Electric vertical take-off and landing aircraft and control method thereof

Abstract

The invention belongs to the technical field of aircrafts, and discloses an electric vertical take-off and landing aircraft and a control method thereof. According to the electric vertical take-off and landing aircraft, a left front rotor wing device and a right front rotor wing device are installed on the front portions of wings on the two sides of an aircraft body correspondingly and are in bilateral symmetry about the aircraft body; the left middle rotor wing device and the right middle rotor wing device are respectively mounted at the end parts of two sides of wings on two sides of the fuselage and are bilaterally symmetrical about the fuselage; the left rear rotor wing device and the right rear rotor wing device are respectively mounted at the end parts of two sides of the empennages on two sides of the fuselage and are bilaterally symmetrical about the fuselage; the left front rotor wing device, the right front rotor wing device, the left rear rotor wing device and the right rear rotor wing device adopt variable-rotating-speed or variable-total-pitch structures; and the left middle rotor device and the right middle rotor device adopt an automatic inclinator to carry out total pitch change and periodic pitch change. According to the technical scheme provided by the invention, the flight quality and the flight performance of the electric vertical take-off and landing aircraft can be effectively improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft, and in particular relates to an electric vertical take-off and landing aircraft and a control method thereof. Background Art

[0002] With the development and progress of battery, motor and electronic control technology, electric vertical take-off and landing aircraft, as a greener and safer low-altitude aircraft, has been greatly promoted. According to forecasts, by 2050, the economic scale of the urban air traffic industry will reach 100 million US dollars.

[0003] In recent years, stimulated by the low-altitude economic policy, many domestic and foreign enterprises / institutions have launched a variety of electric vertical take-off and landing aircraft, for example, there are distributed multi-rotor, composite wing, tilt-rotor and other configurations. With the continuous deepening of experimental research and application, the characteristics of the above-mentioned existing configurations have gradually been discovered; among them, as the take-off weight of electric vertical take-off and landing aircraft continues to increase, the deeper application problems of the existing configurations are constantly exposed, resulting in the slow advancement of its urban application. Specifically, as the take-off weight continues to increase, the distributed multi-rotor configuration responds less and less to attitude control through the differential speed of different rotors, and the robustness control is poor under gust interference; the composite wing configuration uses vertical multi-rotor and fixed-wing propeller, and carries the weight of fixed-wing propeller electric drive when vertically taking off, and carries the weight of vertical multi-rotor electric drive when cruising, which limits the full play of the flight performance of electric vertical take-off and landing aircraft; although the tilt-rotor configuration can greatly improve its flight performance, its mechanism design is complex and the tilt-rotation transition flight control technology is difficult.

[0004] In summary, in order to adapt to the demand for gradually increasing take-off weight and rapidly promote urban applications, a new type of electric vertical take-off and landing aircraft is urgently needed. Summary of the invention

[0005] The purpose of the present invention is to provide an electric vertical take-off and landing aircraft and a control method thereof to solve one or more of the above-mentioned technical problems. The technical solution provided by the present invention can effectively improve the flight quality and flight performance of the electric vertical take-off and landing aircraft, can adapt to the demand for gradually increasing take-off weight, and can enhance the practicality of the electric vertical take-off and landing aircraft in low-altitude urban flight.

[0006] In order to achieve the above object, the present invention adopts the following technical solutions: In a first aspect, the present invention provides an electric vertical take-off and landing aircraft, comprising: a structural platform, a rotor system, and an airborne system; The structural platform includes a fuselage, and wings, a tail wing and a landing gear mounted on the fuselage; The rotor system comprises a left front rotor device, a right front rotor device, a left middle rotor device, a right middle rotor device, a left rear rotor device and a right rear rotor device; the left front rotor device and the right front rotor device are respectively mounted on the front parts of the wings on both sides of the fuselage, and are symmetrical about the fuselage; the left middle rotor device and the right middle rotor device are respectively mounted on the two side ends of the wings on both sides of the fuselage, and are symmetrical about the fuselage; the left rear rotor device and the right rear rotor device are respectively mounted on the two side ends of the tail wings on both sides of the fuselage, and are symmetrical about the fuselage; wherein, the left front rotor device, the right front rotor device, the left rear rotor device and the right rear rotor device adopt a variable speed or variable collective pitch configuration; the left middle rotor device and the right middle rotor device adopt an automatic tilt device to change the collective pitch and cyclic pitch; The airborne system includes a power system and an energy system; wherein the power system is used to receive energy from the energy system and drive the rotor system.

[0007] A further improvement of the present invention is that The airborne system also includes one or more of an avionics system, a flight control system, a lighting system, and an environmental control system; The avionics system includes a sensing system and a communication system; wherein the sensing system includes one or more of air pressure, temperature, altitude, speed, attitude, position, heading, infrared and radiation sensors; The lighting system includes one or both of a high-voltage lighting device and a low-voltage lighting device.

[0008] A further improvement of the present invention is that The left front rotor device and the right front rotor device are respectively installed at the front parts of the wings on both sides of the fuselage, and are symmetrical about the fuselage. Specifically, the left front rotor device and the right front rotor device are respectively installed at the front parts of the wings on both sides of the fuselage through longitudinal components, and are symmetrical about the fuselage.

[0009] A further improvement of the present invention is that The left middle rotor device and the right middle rotor device are respectively installed on the two side ends of the wings on both sides of the fuselage, and are symmetrical about the fuselage. Specifically, the left middle rotor device and the right middle rotor device are respectively installed on the two side ends of the wings on both sides of the fuselage through lateral components, and are symmetrical about the fuselage.

[0010] A second aspect of the present invention provides a control method for an electric vertical take-off and landing aircraft, comprising: When the aircraft is in normal flight, the rotor speeds or collective pitches of the left front rotor device and the right rear rotor device are equal and rotate in opposite directions, the rotor speeds or collective pitches of the right front rotor device and the left rear rotor device are equal and rotate in opposite directions, and the rotor speeds or collective pitches of the left middle rotor device and the right middle rotor device are equal and rotate in opposite directions; When the aircraft performs nose-down pitching motion, the moment of force generated by the front rotor system on the center of mass of the entire aircraft is smaller than the moment of force generated by the rear rotor system on the center of mass of the entire aircraft, and the direction of the resultant moment points to the left side of the transverse section of the entire aircraft; wherein the front rotor system includes a left front rotor device and a right front rotor device, and the rear rotor system includes a left rear rotor device and a right rear rotor device; When the aircraft performs a pitching motion, the moment of force generated by the front rotor system on the center of mass of the entire aircraft is greater than the moment of force generated by the rear rotor system on the center of mass of the entire aircraft, and the direction of the resultant moment points to the right side of the lateral section of the entire aircraft; Among them, when the aircraft performs a nose-down pitch motion, the rotational speed or collective pitch of the left front rotor device and the right front rotor device is less than that of the left rear rotor device and the right rear rotor device; or, the left middle rotor device and the right middle rotor device perform forward cyclical pitch change; when the aircraft performs a nose-up pitch motion, the rotational speed or collective pitch of the left front rotor device and the right front rotor device is greater than that of the left rear rotor device and the right rear rotor device; or, the left middle rotor device and the right middle rotor device perform backward cyclical pitch change.

[0011] A further improvement of the present invention is that it comprises: When the aircraft performs a right rolling motion, the moment of the left rotor system on the center of mass of the whole aircraft is greater than the moment of the right rotor system on the center of mass of the whole aircraft, and the direction of the resultant moment points to the front side of the longitudinal section of the whole aircraft; wherein, the left rotor system includes a left front rotor device, a left middle rotor device and a left rear rotor device, and the right rotor system includes a right front rotor device, a right middle rotor device and a right rear rotor device; When the aircraft is rolling to the left, the moment of force generated by the left rotor system on the center of mass of the whole aircraft is smaller than the moment of force generated by the right rotor system on the center of mass of the whole aircraft, and the direction of the resultant moment points to the rear side of the longitudinal section of the whole aircraft; Among them, when the aircraft performs a right rolling motion, the rotation speed or total pitch of the left front rotor device and the left rear rotor device is greater than that of the right front rotor device and the right rear rotor device; or, the left middle rotor device and the right middle rotor device perform cyclic pitch change to the right; when the aircraft performs a left rolling motion, the rotation speed or total pitch of the left front rotor device and the left rear rotor device is less than that of the right front rotor device and the right rear rotor device; or, the left middle rotor device and the right middle rotor device perform cyclic pitch change to the left.

[0012] A further improvement of the present invention is that it comprises: When the aircraft performs a right yaw motion, the torque on the center of mass of the entire aircraft generated by the left rotor system is greater than the torque on the center of mass of the entire aircraft generated by the right rotor system, and the direction of the resultant torque remains vertically downward; When the aircraft performs a left yaw motion, the torque on the center of mass of the entire aircraft generated by the left rotor system is smaller than the torque on the center of mass of the entire aircraft generated by the left rotor system of the right rotor system, and the direction of the resultant torque remains vertically upward; Among them, when the aircraft performs a yaw motion to the right, the left front rotor device and the right rear rotor device form a vertical downward torque on the speed difference between the right front rotor device and the left rear rotor device; or, the left middle rotor device performs a forward cyclic pitch change, and the right middle rotor device performs a backward cyclic pitch change; when the aircraft performs a yaw motion to the left, the left front rotor device and the right rear rotor device form a vertical upward torque on the speed difference between the right front rotor device and the left rear rotor device; or, the left middle rotor device performs a backward cyclic pitch change, and the right middle rotor device performs a forward cyclic pitch change.

[0013] A further improvement of the present invention is that The control method includes a normal control procedure; Among them, for control requirements with a response speed requirement less than or equal to a preset threshold, the left front rotor device, the right front rotor device, the left middle rotor device, the right middle rotor device, the left rear rotor device and the right rear rotor device are subjected to an average force to balance the attitude of the aircraft; for control requirements with a response speed requirement greater than a preset threshold, the pitch motion adopts a control method in which the left front rotor device, the right front rotor device, the left rear rotor device and the right rear rotor device are mainly used, and the left middle rotor device and the right middle rotor device are auxiliary; the roll and yaw motions adopt a control method in which the left middle rotor device and the right middle rotor device are mainly used, and the left front rotor device, the right front rotor device, the left rear rotor device and the right rear rotor device are auxiliary.

[0014] A further improvement of the present invention is that The control method also includes an emergency control procedure; Among them, when any one of the left front rotor device, the right front rotor device, the left rear rotor device and the right rear rotor device fails, the rotor device on the opposite side of the failed rotor device stops rotating, and the left middle rotor device and the right middle rotor device increase the total pitch to increase the thrust output and perform maneuvering balance control.

[0015] A further improvement of the present invention is that The control method also includes an emergency control procedure; Among them, when any one of the left middle rotor device and the right middle rotor device fails, the rotor device that has not failed stops rotating, and the left front rotor device, the right front rotor device, the left rear rotor device and the right rear rotor device increase the total pitch or rotation speed to increase the thrust output and perform maneuvering balance control.

[0016] Compared with the prior art, the present invention has the following beneficial effects: The electric vertical take-off and landing aircraft provided by the present invention can effectively improve the flight quality and flight performance of the electric vertical take-off and landing aircraft, can adapt to the demand for gradually increasing take-off weight, and can enhance the practicality of the electric vertical take-off and landing aircraft in low-altitude flight in cities; among them, the present invention adopts an integrated design scheme that integrates six sets of rotor devices with variable speed or variable collective pitch and variable period and variable pitch, which can effectively improve the low-altitude flight stability of the aircraft, and can effectively cope with severe wind shear or turbulence problems at low altitudes; in addition, the present invention adopts an integrated design scheme that integrates six sets of rotor devices with variable speed or variable collective pitch and variable period and variable pitch, which can effectively improve the overall flight performance of the aircraft, and can also effectively improve the reliability and redundancy of the aircraft.

[0017] In the control method of the present invention, the stable flight of the aircraft is ensured by optimizing the control strategy; for example, in pitch, roll and yaw movements, different main and auxiliary control methods are adopted according to specific needs to ensure that the aircraft can adjust its attitude quickly and accurately; at the same time, by adjusting the parameters such as the rotation speed and collective pitch of the rotor system, the precise control of the aircraft can be achieved to adapt to the flight requirements under different take-off weights. Specifically, the present invention can quickly respond to changes in wind speed and wind direction in a low-altitude environment by dynamically adjusting the rotation speed of the rotor, especially when encountering wind shear, by increasing or decreasing the rotor speed to balance the external wind force and keep the aircraft attitude stable; combined with variable speed, by adjusting the collective inclination angle (i.e., collective pitch) of the rotor blades, the ratio of lift to drag can be optimized to ensure that a stable flight altitude and speed can be maintained under complex meteorological conditions; the variable period and variable pitch technology is adopted to allow the rotor to independently adjust its periodicity and the relative angle between the blades, so as to finely control the flight direction, pitch, roll and other actions, and maintain precise control even in turbulence. In summary, the combination of variable speed and variable collective pitch technology enables the aircraft to flexibly adjust power output according to the requirements of the flight mission. It can provide sufficient power when high-speed flight is required, and reduce energy consumption and extend flight time when cruising or hovering. The application of variable period and variable pitch technology gives the aircraft higher maneuverability and flexibility, and can quickly perform complex flight maneuvers, such as rapid turns, climbs and descents, which is crucial for performing reconnaissance, rescue and other missions.

[0018] In the present invention, the integrated design of six rotor devices not only provides powerful power support, but also builds a multiple redundant system. Even if some rotor devices fail, the remaining rotor devices can adjust their working status to ensure that the aircraft continues to fly stably, greatly reducing the risk of flight accidents caused by a single failure. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions in the present invention or the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below; obviously, the drawings described below are some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0020] Figure 1 is a schematic diagram of an electric vertical take-off and landing aircraft in an embodiment of the present invention; Figure 2 is a schematic diagram of a normal program control strategy in an embodiment of the present invention; Figure 3 is a schematic diagram of an emergency procedure control strategy in an embodiment of the present invention; The explanation of the reference numerals in the figures is as follows: 1. Left front rotor device; 2. Right front rotor device; 3. Left middle rotor device; 4. Right middle rotor device; 5. Left rear rotor device; 6. Right rear rotor device; 7. Airborne system; 8. Structural platform. DETAILED DESCRIPTION

[0021] In order to make the purpose, technical solutions and advantages of the present invention clearer, the technical solutions in the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention; it is obvious that the described embodiments and technical solutions are only part of the embodiments of the present invention, not all of the embodiments.

[0022] All other embodiments obtained by those of ordinary skill in the art without creative work based on the technical solutions disclosed in the embodiments of the present invention belong to the scope of protection of the present invention. In addition, the terms "include" and "have" and any variations thereof are intended to cover non-exclusive inclusions. For example, a process, method, system, product or device including a series of steps or units is not necessarily limited to those steps or units clearly listed, but may include other steps or units that are not clearly listed or inherent to these processes, methods, products or devices.

[0023] See also Figure 1 An electric vertical take-off and landing aircraft provided in an embodiment of the present invention comprises: a structural platform 8, a rotor system and an airborne system 7; wherein, The structural platform 8 includes a fuselage, and wings, tail wing and landing gear installed on the fuselage; illustratively, the fuselage is provided with different mechanical installation interfaces of the wing, tail wing, landing gear and recording system, and the wing, tail wing and landing gear are installed on the fuselage through the installation interfaces of the wing, tail wing and landing gear respectively; wherein the fuselage is located in the middle of the structural platform 8; the wing is located at the front side of the structural platform 8 and is fixedly connected to the fuselage structure; the tail wing is located at the rear of the structural platform 8 and is fixedly connected to the fuselage structure; the landing gear is located at the lower end of the structural platform 8 and is fixedly connected to the fuselage structure by screwing; The rotor system includes a left front rotor device 1, a right front rotor device 2, a left middle rotor device 3, a right middle rotor device 4, a left rear rotor device 5 and a right rear rotor device 6; wherein the left front rotor device 1 and the right front rotor device 2 are respectively installed at the front of the wings on both sides of the fuselage through longitudinal components, and are symmetrical about the fuselage; the left middle rotor device 3 and the right middle rotor device 4 are respectively installed at the two side ends of the wings on both sides of the fuselage through lateral components, and are symmetrical about the fuselage; the left rear rotor device 5 and the right rear rotor device 6 are respectively installed at the two side ends of the tail wings on both sides of the fuselage, and are symmetrical about the fuselage; the left front rotor device 1, the right front rotor device 2, the left rear rotor device 5 and the right rear rotor device 6 adopt a variable speed or variable collective pitch configuration, and in a further preferred embodiment, a variable speed method is adopted; the left middle rotor device 3 and the right middle rotor device 4 adopt an automatic tilt device to change the collective pitch and cyclic pitch; The airborne system 7 includes a power system and an energy system; wherein, the power system is used to receive energy from the energy system and drive the rotor system; by way of example, the power system is used to provide power for the rotor system, and can be installed respectively under 6 sets of rotor devices, which receive energy provided by the energy system, drive the rotor to rotate through the rotor shaft, and provide lift for the entire aircraft; the energy system can be installed on a structural platform 8, and provide energy to the power system through a cable grid.

[0024] In the technical solution provided by the embodiment of the present invention, the flight quality and flight performance of the electric vertical take-off and landing aircraft are effectively improved through a sophisticated rotor system layout, and the demand for gradually increasing take-off weight is also adapted to; wherein, through a total of six sets of rotor devices on the front and rear sides and the middle left and right sides, three-dimensional attitude control (pitch, roll, yaw) of the entire aircraft is achieved. This multi-rotor layout provides more control redundancy and flexibility, so that the aircraft can maintain a stable attitude in various flight states, and has the advantage of attitude control flexibility; through differential or variable collective pitch control, the rotor system can accurately adjust the thrust, thereby achieving precise control of the aircraft. This control method not only has a fast response speed, but also has high control accuracy, which helps to improve the flight quality and performance of the aircraft; the layout of the six sets of rotor devices enables the load to be distributed, avoiding the problem of excessive force on a single point. As the take-off weight increases, this distributed load design can more effectively disperse the load, reduce the burden of a single rotor device, and improve the carrying capacity and flight stability of the aircraft.

[0025] To explain in further detail, the left front rotor device 1 and the right front rotor device 2 constitute the front rotor system, which are installed at the front of both sides of the wing and connected to the wing through longitudinal components (such as brackets or crossbeams). These rotors are usually used to provide lift for the aircraft during takeoff, landing and hovering, as well as to adjust the attitude. Since they are located at the front of the wing, it is easier to achieve the pitch control of the aircraft, that is, to control the up and down pitch movement of the aircraft around its lateral axis. The left middle rotor device 3 and the right middle rotor device 4 constitute the middle rotor system, which are installed at the ends of both sides of the wing and connected to the wing through lateral components (such as side panels or wingtips). These rotors provide the main lift when the aircraft is cruising and may participate in the steering control of the aircraft. Since they are located at the ends of the wing, they can more effectively utilize the lift generated by the wing while reducing the resistance of the aircraft; in addition, the middle rotor can also be used to adjust the rolling attitude of the aircraft during flight, that is, to control the left and right rolling of the aircraft around its longitudinal axis. The left rear rotor device 5 and the right rear rotor device 6 constitute the rear rotor system, which is installed on both sides of the tail of the aircraft through the tail wing. These rotors are usually used to provide thrust for the aircraft in the forward direction, and to perform yaw control, that is, to control the left and right steering of the aircraft around its vertical axis. The layout and design of the rear rotors are crucial to the stability and maneuverability of the aircraft, and they directly affect the flight direction and speed of the aircraft. In summary, in the technical solution of the embodiment of the present invention, the layout of this multi-rotor system combines the advantages of rotors in different positions, which can achieve more flexible flight control and higher flight performance.

[0026] In the embodiment of the present invention, the left front rotor device 1, the right front rotor device 2, the left rear rotor device 5 and the right rear rotor device 6 adopt a variable speed or variable collective pitch configuration, and the left middle rotor device 3 and the right middle rotor device 4 adopt an automatic tilt device for variable collective pitch and cyclic pitch change. The further explanation is as follows: variable speed technology refers to adjusting the lift generated by the rotor and the flight state of the aircraft by changing the speed of the rotor; variable collective pitch is to adjust the lift generated by the rotor and the flight state of the aircraft by changing the collective pitch (i.e., blade angle) of the rotor blades; the automatic tilt device is a special device in the helicopter for changing the tilt direction and blade inclination angle of the rotor. Through the operation of the pilot, the automatic tilt device can be tilted, thereby driving the periodic change of the blade pitch, generating asymmetry of the rotor aerodynamic force, and then changing the direction of the pull vector of the rotor, so as to realize the lateral and longitudinal flight control of the helicopter; cyclic pitch change refers to the periodic change of the blade pitch within one rotation. For the left front rotor device 1, the right front rotor device 2, the left rear rotor device 5 and the right rear rotor device 6, the variable speed method can more effectively adjust the lift generated by the rotor and the flight state of the aircraft, especially when it is necessary to respond quickly to flight instructions or face a complex flight environment, the variable speed method can provide a more flexible control method. For the left middle rotor device 3 and the right middle rotor device 4, the use of an automatic tilt device for variable collective pitch and cyclic pitch can achieve more accurate flight attitude control. The automatic tilt device can sense and respond to the pilot's control instructions, and achieve the helicopter's roll, pitch and yaw and other actions by adjusting the pitch and tilt direction of the blades. In summary, the technical solution of the embodiment of the present invention combines the advantages of the variable speed and variable collective pitch configuration and the automatic tilt device, providing an efficient and flexible means for flight control and attitude adjustment of the aircraft.

[0027] In a specific embodiment of the present invention, the pulling force range of each rotor device among the left front rotor device 1, the right front rotor device 2, the left middle rotor device 3, the right middle rotor device 4, the left rear rotor device 5 and the right rear rotor device 6 is 300~500kgf, and the rotor blade radius is 1.8~2.4m.

[0028] In a preferred embodiment of the present invention, the airborne system 7 also includes one or more of an avionics system, a flight control system, a lighting system, and an environmental control system; wherein, The avionics system includes a sensing system and a communication system; wherein the sensing system includes sensors such as air pressure, temperature, altitude, speed, attitude, position, heading, infrared, and radiation, which are respectively installed at different positions of the structural platform 8; the communication system includes modules such as Bluetooth, WiFi, 5G, and data link, which are respectively installed at different positions of the structural platform 8 for short-range, medium-range, long-range, and air networking communication and control; The flight control system includes a flight control computer, flight control software and a flight parameter recorder; The lighting system includes strong or weak current lighting, which can be installed on the wings, tail and belly of the aircraft to provide lighting for the entire aircraft; The environmental control system is installed in the fuselage structure to provide a regulating function for the internal ambient temperature of the fuselage.

[0029] The technical solution disclosed in the embodiment of the present invention can be applied to the field of low-altitude urban transportation, and can also be applied to the fields of manned travel and fire fighting in low-altitude cities.

[0030] See also Figure 2 and Figure 3 , a control method for an electric vertical take-off and landing aircraft provided by an embodiment of the present invention comprises the following steps: The front rotor system includes a left front rotor device 1 and a right front rotor device 2, the rear rotor system includes a left rear rotor device 5 and a right rear rotor device 6, the left rotor system includes a left front rotor device 1, a left middle rotor device 3 and a left rear rotor device 5, and the right rotor system includes a right front rotor device 2, a right middle rotor device 4 and a right rear rotor device 6; In the normal flight state, the rotor speeds or collective pitches of the left front rotor device 1 and the right rear rotor device 6, the right front rotor device 2 and the left rear rotor device 5, and the left middle rotor device 3 and the right middle rotor device 4 are equal, and the rotation directions are opposite; When the aircraft needs to perform a nose-down pitching motion, the front rotor system generates a torque on the center of mass of the entire aircraft, which is smaller than the torque generated by the rear rotor system, and the direction of the resultant torque points to the left side of the transverse section of the entire aircraft; in a further exemplary technical solution, the rotation speed or total pitch of the left front rotor device 1 and the right front rotor device 2 may be smaller than that of the left rear rotor device 5 and the right rear rotor device 6, or the left middle rotor device 3 and the right middle rotor device 4 may perform forward cyclic pitch change; When the aircraft needs to perform a nose-up and pitching motion, the front rotor system generates a torque on the center of mass of the entire aircraft, which is greater than the torque generated by the rear rotor system, and the direction of the resultant torque points to the right side of the transverse section of the entire aircraft; in a further exemplary technical solution, the rotation speed or total pitch of the left front rotor device 1 and the right front rotor device 2 may be greater than that of the left rear rotor device 5 and the right rear rotor device 6, or the left middle rotor device 3 and the right middle rotor device 4 may perform a backward cyclic pitch change; When the aircraft needs to perform a right rolling motion, the left rotor system generates a torque on the center of mass of the entire aircraft, which is greater than the torque generated by the right rotor system, and the direction of the resultant torque points to the front side of the longitudinal section of the entire aircraft; in a further exemplary technical solution, the rotation speed or total pitch of the left front rotor device 1 and the left rear rotor device 5 may be greater than that of the right front rotor device 2 and the right rear rotor device 6, or the left middle rotor device 3 and the right middle rotor device 4 perform rightward cyclic pitch change; When the aircraft needs to perform a left rolling motion, the left rotor system generates a torque on the center of mass of the entire aircraft, which is smaller than the torque generated by the right rotor system, and the direction of the resultant torque points to the rear side of the longitudinal section of the entire aircraft; in a further exemplary technical solution, the rotation speed or total pitch of the left front rotor device 1 and the left rear rotor device 5 may be smaller than that of the right front rotor device 2 and the right rear rotor device 6, or the left middle rotor device 3 and the right middle rotor device 4 may perform a left cyclic pitch change; When the aircraft needs to perform a right yaw motion, the left rotor system generates a torque on the center of mass of the entire aircraft, which is greater than the torque generated by the right rotor system, and the direction of the resultant torque remains vertically downward; in a further exemplary technical solution, the left front rotor device 1 and the right rear rotor device 6 may form a vertical downward torque due to the speed difference between the right front rotor device 2 and the left rear rotor device 5, or the left middle rotor device 3 performs a forward cyclic pitch change, and the right middle rotor device 4 performs a backward cyclic pitch change; When the aircraft needs to perform a left yaw motion, the left rotor system generates a torque on the center of mass of the entire aircraft, which is less than the torque generated by the right rotor system, and the direction of the resultant torque remains vertically upward; in a further exemplary technical solution, the left front rotor device 1 and the right rear rotor device 6 may form a vertical upward torque due to the speed difference between the right front rotor device 2 and the left rear rotor device 5, or the left middle rotor device 3 performs a backward cyclic pitch change, and the right middle rotor device 4 performs a forward cyclic pitch change; When the aircraft needs to fly horizontally forward, the torque generated by the front rotor system on the center of mass of the entire aircraft is equal to the torque generated by the rear rotor system, and the total pulling force generated by the six sets of rotor devices in the front, back, left, and middle should maintain a horizontal component and a vertical component equal to the gravity of the entire aircraft; When the aircraft needs to fly horizontally backward, the torque generated by the front rotor system on the center of mass of the entire aircraft is equal to the torque generated by the rear rotor system, and the total pulling force generated by the six sets of rotor devices in the front, back, left, and right should maintain a horizontal component and a vertical component equal to the gravity of the entire aircraft; When the aircraft needs to fly horizontally to the right, the torque generated by the left rotor system on the center of mass of the entire aircraft is equal to the torque generated by the right rotor system, and the total pulling force generated by the left and right rotor systems should maintain a horizontal component to the right, and the vertical component should be equal to the gravity of the entire aircraft; When the aircraft needs to fly horizontally to the left, the torque generated by the left rotor system on the center of mass of the entire aircraft is equal to the torque generated by the right rotor system, and the total pulling force generated by the left and right rotor systems should maintain a horizontal component to the left, and the vertical component should be equal to the gravity of the entire aircraft; When the aircraft needs to move vertically upward, the resultant moments in the three directions of the center of mass of the whole aircraft generated by the left and right rotor systems are all zero, and the total pulling force generated is in the vertical direction and is greater than the gravity of the whole aircraft; in a further exemplary technical solution, the left front rotor device 1 and the right rear rotor device 6, the right front rotor device 2 and the left rear rotor device 5 increase the rotation speed or the total pitch at the same time, or the left middle rotor device 3 and the right middle rotor device 4 increase the total pitch at the same time; When the aircraft needs to move vertically downward, the resultant moments in the three directions of the center of mass of the whole aircraft generated by the left and right rotor systems are all zero, and the total pulling force generated is in the vertical direction and is greater than the gravity of the whole aircraft; in a further exemplary technical solution, the left front rotor device 1 and the right rear rotor device 6, the right front rotor device 2 and the left rear rotor device 5 simultaneously reduce the rotation speed or the total pitch, or the left middle rotor device 3 and the right middle rotor device 4 simultaneously reduce the total pitch; When the aircraft needs to hover in the air, the resultant torques in the three directions on the center of mass of the entire aircraft generated by the left and right rotor systems are all zero, and the total pulling force generated is in the vertical direction and equal to the gravity of the entire aircraft; When the aircraft needs to perform a fixed-point right turn, the resultant torques in the longitudinal and transverse directions on the center of mass of the entire aircraft generated by the left and right rotor systems are zero, the resultant torque in the vertical direction is downward, and the total pulling force generated is equal to the gravity of the entire aircraft; in a further exemplary technical solution, the vertical downward torque may be formed by the speed difference between the left front rotor device 1 and the right rear rotor device 6 and the right front rotor device 2 and the left rear rotor device 5, or the left middle rotor device 3 performs forward cyclic pitch change and the right middle rotor device 4 performs backward cyclic pitch change; When the aircraft needs to make a fixed-point left turn, the resultant torque in the longitudinal and transverse directions on the center of mass of the entire aircraft generated by the left and right rotor systems is zero, the resultant torque in the vertical direction is upward, and the total pulling force generated is equal to the gravity of the entire aircraft; in a further exemplary technical solution, it can be a vertical upward torque formed by the speed difference between the left front rotor device 1 and the right rear rotor device 6 and the right front rotor device 2 and the left rear rotor device 5, or the left middle rotor device 3 performs backward cyclic pitch change, and the right middle rotor device 4 performs forward cyclic pitch change.

[0031] In a further preferred technical solution, in a normal control procedure, when the response speed requirement is not high, the six sets of rotor devices are subjected to an average force to balance the aircraft attitude; when the response speed requirement is high, the pitch motion adopts a control method in which the front and rear rotor systems are mainly controlled and the middle rotor system is supplemented, and the roll and yaw motions adopt a control method in which the middle rotor system is mainly controlled and the front and rear rotor systems are supplemented; In a further preferred technical scheme, in the emergency control procedure, when one of the four rotor devices on the front and rear sides fails, the rotor device on the opposite side thereof enters a failure protection mode, stops rotating and outputting thrust, and the middle rotor system increases the collective pitch, increases the thrust output, and performs control over the control balance; in addition, when one of the two middle left and right rotor devices fails, the rotor device on the middle opposite side enters a failure protection mode, stops rotating and outputting thrust, and the four front and rear rotor devices increase the collective pitch or rotation speed, increases the thrust output, and performs control over the control balance.

[0032] In the technical solution of the embodiment of the present invention, in the normal control procedure, for the case where the response speed requirement is not high, the six sets of rotor devices are subjected to an average force, and the attitude of the UAV aircraft is balanced and controlled, which helps to maintain the stable flight of the aircraft; when a quick response is required, the corresponding main and auxiliary control methods are adopted according to the specific movement type to ensure that the aircraft can quickly adjust its attitude to meet the flight requirements. In addition, the technical solution of the embodiment of the present invention also takes into account the failure of the rotor device and designs a corresponding failure protection mechanism. When a set of rotor systems fails, other rotor systems can quickly adjust the thrust output and perform maneuvering balance control to ensure the safe flight of the aircraft. This design improves the reliability and safety of the aircraft.

[0033] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the above embodiments, ordinary technicians in the relevant field should understand that the specific implementation methods of the present invention can still be modified or replaced by equivalents. Any modification or equivalent replacement that does not depart from the spirit and scope of the present invention should be covered within the scope of protection of the claims of the present invention.

Claims

1. An electric vertical take-off and landing aircraft, characterized in that: include: Structural platform (8), rotor system and airborne systems (7); The structural platform (8) comprises a fuselage, and wings, a tail wing and a landing gear mounted on the fuselage; The rotor system comprises a left front rotor device (1), a right front rotor device (2), a left middle rotor device (3), a right middle rotor device (4), a left rear rotor device (5) and a right rear rotor device (6); the left front rotor device (1) and the right front rotor device (2) are respectively mounted on the front parts of the wings on both sides of the fuselage, and are symmetrical with respect to the fuselage; the left middle rotor device (3) and the right middle rotor device (4) are respectively mounted on the two side ends of the wings on both sides of the fuselage, and are symmetrical with respect to the fuselage; the left rear rotor device (5) and the right rear rotor device (6) are respectively mounted on the two side ends of the tail wing on both sides of the fuselage, and are symmetrical with respect to the fuselage; wherein the left front rotor device (1), the right front rotor device (2), the left rear rotor device (5) and the right rear rotor device (6) adopt a variable speed or variable collective pitch configuration; the left middle rotor device (3) and the right middle rotor device (4) adopt an automatic tilt device to perform variable collective pitch and cyclic variable pitch; The airborne system (7) comprises a power system and an energy system; wherein the power system is used to receive energy from the energy system and drive the rotor system.

2. The electric vertical take-off and landing aircraft according to claim 1, characterized in that: The airborne system (7) further includes one or more of an avionics system, a flight control system, a lighting system, and an environmental control system; The avionics system includes a sensing system and a communication system; wherein the sensing system includes one or more of air pressure, temperature, altitude, speed, attitude, position, heading, infrared and radiation sensors; The lighting system includes one or both of a high-voltage lighting device and a low-voltage lighting device.

3. The electric vertical take-off and landing aircraft according to claim 1, characterized in that: The left front rotor device (1) and the right front rotor device (2) are respectively mounted on the front parts of the wings on both sides of the fuselage, and are symmetrical with respect to the left and right of the fuselage. Specifically, the left front rotor device (1) and the right front rotor device (2) are respectively mounted on the front parts of the wings on both sides of the fuselage through longitudinal components, and are symmetrical with respect to the left and right of the fuselage.

4. The electric vertical take-off and landing aircraft according to claim 1, characterized in that: The left middle rotor device (3) and the right middle rotor device (4) are respectively mounted on the two side ends of the wings on both sides of the fuselage, and are symmetrical with respect to the fuselage. Specifically, the left middle rotor device (3) and the right middle rotor device (4) are respectively mounted on the two side ends of the wings on both sides of the fuselage through lateral components, and are symmetrical with respect to the fuselage.

5. A control method for an electric vertical take-off and landing aircraft according to claim 1, characterized in that: include: When the aircraft is in a normal flight state, the rotor speeds or collective pitches of the left front rotor device (1) and the right rear rotor device (6) are equal and rotate in opposite directions, the rotor speeds or collective pitches of the right front rotor device (2) and the left rear rotor device (5) are equal and rotate in opposite directions, and the rotor speeds or collective pitches of the left middle rotor device (3) and the right middle rotor device (4) are equal and rotate in opposite directions; When the aircraft performs a nose-down pitching motion, the moment generated by the front rotor system on the center of mass of the entire aircraft is smaller than the moment generated by the rear rotor system on the center of mass of the entire aircraft, and the direction of the resultant moment points to the left side of the transverse section of the entire aircraft; wherein the front rotor system includes a left front rotor device (1) and a right front rotor device (2), and the rear rotor system includes a left rear rotor device (5) and a right rear rotor device (6); When the aircraft performs a pitching motion, the moment of force generated by the front rotor system on the center of mass of the entire aircraft is greater than the moment of force generated by the rear rotor system on the center of mass of the entire aircraft, and the direction of the resultant moment points to the right side of the lateral section of the entire aircraft; When the aircraft performs a nose-down pitching motion, the rotation speed or collective pitch of the left front rotor device (1) and the right front rotor device (2) is less than that of the left rear rotor device (5) and the right rear rotor device (6); or, the left middle rotor device (3) and the right middle rotor device (4) perform a forward cyclical pitch change; when the aircraft performs a nose-up pitching motion, the rotation speed or collective pitch of the left front rotor device (1) and the right front rotor device (2) is greater than that of the left rear rotor device (5) and the right rear rotor device (6); or, the left middle rotor device (3) and the right middle rotor device (4) perform a backward cyclical pitch change.

6. The control method of an electric vertical take-off and landing aircraft according to claim 5, characterized in that: include: When the aircraft performs a right rolling motion, the torque generated by the left rotor system on the center of mass of the entire aircraft is greater than the torque generated by the right rotor system on the center of mass of the entire aircraft, and the direction of the resultant torque points to the front side of the longitudinal section of the entire aircraft; wherein the left rotor system includes a left front rotor device (1), a left middle rotor device (3) and a left rear rotor device (5), and the right rotor system includes a right front rotor device (2), a right middle rotor device (4) and a right rear rotor device (6); When the aircraft is rolling to the left, the moment of force generated by the left rotor system on the center of mass of the whole aircraft is smaller than the moment of force generated by the right rotor system on the center of mass of the whole aircraft, and the direction of the resultant moment points to the rear side of the longitudinal section of the whole aircraft; When the aircraft performs a right rolling motion, the rotation speed or collective pitch of the left front rotor device (1) and the left rear rotor device (5) is greater than that of the right front rotor device (2) and the right rear rotor device (6); or, the left middle rotor device (3) and the right middle rotor device (4) perform a rightward cyclic pitch change; when the aircraft performs a left rolling motion, the rotation speed or collective pitch of the left front rotor device (1) and the left rear rotor device (5) is less than that of the right front rotor device (2) and the right rear rotor device (6); or, the left middle rotor device (3) and the right middle rotor device (4) perform a leftward cyclic pitch change.

7. The control method of an electric vertical take-off and landing aircraft according to claim 6, characterized in that: include: When the aircraft performs right yaw motion, the torque on the center of mass of the entire aircraft generated by the left rotor system is greater than the torque on the center of mass of the entire aircraft generated by the right rotor system, and the direction of the resultant torque remains vertically downward; When the aircraft performs a left yaw motion, the torque on the center of mass of the entire aircraft generated by the left rotor system is smaller than the torque on the center of mass of the entire aircraft generated by the left rotor system of the right rotor system, and the direction of the resultant torque remains vertically upward; When the aircraft performs a right yaw motion, the left front rotor device (1) and the right rear rotor device (6) generate a vertical downward torque on the speed difference between the right front rotor device (2) and the left rear rotor device (5); or, the left middle rotor device (3) performs a forward cyclic pitch change, and the right middle rotor device (4) performs a backward cyclic pitch change; when the aircraft performs a left yaw motion, the left front rotor device (1) and the right rear rotor device (6) generate a vertical upward torque on the speed difference between the right front rotor device (2) and the left rear rotor device (5); or, the left middle rotor device (3) performs a backward cyclic pitch change, and the right middle rotor device (4) performs a forward cyclic pitch change.

8. The control method of an electric vertical take-off and landing aircraft according to claim 5, characterized in that: The control method includes a normal control procedure; Wherein, for a control requirement with a response speed requirement less than or equal to a preset threshold, the left front rotor device (1), the right front rotor device (2), the left middle rotor device (3), the right middle rotor device (4), the left rear rotor device (5) and the right rear rotor device (6) are subjected to an average force, and the attitude of the aircraft is balanced and controlled; for a control requirement with a response speed requirement greater than the preset threshold, the pitch motion adopts a control method in which the left front rotor device (1), the right front rotor device (2), the left rear rotor device (5) and the right rear rotor device (6) are mainly controlled, and the left middle rotor device (3) and the right middle rotor device (4) are supplemented, and the roll and yaw motions adopt a control method in which the left middle rotor device (3) and the right middle rotor device (4) are mainly controlled, and the left front rotor device (1), the right front rotor device (2), the left rear rotor device (5) and the right rear rotor device (6) are supplemented.

9. The control method of an electric vertical take-off and landing aircraft according to claim 8, characterized in that: The control method also includes an emergency control procedure; When any one of the left front rotor device (1), the right front rotor device (2), the left rear rotor device (5) and the right rear rotor device (6) fails, the rotor device on the opposite side of the failed rotor device stops rotating, and the left middle rotor device (3) and the right middle rotor device (4) increase their collective pitch to increase the thrust output, thereby performing maneuvering balance control.

10. The control method of an electric vertical take-off and landing aircraft according to claim 8, characterized in that: The control method also includes an emergency control procedure; When any one of the left center rotor device (3) and the right center rotor device (4) fails, the remaining rotor device stops rotating, and the left front rotor device (1), the right front rotor device (2), the left rear rotor device (5) and the right rear rotor device (6) increase their collective pitch or rotation speed to increase the thrust output, thereby performing maneuvering balance control.

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