Abstract: An advanced braking system for an aircraft is disclosed. A controller calculates the braking required from each wheel in terms of force. A constant deceleration is achieved throughout a braking run by calculating the braking from other sources, principally aerodynamic drag, and commanding a complementary total level of braking from the wheel brakes. The performance of each wheel and brake are monitored during the braking run to determine whether their braking performance is limited by the brake discs or by the tire-ground interaction and to see whether the wheel is approaching the maximum slip ratio after which a skid occurs. The controller uses this information to distribute the total demand for braking amongst the wheels. In doing this, it also aims to keep the braking demand symmetrical across the aircraft and not to overheat the brakes. The controller further measures the braking force provided by a wheel and controls its brake pressure accordingly to achieve the force desired.

Abstract: A sealing member for forming a seal in an aircraft, the sealing member comprising: a sealing material; and a stiffening element which provides structural support to the sealing material and comprises a material with a glass transition temperature below +50° C. The stiffening element is relatively flexible when the aircraft is at low altitude (high temperature) but becomes relatively stiff (increasing resistance to seal flutter) when the aircraft is at cruise altitude (low temperature).

Abstract: An aircraft landing gear (9) includes a wheel (1) having a wheel rim (3) on which a tire (4) is held. The gap (6) between the rim (3) and tire (4) is bridged and covered by a sealing element (7), which thereby presents a smooth surface to the air flowing over the wheel during flight of the aircraft (8). Thus, noise that would otherwise be generated by the interaction of air and the parts of the wheel (1) and/or tire (4) defining the gap (6) is reduced. Such noise reduction benefits may also be achieved by providing a tire (4) and wheel (1) so shaped that there is no gap (6) between the tire (4) and wheel rim (3).

Abstract: An aircraft landing gear is arranged such that in its deployed position it is configured to reduce, during the approach on landing, the noise generated by the interaction of the landing gear and the air flowing past the landing gear. Various means are described to reduce the amount of turbulent air flow generated in the region of the landing gear, including inverting and fairing the nose-gear shock-absorbing leg, providing faired twin in-line oleos and providing fairings that produce shielding air curtains.

Abstract: An aircraft landing gear comprising: an arm mounted to a pivot and carrying one or more wheel assemblies; and a composite leaf spring which is coupled to the arm and arranged to provide a resilient biasing force which opposes rotation of the arm about the pivot. The arm can be made relatively stiff in comparison with the leaf spring, so that load is transferred directly into the pivot without a large amount of bending of the leg. The leaf spring shock absorber typically has a reduced part count compared with a conventional oleo-pneumatic shock absorber. By forming the leaf spring from a composite material, its weight is significantly reduced in comparison with a conventional metal spring, making it suitable for use on a relatively large aircraft.

Abstract: This invention relates to aircraft light units, in particular aircraft light units used when landing and taxiing an aircraft. The invention provides an aircraft light unit comprising a first light source with a principal beam axis in a first direction and a second light source with a principal beam axis in a second direction, the second direction being different from the first direction. The first and second light sources are arranged such that they can combine to create a net light beam in a plurality of different directions by variation of the power supplied to the light sources.

Abstract: An aircraft landing gear is described including a noise-reducing means. In one embodiment there is a noise-reducing element that includes an air-deflecting surface, in the form of a fairing, and a perforated airflow-reducing region. In use, some air passes through the perforations, while other air is deflected by the fairing away from a noise-generating region of the landing gear. Noise caused by the passing of the landing gear through the air is therefore reduced. In another embodiment a noise-reducing layer in the form of a mesh is supported by a skeleton structure.

Abstract: According to the invention, the engine pylon for the suspension of a turbo engine under an aircraft wing is such that the trailing edge of the rear lower part of the engine pylon is shifted toward the root of the wing.

Abstract: The invention relates to a fastener assembly 1 for use as a slave bolt. The fastener assembly can be tightened and released from the same side and comprises: a) a shank 2 having a bolt head 3 at one end; b) a sleeve 7 slidably arranged on the shank 2; c) a spring 6 to bias the sleeve on the shank 2; and d) an expander stem 9 which can be screwed into a threaded bore 4 which runs the length of shank 2, the stem 9 having an enlarged portion which spreads the fingers 8 of the sleeve so that they engage and clamp the articles to be fastened.

Abstract: An aircraft comprises a wing (2) defining an aerofoil surface, the wing (2) comprising a drooped leading edge flap (1) being moveable between a stowed position and a deployed position. The wing (2) is so arranged that during flight when the high-lift device is in the deployed position, air may flow through an opening (7) in the wing (2) and over the aerofoil surface. During flight, air preferably flows into the boundary layer (13) on the upper surface of the wing (2). This energises the boundary layer (13), aft of the trailing edge of the drooped leading edge flap increasing its stability allowing the maximum achievable lift coefficient to be increased and hence reducing aircraft take-off and approach speeds.

Abstract: An aircraft structure comprising: a rear spar; and a composite cover attached to the rear spar and having an overhanging portion extending to its rear. The overhanging portion of the cover comprises one or more ramps along which the thickness of the cover reduces. A second cover is attached to the rear spar and overhangs to its rear. A hinge rib is attached to one or both of the covers, and an aerodynamic control element such as an aileron, rudder or elevator is pivotally mounted to the hinge rib.

Abstract: An aircraft landing gear door assembly includes a door, which is moveable to an open position, in which an aperture is exposed, the landing gear being able to be deployed through said aperture, and moveable to a closed position, in which the door is closed across the aperture. The door is arranged for rotational movement of a first type, in which the door rotates about an axis that is substantially parallel to the length of the aircraft, between the closed position and the open position. The door is also arranged for movement of a second type, different from the first type, in which the door rotates, about an axis that is substantially transverse to the length of the aircraft, to a position in which it acts as a fairing to reduce the noise caused by the landing gear or a part thereof. In one embodiment, the second type of movement is rotational movement about an axis that is substantially parallel to the door when in the closed position and the door forms a ramp type fairing.

Abstract: A method of operating an aircraft fuel management system for an aircraft having at least one fuel tank, each fuel tank having an associated fuel quantity indicator arranged to provide an indication of the amount of fuel in the associated fuel tank, the method comprising calculating a value for the amount of fuel on board (FOB_FailedFQI) the aircraft to be utilised by the fuel management system in the event of a failure of at least one of the fuel quantity indicators, the amount of fuel on board being calculated as a value for the initial amount of fuel on board (FOBinit) minus the amount of fuel used. Additionally or alternatively a value for the gross weight centre of gravity of the aircraft is calculated using an assigned value of fuel for each of the fuel tanks having an associated failed fuel quantity indicator.

Abstract: A method of controlling the centre of gravity of an aircraft having a plurality of fuel tanks, the method comprising transferring fuel from one or more of the fuel tanks according to a predetermined sequence, the timing of the sequence being dependent on the decrease in gross weight of the aircraft.

Abstract: A fuel monitoring system for automatically monitoring a fuel transfer in an aircraft fuel system, the fuel system including a plurality of fuel tanks, the fuel monitoring system comprises a fuel quantity sensor arranged to measure the quantity of fuel in a first fuel tank and a data processor arranged to receive a fuel quantity measurement from the sensor, wherein in response to receiving a command to transfer fuel from the first fuel tank to one or more further fuel tanks the data processor is arranged to determine the rate of change of fuel quantity in the first tank from the received fuel quantity measurement and if the rate of change of fuel quantity is less than a threshold value and the received fuel quantity measurement is greater than an expected value then the data processor is further arranged to provide an output indicating that the commanded fuel transfer has failed.

Abstract: A structure comprising a composite laminate having an edge; and an impact indicator which is carried by the edge and comprises a resin which fractures upon impact. The fracture provides permanent visible evidence of impact damage, for instance by cracking or by one or more pieces breaking off from the impact indicator. As well as providing such visible evidence of impact damage, the impact indicator may also provide an element of impact protection. Typically the impact indicator is formed from a resin which is more brittle and less strong than the material forming the composite laminate. For instance the material forming the composite laminate may be reinforced, and the resin forming the impact indicator may be un-reinforced.

Abstract: There is provided a method of manufacturing a rib for an aircraft wing, the method comprising providing a billet of material for forming the rib, machining away material from all sides of the billet to form a rib for use in an aircraft, the rib including a web that extends between first and second flanged portions, each flanged portion having a flange that facilitates fixing of the rib to another component of the aircraft, wherein the material that forms the web extends diagonally across the billet, the web having a first face on a first side and a second face on a second side opposite to the first side of the web, and the first flanged portion extends more to the first side of the web than to the second side and the second flanged portion extends more to the second side of the web than to the first side.

Abstract: This invention relates to aircraft light units, in particular aircraft light units used when landing and taxiing an aircraft. The invention provides an aircraft light unit comprising a first light source with a principal beam axis in a first direction and a second light source with a principal beam axis in a second direction, the second direction being different from the first direction. The first and second light sources are arranged such that they can combine to create a net light beam in a plurality of different directions by variation of the power supplied to the light sources.

Abstract: A method of removing a component from an aircraft includes attaching a device comprising a guide member to the aircraft; attaching the component (such as an actuator) to the guide member; detaching the component from the rest of the aircraft; and moving the guide member so that the component follows a predefined path of removal of the component. A device for use in such a method is also described.

Abstract: Unsafe electric energies can exist in aircraft fuel tanks as a result of static build-up or induced currents from lightning strikes. The invention provides an aircraft fuel tank including two discharge paths for the safe discharge of electricity from the tank before it reaches levels where an arc, spark or other electric discharge, able to ignite fuel in the tank, could occur. The failure of both the discharge paths would significantly increase the risk of an ignition causing electrical discharge occurring. The tank thus also includes a circuit for monitoring the integrity of each discharge path. By having at least two discharge paths, should one discharge path fail (which failure being detected by the monitoring circuit, at least one alternative discharge path exists for the safe discharge of electrical energy from the fuel tank, until the failed discharge path is repaired.