195 results about "Level flight" patented technology

The invention discloses an automatic control method of a four-power-source tiltable aircraft, relates to the technical field of flight control of tiltable aircrafts, is realized based on a PX4 flight control system, and is suitable for a tiltable aircraft provided with four independent tiltable power sources and corresponding tilting execution mechanisms. According to the method, through configuration of a tilting angle control module, a power distribution module and a full-stage flight logic scheduling module, full-stage automatic control of vertical take-off and landing, suspension, horizontal take-off, fixed-point horizontal flight, automatic cruise, one-key return flight, fixed-angle flight, power source fault handling and the like is achieved. Smooth transition of all flight modes is achieved, the stability of the attitude, the height and the position of the tiltable aircraft in different flight stages is maintained, meanwhile, the functions of automatic obstacle avoidance, fault emergency and the like are integrated, flight safety is guaranteed, flight stability and applicability are improved, and the method can be widely applied to various tiltable aircraft operation scenes.

The invention discloses a flight trainee attention and error behavior detection method based on a multi-modal physiological signal, and the method comprises the steps: obtaining the electrocardiogram data, eye movement data and electroencephalogram data of a flight trainee in a flight training process, and subjective evaluation data; based on subjective evaluation data, dividing flight student attention distribution conditions into a high grade, a middle grade and a low grade by adopting a clustering algorithm; performing feature optimization processing on the physiological data, and constructing a physiological feature attention level flight behavior associated data set; constructing an attention distribution level judgment model and a flight error behavior detection and early warning model; and outputting an attention level judgment result and a flight error behavior detection and early warning result. According to the method, association mapping of the multi-modal physiological signals and flight performance is realized, an attention grading standard for a flight training specific scene is established, and conversion from passive response to active early warning is realized.

The invention belongs to the technical field of flight test, and discloses a cockpit vision field evaluation test flight method suitable for a troposphere manned airship, which comprises the following steps: selecting an airship test condition; the airship climbs to a preset field height and keeps level flight; a preset flight test action is completed; and commenting by a flight tester, and filling in a cockpit vision field commenting card. According to the method, the blank of the flight test method for evaluating the visual field of the cockpit of the troposphere manned airship with the vector propulsion system is filled, and the flight test efficiency and the flight test quality of the manned airship are improved.

A vertical take-off and landing aircraft which uses fixed rotors for both VTOL and forward flight operations. Some of the rotors are forward canted, while some of the rotors are forward facing. The forward canted rotors are tilted forward slightly from vertical and provide some forward propulsion during horizontal flight. The forward facing rotors may be fixed in a forward facing configuration. Forward swept wings allow for rotor placement which brings efficiency during hover operations.

The invention discloses an unmanned aerial vehicle route planning method, device and equipment and a medium, and the method comprises the steps: obtaining the real-time transient wind field data of an unmanned aerial vehicle, and generating the flow field information of time-space correlation according to the real-time transient wind field data; a flight energy consumption model is established based on the flow field information, the flight energy consumption model is adopted to calculate energy consumption in different stages, node energy consumption attribute values are obtained, and the different stages comprise a level flight stage, a climbing stage and a descending stage; and the node energy consumption attribute value is input into the heuristic search algorithm model to obtain the current path planning with the lowest energy consumption, so that the accuracy of the route planning is improved, and the route energy consumption is saved.

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.

A flight mode transitioning device, method and system configured to facilitate transitioning between vertical flight mode and horizontal flight mode in a vertical take-off and landing (VTOL) aircraft, the device comprising at least one actuator mediated member capable of limiting the movement the aircraft wing so as to enable both forward fight and vertical flight.

The invention discloses a novel tilting dual-rotor unmanned aerial vehicle, which belongs to the technical field of unmanned aerial vehicles, and comprises a wing body fusion fuselage and two outer wing sections, the two tilting rotor assemblies are used for providing power, controlling rolling, pitching and yawing in a vertical flight mode and controlling yawing in a horizontal flight mode, the two pitching and rolling rudder assemblies are used for controlling pitching and rolling in a horizontal flight state, and the two tilting rotor assemblies are symmetrically arranged in front of the wing body fusion fuselage; the pitching rolling rudder assembly is arranged in the middle of the outer wing section; the outer wing section is of a structure capable of being folded in the tail direction, so that the center of gravity is moved backwards to the position behind a stop point of the tail after the outer wing section is folded, and the unmanned aerial vehicle stably stops on a rod-shaped object with the diameter being 13 cm or less under the unpowered condition. On one hand, the problems that a tailstock type unmanned aerial vehicle is difficult to take off and land and a traditional unmanned aerial vehicle needs an undercarriage to take off and land are solved, and the problems that an existing unmanned aerial vehicle needs to be additionally provided with a parking actuator and needs to be parked with power are solved.

The application belongs to the technical field of aircraft simulation, and particularly relates to a flight simulation method and device under gust load. The method comprises the following steps: S1, performing aircraft 1g level flight trimming to determine the aircraft trimming angle of attack and the aircraft trimming elevator deflection; S2, determining the aerodynamic force of the aircraft; S3, performing dynamic response solution on the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, the flight speed and the pitch angular velocity, and determining the normal overload of the aircraft after superimposing the flight speed on the wind speed of the vertical gust field; S4, obtaining the elevator deflection calculated by the stability augmentation system, superimposing the elevator deflection at a simulation moment to obtain the actual required elevator deflection; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angular velocity, the aerodynamic angle of attack and the actual required elevator deflection until the simulation is completed. The application considers the role of the stability augmentation system in the determination of the aircraft gust load, and the simulation process is simple and has high precision.

1. Name of the product in this design: Graphical User Interface for Adding Flight Information to Electronic Devices. 2. Intended use of this design: for use in an electronic device. 3. The key design feature of this product is its graphical user interface. 4. The image or photo that best illustrates the design points: Design 1 Interface Change State Diagram. 5. Other views are omitted as they are standard design. 6. Design 1 is designated as the basic design. 7. Purpose of the graphical user interface: This interface is used to add flight information in city-level flight service systems, such as operator information or aircraft information. Design 1: The main view is the initial interface. Figure 1 shows the interface that appears after clicking the "+" icon in the upper right corner of the main view of Design 1. Design 2's interaction is consistent with Design 1.

The invention discloses a fixed wing patrol aircraft flight attitude control method, and belongs to the technical field of unmanned aerial vehicles. Aiming at the technical contradiction that high-speed flight maneuverability and weapon shooting stability cannot be compatible, the method achieves breakthrough through the following innovative schemes that a four-wing propeller collaborative framework is designed, and each propeller is provided with a horizontal linear driver to achieve front-back telescopic displacement in the axial direction of a fuselage; in the horizontal flight stage, based on the feedback of a gyro stabilizer, pitching / rolling double-axis attitude accurate control is realized through the cooperative adjustment of propeller position stretching and rotation speed difference; triggering a layered compensation strategy by an ignition event, and synchronously adjusting the horizontal position and the rotating speed of a propeller according to the recoil force direction; and the tail vane intervenes in fine adjustment only when the yaw is out of tolerance. By combining the seamless conversion design of the vertical take-off and landing stage, the three technical problems of height loss without initial-speed deployment, low aerodynamic efficiency of high-speed flight and insufficient shooting precision are solved.

The invention provides a method and a device for judging the flight stage of a transport aircraft conforming to the operation expectation of a pilot, and the method comprises the steps: determining a guided flight stage of a current period according to a preset period, a working mode of an autopilot, a preselected height, a horizontal flight plan and a vertical flight plan; the guided flight stage comprises climbing CLM, cruising CRS and descending DES; the guide flight stage is used for generating a vertical guide instruction of the flight management system; and determining a man-machine interface flight stage according to the working mode of the autopilot, the preselected height, the horizontal flight plan and the vertical flight plan based on the judgment result of the guided flight stage. The method accords with the operation intention of the unit, enhances the vertical situation awareness capability of the unit, and provides a precondition for the output of the vertical guide instruction combination.

The present application relates to the technical field of aircraft operating mechanism, in particular to a drive-control integrated aircraft side stick operating mechanism, which comprises a box body, a controller and a driver, a roll transmission shaft is rotatably supported in the box body, a U-shaped decoupling mounting seat is fixedly connected to one end of the roll transmission shaft, a pitch shaft is horizontally supported in the U-shaped decoupling mounting seat, a base is fixedly sleeved on the pitch shaft, a handle connecting rod and a handle assembly are sequentially connected to the upper end of the base, a roll motor drives a rocker arm I through a rocker arm II and a transmission connecting rod I, the rocker arm I is hinged to the roll transmission shaft, a pitch motor drives a fixed head through a rocker arm III and a transmission connecting rod II, and the fixed head is fixedly arranged at the end of the pitch shaft. The present application provides a drive-control integrated aircraft side stick operating mechanism with high integration, ultra-small volume, double-channel synchronous force sensing, microsecond-level real-time bus communication and plug-and-play, which is particularly suitable for the demand of high-fidelity, high-reliability and small-sized operating device of high-level flight simulator.

To provide a different hybrid aircraft by maximizing the revenue load that the hybrid aircraft can transport.SOLUTION: The present invention relates to a vertical take-off and landing (VTOL) aircraft, a method of controlling a VTOL aircraft, and a control system for controlling a VTOL aircraft. The aircraft comprises a fuselage having wings extending along transverse axes and attached to a fuselage extending between longitudinal axes of the aircraft, and a tail unit or a front tail unit. An array of electric rotors is fixedly mounted to the airframe. The front and rear internal combustion engines are pivotally mounted to the fuselage, and the front and rear rotors are displaceable between a lift position oriented to provide vertical lift to the aircraft for vertical flight and a propulsion position in which the front and rear rotors are oriented to provide forward thrust to the aircraft for horizontal flight. The front and rear rotors provide most or all of the vertical lift for the aircraft during vertical flight.SELECTED DRAWING: Figure 1

This invention discloses a blown-wing vertical takeoff and landing (VTOL) fixed-wing aircraft and its control method, belonging to the field of aircraft technology. It includes a fuselage, main wing, retractable wing, tail-mounted tilting electric duct, and a control system. The fuselage has a built-in pressurized duct and an air intake at the front. An array of nozzles is arranged on the upper surface of the main wing and connected to the pressurized duct. The retractable wing is mounted on the trailing edge of the main wing. Symmetrically arranged tail-mounted electric ducts capable of vertical / horizontal tilting are located at the tail. The entire aircraft has no exposed rotating blades. During VTOL, the retractable wing deploys, and the pressurized duct supplies air to the array of nozzles, generating primary lift through the Coanda effect. The retractable wing guides air to generate secondary lift, and the tail-mounted duct provides vertical thrust for auxiliary lift and attitude adjustment. During level flight, the retractable wing retracts, and the tail-mounted duct provides horizontal rearward cruise thrust. This invention achieves smooth switching between VTOL, transitional flight, and level flight cruise through its control method, making it suitable for emergency rescue and low-altitude operations in complex terrain scenarios.

The application belongs to the technical field of overall design of vertical take-off and landing aircraft, and particularly relates to a single-blade rotor type vertical take-off and landing fixed-wing aircraft power layout structure. The main body of the power layout structure is a rotor with only one blade, one side of the rotor rotating shaft is a load cabin for loading and counterweight, and the other side is a fixed pitch rotor blade, and the rotor blade can swing; the rotor blade is provided with a driving propeller for driving the rotor to rotate or generating a level flight thrust. The single-blade rotor is a main component for generating lift in vertical take-off (rotor mode) and cruising flight (fixed-wing mode). The single-blade rotor system provided by the application can be used as a power system of a vertical take-off and landing aircraft, the power layout can consider cruising level flight efficiency and vertical take-off and landing hovering efficiency, and static balance and dynamic balance can be realized in a steady state hovering state.

The invention discloses a fan blade inspection local dynamic path planning method and system based on an unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicles, and the method comprises the following steps: S1, a template generation stage: based on a blade three-dimensional model, generating a standard inspection template, the standard routing inspection template defines a global optimal routing inspection sequence covering the surface of the blade and a starting point of the global optimal routing inspection sequence; s2, a path mapping stage: mapping the standard inspection template to an actual blade, and generating a three-dimensional space path which keeps a safe distance from the curved surface of the blade based on a preset rule; and S3, an online execution stage: the unmanned aerial vehicle executes the three-dimensional space path. The method solves the problem that the conventional unmanned aerial vehicle path planning method does not fully consider the physical limitation that the unmanned aerial vehicle must keep approximately horizontal flight in the inspection process, so that the unmanned aerial vehicle is forced to frequently adjust the attitude, and the inspection efficiency is greatly reduced.

Symmetrical balance of the helicopter is achieved by selecting a symmetrically-designed novel piston type engine, a fixed horizontal umbrella wing is arranged on the top, and when the helicopter flies normally and horizontally, linear resistance is small; the damping effect of avoiding too high speed is achieved during up-and-down rising and falling. The helicopter is completely balanced in a static state, and when the helicopter flies, the thrust formed by airflow entering the upper portion of the main rotor to the small louver fan wings with a certain inclination angle on the upper portion of the main rotor and the thrust formed by airflow flowing out of the lower portion of the main rotor to the balance wings with the adjustable inclination angle are utilized. And the resultant force equivalently counteracts the counter-acting force of the engine by adjusting the inclination angle, so that the aircraft is stable in direction. A novel piston type engine is adopted, due to the unique structural design, the rotating speed and variable work of the engine can be accurately adjusted in a high-precision mode, and a power shaft of the engine vertically operates. A small-diameter turbofan'main rotor 'with a fixed attack angle is adopted, so that conversion of horizontal power into vertical power through a bevel gear is omitted. The aircraft flies at a low speed through forward or backward force generated by downwash airflow of the main rotor to the angle-adjustable driving wings on the lower portion of the main rotor.

The application relates to the technical field of vector thrust technology and discloses a large-range variable-area-ratio vector exhaust device and an exhaust method thereof. The device comprises a transition section, a guide vane section and a rotating afterbody; the guide vane section is connected with the outlet of the transition section, and the rotating afterbody is connected with the rear side of the transition section; when the device is switched between different working states including a vertical take-off and landing state, a transition state and a level flight state, the airflow direction and the jet angle are adjusted by changing the positions of the guide vane section and the rotating afterbody. The application can realize the continuous deflection of the airflow from 0 DEG to 90 DEG, ensure that an aircraft maintains excellent aerodynamic performance and generates sufficient thrust in different flight states, meet the large-outlet-area-ratio requirements in the vertical take-off and landing state and the level flight state, effectively cope with the large-flow change in the process from the vertical take-off to the level flight, realize a high total pressure recovery coefficient, improve the aerodynamic performance and provide strong thrust in the vertical take-off and landing state and the level flight state.

The invention particularly relates to a flight action recognition method for long-time sequence context modeling, which comprises the following steps of: acquiring original flight parameter time sequence data, and performing data enhancement through Gaussian noise disturbance, random scaling, linear drift and time axis disturbance; adopting dynamic sequence splicing to generate a composite long sequence; establishing a multi-window independent training system based on the spliced data, and setting a plurality of window lengths to capture action features of different time scales; each sub-window model adopts a BiGRU (Big Generation Regulation Unit)-Transform framework; each window model is trained independently, a cross entropy loss function and an RMSprop optimizer are adopted, and Dropout regularization is combined to prevent overfitting; in the reasoning stage, multi-window prediction results are subjected to weighted fusion according to weights, a frame-level flight action recognition result is output, and the problems that in the prior art, a test flight action recognition algorithm is difficult to consider local time sequence dependence and global context information in complex flight time sequence data, key feature extraction is insufficient, and the model generalization ability is insufficient are solved.

The application belongs to the field of aerospace technology, and particularly relates to a vertical take-off and landing horizontal flight combined power unmanned aerial vehicle, which comprises a fuselage, main wings, rotors, horizontal tail wings and movable vertical tail wings; the main wings are symmetrically connected to the middle part of the fuselage, and the rotors are symmetrically connected to the main wings on the two sides; the horizontal tail wings are symmetrically connected to the tail part of the fuselage, and the movable vertical tail wings are connected to the upper part of the tail part of the fuselage; a pressure measuring device is arranged on the movable vertical tail wings; the pressure measuring device collects the pressure on the two sides of the movable vertical tail wings to a flight control system, and the flight control system calculates the pressure difference between the left and right sides to control the deflection of the movable vertical tail wings. The cruising efficiency of the unmanned aerial vehicle is improved, the driving device and structure thereof are relatively simple and light compared with the full-motion vertical tail wing. Meanwhile, the unmanned aerial vehicle also has the cruising efficiency of a fixed-wing aircraft and the low requirement for a take-off and landing site of a rotor vertical take-off and landing aircraft.

The present application discloses an aircraft configured for vertical take-off and landing and horizontal flight. The aircraft includes one or more processors configured to control distribution of fuel among a plurality of aircraft fuel tanks. Controlling the distribution of fuel includes: (a) determining a first fuel level in a first nacelle fuel tank, wherein the first nacelle fuel tank is configured to provide a supply of fuel to a first engine during both (i) a vertical take-off and land operation, and (ii) a horizontal flight operation, (b) determining a fuselage fuel level in a fuselage fuel tank, and (c) determining whether to redistribute fuel across at least the first nacelle fuel tank and the fuselage fuel tank based at least in part on the first fuel level and the fuselage fuel level.

The application discloses a heavy-load long-endurance tilting rotor unmanned aerial vehicle. The unmanned aerial vehicle comprises a wing frame for supporting and tilting, unmanned wings installed on symmetrical two sides of the wing frame, a cargo hold installed on the lower part of the wing frame and used for loading goods, and an unmanned aerial vehicle hydraulic system installed on the cargo hold and used for controlling the unmanned wings. The unmanned aerial vehicle can realize vertical take-off and landing, hovering and fast level flight, is easy to control the moving direction and has strong environmental adaptability. The unmanned aerial vehicle has simple structure, low take-off requirement, small wind resistance, fast level flight speed, low energy consumption and long endurance, so that the voyage and load can be increased. The wing frame overall tilting system has low structural complexity, airflow disturbance is reduced, and the safety of flight mode conversion can be improved.

The invention provides an unmanned aerial vehicle airborne antenna layout optimization method, and relates to the technical field of unmanned aerial vehicle communication, and the method comprises the steps: obtaining an attitude angle time sequence of typical attitudes including level flight, climbing, descending, turning, hovering and the like according to a task demand, discretizing a pitch angle and a roll angle into a plurality of attitude units, and carrying out the optimization of the attitude units; counting the duration of each attitude unit in combination with historical data, and calculating the comprehensive weight of each attitude unit; a plurality of alternative antenna layout schemes are generated under the condition that geometric, engineering and dynamic constraint conditions are met, communication performance evaluation values under all attitude units are obtained through electromagnetic simulation, and the comprehensive communication performance of the schemes is obtained through weighting according to the weights of the attitude units; and finally, with maximization of comprehensive communication performance and minimization of engineering complexity as targets, multi-target sorting is performed by using a TOPSIS method, and an optimal antenna layout scheme is determined in combination with a Pareto solution set, so that collaborative optimization of communication performance and engineering feasibility of the unmanned aerial vehicle in a dynamic flight task is realized.

The invention discloses an unmanned aerial vehicle endurance time evaluation method, device and equipment and a storage medium, and the method comprises the steps: obtaining all index test data of an unmanned aerial vehicle, the index test data comprising endurance time of no-load level flight, full-load level flight, no-load hovering and full-load hovering; distributing the weight of each index test data, and calculating a flight environment correction coefficient; based on the index test data, the weight of the index test data and the flight environment correction coefficient, constructing an unmanned aerial vehicle comprehensive endurance index model; and performing quantitative evaluation on the endurance time of the unmanned aerial vehicle by using the comprehensive endurance index model. According to the invention, the evaluation comprehensiveness can be improved, and the accuracy and reliability of the test result can be improved.

This invention discloses a thin-walled, narrow-edge ducted fan that reduces crosswind interference. The fan comprises a duct shell, a multi-faceted frame, a trailing edge fairing, and a leading edge fairing. A drive motor is housed within the trailing edge fairing, and the bottom of the fairing is connected to the drive motor's base via screws. A motor rod extends from the drive motor, passing through the multi-faceted frame. The duct shell and leading edge fairing are fixedly connected by screws, and the motor rod is rotatably connected to the hub frame of the multi-faceted frame via bearings. The duct shell features a thin-walled, narrow-edge design, with the ratio of the duct shell's side length L to its inner diameter d being L / d ≤ 1 / 6. The interior of the duct shell is filled with foam, and the maximum wall thickness h of the outer wall satisfies either h = 1.5 mm, L ≤ 20 mm or 1.5 mm ≤ h ≤ 0.075L, L ≥ 20 mm. This design significantly reduces the weight of the shell compared to ordinary ducted fans, and when mounted on drones, it can reduce drag during high-speed level flight and mitigate crosswind effects during hovering and vertical takeoff and landing.

The invention belongs to the technical field of aircraft design, and particularly relates to a voyage performance estimation method and device for a hypersonic-speed gliding aircraft, and the method comprises the steps: S1, obtaining the gliding initial-state horizontal flight speed and gliding final-state horizontal flight speed of the hypersonic-speed gliding aircraft, and the aerodynamic lift L and resistance D of the hypersonic-speed gliding aircraft; and S2, calculating the sliding voyage s of the hypersonic-speed gliding aircraft according to the gliding initial-state horizontal flight speed and the gliding final-state horizontal flight speed of the hypersonic-speed gliding aircraft, and the aerodynamic lift L and the resistance D of the hypersonic-speed gliding aircraft. The gliding distance of the aircraft can be quickly estimated, and the gliding design efficiency of the aircraft is improved.

The application belongs to the technical field of aircraft simulation, and relates to a flight simulation method and device under gust load considering the downwash effect. The method comprises the following steps: S1, performing 1g level flight trimming of an aircraft to determine the aircraft trimming angle of attack and the aircraft trimming elevator deflection; S2, superimposing the aerodynamic force coefficient increment caused by the downwash angle of attack on the aerodynamic force coefficient of the aircraft body to determine the aerodynamic force of the aircraft; S3, performing dynamic response solution on the aerodynamic force of the aircraft to obtain the flight speed, superimposing the wind speed of the vertical gust field on the flight speed to determine the corrected speed of the aircraft, and determining the aerodynamic angle of attack and the flight Mach number according to the corrected speed; and S4, re-determining the aerodynamic force of the aircraft according to the flight Mach number, the pitch angular velocity, the aerodynamic angle of attack, the given wind speed of the vertical gust field and the elevator deflection, and performing the above steps cyclically until the simulation is completed. The simulation process of the application is simple and has high precision, and meets the needs of engineering practice.

The utility model relates to a tilting structure which comprises a support, a mounting seat, a connecting piece, a lead screw and a ball nut, one end of the support is rotationally connected with the mounting seat, the other end of the support is rotationally connected with the middle of the connecting piece, a rotating shaft is arranged at one end of the connecting piece, a tilting workpiece is arranged at the other end of the connecting piece, a motor and a bearing are arranged on the mounting seat, and the ball nut is arranged on an inner ring of the bearing. One end of the lead screw is connected with the ball nut in a matched mode, the other end of the lead screw is fixedly connected to the side face of the rotating shaft, and the lead screw pulls the rotating shaft to drive the connecting piece to rotate around a connecting point in the middle of the connecting piece. And the change of the tilting workpiece in the vertical direction and the horizontal direction can be realized. The utility model further relates to a tilting unmanned aerial vehicle which comprises a vehicle body and the tilting structure, the support is fixedly arranged on the vehicle body, and the tilting workpiece is a propeller motor. The tilting unmanned aerial vehicle applying the tilting structure can achieve the conversion purpose of vertical take-off and landing in a rotor wing state and horizontal flight in a fixed wing state.

The invention discloses a compact omnidirectional vector wind-resistant four-rotor aircraft and a control method. The aircraft comprises a fuselage, four vector machine arms distributed on the front, back, left and right sides of the fuselage in a cross shape, a flight controller, thrust motors and propellers. The vector machine arm comprises a machine arm body, an axial tilting steering engine, an axial rotating assembly, a tangential tilting steering engine and a limiting motor base. The axial tilting steering engine drives the axial rotating assembly to continuously rotate around the axis of the aircraft arm within the range of 90 degrees left and right, the tangential tilting steering engine drives the limiting motor base to continuously rotate relative to the axial rotating assembly within the range of 40 degrees inside and outside, and the limiting motor base is used for fixing the thrust motor. After being combined with the propellers, the thrust motor can generate three-dimensional vector thrust along with the two-degree-of-freedom movement. The defect that position control and attitude control of a traditional multi-rotor aircraft are mutually coupled is overcome, and high-wind-resistance-level flight under the condition that spot hover in any attitude at any position in the air is kept and the fuselage is horizontal is achieved.