Multi-rotor integrated aircraft

Abstract

The utility model belongs to the technical field of aviation aircrafts, and discloses a multi-rotor integrated aircraft, which consists of an aircraft body, integrated wings, a flight control system and a power system, and at least two groups of tandem integrated wings are arranged on the aircraft body, so that the problems of vertical take-off and landing, low-speed flight and hovering of the aircraft can be solved like a multi-rotor aircraft. Multi-rotor flying and fixed-wing flying can be achieved by controlling the elevation angle of the integrated wings, endurance is improved by additionally arranging the large foldable fixed wings, economic performance of the integrated wings is improved, maneuverability of the integrated wings is improved by additionally arranging the vector propellers, and the integrated wings are designed to be of a foldable structure, so that storage and transportation are facilitated; through the design that a rudder is additionally arranged on each integrated wing, the aircraft can be controlled to ascend, descend, steer leftwards and rightwards, move forwards and backwards, translate leftwards and rightwards and the like under the condition that a single integrated wing is not provided with a power adapter.

Description

Technical Field

[0001] The utility model belongs to the technical field of aerospace aircraft, and particularly relates to a multi-rotor integrated aircraft. Background Art

[0002] Although traditional multi-rotor UAVs can achieve vertical takeoff and landing and hovering in the air, since there is no fixed wing to generate lift and they rely entirely on motor power to overcome their own weight, they do not make full use of the Bernoulli principle of aerodynamics to increase lift. Therefore, the efficiency ratio of most quad-rotor aircraft is only about 20%. Due to the high drag of the round tube support frame, it is difficult to achieve high-speed flight. Relying on battery energy storage with low density, the flight time is short, and the endurance and economy are not high. Fixed-wing aircraft, although they have improved the efficiency ratio, require an airport runway for takeoff and landing. Compound-wing aircraft, although they have the advantages of the above combination, most of the numerous and heavy multi-rotor motors are only used for a short time during vertical takeoff and landing and hovering. The rotor motors are fixed on several round tube supports, which not only increases the weight of the aircraft, but also rapidly increases the drag during high-speed flight, reduces the efficiency ratio, and greatly reduces the effective payload. The Osprey V22 tilt-rotor aircraft, although it has also solved the vertical and horizontal conversion functions of the rotor at the outer port of the fixed wing, has high requirements for the structural strength of the fixed wing, complex conversion machinery and electric servo structures, difficult production processes, high costs, large vibration amplitudes at the fixed wing ports, and serious destructive forces. At the same time, during the vertical takeoff stage, the downwash airflow of the propeller generates a large downward wind load on the adjacent fixed wing surface and the upper part of the fuselage, which also reduces the takeoff weight and effective load of the aircraft. Often, due to the airflow disorder caused by the conversion between the two working conditions, the overall instability of the fuselage occurs, and there have been many serious flight accidents of plane crashes and deaths.

[0003] In view of the above technical problems, there is an urgent need to design a multi-rotor aircraft with a high efficiency ratio and capable of vertical takeoff and landing, and to innovatively design a multi-rotor integrated aircraft and a flight control method, which can not only meet the requirements of vertical takeoff and landing, low-speed flight, and hovering in the air, but also achieve high-speed flight similar to fixed-wing aircraft without a runway. It is also necessary to solve the smooth conversion between the rotor and fixed-wing flight states, so that the rotor propeller can also play a role in the fixed-wing flight state, and can overcome the problem of the downwash airflow generating a downward load on the wing or the upper part of the fuselage under the rotor working conditions, improve the safety and reliability of flight, etc. At the same time, the economy, convenience of storage and transportation, and landing gear configuration suitable for different environments are considered, as well as the optimized power distribution system and servo control, and emergency safety system protection measures to ensure good stability and controllability in various flight states. Content of the Utility Model

[0004] To solve the problems existing in the background art, the utility model provides a multi-rotor integrated aircraft, which includes a fuselage, integrated wings, a flight control system, and a power system, wherein:

[0005] The integrated wing is a unified component integrating a control wing, a thruster, a rudder, and a control wing transmission shaft interface. The entire elevation angle α of the integrated wing is controlled by an elevator servo through the control wing transmission shaft; the rudder angle β of the rudder on each integrated wing is controlled by a rudder servo; the thruster and the rudder are both installed on the control wing;

[0006] The elevator servo is connected to the control wing transmission shaft, and the control wing transmission shaft is connected to the control wing through the control wing transmission shaft interface;

[0007] At least one control wing, one thruster, and one rudder are provided on each integrated wing; each integrated wing is controlled by at least one elevator servo;

[0008] Two integrated wings symmetrically arranged left and right along the fuselage are a group. The thrusters of the two integrated wings symmetrically arranged left and right have the same specifications and generate torques in opposite directions. The elevator servos of the two integrated wings symmetrically arranged left and right have the same specifications;

[0009] At least two groups of integrated wings are provided front and back on the fuselage; the torques generated by the thrusters on the diagonals of each two groups of integrated wings are in the same direction and have the same power.

[0010] In a preferred solution, the rudder is a bidirectional symmetric rudder surface structure provided in the center of the thruster air flow in the forward direction and based on the control wing.

[0011] In a preferred solution, the integrated wing is provided with a wing folding mechanism.

[0012] In a preferred solution, a fixed wing is provided on the fuselage.

[0013] In a preferred solution, a helicopter rotor structure is provided on the fuselage.

[0014] In a preferred solution, a vector thruster is provided on the fuselage.

[0015] In a preferred solution, two or more layers of longitudinal connected radar radomes are provided on the aircraft.

[0016] In a preferred solution, external attitude balance thrusters are provided on the aircraft. They are provided in pairs. When there is a pair, they are symmetrically distributed left and right along the X center line of the fuselage. When there are two pairs, they are symmetrically distributed front and back along the Y center line of the fuselage.

[0017] In a preferred embodiment, a remotely controllable and automatically detachable external power interface is provided on the aircraft, which consists of an on-board power interface and an on-board interface release device. The on-board power interface is provided with a multi-wire power connection line and a ground command control signal line; this external power interface is connected to the ground power supply system through a matching ground power interface and a special anti-tensile cable.

[0018] The beneficial effects achieved by the present invention compared with the prior art are as follows:

[0019] Through the design of the integrated wing, the aircraft can maintain the advantages of multi-rotor vertical takeoff and landing while achieving the high-speed flight ability of a fixed-wing aircraft. A method for the aircraft to flexibly switch working states in different flight stages is designed to meet diverse flight requirements. In the level flight state, the structure originally used as a rotor support can be transformed into a fixed wing to provide additional lift, thereby reducing drag and improving flight efficiency.

[0020] Through the optimized propeller and wing design, the present invention overcomes the problems of airflow interference and stability encountered by multi-rotor models during high-speed flight, enabling the aircraft to achieve higher-speed cruise flight.

[0021] Through precise power distribution and control, the aircraft can achieve efficient utilization of energy in different flight states, significantly improving the energy utilization rate and the overall power ratio.

[0022] During the conversion process between the rotor and fixed-wing working conditions, through fine control strategies and structural designs, the aircraft can maintain flight stability and reduce the possible unstable phenomena during the conversion process.

[0023] The advantages of multi-rotors, fixed wings, and helicopter rotors are complementary: by integrating the advantages of different types of rotors, the aircraft performs excellently in vertical takeoff and landing, hovering, and high-speed flight, realizing an organic combination of multiple flight modes.

[0024] Through a unique wing and rotor layout, the aircraft avoids the adverse effects of downwash airflow on the upper part of the fuselage in the rotor working condition, increasing the takeoff weight and payload.

[0025] The integrated and standardized rotor system, through standardized design, the key components of the aircraft such as rotors, propellers, and rudders have high versatility and interchangeability, simplifying the production and maintenance processes, reducing costs, and increasing the in-flight rate.

[0026] Adopting a layered layout radome design enables the aircraft to be equipped with multiple radar systems to achieve all-round detection coverage and enhance the aircraft's situational awareness ability.

[0027] By setting the external power interface, the aircraft can obtain the support of ground power during the takeoff phase, which improves the takeoff weight and initial payload capacity and increases the flexibility of the mission.

[0028] The foldable design of the rotors and wings makes the aircraft more convenient during storage and transportation, and also simplifies the on-site deployment process.

[0029] The aircraft provides a variety of power system options, including electric motors, gasoline engines, new energy, etc., to adapt to different environmental and mission requirements.

[0030] The aircraft is designed with a variety of safety assistance means, which improves the redundancy of the system and ensures the safety of flight.

[0031] Through an advanced flight control system, the aircraft can achieve precise control of dynamic flight and hovering, meeting the requirements of high-precision flight missions.

[0032] In summary, the multi-rotor integrated aircraft and its flight control method of the present utility model not only improve the flight performance, but also significantly improve the convenience, economy and safety of the aircraft, have broad application prospects and market potential, and provide an efficient, flexible, safe and innovative flight solution for the aviation field. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 : Large-scale drawing of the fixed-wing working condition of the multi-rotor integrated aircraft;

[0034] Figure 2 : Large-scale drawing of the multi-rotor working condition of the foldable multi-rotor integrated aircraft;

[0035] Figure 3 : Large-scale drawing of the rudder installation structure;

[0036] Figure 4 : Large-scale drawing of the combined deployment of the multi-rotor integrated aircraft and the fixed wing;

[0037] Figure 5 : Large-scale drawing of the combination of the multi-rotor integrated aircraft and the additional propulsion device;

[0038] Figure 6 : Working principle diagram of the on-board external power supply;

[0039] Figure 7 : Large-scale drawing of the combined folding of the foldable multi-rotor integrated aircraft and the additional vector propulsion device;

[0040] Figure 8 : Large-scale drawing of the combined deployment of the multi-rotor integrated aircraft and the coaxial dual rotors;

[0041] Figure 9: General assembly drawing of combined folding of multi-rotor integrated aircraft and coaxial dual-rotors;

[0042] Figure 10 : Schematic diagram of flight control principle for multi-rotor integrated aircraft in α flight control and b flight state;

[0043] Figure 11 : Flight principle of rotor flight state Figure 1 ;

[0044] Figure 12 : Flight principle of fixed-wing flight state Figure 2 。

[0045] Reference numerals in the figure:

[0046] 1, airframe; 1-1, fuselage; 1-2, nose; 1-3, tail;

[0047] 2, integrated wing; 2-1, rudder wing; 2-1-1, right front rudder wing; 2-1-2, left front rudder wing; 2-1-3, left rear rudder wing; 2-1-4, right rear rudder wing; 2-1-5, rudder wing transmission shaft; 2-1-6, external interface of rudder wing transmission shaft;

[0048] 2-2, thruster; 2-2-1, right front thruster; 2-2-2, left front thruster; 2-2-3, left rear thruster; 2-2-4, right rear thruster;

[0049] 2-3, propeller; 2-3-1, right front propeller; 2-3-2, left front propeller; 2-3-3, left rear propeller; 2-3-4, right rear propeller;

[0050] 2-4, rudder; 2-4-1, right front rudder; 2-4-2, left front rudder; 2-4-3, left rear rudder; 2-4-4, right rear rudder; 2-4-5, rudder surface; 2-4-6, rudder servo; 2-4-7, rudder rotating shaft;

[0051] 2-5, elevator servo; 2-5-1, right front elevator servo; 2-5-2, left front elevator servo; 2-5-3, left rear elevator servo; 2-5-4, right rear elevator servo;

[0052] 3, flight control system; 4, power system; 5, vector thruster;

[0053] 6, fixed wing; 6-1, main wing; 6-2, folding wing; 6-3, folding wing interface; 6-4, main wing aileron; 6-5, folding wing aileron;

[0054] 7, helicopter rotor structure;

[0055] 8, external attitude balance thruster;

[0056] 9. Radar radome; 9-1. Radome body; 9-2. Radome bracket; 9-3. Radome hatch;

[0057] 10. External power interface; 10-1. Aircraft power interface; 10-2. Aircraft interface release; 10-3. Ground power interface; 10-4. Ground interface release; 10-5. Tensile cable; 10-6. Cable car. Detailed implementation mode

[0058] The technical solutions in the present invention will be clearly and completely described below with reference to the accompanying drawings in the present invention. In addition, the forms of the various structures described in the following embodiments are merely examples, and the present invention is not limited to the various structures described in the following embodiments. All other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0059] The present invention provides a multi-rotor integrated aircraft, which includes a fuselage 1, an integrated wing 2, a flight control system 3, and a power system 4. The integrated wing 2 integrates a rudder wing 2-1, a propeller 2-2, a rudder 2-4, and a rudder wing transmission shaft 2-1-5 interface into a unified standard component. The elevation angle α of the entire integrated wing 2 is controlled by an elevator 2-5 through the rudder wing transmission shaft 2-1-5, and the rudder angle β of the rudder 2-4 on each integrated wing 2 is controlled by the rudder 2-4 machine; the propeller 2-2 and the rudder 2-4 are both installed on the rudder wing;

[0060] The elevator 2-5 is connected to the rudder wing transmission shaft 2-1-5, and the rudder wing transmission shaft 2-1-5 is connected to the rudder wing through the rudder wing transmission shaft 2-1-5 interface;

[0061] The external interface 2-1-6 of the rudder wing transmission shaft can be designed as one or more of plug-in, screw connection, bolt connection, snap connection, and foldable interface according to the model;

[0062] At least one rudder wing, one propeller 2-2, and one rudder 2-4 are provided on each integrated wing 2; each integrated wing 2 is controlled by at least one elevator 2-5; this can ensure that each rudder wing has independent maneuverability and can increase the redundancy of flight safety by several times;

[0063] Two integrated wings 2 symmetrically arranged along the left and right of the fuselage 1 form a group. The propellers 2-2 of the two integrated wings 2 symmetrically arranged on the left and right have the same specifications, and the torque directions generated are opposite. The elevator specifications of the two integrated wings 2 symmetrically arranged on the left and right are the same;

[0064] At least two sets of integrated wings 2 are arranged front and back on the airframe 1; the torque directions generated by the thrusters 2-2 on the diagonal of each two sets of integrated wings 2 are the same; the power of each set of thrusters 2-2 is the same; such a design can ensure good stability and adaptability during the flight of the aircraft.

[0065] The thruster 2-2 can be any one of a pull thruster 2-2, a push thruster 2-2, and a pull-push combined thruster 2-2, which facilitates the flexibility of the selection of the power system 4 of the aircraft model to suit the aircraft model to complete specific flight environments and tasks;

[0066] This design is a necessary condition that can meet both the rotary-wing flight state and the fixed-wing flight control;

[0067] Obviously, this design can realize the control of the aircraft by using the elevator and rudder 2-4 and the total power control whether in the rotary-wing working state or the fixed-wing working state. It can also enable each rotary-wing motor to complete the flight control without further configuring a power regulator, which not only saves a large number of electrical components, reduces the self-weight, but also reduces the power loss and saves energy;

[0068] A multi-rotor integrated aircraft has multi-rotor flight conditions and fixed-wing flight conditions, and the two conditions can be flexibly switched. Having these two flight modes can solve the problem of taking off and landing whether on an airport runway or a simple helicopter platform, etc. It can fly at high speed in the air and hover in the air, broadening the application scenarios and usage values of the aircraft. The thruster 2-2 takes into account the functions of generating lift for vertical takeoff and landing, horizontal pull or thrust, and traction for flight maneuver, improving their utilization rate, reducing the weight of the airframe 1 itself, optimizing the thrust-to-weight ratio parameter, and is a completely innovative functional integration design concept and a re-improved airframe 1 structural configuration.

[0069] The position setting of the thrusters 2-2 of the integrated wings 2 enables the downwash airflow generated in the multi-rotor flight condition not to constitute a sinking wind load on the upper area of the aircraft airframe 1. This setting condition is designed specifically for the thrusters 2-2 so that the maximum vertical takeoff weight will not be reduced in the multi-rotor flight state, which is an important indicator for improving the flight performance of the aircraft. Such a setting can not only maintain the stability of the airflow around the aircraft, but also improve the wind resistance of the aircraft, especially having very excellent aerodynamic stability in the takeoff, landing, and hovering states. However, in the traditional tilt-rotor aircraft, when taking off and landing vertically, the upper part of the airframe 1 and some fixed-wing 6 segments will block a part of the rotor downwash airflow, resulting in a decrease in the lift of the whole aircraft. Moreover, the airflow under the fixed-wing 6 is disordered during the tilting process, especially the moment of the fuselage 1-1 during the tilting process is difficult to balance, increasing the difficulty of controlling the attitude of the tilt-rotor aircraft fuselage 1-1.

[0070] The rudder 2-4 is a two-way symmetric rudder surface structure arranged in the forward direction of the airflow of the thruster 2-2 and based on the rudder wing. It consists of a rudder servo 2-4-6, a rudder 2-4 transmission shaft, and two rudder surfaces 2-4-5. This design ensures a high rudder efficiency of the rudder 2-4 by means of the forward airflow of the thruster 2-2, improving the heading control sensitivity of the aircraft. The two-way symmetric rudder surface structure based on the rudder wing can cancel out the deformation torque generated on the rudder wing surface, overcoming the torque generated by a single-sided rudder and affecting the structural deformation of the rudder wing.

[0071] The integrated wing 2 is provided with a wing folding mechanism. This design is beneficial for storage and transportation, saving hangar space, and facilitating on-site deployment and application operations.

[0072] The fuselage 1 is provided with a fixed wing 6. The fixed wing 6 can also be one or more of a rotatable retractable fixed wing 6 and a folding fixed wing 6. Among them, the folding wing 6-2 is divided into a main wing 6-1 close to the fuselage 1 and a folding wing 6-2 far from the fuselage 1. A folding wing interface 6-3 is provided between the main wing 6-1 and the folding wing 6-2. A main wing aileron 6-4 is provided on the main wing 6-1. A folding wing aileron 6-5 is provided on the folding wing 6-2.

[0073] This design is mainly to improve the aerodynamic performance of the aircraft, increase lift, improve the service ceiling, increase the range, greatly enhance the economy, and also improve the handling stability and wind resistance of the aircraft. It also compensates for the safety during the conversion between the rotor flight state and the fixed wing 6 flight state. The folding wing 6-2 is convenient for storage and transportation, and electric folding reduces the labor intensity of people.

[0074] The fuselage 1 is provided with a helicopter rotor structure 7. The helicopter rotor structure 7 is one of a single rotor, a coaxial dual rotor structure, and a tandem dual rotor. When the aircraft needs to perform large-load vertical takeoff and landing, it is necessary to install the helicopter rotor structure 7 to ensure. The use of a coaxial dual rotor can cancel out the torque generated by the forward and reverse rotation of the blades.

[0075] A vector thruster 5 is provided on the aircraft. When designing, factors such as non-interference between the vector thruster 5 in the deflection working condition and the fuselage 1 and external equipment also need to be considered. The setting of the vector thruster 5 is mainly to improve its control performance, used to supplement and correct the load redundancy against instability, and reduce the impact of sudden airflow on flight safety. When the vector thruster 5 is added, a helicopter single-axis rotor structure can also be provided on the fuselage 1 to jointly form a complementary scheme for flight stability. The vector thruster 5 can also assist in completing the climbing, diving, turning, dynamic balance and other control functions of the aircraft.

[0076] Two or more layers of longitudinally connected radar radomes 9 are provided on the aircraft, which are composed of a radome main body 9-1, a radome support 9-2 and a radome hatch 9-3. The radome main body 9-1 is installed on the unobstructed part of the airframe 1 through the radome support 9-2, or can also be installed on the unobstructed part of the top of the fixed wing 6 or helicopter rotor 7 connected to the airframe 1. A radome hatch 9-3 is provided on the top radome main body 9-1 to facilitate maintenance and equipment installation; this design can solve the problems that each layer of radome is all-round, unobstructed, and does not interfere with each other in transmission and reception. For example, for the layered radomes arranged on the upper part of the airframe 1 or the top of the main shaft of the helicopter rotor 7, the top layer radome can be provided with satellite receiving and interconnection antennas facing the sky, and the lower layers can be arranged with air-to-air and air-to-sea radar antennas in sequence; of course, a parachute can also be built into the radome for emergency in special situations of the aircraft.

[0077] External attitude balance thrusters 8 are provided on the aircraft, which are arranged in pairs. When there is a pair, they are symmetrically distributed left and right along the X center line of the airframe 1. When there are two pairs, they are symmetrically distributed front and back along the Y center line of the airframe 1; they are mainly used to balance the flight attitude of the airframe 1, automatically connect to the external power supply through the switch circuit during takeoff and landing, and can also be used to increase the takeoff lift.

[0078] A remotely controllable and automatically detachable external power interface 10 is provided on the aircraft. This external power interface 10 is connected to the ground power supply system through a special anti-tensile cable 10-5, and is mainly composed of an airborne power interface 10-1 and an airborne interface release device 10-2; it is used in combination with the ground power interface 10-3, the anti-tensile cable 10-5, the cable car 10-6 and the ground power connection. This design mainly ensures that the power supply of the aircraft during the takeoff stage is supported by the ground power supply, and is transmitted to each electric thruster 2-2 through the switch control circuit, increasing the takeoff weight and effective load of the aircraft, and controlling the detachment after the aircraft takes off to a certain height through cable connection, so as to increase the potential energy of the aircraft and increase the effective load. Especially when performing delivery tasks, its superiority can be more reflected; the ground power supply system is supplied by shore power or ship power.

[0079] The power system 4 of the thruster 2-2 is one or more of an electric motor, a gasoline engine, new energy, and a hybrid power system 4; the propeller pitch of the propeller 2-3 is one or more combinations of a fixed pitch and a variable pitch; these selections are mainly for the optimal combination of the power system 4 when designing aircraft for different types of tasks. Especially the introduction and application of new energy can improve the power adaptability and environmental protection performance of the aircraft, opening up a direction for the popularization of new technologies.

[0080] Flight control method for a multi-rotor integrated aircraft:

[0081] It is divided into a rotor working state and a fixed-wing flight state. The position setting of the propeller 2-2 of the integrated wing 2 makes the disk surface of the propeller 2-3 lower than the upper surface of the aircraft body 1 under the multi-rotor flight condition, and the downwash airflow of the propeller 2-3 does not cause a sinking wind load on the upper area of the aircraft body 1. This aircraft has a rotor working state and a fixed-wing flight state. The flight control in the multi-rotor flight condition and the fixed-wing flight condition is completed by changing the horizontal rudder angle of the rudder wing, the rudder angle β of the rudder 2-4 and the power combination control of the propeller 2-2;

[0082] The elevator servo 2-5 is arranged inside the body 1; the servo is provided with an over-limit safety protection device. In the rotor flight state, the elevation angle α of the rudder wing of the integrated wing 2 is controlled between 75° and +105°; in the fixed-wing flight state, the elevation angle α of the rudder wing of the integrated wing 2 is controlled between -15° and +25°, and the conversion between the rotor and fixed-wing flight states is carried out with α controlled between +25° and +75°. This small-angle design is mainly to control the stability of the rotor flight state and the fixed-wing flight state, prevent overload caused by large-angle operations and cause flight accidents. The data is the corrected value after structural design and big data statistics. Within this range, it can not only ensure the maneuverability of the aircraft, but also minimize the flight accident rate, which is an optimized combination parameter.

[0083] For the convenience of describing the flight control method, the two groups of integrated wings 2 are respectively named the right front rudder wing 2-1-1, the left front rudder wing 2-1-2, the left rear rudder wing 2-1-3, and the right rear rudder wing 2-1-4; the rudder 2-4 is respectively named the right front rudder 2-4-1, the left front rudder 2-4-2, the left rear rudder 2-4-3, and the right rear rudder 2-4-4. The direction of the rudder angle β is the deflection direction of the tail of the rudder 2-4, positive to the right and negative to the left; the propellers 2-3 are respectively named the right front propeller 2-3-1; the left front propeller 2-3-2; the left rear propeller 2-3-3; the right rear propeller 2-3-4; among them, the right front propeller 2-3-1 and the left rear propeller 2-3-3 both rotate counterclockwise, and the reverse torque is in the clockwise direction. The left front propeller 2-3-2 and the right rear propeller 2-3-4 rotate clockwise, and the reverse torque is in the counterclockwise direction. The propellers 2-2 are all set as electric motors, and the thrust is represented by F, positive for power increase and negative for power reduction. Since the total power of each propeller 2-3 is controlled synchronously and the power of each propeller 2-3 is also the same, the torques generated cancel each other out and the total torque is zero. The flight control is mainly achieved by the elevator, the rudder 2-4 and the total power adjustment. The specific steps are as follows:

[0084] S1. Pre - flight system check: First, move the aircraft to the take - off area and unfold it. Check that all systems are in good condition. Power on the systems to restore them to the initial state. The elevation angle α of the two integrated wings 2 and the rudder angle β of the rudder 2 - 4 are restored to the unified initial reference zero - position state. If it is not at the zero - position, it should be fine - tuned to the zero - position. Then press the reset wing button. At this time, the flight control system 3 automatically memorizes the wing reference zero - position as the control reference plane for the entire flight process. Through the flight control system 3, remotely control the servo to check that the elevation angle α of the two integrated wings 2 is between - 15° and + 105°, and the rudder angle β of the rudder 2 - 4 operates flexibly within the set range, and the over - limit limit control is accurate.

[0085] Rotary wing working state: This aircraft also has the control mode usually adopted by quad - rotors, which relies on the difference in lift and torque generated by changing the rotation speed of the corresponding propellers 2 - 3 by each power regulator. However, it requires a power regulator to be configured for each thruster 2 - 2 of each propeller 2 - 3 to complete. This aircraft can complete the flight control method only with a total power synchronous control, that is, it can also complete the flight control without configuring a power regulator for each propeller 2 - 3. The flight control is completed by controlling the elevation angle α of the integrated wing 2 between 75° and + 105° through the elevator servo 2 - 5, the rudder angle β of the rudder 2 - 4 within the set range, and the overall change of the power output by the power system 4 to the thrusters 2 - 2 to generate force changes in different directions of the aircraft. In this flight state, the lift F generated by the thrusters 2 - 2 mainly overcomes the gravity W of the aircraft.

[0086] S2. Vertical take - off, landing and hovering:

[0087] ① Start and reset: The flight control system 3 drives the two integrated wings 2 through the servo so that the wing surfaces are in a vertical state, the elevation angle α≈90 degrees, and the rudder angle β = 0 degrees. This state is achieved in one step by setting a dedicated button. The propeller disks of the propellers 2 - 3 are all in a vertically upward state. If it is a variable - pitch thruster 2 - 2, the total pitch is further reduced to the vertical take - off start state.

[0088] ② Start: The flight control system 3 supplies small - current power to the motors of each thruster 2 - 2 of the integrated wing 2 through the power system 4, and ensures that the rotation speeds of each group of propellers 2 - 3 are the same and the rotation directions are opposite.

[0089] ③ Vertical take - off: After pre - heating for two minutes and observing stable operation, synchronously increase the output power. If it is a variable - pitch thruster 2 - 2, the total pitch also needs to be synchronously increased to the vertical take - off state. When the total lift F generated by all thrusters 2 - 2 is greater than the take - off weight W of the aircraft, it is in the state of vertical take - off and ascending.

[0090] ④Hovering: After taking off to a certain height, synchronously reduce the output power of each thruster 2-2. If it is a variable-pitch thruster 2-2, it is also necessary to synchronously reduce the total pitch to the hovering state. When the total lift force F generated by all thrusters 2-2 is equal to the current weight W of the aircraft, it is in the hovering state;

[0091] ⑤Descending: Synchronously reduce the output power further. If it is a variable-pitch thruster 2-2, it is also necessary to synchronously reduce the total pitch to the descending state. When the total lift force F generated by all thrusters 2-2 is less than the current weight W of the aircraft, it is in the vertical descending state;

[0092] ⑥Forward and backward movement: After taking off to a certain height, the aircraft is translated forward and backward by the flight control system 3 by synchronously controlling the wing surface elevation angle α of the two sets of integrated wings 2. When the wing surface elevation angles α of the left and right integrated wings 2 in a group in the nose 1-2 direction synchronously decrease from the vertical 90° to 75°, a component force in the direction of the nose 1-2 is generated in the horizontal direction by the thrusters 2-2, and the aircraft moves forward; conversely, when the wing surface elevation angles α of the left and right integrated wings 2 in a group in the tail 1-3 direction synchronously increase from the vertical 90° to 105°, a component force in the direction of the tail 1-3 is generated in the horizontal direction by the thrusters 2-2, and the aircraft moves backward; when the wing surface elevation angles α of the front and rear sets of wings are all tilted forward, the fuselage 1 will accelerate forward, and conversely, when the wing surface elevation angles α of the front and rear sets of wings are all tilted backward, the fuselage 1 will accelerate backward;

[0093] ⑦Left and right lateral movement: After taking off to a certain height, synchronously deflect the rudder angles β of the rudders 2-4 on the two sets of rudder wings to the right. Due to the airflow generated by the propellers 2-3, all rudder surfaces are subjected to a thrust from right to left, causing the entire fuselage 1 to translate to the left. Similarly, synchronously deflect the rudder angles β of the rudders 2-4 on the two sets of rudder wings to the left, and all rudder surfaces are subjected to a thrust from left to right, causing the entire fuselage 1 to translate to the right. When the rudder angle β of the rudder 2-4 is at the best rudder efficiency, the lateral translation efficiency of the fuselage 1 is the highest;

[0094] ⑧Left and right turning: After taking off to a certain height, the aircraft generates a rotational torque around the Z-axis by the flight control system 3 through synchronously controlling the changes in the rudder angles β of the front and rear sets of rudders 2-4, causing the nose 1-2 to deflect around the Z-axis to turn; Right-turning control method: When the rudder angles β of the left and right integrated wings 2 on the rudders 2-4 in a group in the tail 1-3 direction are synchronously deflected to the right or the rudder angles β of the left and right integrated wings 2 on the rudders 2-4 in a group in the nose 1-2 direction are synchronously deflected to the left, according to the principle of the rudder surface being affected by the wind force, it can be analyzed that torques in the clockwise direction around the Z-axis are generated in the tail 1-3 and nose 1-2 directions respectively, causing the nose 1-2 direction to deflect to the right to turn, and conversely to turn to the left.

[0095] Fixed-wing working state: It refers to the flight control completed by changing the forces acting on the aircraft in different directions through the servo control of the wing surface elevation angle α of the integrated wing 2 within the range of -20° to +30°, the rudder angle β of the rudder 2-4 within the set range, and the power distribution system to change the power output by the thruster 2-2. In this flight state, the power generated by the thruster 2-2 mainly pulls the aircraft forward to overcome air resistance;

[0096] ⑨Conversion from the rotary-wing working state to the fixed-wing working state: When the aircraft is flying forward in the rotary-wing working state, synchronously reset the rudder angle β of the two integrated wings 2 in the front and rear groups to zero, synchronously reduce the wing surface elevation angle α of the integrated wing 2 from the vertical 90° to 75°, and increase the power of each thruster 2-2 to accelerate the forward flight speed. At the same time, continue to synchronously change the wing surface elevation angle α of the integrated wing 2 from 75° to 0°. Then, the thruster 2-2 generates a greater component force in the direction of the nose 1-2, and the aircraft accelerates forward. The lift generated by the two integrated wings 2 in the two groups gradually increases. When the total lift F generated by the integrated wing 2 is equal to the weight W of the airframe 1, the aircraft is in the level flight state, thus completing the conversion from the rotary-wing working state to the fixed-wing working state;

[0097] ⑩Level flight, climb, dive, and glide in the fixed-wing working state:

[0098] Level flight: When the total lift F generated by the integrated wing 2 is equal to the weight W of the airframe 1 in the fixed-wing working state, the aircraft is in the level flight state;

[0099] Climb: When the wing surface elevation angle α of the pair of integrated wings 2 on the left and right in the front group in the direction of the nose 1-2 is synchronously increased from 0° to 25° or the wing surface elevation angle α of the pair of integrated wings 2 on the left and right in the rear group in the direction of the tail 1-3 is synchronously decreased from 0° to -15°, the wing surface of the integrated wing 2 is subjected to the resistance of the airflow, causing the airframe 1 to rotate backward around the Y-axis. After the nose 1-2 is lifted, the aircraft is in the climbing process;

[0100] Dive: When the wing surface elevation angle α of the pair of integrated wings 2 on the left and right in the front group in the direction of the nose 1-2 is synchronously changed from 0° to -15° or the wing surface elevation angle α of the pair of integrated wings 2 on the left and right in the rear group in the direction of the tail 1-3 is synchronously increased from 0° to 25°, the wing surface of the integrated wing 2 is subjected to the resistance of the airflow, causing the airframe 1 to rotate forward around the Y-axis. After the nose 1-2 dips, the aircraft is in the diving process;

[0101] Glide: While maintaining the level flight state, synchronously reduce the power of each thruster 2-2 to reduce the level flight speed. When the total lift F generated by the integrated wing 2 is less than the weight W of the airframe 1, the aircraft is in the gliding state;

[0102] ⑪ Left and right turning in the fixed-wing working state: The aircraft is rotated around the Z-axis by the flight control system 3 through synchronously controlling the rudder angles β of the front and rear groups of rudders 2-4, causing the nose 1-2 to deflect around the Z-axis for turning. For the left-turning control method, when the rudder angles β of the left and right two integrated wing rudders 2-4 on one side of the tail 1-3 direction are synchronously deflected to the left or the rudder angles β of the left and right two integrated wing rudders 2-4 on one side of the nose 1-2 direction are synchronously deflected to the right, according to the principle of the rudder surface being affected by the wind force, it can be analyzed that torques in the counterclockwise direction around the Z-axis are generated in the tail 1-3 and nose 1-2 directions respectively, causing the nose 1-2 to deflect to the left for turning, and vice versa for turning to the right.

[0103] ⑫ Rolling and lateral movement in the fixed-wing working state: By differentially changing the angle of attack α of a pair of integrated wing surfaces 2 on the left and right sides of the airframe 1, a torque for rotating around the X-axis is generated, causing the airframe 1 to roll left and right and shift to the side of the roll. For example, if the angle of attack α of the right integrated wing 2 is increased and the angle of attack α of the left integrated wing 2 is decreased synchronously, a left roll is generated, and at the same time a lateral shift to the left is generated, and vice versa for a right roll and a synchronous lateral shift to the right;

[0104] Horizontal lateral movement: Synchronously deflect the rudder angles β of the two groups of rudders 2-4 to the right, and all the rudder surfaces are subjected to a thrust from right to left, causing the entire airframe 1 to translate to the left. Similarly, synchronously deflect the rudder angles β of the two groups of rudders 2-4 to the left, and all the rudder surfaces are subjected to a thrust from left to right, causing the entire airframe 1 to translate to the right. When the rudder angle β of the rudders 2-4 is at the optimal rudder efficiency, the lateral translation efficiency of the airframe 1 is the highest;

[0105] Conversion from the fixed-wing working state to the rudder-rotor working state: In the fixed-wing level flight working state, while synchronously changing the angle of attack α of the integrated wing 2 from 0° to between 75° and +105°, increase the power of each thruster 2-2, so that the lift F generated by the thrusters 2-2 is not less than the weight W of the aircraft. After all the flight indicators are stable, it is considered a successful conversion. The vector thruster 5 can also assist in completing the climbing, diving, turning, dynamic balance and other control functions of the aircraft.

[0106] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principle of the present invention shall be included within the protection scope of the present invention.

Claims

1. A multi-rotor integrated aircraft, characterized in that, It includes a fuselage, an integrated wing, a flight control system, and a power system, where: The integrated wing is a unified component integrating a control wing, a thruster, a rudder, and a control wing drive shaft interface. The elevation angle α of the entire integrated wing is controlled by an elevator servo through the control wing drive shaft. The rudder angle β of the rudder on each integrated wing is controlled by a rudder servo. The thruster and the rudder are both installed on the control wing. The elevator servo is connected to the control wing drive shaft, and the control wing drive shaft is connected to the control wing through the control wing drive shaft interface. At least one control wing, one thruster, and one rudder are provided on each integrated wing. Each integrated wing is controlled by at least one elevator servo. Two integrated wings symmetrically arranged left and right along the fuselage form a group. The thrusters of the two integrated wings symmetrically arranged left and right have the same specifications and generate torques in opposite directions. The elevator servos of the two integrated wings symmetrically arranged left and right have the same specifications. At least two groups of integrated wings are provided front and back on the fuselage. The torques generated by the thrusters on the diagonals of every two groups of integrated wings are in the same direction and have the same power.

2. The multi-rotor integrated aircraft according to claim 1, wherein: The rudder is a two-way symmetric rudder surface structure provided in the center of the thruster air flow in the forward direction and based on the control wing.

3. A multi-rotor integrated aircraft according to claim 1, characterized in that: The integrated wing is provided with a wing folding mechanism.

4. A multi-rotor integrated aircraft according to claim 1, wherein: The fuselage is provided with a fixed wing.

5. A multi-rotor integrated aircraft according to claim 1, characterized in that: The fuselage is provided with a helicopter rotor structure.

6. The multi-rotor integrated aircraft according to claim 1, characterized in that: A vector thruster is provided on the fuselage.

7. A multi-rotor integrated aircraft according to claim 1, characterized in that: Two or more layers of longitudinally connected radar radomes are provided on the aircraft.

8. A multi-rotor integrated aircraft according to claim 1, characterized in that: External attitude balance thrusters are provided on the aircraft. They are provided in pairs. When there is a pair, they are symmetrically distributed left and right along the X center line of the fuselage. When there are two pairs, they are symmetrically distributed front and back along the Y center line of the fuselage.

9. The multi-rotor integrated aircraft according to claim 1, wherein: A remotely controllable and automatically detachable external power interface is provided on the aircraft. It consists of an on-board power interface and an on-board interface releaser. The on-board power interface is provided with a multi-strand power connection wire and a ground command control signal wire. This external power interface is connected to the ground power supply system through a matching ground power interface and a special anti-tensile cable.

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