resiliently flexible. The socket or the pipe has a recess which houses the seal. The recess has a ridge (51a) in a base of the recess and the seal has a groove (42a) which extends around a radial periphery of the seal. The ridge fits into the groove to fool proof the installation, such that a standard non-conductive 0-ring seal is more difficult to install in error.
F16L 21/02 - Joints with sleeve or socket with elastic sealing rings between pipe and sleeve or between pipe and socket, e.g. with rolling or other prefabricated profiled rings
F16L 21/03 - Joints with sleeve or socket with elastic sealing rings between pipe and sleeve or between pipe and socket, e.g. with rolling or other prefabricated profiled rings placed in the socket before connection
F16L 21/035 - Joints with sleeve or socket with elastic sealing rings between pipe and sleeve or between pipe and socket, e.g. with rolling or other prefabricated profiled rings placed around the spigot end before connection
F16L 25/01 - Construction or details of pipe joints not provided for in, or of interest apart from, groups specially adapted for realising electrical conduction between the two pipe ends of the joint or between parts thereof
B64D 37/32 - Safety measures not otherwise provided for, e.g. preventing explosive conditions
A wing tip device (120) for a fixed wing aircraft (4). The wing (2, 3) or wing tip device (120) has an alular-like projection. In particular the wing (2, 3) or wing tip device (120) has a first leading edge (26) region (131) having a first sweep angle, a second leading edge region (132) outboard of the first leading edge region (131) in a spanwise direction and having a second sweep angle greater than the first sweep angle, a third leading edge region (133) outboard of the second leading edge region (132) in the spanwise direction and adjacent a tip end (124) of the wing tip device (120) and having a third sweep angle greater than the first sweep angle. The second leading edge region (132) is adapted to generate a first vortex (173), and the third leading edge region (133) is adapted to generate a second vortex (174) which builds towards the tip end (124) of the wing tip device (120).
Disclosed is an aircraft engine nacelle (100) for coupling to a wing (400) of an aircraft (1,2,3), the nacelle (100) comprising a fore end (10); and an aft end (30) that is immoveable relative to the fore end (10); wherein the aft end (30) comprises a major axis (Mj) and a mino axis (Mi); and wherein the nacelle (100) is configured such the minor axis (Mi) is closer to vertical (V) than the major axis (Mj), when the nacelle (100) is coupled to the wing (400) and the aircraft (1,2,3) is stationary on the ground. Also disclosed is an aircraft system and an aircraft (1,2,3), each comprising the aircraft engine nacelle (100).
The invention provides an aircraft landing gear (100) comprising a first oleo strut (10) comprising a sleeve portion (11) and a slider portion (13), the slider portion being slidable within a hydraulic fluid chamber (12) of the sleeve portion, and a second, similar oleo strut (30). The landing gear also comprises a hydraulic fluid balancer (50) comprising a balance chamber separated into first (55) and second (57) end sections, wherein the hydraulic fluid chamber of the sleeve portion of the first oleo strut is fluidly connected to the first section of the balance chamber and the hydraulic fluid chamber of the sleeve portion of the second oleo strut is fluidly connected to the second section of the balance chamber of the hydraulic fluid balancer. The invention also provides a "live axle" landing gear with a drive mechanism or a braking device (22) being mounted on the axle, and various associated methods.
Conventionally, spars have been manufactured so as to have box sections to provide good torsional stiffness. However, this design makes them difficult to manufacture from composite materials using known techniques, such as Resin Transfer Moulding (RTM), because of the difficulties of removing the mandrel from the interior of the spar during demoulding. Certain desired configurations of spar are not possible to make by means of RTM because of the difficulty of removing the mandrel. The invention provides a composite spar (1) comprising a main body (2) and a pair of spaced apart leg (3,4). The space between the legs permits access to the mandrel assembly within the spar in order to ease its removal. This arrangement permits a spar of any length and contour to be manufactured.
B64C 23/06 - Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
Disclosed is a landing gear system controller for controlling a landing gear system of an aircraft. The controller is configured to receive a request to operate a non-landing gear system element of the aircraft, and to cause operation of at least part of the landing gear system of the aircraft during a take-off or landing procedure on the basis of the request.
A system is provided that includes a blockchain computer system, an issuer computer system, and a verifier computer system. Different nodes of the blockchain computer system are associated with the issuer and verifier computer systems. The issuer computer system and the node of the blockchain computer system associated therewith generate and submit blockchain transactions that contain data for aircraft personnel certifications. The aircraft personnel certifications are stored on the blockchain and may be later retrieved by the verifier computer system and its associated node of the blockchain computer system. Accordingly, the certifications of aircraft personnel can be maintained and verified. The system ensures the digital certifications are authentic and have not been tampered with.
An aerodynamic body for use on an aircraft is presented. The aerodynamic body includes at least a first perforated surface portion (25) and an ice-protection system (31). The first perforated surface portion (25) has perforations, the ice-protection system (31) includes an actuatable element (33) and the actuatable element (33) is movable or deformable between a first position and a second position. In the first position at least a section of the actuatable element (33) is thermally coupled to the first perforated surface portion (25) and configured to prevent an inflow or outflow between a boundary layer of an outer aerodynamic airflow and the aerodynamic body through at least one of the perforations. In the second position at least a section of the actuatable element (33) is distanced from the first perforated surface portion (25) and configured to allow an inflow from a boundary layer of an outer aerodynamic airflow through at least one of the perforations into the aerodynamic body.
An uplock (101) for use with an aircraft component. The uplock comprises a hook (102) configured to engage a capture pin (126) mounted on the aircraft component. The hook (102) is mounted for movement between a closed position and an open position. The uplock (101) further comprises an indicator system (130) configured to detect whether a pin (126) is engaged with the hook (102) when the hook (102) is in the closed position.
A control system (20) has both a first and second processing modules (22, 24) for controlling retraction and extension of a landing gear assembly (14) on an aircraft (10). The processing modules can operate independently of each other and thus provide redundancy. Each processing module is configured to perform a first sequence of steps for retracting the landing gear assembly and a second sequence of steps for extending the landing gear assembly. A step of switching control from one processing module to the other (e.g. an avionics side changeover step) is performed (a) as part of the first sequence of steps, but only after the landing gear assembly has been retracted (e.g. step 320 of Fig 8), or (b) as part of the second sequence of steps (e.g. step 420 of Fig 9 or step 520 of Fig 10). By ensuring that the step of switching control is performed at such times, and not for example between initiation of the first sequence of steps and the retraction of the landing gear assembly, the landing gear assembly (14) may be retracted sooner after such initiation.
An aircraft wing assembly (5) is presented comprising a main wing portion (21), a high-lift device (9) comprising a flow surface with an upper skin portion (25) and a lower skin portion (27), wherein the high-lift device is movable between a retracted position at the main wing portion and a deployed position defining a gap (37) between the high-lift device and the main wing portion, wherein the main wing portion and the high-lift device in the retracted position define an airfoil having a local chord (c(y)) between a trailing edge (15) and a leading edge (11), wherein the leading edge is on the flow surface of the high-lift device between the upper skin portion and the lower skin portion, wherein the flow surface of the high-lift device comprises a first flow surface portion (23), a second flow surface portion (25) and a third flow surface portion (27), wherein the first flow surface portion is micro-perforated for an air inflow, wherein the first flow surface portion extends on the upper skin from the leading edge in chordwise direction for 2% or less of the local chord and extends on the lower skin from the leading edge in chordwise direction for 2% or less of the local chord, and wherein the second flow surface portion is not micro-perforated and extends over the rest of the upper skin portion, and wherein the third flow surface portion is not micro-perforated and extends over the rest of the lower skin portion.
Disclosed is an apparatus for dissipating braking energy generated by the operation of a vehicle brake. The apparatus comprises a heating element configured to receive electric current generated by operation of the vehicle brake, the heating element arranged in a chamber for containing liquid. The apparatus is configured such that liquid contained in the chamber is heated in response to operation of the vehicle brake.
The present invention is concerned with actuating aircraft components to tailor their configuration to achieve improved performance in one or more flight phases. In general terms, the invention provides a method and system for actuating an aircraft component comprising an actuating material in which an actual deformation change undergone by the actuating material in response to an activation input signal is determined by analysis of an output signal generated by the actuating material in response to the actual deformation. A first aspect of the invention provides a method of actuating an aircraft component, at least a portion of the aircraft component comprising an actuating material which undergoes deformation in response to the application of an electrical signal thereto, and which generates an electrical signal in response to a deformation of the actuating material, the method comprising the steps of: applying an activation input signal to the actuating material of the aircraft component, the activation input signal corresponding to a desired deformation of the actuating material, the actuating material of the aircraft component undergoing an actual deformation in response to the activation input signal; and generating an output signal representative of the actual deformation of the actuating material.
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