Abstract: A reading device, such as a reading lamp or reading spotlight, is provided for use in a vehicle, such as a motor vehicle with immovable properties or ships. The reading device for example can be used in passenger cabins and/or cockpits of aircraft. The reading device includes, but is not limited to an LED light source with a color temperature that can be adjusted with the color “warm-white” to “cold-white”, a primary optics or lens advantageously designed as a converging lens, and a secondary optics or lens to focus the light emitted by the light source onto a surface, such as a reading plane to be illuminated.

Abstract: The present invention provides a fault diagnosis device for optimizing maintenance measures in a system, in particular a flying object. The diagnosis device has a storage unit for storing diagnosis rules, which each have a fault condition indicating the presence of a fault, and a fault conclusion formed by logical combinations of potential fault causes of the respective fault. Furthermore, a logic unit is provided, which logically combines those stored fault conclusions whose associated fault conditions are met and, on the basis of the combined fault conclusions, determines a logical diagnosis result having at least one fault cause hypothesis. A calculation unit subsequently calculates an occurrence probability for each fault cause hypothesis of the logical diagnosis result determined.

Abstract: A cooling system (10) for cooling an aircraft device (12) comprises an evaporator (14, 16, 18) for receiving a fluid (F) which is to be evaporated, a first adsorber (24) which contains a medium (28) for the adsorption of the fluid (F) which is evaporated in the evaporator (14, 16, 18), as well as a second adsorber (26) which contains a medium for the adsorption of the fluid (F) which is evaporated in the evaporator (14, 16, 18). A control system (22) is adapted to establish or to interrupt a fluid connection between the evaporator (14, 16, 18) and the first and/or the second adsorber(s) (24, 26).

Abstract: A method for tolerance compensation in an overlap joint between a first and a second body section of adjacent fuselage rings includes introducing a plurality of holes into a first body section corresponding to a pattern of rivets; arranging the first and second body section in an overlapping configuration so as to form an overlap; injecting a filler mass through the plurality of holes into a gap formed in the overlap; and hardening the filler mass.

Abstract: A flight management verification apparatus adapted to verify at least one flight management, wherein the flight management in use outputs a warning, disables an aircraft and/or inhibits the activation thereof in response to a power fail condition. The flight management verification apparatus comprises a first connection means adapted to be disposed in a first power indicator line to a first flight management, wherein the first power indicator line provides a signal representative of a condition of a first power supply, and further comprises a first switch connected in use such that the signal of the power indicator line for the first flight management means is switched. Actuating the switch causes the signal on the power indicator line to change indicating a distortion or failure of the power supply of the aircraft and/or a part thereof. The response of the flight management can be verified by the signal output by the flight management, e.g.

Abstract: The present invention provides a device and method for determining aerodynamic characteristics of an aircraft, in particular a pitch moment coefficient at an ascending force of the aircraft without a horizontal tail plane of zero and an aerodynamic neutral point of the aircraft without a horizontal tail plane. In the method according to the invention, forces and mechanical flight parameters on aerofoils and a horizontal tail plane of the aircraft are detected at various detection instants during at least one non-stationary flight maneuvre of the aircraft. In this case, a linear system of equations is formed, which comprises a pitch movement moment balance equation for each detection instant, the pitch movement moment balance equation having the detected forces and mechanical flight parameters. This linear system of equations is evaluated to calculate the aerodynamic characteristics.

Abstract: A positioning system is provided for mounting a wing of an aircraft to a fuselage of the aircraft. The positioning system includes, but is not limited to a programmable mounting unit that can perform a final adjustment and alignment of the wing with respect to the fuselage without user interaction.

Abstract: An aircraft leading edge is extended by an aerodynamic surface providing an aerodynamic flow and where air discharges are arranged preventing separation of the aerodynamic flow. The air discharges are arranged in at least two rows essentially parallel to the leading edge and in an offset manner for at least two consecutive rows. At least one block is inserted between two walls that form the aerodynamic surface. The block includes, on the one hand, an outside surface extending the aerodynamic surface, a first inclined surface contacting the first wall forming the aerodynamic surface, and a second inclined surface contacting the second wall that forms the aerodynamic surface, and, on the other hand, projecting and/or hollow shapes that are made at the inclined surfaces and that are arranged in an alternating fashion from one surface to the next, allowing air to pass on both sides of the aerodynamic surface.

Abstract: A footrest is provided that includes a footplate element, an upper swing arm, and a lower swing arm, which footrest can be arranged on a wall that faces a seat, and allows several seating positions.

Abstract: An aircraft fuel tank ventilation system, comprising a desiccative dehumidifying device including a desiccant medium disposed in flow communication between a vent open to the atmosphere and a fuel tank, and a microwave energy transmitter for energizing liquid water in the desiccant medium to facilitate regeneration of the medium. Also, a method of regenerating a desiccant medium of a dehumidifying device of an aircraft fuel tank ventilation system, the method comprising directing air through the desiccant medium, and transmitting microwave energy into the desiccant medium for energizing liquid water in the desiccant medium to facilitate regeneration of the medium.

Abstract: An aircraft tank system includes a tank cell for receiving fuel, a fuel line that connects the tank cell to a fuel inlet connection, and an aeration/deaeration device that for aeration/deaeration of the tank cell connects the tank cell to the environment. A shut-off valve of the aeration/deaeration device is adapted in an open position to open a connection between the tank cell and the environment and in a closed position to interrupt the connection between the tank cell and the environment. The aircraft tank system further includes an exhaust air outlet connection, which is connected to the aeration/deaeration device and which is adapted during filling of the tank cell with fuel to be connected to an aeration/deaeration connection of a storage tank of a refueling unit.

Abstract: The present invention provides test equipment for testing an additional center tank (ACT) system of an aircraft. The ACT system comprising control means configured to perform functions of an Auxiliary Fuel Management Computer (AFMC) and an Auxiliary Level Sensing Control Unit (ALSCU). The test equipment comprises at least one test module adapted for coupling to the ACT system in place of at least a portion of the control means, the test module comprising means for generating one or more test signals and outputting the test signals to the ACT system.

Abstract: An aerodynamically shaped profile member for aircraft and wind power stations includes, for example, phased-array ultrasonic generator arrangement arranged therein. During operation of the profile member, the ultrasonic generator emits ultrasonic waves in a targeted manner in multiple directions to determine a profile of the thickness of an ice layer on the surface of the aerodynamic profile. The structure may be, for example, a composite fiber material arranged around a foam core, with the ultrasonic generator arrangement being laminated into the composite fiber material. For measurement of the ice thickness, ultrasonic waves are transmitted in a targeted manner to different positions of the surface of the aerodynamic profile, and the ultrasonic waves reflected on the interfaces of the ice layer are detected. At least one region of the surface is scanned by means of the targeted ultrasonic waves, to determine an ice thickness profile.

Abstract: A method and system for improving aircraft performances during take-off is described. The system (1) can include means (4, 6, 7) for determining an optimised take-off position of the control surfaces (S1-Sn) of the aircraft, in the case where a regulatory safety criterion relating to the minimum gradient of climb with a breakdown engine is predominant.

Abstract: An anti-lightning system for a composite structure having an external surface liable to be subjected to an air flow, the aforementioned system including at least one electrically conducting strip and fasteners fastening the aforementioned strip to the structure, the aforementioned strip being electrically connected to an electrical earth for removing current. According the disclosed embodiments, the aforementioned strip has a shape intended to collaborate with a housing created at the aforementioned external surface so that the external surface of the strip lies flush with the aforementioned external surface of the structure, and the aforementioned fasteners include an upper part flush with the external surface of the strip.

Abstract: An aircraft front portion includes a sealed bulkhead separating a non pressurized radome area from a pressurized area. The sealed bulkhead is delimited by two surfaces opposite to each other, one being substantially concave when seen from the radome area and the other being substantially convex when seen from the pressurized area.

Abstract: An aircraft landing gear for an aircraft having an engine, an outer surface, a cavity formed in the outer surface, a cover arranged to close the cavity, the landing gear being arranged to be movable between a stowed configuration, in which it is stowed in the cavity, and a deployed configuration, the landing gear comprising an engine debris guard panel arranged to inhibit tire propelled debris striking the engine whilst the landing gear is in the deployed configuration, wherein the guard panel is arranged to form at least a part of the cover when the landing gear is in the stowed configuration.

Abstract: The invention relates to a pressurized aircraft door that minimizes risk of ingress of precipitation inside the aircraft, while maintaining venting of the aircraft before unlocking the door. The purpose of the invention is achieved using a vent flap controlled by a door control member and designed such that the flap successively occupies the following positions during the opening movement of the control member: a closed position when the control member occupies a first end position putting the door in a closed/locked position; at least one partially open position before the door is in a closed/unlocked position; then once again in the closed position when the control member occupies a second end position putting the door in an open position.

Abstract: The invention relates to an aircraft comprising a force transmission element which detachably connects a cabin structural segment to an aircraft primary structure and which comprises a cabin bearing element and a structure bearing element, the cabin bearing element being connected to the cabin structural segment and the structure bearing element being connected to the aircraft primary structure. The force transmission element is designed in such a manner that a force transmission can take place between the cabin structural segment and the aircraft primary structure with at least one degree of freedom of movement.

Abstract: Power takeoff means (46, 47, 60, 61) for this flight control system adjacent to the aircraft engines (41, 42) are all-electric and this makes it possible to reduce the length of hydraulic connections leading to actuators of the controlled devices (54, 55, 66). Electrical generators are of different types to minimize the risk of a generic failure. Characteristically of the invention, some generators (60, 61) have a constant voltage to frequency ratio and directly or almost directly control electro-hydraulic pumps (64, 65), without modifying the current characteristic and with no power electronics. Furthermore, system reconfigurations are made possible by additional electrical connections (69, 70, 71, 72) in case of failure.