125results about "Movable aircraft element position indicator" patented technology

A skew sensing apparatus for detecting skew in at least one, or a plurality of control surfaces of a vehicle. The apparatus includes a tension cord which is held under tension between a) a first spring associated with a sensor and a sensor target and b) an opposing second spring. During normal operation, the apparatus is configured to maintain said sensor target in a first position relative to said sensor, thereby indicating said normal operation. During an abnormal operation, the apparatus is configured to shunt said tension cord such that said sensor target is translated to a second position, relative to said sensor thereby indicating said abnormal status.

The present invention relates to a locking system (20) for locking a door (15) comprising at least two locks (31). The locking system (20) has a position sensor (40) per lock (31), each position sensor (40) emitting a measurement signal carrying a locked or unlocked state of the associated lock (31), the signaling system (30) comprising a controller (50) receiving the measurement signals and being connected to an alarm (60), the controller (50) being configured to emit either an unlock signal (SDV) when at least one measurement signal carries an unlocked state of the associated lock or a locking signal if each measurement signal carries a locked state of the associated lock, the alarm (60) generating a visual unlock alert upon receipt of the unlock signal and a visual lock alert upon receipt of the lock signal.Abbreviated figure: figure 1.

In order to provide an improved monitoring, a monitoring arrangement (10) for a plurality of adaptable segments of an airfoil is provided. The arrangement comprises a plurality of sensor cable segments (12), a plurality of cable connections (14); and a signal data processing unit (16). The sensor cable segments are each configured: to be attached between adjacent adaptable segments of the plurality of the adaptable segments; and to detect a spatial relative movement of the adjacent adaptable segments and to generate a signal representing the relative movement. The cable connections are configured to supply the signals of the sensor cable segments to the data processing unit. The data processing unit is configured to determine relative changes in a spatial arrangement of the plurality of adaptable segments based on the supplied signals.

A folding wingtip aircraft (100) includes a wing (104) that includes a primary portion (106) and a wingtip (102). The folding wingtip aircraft (100) includes a folding wingtip (102) system configured to rotate the wingtip (102) relative to the primary portion (106). The folding wingtip aircraft (100) also includes a flight control computer (202). The flight control computer (202) is configured to, in response to a first determination that a wingtip (102) position of the wingtip (102) is not in a folded position when a second determination indicates that a last completed command by the folding wingtip (102) system was a fold command, set error codes for the folding wingtip (102) system. The flight control computer (202) is also configured to change a setting to indicate that the folding wingtip aircraft (100) is grounded based on the error codes.

This device for detecting the position of a movable piece able to move between a first extreme position, a second extreme position and at least one intermediate position, comprises a fixed antenna and a plurality of radio-frequency identification tags arranged along the path of the movable piece between said extreme and intermediate positions, the device further comprising a mechanical contactor configured to establish an electrical connection between the antenna and a unique combination of at least one radio-frequency identification tag for each position of the movable piece.

To provide a system and method for dynamically measuring blade position during flight of a rotorcraft.SOLUTION: A blade 14 of a rotorcraft 10 is repeatedly illuminated by a light source during flight of the rotorcraft 10 while the blade is rotating. The method also includes a step of detecting radiation scattered from the blade in response to illumination of the blade 14. The method further includes a step of determining at least one of a blade pitch angle (ΔθP), a blade flap angle (ΔθF), a blade leading position or a blade lagging position on the basis of the radiation that is scattered from the blade 14 and detected. A rotorcraft 10 is also provided that includes a chip-scale light detection and ranging (LIDAR) sensor configured to illuminate a plurality of blades 14 while the blades 14 are rotating in order to permit blade 14 position to be measured or to illuminate terrain beneath the rotorcraft 10 in order to perform an altitude measurement.SELECTED DRAWING: Figure 1

An aircraft includes actuators coupled to a lift device. The aircraft also includes a lift device control interface disposed on a flight deck of the aircraft. The lift device control interface is configured to receive a pilot input. The aircraft further includes a flight control system configured to generate a command corresponding to a target lift device position. The flight control system is also configured to send the command to the lift device control interface to cause the lift device control interface to indicate the target lift device position. The flight control system is further configured to send a control signal to the actuators to set the lift device to the target lift device position.

A method for extending a landing gear system for a vehicle is disclosed. The landing gear system includes a first landing gear assembly and a second landing gear assembly. The method includes the steps of sensing a first load on the first landing gear assembly after the first landing gear assembly has reached a wheels-on-ground state and comparing the first load to a first target value. The method further includes the step of controlling an extension speed of the first landing gear assembly according to a difference between the first load and the first target value.

A bush is provided including a sleeve for receiving a pin; a shoulder extending from the sleeve, the shoulder for delimiting insertion of the sleeve into an aperture; and a sensor mount for mounting a sensor such that the sensor may detect the presence of a pin within the sleeve. Installing such a bush within an aperture of a component may negate the requirement for lugs and / or brackets on the component. This may enable reduction of the weight, size and / or complexity of the component.

An aircraft system for an aircraft is disclosed having a controller that is configured to receive at least one signal during a landing procedure of the aircraft and to determine that the aircraft 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 of the aircraft is issued during the landing procedure. The controller is configured to cause initiation of at least a part of a procedure to interrogate an aircraft braking system, on the basis of the determination.

A method of performing health monitoring for an aircraft having a variable pitch propeller system, the method comprising: receiving aircraft flight information; performing a virtual pitch calculation to determine a virtual propeller pitch or virtual blade angle from the aircraft flight information; receiving a measured propeller pitch or blade angle from at least one physical sensor associated with the propeller; and using the virtual propeller pitch or virtual blade angle and the measured propeller pitch or blade angle to perform a health monitoring process to determine a health condition of at least one component of the aircraft, wherein the at least one component is distinct from the blades.

Method of monitoring motion of an aircraft (100) during flight. The method includes determining sensed motion of the aircraft (100) about an axis during the flight and determining predictive motion of the aircraft (100) about the axis during the flight (based on a digital model (200)). An error is determined based on the sensed motion and the predictive motion. An effector (110) on the aircraft (100) is determined that controls the motion of the aircraft about the axis. The error is converted into an amount of necessary movement of the effector (110) to correct the error.

A computer-implemented method of determining a steering angle of an aircraft landing gear including: obtaining an input image of the aircraft landing gear; comparing the input image against a plurality of reference images, the plurality of reference images comprising images of the aircraft landing gear at known steering angles; determining a most similar reference image, the most similar reference image comprising the reference image of the plurality of reference images most closely matched to the input image; and determining, based at least in part on the most similar reference image, the steering angle of the aircraft landing gear.

The invention proposes a method for detecting a fault in a high-lift system (13) of an aircraft (1) and / or for monitoring the integrity of a high-lift system of an aircraft. The method comprises temporarily deploying at least one spoiler surface (23), performing a detection or measurement at least when the spoiler surface is temporarily deployed, and evaluating whether the high-lift system is faulty and / or evaluating the integrity of the high-lift system based on the detection or measurement.

This device (2) for detecting the position of a mobile part (1) able to move between a first extreme position, a second extreme position and at least one intermediate position, comprises a fixed antenna (3) and a plurality of radio-frequency identification tags (4, 4a, 4b, 4c) arranged along the path of the mobile part (1) between said extreme and intermediate positions, the device (2) further comprising a mechanical contactor (5) configured to establish an electrical connection between the antenna (3) and a unique combination of at least one radio-frequency identification tag (4, 4a, 4b, 4c) for each position of the mobile part (1).

A malfunction indicator assembly for a suite door slidable along a redundant rail system. The malfunction indicator assembly includes a weighted malfunction indicator pivotally mountable to a stationary wall and carrying a first magnet, and a second magnet mountable to one end of a secondary rail and carriage subassembly proximal to the weighted malfunction indicator. In use, in the absence of primary rail and carriage subassembly failure, a magnetic connection is maintained between the first and second magnets such that the indicator does not present outside of a vertical plane of the stationary wall, and in the presence of the primary rail and carriage failure the magnetic connection is broken to present the indicator outside of the vertical plane of the stationary wall. In embodiments, upper and lower malfunction indicators operate independently to report failures.

A system for detecting the state of a mechanism includes a processing circuit. The mechanism includes a plurality of components at least some of which are movable relative to each other. At least one state is specified for the mechanism based on the alignment of the mechanism components. The processing circuit identifies reference points in at least two of the components within the image data of the mechanism. Based on the respective positions of the reference points, it determines whether the mechanism components are in one of the specified alignments (i.e., states) and outputs an indicator.

Apparatuses and methods for shape-memory alloy elements and sensing. An aircraft assembly comprising a sensor, a cable connected to the sensor, and a shape-memory alloy (SMA) element connected to the cable. The SMA element is configured to change shape and control tension in the cable. A method of using an aircraft assembly by changing the shape of the SMA element by an amount corresponding to a change in a length of the cable.

A mechanical architecture is provided for aircraft distributed flap actuation and includes a shaft (310) and first and second actuation units (320, 330). The first actuation unit (320) includes a motor, a first actuator for translating and rotating a flap and being connected to a first side of the flap, a first gearbox to transfer and increase torque from the motor to the first actuator and a first brake configured to hold the flap in place. The second actuation unit (330) includes a second actuator that receives the torque from the first gearbox via the shaft (310) for translating and rotating the flap and being connected to a second side of the flap, a second gearbox that receives the torque from the first gearbox via the shaft (310) and reduces the torque and a second brake receptive of the torque from the second gearbox and configured to hold the flap in place through the second actuator.

The present invention discloses a landing gear position indication system and indication method, which obtains signals collected by multiple position sensors at various parts of the landing gear; performs an OR operation on the analog-to-digital converted signals of the position sensors located at the same part of the landing gear to obtain a first indication result; and / or performs an AND operation on the analog-to-digital converted signals of the position sensors located at different parts of the landing gear to obtain a second indication result. The lower position indication result of the landing gear related to the safety type I failure event is connected to the indicator light box and displayed through a hard wire, and the landing gear position indication result is connected to the integrated indicator through a bus and displayed. The present invention meets the requirements of independence while realizing the redundancy function, and has the characteristics of simple system architecture, high reliability and high safety, providing a strong guarantee for the accurate landing gear position indication and the flight safety of the aircraft.

An aircraft landing gear shock absorber strut (50) comprising: an outer cylinder (52) and a sliding tube (54) being movably mounted within the outer cylinder so as to be movable between a compressed position and an extended position, further comprising a first piston head (56) including a first axial surface upon which pressurised oleo-pneumatic shock absorber fluid within an oleo chamber acts to force the sliding tube to move from the compressed position to the extended position; a second piston head (60) slidably mounted within the sliding tube bore; a piston restraining arrangement (62) configured to limit axial separation between the second piston and the outer cylinder; and a first abutment (68) within the sliding tube bore on the opposite side of the second piston head with respect to the oleo chamber, the first abutment being spaced from the second piston head by a first distance when the second piston head is in the first position and the sliding tube is in the extended position, the first abutment being arranged to engage the second piston head as the sliding tube moves from the extended position to the compressed position.

To provide a system and method for dynamically measuring a blade position during flight of a rotorcraft.SOLUTION: The invention provides a blade positioning system and method in order to dynamically measure a position of a blade 14 during flight of a rotorcraft 10. The blade is repeatedly illuminated by a light source while the blade is rotating. The method also includes detecting radiation scattered from the blade in response to the illumination of the blade. The method further includes determining at least one of a blade pitch angle (ΔθP), a blade flap angle (ΔθF), a blade leading position and a blade lagging position on the basis of the radiation that is scattered from the blade and detected. Also provided is a chip-scale light detection and ranging (LIDAR) sensor configured to illuminate a plurality of blades while the blades are rotating in order to permit the blade position to be measured or to illuminate terrain beneath the rotorcraft in order to perform altitude measurement.SELECTED DRAWING: Figure 1

A method for unlocking a door (20) of an aircraft landing gear movably mounted between an open position and a closed position, the door (20) being associated with attachment means comprising a first attachment housing (23.1) and a second attachment housing (23.2) arranged to individually lock the door (20) in the closed position, a first unlocking actuator (22.1) and a second unlocking actuator (22.2) arranged to respectively control the first attachment housing (23.1) and the second attachment housing (23.2) with a view to deactivating them, and means for detecting (40) the door in the closed position, the method comprising the step of sequentially deactivating the attachment housing in order to detect whether the last attachment housing has failed to deactivate. A device and an aircraft using said method.

Methods and system of transporting cargo in a vehicle. A method includes after first cargo has been loaded into a cargo hold of a vehicle at a first location, determining available space within the cargo hold of the vehicle with the available space being unoccupied by the first cargo. After the vehicle leaves the first location, determining second cargo that is sized to fit in the available space. Loading the second cargo into the available space of the vehicle at a second location. The method also includes transporting the first cargo and the second cargo to a third location.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In one embodiment, a computer-implemented method is disclosed, comprising: measuring one or more state variables of the aircraft; inputting the one or more measured state variables to a prioritization scheme configured to determine an optimized actuator setting; determining one or more actuator commands based at least in part on inputting the one or more measured state variables to the prioritization scheme; and automatically controlling at least one actuator of the aircraft based on the determined one or more actuator commands.

Disclosed are methods and devices for determining a load condition of an aircraft. A sensing device is mountable relative to an aircraft strut during loading operations and demountable after loading operations. A data processing device is operable to process output from the sensing device together with aircraft specific data, to determine a load condition of the aircraft.

The present invention discloses a front wheel turning hand wheel deflection angle measuring device, comprising a scale plate, a front splint, a rear splint and a T-shaped connector, wherein: the scale plate is a disk-shaped structure with a fan-shaped notch, a semicircular notch is left at the center of the scale plate, and scales are arranged on the circumference of the scale plate; the front splint and the rear splint are splicable structures, the middle parts of the two are both arc-shaped structures, and after splicing, a clamping hole for clamping a front wheel turning hand wheel transmission rod is formed; one end of the front splint and the rear splint are connected by a pin shaft, and the other ends of the front splint and the rear splint are openable or locking structures; a T-shaped connector is fixed on the side of the front splint, the T-shaped connector has an upper pointer and a lower pointer, the lower pointer is connected and fixed to the 0° position of the scale on the scale plate; the upper pointer and the lower pointer point in opposite directions, the upper pointer is used to align with the triangular neutral mark on the front wheel turning hand wheel, and the lower pointer is used to align with the neutral mark line on the light guide plate in the cockpit.

The present invention relates to an arrangement (36; 36'; 36") configured to detect failure of and / or monitor integrity of a drive load path (P) for driving a high lift surface (17; 19) of an aircraft (1). The arrangement comprises a detection device (71; 71'; 71") arranged and adapted to detect an interruption of the drive load path or at least partial loss of load transmitted along the same. The arrangement is configured to provide a signal indicating failure of the drive load path and / or a signal indicating integrity of the drive load path depending on a condition, value or measure detected by the detection device. Moreover, the invention proposes a high lift system of an aircraft including an arrangement in accordance with the invention. Still further, the invention relates to a method of detecting failure and / or monitoring integrity of a drive load path (P) for driving a high lift surface (17; 19) of an aircraft (1).

A method for controlling an overdetermined system having multiple actuators, such as an aircraft (1) having multiple drive units (3), the actuators performing at least one primary task and at least one non-primary task, the method comprising: a) determining pseudo-control commands based on a physical model of the system, the pseudo-control commands specifically representing torques (L, M, N) and total thrust (F) acting on the system; b) based on u p =D u Determine the control matrix D, so that u1 = D ‑1 u p , represents the control instructions for the actuator to perform the main task; c) map the non-main task into the null space N(D) of the main task, so that when u2, represents the control instructions for the actuator to implement the non-main task, Du2 = 0, and d) provide the control instructions from b) and c) to the actuator.

The invention concerns an uplock for the selective retention of an olive (200) of a moving element, the uplock comprising a hook (3) that can move between a release position and a retaining position of the olive of the moving element, and a locking member (10) that can move between a locking position, in which it immobilises the hook in the retaining position, and an unlocking position, in which the hook is free to move under the thrust of the olive of the moving element, the uplock being equipped with detection means which include a sensor (31) associated with a single target (32) that can move between two positions so that the sensor emits a signal which is modified when the target moves from one of the positions to the other, the target being in kinematic interaction with the hook and the olive so that the target is maintained in one of its positions and can only be brought into the other of its positions if the hook is in the retaining position and the olive is actually present in the hook.