A tilt multi-rotor unmanned aerial vehicle and a tilt transition method

Abstract

The present invention discloses a tilt multi-rotor unmanned aerial vehicle and a tilt transition method, belonging to the technical field of unmanned aerial vehicles. The tilt multi-rotor unmanned aerial vehicle includes a fuselage, wings are symmetrically arranged on both sides of the fuselage, winglets with rudders are arranged at the ends of the wings, a front tilt mechanism symmetrically arranged with respect to the fuselage is arranged at the front end of the fuselage, a rear tilt mechanism symmetrically arranged with respect to the fuselage is arranged at the rear end of the fuselage, and a propulsion propeller is arranged in the middle of the rear end of the fuselage; the front tilt mechanism includes a front propeller, and the front propeller is connected to the fuselage through a tilt power structure, and the rear tilt mechanism includes a rear propeller, and the rear propeller is connected to the fuselage through a tilt power structure. By adopting the above-mentioned tilt multi-rotor unmanned aerial vehicle and tilt transition method, the flight stability is improved through the fuselage layout in cooperation with the front tilt mechanism and the rear tilt mechanism, and stronger air mobility is achieved.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aerial vehicles, and in particular to a tilting multi-rotor unmanned aerial vehicle and a tilting transition method. Background Art

[0002] Currently, unmanned aerial vehicles are mainly divided into rotor unmanned aerial vehicles and fixed-wing unmanned aerial vehicles. Rotor unmanned aerial vehicles can achieve vertical takeoff and landing without relying on a runway and other ground auxiliary facilities, and can adapt to various takeoff environments. However, due to their short endurance and low flight speed, their application scenarios are limited to a certain extent. Fixed-wing unmanned aerial vehicles can fly at high speeds and over long distances. However, they need to rely on a runway or ground launch and recovery equipment to assist in takeoff and landing, which also limits their scope of application to a certain extent. Tilt-rotor unmanned aerial vehicles integrate the advantages of both, and can not only take off and land vertically, but also fly at high speeds and over long distances. They can be applied to tasks such as short-distance transportation, medical rescue, aerial photography, agricultural plant protection, wildlife observation, surveying and mapping, power line inspection, and aerial search and rescue. Therefore, tilt-rotor unmanned aerial vehicles are regarded as one of the trends in the future development of unmanned aerial vehicles.

[0003] At present, in most traditional vertical takeoff and landing technologies, the power device during takeoff and landing will contract and fold after takeoff, resulting in redundancy in the power system and increasing the weight of the unmanned aerial vehicle. The wing tip effect of existing conventional layout unmanned aerial vehicles is relatively large, which will affect the lift of the wing surface.

[0004] Currently, the transition of traditional tilt-wing unmanned aerial vehicles has high requirements for the flight control system. During the tilting process, the lift is still provided by the rotors. It is difficult to control the flight attitude during the tilting process, and the torsional load of the tilting mechanism is relatively large, and the stress on the torsion bar is relatively large. Once a problem occurs in the tilting mechanism, the unmanned aerial vehicle cannot continue to fly, and problems are likely to occur during the transition process. Summary of the Invention

[0005] The purpose of the present invention is to provide a tilting multi-rotor unmanned aerial vehicle and a tilting transition method to solve the problems raised in the above background art.

[0006] To achieve the above purpose, the present invention provides a tilting multi-rotor unmanned aerial vehicle, including a fuselage. Wings are symmetrically arranged on both sides of the fuselage. Winglet with a rudder is arranged at the end of the wing. A front tilting mechanism symmetrically arranged with respect to the fuselage is arranged at the front end of the fuselage. A rear tilting mechanism symmetrically arranged with respect to the fuselage is arranged at the rear end of the fuselage. A propulsion propeller is arranged in the middle of the front and rear ends of the fuselage.

[0007] The front tilting mechanism includes a front propeller, and the front propeller is connected to the fuselage through a tilting power structure. The rear tilting mechanism includes a rear propeller, and the rear propeller is connected to the fuselage through a tilting power structure.

[0008] Preferably, a fuselage frame for support is arranged inside the fuselage.

[0009] Preferably, ailerons and flaps are provided at the rear end of the wing. The ailerons are used for the roll maneuver of the UAV, and the flaps are used for the takeoff and landing of the UAV.

[0010] Preferably, a wing skeleton is provided inside the wing. The wing skeleton includes a main carbon fiber tube beam frame, an auxiliary carbon fiber tube beam frame, and several wing ribs. The main carbon fiber tube beam frame and the auxiliary carbon fiber tube beam frame are respectively arranged along the length direction of the wing. The wing ribs are arranged between the main carbon fiber tube beam frame and the auxiliary carbon fiber tube beam frame and are arranged at intervals along the length direction of the wing.

[0011] Preferably, the rudder is provided at the rear end of the winglet. A winglet skeleton is provided inside the winglet. The winglet skeleton includes winglet ribs and a winglet carbon fiber beam frame. The winglet carbon fiber beam frame is arranged along the length direction of the winglet, and the winglet ribs are arranged along the width direction of the winglet.

[0012] Preferably, the fuselage is provided with a tilting device fixing frame. The front propeller and the rear propeller are respectively connected to the tilting device fixing frame through a tilting power structure. Fairings are provided at the front ends of the front propeller and the rear propeller. The tilting power structure includes a tilting frame. A servo motor is provided on the mounting surface of the tilting frame. The output end of the servo motor is connected to the front propeller or the rear propeller. The frame plates on both sides of the tilting frame are connected to the box body through a tilting connection assembly, and the box body is connected to the tilting device fixing frame.

[0013] Preferably, the tilting connection assembly includes a worm provided inside the box body. One end of the worm is connected to the output shaft of the tilting motor. A fixing box for fixing the tilting motor is provided inside the box body. The other end of the worm is rotationally connected to the box body through a bearing. The middle of the worm is connected to the middle of the tilting gear rod. The two ends of the tilting gear rod respectively pass through the box body and are connected to the frame plates on both sides of the tilting frame.

[0014] Preferably, a front landing gear is provided at the front end of the bottom of the fuselage. A left landing gear and a right landing gear are symmetrically provided on both sides of the bottom of the fuselage. A flight control device, a mission payload, and a battery are provided inside the fuselage. A propulsion motor is provided at the rear end of the fuselage. The output shaft of the propulsion motor is connected to a propulsion propeller. A displacement sensor, a phase sensor, an acceleration sensor, and a gyroscope are integrated in the flight control device.

[0015] The present invention also provides a tilting transition method for a tilting multi-rotor UAV, including the following steps:

[0016] Step 1, perform signal processing in real time and in sequence: Taking the center of gravity of the UAV as the coordinate origin O, taking the chord direction of the UAV as the x-axis, taking the plane passing through the chord and horizontal to the forward direction of the UAV as the chord plane, taking the direction perpendicular to the chord plane of the UAV as the y-axis, and taking the direction perpendicular to both the x-axis and the y-axis as the z-axis;

[0017] According to the acceleration sensor, obtain the x-axis acceleration u, y-axis acceleration v, and z-axis acceleration w of the drone. The gyroscope obtains the pitch angular velocity ω3, pitch angle θ3, roll angular velocity ω1, roll angle θ1, yaw angular velocity ω2, and yaw angle θ2 of the flight state of the drone; calculate the components F x 、F y 、F z of the resultant force acting on the drone during flight on the x-axis, y-axis, and z-axis, and the pitch resultant moment M3, roll resultant moment M1, and yaw resultant moment M2 generated by the sum of the rotor thrust and flight aerodynamic forces; the front propeller and the rear propeller are rotors;

[0018] By integrating the acceleration, further obtain the forward flight speed V X in the horizontal direction, the vertical speed V Y 、yaw speed V Z ;

[0019] Step 2, adjust the attitude of the drone according to the above calculations. By adjusting the servo motors on the four tilting power structures, further adjust the vertical speed V Y and the pitch resultant moment M3, so that when the drone is not flying horizontally, the forward flight speed V X tends to 0 until it reaches a flight altitude of 50 m;

[0020] Step 3, after reaching 50 m according to the above flight process, start the propulsion motor, and change the forward flight speed V X of the drone in the horizontal direction by changing the thrust. As V X increases, gradually reduce the rotational speed of the servo motors on the four tilting power structures. Judge the lift change of the drone through the change of the forward flight speed V X in the horizontal direction, and adjust the balance of the drone in the overall y-axis and z-axis directions according to the flight control device by controlling the rudder on the winglet, the aileron on the wing, and the rotational speed of the servo motors on the four tilting power structures until the rotational speed of the servo motors on the four tilting power structures drops to 0;

[0021] Step 4, drive the tilting frame to rotate through the worm, and then adjust the positions of the front propeller and the rear propeller, so that the drone transitions from the vertical state to the horizontal state. Restart the servo motors on the four tilting power structures through the flight control device, and drive the drone to fly through a total of five power sources of two pulls and three pushes.

[0022] Preferably, in step 1, the following formulas are used to calculate and obtain the flight parameters of the drone:

[0023] F x =F 推 +mgsin(θ3)+F 阻 cos(β);

[0024] Fy = F 升 + mgcos(θ3) + F 阻 sin(β);

[0025]

[0026] u = F x / m;

[0027] v = F y / m;

[0028] w = F z / m;

[0029] V X = ∫u cos(θ3) + v sin(θ3)dt;

[0030] V Y = ∫u sin(θ3) + v cos(θ3)dt;

[0031] V Z = ∫wcos(θ1) + v sin(θ1)dt;

[0032]

[0033] H = ∫V Y dt;

[0034]

[0035] Wherein, M 偏 is the yaw moment, F x , F y , F z respectively represent the components of the resultant force received by the UAV during vertical takeoff and landing on the x-axis, y-axis, and z-axis. The angles within the trigonometric functions in the above formulas are all implicit functions. F 推 is the thrust of the UAV along the chord direction; F 升 is the lift force in the plane perpendicular to the chord; F 阻 is the drag force in the direction of the actual oncoming flow. The β angle is the angle between the actual oncoming flow direction and the chord. I1, I2, and I3 are the moments of inertia of the airframe about the x-axis, y-axis, and z-axis respectively, is the forward displacement of the UAV in the V X direction, H is the forward displacement of the UAV in the V Y direction, that is, the flight altitude, is the forward displacement of the UAV in the V Z direction, is the sum of and The modulus represents the actual flight distance; R is the projection of the actual distance between the wing lift center and the center of gravity on the x-axis. In the unmanned aerial vehicle, the lift center is on the negative half-axis of the x-axis;

[0036] In step two, the propulsion motor is not started. The component of the resultant force on the y-axis, excluding gravity, acting on the unmanned aerial vehicle during vertical takeoff and landing is the sum of the gravity mg and the downward resistance F 阻 and thus the vertical acceleration v of the unmanned aerial vehicle at this time is obtained. The formula is as follows:

[0037] F 阻 = kV Y 2 ;

[0038] F y = mg + kV Y 2 ;

[0039]

[0040] where m is the fuselage mass and k represents the resistance constant;

[0041] According to the Newton-Euler method in classical mechanics, a suitable dynamic mathematical model is constructed:

[0042]

[0043] In the above formula, the control of the unmanned aerial vehicle depends on the roll pitch and yaw angular velocities and their influence on the attitude. The front propellers and the rear propellers are all propellers with the same structure. The rotational speed of the propeller Ω(t) = -Ω1(t) + Ω2(t) - Ω3(t) + Ω4(t) determines the lift and torque. The driving torques T1(t), T2(t), T3(t) determine the horizontal, vertical movement and turning of the unmanned aerial vehicle, and the moment of inertia I R reflects the anti-rotation ability of the propeller;

[0044] According to the above dynamic mathematical model, in step two, the rotational speeds Ω1(t), Ω2(t), Ω3(t), Ω4(t) of the two front propellers and the two rear propellers are adjusted to adjust the flight attitude by changing the rotational speeds of the front propellers and the rear propellers;

[0045] In step three, the propulsion motor is started. The speed of the servo motor finally becomes 0. The lift received by the unmanned aerial vehicle is equal to the gravity, and the horizontal flight speed V of the unmanned aerial vehicle at this time is obtained X .

[0046] Therefore, the present invention adopts the above-mentioned tilt multi-rotor unmanned aerial vehicle and tilt transition method, and has the following beneficial effects:

[0047] (1) The present invention can smoothly transition the take-off process. During the tilting process of the unmanned aerial vehicle (UAV), the front propellers and the rear propellers do not rotate, greatly reducing the stress on the tilting link, which has the highest failure rate in the tilting-rotor UAV, thereby significantly reducing the failure rate of the tilting mechanism. At the same time, the design of driving the gear to rotate through the rotation of the worm structure has a self-locking effect and will not be affected by disturbances before and after tilting, thus ensuring the smoothness of the tilting process.

[0048] (2) Compared with traditional UAVs, the present invention adds a power source, has stronger load-carrying capacity and a longer flight range, and improves the reliability of the aircraft in level flight. Assuming that an accident occurs to one propeller during level flight, it can be handled through the cooperation of other propellers. For example, if the left front propeller has a problem, the yaw moment can be balanced by closing the right front propeller, and the control of the UAV can be completed through the two tilting propellers at the rear and the propulsion motor, and the flight mission can continue to be completed.

[0049] (3) Through the design of two front propellers and two rear propellers, the present invention enables the UAV to have two take-off and landing methods, namely vertical take-off and landing and horizontal take-off and landing. When there is a runway, the UAV can choose the horizontal take-off and landing method with less energy consumption, and the load-carrying capacity is greatly improved. Compared with traditional UAVs, it can fly at high speed in level flight and has stronger air maneuverability.

[0050] (4) The present invention is provided with a rudder at the wingtip winglet, which can correct the disturbance of the UAV in the case of no tail wing. It can not only adjust the flight attitude by adjusting the rotation speeds of the front propeller and the two rear propellers, but also correct the flight attitude through the wingtip rudder surface, improving the flight stability. The design of the wingtip winglet can effectively reduce and alleviate the wingtip effect, ensuring that the lift loss at the wingtip is small.

[0051] (5) The present invention adopts a tailless design and unique front tilting mechanism and rear tilting mechanism. Compared with traditional tilting UAVs that tilt the entire wing surface, the wing area is reduced, and the geometric cross-sectional area is reduced, resulting in a lower radar cross-sectional area of the aircraft, which is beneficial to the stealth performance and reduces the possibility of being detected by the enemy, with good stealth effect.

[0052] The technical solution of the present invention will be further described in detail below through the attached drawings and embodiments. Description of the Drawings

[0053] Figure 1 It is a schematic diagram of the overall structure of an embodiment of the present invention;

[0054] Figure 2 It is a schematic diagram of the internal structure of the UAV in an embodiment of the present invention;

[0055] Figure 3 It is a top view of an embodiment of the present invention;

[0056] Figure 4Schematic diagram of the forward tilting mechanism according to an embodiment of the present invention;

[0057] Figure 5 Schematic diagram of the wing skeleton structure according to an embodiment of the present invention;

[0058] Figure 6 Schematic diagram of the coordinate axes established according to an embodiment of the present invention;

[0059] Figure 7 Schematic diagram of the method flow according to an embodiment of the present invention.

[0060] Reference numerals

[0061] 1, fuselage; 11, tilting device fixing bracket; 12, fuselage skeleton; 2, wing; 21, main carbon fiber tube beam frame; 22, auxiliary carbon fiber tube beam frame; 23, aileron; 24, flap; 25, wing rib; 26, winglet; 261, winglet rib; 262, winglet carbon fiber beam frame; 263, rudder; 3, forward tilting mechanism; 31, tilting motor; 32, worm; 33, tilting gear rod; 34, bearing; 35, fairing; 36, fixing box; 37, tilting frame; 38, front propeller; 39, servo motor; 310, rear propeller; 4, rear tilting mechanism; 5, propulsion motor; 6, front landing gear; 7, left landing gear; 8, right landing gear; 9, flight control equipment; 10, battery; 20, mission payload. Detailed implementation manners

[0062] In order to make the objectives, technical solutions and advantages of the embodiments disclosed in the present invention clearer and more understandable, the following further describes the embodiments of the present invention in detail with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the embodiments of the present invention, and are not used to limit the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments in this application without creative efforts belong to the scope of protection of this application. Examples of the embodiments are shown in the accompanying drawings, where the same or similar reference numerals represent the same or similar elements or elements with the same or similar functions throughout.

[0063] It should be noted that the terms "include" and "have" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, system, product or server that includes a series of steps or units does not necessarily have to be limited to those steps or units clearly listed, but may include other steps or units that are not clearly listed or are inherent to these processes, methods, products or devices.

[0064] Similar reference numerals and letters represent similar items in the following drawings. Therefore, once an item is defined in one drawing, it does not need to be further defined and explained in subsequent drawings.

[0065] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by the terms "upper", "lower", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings, or the orientation or positional relationship in which the inventive product is usually placed during use. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation to the present invention.

[0066] In the description of the present invention, it should also be noted that unless otherwise clearly specified and defined, the terms "set", "installed", "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0067] Embodiment

[0068] As Figure 1 shown, a tilting multi-rotor unmanned aerial vehicle according to the present invention includes a fuselage 1, and wings 2 are symmetrically arranged on both sides of the fuselage 1. As Figure 2 、 Figure 6 shown, a fuselage frame 12 for support is arranged inside the fuselage 1, and ailerons 23 and flaps 24 are arranged at the rear ends of the wings 2. The ailerons 23 are used for the roll maneuver of the unmanned aerial vehicle, and the flaps 24 are used for the takeoff and landing of the unmanned aerial vehicle.

[0069] As Figure 2 、 Figure 5 shown, a wing frame is arranged inside the wing 2. The wing frame includes a main carbon fiber tube beam frame 21, an auxiliary carbon fiber tube beam frame 22, and a plurality of wing ribs 25. The main carbon fiber tube beam frame 21 and the auxiliary carbon fiber tube beam frame 22 are respectively arranged along the length direction of the wing 2, and the wing ribs 25 are arranged between the main carbon fiber tube beam frame 21 and the auxiliary carbon fiber tube beam frame 22 and are arranged at intervals along the length direction of the wing 2.

[0070] A winglet 26 with a rudder 263 is arranged at the end of the wing 2. The rudder 263 is arranged at the rear end of the winglet 26, and a winglet frame is arranged inside the winglet 26. The winglet frame includes winglet ribs 261 and a winglet carbon fiber beam frame 262. The winglet carbon fiber beam frame 262 is arranged along the length direction of the winglet 2, and the winglet ribs 261 are arranged along the width direction of the winglet 2.

[0071] As Figure 1 、 Figure 3 、Figure 4 As shown in the figure, a forward tilting mechanism 3 symmetrically arranged with respect to the fuselage 1 is provided at the front end of the fuselage 1, and a rear tilting mechanism 4 symmetrically arranged with respect to the fuselage 1 is provided at the rear end of the fuselage 1. The forward tilting mechanism 3 includes two front propellers 38 symmetrically arranged with respect to the fuselage 1, and the front propellers 38 are connected to the fuselage 1 through a tilting power structure. The rear tilting mechanism 4 includes two rear propellers 310 symmetrically arranged with respect to the fuselage 1, and the rear propellers 310 are connected to the fuselage 1 through a tilting power structure. The fuselage 1 is provided with a tilting device fixing frame 11, and the front propellers 38 and the rear propellers 310 are respectively connected to the tilting device fixing frame 11 through a tilting power structure. Fairings 35 are provided at the front ends of the front propellers 38 and the rear propellers 310. The tilting power structure includes a tilting frame 37. A servo motor 39 is provided on the mounting surface of the tilting frame 37. The output end of the servo motor 39 is connected to the front propeller 38 or the rear propeller 310. The frame plates on both sides of the tilting frame 37 are connected to the box body through a tilting connection assembly, and the box body is connected to the tilting device fixing frame 11. The tilting connection assembly includes a worm 32 provided inside the box body. One end of the worm 32 is connected to the output shaft of the tilting motor 31, and the other end of the worm 32 is rotatably connected to the box body through a bearing 34. A fixing box 36 for fixing the tilting motor 31 is provided inside the box body. The middle of the worm 32 is connected to the middle of the tilting gear rod 33. The two ends of the tilting gear rod 33 respectively pass through the box body and are connected to the frame plates on both sides of the tilting frame 37. The two ends of the tilting gear rod 33 are connected to the box body through bearings 34.

[0072] A front landing gear 6 is provided at the front end of the bottom of the fuselage 1. A left landing gear 7 and a right landing gear 8 are symmetrically arranged on both sides of the bottom of the fuselage 1. A flight control device 9, a mission payload 20 and a battery 10 are provided inside the fuselage 1. A propulsion motor 5 is provided at the rear end of the fuselage 1, and the output shaft of the propulsion motor 5 is connected to a propulsion propeller in the middle of the rear end of the fuselage 1. The flight control device 9 integrates a displacement sensor, a phase sensor, an acceleration sensor, a gyroscope, etc.

[0073] The flight control device 9 and the mission payload 20 adopt the conventional settings of existing unmanned aerial vehicles. The battery 10, the propulsion motor 5, the tilting motor 31, the servo motor 39 and other electric control components on the unmanned aerial vehicle are electrically connected to the flight control device 9 by using an existing structure.

[0074] The front propellers 38 and the rear propellers 310 of the unmanned aerial vehicle can be driven to rotate by the servo motor 39. The tilting motor 31 can drive the tilting frame 37 to rotate through the worm 32 and the tilting gear rod 33, so as to realize the individual adjustment of the tilting angle of each of the two front propellers 38 and the two rear propellers 310.

[0075] As Figure 7 shown, a tilting transition method for a tilting multi-rotor unmanned aerial vehicle according to the present invention includes the following steps:

[0076] Step 1: Perform signal processing in real time and sequentially: As Figure 6 shown, taking the center of gravity of the drone as the coordinate origin O, the chord direction of the drone as the x-axis, the plane passing through the chord and horizontal to the forward direction of the drone as the chord plane, the direction perpendicular to the chord plane of the drone as the y-axis, and the direction perpendicular to both the x-axis and the y-axis as the z-axis.

[0077] Obtain the x-axis acceleration u, y-axis acceleration v, and z-axis acceleration w of the drone from the acceleration sensor, and the pitch angular velocity ω3, pitch angle θ3, roll angular velocity ω1, roll angle θ1, yaw angular velocity ω2, and yaw angle θ2 of the flight state of the drone from the gyroscope. Calculate the components F x 、F y 、F z of the resultant force acting on the drone during flight on the x-axis, y-axis, and z-axis, and the pitch resultant moment M3, roll resultant moment M1, and yaw resultant moment M2 generated by the sum of the tensions of the rotors (front propeller 38 and rear propeller 310) and the flight aerodynamic force.

[0078] By integrating the acceleration, further obtain the forward flight speed V X in the horizontal direction, the vertical speed V Y , and the yaw speed V Z .

[0079] In Step 1, the following formulas are used to calculate the flight parameters of the drone:

[0080] F x = F 推 + mgsin(θ3)+ F 阻 cos(β);

[0081] F y = F 升 + mgcos(θ3)+ F 阻 sin(β);

[0082]

[0083] u = F x / m;

[0084] v = F y / m;

[0085] w = F z / m;

[0086] V X = ∫ucos(θ3)+ vsin(θ3)dt;

[0087] V Y = ∫usin(θ3)+ vcos(θ3)dt;

[0088] V Z = ∫ w cos(θ1) + v sin(θ1) dt;

[0089]

[0090] H = ∫ V Y dt;

[0091]

[0092] Wherein, M 偏 is the yaw moment, F x , F y , F z respectively represent the components of the resultant force received by the UAV during vertical takeoff and landing on the x-axis, y-axis, and z-axis. The angles within the trigonometric functions in the above formulas are all implicit functions, and their function values change with time. Therefore, implicit function integration should be performed on them during integration. F 推 is the thrust of the UAV along the chord direction (x-axis); F 升 is the lift force in the direction perpendicular to the chord plane (y-axis); F 阻 is the drag force in the direction of the actual oncoming flow, the β angle is the angle between the actual oncoming flow direction and the chord (flight angle of attack), I1, I2, and I3 are the moments of inertia of the airframe on the x-axis, y-axis, and z-axis respectively, is the forward displacement of the UAV in the V X direction, H is the forward displacement of the UAV in the V Y direction, that is, the flight altitude, is the forward displacement of the UAV in the V Z direction, is and sum, S is modulus, representing the actual flight distance; R is the projection of the actual distance between the lift center of wing 2 and the center of gravity on the x-axis. In this UAV, the lift center is on the negative half-axis of the x-axis, that is, the lift center is closer to the tail compared to the center of gravity. When the flap 24 is adjusted, the relative positions of the lift center and the center of gravity are not changed.

[0093] Step 2, adjust the attitude of the UAV according to the above calculations. By adjusting the servo motors 39 on the four tilting power structures, the vertical speed V Y and the pitching resultant moment M3 are adjusted, so that when the UAV is not flying horizontally, the forward flight speed V X tends to 0 until it reaches a flight altitude of 50 m.

[0094] In Step 2, the propulsion motor 5 is not started. The component of the resultant force received by the UAV during vertical takeoff and landing on the y-axis except for gravity is the gravity mg and the downward drag force F 阻The sum is then used to obtain the vertical acceleration v of the drone at this time. The formula is as follows:

[0095] F 阻 = kV Y 2 ;

[0096] F y = mg + kV Y 2 ;

[0097]

[0098] where m is the mass of the fuselage 1 and k represents the drag constant;

[0099] According to the Newton - Euler method in classical mechanics, a suitable dynamic mathematical model is constructed:

[0100]

[0101] In the above formula, the control of the drone depends on the roll pitch and yaw angular velocities, and their influence on the attitude. The front propellers 38 and the rear propellers 310 are both propellers with the same structure. The rotational speed of the propeller Ω(t)= -Ω1(t)+Ω2(t)-Ω3(t)+Ω4(t) determines the lift and torque. The driving torques T1(t), T2(t), T3(t) determine the horizontal, vertical movement and turning of the drone. The moment of inertia I R reflects the anti - rotation ability of the propeller. I1, I2, I3 are the moments of inertia of the fuselage about the x - axis, y - axis, and z - axis respectively. These factors together determine the dynamic characteristics and response ability of the drone, especially during rapid maneuvers or against wind resistance.

[0102] According to the above dynamic mathematical model, in step two, the rotational speeds Ω1(t), Ω2(t), Ω3(t), Ω4(t) of the two front propellers 38 and the two rear propellers 310 are adjusted. The flight attitude is adjusted by changing the rotational speeds of the front propellers 38 and the rear propellers 310. By changing the rotational speeds of the propellers, the balance of the moments in the left - right, front - back directions of the drone can be achieved; at the same time, on the basis of changing the rotational speeds, the rotational speeds of the servo motors 39 are further synchronously changed to enable the drone to obtain the vertical acceleration v, so as to achieve the purpose of gradually raising the drone.

[0103] During this process, the propulsion motor 5 does not start, and the overall state is vertical take - off and landing. Except for disturbances, the forward flight acceleration u and the horizontal flight speed V X are 0, and the pitch angle θ3 and the roll angle θ1 are also made as close to 0 as possible after control and adjustment.

[0104] Step 3: After reaching 50 m according to the above flight process, start the propulsion motor 5, and change the forward flight speed V of the UAV in the horizontal direction by changing the thrust. X , as V X increases, the lift of the wing surface of the wing 2 increases accordingly. Gradually reduce the rotation speed of the servo motors 39 on the four tilt power structures. Judge the lift change of the UAV through the change of the forward flight speed V X in the horizontal flight direction, and adjust the balance of the UAV in the overall y-axis and z-axis directions according to the flight control device 9 by controlling the rudder 263 on the winglet 26, the aileron 23 on the wing 2, and the rotation speed of the servo motors 39 on the four tilt power structures until the rotation speed of the servo motors 39 on the four tilt power structures drops to 0.

[0105] There are differences between the processes of Step 2 and Step 3. In this process, with the start of the propulsion motor 5, the lift gradually changes from being completely provided by the original propellers to being provided partly by the propellers and partly by the wing surface of the wing 2, which causes changes in the control and power of the UAV accordingly.

[0106] In Step 3, when the propulsion motor 5 starts and the speed of the servo motor 39 finally becomes 0, the lift received by the UAV is the gravity, and the flight horizontal direction speed V of the UAV at this time is obtained. X .

[0107] Step 4: Drive the tilt frame 37 to rotate through the worm 32, and then adjust the positions of the front propeller 38 and the rear propeller 310, so that the UAV transitions from the vertical state to the horizontal state. Restart the servo motors 39 on the four tilt power structures through the flight control device 9, and drive the UAV to fly through a total of five power sources of two pulls and three pushes.

[0108] In the above steps, the transition balance controller adopts a PID controller. Among them, the tilt motors 31, the servo motors 39, and the rudders 263 all adopt the PID strategy for control. Real-time compensation is carried out for the control signal vibration caused by the transition switching of the tilt-wing UAV. The above method can suppress the signal jitter caused by the switching. The UAV of the present invention has three modes, namely the vertical takeoff and landing mode, the transition flight mode, and the cruise flight mode, which are specifically as follows:

[0109] Vertical takeoff and landing mode: Before flight, during vertical takeoff and landing, the output shafts of the four propellers are perpendicular to the longitudinal axis y of the UAV to provide sufficient lift, that is, perpendicular to the ground. When the UAV vertically ascends, the lift of the UAV is generated by the four tilt-rotor propellers (two front propellers 38 and two rear propellers 310). The pitch stability of the UAV is achieved by PID controlling the y-axis thrust difference between the front and rear groups of two tilt motors 31 of the four tilt-rotor propellers, and the roll stability control is achieved by PID controlling the y-axis thrust difference between the left and right groups of two tilt motors 31 of the tilt-rotor propellers.

[0110] The heading control of the UAV is achieved by controlling the rotational speed and direction of the diagonal servo motors 39 of the four propellers, that is, one set of diagonal servo motors 39 reduces the rotational speed, and the other set of diagonal servo motors 39 increases the rotational speed to generate a z-axis torque while compensating for the lift.

[0111] Transition flight mode: After the UAV vertically takes off and reaches the specified altitude of 50 m, it enters the transition flight mode. The propulsion motor 5 starts, and the four tilting-rotor propellers do not tilt. The thrust direction of the propellers still remains along the Y-axis. The UAV generates an X-axis acceleration and speed. At the same time, the wings 2 gradually start to provide lift, and the four tilting motors 31 gradually decelerate to complete the transition of the lift source. At this time, the pitch stability of the UAV is adjusted by controlling the left and right flaps 24 of the UAV through PID to adjust the lift magnitude of the wings 2, and at the same time, a pitch moment is formed with the lift center of the four-rotor propellers (the lift center of the four-rotor propellers coincides with the center of gravity of the UAV), so as to dynamically adjust the pitch angle. The roll stability of the UAV obtains the moment in the x-axis direction of the flight direction by controlling the left and right ailerons 23 of the UAV through PID, so as to dynamically adjust the roll angle. The heading stability of the UAV obtains the moment in the normal direction of the wings 2 by controlling the left and right winglets of the UAV through PID, so as to dynamically adjust the heading angle.

[0112] In the transition flight mode, the four-rotor tilting motors 31 simultaneously execute the PID control program of the vertical takeoff and landing mode. In this way, during the entire transition flight mode, the stable control of all directions of the UAV is hyperstatically stable. In the transition flight mode, the lift of the UAV in the early stage mainly comes from the four tilting-rotor propellers.

[0113] Then the PID stable control effect of the four tilting-rotors is better than that of the PID stable control of the wings 2, that is, the PID stable control of the four tilting-rotors is the main stable control, and the PID stable control of the wings 2 is the auxiliary stable control; in the later stage, the lift of the UAV mainly comes from the wings 2, then the PID stable control effect of the wings 2 is better than that of the PID stable control of the four tilting-rotors, that is, the PID stable control of the wings 2 is the main stable control, and the PID stable control of the four tilting-rotors is the auxiliary stable control; in this way, the transition flight mode completes the transition of the lift source of the UAV and also completes the transition of the stable control of the UAV.

[0114] Cruise flight mode: After the UAV completes the transition flight mode, the wings 2 provide all the lift, and the four tilting-rotor propellers stop rotating. At this time, the tilting frame 37 is rotated by the rotation of the worm gear so that the thrust direction of the four tilting-rotor propellers is the same as that of the propulsion motor 5, both along the X-axis of the UAV flight direction. After the tilting is completed, the four-rotor tilting motors 31 are started, and the propulsion motor 5 and the four tilting motors 31 jointly provide the thrust in the X-axis direction to complete the thrust distribution.

[0115] In the cruise flight mode, the PID stabilization control system of the four tilt-rotor propellers stops operating, and the wing 2 PID stabilization control manages all the stabilization control of the UAV.

[0116] Assume that a disturbance is encountered. The flight control device 9 adjusts the lift of the aircraft through the two flight control systems to ensure its flight stability. The two flight control systems are the fixed wing and the four tilt-rotor propellers. The fixed wing refers to the wing structure part, which is mainly controlled by adjusting the control surfaces on the flaps 24, ailerons 23 and winglets 26 at the wingtips. The lift is changed by adjusting the propeller speed, so as to adjust the flight attitude by using the lift difference brought by the four power sources.

[0117] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them. Although the present invention has been described in detail with reference to the preferred embodiments, those of ordinary skill in the art should understand that they can still modify or equivalently replace the technical solutions of the present invention, and these modifications or equivalent replacements cannot make the modified technical solutions deviate from the spirit and scope of the technical solutions of the present invention.

Claims

1. A tilting transition method for a tilting multi-rotor unmanned aerial vehicle, characterized in that: The tilt-rotor multi-rotor UAV includes a fuselage, wings symmetrically arranged on both sides of the fuselage, winglets with rudders at the ends of the wings, a front tilt-rotor mechanism symmetrically arranged about the fuselage at the front end of the fuselage, a rear tilt-rotor mechanism symmetrically arranged about the fuselage at the rear end of the fuselage, and a propeller at the middle of the rear end of the fuselage. The front tilt-rotor mechanism includes a front propeller, and the front propeller is connected to the fuselage through a tilt power structure. The rear tilt-rotor mechanism includes a rear propeller, and the rear propeller is connected to the fuselage through a tilt power structure. The tilt transition method of the tilt-rotor multi-rotor UAV includes the following steps: Step 1, perform signal processing in real time and sequentially: Take the center of gravity of the UAV as the coordinate origin O, the chord direction of the UAV as the x-axis, the plane passing through the chord and horizontal to the forward direction of the UAV as the chord plane, the direction perpendicular to the chord plane of the UAV as the y-axis, and the direction perpendicular to both the x-axis and the y-axis as the z-axis. According to the acceleration sensor, obtain the x-axis acceleration u, y-axis acceleration v, and z-axis acceleration w of the drone. The gyroscope obtains the pitch angular velocity ω3, pitch angle θ3, roll angular velocity ω1, roll angle θ1, yaw angular velocity ω2, and yaw angle θ2 of the drone's flight state; calculate the components F x , F y , F z on the x-axis, y-axis, and z-axis of the resultant force received by the drone during flight, and the pitch resultant moment M3, roll resultant moment M1, and yaw resultant moment M2 generated by the sum of the rotor thrust and flight aerodynamic force; the front propeller and the rear propeller are rotors; By integrating the acceleration, the forward flight velocity V in the horizontal direction is further obtained X , the velocity V in the vertical direction Y , and the yaw velocity V Z ; Step 2: Adjust the attitude of the UAV according to the above calculations. Adjust the vertical velocity V by adjusting the servo motors on the four tilt power structures Y and the pitching resultant moment M3, so that when the UAV is not flying horizontally, the forward flight velocity V X tends to 0 until it reaches a flight altitude of 50 m; Step 3: After the flight altitude reaches 50 m, start the propulsion motor and change the forward flight speed V of the UAV in the horizontal direction by changing the thrust. X , as V X increases, gradually reduce the rotational speeds of the servo motors on the four tilting power structures. Judge the change in the lift of the UAV through the change in the forward flight speed V X in the horizontal direction of flight, and adjust the balance of the UAV in the overall y-axis and z-axis directions according to the flight control equipment by controlling the rudder on the winglet, the aileron on the wing, and the rotational speeds of the servo motors on the four tilting power structures until the rotational speeds of the servo motors on the four tilting power structures drop to 0. Step 4, drive the tilt frame to rotate through the worm, thereby adjusting the positions of the front propeller and the rear propeller, so that the UAV transitions from the vertical state to the horizontal state. Restart the servo motors on the four tilt power structures through the flight control device, and drive the UAV to fly through a total of five power sources, two pulling and three pushing.

2. The tilting transition method of the tilt-rotor multi-rotor UAV according to claim 1, wherein: The fuselage is internally provided with a fuselage skeleton for support.

3. The tilting transition method of the tilt-rotor multi-rotor UAV according to claim 1, characterized in that: Ailerons and flaps are provided at the rear end of the wing. The ailerons are used for the roll maneuver of the UAV, and the flaps are used for the takeoff and landing of the UAV.

4. The tilting transition method of the tilt-rotor multi-rotor UAV according to claim 1, characterized in that: The wing is internally provided with a wing skeleton, which includes a main carbon fiber tube beam frame, an auxiliary carbon fiber tube beam frame, and several wing ribs. The main carbon fiber tube beam frame and the auxiliary carbon fiber tube beam frame are respectively arranged along the length direction of the wing. The wing ribs are arranged between the main carbon fiber tube beam frame and the auxiliary carbon fiber tube beam frame and are arranged at intervals along the length direction of the wing.

5. The tilting transition method of the tilt-rotor multi-rotor UAV according to claim 1, characterized in that: The rudder is provided at the rear end of the winglet. The winglet is internally provided with a winglet skeleton, which includes winglet ribs and a winglet carbon fiber beam frame. The winglet carbon fiber beam frame is arranged along the length direction of the winglet, and the winglet ribs are arranged along the width direction of the winglet.

6. The tilt transition method of the tilt-rotor multi-rotor UAV according to claim 1, characterized in that: The fuselage is provided with a tilt device fixing frame. The front propeller and the rear propeller are respectively connected to the tilt device fixing frame through a tilt power structure. Fairings are provided at the front ends of the front propeller and the rear propeller. The tilt power structure includes a tilt frame. A servo motor is arranged on the mounting surface of the tilt frame. The output end of the servo motor is connected to the front propeller or the rear propeller. The frame plates on both sides of the tilt frame are connected to the box body through a tilt connection component, and the box body is connected to the tilt device fixing frame.

7. The tilting transition method of the tilt-rotor multi-rotor UAV according to claim 1, wherein: The tilt connection component includes a worm arranged inside the box body. One end of the worm is connected to the output shaft of the tilt motor. A fixing box for fixing the tilt motor is arranged inside the box body. The other end of the worm is rotatably connected to the box body through a bearing. The middle of the worm is connected to the middle of the tilt gear rod. The two ends of the tilt gear rod respectively pass through the box body and are connected to the frame plates on both sides of the tilt frame.

8. The tilt transition method of the tilt-rotor multi-rotor UAV according to claim 1, wherein: The front landing gear is provided at the front end of the bottom of the fuselage. The left landing gear and the right landing gear are symmetrically provided on both sides of the bottom of the fuselage. The flight control equipment, mission payload and battery are arranged inside the fuselage. The propulsion motor is provided at the rear end of the fuselage, and the output shaft of the propulsion motor is connected to the propulsion propeller. The flight control equipment integrates a displacement sensor, a phase sensor, an acceleration sensor and a gyroscope.

9. The tilting transition method of the tilting multi-rotor UAV according to claim 1, characterized in that: In step one, the following formula is used to calculate the flight parameters of the UAV: F x = F 推 + mgsin(θ3) + F 阻 cos(β); F y = F 升 + mg cos(θ3) + F 阻 sin(β); u = F x / m; v = F y / m; w = F z / m; V X = ∫ u cos(θ3) + v sin(θ3) dt; V Y = ∫ u sin(θ3) + v cos(θ3) dt; V Z = ∫ w cos(θ1) + v sin(θ1) dt; H = ∫V Y dt; Among them, M 偏 is the yaw moment, F x , F y , F z respectively represent the components of the resultant force received by the UAV during vertical takeoff and landing on the x-axis, y-axis, and z-axis. The angles within the trigonometric functions in the above formula are all implicit functions. F 推 is the thrust of the UAV along the chord direction; F 升 is the lift in the direction perpendicular to the chord plane; F 阻 is the drag in the direction of the actual oncoming flow. The β angle is the angle between the actual oncoming flow direction and the chord. I1, I2, and I3 are the moments of inertia of the airframe about the x-axis, y-axis, and z-axis respectively, is the forward displacement of the UAV in the V X direction, H is the forward displacement of the UAV in the V Y direction, is the forward displacement of the UAV in the V Z direction, is plus , S is the modulus of , representing the actual flight distance; R is the projection of the actual distance between the lift center and the center of gravity of the wing on the x-axis. The lift center of the UAV is on the negative half-axis of the x-axis; In step 2, when the propulsion motor is not started, the component of the resultant force on the UAV during vertical takeoff and landing in the y-axis direction, excluding gravity, is the sum of the gravity mg and the downward resistance F, and then the vertical acceleration v of the UAV at this time is obtained. The formula is as follows: 阻 The sum, and then the vertical acceleration v of the UAV at this time is obtained. The formula is as follows: F 阻 = kV Y 2 ; F y = mg + kV Y 2 ; where m is the mass of the fuselage and k represents the drag constant; According to the Newton-Euler method in classical mechanics, a suitable dynamic mathematical model is constructed: In the above formula, the control of the drone depends on roll pitch and yaw angular velocities, and their influence on the attitude. The front and rear propellers are of the same structure. The rotational speed of the propeller Ω(t) = -Ω1(t) + Ω2(t) - Ω3(t) + Ω4(t) determines the lift and torque. The driving torques T1(t), T2(t), T3(t) determine the horizontal, vertical movement and turning of the drone, and the moment of inertia I R reflects the anti-rotation ability of the propeller; According to the above dynamic mathematical model, in step two, the rotational speeds Ω1(t), Ω2(t), Ω3(t), and Ω4(t) of the two front propellers and the two rear propellers are adjusted, and the flight attitude is adjusted by changing the rotational speeds of the front propellers and the rear propellers. In step 3, the propulsion motor starts, the speed of the servo motor finally becomes 0, the lift force on the drone is equal to the gravity, and the horizontal flight speed V of the drone at this time is obtained. X .

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