250 results about "Thrust vectoring" patented technology

The invention relates to a control method and device for wind load resistance of a mooring unmanned aerial vehicle and a storage medium. The method comprises the following steps: collecting and modeling a mooring unmanned aerial vehicle to obtain a system dynamics model; state estimation is carried out to obtain system wind load disturbance data and expected attitude data; performing optimal control solution on the plurality of mooring points to obtain tension distribution control data of each mooring point; inputting into a discrete time neural network controller for attitude compensation calculation to obtain thrust vector compensation control data; performing adaptive robust control on the attitude of the mooring unmanned aerial vehicle to obtain a rotating speed control instruction of each motor of the four rotors; and performing safety threshold judgment, and when any monitoring data exceeds a preset threshold, outputting a working mode switching instruction and executing safe landing control. According to the invention, a multi-stage safety protection mechanism is established, and reliable operation of the mooring unmanned aerial vehicle under extreme conditions is ensured through real-time monitoring and mode switching.

A rotating pylon assembly for aircraft thrust vectoring control. In embodiments, the rotating pylon assembly of embodiments includes a rotating sleeve disposed within a fixed boom of the aircraft, and configured to support a rotor assembly. The rotating pylon assembly includes a bearing assembly configured to functionally couple the rotating sleeve against an inner surface of the fixed boom. The bearing assembly enables the rotating sleeve to rotate against the inner surface of the fixed boom and about a longitudinal axis of the fixed boom. An actuator is configured to rotate the rotating sleeve about the longitudinal axis of the fixed boom. The rotating sleeve may cause the rotor assembly to be rotated or swung about the longitudinal axis of the fixed boom to position the rotor assembly in a position to vector the thrust provided by the rotor assembly based on a control command.

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on-off state of the engine and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

The invention discloses a full-process closed-loop quick response satellite collision avoidance method, which comprises the following steps of: 1, observing and acquiring an orbit state of a space object, and then carrying out multi-source data fusion so as to construct a space object database; 2, screening out high-risk collision sources which have the possibility of collision with the own satellite; 3, obtaining a determined high-risk collision source according to the track state historical record of the possible high-risk collision source and the own track state; 4, generating an initial evasion maneuvering scheme, and optimizing the initial evasion maneuvering scheme through a multi-objective optimization algorithm; and obtaining a final evading maneuvering scheme for subsequent execution. And 5, according to the engine startup and shutdown time and the thrust vector contained in the final evasion maneuver scheme, the satellite engine is controlled to work with the specified thrust vector within the given startup and shutdown time, and collision evasion maneuver is executed. Compared with a conventional method, the method has the advantage that the response speed is higher.

The invention relates to the field of unmanned aerial vehicles, and discloses a high-wind-resistance multi-rotor unmanned aerial vehicle which comprises a vehicle body and vehicle arm supports arrayed on the side face of the vehicle body, a first motor is arranged at the tail end of each vehicle arm support, a propeller part is installed on an output shaft of each first motor, and the propeller parts are driven by steering engines to change the inclination angle; when the unmanned aerial vehicle encounters strong crosswind, the resistance of the unmanned aerial vehicle to the crosswind is improved by changing the rotation angles of the corresponding propellers and the propeller parts, so that the stability of the unmanned aerial vehicle encountering the strong crosswind in the flying process is improved, the thrust vector direction is changed, and the compensation force opposite to the wind pressure direction is generated. According to the dynamic adjustment mechanism, the unmanned aerial vehicle can keep the flight path stable under the strong crosswind condition, and the problem of flight path deviation caused by the fixed thrust direction of a traditional unmanned aerial vehicle is effectively avoided.

The invention relates to the technical field of ship power and propulsion control. The invention provides a multi-source cooperative propulsion control method and system based on hybrid power and a straight wing steering oar. The method comprises the following steps: collecting ship motion state parameters and environment parameters in real time; constructing a multi-objective joint optimization model, and solving an optimal power distribution proportion and an optimal thrust vector under the current working condition by adopting a dynamic programming algorithm; based on the optimal power distribution proportion, a fuzzy PID controller is adopted to dynamically adjust the torque coupling ratio of the diesel engine and the motor; according to the optimal thrust vector calculation result, vector synthesis is carried out on the propulsive force and the steering torque; and predicting the motion state of the ship based on an extended Kalman filter, and correcting the weight coefficient of the multi-target joint optimization model on line in combination with the change rate of the channel boundary distance. The problem that adaptability to complex working conditions is insufficient due to the fact that a traditional control strategy does not fully fuse ship motion states and environment parameters is solved.

The invention belongs to the technical field of three-dimensional vision measurement, and relates to a deformation measurement method and system of a composite vision measurement unit. The internal and external parameters of the binocular measurement unit are obtained according to the image of the thrust vectoring nozzle target area calibration plate; obtaining coordinates of all binocular measurement units in the same reference coordinate system according to the internal and external parameters and the images of the thrust vectoring nozzles with the first coded mark points, and obtaining coordinates of all binocular measurement units in a unified coordinate system; and acquiring full-field deformation of all target areas of the thrust vectoring nozzle according to the speckle images of the target areas of the thrust vectoring nozzle in the working state and the coordinates of the unified coordinate system of all binocular measurement units. According to the invention, a digital image correlation method and a close-range photogrammetry technology are utilized to realize coordinate system unification and full-field deformation measurement. Compared with a traditional method, the measurement times and data processing steps are reduced, and the measurement efficiency is improved.

The invention relates to the technical field of hovercars, and discloses a vertical-lifting flat-flying car which comprises a power system used for providing power, a wing-shaped car body, a multifunctional rotor wing, a tilting ducted fan engine and an empennage assembly, the wing-shaped car body comprises a car body front edge part and a car body rear edge part, and the tilting ducted fan engine provides power for lifting and cruising. The tilting ducted fan engine is arranged on the rear edge portion of the vehicle body and integrated with the empennage assembly, the empennage assembly is arranged on the rear edge portion of the vehicle body and works in cooperation with the tilting ducted fan engine, and aircraft yawing is corrected through thrust vectors of the two fan engines. According to the multifunctional rotor wing, the aircraft can stably take off and land vertically or hover, when the aircraft flies horizontally, the rotating rotor wing can be converted into a fixed wing, meanwhile, the thrust vector of the tilting ducted fan engine is rotated to the horizontal direction, the whole aircraft is changed into a fixed wing structure, and the cruising performance of the aircraft is improved.

An aerial and preferably marine vehicle intended to operate in its principal mode near the surface of water or land employing the Wing-in-Ground Effect (WIG) and capable to take-off and land vertically (VTOL) by means of thrust vectoring, which vehicle has a wing arranged at the lowermost part of the fuselage, propulsion units comprising four rotatable pylons extending transversely in pairs on both sides of the upper part of the fuselage, four elongated nacelles mounted on the outer tips of said pylons, which nacelles contain electric motors (E) and provided with propellers at their extremities, so that the rotation of the pylons results in turning the nacelles and thrust of propellers from substantially vertical, ensuring take-off and landing, to substantially horizontal providing a flight mode, retractable hydroskis for emergency landing on water, while said propulsion units and said wing are spaced apart vertically and horizontally and do not overlap.

The invention discloses a flying car vertical take-off and landing method, system and device based on attitude compensation and a medium, and the method comprises the steps: obtaining three-dimensional point cloud data and binocular parallax data of a take-off and landing region, and determining the terrain gradient information of the take-off and landing region according to the three-dimensional point cloud data and the binocular parallax data; vehicle body posture information of the hovercar is obtained, vehicle body posture changes in the future time period are predicted according to the vehicle body posture information and the terrain gradient information, and a first target thrust vector is determined according to the vehicle body posture changes; and performing multi-rotor power distribution on the first target thrust vector to obtain a first optimal thrust matrix, and performing attitude compensation on the hovercar according to the first optimal thrust matrix. The safety and reliability of vertical take-off and landing of the hovercar are enhanced, the dependence on ground conditions and the construction and maintenance cost of a take-off and landing area are reduced, the flexibility and adaptability of deployment of vertical take-off and landing in various environments are improved, and the method can be widely applied to the technical field of hovercars.

The invention relates to the technical field of flight control, in particular to a thrust vector type short-range vertical take-off and landing aircraft hovering control method and device, and the method comprises the steps: building a thrust vector type short-range vertical take-off and landing aircraft hovering dynamic model based on the nonlinear relation of the aircraft speed, the Euler angle, the angular speed, the thrust and the torque; designing a reconfiguration controller of a power system of the thrust vector type short-range vertical take-off and landing aircraft, and designing a hovering controller of the thrust vector type short-range vertical take-off and landing aircraft; designing an extended state observer used for estimating a nonlinear term and a state derivative in the hovering dynamics model; and feeding back the estimated nonlinear term and the estimated state derivative of the extended state observer to a hovering controller so as to carry out hovering control on the thrust vector type short-distance vertical take-off and landing aircraft. Therefore, the problem that efficient decoupling and robust control of multiple execution mechanisms in an aircraft system are difficult to realize by a dynamic inverse-based hovering control method in related technologies is solved.

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.

The invention discloses a plume baffle device and method for attitude control nozzle thrust non-contact measurement, and belongs to the field of engine test. The device comprises a plume baffle, a dynamic indirect force transducer and a mounting bracket. The mounting bracket is fixed on one side of the multi-component test bed, the plume baffle is coaxially and fixedly connected with the dynamic indirect force transducer, the inclined surface of the plume baffle faces the attitude and orbit control engine to be tested and is coaxial with the attitude control spray pipe, and the dynamic indirect force transducer is fixed on the mounting bracket. The plume baffle bears plume impact of the attitude control spray pipe, the dynamic indirect force measurement sensor measures the impact force borne by the plume baffle, and a measurement result is fed back to the upper computer for processing. The device is used in cooperation with a multi-component thrust vector test system, and by establishing the corresponding relation between the thrust of the attitude control spray pipe and the impact force of the plume baffle, the test problems that in the thrust vector test of the solid attitude and orbit control engine, the thrust of the orbit control spray pipe and the thrust of the attitude control spray pipe influence each other, and the thrust test error of the attitude control spray pipe is large are solved.

The invention provides a vertical take-off and landing aviation aircraft driven by tandem double vectors, which belongs to the technical field of aviation aircrafts and comprises an aircraft body, rotors, front vector drivers and rear vector drivers. The rotor wing is rotationally connected to the top end of the middle of the fuselage through a rotor wing shaft, and the front vector driver and the rear vector driver are rotationally connected to the front portion and the rear portion of the fuselage through mounting shafts perpendicular to the horizontal mounting face of the fuselage correspondingly and located below the rotating plane of the rotor wing; the center line of the mounting shaft and the center line of the rotor shaft are positioned in the longitudinal vertical symmetry plane of the fuselage; thrust vectors generated by the front vector driver and the rear vector driver balance reverse torque brought by rotation of the rotor wings in the hovering / vertical lifting and cruising stages of the aircraft, and the aircraft is pushed to fly. The aircraft is simple and easy to implement, high in safety, low in cost, long in endurance time and large in load capacity.

The invention relates to a gas vane thrust vectoring device servo driving scheme of a plug nozzle engine, which is specifically divided into a gas vane, a fixing device of the gas vane and an integrated electro-hydraulic servo control system, so that the gas vane is fixed on the surface of a plug cone, and the gas vane can be controlled to act at a specified angle according to requirements. In order to enable the gas vanes to generate large axial force and to have small axial loss at the same time, the gas vanes are evenly arranged at the 50% length position of the plug cone in the circumferential direction. In order to solve the problems of feasibility and reliability of axial fastening of the gas rudder shaft on the curved surface of the plug cone, a split type design and dual-bearing fastening are used. In order to guarantee the transmission rotation stability and accuracy of the gas vane, a servo motor is used for driving a hydraulic cylinder, and a double-rocker mechanism is formed to achieve driving. In order to ensure the flexibility and reliability of the thrust vector of the engine, a unit integrated electro-hydraulic servo system is adopted, so that each gas vane can independently realize vector control. The invention has the advantages of high structural reliability, high control accuracy and high vector maneuverability.

The invention relates to the field of unmanned aerial vehicles, and discloses a tilting type composite thrust system of a small fixed-wing aircraft, comprising: a physical bearing and power unit module comprising wings and six groups of tilting thrust units on the wings, each group of unit being composed of a turboprop engine and a variable pitch propeller and generating thrust; the thrust vector adjusting module comprises six sets of electric tilting joints and drives the tilting thrust unit to rotate so as to adjust the thrust direction. The state sensing and feedback module is provided with an angle sensor for monitoring the tilting angle in real time; the cooperation and synchronization control module comprises a synchronization controller which is connected with the angle sensor and the electric tilting joints and controls the six sets of electric tilting joints to be converted in batches according to the monitored angle. According to the invention, the integrated electric tilting joint is arranged, and the joint contains a main tilting shaft for wide-angle mode switching and an auxiliary fine tuning shaft for small-range active compensation, so that two adjusting mechanisms with different functions are integrated in a single structure.

The invention relates to the technical field of aircrafts, and discloses a balanced multi-rotor aircraft which comprises a framework structure, a control part and an energy storage part are arranged on the framework structure, the framework structure comprises a supporting part, a flight arm is rotationally connected to the supporting part, a rotor part is arranged at the free end of the flight arm, and the control part comprises a connecting rod mechanism installed on the supporting part. The connecting rod mechanisms are used for controlling the rotation angles of the rotor wing parts around the axes of the flight arms, and when the two rotor wing parts located on the same axis rotate around the axes in the same direction, the aircraft translates in a horizontal posture; according to the unmanned aerial vehicle, the flying arms are arranged by using a plus sign or cross-shaped frame, the rotor wing parts incline and enable the aircraft to move forwards or backwards when the flying arms are rotated, the rotor wing parts which are not inclined can concentrate on stability, pitching and yawing of the main body fuselage cannot be caused, and yawing can be achieved by changing the rotating speed or thrust vector of the rotor wing parts in the same direction.

The application discloses a shaft-symmetrical rocket-based combined engine with a variable geometry throat, which comprises a combustion chamber and a square flow channel section in communication in an axial direction from front to back; the shell of the combustion chamber is surrounded by a bottom support bottom plate and a top plate, the bottom support bottom plate is horizontal, and the top plate is a semicircular arc shape which is arched upward; the square flow channel section is a cuboid cavity, and an upper variable structure geometry throat and a lower variable structure geometry throat which are consistent with the size of corresponding holes are embedded in the holes at the upper part and the lower part; the upper variable structure geometry throat and the lower variable structure geometry throat are both plate bodies in the shape of an inclined wedge, the tip of the inclined wedge faces the rear, and the inclined surface sides of the inclined wedge both face the cavity; the upper variable structure geometry throat and the lower variable structure geometry throat are used for changing the flow channel geometry throat area in the square flow channel section. The engine is matched with the flight state by reducing the throat area with the increase of the flight Mach number, the engine performance is improved, and the engine thrust vector adjustment can be realized.

The invention belongs to the technical field of aero-engines, and particularly relates to a thrust vectoring nozzle actuator cylinder telescopic mechanical adjusting tool. Comprising a force-bearing rod, one end of the force-bearing rod is a threaded section with an external thread, and the other end of the force-bearing rod is provided with a clamping jaw; the movable bearing sleeve is sleeved on the inner side of the threaded section of the bearing rod in a sliding manner and is provided with a clamping jaw; the contraction nut is mounted on the threaded section; the two clamping jaws act on the piston rod of the actuating cylinder and the fixed end of the actuating cylinder respectively, the adjusting nut is screwed in to enable the movable force bearing sleeve to move towards the other end, and the two clamping jaws enable the piston rod to retract into the actuating cylinder. The expansion rod tool and the contraction rod tool are similar in structure, the two clamping jaws act on a piston rod of the actuator cylinder and the fixed end of the actuator cylinder respectively, the expansion nut is screwed out to enable the right supporting rod to move in the direction away from the left supporting rod, and the two clamping jaws enable the piston rod to stretch out.

The present invention relates to a method for determining the amount of non-isobaric air cushion propellant filling in parallel tanks and the satellite mass center, and the steps include: obtaining input parameters through experiments, tests and simulations; carrying out balanced filling to obtain the propellant filling mass, propellant mass center and satellite mass center coordinates of each tank in the launch state during balanced filling; calculating the mass center adjustment capacity during unbalanced filling, and sequentially obtaining the propellant filling mass, propellant mass center and satellite mass center coordinates of each tank when each tank is filled according to the maximum filling mass; calculating the intersection of the engine thrust vector and the XY plane at the satellite mass center; based on the determined propellant filling amount, air cushion pressure and satellite mass center, performing propellant and helium filling and subsequent on-orbit operations. The present invention introduces a satellite mass center relationship diagram, which can vividly display the positional relationship between the satellite mass center and the constraint conditions at different times, which is easy to understand and quickly guide engineering applications.

The invention relates to a ball thrust vector nozzle (BTN) system, which is a kind of tail rudder system, especially in liquid-fuelled rockets, for steering a missile, rocket or any thrusted aircraft by directing the exhaust gas output.

In the aerospace industry, the Laval nozzle is primarily used. This nozzle has the problem of not being able to adapt to changes in ambient pressure. Therefore, it loses efficiency as the vehicle ascends. Aerospike nozzles solve this problem, but are more difficult to cool, so they are not a viable alternative to the Laval nozzle. By means of a fixed arrangement of several Laval nozzles, consisting of nozzles (16) directed parallel to the object's flight direction and several nozzles (17) directed at an angle to this jet, the nozzle assembly creates an effect similar to an aerospike. Therefore, at decreasing pressure, such a nozzle assembly exhibits greater efficiency than conventional Laval nozzles. Since the assembly consists entirely of Laval nozzles, it is no more difficult to manufacture and cool than these. Furthermore, the thrust vector exerted by the nozzle assembly, and thus the direction of motion of the object being accelerated, can be controlled by throttling certain secondary nozzles (e.g., 17.2). The resulting fixed connection between the nozzles and combustion chambers allows for higher operating pressure within the combustion chambers, allowing for more efficient fuel use. Since the fixed connection allows for better thrust transfer, the thrust vector can be changed more radically and quickly.

The invention discloses a solid attitude and orbit control engine thrust vector in-situ calibration device and calibration method. The device is characterized by comprising a rack, a gantry type calibration frame, a standard force source mechanism and a force measuring system, the gantry type calibration frame is fixed above the rack, the standard force source mechanism is fixed on the gantry type calibration frame, and the force measuring system is arranged below the gantry type calibration frame and fixed on the rack; the standard force source mechanism comprises a transverse force generating unit and a vertical force generating unit; the force measuring system comprises a piezoelectric force measuring table, an adapter plate and a solid attitude and orbit control simulation engine, the piezoelectric force measuring table is fixed on the rack, and the solid attitude and orbit control simulation engine is mechanically positioned on the piezoelectric force measuring table through the adapter plate. The structure is simple, the loading position is flexibly adjusted, the influence of the zero drift of the piezoelectric sensor and the charge amplifier on the calibration precision is effectively reduced, and the thrust vector test precision of the solid attitude and orbit control engine is further improved.

The application provides a vertical take-off and landing aircraft driven by longitudinal double vectors, and belongs to the technical field of aircrafts. The aircraft comprises a fuselage, a rotor, a front vector driver and a rear vector driver. The rotor is rotatably connected to the middle top end of the fuselage through a rotor shaft. The front vector driver and the rear vector driver are rotatably connected to the front part of the fuselage and the rear part of the fuselage respectively through installation shafts which are perpendicular to the horizontal installation surface of the fuselage. The installation shafts and the rotor shaft are located in the longitudinal vertical symmetry plane of the fuselage. The thrust vectors generated by the front vector driver and the rear vector driver balance the reverse torque caused by the rotation of the rotor and drive the aircraft to fly in the hovering / vertical lifting and cruising stages. The aircraft has the advantages of simple and feasible implementation, high safety, low cost, long endurance time and large load capacity.

The invention discloses a novel turbomachinery vector sealing structure. A vector sealing cavity is formed between a stator shell containing sealing teeth and a rotor. The vector sealing cavity is of a spray pipe type structure in mirror symmetry, and the spray pipe type structure comprises multiple stages of sealing cavities which are sequentially connected end to end and are provided with double throat parts; the sealing cavity is sequentially divided into an expansion section and a contraction section in the airflow flowing direction, and a tooth tip jet flow bypass is arranged between the contraction section and the throat part adjacent to the downstream of the contraction section; the sealing cavity is designed to be of a thrust vectoring nozzle structure with a bypass channel, by means of the tooth cavity pressure difference generated by the throttling effect, bypass jet flow is spontaneously driven to be mixed with main leakage flow, a leakage path is guided to directionally deflect in the circumferential direction and the radial direction, multi-dimensional reconstruction of a sealing flow field is achieved, and the dual purposes of reducing leakage and restraining fluid excitation are achieved. And a new method is provided for reducing sealing leakage and improving the excitation suppression performance of sealing fluid.

The invention discloses a thrust vector compensation control method suitable for a nonlinear friction load, which adopts a current positive feedback compensation correction network, combines low-pass filtering with proportional gain, takes iq current as input and takes a compensation instruction delta r as output, and is simple, convenient and effective. The problems of amplitude attenuation and phase lag caused by inherent nonlinear load characteristics of the nozzle are solved, and the tracking capability of an electromechanical servo system to a control instruction is improved. The algorithm complexity is low, parameter adjustment is simple and easy to operate, and parameter adjustment time is saved.

This invention discloses an aerodynamic optimization design method for scaled-down model blades of large vertical axis wind turbines based on DMST theory. The method includes: determining scaled-down parameters according to wind tunnel similarity criteria; selecting an optimized airfoil for aerodynamic characteristic matching at low Reynolds numbers based on the lift-to-drag ratio characteristics of the prototype blades; constructing an aerodynamic performance model based on dual-disk multi-flow theory, calculating and verifying the theoretical thrust coefficients of the prototype and optimized airfoils; conducting force-pressure tests in a wind tunnel, extracting the pure thrust of the blades by subtracting the total thrust of the turbine from the tower thrust vector; and comparing the measured thrust coefficients with the theoretical thrust coefficients to verify the aerodynamic load matching degree. This invention effectively eliminates the Reynolds number effect in scaled-down tests by combining airfoil optimization, DMST theory verification, and wind tunnel test verification, achieving high-precision matching of aerodynamic loads on scaled-down model blades of vertical axis wind turbines and improving the reliability of wind tunnel test data.

A heat exchanger configured to perform heat exchange for a heat source of a fuselage and provide thrust or thrust vectoring to the fuselage, and an aircraft including the same are disclosed. The aircraft includes a fuselage, at least one inlet provided at a side of the fuselage, at least one outlet provided at a bottom of the fuselage, and a fan unit having at least one fan disposed between the inlet and the outlet along a fluid flow direction and drawing air inward through the at least one inlet to discharge through the at least one outlet.

The application relates to a force couple balance and longitudinal array push-pull vertical take-off and landing high-speed aircraft, belonging to the technical field of aircrafts, which comprises a fuselage, a rotor, a fixed wing, a steering controller and a propeller. The rotor is installed at the top of the middle part of the fuselage through a rotor shaft, and a clutch is arranged between the rotor and a power system and / or a transmission system matched with the fuselage, so as to be controlled to be disconnected or to transmit power to the rotor. The steering controller comprises a fixed base and a rotating base, and each of the nose and the tail has a set of the rotating base. The propeller is provided with two propellers which are respectively installed on the rotating bases of the nose and the tail, and are controlled to rotate with the rotating bases relative to the fixed base. The force couple formed by the front and rear thrust vectors is used to balance the reverse torque brought by the rotor. When the propeller is coaxial with the middle axis of the fuselage, the wind area of the single propeller obtains double thrust. The application can improve the energy utilization rate, improve the carrying capacity, improve the flight safety performance and obtain a longer hovering time.

The application provides a ball socket, pull rod and pin shaft friction and wear test bench in a thrust vector nozzle adjusting mechanism and a test method thereof, belonging to the technical field of aero-engine testing, and comprising: a first driving device, a ball socket mechanism, a force adding device, a pressure sensor, an ear seat mechanism, a second driving device, a third driving device and a sensor assembly arranged on a workbench; the ball socket mechanism comprises a ball socket support seat, a ball socket shaft, a first pull rod, a ball socket, a first ceramic heating sheet and two second ceramic heating sheets; the method can truly reproduce the combined space movement of the ball socket and the pull rod and the pull rod and the pin shaft under heavy load; the independent or cooperative temperature rising and accurate temperature control of the ball socket, the first pull rod and the second pull rod can simulate the non-uniform high-temperature working condition of the temperature gradient existing in different parts in a real nozzle.