2104 results about "Airframe" 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.

An undercarriage, a middle large rotor and a middle large motor are connected to the lower portion of a fuselage body and sequentially connected to the top of a middle small tower, and the right front portion, the left front portion, the left rear portion and the right rear portion of the fuselage body are correspondingly connected with a right front fuselage arm, a left front fuselage arm, a left rear fuselage arm, a right rear fuselage arm and a power assembly. The connecting line of the rotating centers of the rotors of the four power assemblies is square or rectangular, the rotating directions of the rotors of the power assemblies at the same opposite angles are the same, the rotating directions of the rotors of the power assemblies at different opposite angles are opposite, and the rotating faces of the two rotors opposite to the middle large rotor in rotating direction incline around the axes of the respective arms. The included angle between the rotating surfaces of the two rotors and the horizontal plane is an acute angle beta, the rotating surfaces of the two rotors with the same rotating direction as the middle large rotor incline around respective arm axes, the included angle between the rotating surfaces of the two rotors and the horizontal plane is an acute angle theta, beta is larger than theta, the reaction torque and the horizontal component torque of the rotors jointly drive the aircraft to steer, and the aircraft has the advantages of being long in endurance time, large in load and high in wind resistance. And a universal flight platform is formed.

The invention relates to the technical field of unmanned aerial vehicle control, in particular to an unmanned aerial vehicle dynamic obstacle avoidance method based on multi-sensor fusion. Comprising the steps of receiving an image sequence, a depth point cloud set, a pose state parameter and a flight speed parameter; calculating an illumination distortion gradient value and a speckle noise entropy value based on the image sequence and the depth point cloud set; performing coordinate registration and feature extraction on the image sequence, the depth point cloud set and the pose state parameter, and outputting a visual feature set and a radar feature set; generating a weighted visual feature set and a weighted radar feature set through a weight adjustment function, and generating a joint heterogeneous feature tensor through alignment splicing; and generating a body attitude and thrust control instruction by using the combined heterogeneous feature tensor through a probability prediction model and a model prediction control algorithm device. According to the method, through cross-domain multiplexing and deep coupling of the body physical parameters in the data stream, error accumulation and decision delay caused by multi-stage series calculation are avoided, and global performability of an obstacle avoidance task is facilitated.

The invention discloses a deformable wing with the characteristics of large unfolding-folding ratio and high bearing capacity. The deformable wing comprises a bearing structure, a driving mechanism, a wing surface dimensional structure and an accessory structure, the bearing structure comprises a plurality of groups of telescopic mechanisms, wing ribs and corner boxes; the accessory structure comprises a mounting plate, and the mounting plate is connected with the machine body; each group of telescopic mechanisms comprises a plurality of stages of telescopic rods which are arranged along the chord length direction and are fixed on the mounting plate; the wing ribs are arranged in the wingspan direction and are connected with the multi-stage telescopic rods through the angle boxes to form a wing framework; the driving mechanism penetrates through the hollow parts of the wing ribs, and two ends of the driving mechanism are respectively connected with the mounting plate and the wing rib on the outermost side of the wing; the wing skeleton is externally coated with the wing surface dimensional structure to form a wing shape.

A wing fuselage joint for an aircraft having joining members which include a first configuration and a second configuration. The joining members of the first configuration include a platform and bracket configured to pivot in a lateral plane and translate in a longitudinal direction. The joining members of the second configuration include a platform and bracket configured to pivot in a longitudinal plane and translate in a lateral direction. The joining members of either configuration join the fuselage structure and the spar of an aircraft together.

The invention relates to a foldable tilting four-rotor aircraft concept scheme design method based on task requirements. According to the method, task demand indexes such as voyage, cruising speed, task load, cruising height and hovering height are used as input, and the total design weight, power device parameters, rotor wing geometric parameters and fuselage and wing parameters of the aircraft are determined in sequence; and performance indexes such as hovering lift limit, level flight speed, climbing performance, practical lift limit, voyage and endurance in a helicopter mode and a fixed wing mode are calculated and analyzed respectively. And performing matching judgment on a performance calculation result and a task requirement, and performing iterative adjustment on key parameters such as a paddle disc load when the task requirement is not met, so as to form a closed-loop concept design process driven by the task requirement. According to the method, the vertical take-off and landing capacity and the fixed wing cruising performance can be systematically considered in the conceptual design stage, and systematicness, repeatability and engineering feasibility of scheme design of the folding type tilting four-rotor aircraft are improved.

The invention discloses a rotorcraft flutter model design and test method for a low-speed wind tunnel test, and belongs to the technical field of low-speed wind tunnel tests. The method solves the difficult problem of how to separate key nacelle modes in rotation flutter of the tilt rotorcraft. A fuselage is connected with the side wall of a wind tunnel, wings are of a double-girder straight wing structure, one end of each wing is connected with the fuselage, the other end of each wing is connected with a nacelle mechanism, a rotating paddle is connected with the nacelle mechanism, and a motor, a yawing mechanism and a pitching mechanism are arranged in the nacelle mechanism. The yawing mechanism and the pitching mechanism are connected in a matched mode to achieve decoupling in the yawing direction and the pitching direction of the model, the motor is installed on the pitching mechanism, and the execution end of the motor is connected with a main shaft of the rotating paddle. According to the rotorcraft flutter model design and test method for the low-speed wind tunnel test, decoupling of modes in the pitching direction and the yawing direction of the model can be achieved, the flutter characteristic of the model is shown as coupling of the pitching mode and the yawing mode of a nacelle, and the test requirement of the model wind tunnel test is met.

The invention provides a system and a method for realizing air end tool switching, and belongs to the technical field of air docking of unmanned aerial vehicles. The system provided by the invention comprises an operation unmanned aerial vehicle, the operation unmanned aerial vehicle is provided with a terminal tool, the operation unmanned aerial vehicle is configured to obtain positioning information of a toolbox unmanned aerial vehicle below and size information of a plurality of replacement tools on the toolbox unmanned aerial vehicle in the air, and a target replacement tool is determined based on the size information, realizing end tool switching based on contact between the positioning information and the target replacement tool; a plurality of replacement tools are arranged on the top plane of the tool box unmanned aerial vehicle, and when the operation unmanned aerial vehicle makes contact with the tool box unmanned aerial vehicle to replace the tail end tools, the tool box unmanned aerial vehicle body and the operation unmanned aerial vehicle body are in flexible butt joint. According to the system and the method for realizing air terminal tool switching provided by the invention, the air flexible docking of the unmanned aerial vehicle can be realized.

The utility model belongs to the technical field of flexible structure design, and particularly relates to a perimeter-variable fuselage frame structure which comprises sliding groove sections, sliding key sections, corrugated beams, height-variable supporting columns and length-variable supporting rods. The corrugated beams are arranged on the outer sides of the sliding key sections and the sliding groove sections and are provided with wave crest and wave trough structures; the sliding key sections and the sliding groove sections are located on the inner sides of the corrugated beams, arranged at intervals and distributed in a staggered mode. The corrugated beams, the sliding key sections and the sliding groove sections are all of arc-shaped structures, and wave crests and wave troughs of the corrugated beams are arranged in the radial direction of the arc-shaped structures; the sliding key section and the sliding groove section are mutually nested to form an arc-shaped supporting sliding way, and the wave trough of the corrugated beam slides on the arc-shaped supporting sliding way; the height-variable supporting column and the length-variable supporting rod are hinged to the two ends of the sliding key section respectively. The shape of the fuselage can be greatly changed, the large-size fuselage can provide a large space to carry more fuel, and the small-size fuselage is beneficial for reducing flight resistance and improving the fuel utilization rate.

The invention discloses a land leveler automatic control and guidance method and system based on Beidou high-precision positioning. The method comprises the following steps: acquiring comprehensive state information of a land leveler including a three-dimensional position, a course, an attitude and a real-time elevation of a scraper knife; outputting future state parameters of a scraper knife elevation adjusting mechanism and a machine body steering mechanism of the land leveler in the rolling time domain through a preset land leveler kinematics model on the basis of the comprehensive state information and the current target working plane design information; based on the future state parameters, determining an optimal pre-control instruction sequence for a scraper knife elevation adjusting mechanism and a fuselage steering mechanism through a multi-objective optimization mode; and the optimal pre-control instruction sequence is sent to an electro-hydraulic control system to drive an execution mechanism to act. Complex factors such as topographic changes and fuselage dynamics coupling can be effectively coped with by predicting future states by adopting a kinematic model and prospectively generating a coordination control instruction, and the accuracy and smoothness of scraper knife control in the dynamic operation process are ensured.

The utility model discloses a water area propeller and water area movable equipment. The water area propeller comprises a machine body, an operation device and a locking module. A propelling device used for outputting propelling force for propelling the water area carrier to move is arranged at the end of the machine body. The operation device is rotationally connected to the end, away from the propelling device, of the fuselage body through a transverse connecting shaft, can be folded or unfolded with the fuselage body and is used for being operated by a user to control the water area propeller to operate. The locking module comprises a first locking unit and a second locking unit, and the first locking unit and the second locking unit are arranged in the operation device and the machine body respectively. The suction force between the first locking unit and the second locking unit of the locking module locks the operating device and the fuselage body at the position between the transverse connecting shaft and the propelling device, and the operating device forms a cantilever beam capable of loading the fuselage body between the transverse connecting shaft and the locking module.

The present invention relates to a vertical take-off and landing (VTOL) aircraft featuring a central fuselage and at least two power units. Each power unit is connected to at least two separate lifting rotors via transmission lines, ensuring that each power unit is linked to at least two rotors diametrically opposed relative to the centre of gravity of the aircraft. The lifting rotors might be driven by power transmission lines of hydrostatic type. The power units on either side might be housed in separate pods which might be incorporated into the main fuselage or arranged as separate pods. These separate pods might also function as floats / pontoons for the aircraft to land, take-off and manoeuvre on water.

The invention discloses a fixed-wing unmanned aerial vehicle rocket separation mechanism which comprises a boosting limiting device arranged on a boosting rocket, the boosting limiting device comprises a boosting limiting block fixedly connected with the bottom of an unmanned aerial vehicle and provided with a through hole, and a boosting limiting taper pin fixed in the through hole, and one end of the boosting limiting taper pin forms a conical end face; the boosting limiting device further comprises a conical groove which is formed in the boosting rocket and matched with the conical end face. In the boosting process, the conical end face and the conical groove form surface contact and abut against and are attached, it is guaranteed that thrust is evenly transmitted, the stress concentration risk of the fuselage is reduced, meanwhile, interference of non-directional airflow in the boosting process is counteracted, and the situation that the attitude of the unmanned aerial vehicle is out of control is avoided. And after boosting is completed, unpowered natural separation of the unmanned aerial vehicle and the boosting limiting device is achieved through the resistance difference of air to the unmanned aerial vehicle and the boosting limiting device, the conical groove guides the conical end face, additional driving parts are not needed, the structure is simplified, and the flight safety of the separated unmanned aerial vehicle is improved.

The invention provides a modular free-combination rotor unmanned aerial vehicle and a combination method, and the unmanned aerial vehicle comprises a plurality of independent rotor unmanned aerial vehicle units, and each rotor unmanned aerial vehicle unit comprises a fuselage and a rotor; a battery, a motor, a communication module and a power control module are arranged in the machine body; a standard socket is arranged outside the fuselage and used for being fixedly connected with other rotor wing unmanned aerial vehicle units or rods with the same interfaces, and sharing of power supply and communication is achieved through the standard socket. The power control module dynamically adjusts rotor power distribution in real time according to the form of the combined unmanned aerial vehicle, and the communication module automatically establishes a master-slave control relation after combination to achieve unified coordination of flight control. According to the technical scheme provided by the invention, through free combination of the modular rotor unmanned aerial vehicle units, a specific unmanned aerial vehicle can be quickly formed, the purchase and use cost is reduced, and the flight stability and safety are improved.

Aircraft capable of hovering comprising a fuselage at least one rotor that is rotatable with respect to the fuselage, electrical drive means adapted to rotate the rotor and electrically powered by a battery, and a cooling system of the battery. Said cooling system comprises, in turn, a first opening for the air to enter, second openings for the air to escape and a passage which places the first opening in fluidic communication with the second openings and within which the battery is placed. The cooling system further comprises a fan adapted to increase the kinetic energy of the air contained in the passage and which is operated when the forward speed of the aircraft with respect to the ground is lower than a speed threshold value and / or when the temperature of the battery exceeds a temperature threshold value.

The invention belongs to the technical field of aircraft structure design, and discloses a fuselage frame preparation process and a fuselage frame. According to the fuselage frame preparation process, preparation of the first woven preformed body and the second woven preformed body is achieved through the weaving operation of the weaving machine, the preparation efficiency and the structural integrity of the first woven preformed body and the second woven preformed body are improved, and the influence of manual operation on the fuselage frame forming quality is avoided. The preparation of the first weaving body and the second weaving body is realized by cutting operation of cutting guide parts arranged on the first weaving auxiliary core mold and the second weaving auxiliary core mold in advance, so that the preparation precision is guaranteed, and the forming quality of the fuselage frame is improved. The invention further provides the machine body frame, the machine body frame is prepared through the machine body frame preparation technology, the manual operation participation degree is reduced in the machine body frame preparation process, the influence of manual operation errors on the forming quality of the machine body frame is avoided, the forming quality and forming efficiency of the machine body frame are improved, and the manual labor intensity is relieved.

A linear shaped charged Electro-Explosive (LSCe) Device includes an initiator in direct and intimate contact with an exposed portion of the LSC's main charge. In different configurations, the initiator's package, initiation charge or initiation circuit (acting over a gap) are in direct and intimate contact with the main charge. The LSCe device may be manufactured by connecting a packaged initiator to an opening in the LSC housing through which a portion of the main charge is exposed or by forming the initiator in situ in a receptacle coupled to the housing around the opening. The LSCe device can be used for such applications as thrust or mechanical termination of an airframe, cable cutting or initiation of a LSC.

The invention belongs to the technical field of aircrafts, and discloses a trigger-containing aircraft composite material energy absorption structure analysis method, which comprises the following steps: establishing an external trigger composite material structure model according to actual working conditions; constructing an intra-layer and inter-layer damage coupled progressive damage model; configuring finite element model parameters adaptive to the composite material; integrating simulation results; the crashworthiness of the coupling trigger structure is adjusted and optimized in combination with basic parameters; repeating the steps to verify the optimization effect; and predicting the strength of the composite material by coupling stress strain and damage transfer, revealing a damage failure mechanism and determining an optimal design. The method has the beneficial effects that the failure mode and the load response of the aircraft structure under different impact loads can be accurately captured, and low initial peak load and high specific energy absorption requirements are considered through multi-type triggering collaborative design; reliable technical support is provided for crashworthiness design of composite aircraft structures of aircraft fuselage frames, wing supports and other key structures in the aerospace field.

An assembly is provided for an aircraft. This aircraft assembly includes an airframe structure, a propulsion system and a vane. The airframe structure includes a fuselage and at least two wings. The fuselage extends longitudinally along a longitudinal centerline from a forward end of the fuselage to an aft end of the fuselage. The wings extend laterally from opposing sides of the fuselage. The propulsion system is mounted to the airframe structure. The propulsion system includes an airflow inlet into the propulsion system. The airflow inlet is spaced vertically out from a top side of the airframe structure. The vane projects vertically out from the top side of the airframe structure to a distal end of the vane. The vane laterally overlaps and is longitudinally forward of the airflow inlet.

The application discloses a kind of based on six degrees of freedom parallel mechanism's aircraft cabin section butt joint method, comprising: S1: construct global assembly coordinate system, ground reference point is laid to specified position;S2: numerical control positioner is hoisted to specified work station;S3: fuselage shape frame is hoisted to specified work station;S4: middle and rear fuselage is connected with fuselage shape frame;S5: middle and rear fuselage pose is adjusted;S6: rear fuselage is connected with fuselage shape frame;S7: rear fuselage pose is adjusted;S8: by numerical control positioner, middle and rear fuselage is connected with rear fuselage;The application compact structure is stable, with good rigidity and carrying capacity, suitable for large, heavy fuselage cabin body butt joint work, process planning efficiency is high, reliable high strength.

The invention discloses a wide-speed-range variant high-speed aircraft aerodynamic layout, and belongs to the technical field of aircrafts, the wide-speed-range variant high-speed aircraft aerodynamic layout comprises a fuselage, the front end of the fuselage is provided with a drag reduction rod, the rear end is provided with a thrust vectoring nozzle, the left and right sides are symmetrically provided with triangular sweepback wings, and the top is provided with a supporting structure; a high-pressure capturing wing is arranged on the top of the supporting structure. The shock wave separation effect of the drag reduction rod is combined with the high-pressure capture and reflection shock wave induction mechanism of the high-pressure capture wings, so that the aircraft can achieve effective drag reduction and lift augmentation effects in different flight states, and the lift-drag ratio and adaptability of the high-speed blunt cone aircraft are remarkably improved.

An aircraft having a plurality of independent yaw control mechanisms includes an airframe having a central fuselage sized to hold at least one operator or payload and a wing extending from the fuselage. The aircraft includes a distributed thrust array coupled to the airframe, the thrust array having at least a first, second, and third pair of propulsion assemblies. Each propulsion assembly includes a rotor and is operable for at least single-axis thrust vectoring. The aircraft further includes a flight control system operable to independently control and combine each yaw mechanism of the propulsion assemblies. The yaw mechanisms for inducing a yaw moment include: canting at least one pair of propulsion assemblies away from the fuselage, selectively tilting at least one pair of propulsion assemblies forwards and backwards, varying an aerodynamic control surface of at least one propulsion assembly, and varying a rotational speed of at least one propulsion assembly.

Examples relate to uncrewed aerial vehicles (UAVs) and methods for controlled descent during control tier failures. A computing device may initially detect a control tier failure at an UAV. In some examples, the UAV includes a fuselage, a pair of wings extending outwardly from the fuselage, and a pair of stabilizers arranged in a V-shape configuration. Each stabilizer has a control surface that is adjustable relative to a fixed portion of the stabilizer. Based on detecting the control tier failure at the UAV, the computing device may adjust the control surface of each stabilizer from a first angle to a second angle relative to the fixed portion of the stabilizer. By adjusting the angle between the control surfaces and fixed portions of one or both stabilizers, the UAV may induce a deep stall maneuver that can enable a controlled descent of the UAV.

The invention belongs to the field of aviation structure strength tests, and particularly relates to a testing machine pitching and rolling angle adjusting method and system for a full-aircraft drop test, and the method comprises the steps: adjusting the lengths of a front lifting rope l1 and a rear lifting rope l2 through a first servo motor; the lengths of the adjusted front lifting rope l1 and the adjusted rear lifting rope l2 are calculated through the cosine law according to the geometrical relationship; the lengths of the left lifting rope 13 and the right lifting rope 14 are adjusted through a second servo motor; and the lengths of the adjusted left lifting rope 13 and the right lifting rope 14 can be calculated through the cosine law according to the geometrical relationship. The pitching and rolling postures can be automatically adjusted, and the problems that an existing method is low in efficiency and poor in precision are solved. According to test experience, according to an existing posture adjusting method, replacement needs about half a day to one day every time, the number of times of putting every day does not exceed ten at most, and manual operation is needed on wings or a fuselage every time. And after automatic adjustment is achieved, the target can be achieved in several minutes through adjustment of the motor, efficiency is high, and safety is higher.

An aircraft fuselage for releasing a fuel tank from inside the aircraft fuselage in case of emergency includes an outer structure, a mounting device to releasably secure a sled device for carrying a fuel tank in an inside of the outer structure, and a first openable release door separating the inside from an outside of the outer structure, wherein the first release door is positioned in relation to the mounting device such that the fuel tank is released through the first release door when the mounting device releases the sled device. A method is disclosed for releasing a fuel tank from inside the aircraft fuselage in case of emergency as well as an aircraft including such an aircraft fuselage.

The invention discloses a multi-rotor unmanned aerial vehicle carrying satellite communication equipment and an aircraft thereof, the unmanned aerial vehicle adopts a six-rotor layout and has flight stability and redundancy control capability, and when a certain rotor fails, flight is maintained through the remaining rotors; a lightweight satellite communication module is integrated, and a plurality of antennas are arranged below the unmanned aerial vehicle body; peripheral obstacle avoidance sensing devices are mounted at the tail ends of connecting rods of the six groups of rotors, and vertical obstacle avoidance sensing devices are mounted above and below the head of the fuselage respectively. A flight control program adopts an intelligent filtering algorithm, small target interference data is automatically filtered out, and false triggering is prevented. Compared with a conventional multi-rotor unmanned aerial vehicle, the unmanned aerial vehicle has the characteristic of carrying satellite communication equipment, and unmanned aerial vehicle control communication can be carried out nationwide through the satellite communication equipment; meanwhile, the unmanned aerial vehicle has the characteristic of long endurance, and can fly for 2 hours and above 60km; the problems of limited remote communication, short endurance, short voyage and the like in the prior art can be solved.

The invention provides a heavy-load unmanned aerial vehicle landing point identification processing method and system based on visual identification, and relates to the technical field of computer vision, and the method comprises the steps: obtaining the airflow speed direction and airflow change data above a landing area of a heavy-load unmanned aerial vehicle, real-time flight data, and an enhanced panoramic image; based on the initial positioning position of the heavy-load unmanned aerial vehicle, the initial coordinates of the preset landing area and the flight task parameters, generating an initial landing trajectory to determine a qualified landing point position; based on the qualified landing point position, the enhanced panoramic image, the airflow velocity direction and the airflow change data, simulating a fuselage stress simulation result of the heavy-load unmanned aerial vehicle in the landing process; based on the qualified landing point position and the fuselage stress simulation result, obtaining corrected flight data; and based on the qualified landing point position and the corrected flight data, an optimized landing trajectory is obtained, and high-precision and high-stability landing control of the heavy-load unmanned aerial vehicle in a complex environment is realized.

The embodiment of the invention discloses an aircraft height measurement method based on binocular vision, and the method comprises the steps: obtaining internal parameters and distortion parameters of a first camera and a second camera, and obtaining a first image and a second image collected by the first camera and the second camera; calibrating the first image and the second image according to the internal reference and the distortion parameter of the first camera and the second camera to obtain a first calibration image and a second calibration image; binocular matching is carried out on the first calibration image and the second calibration image, and a full-image disparity map is obtained based on the first calibration image; performing airport runway semantic segmentation based on the first calibration image to obtain a semantic segmentation map of the airport runway; performing 3D coordinate reduction operation on the 2D pixels according to the full-image disparity map and the semantic segmentation map to obtain a 3D point cloud of the airport runway; fitting a runway plane equation according to the 3D point cloud, converting the runway plane equation to an airframe coordinate system according to the relative position of the first camera and the airframe, and obtaining an aircraft height measurement result according to the runway plane equation under the airframe coordinate system.

The invention relates to the technical field of unmanned aerial vehicles, and particularly discloses an unmanned aerial vehicle capable of landing and buffering, which comprises a fuselage shell, connecting arms are fixedly connected to the opposite angles of the side surfaces of the fuselage shell, and fan blade protection rings are fixedly connected to the ends, away from the fuselage shell, of the connecting arms. A driving motor is fixedly connected to the inner wall of the fan blade protection ring through a support, a driving shaft of the driving motor is fixedly connected with flying fan blades, a mounting frame is fixedly connected to the bottom of the fuselage shell, a damping device is fixedly connected to the bottom of the inner wall of the mounting frame, and a buffering mechanism is fixedly connected to the end, away from the supporting frame, of a fixing rod. According to the unmanned aerial vehicle capable of landing and buffering, the buffering mechanism is arranged, so that the impact force generated when the unmanned aerial vehicle descends can be buffered, and the damping device is arranged, so that the buffered vehicle body can be further damped; damage of parts in the unmanned aerial vehicle caused by violent vibration during descending is avoided.

A vertical take-off and landing aircraft includes: a fuselage and 2N tilting rotors. Both sides of the fuselage are provided with a wing symmetrically, a tail of the fuselage is provided with an empennage, and ruddervators are provided at the empennage. 2N tilting rotors are symmetrically provided at both sides of the fuselage, and a part of the 2N tilting rotors are provided on the empennage. N is a natural number greater than or equal to 2; in a vertical take-off and landing state, projections of propellers of the 2N tilting rotors on a horizontal plane are centrally symmetrical about point B, point B and a center of gravity point G of the vertical take-off and landing aircraft are both provided in a symmetry plane of the fuselage.