69results 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.

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.

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 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.

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.

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.

A method and system for automatically determining the actual relative mechanical backlash between at least two actuators of an aircraft control surface is proposed herein. The method comprises: executing (302) a position control loop to position said at least two actuators, collecting, for at least a predetermined period of measurements: first data and second data, determining, from the first data, a theoretical relative displacement without mechanical backlash between said at least two actuators and determining (303), from said second data and the theoretical relative displacement, an actual relative mechanical backlash between said at least two actuators, and, if the actual relative mechanical backlash is greater than or equal to a predetermined threshold, then issuing (305) an alert message.Thus, this disclosure makes it possible to significantly reduce the time required to diagnose actual mechanical play in aircraft control surfaces.

A tertiary lock system for an aircraft thrust reverser, including: a lock member actuatable between a locked position, in which a load path of the lock member engages a movable component to prevent deployment of the thrust reverser, and an unlocked position, in which the lock member permits deployment of the movable component of the thrust reverser; wherein the lock member comprises an element arranged such that at least a portion of the element extends into or through the load path of the lock member; and a locking verification system including a sensor module arranged for interrogating the element when the lock member is in the locked position; wherein the locking verification system is configured to: confirm that the tertiary lock system has been commanded to actuate the lock member into the locked position; instruct the sensor module to interrogate the element; and verify that the lock member is in the locked position and the load path is uncompromised in response to determining that the element is intact.

A flight control surface actuation system (10) includes a rotary actuator (12) having an output part configured to rotate about an axis (R-R), a connecting rod (16) for connecting a flight control surface to the rotary actuator, and a connection mechanism connecting a first end (18) of the connecting rod to the output part such that rotation of the output part about the axis causes at least part of the connection rod to rotate about the axis. The connection mechanism comprises a connection structure (22) fixed to the output part and an engagement part (50) mounted for rotation about a longitudinal axis. The engagement part extends through the connecting rod and is fixed to it. A position sensor (30) is arranged to measure the angular displacement of the first end of the connecting rod about the axis. A first part of the position sensor is secured to the connection structure in a fixed position. A second part of the position sensor is rotatably mounted to the first part and is configured to rotate with the engagement part and the connecting rod. The position sensor is configured to measure the relative rotation between the first and second parts of the position sensor.

Method for determining at least one characteristic of the motion of a moving structure of an aircraft and aircraft comprising a device for its implementation. The invention relates to a method for determining, at a given instant, at least one characteristic of the motion of a moving structure of an aircraft, characterized in that the method comprises: a step of determining a pair of coordinates in a planar reference frame corresponding to a projection of a frame (R1 to R5) fixed to the moving structure in said planar reference frame; a step of determining the characteristic of the motion of the moving structure by comparing the previously determined pair of coordinates with reference pairs of coordinates, each reference pair of coordinates being associated with at least one characteristic of the motion of the moving structure. The invention also relates to an aircraft comprising a device for implementing said method. Figure 1

Systems and methods for backlash measurement and monitoring are disclosed for an actuator having a motor end and an output end. The method may include: driving the motor end of the actuator to one endpoint; measuring, using a first sensor, a motor-end initial position value associated with the motor end; measuring, using a second sensor, an output-end initial position value associated with the output end; driving the motor end of the actuator to another endpoint; measuring a motor-end final position value using the first sensor and an output-end final position value using the second sensor; determining a backlash value based on the motor-end initial position value, output-end initial position value, motor-end final position value, and output-end final position value; generating a signal based on comparing the backlash value to a predetermined threshold; and displaying a notification indicative of an operational condition of the actuator based on the signal.

A bush (100, 200, 300) is provided comprising a sleeve (102, 214) for receiving a pin; a shoulder (104, 216) extending from the sleeve, the shoulder for delimiting insertion of the sleeve into an aperture; and a sensor mount (108) 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.

The invention relates to a locking system for locking a door in a closed position relative to a carrier, a door and an aircraft. A locking system (20) for locking a door (15) comprising at least two locks (31). The locking system (20) has a position sensor (40) for each lock (31), each position sensor (40) emitting a measurement signal carrying a locked state or an unlocked state of the associated lock (31), the signal system (30) comprising a controller (50) receiving the measurement signal and connected to an alarm (60), the controller (50) is configured to issue an unlock signal (SDV) when at least one measurement signal carries an unlocked state of the associated lock, or to issue a lock signal if each measurement signal carries a locked state of the associated lock, and the alarm (60) generates a visual unlock alarm after receiving the unlock signal, and a visual locking alarm is generated after the locking signal is received.

A system and a method include an emergency exit door for an internal cabin of an aircraft. The emergency exit door includes a lever configured to be engaged to open the emergency exit door. A solenoid is operatively coupled to the lever. The solenoid in a deactivated state is coupled to the lever to lock the lever and prevent the lever from being pulled to open the emergency exit door. The solenoid in an activated state is uncoupled from the lever and allows the lever to be pulled to open the emergency exit door.

A firefighting float plane (10) having a fuselage (112), one or more wings (110), two floats (114, 116), a water scoop (120, 122) associated with each float for filling a water tank (300) of the firefighting float plane, and systems and / or methods for deploying the water scoops to prevent (and / or mitigate the effects of) asymmetric deployment and / or retraction of the water scoops.

A firefighting float plane having a fuselage, one or more wings, two floats, a water scoop associated with each float for filling a water tank of the firefighting float plane, and systems and / or methods for deploying the water scoops to prevent (and / or mitigate the effects of) asymmetric deployment and / or retraction of the water scoops.

A mechanical architecture is provided for aircraft distributed flap actuation and includes a shaft and first and second actuation units. The first actuation unit 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 includes a second actuator that receives the torque from the first gearbox via the shaft 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 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.

A flight control surface actuation system includes a rotary actuator having an output part configured to rotate about an axis (R-R), a connecting rod for connecting a flight control surface to the rotary actuator, and a connection mechanism connecting a first end of the connecting rod to the output part such that rotation of the output part about the axis causes at least part of the connection rod to rotate about the axis. The connection mechanism comprises a connection structure fixed to the output part and an engagement part mounted for rotation about a longitudinal axis. The engagement part extends through the connecting rod and is fixed to it. A position sensor is arranged to measure the angular displacement of the first end of the connecting rod about the axis.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft and other types of aircraft driven by an electric propulsion system. 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, the 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.

A drive assembly for driving a movable flow body of an aircraft comprises an electric motor having a two sets of independent windings, two motor control electronics units coupled with the windings and a control computer, two actuators couplable with a first or second section of the flow body, a first and a second transmission shaft, wherein the transmission shafts each have a first and a second end, wherein the electric motor is coupled with the first ends of the transmission shafts, wherein the second ends of the transmission shafts are coupled with the respective actuator, and wherein the drive assembly is designed to selectively move and hold the movable flow body into a plurality of extended positions and a retracted position relative to a fixed structural component of the aircraft by selectively moving and holding the first actuator and the second actuator.