198results about "Movable aircraft element position indicator" patented technology

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

The present disclosure relates to a system for detecting the position and / or the operating state of a movable component of an aircraft, wherein the system has one movable component, two actuators, two motion sensors, one asymmetry sensor, one drive train, and one monitoring device; wherein the component can be actuated by the at least two mutually offset actuators; wherein at least one motion sensor is provided for one respective actuator; wherein the motion sensor can monitor a mechanical movement of the drive train; wherein the monitoring device is connected to the motion sensors and to the asymmetry sensor, and wherein the system is configured such that an initialization of the monitoring device and / or of the motion sensors and / or of the asymmetry sensor can take place after a booting of the system and / or when the component has adopted a fully retracted and / or a fully extended position.

A digital surface position sensor includes a position sensor, an adaptable hardware interface, and a processing circuit. The position sensor is adapted to be coupled to an aircraft flight control surface and is configured to sense a position of the aircraft flight control surface and supply a position signal representative thereof. The adaptable hardware input supplies an identification signal that identifies the aircraft flight control surface to which the position sensor is coupled. The processing circuit is coupled to receive the position signal and the identification signal. The processing circuit is configured, upon receipt of the signals, to process the position signal and the identification signal and generate (i) a first digital position signal representative of the position of, and the identification of, the aircraft flight control surface and (ii) and independent second digital signal representative of the position of, and the identification of, the aircraft flight control surface.

Systems and methods for measuring an axial position of a phonic wheel or other rotating component are provided. The system includes a phonic wheel rotatable about a rotation axis and translatable along the rotation axis, a first sensor, a second sensor and a computer. The phonic wheel includes an inclined tooth having an axially non-uniform radial height and a reference tooth having an axially uniform radial height. The first sensor generates a positioning signal indicative of a gap between the inclined tooth and the first sensor. The second sensor generates a reference signal indicative of a gap between the reference tooth and the reference sensor. The first and second sensors have different orientations. The computer generates an output indicative of the axial position of the phonic wheel based on the positioning signal and the reference signal.

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.

An aircraft notification system, the aircraft notification system having a processor and a memory communicatively connected to the processor, the memory containing instructions configuring the processor to receive aircraft angle data from a blended wing body aircraft (BWB) having a main body and a wing with no clear demarcation between the wings and the main body along a leading edge of the aircraft, determine an aircraft orientation as a function of the aircraft angle data, wherein the aircraft orientation comprises at least a pitch angle of the BWB while the BWB is in contact with a ground surface, transmit the aircraft orientation to an aircraft display system and visually display the aircraft orientation through the aircraft display system.

In an aspect, a system for propeller parking control for an electric aircraft. The system include at least a sensor and a computing device. A sensor may be configured to generate angular datum. The computing device may be configured to generate a trajectory as a function of angular datum. The computing device may also be configured to initiate the transition from hover to fixed-wing flight as a function of a trajectory.

A system monitors an operational state of a movable component, e.g. flaps 1, of an aircraft. The system uses a combination sensor 3 to sense both position and torque e.g. a kinematic and / or kinetic state of the flap 1. The operational state may be the skew or disconnect of the flaps 1.

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

The invention relates to a method for determining at least one motion characteristic of a movable structure of an aircraft at a given point in time, characterized in that it comprises the following steps: determining a coordinate pair in a planar reference frame corresponding to a projection of a marking (R1 to R5) integral with the movable structure in said planar reference frame; -determining a motion characteristic of the movable structure by comparing previously determined coordinate pairs with reference coordinate pairs, each pair of reference coordinate pairs being associated with at least one motion characteristic of the movable structure. The invention also relates to an aircraft comprising means for carrying out said method.

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 of determining the position of one or more components of a landing gear assembly of an aircraft is disclosed including scanning the one or more components with a lidar system to generate a set of position data points, each position data point comprising a set of three orthogonal position values. The position data points are partitioned into one or more clusters using a distance metric. Each cluster is determined to represent a component of the landing gear assembly. The position of the components are then determined from the position data points in the clusters. The value of the distance metric for a first position data point and a second position data point is representative of the difference between a first position value of the three orthogonal position values of the first position data point and the corresponding first position value of the three orthogonal position values of the second position data point.

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.

The invention relates to methods for operating an inductive proximity sensor, comprising the steps of exciting the inductive proximity sensor by applying a voltage signal in the form of a step function to the voltage input of the proximity sensor, in particular by applying a square-wave voltage, Sampling the voltage signal applied to the proximity sensor in order to obtain corresponding voltage values, Sampling the current flowing in the proximity sensor in response to the application of the voltage signal in order to obtain corresponding current values, Summing up the sampled current values ​​of a first time interval of the step response, Calculating an average current value based on at least one current value from a second time interval of the step response following the first time interval, Calculating an average voltage value based on at least one voltage value from the second time interval, Determining a total system resistance (Rtot ) based on the calculated mean current value and the calculated mean voltage value, and determining a normalized current signal (i Σ,norm ) by dividing the summed sampled current values ​​of the first time interval by the calculated current mean value of the second time interval. The invention is characterized by looking up a lookup table or by using a description of a two-dimensional relationship, e.g. by a polynomial, to determine the inductance (L p ), where as input values ​​to look up a corresponding inductance (L p ) in the lookup table the determined total system resistance (R tot ) and the determined standardized current signal (i Σ,norm ) and comparing the determined inductance (L p) with at least one predetermined threshold to infer the presence or absence of a nearby target object.

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.

Disclosed is a wing (5) for an aircraft (1), comprising a main wing (11), a high lift body (15), and a connection assembly (17) movably connecting the high lift body (15) to the main wing (11), such that the high lift body (15) can be moved between a retracted position and at least one extended position, wherein the connection assembly (17) comprises a drive system (25) comprising a first drive unit (27) and a second drive unit (29), wherein the first drive unit (27) has a first input section (31) coupled to a drive shaft (33), a first gear unit (35) and a first output section (37) drivingly coupled to the high lift body (15), wherein the second drive unit (29) has a second input section (39) coupled to the drive shaft (33), a second gear unit (41), and a second output section (43) drivingly coupled to the high lift body (15), wherein the first output section (37) comprises a first drive arm (45) drivingly coupled to the high lift body (15) via at least one first link element (49) rotatably coupled to the first drive arm (45) and mounted to the high lift body (15), and the second output section (43) comprises a second drive arm (47) drivingly coupled to the high lift body (15) via at least one second link element (51) rotatably coupled to the second drive arm (47) and mounted to the high lift body (15). The object to prevent excessively high torque loads during skew cases of the high lift body, is achieved in that the first link element (49) comprises a first linear deformation element (53), and / or the second link element (51) comprises a second linear deformation element (55), and the first linear deformation element (53) and / or the second linear deformation element (55) have a non-linear force-displacement characteristic (62).

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.