128results about "Aircraft transmission means" patented technology

A crosswind kite system adapted to operate in an alternate mode in high winds. The system may operate at reduced efficiency in high winds in order to moderate loading on the system during those high winds. The system may use multi-element airfoils which are actuated to reduce the coefficient of lift of the airfoils in order to moderate loading in high wind conditions. Other flight aspects may be controlled, including flying the crosswind kite in side slip to induce drag which may lower loading on the system.

An electromechanical actuator (EMA) is provided. The EMA includes a threaded output ram connectable to a mechanical component and at least one motor module engageable with the output ram for controllably translating the output ram along a linear axis of the output ram. The actuator further includes a torque sensing adaptive control (TSAC) system for monitoring torque within the motor module. The TSAC generates a disengagement command signal when the TSAC system determines torque within the motor module is outside an allowable motor module torque range. The disengagement command signal initiates disengagement of the motor module from the output ram.

An electrical actuator comprises two or more electrical motors to drive a threaded screw ram. Each motor has an armature that drives a threaded roller screw that is engaged to the threaded screw ram. Each armature of each motor is independently engageable and / or dis-engageable with the ram. The motor armatures are engageable and disengageable by way of threaded roller screws, the ends of which are tapered to enable them to be lifted away from or lowered into engagement with the threaded ram.

The invention pertains to the field of security for aircraft and other vehicles, and more particularly to systems for preventing the hijacking, commandeering or suicide bombing of aircraft, or other vehicle(s). Aircraft system and vehicle system embodiments of the invention have one or more kinds of automating computers capable of safely controlling either an aircraft or vehicle in one or more types of common, well-proven, or yet-to-be-developed, computer-automated modes. The system provides mechanical control linkage disabling means interfaced with control signal receiving means responsive to wireless, or hard-wired, transmitted security-related control signal(s). Control linkage disabling means are located within a series of physical control linkage components of a vehicle at a point subsequent to where manual control input is initiated and prior to where computer automated control is provided. The disabling means renders ineffectual the mechanical control needed for one or more humans to control or direct a vehicle.

According to a first aspect of the invention, there is provided a method for operating a multicopter experiencing a failure during flight, the multicopter comprising a body, and at least four effectors attached to the body, each operable to produce both a torque and a thrust force which can cause the multicopter to fly when not experiencing said failure. The method may comprise the step of identifying a failure wherein the failure affects the torque and / or thrust force produced by an effector, and in response to identifying a failure carrying out the following steps, (1) computing an estimate of the orientation of a primary axis of said body with respect to a predefined reference frame, wherein said primary axis is an axis about which said multicopter rotates when flying, (2) computing an estimate of the angular velocity of said multicopter, (3) controlling one or more of said at least four effectors based on said estimate of the orientation of the primary axis of said body with respect to said predefined reference frame and said estimate of the angular velocity of the multicopter. The step of controlling one or more of said at least four effectors may be performed such that (a) said one or more effectors collectively produce a torque along said primary axis and a torque perpendicular to said primary axis, wherein (i) the torque along said primary axis causes said multicopter to rotate about said primary axis, and (ii) the torque perpendicular to said primary axis causes said multicopter to move such that the orientation of said primary axis converges to a target orientation with respect to said predefined reference frame, and (b) such that said one or more effectors individually produce a thrust force along said primary axis.

An aircraft flap (or slat) drive apparatus includes first and second centralized drive units that respectively rotationally drive first and second drive lines in both wings. The first drive line includes shaft segments connected end-to-end by selectively engageable separating devices. The outboard ends of the first and second drive lines are coupled through one of the separating devices, in each wing. At least one actuator mechanism connects each flap to a respective shaft segment of the first drive line, and converts the drive line rotational motion to a flap translational motion. If a component breaks, jams, or otherwise fails, it is isolated by disengaging the two adjacent separating devices with the faulty component therebetween, so that the rest of the apparatus remains functional. Each one or both of the drive units can drive one or more or all of the flaps through the interconnected first and second drive lines.

An electrical actuator comprises two or more electrical motors to drive a threaded screw ram. Each motor has an armature that drives a threaded roller screw that is engaged to the threaded screw ram. Each armature of each motor is independently engageable and / or dis-engageable with the ram. The motor armatures are engageable and disengageable by way of threaded roller screws, the ends of which are tapered to enable them to be lifted away from or lowered into engagement with the threaded ram.

A system for setting a span load distribution of a wing of an aircraft with a base flap system comprises at least one inboard flap element and one outboard flap element, which elements in the direction of the span are arranged on the trailing edge of the wing, and can be positioned relative to the span direction of the wing. The flap elements are not mechanically coupled with each other and are controlled independently of each other for the purpose of setting the span load distribution.

An aerodynamic component such as a helicopter rotor blade has an aerodynamic flow profile, a free end and a mounting end forming a blade root for attachment to a rotor mast. A blade section between the root and the free end has a leading edge and a trailing edge as viewed in the flow or chord direction. The blade is enclosed on its suction side and its pressure side with a respective skin. A nose flap or leading edge flap is secured to the leading edge by a bearing or hinge. The blade tilting angle is adjustable by piezo-drive elements arranged in at least one pair forming an actuator for each nose flap. The piezo-elements of a pair are arranged in the chord direction one behind the other and the pair is secured to the body of the blade by a fixed point positioned between the elements of a pair. The expansion and contraction from the piezo-element closer to the trailing edge is transmitted to the flap by a push rod. The expansion and contraction of the piezo-element closer to the leading edge is transmitted to the flap by a pronged pull-fork. The push rod and the pull-fork drive a free end of a lever secured to the flap, whereby a displacement of the push rod or of the pull fork causes the tilting adjustment of the flap by a push-pull action.

According to a first aspect of the invention, there is provided a method for operating a multicopter experiencing a failure during flight, the multicopter comprising a body, and at least four effectors attached to the body, each operable to produce both a torque and a thrust force which can cause the multicopter to fly when not experiencing said failure. The method may comprise the step of identifying a failure wherein the failure affects the torque and / or thrust force produced by an effector, and in response to identifying a failure carrying out the following steps, (1) computing an estimate of the orientation of a primary axis of said body with respect to a predefined reference frame, wherein said primary axis is an axis about which said multicopter rotates when flying, (2) computing an estimate of the angular velocity of said multicopter, (3) controlling one or more of said at least four effectors based on said estimate of the orientation of the primary axis of said body with respect to said predefined reference frame and said estimate of the angular velocity of the multicopter. The step of controlling one or more of said at least four effectors may be performed such that (a) said one or more effectors collectively produce a torque along said primary axis and a torque perpendicular to said primary axis, wherein (i) the torque along said primary axis causes said multicopter to rotate about said primary axis, and (ii) the torque perpendicular to said primary axis causes said multicopter to move such that the orientation of said primary axis converges to a target orientation with respect to said predefined reference frame, and (b) such that said one or more effectors individually produce a thrust force along said primary axis.

This invention is constant speed-differential speed, double shaft reverse output speed changing method and its device. A set of driving mechanism drive two groups of planet gear trains. The said planet gear trains drive two output shafts with same rotating speed and reverse direction. Another driving mechanism drives two control gears to rotate in same direction. The said control gears are assembled on the same control shaft and are respectively joggled to gear ring outer tooth of the said two sets of planet gear trains. Thus, the control gear can drive the gear ring of planet gear train to rotate in same speed. When the rotating direction of one gear ring is the same as the output direction of its corresponding planet gear train, the rotating direction of the other gear ring will be opposite to its corresponding planet gear train. This invention can be applied widely in the field of military affairs and civil aviation.

An electromechanical actuator (EMA) is provided. The EMA includes a threaded output ram connectable to a mechanical component and at least one motor module engageable with the output ram for controllably translating the output ram along a linear axis of the output ram. The actuator further includes a torque sensing adaptive control (TSAC) system for monitoring torque within the motor module. The TSAC generates a disengagement command signal when the TSAC system determines torque within the motor module is outside an allowable motor module torque range. The disengagement command signal initiates disengagement of the motor module from the output ram.

A modularized airplane includes two separably interconnected modules. The first module, incorporating substantial airplane style characteristics, comprises wings and stabilizers having at least one associated control surface joined with a fuselage portion. The second module carries a set of essential airplane components interconnected to suffice airplane operations including a propulsion unit, radio receiver and / or auto-piloting electronics unit, control servo devices, and power means mounted on provided support structures. At least one pair of mutually magnetically attractive connectors oppositely affixed on said two module members provide means for inter-modular structural connection. An alignment mechanism having oppositely provided complementary structures on said two modules facilitate alignment as well as lateral interlocking for inter-modular structural connection. A linkage means includes two linkage portions separably connected by two mutually magnetically attractive elements oppositely mounted thereon to form a linkage assembly for linking between control surfaces and associated servo devices, a linkage guide means, and a stress isolation mechanism. The structural connection and the control linkage means facilitate substantially effortless inter-modular connections to form a functional modularized airplane as well as excessive parting tension induced nondestructive inter-modular disconnection. Multiple modules of differing styles and aerodynamic specifications can be made as the first module to be interconnected with the second module to form differing airplanes, sharing essential components, for various applications.

A crosswind kite system adapted to operate in an alternate mode in high winds. The system may operate at reduced efficiency in high winds in order to moderate loading on the system during those high winds. The system may use multi-element airfoils which are actuated to reduce the coefficient of lift of the airfoils in order to moderate loading in high wind conditions. Other flight aspects may be controlled, including flying the crosswind kite in side slip to induce drag which may lower loading on the system.

The invention concerns an aircraft high-lift system with a drive system, components for transmitting the drive energy over the entire wing span to drive stations of individual segments of landing flap / slat flap systems, and with overload protection. According to the invention, the overload protection consists of electrical load sensors positioned at the drive-energy in-take points of the individual power trains on the flaps.

A solar powered air vehicle that can stay aloft for indefinite periods of time. The vehicle employs photovoltaic solar cells for primary power and high speed counter-rotating flywheels for energy storage and steering of the vehicle. The flywheels are placed in the wing to reduce airfoil drag. A control law provides three-axis stabilized control of the vehicle by controlling propeller pitch to vary the speeds of the flywheels.

The invention provides a trailing edge with adjustable degrees of curvature for an aircraft wing. The aircraft wing comprises a middle wing part with the fixed degree of curvature. The middle wing part comprises a main beam. The trailing edge with the adjustable degrees of curvature comprises a rear beam, a plurality of trailing edge pole group structures and a trailing edge driving mechanism, wherein the trailing edge pole group structures are arranged in parallel. The rear beam is connected with the rear end of the main beam. The trailing edge driving mechanism comprises rockers driven by a driving device. One ends of the rockers are connected with the rear beam in a rotatable mode, and the other ends of the rockers are connected with the trailing edge pole group structures in a rotatable mode. The trailing edge pole group structures are plane pole group structures based on 5R closed-loop units and respectively have one degree of freedom. Each trailing edge pole group structure comprises a plurality of pairs of 5R closed-loop units, a plurality of networking triangular units, a tail-end 4R closed-loop unit and a tail-end triangular unit. Each pair of 5R closed-loop units comprises the upper 5R closed-loop unit and the lower 5R closed-loop unit. Meanwhile, the invention provides a leading edge with adjustable degrees of curvature for the aircraft wing.

This invention relates to a device for monitoring the synchronism of one or more flaps of aircraft wings, wherein the device includes a control cable which is connected with the flaps such that the control cable follows the flap movement. In accordance with the invention, the path of installation of the control cable extends from a first point to a second point, one or both of which are arranged on non-movable structural components of the aircraft wing.

Pivoting coupling system of a large dihedral empennage to the tail fuselage of an aircraft, in which the empennage (1) comprises a right lateral box (la) and a left lateral box (1b) arranged at a large dihedral angle in the tail fuselage (6) of the aircraft, characterised in that it comprises a horizontal central box (1c) joining the lateral boxes (1a, 1b), and which comprises a rear spar (11); and means of linkage (2, 3, 4) for horizontally linking the central box (1c) to a frame structure (7, 8, 9) of the tail fuselage (6) of the aircraft, which allow the empennage (1) to rotate vertically about a horizontal linkage shaft (5) between a negative maximum angle of incidence (−θ1) and a positive maximum angle of incidence (+θ2) in response to the actuation of actuator means which are connected to the empennage (1) and to a structural element of the fuselage of the aircraft.

The present invention discloses an airplane active side lever system. The airplane active side lever system comprises a monitoring module and a side level module. The monitoring module communicates the side level module through a serial port communication mode, and the format and the length of a communication data packet are constant. The monitoring module is able to obtain the real-time state of data of a control stick and set the work modes of the side level module, wherein the work modes include a follow-up mode, a balancing mode, an active mode and a passive mode. The balancing mode, the active mode and the passive mode can be switched by inputting instructions through the monitoring module or be switched by buttons on a side lever handle, the balancing mode and the active mode are switched by a balancing switch at the top of the side level handle, and the passive mode and the active mode are switched by a switching switch at the side surface of the side lever handle; and the monitoring module sends instructions to the follow-up mode which is switched in the active mode and the balancing mode.

A method and apparatus are disclosed for disabling on-board pilot operation of an aircraft and transferring aircraft operation to an alternate source of control. The aircraft has at least one manually actuated control device for controlling at least one mechanical actuator, with the actuator causing movement of an aircraft attitude control surface or an aircraft engine throttle. The control device is mechanically connected to the actuator(s). The alternate source of control may be one or more of an autopilot, a flight control system and an off-aircraft human pilot. The method and apparatus for disabling on-board pilot operation provides for (a) receiving a signal indicative of an emergency condition requiring the disabling of on-board pilot control of the aircraft; (b) disconnecting the one or more control devices from their respective actuator(s) in response to the receipt of the emergency condition signal; and (c) connecting the actuator(s) to the alternate source of control.

Described is an interconnection system, in particular for flaps which are located side-by-side, on an aircraft wing, for example landing flaps or high-lift flaps. The interconnection system is in particular characterised in that it comprises a correspondingly guided compensating profile for a gap between the flaps, which gap changes when the flaps are moved. Preferably, the interconnection system is combined with a failsafe device for adjacent flaps in the form of a longitudinally adjustable interconnection strut which is connected to the adjacent flaps by way of corresponding load introduction fittings. The interconnection system is in particular suitable for large aircraft in which the gap between adjacent flaps, due to the relatively large lateral movement of the flaps during retraction and extension, becomes relatively wide, and the forces acting on the flaps are relatively great.

Autopilot control of an aircraft is accomplished using a servo wherein the input shaft and output shaft are collinear. In one embodiment, the output shaft is supported, in part, by the input shaft. In further embodiments an engage clutch mechanism is provided to allow decoupling of the motor from the output shaft, and a slip clutch mechanism is provided to limit the torque output of the servo to enhance the safety of operation. The collinear shaft arrangement enables the engage clutch mechanism and slip clutch mechanism to synergistically utilize components for multiple functions.

The present invention provides a drive system which can automatically deploy a ground spoiler even when a wheel is locked. A drive system for a ground spoiler deploys closed ground spoilers 4A to 4D when a wheel speed VW exceeds a wheel reference speed VW1 or when an air speed VB exceeds an air reference speed VB1 after a main gear 6 touches down. The system may also close the deployed ground spoilers 4A to 4D when the wheel speed VW is smaller than a wheel reference speed VW2 or when the air speed VB is smaller than an air reference speed VB2.

A device (22) fitted with ribs is placed on a cylindrical rod (20) so that it is fixed to it in bending. The device (22) is formed of at least three ribs connected to each other at their ends, and at least one pair of elements in the form of a star connecting the ribs together, at least close to its central region. This device enables the rod (20) to apply an approximately radial force to the center of the rod through an actuator onto webs (24) bearing on its ends, minimizing the mass for a given spacing of the webs (24). The device is particularly applicable to the control of aircraft control surfaces.

The invention discloses a long-service-life electric steering engine which comprises a potentiometer, a motor, a bevel gear pair driven by the motor, and a speed reducer driven by the bevel gear pair, wherein the speed reducer comprises a first small-teeth-difference internal gear pair consisting of a first external gear and a first internal gear, a second small-teeth-difference internal gear pair consisting of a second external gear and a second internal gear, and an eccentric shaft driven by the bevel gear pair, wherein the first external gear and the second external gear are fixedly connected to the eccentric shaft to form two stages of planet transmission, the second internal gear is fixedly connected to a housing of the long-service life electric steering engine, the first internal gear is used for driving a load to rotate, and the transmission ratio of the first small-teeth-difference internal gear pair is more than that of the second small-teeth-difference internal gear pair. The long-service life electric steering engine disclosed by the invention realizes the long service life and high reliability based on a small-teeth-difference planet gear speed reducer, and meets the requirement of an intermediate and long-range aircraft on the long service life. The invention also discloses an aircraft employing the long-service life electric steering engine.

A bearing outer shell encases a self-lubricating polymeric bearing pad material concentrically therein. A helical spring applies a force through from the outer shell, through the bearing pad material, and onto a shaft when the bearing is installed on a shaft. An elastomeric base material may be interposed between the outer shell and the bearing pad to provide additional resiliency. The shaft may be sized to create an interference fit, causing a radial preload that can be designed to exceed the expected external vibration or shock forces. Proper selection of the elastomeric base material affords damping of these environments. The bearing provides the necessary guidance with minimal loss in power consumption. Therefore, not only is the useful life of the shaft and housing extended, but the life of the associated gears, bearings, and components that require proper alignment to minimize wear the load concentrations is further extended.

The invention discloses an unmanned aerial vehicle, a detection method of flight states thereof and a wearable device. In the detection method, a propeller operation state collector is arranged at least one support arm of the unmanned aerial vehicle, and is used for acquiring operation state signals of a propeller. The method comprises steps of acquiring the operation state signals, acquired by the propeller operation state collector, of the propeller; processing the operation state signals to obtain operation states of the propeller; and according to the operation states of the propeller, determining the flight states of the unmanned aerial vehicle. In this way, by setting the propeller operation state collectors on the support arms of the unmanned aerial vehicle to acquire the operation state signals of the propeller, and calculating the flight states of the unmanned aerial vehicle according to the operation state signals, better flight control can be performed on the unmanned aerial vehicle by use of the detected flight states of the unmanned aerial vehicle, expected flight paths can be acquired and controllability and safety in the unmanned aerial vehicle flight process can be improved.

The invention discloses a vertical axis propeller eccentric circle control mechanism. The vertical axis propeller eccentric circle control mechanism is characterized in that a vertical axis propeller rotating shaft is installed in the center of a base; a chassis, an upper end plate and a lower end plate are sleeved respectively with the vertical axis propeller rotating shaft; the upper end plate is connected with the base; the low end plate is connected with the chassis; the upper end plate and the lower end plate are connected through a first connection plate and a second connection plate to form a second revolute joint composed of a part of the upper end plate and a part of the first connection plate, a third revolute joint composed of a part of the lower end plate and a part of the first connection plate, a fourth revolute joint composed of a part of the lower end plate and a part of the second connection plate, and a fifth revolute joint composed of a part of the upper end plate and a part of the second connection plate; and the upper end plate, the lower end plate, the first connection plate, the second connection plate, the second revolute joint, the third revolute joint, the forth revolute joint and the fifth revolute joint compose a parallelogram mechanism. The vertical axis propeller eccentric circle control mechanism can control the position of the parallelogram mechanism through a control rocker arm of the upper end plate and a control rocker arm of the first connection plate to drive orderly the lower end plate and the chassis so that relative positions of the vertical axis propeller rotating shaft and the chassis are changed thus a blade pitch angle can be controlled.