A rotor mounting assembly, a multi-rotor aircraft and a control method thereof

Abstract

The present invention provides a rotor mounting assembly, which includes a rotating shaft, a pitch control module, and two blade mounting bases symmetrically mounted on the rotating shaft, wherein: the pitch control module includes an electromagnetic eddy current brake and a first gear coaxially mounted on the rotating shaft; the blade mounting base includes a blade mounting clip, a mounting sleeve perpendicular to the rotating shaft, and a second gear provided on the outer peripheral wall of the mounting sleeve; the rotation axis of the second gear is perpendicular to the rotation axis of the first gear. The present invention also provides a multi-rotor aircraft, which includes a main body, at least two of the above-mentioned rotor mounting assemblies, and blades mounted on the rotor mounting assemblies. The present invention also provides a control method for a multi-rotor aircraft. The present invention realizes pitch control through an electromagnetic eddy current brake, a first gear, and a second gear whose rotation axis is perpendicular to the rotation axis of the first gear.

Description

Technical Field

[0001] The present invention relates to an aircraft, and in particular to a rotor mounting assembly, a multi-rotor aircraft and a control method thereof. Background Art

[0002] A multi-rotor aircraft refers to a helicopter-like aircraft with three or more rotors, which can take off and land vertically. It belongs to a kind of helicopter aircraft and is generally called a multi-rotor aircraft. A multi-rotor aircraft belongs to an unstable system and cannot achieve self-stabilization during flight and hovering.

[0003] Multi-rotor aircraft can be divided into fixed-pitch aircraft and variable-pitch aircraft; the so-called variable pitch means adjusting the flight attitude of the aircraft by adjusting the angle of the blades. Since it can more stably control the flight attitude of the aircraft, it has gradually become a research hotspot.

[0004] The existing variable-pitch aircraft structure usually uses a series of mechanical transmission structures such as hydraulic tie rods or electric servo tie rods to drive, but hydraulics is not suitable for quickly adjusting the pitch, and the electric servo tie rod transmission structure is prone to failure during long-term use, has poor reliability, and it is difficult to achieve maintenance-free. Summary of the Invention

[0005] The features and advantages of the present invention are partially stated in the following description, or may be obvious from the description, or may be learned by practicing the present invention.

[0006] To overcome the problems of the prior art, the present invention provides a rotor mounting assembly, including a rotating shaft, a pitch control module, and two blade mounting bases symmetrically mounted on the rotating shaft, wherein:

[0007] The pitch control module includes an electromagnetic eddy current brake and a first gear coaxially mounted on the rotating shaft;

[0008] The blade mounting base includes a blade mounting clip, a mounting sleeve perpendicular to the rotating shaft, and a second gear provided on the outer peripheral wall of the mounting sleeve; the rotation axis of the second gear is perpendicular to the rotation axis of the first gear.

[0009] Preferably, the first gear and the second gear are bevel gears.

[0010] Preferably, the first gear is an annular gear track.

[0011] Preferably, the electromagnetic eddy current brake includes an eddy current brake disc and an excitation coil, and the first gear is fixedly connected to the eddy current brake disc.

[0012] Preferably, the blade mounting base further includes a deflecting member, and the deflecting member is fixed on the mounting sleeve.

[0013] Preferably, the deflector is disposed between the second gear and the blade mounting clip.

[0014] Preferably, the center line of the deflector is perpendicular to the center line of the mounting sleeve.

[0015] The present invention also provides a multi-rotor aircraft, including a main body, at least two of the above-mentioned rotor mounting assemblies mounted on the main body, and blades mounted on the rotor mounting assemblies.

[0016] The present invention also provides a control method for a multi-rotor aircraft. The multi-rotor aircraft includes at least two of the above-mentioned rotor mounting assemblies and blades mounted on the rotor mounting assemblies, and the method includes the steps of:

[0017] Outputting the thrust corresponding to each axis according to the target attitude and the actual attitude of the aircraft during flight;

[0018] Obtaining the thrust change amount according to the thrust;

[0019] Changing the excitation intensity of the electromagnetic eddy current brake according to the thrust change amount.

[0020] Preferably, after changing the excitation intensity of the electromagnetic eddy current brake according to the thrust change amount, it further includes:

[0021] Obtaining the attitude information of the blade and adjusting the excitation intensity of the electromagnetic eddy current brake accordingly.

[0022] The present invention realizes pitch control through the first gear and the second gear that are vertically meshed and transmitted, thereby meeting the needs of attitude control.

[0023] Those of ordinary skill in the art will better understand the features and contents of these technical solutions by reading the specification. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] The present invention will be specifically described below with reference to the accompanying drawings and in combination with examples. The advantages and implementation manners of the present invention will become more obvious. The content shown in the accompanying drawings is only used for the explanation of the present invention and does not constitute any limitation to the present invention. In the accompanying drawings:

[0025] Figure 1 is a schematic structural diagram of a rotor mounting assembly according to an embodiment of the present invention.

[0026] Figure 2 is a schematic diagram of a blade according to an embodiment of the present invention.

[0027] Figure 3 is a schematic structural diagram of a multi-rotor aircraft according to an embodiment of the present invention.

[0028] Figure 4 Schematic flow chart of the control method for a multi-rotor aircraft according to an embodiment of the present invention.

[0029] Reference numerals in the drawings:

[0030] 11. Rotation axis; 12. Cross bar; 20. Blade mounting base; 21. Mounting sleeve; 22. Second gear; 23. Blade mounting clip; 24. Deflection member; 25. Blade; 30. Pitch control module; 31. Excitation coil; 32. Electromagnetic eddy current brake; 33. Eddy current brake disc; 34. First gear; 41. Attitude detection module; 42. Wireless communication module; 43. Flight controller; 44. Pitch controller. Detailed implementation manners

[0031] In the description of the present invention, unless otherwise specified, "a plurality of" means two or more; the terms "upper", "lower", "inner", "outer", "top", "bottom", etc. indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, and are only for facilitating the description of 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 thus should not be construed as a limitation to the present invention. In addition, the terms "first", "second", etc. are only used for descriptive purposes and cannot be construed as indicating or implying relative importance.

[0032] In the description of the present invention, it should be noted that unless otherwise clearly defined and limited, the terms "connected" and "coupled" 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. 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.

[0033] As Figure 1 , Figure 2 shown, the present invention provides a rotor mounting assembly, including a rotation axis 11, a pitch control module 30, and two blade mounting bases 20 symmetrically mounted on the rotation axis 11.

[0034] The rotation axis 11 is connected to a motor (not shown in the figure), and the rotation axis 11 can rotate at a preset speed under the drive of the motor.

[0035] The blade mounting base 20 is mounted on the rotation axis 11. More specifically, the blade mounting base 20 is mounted on the rotation axis 11 through a cross bar 12, the cross bar 12 is perpendicular to the rotation axis 11 and fixedly mounted on the rotation axis 11.

[0036] The variable pitch control module 30 includes an electromagnetic eddy current brake 32 and a first gear 34 coaxially mounted on the rotating shaft 11. The electromagnetic eddy current brake 32 includes an exciting coil 31 and an eddy current brake disc 33. The eddy current brake disc 33 is rigidly connected to the first gear 34. The present invention uses electro-eddy current electromagnetic braking. The rotating shaft 11 rotates, and the exciting coil 31 can be installed on the aircraft body.

[0037] In this embodiment, the blade mounting base 20 includes a mounting sleeve 21, a second gear 22, and a blade mounting clamp 23. The mounting sleeve 21 is perpendicular to the rotating shaft 11. In specific implementation, one end face of the mounting sleeve 21 is fixedly installed on the cross bar 12, and the other end face is fixedly connected to the blade mounting clamp 23. The second gear 22 is arranged on the outer peripheral wall of the mounting sleeve 21, and the rotation axis of the second gear 22 is perpendicular to the rotation axis of the first gear 34. That is to say, the first gear 34 and the second gear 22 are perpendicularly meshed. In this embodiment, the first gear 34 and the second gear 22 are bevel gears meshing with each other. In another embodiment of the present invention, the first gear 34 is an annular gear track, and the second gear 22 is a gear meshing with the annular gear track.

[0038] The blade mounting clamp 23 is used to mount the blade 25, and the blade 25 can be mounted on the blade mounting clamp 23 through bolts.

[0039] When the exciting coil 31 is energized, the magnetic field generated on the exciting coil 31 runs relative to the first gear 34, generating eddy currents, and these eddy currents generate a braking torque that prevents the first gear 34 from rotating. This braking torque causes the blade 25 to rotate in the direction of decreasing pitch.

[0040] In this embodiment, the blade mounting base 20 further includes a deflecting member 24, which is fixed on the mounting sleeve 21. The deflecting member 24 is used to give the blade 25 a deflecting torque. Preferably, the center line of the deflecting member 24 is perpendicular to the center line of the mounting sleeve. In this embodiment, the deflecting member 24 is arranged between the second gear 22 and the blade mounting clamp 23. When the rotating shaft 11 rotates at a certain speed, due to the non-coincidence of the centripetal force and the structural tension line, the deflecting member 24 will give the blade 25 a deflecting torque. It can be seen that the magnitude of the deflecting torque is related to the rotational speed. When the rotational speed is constant, the deflecting torque is a fixed value, so that the blade 25 rotates in the direction of increasing pitch. The weight of the deflecting member 24 is determined according to the deflecting torque required by the blade 25 and the operating rotational speed.

[0041] It can be seen that what ultimately determines the movement direction of the blade 25 is the difference between the deflecting torque generated by the deflecting member 24 and the torque generated by the eddy current brake disc 33. When the weight of the deflecting member 24 is the same, by changing the exciting current intensity, the braking torque of the eddy current brake disc 33 can be changed to achieve the purpose of controlling the movement of the blade 25.

[0042] As shown Figure 3 The present invention provides a multi-rotor aircraft, including a main body, at least two rotor mounting assemblies provided in any embodiment of the present invention, and blades 25 mounted on the rotor mounting assemblies; for example, 3, 4, 6 or 8. Generally, 4 rotor mounting assemblies can be mounted on the main body to form a quad-rotor aircraft, or 8 rotor mounting assemblies can be mounted on the main body to form an octa-rotor aircraft. The present invention does not limit this.

[0043] An attitude detection module 41, a wireless communication module 42, a flight controller 43, and a pitch controller 44 are provided on the main body. The pitch controller 44 is used to control the current intensity of the excitation coil 31. Although not shown in the figure, the flight controller 43 is also connected to a plurality of motors, and the motors are connected to the rotating shaft 11. The rotation speed of the rotating shaft 11 can be controlled by the flight controller.

[0044] Attitude is used to describe the angular and positional relationships between the body-fixed coordinate system of a rigid body and a reference coordinate system. Common description methods include Euler angles, namely the three rotation angles of Pitch, roll, and yaw. In addition, there are also description methods such as quaternions and direction cosine matrices.

[0045] Common attitude measurement devices include gyroscopes, accelerometers, magnetometers, barometers, ultrasonic sensors, GPS, cameras, infrared sensors, optical flow sensors, and so on. In this embodiment, the attitude detection module 41 is composed of a gyroscope, an accelerometer, a barometer, and GPS.

[0046] A gyroscope is an instrument for indirectly measuring angles, and it outputs the angular velocity of the carrier's motion. The angle needs to be obtained by integrating the angular velocity in the time domain. An acceleration sensor is a device for measuring the linear acceleration of a vehicle. By measuring the acceleration caused by gravity, the inclination angle of the vehicle relative to the horizontal plane can be calculated. Continuously correcting the gyroscope with the accelerometer and fusing the attitude data of both can more accurately calculate the current attitude of the aircraft.

[0047] In specific implementation, the data measured by the gyroscope and the accelerometer can be used as inputs; the outputs are the inclination angle and angular velocity relative to the earth coordinate system. The algorithm is specifically introduced as follows:

[0048] Low-Pass Fllter: The original angular velocity data of the accelerometer is low-pass filtered. The purpose is to filter the fluctuations of the acceleration data in the short term, remove the burrs, smooth the data, and in the implementation, the new measurement position can be weighted with the previous calculated value, for example:

[0049] angle=(0.98)*angle+(0.02)*x-acc,

[0050] Among them, x-acc is the currently measured angle, and angle is the angle value calculated each time.

[0051] Numeric Integration: The output of the gyroscope is the angular velocity of the object's rotation. By integrating, a smooth angle can be obtained.

[0052] High-PassFilter: Since the gyroscope measures instantaneous values, for the angle change obtained by integration, high-frequency components need to be filtered out.

[0053] Finally, the angle data obtained after processing the data of the gyroscope and the accelerometer is subjected to a summation operation in a certain proportion to finally obtain the desired attitude angle.

[0054] In an embodiment of the present invention, the attitude monitoring module further includes a pitch encoder. The pitch encoder can transmit the angular position of the blade in real time. The flight controller adjusts the current intensity of the excitation coil according to the angular position of the blade and uses closed-loop control to control the angular position of the blade to the expected value.

[0055] The wireless communication module 42 is used to receive the control instructions sent by the remote controller and thus obtain the target attitude. After receiving the target attitude data and the actual attitude data transmitted by the attitude monitoring module 41 and the wireless communication module 42, the flight controller 43 completes a series of complex algorithms and finally outputs the thrust corresponding to each axis. The pitch controller 44 is used to obtain the thrust change amount according to the thrust and accordingly change the excitation intensity of the electromagnetic eddy current brake, that is, the current intensity of the excitation coil 31. The thrust change amount refers to whether the thrust corresponding to the target attitude needs to be increased, decreased, or maintained compared to the existing thrust.

[0056] As Figure 4 shown, the present invention also provides a control method for a multi-rotor aircraft. The multi-rotor aircraft includes the rotor mounting assembly described in any one of the above and the blades 25 mounted on the rotor mounting assembly. The control method includes the steps:

[0057] 110. Output the thrust corresponding to each axis according to the target attitude and the actual attitude of the aircraft flight;

[0058] In specific implementation, a gyroscope, an accelerometer, and a GPS can be used to monitor the actual attitude of the aircraft flight in real time and obtain the actual attitude data. When calculating the control quantity in combination with the target attitude data, a Kalman filter algorithm, a flight control PID algorithm, a strapdown inertial algorithm, etc. can be used to output the thrust corresponding to each axis. The present invention does not limit this.

[0059] 111. Obtain the thrust change amount according to the thrust;

[0060] The thrust change refers to whether the thrust corresponding to the target attitude needs to be increased, decreased, or maintained compared to the existing thrust.

[0061] 112. Change the excitation intensity of the electromagnetic eddy current brake according to the thrust change.

[0062] The thrust change is converted into a corresponding PWM signal and then passed through the drive circuit to control the current intensity of the excitation coil, thereby maintaining the stable flight of the multi-rotor aircraft.

[0063] In this embodiment, it is not necessary to brake the actual propeller torque. Instead, the thrust output by the flight controller needs to be increased or decreased, and the thrust is maintained to control the width of the PWM pulse of the excitation coil, thereby changing the magnetic field intensity and the braking torque to achieve the purpose of changing the propeller thrust.

[0064] Preferably, after step 112, the method further includes:

[0065] The attitude information of the blade is obtained and the excitation intensity of the electromagnetic eddy current brake is adjusted accordingly.

[0066] In specific implementation, a pitch encoder can be used to transmit the angular position of the blade in real time, the thrust change can be adjusted according to the angular position of the blade, and the current intensity of the excitation coil can be adjusted according to the adjusted thrust change, so as to use closed-loop control to control the angular position of the blade to the expected value.

[0067] The rotor mounting assembly, multi-rotor aircraft and control method thereof provided by the present invention realize variable pitch control through an electromagnetic eddy current brake, a first gear and a second gear perpendicular to the rotation axis of the first gear, and are reliable and easy to maintain.

[0068] The preferred embodiments of the present invention are described above with reference to the accompanying drawings. A person skilled in the art may implement the present invention in a variety of variations without departing from the scope and essence of the present invention. For example, a feature shown or described as part of one embodiment may be used in another embodiment to obtain another embodiment. The above are only preferred feasible embodiments of the present invention, and do not limit the scope of rights of the present invention. Any equivalent changes made using the contents of the present specification and the accompanying drawings are included in the scope of rights of the present invention.

Claims

1. A rotor installation assembly, characterized in that, it includes a rotating shaft, a pitch control module, and two blade mounting bases symmetrically mounted on the rotating shaft, wherein: the pitch control module includes an electromagnetic eddy current brake and a first gear coaxially mounted on the rotating shaft; the blade mounting base includes a blade mounting clip, a mounting sleeve perpendicular to the rotating shaft, and a second gear provided on the outer peripheral wall of the mounting sleeve; the rotation axis of the second gear is perpendicular to the rotation axis of the first gear; the blade mounting base further includes a deflecting member fixed to the mounting sleeve; the movement direction of the blade is determined by the difference between the deflecting moment generated by the deflecting member and the moment generated by the electromagnetic eddy current brake.

2. The rotor installation assembly according to claim 1, characterized in that, the first gear and the second gear are bevel gears.

3. The rotor installation assembly according to claim 1, characterized in that, the first gear is an annular gear track.

4. The rotor installation assembly according to claim 1, characterized in that, the electromagnetic eddy current brake includes an eddy current brake disc and an exciting coil, and the first gear is fixedly connected to the eddy current brake disc.

5. The rotor installation assembly according to claim 1, characterized in that, the deflecting member is arranged between the second gear and the blade mounting clip.

6. The rotor installation assembly according to claim 1, characterized in that, the center line of the deflecting member is perpendicular to the center line of the mounting sleeve.

7. A multi-rotor aircraft, characterized in that, it includes a body, at least two rotor installation assemblies as described in any one of claims 1 to 6 mounted on the body, and blades mounted on the rotor installation assemblies.

8. A control method for a multi-rotor aircraft, the multi-rotor aircraft includes at least two rotor installation assemblies as described in any one of claims 1 to 6 and blades mounted on the rotor installation assemblies, characterized in that, it includes the steps of: outputting the thrust corresponding to each axis according to the target attitude and the actual attitude of the aircraft during flight; obtaining the thrust change amount according to the thrust; changing the exciting intensity of the electromagnetic eddy current brake according to the thrust change amount.

9. The control method for a multi-rotor aircraft according to claim 8, characterized in that, after changing the exciting intensity of the electromagnetic eddy current brake according to the thrust change amount, it further includes: obtaining the attitude information of the blade and adjusting the exciting intensity of the electromagnetic eddy current brake accordingly.

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