28 results about "Vertical stabilizer" patented technology

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.

A flying vehicle with folding wings that drives well on the ground and flies well in the air; controlled from inside, it changes from automotive to aircraft configuration without the operator needing to get out of the vehicle. The balance point of the vehicle is located midway between the front and back wheels, providing good handling on the road. In the aircraft configuration with wings unfolded, there is a front wing and a back wing. The incidence of the front wing is controllable, enabling the craft to rotate and take off with a center of gravity located well ahead of the rear wheels. The back wing is fitted with foldable vertical stabilizers near its wing tips. In automobile configuration the wings are folded on top of the body; the wings resemble a roof rack with large surfboards on it. Driven wheels provide motive power on the ground; ducted fans provide thrust for air travel. An autopilot system provides stability and navigation aid, particularly in bad weather or poor visibility. In the roadable configuration, filling water ballast tanks provides additional crosswind stability and crash protection. Also in a roadable configuration, the present invention may be licensed as a three-wheeled motorcycle towing a two-wheeled trailer. In its flying configuration, the invention can be licensed as a Light Sport Aircraft.

An electric vertical take-off and landing aircraft comprises a fuselage and a tilt wing. The tilt wing is tiltable relative to the fuselage in a range of tilt wing angles. Control surfaces are arranged at the tilt wings. The aircraft comprises a horizontal stabilizer with an elevator, in particular a stabilator, a vertical stabilizer with a rudder, in particular a fully movable stabilizer, at least one electric engine for powering main propellers arranged at the tilt wing and at least one software control device. The control device is configured to control at least the control surfaces , the main propellers and the tilt wing angle during hover and transition between hover and cruise flight.

This invention relates to a horizontal stabilizer load transfer device, comprising: a support fixed to the horizontal stabilizer and including a first support and a second support spaced apart from the first support; a roller frame having a first joint and an opposing second joint, wherein the first joint is pivotally fitted to the first support; a support arm including a third joint and an opposing fourth joint, wherein the third joint fits to the second support, and the fourth joint fits to the roller frame; a roller pivotally supported by the second joint of the roller frame; and a guide rail fixed to the aft fuselage or vertical stabilizer, such that the roller faces the guide rail and a gap is formed between the guide rail and the roller. This arrangement allows for flexible and convenient adjustment of the fit gap between the roller and the guide rail, enabling gap calibration without complex tooling, significantly reducing on-site assembly and debugging time, and effectively improving the overall assembly efficiency of the aircraft's horizontal stabilizer components. Furthermore, this invention also relates to a horizontal stabilizer and an aircraft.

The invention discloses an aircraft vertical fin damping device based on shape memory alloy net-shaped non-planar weaving and an integrated compression molding method of the aircraft vertical fin damping device, and relates to the technical field of aviation composite material structures, the aircraft vertical fin damping device comprises a composite material combined thin-wall plate, and the appearance of the composite material combined thin-wall plate is matched with a curved surface structure of a vertical stabilizer of a new energy aircraft; the non-planar net-shaped woven shape memory alloy steel wire rope vibration reduction unit is formed by weaving horizontal memory alloy steel wire ropes and vertical memory alloy steel wire ropes in a transverse and longitudinal staggered mode. The integral forming mould pressing system comprises an inner surface forming mould, an outer surface forming mould, a bottom plate and a steel wire rope clamping device; the epoxy resin adhesive is used for bonding and curing the vibration reduction unit and the composite material thin-wall plate, the vibration reduction device and the composite material vertical fin are integrally formed through a mold pressing process, the additional weight and installation space occupation of a traditional externally-hung device are avoided, and additional energy supply is not needed based on the passive vibration reduction characteristic of the shape memory alloy.

A VTOL aircraft includes a pod having a pod forward portion and a pod aft portion. The pod forward portion is connected to an inboard wing of the VTOL aircraft and the pod aft portion is connected to an outboard wing and a vertical stabilizer of the VTOL aircraft. The pod aft portion rotates, together with the outboard wing and the vertical stabilizer, about a centerline of the pod relative to the pod forward portion.

The application discloses a vertical stabilizer conical bolt dismounting device and a dismounting force measuring method, which comprises a main plate, a back plate, a pressure sensor assembly, a connecting plate, a ball head bolt, a guide bolt, a wrench, and an industrial computer. The main plate and the back plate are fixedly connected, and an airplane tail wing product provided with the vertical stabilizer conical bolt is fixed between the main plate and the back plate. The ball head bolt is arranged on the main plate. The connecting plate is connected to the main plate through the guide bolt on the side close to the back plate. The pressure sensor assembly is fixed on the connecting plate on the side close to the back plate. The back plate is provided with a stepped through hole. The industrial computer is connected to the pressure sensor assembly. The wrench clamps the nut end of the ball head bolt and applies a torque to the nut end. The ball head bolt rotates to drive the connecting plate to move close to the airplane tail wing product, pushes the pressure sensor assembly to tightly press the vertical stabilizer conical bolt, and then pushes the vertical stabilizer conical bolt out of the mounting hole and makes the vertical stabilizer conical bolt fall into the stepped through hole on the back plate. Meanwhile, the thrust value of the process is recorded and transmitted to the industrial computer for reading and analysis.

This utility model discloses a parachute cover and a parachute landing device. The parachute cover includes: a parachute pouch with a cavity for storing the parachute canopy; and a parachute fabric connected to the outside of the pouch. An inflation cavity is formed between the fabric and the pouch, allowing external airflow to enter. When airflow enters the inflation cavity, at least a portion of its kinetic energy is converted into upward lift force acting on the fabric, thus generating upward buoyancy for the parachute cover. In this utility model, the parachute cover, through the inflation cavity formed between the fabric and the pouch, allows external airflow to enter when the cover is deployed. The airflow flows within the inflation cavity and converts part of its kinetic energy into upward lift force acting on the fabric. This lift force causes the entire parachute cover to experience upward buoyancy, altering the trajectory and position of the parachute canopy inside, enabling it to move upward after release and smoothly pass over the vertical stabilizer or tail propeller area, avoiding collision or entanglement.

This invention discloses a wide-speed-range aircraft with a large payload bay, belonging to the field of supersonic aircraft design technology. It addresses the problem that when a large payload bay is installed on a carrier aircraft and enters supersonic cruise, the strong shock wave generated at the nose causes a significant increase in overall drag, resulting in the ineffective utilization of the resulting localized high-pressure flow. The invention includes a fuselage, structural connectors, a large payload bay, a left wing, a right wing, a horizontal stabilizer, and a vertical stabilizer. The fuselage comprises a nose, a mid-fuselage section, and a tail section connected sequentially. The large payload bay is connected to the upper side of the mid-fuselage section via structural connectors. The left wing includes a left inner wing section and a left outer wing section, and the right wing includes a right inner wing section and a right outer wing section. During high-speed cruise, the payload bay of this invention can compress the incoming airflow and reflect the shock wave from the aircraft's nose, thereby generating a localized high-pressure zone below the large payload bay. This significantly improves transport efficiency and single-load capacity, resulting in excellent supersonic aerodynamic performance.

A connection structure for the fuselage and tail of a wooden model aircraft that facilitates quick assembly and disassembly relates to the field of aircraft model component technology. The structure comprises two main parts: a truss-type tail boom and a tail fin mounting bracket. The main body of the truss-type tail boom is a wooden truss structure with a connecting box at the front and a connecting box slot at the rear, the outer wall of which is glued with carbon fiber reinforcement plates. The tail fin mounting bracket is also a wooden truss structure. The connecting box at its front is used to connect the tail boom, and pre-drilled holes on the outer side of the bracket are used to connect the horizontal stabilizer and vertical stabilizer. A servo mounting component is glued to its bottom. This invention achieves efficient assembly and disassembly through the connecting box design. After disassembly, the components can be stored separately, significantly improving portability. It also allows for quick replacement of individual faulty parts, significantly reducing maintenance costs. Furthermore, it balances overall lightweight design with structural stability, exhibits excellent re-assembly and disassembly performance, and extends the overall lifespan of the model aircraft.

The utility model discloses a fixed-wing unmanned aerial vehicle based on flexible perovskite solar cells, which adopts a wing body fusion design and comprises an equipment compartment in the middle of a vehicle body and wings on two sides, and a pair of ailerons with rolling and pitching control functions is arranged on the rear edge of the outer section of each wing; a pair of vertical stabilizers are arranged at the tips of the wings so as to enhance the course stability during flight; two flexible perovskite solar cells which accord with the shapes of the wings and are tightly attached to the wings are mounted on the surfaces of the fuselage and the wings, and power is provided for the unmanned aerial vehicle. According to the utility model, the sweepback wing flying wing type layout is adopted, so that the effective lift area of the wings and the laying area of the solar cells are increased while the structural weight is reduced; the battery is laid on the surface of the wing, damage to the battery in the laying process can be reduced, and the installation difficulty is lowered; meanwhile, the flexible perovskite solar cell has the advantages of low cost, good folding and bending performance and the like, so that the flight efficiency of the unmanned aerial vehicle is improved, the effective load is increased, and the endurance is prolonged.

This invention discloses a wide-speed-range aircraft with a forward-mounted payload bay, belonging to the field of supersonic aircraft design technology. It addresses the problems of supersonic transport aircraft employing high aspect ratio fuselage designs, which limit internal height and volume, resulting in lower overall loading performance due to the slender fuselage, and the ineffective utilization of localized high-pressure flow during supersonic cruise. The invention comprises a fuselage, left wing, right wing, horizontal stabilizer, and vertical stabilizer. The fuselage includes a payload bay, a blending transition section, a mid-fuselage section, and a tail section. The payload bay is connected to the front of the mid-fuselage section via the blending transition section, and the rear of the mid-fuselage section is connected to the tail section. During supersonic cruise, the lower surface of the payload bay acts as a compressive surface, effectively compressing the incoming flow and generating a localized high-pressure zone below the payload bay's outer envelope. This effectively improves the overall lift-to-drag characteristics of the aircraft and alleviates the nose-down moment problem caused by the forward-mounted payload bay.

An aircraft tail section comprises: - a vertical tail; - an aft fuselage section attached to the vertical tail and comprising a skin and internal stiffening members; - a horizontal tail comprising two lateral torsion boxes and a frame structure between the two lateral torsion boxes, the frame structure comprising a front spar, a rear spar and two ribs extending between the front spar and the rear spar, and each rib being adjacent a lateral torsion box. The frame structure encloses a portion of the vertical tail along a spanwise direction of the vertical tail. The aircraft tail section comprises an attachment assembly attaching the frame structure to the aft fuselage section, the attachment assembly crossing the skin and extending between the internal stiffening members of the aft fuselage section and the frame structure.

The invention discloses an aircraft vertical stabilizer repeated impact resistance testing device, which relates to the technical field of performance testing equipment and comprises a supporting mechanism, a clamp mechanism and an impact mechanism. The supporting mechanism is provided with a base and a top cross beam, the base is provided with a containing groove, and the top cross beam is located above the containing groove; the clamp mechanism is fixed in the containing groove, two opposite clamping devices are arranged on the clamp mechanism, and the clamping devices are fixed to the clamp mechanism and provided with clamping gaps for clamping the test piece; a fixed flexible block is arranged at the lower end of the gap, and a movable flexible block is arranged at the upper end of the gap; the impact mechanism comprises a base, a drop hammer plate and two guide columns, the base is fixed above the containing groove, and the guide columns are fixed to the base in parallel; the sleeve assembly of the drop hammer plate vertically slides along the guide column, the hammer head is detachably fixed on the sleeve assembly, and the base is provided with an avoiding hole for the hammer head to penetrate through the test piece; the guide columns are sleeved with buffer springs, so that the hammer head can be prevented from impacting the test piece when the drop hammer plate rebounds. Damage to a test piece can be reduced, and the test piece is prevented from being interfered by secondary impact.

This application relates to a cross-rotor unmanned helicopter and its fairing assembly. The cross-rotor unmanned helicopter includes a fuselage, and the fairing assembly includes a fairing body and a horizontal stabilizer and a vertical stabilizer separately disposed from the fairing body. The rear end and the rear bottom side of the fairing body are open structures, allowing the fairing body to cover the nose of the fuselage. The open structure allows the rear part of the fuselage to be exposed outside the fairing body. The horizontal stabilizer and the vertical stabilizer are respectively fixed to the exposed parts of the fuselage. The fairing body includes a first fairing body and a second fairing body, respectively fixed to the fuselage. The second fairing body is detachably connected to the upper rear end of the first fairing body via a connector. This fairing assembly can reduce the overall weight of the unmanned helicopter, facilitating lightweight design, while also improving the maintenance efficiency and simplifying operation.

The application provides a method for improving the docking efficiency of vertical stabilizer and rudder products, which is used for the assembly of tail wing large components made of composite materials of domestic large aircraft, realizes the rapid docking of the rudder assembly and the vertical tail assembly in the flight attitude under the rudder hanging point connection state, and belongs to the technical field of aircraft assembly. The pressure maintaining gas tank on the lifting clamp and the lifting connection mechanism ensure that the lifting joint can be connected, the design form that does not need to be connected to the gas source at all times ensures the flexibility of the lifting clamp movement, a storage battery is installed on the lifting device, and the angle of the product can be adjusted at any time by using a motor device and a wireless remote control device. The method of using the cylindrical bearing stop device improves the docking efficiency and protects the parts from damage, and the docking work can be completed by 4 people in 2 hours, and the efficiency is improved by more than 100%. It belongs to a semi-automatic docking scheme.

The present application relates to the technical field of aviation, and in particular to a supersonic civil aircraft configuration. In the supersonic civil aircraft configuration of the present application, engine nacelles are mounted above a rear fuselage between wings and horizontal stabilizers; the lower portions of inlets are shielded by the trailing edges of the wings, and the lower portions of outlets are shielded by the leading edges of the horizontal stabilizers; by means of the shielding effect of the wings and the horizontal stabilizers, shock waves generated by engines are prevented from propagating to the ground, effectively reducing the ground sonic boom loudness of such an aircraft; the engines and pylons can perform small-angle pitch deflection around pivot shafts in the pylons, thereby adjusting mounting angles of the engines and improving the propulsive efficiency of the engines under off-design operating conditions; during supersonic cruising, the tips of the horizontal stabilizers can be folded upward to provide an additional vertical stabilizer volume, so as to compensate for the loss of vertical stabilizer efficiency during supersonic cruising, and preventing expansion waves generated by engine exhaust nozzles and the horizontal stabilizers from propagating to the ground; and during low-speed takeoff / landing, the tips are unfolded to ensure a large vertical stabilizer area, so as to meet the requirements of trim in takeoff / landing configuration.

The invention discloses an unmanned aerial vehicle heat dissipation system carrying a heavy oil piston engine and an unmanned aerial vehicle, relates to the technical field of aviation aircrafts, and aims to solve the problems of resistance increase and gravity center backward movement caused by increase of the heat dissipation requirement of the heavy oil piston engine. The cooling system comprises ventral fin type water-cooling radiators, plate-fin type water-cooling radiators and plate-fin type intercooling radiators. The ventral fin type water-cooling radiators are symmetrically arranged on the two sides of an unmanned aerial vehicle rear body. The plate-fin type water-cooling radiator is connected with the pelvic fin type water-cooling radiator and is used for carrying out liquid cooling on the heavy oil piston engine; the plate-fin type intercooling radiator is used for conducting air cooling on the heavy oil piston engine, and radiating fins of the plate-fin type intercooling radiator are of a rectangular straight-through fin structure. By means of the arrangement, the requirement for high heat dissipation of the heavy oil piston engine is met, the aerodynamic resistance is remarkably reduced, the focus moves backwards, the equivalent vertical stabilizer area is increased, and therefore the aerodynamic performance and the course stability of the unmanned aerial vehicle are greatly improved.

A modular UAV comprising a fuselage, wing panels, a stabilizer, and two booms with vertical stabilizer, each of the booms equipped with an adapter. The wing panels are configured so the adapters of the boom are connectable thereto to form a lock connection. A spar passes through the fuselage, and the wing panels include holders for the spar. Ends of the spar comprise adapters formed as tabs. The lock connection is formed by connection of a turn bushing having a transverse notch to a spring-loaded fastener positioned in each of the wing panels, and of each spar adapter to a respective boom adapter that is hook-shaped, while being fixable by an external lever rigidly coupled to the turn bushing. Each wing panel includes a torque pin coupled to a corresponding fuselage hole displaced along a fuselage axis relative to the spar.

The present invention relates to the field of aircraft technology and discloses an aircraft and its arm. The aircraft arm, applicable to an aircraft, comprises an arm main body and an external extension part, which are integrally formed. The external extension part extends radially outward and forms a structure that gradually tapers along a first direction, wherein the external extension part comprises a streamlined surface structure to reduce aerodynamic drag during flight. The integral molding and connection of the arm main body and the vertical stabilizer in this disclosure can improve the stiffness of the vertical stabilizer surface, reduce the risk of loosening or of detaching of the vertical stabilizer, decrease the number of parts, and reduce weight.

A VTOL aircraft includes a pod having a pod forward portion and a pod aft portion. The pod forward portion is connected to an inboard wing of the VTOL aircraft and the pod aft portion is connected to an outboard wing and a vertical stabilizer of the VTOL aircraft. The pod aft portion rotates, together with the outboard wing and the vertical stabilizer, about a centerline of the pod relative to the pod forward portion.

The invention discloses a composite configuration aircraft which comprises a fuselage, an undercarriage arranged below the fuselage, a main rotor wing unit arranged above the fuselage and a horizontal tail arranged at the rear end of the fuselage and further comprises tilting rotor wing units arranged on the two sides of the fuselage and located below the main rotor wing unit in space. The vertical fins are longitudinally arranged in space and are positioned at the tail end of the fuselage; and the tail rotor is arranged at the tail end of the fuselage and is spatially vertical to the vertical fin. Compared with the prior art, on the basis of an existing helicopter structure, the tilting rotor wing unit is additionally arranged to serve as an auxiliary power source, and therefore the using load of the helicopter can be effectively improved when the helicopter vertically takes off and lands and flies forwards at a low speed; when the helicopter flies forwards at a high speed, the forward flying tension can be effectively improved, so that the maximum use load of a conventional helicopter can be increased, and the maximum forward flying speed of the conventional helicopter is improved.

To enable more accurate identification of the subject being photographed. [Solution] A video signal is acquired from a camera-equipped tripod head device, and the aircraft and its vertical stabilizer are detected from the acquired video signal. Then, the distance between the aircraft and the tripod head device is calculated based on the actual dimensions of the detected vertical stabilizer, and it is determined whether or not that distance is within the effective range of the shooting distance. If it is determined that the distance between the aircraft and the tripod head device is within the effective range of the shooting distance, a control signal is sent to the tripod head device to track the aircraft.

The utility model discloses a kind of unmanned aerial vehicle for realizing vertical take-off and attitude control using small turbojet engine, small turbojet engine is installed at wing tip of wing, small turbojet engine is installed at the back of fuselage (or in front of fuselage) two sides;Small turbojet engine is made to produce upward thrust by vertical rotation 90 ° using turbine worm drive mechanism or motor with self-locking holding torque, and unmanned aerial vehicle vertical take-off is realized;Unmanned aerial vehicle has aileron installed at the trailing edge of left and right wing tips, and the differential deflection of aileron makes unmanned aerial vehicle tilt rotation around longitudinal axis, and the direction of thrust of small turbojet engine is deflected using turbine worm drive mechanism, to realize the flight state stabilization and flight attitude control of unmanned aerial vehicle, so that unmanned aerial vehicle has good maneuverability;Horizontal stabilizer and elevator are omitted, vertical stabilizer and rudder are omitted, so as to reduce flight resistance, which is conducive to improving the flight speed and effective load of unmanned aerial vehicle.

The present invention relates to an aircraft (1). More specifically, the present invention relates to an aircraft (1) having a main body (2), that conserves fuel by utilizing wind surfing. The aircraft comprises a nose section (3) that tapers towards the front end of the main body (2), and a tail section (4) that tapers towards the rear end of the main body (2). It also includes a vertical stabilizer (5) positioned centrally in the upper part, close to the tail section (4), two horizontal stabilizers (6) located on the lateral sides of the main body (2) near the nose section (3), at least two flaps (7) situated on the tail section (4), and at least two lateral engines (8) located on the lateral sides of the main body (2) that can rotate to provide thrust in different directions. Additionally, there is a main engine (9) located on the tail section (4) for forward thrust and at least two air trapping structures (10) that extend from the lower part of the nose section (3) to the beginning of the tail section (4), protruding towards the lower part of the main body (2). The surface area of the lower inclined section (3b) at the lower part of the nose section (3) is greater than the surface area of the upper inclined section (3a) at the upper part of the nose section (3) to allow air to be gathered towards the air trapping structures (10) when the aircraft is in motion. The present invention also relates to an aircraft (1) in which the nose section (3) is adapted to direct all incoming air towards the air trapping structures (10).

The utility model discloses a kind of unmanned aerial vehicle for realizing vertical take-off and attitude control using rotor steering, there are two across installation in left and right wings in flight direction on hanger, hanger front end and rear end are equipped with rotors and power device that can be steered;When horizontal flight, front rotor generates pull, rear rotor generates thrust;Front rotor and power device vertically upward rotate 90 ° and generate upward pull, rear rotor and power device downward rotate 90 ° and generate upward thrust, realize unmanned aerial vehicle vertical take-off;Unmanned aerial vehicle flight attitude is controlled using deflection combination of front two rotors and rear two rotors, and using deflection combination of same side hanger front and rear rotors;Make unmanned aerial vehicle have good maneuverability, save horizontal stabilizer and elevator, save vertical stabilizer and rudder, thereby reduce flight resistance, it is favorable to improve the effective load of unmanned aerial vehicle.

This invention discloses a full-size carbon fiber composite material demonstration prototype and its manufacturing method. The demonstration prototype includes a fuselage, a wing structure, and a landing gear. The fuselage includes a fuselage frame and fuselage skin covering the fuselage frame. The wing structure includes a wing, a horizontal stabilizer, a vertical stabilizer, a canard, and a tiltrotor. The fuselage frame comprises several beams, frames, and ribs of different dimensions. The wing structure includes wing skin and several beams and ribs of different dimensions disposed within the wing skin. The beams, frames, and ribs of the fuselage frame and wing structure are all obtained by cutting and processing standardized carbon fiber foam sandwich composite material using a large flatbed machine.

The application discloses a kind of adaptive fusion type emergency glide parachute for rotor manned aircraft, including fuselage fusion type umbrella cabin, airfoil glide parachute body, multistage trigger opening parachute mechanism, gravity center adaptive adjustment system, attitude stabilizing assembly and emergency power auxiliary module, fuselage fusion type umbrella cabin inside is equipped with throwable cabin cover and umbrella bag storage cavity;Glide parachute body is stored in umbrella bag storage cavity, and airfoil parachute cloth structure is used;Gravity center adaptive adjustment system includes parachute rope length adjustment mechanism, weight detection unit and system controller;Multistage trigger opening parachute mechanism uses multistage trigger mechanism of automatic+manual+independent backup, attitude stabilizing assembly includes vertical stabilizer and horizontal stabilizer, and emergency power auxiliary module is symmetrically installed on both sides of fuselage, and is linked with multistage trigger opening parachute mechanism.The application solves the problems of low opening parachute reliability, easy to entangle with rotor, poor glide performance and unable to adapt to the change of passenger gravity center of existing pure rotor aircraft emergency parachute.