Abstract: A multimodal wearable band which uses mechanical vibrations to stimulate sensory neurons in the wrist or ankle in order to reduce the severity of tremors, rigidity, involuntary muscle contractions, and bradykinesia caused by neurological movement disorders and to free users from freezing induced by movement disorders. The device uses sensors to provide output used by a processing unit to determine the optimal stimulation pattern for each user and to determine when stimulation is necessary, and then uses one or more vibration motors to accordingly stimulate the user's neurological pathways to lessen the severity of a user's symptoms. The device can also be adapted to integrate with 3rd party devices.

Abstract: A system of fall back flight control configured for use in electric aircraft includes an input control configured to receive a pilot input and generate a control datum. System includes a flight controller communicatively coupled to the input control and configured to receive the control datum and generate an output datum. The system includes the actuator having a primary mode in which the actuator is configured to move the at least a portion of the electric aircraft as a function of the output datum and a fall back mode in which the actuator is configured to move the at least a portion of the aircraft as a function of the control datum. The actuator configured to receive the control datum, receive the output datum, detect a loss of communication with the flight controller, and select the fall back mode as a function of the detection.

Abstract: A dust extractor (1) includes a motor (30) and a blower fan (40), which is disposed downward of the motor and is connected to the motor. The dust extractor further includes a flexible (pliable) support member (50; 500), which has a lower-end portion (52) connected to the motor and suspends the motor.

Abstract: A crossbar assembly for facilitating isolation of a sensor assembly from vibration comprises an outer crossbar segment, an inner crossbar segment, and an isolator. The outer crossbar segment comprises a payload mount interface and an outer isolator interface operable to mount to an isolator. The inner crossbar segment comprises a structure interface and an inner isolator interface operable to mount to the isolator. The isolator can be supported by the outer and inner crossbar segments. The isolator comprises a first wire rope assembly comprising wire ropes extending longitudinally from the outer crossbar segment to the inner crossbar segment, and a second wire rope assembly comprising a wire rope extending circumferentially between the outer and inner crossbar segments. The isolator operates to partially decouple the outer crossbar segment from the inner crossbar segment and dampen vibrations propagating between the outer and inner crossbar segments.

Abstract: An assembly for manual control of an HSTA for controlling the position of a moveable surface, the assembly comprising a user-operated manual control element (1?) e.g. a trim wheel in the cockpit, a first motor and a first resolver connected to the manual control element and a second motor and a second resolver arranged to communicate with the first motor and the first resolver and to cause corresponding movement of the actuator, in use.

Abstract: A system of fall back flight control configured for use in electric aircraft includes an input control configured to receive a pilot input and generate a control datum. System includes a flight controller communicatively coupled to the input control and configured to receive the control datum and generate an output datum. The system includes the actuator having a primary mode in which the actuator is configured to move the at least a portion of the electric aircraft as a function of the output datum and a fall back mode in which the actuator is configured to move the at least a portion of the aircraft as a function of the control datum. The actuator configured to receive the control datum, receive the output datum, detect a loss of communication with the flight controller, and select the fall back mode as a function of the detection.

Abstract: A roll control system controls roll control surfaces of an aircraft that are capable of causing the aircraft to perform a roll maneuver by respectively deflecting in upward and downward directions. The roll control system includes deflection limiter units for respectively limiting angles of deflection of the roll control surfaces, and further includes deflection rate limiter units for respectively limiting rates of deflection of the roll control surfaces. The deflection limiter unit limits the roll control surfaces to deflection distances and deflection rates based at least in part on the deflection direction of the roll control surfaces. For a given set of flight conditions, such as airspeed, if a roll control surface is deflecting upwardly, it is less limited by the roll control system in terms of deflection distance and deflection rate than if the roll control surface is deflecting downwardly.

Abstract: An actuator-link assembly includes an actuator having first and second opposing ends in order to drive a control surface of an aircraft and a link coupled to the actuator. The actuator has a rod portion having a tip end, located at first end of the actuator, said tip end pivotably attach to a control surface of an aircraft. The link includes a pair of linear portions, each having first and second opposing ends, wherein the linear portions are disposed alongside each other; a coupling portion connecting the first ends of linear portions, a fulcrum shaft attachment portion protruding from the coupling portion configured to pivotably attach to a fulcrum shaft for rotatably supporting the control surface; and an actuator attachment portion configured to pivotably attach to the second end of the actuator. A material constituting the linear portions and the coupling portion contains fiber reinforced plastic.

Abstract: A method, apparatus, and computer program product for managing movement of a flight control surface on an aircraft. A control signal is received to move the flight control surface to a position. Travel in a number of actuators in a plurality of actuators coupled to the flight control surface is limited to an amount that reduces a load on the number of actuators in the plurality of actuators in response to receiving the control signal while the aircraft is on the ground and the speed of the aircraft is less than a threshold.

Abstract: An aircraft is provided with a mechanical linkage that mechanically transmits an amount of manipulation of a control stick to a rudder controlling mechanism, an amount-of-manipulation sensor that detects the amount of manipulation, and an FBW control-law computing unit that calculates an FBW control-law command indicating an amount of control for a craft in accordance with the detected amount of manipulation. Also, with the aircraft, a high-pass filter outputs, among the control-stick manipulation amounts detected by the amount-of-manipulation sensor, an amount of change in the control-stick manipulation amounts; by subtracting the amount of change that has passed through the high-pass filter by means of the subtracting portion, the FBW control-law command serves as an SAS command; and the rudder controlling mechanism is driven by adding an amount of movement of an SAS actuator in accordance with the SAS command to the control-stick manipulation amount transmitted by the mechanical linkage.

Abstract: Secondary air flow is provided for a ducted fan having an engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides a first flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades.

Abstract: The present invention relates to haptic feedback for operating at least one manual flight control device (21) for controlling the cyclic pitch of the blades (5) of a rotary wing (4) of a hybrid helicopter (1) via power assistance (27). Said operations of said control device (21) are defined according to a predetermined damping force relationship (A) that is a function of an instantaneous load factor of the hybrid helicopter (1) in such a manner that the instantaneous load factor is maintained between its minimum and maximum limit values in proportion to a position of a thrust control member (20) between its minimum and maximum thrust values (23, 22).

Abstract: A method, apparatus, and computer program product for managing movement of a flight control surface on an aircraft. A control signal is received to move the flight control surface to a position. Travel in a number of actuators in a plurality of actuators coupled to the flight control surface is limited to an amount that reduces a load on the number of actuators in the plurality of actuators in response to receiving the control signal while the aircraft is on the ground and the speed of the aircraft is less than a threshold.

Abstract: A method and device for generating a speed profile for an aircraft rolling on the ground. The device (1) comprises means (8) for automatically determining a speed profile which is suited to successive elements of a ground rolling trajectory and which complies with maximum speeds and particular constraints.

Abstract: Rudder assist mechanisms and methods are capable of being operably connected to an aircraft's rudder control system. The rudder assist mechanisms most preferably have over-the-center spring biasing functions so as to cause either substantially no spring force (i.e., when the linkage is over the spring-bias center) or substantially all spring force (i.e., when the linkage is left or right of the spring-bias center) to be exerted on the rudder control system. The rudder assist mechanism may include a control spring assembly, and a linkage assembly which operably connects the control spring assembly to the rudder control assembly. The linkage assembly is moveable operably between a null position wherein substantially no spring force of the control spring assembly is transferred to the rudder control system by the linkage assembly, and right and left spring-biased positions wherein right and left spring forces of the control spring assembly are transferred to the rudder control system, respectively.

Abstract: The present invention relates to a bellcrank (40) for a flight-attitude-changing linkage (27, 25A, 28A) of the variable gain type. The linkage (27, 25A, 28A) connects a manual flight control device of a rotary wing aircraft (1), e.g. a hybrid helicopter, to at least one airfoil of said aircraft (1) suitable for causing a change in flight attitude. Said bellcrank (40) includes first and second sliders (50, 52). Said first slider (50) is connected to a movable portion (57, 61) of a setpoint flight-attitude-changing linkage (24) that delivers a setpoint, with movements of said linkage causing the gain of said variable-gain linkage (27, 25A, 28A) to be adjusted, which gain is proportional to the position of said movable portion (57, 61).

Abstract: The present invention provides systems and methods for controlling the speed of flap retraction on aircraft, and alerts to the pilot of potentially unsafe flap position. The invention accepts direction from a pilot and senses operation of the aircraft to determine appropriate flap position and flap retraction speed.

Abstract: A system and method for providing variable motion-resisting loads for a side stick controller on an aircraft where the primary resistive load is provided by a mechanical spring mechanism which is augmented or supplemented by an electric servo motor and electric servo motor linkage system.

Abstract: A rotor aircraft has an adjusting mechanism that controls the sensitivity of the control stick relative to fore and aft tilt of the rotor. The control stick is pivotable between fore and aft directions as well as laterally about a control stick pivot point. A rotor linkage is connected between the control stick and the rotor rotor head. The rotor linkage assembly tilts the rotor head in response to tilting movement of the control stick. The linkage assembly has a control point that rotates at a radius about the control stick pivot point. An adjusting member located between the control stick and the linkage assembly can be moved to change the radius of the control point to the control stick pivot point. The change in radius corresponds to the amount of tilt that the rotor head make while the control stick moves between full aft and full forward positions. The control stick also controls ailerons and a horizontal stabilizer but these control services are not affected by the adjusting member.

Abstract: The present invention relates to a flight control device for an aircraft comprising a control (2) and a means for actuating a controlled member (RP, RQ) to which a command is applied. According to the invention, said device (1) additionally includes a sensor (E) for determining a second value that is representative of the control executed by the aircraft (He) with respect to the control axis, and second means (M2) which determine: as long as the second value is lower than or equal to a reference value, a first trim command that is proportional to the actuation of the control (2); and when the second value is higher than the reference value, in addition to a second trim command, a speed command that is proportional to the additional actuation over and above said second value.

Abstract: An aircraft rudder command system for allowing a pilot to directly input a sideslip command for yaw-axis control through use of the rudder pedals is disclosed. The aircraft rudder command system includes a pedal input device for receiving a pedal input signal indicative of pilot rudder pedal input, a signal-receiving device for receiving feedback signals indicative of the current state of aircraft operation, a command generator system responsive to the pedal input signal and at least one of the feedback signals for generating a sideslip angle command, command limiting means for generating a limited sideslip angle command, and a feedback control system for transmitting a sideslip control rudder command to a rudder actuation system. The rudder actuation system causes the rudder to move in such a manner so as to produce an actual aircraft sideslip angle which follows the limited sideslip angle command. The aircraft rudder command system may also include a sideslip estimator.

Abstract: Control apparatus, such as an aircraft throttle box, includes a manual control member 8 linearly movable on a first shaft 2. The position of the control member is sensed by rotation of the shaft 2 to generate an output control signal. In order to balance the apparatus, especially against high g-forces, a second shaft 3 is provided, parallel to the first carrying a counterweight 13 that moves together with the control member 8, but in the opposite direction.

Abstract: Aircraft supported by rotors have an elastic device (28) that helps to angularly place the body of the aircraft in relation to the rotation axis of the rotor (1). This device has adjusting means (39-40-27) which the pilot can use during flight. The invention is in particular applicable to autogyros.

Abstract: A rudder control linkage having thermal compensation features employs a stationary thermal compensation rod, a lever with a fixed pivot, a reversing link with a movable pivot, a summing link, pilot and yaw damper input rods, and output rods. In an airplane having a fin and a rudder, the linkage is utilized to cancel undesired yaw inputs to the rudder caused by fin bending and differential thermal expansion which, if not corrected, would deleteriously affect airplane trim.

Abstract: Hand movement assistance apparatus of both the active and passive types and usable for movement stabilization by either a normal human subject seeking to perform a task of great precision or a neurologically impaired human subject having Parkinson's disease or similar dysfunction. Both the active and passive apparatus provide a glove-like receptacle for the subject's hand; the passive system applies closed-loop feedback system generated forces to the glove and the active system provides a hand guidance track arrangement for the subject's hand. Three axis stabilization forces and rotationally oriented forces are alternatively provided by the active system, along with responses particularly adopted to the Parkinson's tremor frequency.

Abstract: An all-electric powered flight control system for an aircraft incorporating fault-tolerant avionics combined with a redundant power actuation control system utilizing electro-mechanical devices for actuating aerodynamic control surfaces; and wherein, the redundancy is implemented through the division of the total aerodynamic surface area to be actuated into multi-segments, with each control segment having its own channel in the avionic control system. The avionic control system is effective in minimizing the effects of power actuated control loads reacting through aircraft structure by the incorporation of an electronic gain control power module which is positioned between the power output side of an actuator (e.g., an electro-mechanical or electro-hydraulic device) and the aerodynamic control surface or segment being actuated.

Abstract: A biomechanical feedback arrangement wherein a varying force tending to improve the neuromotor tracking response of a human subject, particularly in the presence of lateral or front-back G force fields, is added to the test subject input member of a feedback control system. Use of the biomechanical feedback in a high-performance aircraft and in a ground-based simulator apparatus is also disclosed, along with comparison results from simulator testing of human subjects.

Abstract: An all-electric powered flight control system for an aircraft incorporating fault-tolerant avionics combined with a redundant power actuation control system utilizing electro-mechanical devices for actuating aerodynamic control surfaces; and wherein, the redundancy is implemented through the division of the total aerodynamic surface area to be actuated into multi-segments, with each control segment having its own channel in the avionic control system. The avionic control system is effective in minimizing the effects of power actuated control loads reacting through aircraft structure by the incorporation of an electronic gain control power module which is positioned between the power output side of an actuator (e.g., an electro-mechanical or electro-hydraulic device) and the aerodynamic control surface or segment being actuated.

Abstract: In a control system having a manipulatable unit, such as an aircraft stick, which is biased toward a neutral position to provide control "feel", an input which at each instant signifies the position of displacement of the stick from its neutral position is processed to provide an output signifying the magnitude of the biasing force to be applied to the stick. In a range of positions spaced to one side of its neutral position (or in such a range at each side of neutral) the output provides for a steep biasing force gradient as the stick is moved through a predetermined distance in either direction from a turning point to which the stick had been brought by movement in the opposite direction. For all other movements of the stick within that range the biasing force has a low gradient. The higher gradient of bias for small trimming movements provides better feel for trimming without requiring stick bias to be unduly stiff during coarser stick displacements.

Abstract: An improvement to an aircraft control stick is disclosed. Movement of an aircraft may generate a force which undesirably causes the aircraft pilot to deflect the aircraft's control stick, which thereby results in the aircraft deviating from the desired flight path. The present invention includes a processor-based system which employs an algorithm that generates a signal for causing the control stick to resist such forces. The invention includes a spring and damper connected to the control stick so that the spring constant and damping ratio may be varied.

Abstract: The present aircraft steering mechanism provides different steering characteristics for different aircraft speeds, especially at low speeds on the one hand and at high speeds on the other hand. For this purpose the steering mechanism, which is operatively interposed between a control input and a controlled member, such as a rudder or flap, forms a transmission system including a walking-beam type lever with a free-end member and a forked member. The forked member is movable relative to the free-end member in response to a flight parameter, such as the dynamic pressure, whereby the free-end member dips more or less into the forked member. Thus, the excursion of the free-end member is increasingly more limited as the speed of the aircraft increases. The see-saw type lever forms a control input and a control output and the degree of excursion is determined by the extent to which the free-end member dips into the forked member.