190 results about "Aileron" patented technology

The invention belongs to the field of vertical take-off and landing aircrafts, and relates to a distributed contra-rotating rotor hybrid tilting rotorcraft. Comprising a front tilting electric contra-rotating rotor nacelle, a front cabin body, a wing end tilting electric contra-rotating rotor nacelle, a V-shaped empennage, a V-shaped empennage aileron, a vertical empennage, a retractable undercarriage, a green hybrid power energy system and an empennage tilting electric contra-rotating rotor nacelle, the front tilting electric contra-rotating rotor nacelle is hinged with the other end of the front cabin body; an empennage tilting electric contra-rotating rotor nacelle is arranged at the end part of the V-shaped empennage; wing end tilting electric contra-rotating rotor wing nacelles are arranged at the ends of the wings; and rotor wing units are arranged in the rotor wing nacelles, the six rotor wing units all adopt a coaxial reverse upper and lower two-paddle structure, and the rotating directions of the rotor wings in the six rotor wing units are that the adjacent rotating directions and the opposite rotating directions are opposite.

The utility model discloses a reversible control surface of an unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicles. The reversible control surface of the unmanned aerial vehicle comprises an unmanned aerial vehicle body, the number of the fixed control surfaces is two, and one end of each fixed control surface is fixedly connected to the outer surface of one side of the unmanned aerial vehicle body; and the number of the disassembling mechanisms is four, and each disassembling mechanism comprises a mounting box, a rotating component, an adjusting component, a moving component and two pushing components. According to the turnover control surface of the unmanned aerial vehicle, the positions of the two moving plates are adjusted through cooperation of the adjusting component and the moving component, the fixing blocks are separated from the fixing grooves through arrangement of the pushing component, and the movable control surface is taken down from the mounting box and replaced with a new movable control surface; the fixing grooves in the movable control surfaces are connected to the mounting box in a butt joint mode, the movable control surfaces are fixed through the pushing parts, and under the condition that the aileron and the empennage are not replaced, the maintenance cost and the maintenance time are reduced by replacing the single movable control surface.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

The invention belongs to the technical field of unmanned aerial vehicles, and particularly relates to a tailstock type overwater vertical take-off and landing fixed-wing unmanned aerial vehicle which can float and park on the water surface, can vertically take off to leave the water and is switched into a fixed-wing mode for cruising and level flight. The unmanned aerial vehicle comprises a four-rotor support buoy and a wing body fusion body fixed to the upper portion of the four-rotor support buoy, small wing type buoys are arranged on wing tips of the two sides of the wing body fusion body respectively, and lifting ailerons capable of deflecting up and down are arranged on the rear edges of wings on the two sides of the wing body fusion body respectively. And four motor rotor propeller propelling devices are uniformly arranged at the tip part of the four-rotor bracket buoy.

The invention discloses a fixed wing type cross-domain aircraft capable of realizing self-adaptive load replacement, belongs to the technical field of cross-domain aircrafts, and is used for a water-air double-domain operation cross-domain aircraft. Comprising an aircraft body, a main wing, an aileron, a tail assembly, a vertical take-off and landing support, a vertical take-off and landing rotor wing, a fixed wing propeller, an underwater propeller, a load replacement device and a control module. The control module comprises a decision making system, a telemetering communication system, a state sensing system, an underwater water surface load replacement system, a water surface four-rotor vertical take-off and landing control system, an air fixed-wing cruise control system and an underwater vehicle control system. According to the fixed wing type cross-domain aircraft, through collaborative innovation of a multi-mode power system and self-adaptive buoyancy control, the dynamic characteristic adaptation problem in cross-domain navigation is solved, and the fixed wing type cross-domain aircraft has the vertical take-off and landing capacity, the air cruising capacity and the underwater maneuvering capacity.

A control system for rotating an airfoil along one axis that is offset from a perpendicular axis that varies the angle of attack (AOA), dihedral, and center of lift based on the position of the attachment to precisely maneuver an aircraft. This technique can eliminate the need for traditional control surfaces such as ailerons, elevators, and rudders. allowing for an aerodynamically advantaged shape.

Provided are a method and a system for cooperative airborne meteorological perception with multiple aircraft. The method includes: determining a training set based on a spatial position, motion data, weather, an observation image, and static information of each of multiple aircraft; determining a fusion performance function based on an objective of a cooperative perception network; generating a fusion model based on the training set and the fusion performance function; generating a perception result image based on the fusion model; determining flight risk information based on the perception result image; determining a route alarm command and sending the route alarm command to a management user interface and an interactive device of a target aircraft; after receiving the route alarm command, generating, by the management user interface, a route adjustment parameter; in response to obtaining the route adjustment parameter, performing at least one of the following operations: adjusting an engine velocity, adjusting an aileron pressure, adjusting an elevator deflection, and adjusting a rudder deflection.

The invention discloses a tilt rotor aircraft depending on aerodynamic force, and relates to the technical field of aircrafts. The four-axis unmanned aerial vehicle comprises a fuselage, wing assemblies, four-axis unmanned aerial vehicle modules and empennage assemblies. The wing assembly comprises a left wing, a right wing and a wing cross beam which are all provided with ailerons; the four-axis unmanned aerial vehicle module is I-shaped, is hinged to a wing cross beam through a bearing, and can tilt bidirectionally; the empennage assembly comprises a tail support, a horizontal electric variable pitch propeller, a horizontal empennage with a control surface and a vertical empennage. The four-axis unmanned aerial vehicle module serves as a tilting unit, and the structure is simplified; attitude control is realized by virtue of aerodynamic difference, and the efficiency is improved; the tail horizontal electric variable-pitch propeller realizes rapid pitching adjustment; vertical take-off and landing, low-altitude maneuvering and high-speed cruising capabilities are integrated, the application scene is wide, and safe forced landing can be achieved when faults occur.

[Problem] To provide a flying object, a control method, and a program capable of controlling to achieve an increase in the cruising distance of the flying object and the stability of the flying object during hovering and landing by providing rotatable auxiliary wings to the flying object and controlling the angle of attack. [Solution] Provided is a flying object comprising: a main body part; a plurality of rotary wing parts provided to the main body part; an auxiliary wing part provided to the main body part; and a control part for controlling the operation of the rotary wing parts and the auxiliary wing part. The control part executes a change control of the lift of the auxiliary wing part in accordance with a comparison result between an acquired thrust-related parameter and one or more reference parameters.

The invention relates to the technical field of aircrafts, and discloses a triphibian aircraft and a triphibian form switching method thereof. Comprising an aircraft main body, a left auxiliary body and a right auxiliary body, retractable wings are arranged between the aircraft main body and the left auxiliary body and the right auxiliary body, sliding wheel assemblies are arranged at the bottoms of the three bodies respectively, foldable rotor wing mechanisms are arranged at the tops of the left auxiliary body and the right auxiliary body, a propeller is arranged at the tail of the aircraft main body, and foldable small wings are arranged on the outer sides of the left auxiliary body and the right auxiliary body. During take-off, the wings and the winglets are unfolded, the rotor mechanisms are started, and rising is achieved; during air cruise, the sliding wheel assemblies and the rotor mechanisms are stored, the propellers are started, and smooth flight is ensured; when the unmanned aerial vehicle runs on land, the sliding wheel assemblies are opened, the rotor mechanisms are stored, the wings and the winglets are folded, wind resistance is reduced, and the propellers are started to run; when sailing on the water, the three-body structure ensures buoyancy, and the propeller is pushed to move forwards; through the three-body structure, seamless switching of air, land and water triphibian scenes is achieved, the system adapts to multiple environments, the use flexibility is improved, and the application range is widened.

Various control surfaces are disclosed that can cause a side force along a lateral axis running through a center of mass of a craft when the control surfaces are deployed in a same direction. The side force can be created without creating a yawing moment on the craft. The control surfaces can be on a winglet of the craft. Also disclosed are a split-flap control surface and stand-alone ailerons.

The utility model discloses an integrated aileron driving system, which belongs to the field of aircraft control systems and comprises a fixed wing, an aileron is arranged on one side of the fixed wing, the aileron is rotatably connected with the fixed wing through a first linkage shaft, the upper surface of the aileron is rigidly connected with a first speed reducing plate, and a second speed reducing plate is arranged on the other side of the fixed wing. A power rotating assembly is arranged on the upper surface of the fixed wing, the power rotating assembly is connected with the upper surface of the first air brake and used for driving the first air brake to move, fixed limiting pieces are symmetrically arranged at the two ends of the surface of the side, close to the aileron, of the fixed wing, and a second air brake is arranged between the fixed limiting pieces. The second air brake is located on the lower surface of the aileron, and the first air brake and the second air brake are connected through a transmission assembly. According to the integrated aileron driving system, through the linkage design of the aileron and the air brake, the structure is simplified, pneumatic coordination control is achieved, and the control performance at different flight speeds is improved.

The present invention provides an extreme STOL airplane comprising: a fuselage and a wing having port and starboard wing sections. Each wing section has a body with top and bottom surfaces, leading and trailing edges, and a span, i.e., the distance between a wing section tip and root connected to the fuselage, and flap coves. Aerodynamic elements integrated into the wing sections include Fowler flaps, having leading and trailing edges and a span, flap tracks, wherein the flap tracks are external to the wing section body, extend aftward beyond the trailing edge of the wing section, and are configured to enable the Fowler flaps to rotate or deflect and to translate or extend and retract, Frise ailerons, wherein the Frise ailerons are located outboard of the Fowler flaps, and spoilerons located over the leading edge of the Fowler flaps when the flaps are in the fully extended position.

The present application relates to the field of aircraft technology, and discloses a vertical take-off and landing aircraft, which comprises a wing-body integrated lifting body fuselage, a front wing and a rear wing, and a stabilizing tail, and is composed of a two-degree-of-freedom propulsion Y3 configuration formed by the front wing, the rear wing and the fuselage, and is composed of two fixed lifting propellers installed below the front wing and in the front part of the gravity center and a two-degree-of-freedom tilting thruster installed at the tail to solve the problem that the existing Y3 configuration cannot simultaneously consider the low-speed hovering mode and the cruising mode; the tandem-lifting body aerodynamic configuration is combined with the new Y3 power configuration to solve the problem that the aircraft is difficult to balance the contradiction between high aerodynamic efficiency and large cargo capacity; and finally, the side aerodynamic optimization of the fuselage, the up-down staggered design of the front and rear wing tip ailerons, and the ring wing 21 design are adopted to solve the problem that the existing configuration is prone to generate large aerodynamic resistance.

The present invention provides a method and device for inhibiting the trim mechanism of a fly-by-wire flight control system. The method comprises: determining the current operating state of the flight control system; determining whether the cockpit control system is in a non-disengaged state based on a redundancy signal output by a displacement sensor; determining whether the lateral trim position and the longitudinal trim position are at zero; the flight control module issuing a trim inhibition enable signal to cut off power to the lateral and longitudinal trim mechanisms, thereby successfully activating trim inhibition; and after trim inhibition is successfully activated, toggling the aileron trim switch and the elevator trim switch disables the corresponding lateral and longitudinal trim mechanisms. The present invention is intended to address the problem of the normal function of a fly-by-wire flight control system being affected by erroneous operation of the cockpit manual trim system.

The application provides a four-rudder control coaxial rotor system, which comprises a rotor shaft, an upper rotor and a lower rotor installed on the rotor shaft, a gear box arranged on the rotor shaft, the gear box being located below the lower rotor, a first control mechanism for controlling the movement of the upper rotor to make the aircraft produce pitching movement and a second control mechanism for controlling the movement of the lower rotor to make the aircraft produce aileron movement, the first control mechanism being located between the gear box and the aircraft, the second control mechanism being located between the gear box and the lower rotor, and the first control mechanism comprising an upper tilt disc, a first rudder set and an upper variable-pitch pull rod. The application has the functions of using two sets of rudders to make the aircraft achieve elevator, forward flight and left-right steering, the rotor control functions are distinct and do not interfere with each other, the rotor control efficiency is improved, the structural complexity is reduced, the overall reliability of the device is improved, and the device is easy to maintain.

The application provides a kind of aircraft servo elastic dynamics model correction method considering operating surface gap, first aircraft structure is divided into substructure, and the aircraft dynamics model containing aileron gap nonlinearity is established according to substructure branch modal;And additional inertia force caused by aircraft operating aileron deflection is introduced in dynamics model, and the aircraft open-loop servo elastic motion equation containing aileron gap nonlinearity under independent generalized coordinate is obtained;According to sensor output equation, rudder transfer function and flight control system equation, the aircraft nonlinear servo elastic closed-loop system model containing aileron gap nonlinearity is established, and the model is corrected based on ground vibration test, rudder frequency response test and gap measurement test result, and the stability of the model is judged based on simulation analysis result and open-loop frequency response test result and is further corrected, which improves the accuracy of gap nonlinearity aircraft aerodynamic servo elastic characteristic evaluation, and is helpful to guide aircraft servo elastic ground test engineering practice.

An aircraft featuring at least one lifting rotor, one or more lifting airfoil-shaped wings, and at least one thrust-generating device capable of spinning the wings or the aircraft with the wings around its centre axis, characterised by the aircraft being able to change its orientation during flight by varying the aerodynamic lift of at least one of the airfoil-shaped wings by a device or a combination of devices including but not limited to, aileron or wing tip control surfaces, orwing pitch modulation, or wing profile modulation, or by modulating the lift produced by the lifting rotors around the aircraft centre of rotation, or a combination of said devices, and with a controller for the aerodynamic lift modulating device.

The application provides a swept-back three-beam thin wing composite wing structure, which comprises a wing main body and an aileron installed in a wing main body mounting groove, the wing main body comprises a wing skeleton and a wing outer skin structure outside the wing skeleton; the wing skeleton comprises a longitudinal member unit and a transverse member unit, the longitudinal member unit comprises a front beam, a middle beam and a rear beam, the front beam and the middle beam and the middle beam and the rear beam are arranged at acute angles, the transverse member unit comprises a plurality of wing ribs arranged between the front beam and the middle beam, between the middle beam and the rear beam and on one side of the front beam towards the front of the fuselage; the front beam, the middle beam and the rear beam are all provided with a spar connecting head for connecting with the fuselage. The wing provided by the application adopts a three-beam structure, each wing beam has a mouth-shaped cross section, so that each wing beam forms a multi-wall form, can bear the bending moment and the torque brought by a large wing effective load, and the wing structure is light in weight, high in strength and good in integrity.

Wheels are arranged on the lower portion of a ground carrier in a bilateral symmetry mode, a power system and a communication and automatic driving control system are arranged in a vehicle head on the front portion of the ground carrier, and a connecting clamping groove is formed in a floor of the ground carrier. Electromagnetic spring buckles are arranged on the lower portion of the flight cabin and correspond to the four connecting clamping grooves formed in the ground carrier floor in position. Wings with upward inverted inner sections and straight outer sections are respectively arranged on the left and right sides of the upper part of the flight cabin; the inner section and the outer section of each wing are averagely divided into two sections; when the hydraulic mechanism is opened and closed, the inner section and the outer section can be folded or unfolded around the hinge connecting mechanism, and after folding, the spring positioning pins are inserted into the positioning holes. Oblique empennages are arranged on the left side and the right side of the tail portion of the upper portion of the flight cabin respectively, and deflection rudders are arranged on the rear edges of the outer ends of the oblique empennages. The hovercar can run on the ground, can take off and land vertically, can hover in the air or fly horizontally and can sail underwater; the safety and reliability are high, and the use efficiency is high.

The systems, methods, and other embodiments described herein relate to controlling the elevons of an aircraft to stabilize the fuselage of a free-wing aircraft. In one embodiment, the system includes a processor and a memory storing machine-readable instructions. When an instruction is executed by the processor, the processor 1) measures the fuselage pitch attribute of the aircraft, 2) measures the wing pitch angle of the wing assembly, and 3) controls the angles of the elevons of the wing assembly based on the fuselage pitch attribute and the wing pitch angle, so that the fuselage and wing assembly of the aircraft rotate independently of each other and freely around the pitch axis of the aircraft.

The application discloses a wing, rib and skin integrated forming auxiliary wing mold and a forming control method thereof, and belongs to the technical field of unmanned aerial vehicle composite auxiliary wings, and comprises an upper mold and a lower mold, characterized in that a forming surface of the upper mold is matched with a cavity surface of the lower mold, and a three-dimensional closed auxiliary wing cavity for overall forming of the auxiliary wing is formed after the mold is closed; the auxiliary wing cavity comprises a main beam cavity platform extending along the length direction of the auxiliary wing, a plurality of rib cavity platforms arranged at intervals with the main beam cavity platform and arranged in a vertical direction, and a continuous skin transition curved surface connecting the main beam cavity platform and the rib cavity platform; and a plurality of glue injection ports and exhaust ports in communication with the auxiliary wing cavity are arranged on the lower mold. The application can complete overall manufacturing of the auxiliary wing in one forming process, significantly reduces assembly links, and the formed product has a continuous fiber reinforced overall structure, no mechanical connection holes, uniform stress distribution, and significantly improved structural strength and service life.

This aims to create an attitude control system that can respond appropriately to pilot inputs during maneuvers such as turning and landing. [Solution] An attitude control device installed in a radio-controlled model airplane, which is a controlled object that is operated based on control signals received from an external device, includes a calculation unit that selectively outputs a first aileron control signal input as a control signal and a second aileron control signal calculated to maintain the horizontal position of the controlled object based on the roll angle of the controlled object. The calculation unit selects and outputs the second aileron control signal when the function is instructed to be ON by a function setting signal received from an external source, the first aileron control signal is a signal indicating neutral, and the roll angle of the controlled object is within a set angle.

This invention discloses a control system and launch method for an airborne folding-wing unmanned aerial vehicle (UAV). Before the launch of the sub-unit, the mother aircraft sends a voltage signal to the sub-unit, which is continuously monitored by the sub-unit's flight control system. Upon reaching the launch point, the mother aircraft is powered off, the unlocking device is activated, and the flight control system initiates control after the voltage monitoring disappears for a period of time. At this moment, the V-tail controls pitch and participates in roll control. After a certain period, the flight control system unlocks the locking servo, causing the wings to begin unfolding until they are fully extended. At this time, the ailerons participate in control, entering the unpowered dive-pull control phase. After a certain period, the wing control is locked. Subsequently, the flight control system determines that the attitude and airspeed are stable. Once stable, the engine is started, and the engine speed is determined. If the start fails, the start operation is repeated three times. After the engine starts, the flight path is executed normally, the ground station launch process is exited, and normal flight control logic is entered. This invention enables full-process control of an airborne folding-wing UAV.

The utility model provides an aircraft airfoil skin repair training rack. The aircraft airfoil skin repair training rack comprises a mounting rack, an aileron skin assembly structure, a wing skin assembly structure and an empennage skin assembly structure. An operation table is arranged on the top of the mounting frame. The aileron upper skin wraps the upper surfaces of the aileron framework and the flap accessory, the aileron lower skin wraps the lower surfaces of the aileron framework and the flap accessory, the wing skin wraps the upper and lower surfaces of the wing framework through the outer end of the wing framework, and the empennage skin wraps the upper and lower surfaces of the empennage framework through the outer end of the empennage framework. The device is complete in parts, rich in structure, convenient to disassemble and assemble, and convenient for teaching and daily airplane airfoil skin repair practical training operation of trainees.

The present invention belongs to the field of mechanical connection and bonding, and specifically relates to a device and method for translating the rotation axis of an aircraft aileron. The device comprises: an I-shaped material (1), three joint locator supports (3), a new joint locator (4-1) at a first joint, an old joint locator (4-2) at a first joint, a new joint locator (10-1) at a second joint, an old joint locator (10-2) at a second joint, a new joint locator (11) at a third joint, and a movable latch (5). During operation, the device is positioned by utilizing the mutual positional relationship between the axis translation joints, and with the assistance of the positioning function of an auxiliary positioning body, the device is accurately positioned, thereby performing replacement of the new and old joints and realizing translation of the aileron axis.

The application discloses a device for inhibiting the rotating movement of the aileron of an aircraft wing, and belongs to the field of aircraft structure design. The device can inhibit the rotating movement of the aileron of the aircraft wing. The device comprises a support frame, a limiting block, a limiting rod, a driving rod, a control motor and a support arm. The support frame and the support arm are installed on the rear wall of the wing through the support arm. When the device is started, the limiting block contains the limiting rod; when the device is restored, the limiting block is separated from the limiting rod. When the motor is started, the limiting block moves to the limiting rod along the axis of the rotating shaft. In the process that the limiting block moves to contain the limiting rod, the rotating angle of the limiting rod (the aileron) is gradually reduced until the rotation is stopped. The aileron flutter of the aircraft is caused by the coupling of the rotating mode of the aileron and other modes of the structure, so that the rotating movement of the aileron is inhibited, and the rotating flutter of the aileron of the aircraft is inhibited. The application can be used in the aircraft flutter test to inhibit the aileron flutter of the wing.

This invention belongs to the field of aircraft manufacturing technology, and provides an integral molding die for an aileron, a molding method, and an aileron. The die includes a lower die, a core die, side dies, and an upper die. The lower die includes a lower die pressing part and a lower fixing part; the lower fixing part is located on both sides of the lower die pressing part and is a protruding structure relative to the lower fixing part. The upper die includes an upper die pressing part and an upper fixing part; the upper fixing part is located on both sides of the upper die pressing part and is a protruding structure relative to the upper die pressing part. The core die is installed at the lower die pressing part and corresponds to the mounting shaft of the aileron. The side die is installed at the lower fixing part. After the lower die, upper die, and side die are closed, the inner surface of the side die is in contact with the ends of the lower die pressing part and the upper die pressing part, and the outer surface is in contact with the inner surface of the upper fixing part, forming a cavity structure corresponding to the shape of the aileron. This invention has the advantages of low cost, simple process, short time, and lighter aileron weight in aileron manufacturing.

An attitude control device mounted on a radio-controlled model airplane, which is a controlled object controlled based on a control signal received from an external device, comprises a calculation part configured to selectively output a first aileron control signal, which is input as the control signal, and a second aileron control signal, which is calculated based on a roll angle of the controlled object to maintain the controlled object in a horizontal state. The calculation part selects and outputs the second aileron control signal when the following conditions are satisfied: turning on of a function is instructed by a function setting signal received from the external device; the first aileron control signal is a signal indicating neutral; and the roll angle of the controlled object is within a set angle.

------ Vector-Tripping Aircraft The present invention relates to a vector-tripping aircraft (1), comprising: a wing in the form of a wing (2); a tail in the form of at least one rear aileron (3) carried by said wing (2); and an engine assembly (4) connected to said wing (2) for thrusting the aircraft (1), characterized in that said engine assembly (4) is disposed in front of said wing (2) and is connected to it by at least one arm (6), and that it comprises at least one turbine (10) configured to create, in operation, a jet of air discharged towards said wing (2) by a steerable nozzle (11). Figure to be published with the abbreviation: Figure 1