110results about "Wing adjustments" patented technology

The invention discloses a folding wing structure of an unmanned aerial vehicle, which comprises a rotating shaft embedded in a vehicle body, wherein a lower wing elastic sheet and an upper wing elastic sheet are mounted in an inner cavity of the rotating shaft from bottom to top, a first locking block and a second locking block are respectively arranged at the free ends of the lower wing elastic sheet and the upper wing elastic sheet, the first locking block and the second locking block are respectively matched with a first notch and a second notch which are formed in the rotating shaft, the outer side of the rotating shaft is sequentially sleeved with a lower wing and an upper wing from bottom to top, the upper surface of the lower wing and the lower surface of the upper wing are each provided with a groove, the two grooves are oppositely formed to form a torsional spring groove, a torsional spring is arranged in the torsional spring groove, and the upper inner sides of the lower wingand the upper wing are provided with a first locking groove and a second locking groove correspondingly; and the outer side of the rotating shaft is further sleeved with a roller bearing, a baffle isinstalled at the top of the rotating shaft, and the roller bearing is located between the upper wing and the baffle. The processing cost of the unmanned aerial vehicle is reduced, the folding wing structure is insensitive to the structural size, and the folding wing structure can adapt to a more miniaturized barrel-shooting unmanned aerial vehicle.

A wing assembly for an aircraft including a fixed wing and a wing tip device rotatable about a joint assembly. The wing assembly is configurable between a flight configuration for use during flight and a ground configuration for use during ground-based operations. The wing assembly includes joint-bordering structure located adjacent the joint assembly, the joint-bordering structure having a rib, an upper cover and a lower cover. The joint assembly includes a moveable panel. The panel includes a first hard stop element, and the joint-bordering structure includes a second hard stop element. The first hard stop element is integrally formed in the panel, and the second hard stop element is integrally formed in one of the rib, the upper cover or the lower cover of the joint-bordering structure. The hard stop elements may be configured to self-align.

A hinge arrangement for an aircraft wing is provided, for providing a folding wingtip. The hinge arrangement comprises a first mount for securing in a first wing section and a second mount for securing to a second relatively foldable wing section, each having a plurality of hinge elements arranged along a hinge axis. The arrangement comprises a latching arrangement for securing the mounts against folding movement, the latch arrangement having a latch axis being spaced from the hinge axis in a substantially spanwise direction for example so as to extend into the wing box. This can provide an efficient arrangement which may be contained within the outer mould line of the wing which can reduce the need for a fairing.

A wing assembly 10 for an aircraft, comprising a fixed wing 18 and a wing tip device 12 rotatable about a joint assembly 16. The wing assembly is configurable between a flight configuration for use during flight and a ground configuration for use during ground-based operations. The wing assembly 10 comprises joint-bordering structure located adjacent the joint assembly 16, the joint-bordering structure comprising a rib 34, an upper cover 26 and a lower cover. The joint assembly comprises a moveable panel 32. The panel 32 comprises a first hard stop element 38a, and the joint-bordering structure comprises a second hard stop element 40a. The first hard stop element 38a is integrally formed in the panel 32, and the second hard stop element 40a is integrally formed in one of the rib 34, the upper cover 26 or the lower cover of the joint-bordering structure. The hard stop elements 38a, 40a may be configured to self-align.

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.

The invention relates to an aerodynamic profile body with an outer flow surface for flow through a fluid, wherein the aerodynamic profile body comprises:a first surface layer and a second surface layer, which form at least part of the outer flow surface and are designed to be flexible in an end section of the aerodynamic profile body,a connection arrangement that flexibly connects the first surface layer to the second surface layer of the profile body in the end section, andan actuator device with at least one actuator provided in an interior of the aerodynamic profile body which interacts with at least one of the surface layers in the flexible end section in such a way that the cross-section of the aerodynamic profile body in the flexible end section can be changed by the at least one actuator,characterized in thatthe outer flow surface of at least one of the surface layers in the flexible end section has, at least in sections, a surface profile that extends parallel to the direction of flow.

According to at least one example a launch system is provided, including a carrier vehicle and at least one payload vehicle. The payload vehicle includes a payload propulsion system and aerodynamic lift surfaces designed for providing the payload vehicle with aerodynamic powered flight at a design subsonic cruise speed at a desired altitude. The carrier vehicle is configured for carrying the at least one payload vehicle at least up to the desired altitude, and further includes a solid rocket propulsion system for propelling the launch system to the desired altitude. The carrier vehicle is configured for providing a predetermined forward speed at the desired altitude, correlated to the design subsonic cruise speed. The carrier vehicle is configured for releasing the payload vehicle with respect to the carrier vehicle at the desired altitude and the predetermined forward speed. The design subsonic cruise speed is less than 0.7 Mach number, and, the desired altitude is greater than 3km.

The application discloses a folding wing unfolding device and relates to the technical field of flight equipment. The device comprises a fixed shaft, a steel ball assembly fixedly connected to the fixed shaft, a lower piston rotating assembly and an upper piston rotating assembly rotatably connected to the side wall of the fixed shaft, the upper piston rotating assembly and the lower piston rotating assembly are in plug-in cooperation and form a working cavity, the steel ball assembly is located in the working cavity, and the bottom end of the lower piston rotating assembly is connected to an ignition driving assembly. The upper piston rotating assembly comprises an upper piston, the upper piston is provided with an upper spiral slide rail, the lower piston rotating assembly comprises a lower piston, the lower piston is provided with a lower spiral slide rail, the rotation directions of the upper spiral slide rail and the lower spiral slide rail are opposite, the upper piston and the lower piston are both sleeved on the fixed shaft, and the upper piston and the lower piston are both movably connected to the steel ball assembly. The folding wing unfolding device of the application has the advantages of reducing impact overload without affecting the storage life and being more practical.

A vertical take-off and landing aircraft (1) comprises a fuselage (2), at least one propulsion system (5), at least four lift propellers (10), and at least two fixed wing planes (20, 30, 40), the fixed wing planes (30, 40) located rearward of the forwardmost lift propellers (10) being located above the lift propellers (10). The lift propellers (10) are distributed on either side of the fixed wing planes (30) and on either side of the fuselage (2) such that the lift propellers (10) are separated in the longitudinal direction by at least the size of the chord (Cl, C2) of the fixed wing planes located between the lift propellers (10) and in the lateral direction by at least the width (L1, L2) of the fuselage (2) located between the lift propellers (10). At least one of the fixed wing planes (30) has a variable wingspan and the aircraft (1) comprises a control system for varying the wingspan in flight.

This invention relates to an integrated variable camber wing based on a piezoelectric-driven torsional cubic superstructure, comprising a wing skeleton and a skin. The wing skeleton includes multiple torsional cubic superstructures arrayed along the wing chord direction. Each torsional cubic superstructure is a cube formed by an array of multiple unit-cell torsional superstructures. Adjacent unit-cell torsional superstructures in the wing length and thickness directions share a straight rod, while adjacent unit-cell torsional superstructures in the wing width direction share four Z-shaped piezoelectric rods. Each unit-cell torsional superstructure is a cube formed by an array of four cell units, with adjacent cell units sharing a straight rod. Each cell unit includes Z-shaped piezoelectric rods and straight rods. Two Z-shaped piezoelectric rods are connected to the two ends of two straight rods. Each Z-shaped piezoelectric rod at both ends of the torsional cubic superstructure is connected to the skin. When all the Z-shaped piezoelectric rods of the unit-cell torsional superstructure undergo compressive deformation in the height direction, the unit-cell torsional superstructure undergoes torsional deformation, thus causing the torsional cubic superstructure to undergo torsional deformation. By converting the compressive deformation of the cellular unit into the torsional deformation of the compressive-torsional cubic superstructure, the wing camber is changed, and the drive components are eliminated, thus realizing the lightweight design of the morphing wing.

A fixed-wing aircraft (10) includes: a fuselage (14), a closed wing (16), a first actuator device (22), and a first connector device (24). The closed wing (16) is coupled to the fuselage (14) and has a wingtip (18) and a first deformable wing portion (20). The first actuator device (22) is arranged to supply a first input torque to the first connector device (24) when the first actuator device (22) is actuated, and the first connector device (24) is arranged to apply a first output torque to the first deformable wing portion (20) when the first input torque is supplied to the connector device. The first deformable wing portion (20) is arranged to elastically deform when the first output torque is applied to the first deformable wing portion (20). The closed wing (16) forms an outer support (26) at the wingtip (18), which is arranged to counteract the first output torque.

The invention relates to an aircraft wing and to an aircraft comprising such an aircraft wing. The aircraft wing comprises a wing structure supporting a wing cover having an outer aerodynamic surface forming part of an airfoil profile of the wing and a recess stepped back from the wing airfoil profile, and a spoiler device comprising a spoiler coupled to the wing structure by a spoiler actuation mechanism for moving the spoiler between a retracted position and a deployed position, the spoiler having a surface following the outer profile of the wing airfoil profile and the spoiler being fully accommodated in the recess when the spoiler is in the retracted position, and wherein the spoiler surface extends away from the wing cover when the spoiler is in the deployed position.

The application discloses an aircraft wing adjusting mechanism and adjusting method, and relates to the technical field of aircrafts. The aircraft wing adjusting mechanism comprises a guide part, a moving part, a first elastic temperature control assembly and a locking part. The guide part is fixedly arranged at both ends in the body. The moving part is fixedly connected with the moving end of the wing and can move along the guide part. The first elastic temperature control assembly is arranged between the guide part and the moving part. The first elastic temperature control assembly is deformed after being powered on, can drive the moving part to move in the first direction, and can drive the moving part to move in the second direction under the action of the elastic restoring force after being powered off. The locking part is arranged on one side of the guide part. The contraction and expansion of the wing can be realized through the power-on and power-off of the first elastic temperature control assembly and the cooperation of the locking part. The design requirements of the lightweight and compact design of the morphing aircraft can be met. Moreover, the transmission stage number is small, and the control mode of power-on and power-off has higher reliability and economy.

The application relates to the technical field of aircraft design, and discloses a continuous variable-camber rigid wing driven by multiple rudders and a control method thereof, wherein the continuous variable-camber rigid wing comprises a wing leading edge section, multiple rigid wing sections, a distributed rudder driving mechanism and a transmission mechanism; the multiple rigid wing sections comprise a first wing section, a second wing section and a third wing section which are sequentially hinged through rotating shafts; the distributed rudder driving mechanism is arranged in the wing leading edge section and is used for independently driving the deflection of each rigid wing section; and the transmission mechanism is connected between the distributed rudder driving mechanism and the corresponding rigid wing section and is used for converting the output motion of the distributed rudder driving mechanism into the deflection motion of the corresponding rigid wing section. According to the application, the rigid wing is segmented, multiple rudders are used for independent driving, an advanced aerodynamic analysis model and a closed-loop feedback control are combined, the wing camber is self-adaptively, continuously and accurately regulated within a full flight envelope, and thus the optimal aerodynamic performance can be maintained at all times.

To provide an improved wing assembly comprising a fixed wing and a wingtip device that can rotate around a joint assembly at the tip of the fixed wing. [Solution] The main wing assembly 10 for an aircraft can be configured between a flight configuration for use in flight and a ground configuration for use during ground-based operations. The main wing assembly 10 includes a joint boundary structure positioned adjacent to a joint assembly 16, the joint boundary structure comprising ribs, an upper cover, and a lower cover. The joint assembly 16 includes a movable panel 32. The movable panel 32 includes a first hard stop element, and the joint boundary structure includes a second hard stop element. The first hard stop element is integrally formed with the panel 32, and the second hard stop element is integrally formed with one of the ribs, the upper cover, or the lower cover of the joint boundary structure. The hard stop elements may be configured to self-align.

A kinematics assembly for a high lift body, in particular a droop nose or leading-edge-slat, configured to be mounted to a main body of a wing of an aircraft, a corresponding support station for supporting the high lift body, and a high lift body arrangement includes an actuation organ providing a drive mounting point configured to mount the kinematics assembly to a drive organ for moving the high lift body with respect to the main body between a stowed position and a deployed position, and a deflection organ configured to be connected to the actuation organ and providing a central hinge point for hinging the high lift body at the main body movably between the stowed position and the deployed position; wherein, at least in the stowed position, the drive mounting point is located aft of the central hinge point with respect to a travel direction of the aircraft.

An aircraft capable of transonic flight is equipped with a slotted airfoil to significantly improve the cruise performance during flight. The aircraft includes a plurality of slotted airfoils uniquely configured to maximize natural laminar flow across the aircraft wings. In addition to reduced wing drag, the unique configuration minimizes shocks at various conditions over the course of a flight. Each slotted airfoil includes a fore element, an after element, and a slot disposed between the fore element and aft element. The aft element is maneuverable relative to the fore element. The slot is configured to promote a substantially constant air flow field onto the aft element during all flight angles of attack.

A wing hinge assembly 13 comprising a wing tip 11(fig.1) pivotally mounted to a fixed wing 5(fig.1); wherein one of the wing tip and the fixed wing comprises a pair of opposing walls 28(fig.4) that de

Stretchable shape-memory polymer composite skin with a Poisson's ratio of zero, comprising the following parts: a main body made of conductive fibers; several metal frames made of metallic aluminium material, arranged under the main body, from which the electrical power is transferred to the main body; and a multitude of reinforcing elements made of synthetic resin or plastic material, each connected at intervals to the lower part of the main body by the several metal frames, the temperature in the main body can change based on the electrical power transmitted by the multiple metal frames.

An aircraft (1, 1', 1") is described with a fuselage (2) extending along a first longitudinal axis (Y) of the aircraft (1, 1', 1"); a pair of half-wings (3, 8) projecting sideways, cantilevered, from the fuselage (2); a pair of first rotors (20a, 20b; 21a, 21b; 22a, 22b) that revolve about a second axis (B, C; D, E; F, G) and can be tilted independently of each other about a third axis (H; L; J); a pair of first motors (30a, 30b; 31a, 31b; 32a, 32b) that can be operated to rotate first rotors (20a, 20b; 21a, 21b; 22a, 22b) independently of each other; the aircraft (1, 1', 1") can be moved between: a first forward flight configuration and a second take-off / landing configuration in which the second axes (B, C; D, E; F, G) are arranged transverse to the first axis (Y); first wing areas (60a, 60b; 61a; 62a, 62b; 70a, 70b; 72a, 72b) operationally connected to respective first rotors (20a, 20b; 21a; 21b; 22a, 22b); the first wing areas (60a, 60b; 61a, 61b; 62a, 62b; 70a, 70b; 72a, 72b) being tiltable with respect to the third axis (H; L; J) as a result of the inclination of the respective first rotors (20a, 20b; 21a; 21b; 22a, 22b) about the third axis (H; L; J) independently of each other.

A lifting member motion device for moving a lifting member relative to a fixed wing includes: a first linkage group connecting the lifting member and the fixed wing, driving the lifting member to displace relative to the fixed wing; and a second linkage group simultaneously connected to the first linkage group and the lifting member, constraining the lifting member to move along a predetermined trajectory, wherein the predetermined trajectory includes one or more locking positions, and when the lifting member moves to a locking position, the trailing edge of the lifting member is in close contact with the fixed wing; when the lifting member is not in a locking position, the trailing edge of the lifting member is spaced apart from the fixed wing by a certain distance. The lifting member motion device of the present invention ensures that the lifting member is in close contact with the fixed wing when in a locking position, and moves away from the fixed wing when moving between locking positions, thereby avoiding friction between the lifting member and the fixed wing and undesirable component wear.

The invention relates to a drone system (2) for transporting a payload (4), comprising a fuselage (6) for receiving the payload (4) and a wing assembly (8) which is connected to the fuselage (6). The drone system (2) has a first aerodynamic configuration (A1) formed by at least one structural element (10). According to the invention, the drone system (2) can be transferred into a second aerodynamic configuration (A2), which differs from the first aerodynamic configuration (A1), by releasing the at least one structural element (10), in particular by ejecting same, the structural element (10) forming a first wing structure (18) in the first aerodynamic configuration (A1), in which the wing assembly (8) has a first aerodynamic profile (20), and the drone system (2) can be transferred into the second aerodynamic configuration (A2) by ejecting the first wing structure (18), the wing assembly (8) having, in the second aerodynamic configuration, a second aerodynamic profile (24) which is formed by a second wing structure (22) and differs from the first profile (20), and the first aerodynamic profile (20) at least partly enclosing the second aerodynamic profile (24) in the first aerodynamic configuration (A1). The invention also relates to a corresponding designing method (100), a corresponding production method (200), and a corresponding operating method (300).

This slat cove filling device (24) for a slat (18) of an aircraft (10) comprises a sheltering panel (46) extending along a spanwise direction (Y) and comprising a proximal edge (50), configured for being fastened to a slat body (22) of the slat (18), and a distal edge (52), configured for being slidable relative to the slat body (22), the sheltering panel (46) comprising a plurality of guiding passages, each guiding passage being configured for cooperating with a guiding element of the slat body (22) for guiding the distal edge (52) along a sliding direction (S) relative to the slat body (22).

Aircraft high-lift device (1) of a lifting surface (2). The high-lift device (1) comprising a skin comprising an upper skin (3), a lower skin (4) and a nose (5), the high-lift device (1) comprising: - a flexible skin allowing for a morphing high-lift device (1), having a rear upper edge (7.1) and a rear lower edge (7.2), at least the rear upper edge (7.1) sliding along the chord direction between a retracted position and an extended position, - a linear actuator (6) connected to the nose (5) and pivotably connected to the lifting surface (2) between the retracted position and the extended position, so that in the extended position the nose (5) is located in a downward location with respect to its location in the retracted position and at least the rear upper edge (7.1) is located in an anterior location with respect to its location in the retracted position.

An aircraft including one or more sensors configured to obtain sensor data, a wing that includes a primary portion and a wingtip, and a folding wingtip system that includes a latch pin system, a first electronic controller, and a second electronic controller. The first electronic controller configured to receive a fold command, send instructions to use the latch pin system according to the fold command, and make a first determination that a wingtip position of the wingtip is not in a folded position. The aircraft further including a flight control computer that is configured to send the fold command to the first electronic controller, in response to the first determination, set first error codes for the folding wingtip system, change a setting to indicate that the aircraft is grounded based on the first error codes, and schedule maintenance of the aircraft based on the first error codes.

An aerial mobility vehicle includes a wing extending from a fuselage of the aerial mobility vehicle; a folding portion disposed at an edge portion of the wing, configured to extend from the wing to form a portion of the wing during unfolding of the folding portion, and configured to move to coincide with the wing to reduce an air resistance area of the wing in a vertical direction during folding of the folding portion; an actuator connected to the folding portion and configured to power the folding portion to unfold or fold to the wing; and a controller connected to the actuator and configured to control the actuator to fold the folding portion during vertical takeoff and landing of the fuselage, and configured to control the actuator to unfold the folding portion during cruising of the fuselage.

This application provides a rapid wing deployment and locking mechanism for swept-wing aircraft. The mechanism includes: a base, a slider, a guide mechanism, a linkage mechanism, a rotating mechanism, and a locking mechanism. The guide mechanism, linkage mechanism, rotating mechanism, and locking mechanism are all mounted on the base. The slider is slidably connected to the guide mechanism and connected to the linkage mechanism. The end of the linkage mechanism away from the slider is connected to the rotating mechanism. The wing is fixedly connected to the rotating mechanism. The locking mechanism is located on the side of the rotating mechanism and is used to lock the rotating mechanism. When the slider is impacted by a pyrotechnic device, the slider moves along the guide mechanism, causing the linkage mechanism to swing. The linkage mechanism drives the rotating mechanism to rotate, which in turn causes the wing to deploy. After the wing is fully deployed, the locking mechanism locks into place with the rotating mechanism. This application eliminates the need for wing reassembly, reduces the need for large packaging space, and allows for rapid wing deployment from under the fuselage. Locking the deployed wing ensures stable flight.

The invention refers to an aircraft, particularly an aircraft capable of vertical take-off and landing, comprising a fuselage, and a variable lift body defining an aerofoil and being moveably attached to the fuselage, wherein the variable lift body is pivotable around a first axis extending in the wing span direction; wherein a rotary actor adapted to cause the variable lift body to pivot in relation to the fuselage arranged within the aerofoil.

The detachable unmanned aerial vehicle comprises a vehicle body; the two front wings respectively extend towards two sides at the front part of the fuselage; the two rear wings respectively extend towards two sides at the rear part of the fuselage; the two supporting rod assemblies are respectively connected to the outer ends of the corresponding front wing and the rear wing; a plurality of propellers mounted to the strut assembly; wherein the front wing and the rear wing are detachably mounted to the fuselage, and / or the strut assembly is detachably mounted to the front wing and the rear wing.