530results about "Aircraft control" patented technology

A wing for an aircraft, including a main wing and a slat assembly including a slat and at least one connection assembly movably connecting the slat to the main wing, such that the slat is movable between a retracted position and at least one extended position, such that the connection assembly has an elongate slat track, and the front end of the slat track is mounted to the slat, such that the rear end and / or the intermediate portion of the slat track are mounted to the main wing by a roller or slide bearing such that the slat track is movable along the track longitudinal axis, such that the connection assembly has a drive unit configured for driving the slat track, such that the drive unit includes a rotary actuator having at least one drive arm rotatably driven about a rotation axis and drivingly engaging the slat track.

An aircraft has a boom, a propulsion assembly coupled to a first end of the boom, and a first wing coupled to a second end of the boom. The propulsion assembly is coupled to the boom by a rotating joint. A second wing is optionally coupled to the rotating joint. The first wing is coupled to the boom by a rotating joint. The first wing is coupled to the rotating joint by a hinge. A vehicle with roll, pitch, and yaw maneuverability able to mirror the aircraft movements may be coupled to the second end of the boom. The vehicle body may be picked up with a vehicle chassis disconnected from the vehicle body. The boom houses an energy source to power the propulsion assembly. A rudder is coupled to the second end of the boom. A paddle is disposed between the propulsion assembly and the boom.

The invention discloses a vertical take-off and landing aircraft with improved linked wing tilting ducted fan layout, and belongs to the technical field of pneumatic and power systems of vertical take-off and landing aircrafts. The aircraft comprises an improved joined wing structure layout and a tilting ducted fan system integrally designed with the improved joined wing structure layout. The improved joined wing structure layout comprises a front wing, a rear wing, a vertical connecting wing surface, a vertical stabilizer and a fuselage. The front wing is divided into a front wing inner side wing surface and a front wing outer side wing surface by a vertical connecting wing surface, and adopts a sweepback wing belt dihedral structure; two ends of the rear wing are respectively arranged above the front wing through a vertical connecting wing surface, the middle part of the rear wing is arranged above the tail part of the fuselage through a vertical stabilizer, and the rear wing adopts a forward swept wing structure; the tilting ducted fan system comprises a front wing integrated ducted fan flap 7, a rear wing integrated ducted fan lifting control surface 9 and a tail propelling ducted fan 16. The core problems that pitching of a traditional connected wing is out of control, the weight of a tilting structure is large, and the energy efficiency is low are solved.

An aircraft wing includes a fixed structure and a linear hydraulic slat actuator attached to the fixed structure. The linear hydraulic slat actuator is oriented spanwise along the aircraft wing. The aircraft wing also includes an actuator output link and a slat attached to the actuator output link. The aircraft wing also includes a bell-crank coupled to the linear hydraulic slat actuator and to the actuator output link. In response to spanwise movement of the linear hydraulic slat actuator, the bell-crank is configured to move the actuator output link in a different direction that enables the slat to move between a retracted position and an extended position.

The invention relates to the technical field of aircraft control, in particular to a control-surface-free aircraft control system based on synthetic jet and a momentum wheel, which comprises a lift force and attitude control subsystem, a steering control subsystem, a propulsion subsystem, a cooperative control unit and an attitude and navigation sensor. The lift force and attitude control subsystem independently controls different areas to generate local lift force difference through a synthetic jet generator matrix integrated on the surface, and rolling and pitching control is achieved. And the steering control subsystem changes the angular momentum direction of a flywheel through a control moment gyroscope arranged in the steering control subsystem, and generates reaction moment to control yawing. The propulsion subsystem provides a primary forward thrust. The cooperative control unit receives a sensor signal, generates an instruction according to a preset strategy, coordinates the subsystems to work cooperatively, and achieves whole-airspace flight control without an exposed control surface.

A convertiplane (1) is described, which comprises a fuselage (2) extending along a first axis (A), a pair of half-wings (3) projecting cantilevered from respective sides of the fuselage (2), a pair of nacelles (4) housing respective engines and operatively connected to the half-wings (3); and a pair of rotors (5) associated with the half-wings (3), rotatable around respective second axes (B) and tiltable so as to arrange the convertiplane (1) in an airplane configuration or in a helicopter configuration. Each half-wing (3) further comprises a wingbox (20) and an appendage (21) articulated on the wingbox (20) and selectively movable between a relative first neutral operating position taken when the convertiplane (1) is in the airplane configuration and in which it defines an extension of the wingbox (20) and a relative second position taken when the convertiplane (1) is in the helicopter configuration and in which it defines with the wingbox (20) a compartment (50) through which the air flow generated by the respective rotor (5) can pass. The convertiplane (1) comprises a closing system (60) configured to elastically preload each appendage (21) towards the respective first neutral operating position.

An aircraft wing assembly includes a wing box including an upper surface, a lower surface, and a leading edge structure connecting the upper surface to the lower surface and secured to the upper surface at a connection point. The aircraft wing assembly includes a drive mechanism secured to the wing box and a high lift device mounted to the drive mechanism. The drive mechanism is configured to move the high lift device between a retracted position and an extended position in use, and wherein the high lift device is configured to overlap the connection point when in the retracted position.

Systems, devices, and methods including: an assembly configured to maintain a wing on an aircraft, the assembly comprising: a connector configured to removably connect to a first portion of the wing; and an actuator having a mechanism configured to removably engage a second portion of the wing, the second portion of the wing is configured to be moveable; wherein the actuator is configured to control the second portion of the wing.

A wing for an aircraft, comprising a main wing and a high lift assembly which includes a body, and an assembly connecting the body to the main wing, so the body is movable between a retracted position and an extended position. The assembly comprises a track extending along a longitudinal axis between first and second ends, with an intermediate portion therebetween. A roller bearing ensures that the track moves along the longitudinal axis. The roller bearing comprises bearing units, each mounted to the main wing and having a bearing surface for engaging the elongate track. The units include a first and second end units and an intermediate unit wherein, relative to the longitudinal axis, the intermediate unit is positioned between the first end unit and the second end unit. The elongate track comprises a recessed portion which, in the retracted position, aligns the intermediate unit.

The invention discloses a dismounting and mounting device for an output rudder of an aircraft, and aims to solve the problems of low dismounting and mounting efficiency of the output rudder and high equipment production and maintenance cost caused by complex dismounting and mounting structure and tedious dismounting and mounting steps of an existing device. The dismounting and mounting device comprises a connecting assembly, a locking assembly and a limiting assembly. The steering engine is connected with a rudder shaft of the output rudder through a connecting assembly, and the connecting assembly is configured to be capable of being driven by the steering engine and driving the rudder shaft to rotate; the locking assembly is arranged on the connecting assembly, and the locking assembly is configured to be capable of achieving locking of the connecting assembly so that the rudder shaft can be connected with the steering engine; the limiting assembly is matched with the locking assembly, and the limiting assembly is configured to be capable of controlling the locking state and the unlocking state of the locking assembly and limiting the axial displacement of the output rudder; the locking assembly can be switched from the locking state to the unlocking state under the action of external force, at the moment, the rudder shaft is allowed to be inserted into or pulled out of the connecting assembly, and when the external force is removed, the locking assembly recovers to the locking state so that the rudder shaft can be connected with the steering engine.

A compliant mount for an aircraft, for attaching equipment (such as a latching unit, for example) to a structural member of the aircraft, the compliant mount comprising: a mounting member for attachment to the structural member of the aircraft; a receptacle for supporting the equipment, the receptacle being coupled to the mounting member along a longitudinal axis; and one or more spring elements arranged to resiliently bias the receptacle into alignment with the longitudinal axis and to allow angular movement of the receptacle away from the longitudinal axis. Also provided is an aircraft equipment assembly comprising such a compliant mount, and aircraft equipment supported by the receptacle, and an aircraft having one or more such equipment assemblies. The aircraft may be a folding wingtip aircraft and the equipment assembly may be part of a folding wingtip actuation mechanism.

The invention discloses a multi-layer combined unmanned aerial vehicle and a control method, and relates to the field of high-altitude long-endurance unmanned aerial vehicles, the multi-layer combined unmanned aerial vehicle comprises a plurality of sub-aircrafts, each sub-aircraft comprises a fuselage, an empennage, wings, wingtips and protrusions, the wingtips are arranged at the ends, away from the fuselage, of the left and right wings, and the wingtips can be in butt joint with the wingtips of the other aircraft; protrusions are arranged on the upper surface and the lower surface of each wing, the end faces of the protrusions can be in butt joint with the end faces of the protrusions of another airplane, and the protrusions can be used for containing an aircraft undercarriage at the same time. Through aggregation and recombination of a plurality of small unmanned aerial vehicles, a multi-layer combined unmanned aerial vehicle with a high aspect ratio is formed, and the combined unmanned aerial vehicle can effectively overcome the defect that the rigidity of wings of a traditional unmanned aerial vehicle with a high aspect ratio is insufficient and can carry an oversized radar conformal antenna with long wave bands such as meter waves and the like. And long-time air-hang detection of high-value targets such as stealth aircrafts and the like is realized.

The utility model discloses a control surface control mechanism of a V-tail unmanned aerial vehicle, which comprises a pair of control surface main bodies which are arranged in a V shape, and the included angle between the two control surface main bodies is 90 degrees; the main shaft is vertically arranged on the end faces of the ends, close to each other, of the control surface main bodies; and the control assembly is used for controlling the pair of main shafts to rotate synchronously. The utility model has the following advantages and effects: after being assembled, the control mechanism is completely hidden in the fuselage of the aircraft, so that the flight resistance is not additionally increased; differential motion and homodromous motion of the two control surface main bodies are respectively driven by the same motor, so that the deflection consistency of the control surface main bodies on the two sides can be ensured.

A harmonic gear system, having: a housing; an input shaft having an aft end within the housing that defines a wave generator; a wave generator bearing; first and second nested cup-shaped flex splines that are grounded to the housing, the first and second flex splines are biased by the wave generator via the wave generator bearing; an output shaft having a forward end that is within the housing, and a radial flange at the forward end; first and second nested circular splined shafts within the housing, the first and second circular splined shafts are connected to the radial flange of the output shaft and are driven by ones of the first and second flex splines, to thereby drive the output shaft; and a bearing between the first and second circular splined shafts.

An avionics system can provide data to first receiving components associated with a proprietary first protocol and second receiving components associated with a non-proprietary second protocol. A configuration tool generates a first configuration file including instructions for a data access component to provide a first data output to the first receiving components and a second data output to the second receiving components. The configuration tool can generate a second configuration file including instructions to translate the second data output. The data access component can receive a definition database defining the first protocol to translate data according to the non-proprietary second protocol. The data access component can provide requested data to the first and second receiving components according to the non-proprietary second protocol. The second receiving components are configured to translate data received based on the second configuration file.

This actuation system (22) comprises: an actuating device (26) comprising a first end (30) configured to be connected to a main section (18) and a second end (32), which is designed to be connected to a control surface (20), and at least one hinge device (28) comprising at least one hinge element (34) and designed to guide a movement of the control surface (20) relative to the main section (18), wherein the actuating device (26) is configured to move its first (30) and second end (32) translationally along an actuator axis (A-A'), wherein the hinge device (28) defines a plurality of hinge axes (H-H') which intersect to define a pivot axis (P-P') wherein the pivot axis (P-P') is offset from the hinge device (28) wherein the hinge device (28) is configured to guide the movement of the control surface (20) in a rotation about the pivot axis (P-P').

The wide-speed-range tailless variable-structure aircraft comprises an aircraft body, the aircraft body comprises an aircraft body-fixed wing fusion body and variable-structure wing rudders, the variable-structure wing rudders are symmetrically arranged on the two sides of the aircraft body-fixed wing fusion body, and a wing body fusion body layout is integrally formed; the fuselage-fixed wing fusion body is connected with the variable-structure wing rudder through a driving mechanism, and the variable-structure wing rudder can deflect in the axial direction of the fuselage; the tail of the fuselage-fixed wing fusion body is provided with a trailing edge rudder, the trailing edge rudder of the fuselage-fixed wing fusion body achieves pitching channel control through same-direction deflection, and under the condition that the channel coupling coefficient is smaller than 1.5, rolling channel control is achieved through differential deflection. The problem that the yaw channel operation stability characteristic is insufficient under the tailless layout condition is solved, and the wide-speed-range operation performance of the aircraft is effectively improved; compared with other variable-structure aircrafts in which additional variables are introduced, the combination of the variable-structure wing rudder and the trailing edge rudder is simple and efficient, and the burden on a control system is small.

An aircraft wing with an array of moveably flight control surfaces, each flight control surface having a trailing edge with a moveable tab attached to the flight control surface trailing edge. A flight control system coupled to a flight control surface drive system which moves the flight control surfaces; a tab drive system which moves the moveable tabs and one or more aircraft angle of attack sensors. The flight control system stores a set of angular deflections to deflect the tabs upwardly when the angle of attack reaches a threshold.

The invention provides a rudder for a supersonic aircraft and a manufacturing method, and relates to the field of supersonic aircraft design, the rudder is integrally formed through co-curing, the rudder comprises a rudder body, the rudder body comprises a supporting plate, the supporting plate is formed by co-curing a carbon fiber composite material, and the supporting plate is made of a carbon fiber composite material; the supporting plate is provided with at least three hollow supporting pipes, and the cross section of each supporting pipe is in a regular hexagon shape. The supporting inner core is arranged in the hollow position of the interior of the supporting pipe, and the cross section of the supporting inner core is in a regular hexagon shape; the supporting outer layer covers the surface of the supporting plate; a heat protection layer; the supporting inner core and the supporting outer layer are made of foam materials. According to the rudder for the supersonic aircraft and the manufacturing method, the overall structure of the rudder is co-cured and integrally formed, the supporting plate made of the carbon fiber composite material is adopted, the stability of the structure of the rudder is guaranteed, meanwhile, the overall mass of the rudder is effectively reduced, and meanwhile mass production is facilitated.

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.

The invention belongs to the technical field of aerospace, and particularly relates to a deflected and shrunk flexible dorsal fin suitable for heading control of a tailless aircraft and the aircraft, and the deflected and shrunk flexible dorsal fin comprises a middle separation beam, a flexible skin, an air bag and an air pressure control device; wherein the middle partition beam is of a rigid supporting structure and extends to the front end of the dorsal fin to form a partition, and the dorsal fin is divided into a left independent structure and a right independent structure; the flexible skin covers the two sides of the middle separation beam to form a flexible deformable surface; the air bags are arranged on the inner side of the skin and connected with the air pressure control device, the air pressure control device fills the air bag on one side with air, the air bag on the other side sucks air, and the asymmetric state of the wing face of the dorsal fin is formed. The heading control requirement of the tailless aircraft is met, and meanwhile the body heading stability of the tailless aircraft is improved.

Disclosed are a wing for an aircraft, a leading edge high lift assembly and an aircraft, the wing comprising a main wing and a leading edge high lift assembly, the leading edge high lift assembly comprising a slat and a connecting assembly movably connecting the slat to the main wing, and a connecting assembly for connecting the slat to the main wing such that the slat is movable between a retracted position, a first extended position, and a second extended position, where the connecting assembly comprises an elongate slat track, a first end and / or a middle portion of which is mounted to the slat, the slat track being mounted to the main wing by means of a roller support. The aim of providing a wing with a beneficial ratio of aerodynamic performance to noise resulting from the extension of the slat is achieved, in which the slat track and the roller support are configured such that, in a retracted position and a first extended position, the slat rests against the main wing and, in a second extended position, the slat rests against the main wing. A gap is formed between the trailing edge of the slat and the leading edge of the main wing.

A rotorcraft having a main rotor with at least one main rotor actuator; a series actuator arranged upstream of the main rotor actuator; engines for powering the main rotor; and a flight control system for controlling flight operation in an AEO mode or in an AFR mode, comprising: a control logic for pre-conditioning the series actuator upon entering the AFR mode for enabling rotorcraft control upon occurrence of an IFSD, wherein pre-conditioning of the series actuator comprises shifting a reduced nominal stroke range authorized for the series actuator stroke of the series actuator in the AEO mode into a reduced shifted stroke range authorized for the series actuator stroke of the series actuator in the AFR mode to enable authorization of an increased stroke range upon occurrence of an IFSD to provide sufficient control authority for a recovery maneuver.

In order to simplify assembly, the invention proposes a forward fairing 8 for an aircraft is provided. The forward fairing 8 comprises a fairing shell 11 and an attachment device 10. The attachment device 10 comprises a protrusion 14 that is integrally formed with the fairing shell 11 and a spigot 16 that is coupled to the protrusion 14.

An aircraft assembly includes an aircraft structure, a flexible panel having a supported end attached to the aircraft structure and configured to bend about the supported end relative to the aircraft structure. A rigid stopper extends from and fixedly attached to the aircraft structure and is configured to restrict bending of the flexible panel about the supported end in a first direction and permit bending of the flexible panel in an opposing second direction. The flexible panel has a first deformed configuration wherein the flexible panel is bent towards the rigid stopper in the first direction and abuts the rigid stopper and the rigid stopper restricts bending of the flexible panel in the first direction. The flexible panel has a second deformed configuration wherein the flexible panel is bent away from the rigid stopper in the second direction and spaced from the rigid stopper in the second direction.

Human-Machine Interface This human-machine interface comprises: - a female part (58) having a cylindrical face (154) opposite a cylindrical face (170) of a male part (56) only between a first plane (164) and a second plane (166), - two cylindrical channels (156, 158) formed only in one of the male and female parts and a groove (174) cut into the cylindrical face of the other male and female parts. The groove (174) has a locking section (174) extending in a plane perpendicular to an insertion axis (10) such that a hardenable material filling both the groove and the channels forms a hardenable pin that locks the male part in translation inside the female part independently of its adhesion force on the male and female parts. The locking section is located away from the first and second planes (164, 166). Fig. 6

A roller arrangement for a moveable control surface of an aircraft component includes a roller, a fitting, and an adapter sleeve. The roller component has a shaft and a wheel coupled to a roller end of the shaft, wherein the wheel is configured to rotate about a major longitudinal axis of the shaft. The fitting is couplable to the moveable control surface, and it includes a support opening defined therein. The support opening is shaped, sized, and configured to accommodate passage of the wheel therethrough. The adapter sleeve includes an interior passageway defined therein, wherein the interior passageway is shaped, sized, and configured to removably receive and couple with the shaft of the roller component. The adapter sleeve is removably secured to the fitting, to inhibit lateral movement of the adapter sleeve relative to the fitting.

The invention relates to a training system (1) for combating naval surface drones, this training system comprising: a remote operation system (2), a telecommunications system (3), one or more naval surface drones (5) which are configured to be operated simultaneously by the remote operation system (2), via the telecommunications system (3), to simulate an attack mission against one or more target ships (50), the number of naval surface drones (5) being in particular between 1 and 15, in particular between 3 and 10. Figure for the abstract: Fig. 1

An electrical flight control system for controlling an aircraft, the flight control system comprising at least one subassembly for acquiring a piloting command in order to acquire the piloting command from an operation of an associated piloting member that can be operated by a pilot of the aircraft along at least one piloting axis, the flight control system comprising at least three position sensors per piloting axis generating position information of the piloting member and being in communication with the at least one acquisition subassembly. Each acquisition subassembly comprises at least three simplex computers each having a single calculation channel connected to one of the at least three position sensors.

A pulley frame assembly includes a pulley frame and rollers adjustablely mounted on the pulley frame. The pulley frame assembly further includes a roller adjustment device comprising: a transition frame rotatably supporting the rollers and movably connected to the pulley frame via a rotating connector; an adjustment member including an eccentric shaft passing through the transition frame and the pulley frame, the eccentric shaft including a first shaft section and a second shaft section eccentrically arranged, wherein the first shaft section is connected to the transition frame, and the second shaft section is connected to the pulley frame; and a rotating connector passing through the transition frame and the pulley frame at a distance spaced relative to the eccentric shaft, allowing the transition frame to rotate about the rotating connector. This device reduces the adverse effects of roller stress on the overall lifespan of the device, allows for quick confirmation of whether the rollers are in the correct position, and enables adjustment of the roller position over a wide range.