WO2023246424A1 - Electrically controlled variable pitch propeller, and multirotor and control method therefor - Google Patents

Abstract

An electrically controlled variable pitch propeller, and a multirotor and a control method therefor. The electrically controlled variable pitch propeller comprises a main shaft (10), a variable pitch blade (20) and a brake module (30), wherein the main shaft (10) comprises a vertical rod (11) and a cross rod (12) which are arranged perpendicularly; the brake module (30) comprises an excitation coil (31), a brake vortex disk (32) and a brake gear (33) which are arranged coaxially, and the brake vortex disk (32) and the brake gear (33) are mounted on the vertical rod (11); and the variable pitch blade (20) comprises a variable pitch bevel gear (21) and a blade mounting clamp (25) connected to the variable pitch bevel gear (21), and the variable pitch bevel gear (21) is connected to the cross rod (12) and meshes with the brake gear (33).

Description

An electronically controlled variable pitch propeller, a multi-axis aircraft and a control method thereof Technical field

The present invention relates to aircraft, in particular to electronically controlled variable pitch propellers, multi-axis aircraft and control methods thereof.

Background technique

At present, multi-copter aircraft are widely used in civilian, commercial and military fields. In the civilian field, more and more extreme sports enthusiasts use multi-copter aircraft to record. In the commercial field, camera equipment is equipped to record various sports events or various places. Tracking aerial photography of the scenery shows that multi-copter aircraft have a wide range of applications and broad market prospects.

A multicopter refers to a helicopter-type aircraft with three or more rotors that can take off and land vertically. It is a type of helicopter aircraft. Multi-copter aircraft are unstable systems and cannot achieve self-stabilization during flight and hovering. Multicopter aircraft can be divided into fixed-pitch aircraft and variable-pitch aircraft; the so-called variable pitch refers to adjusting the flight attitude of the aircraft by adjusting the angle of the blades. Since it can control the flight attitude of the aircraft more stably, it has gradually become Research hotspots.

The take-off weight of electric multicopters is getting larger and larger, and in order to achieve higher efficiency, the diameter of the propellers is getting larger and larger. Under the restriction of ensuring that the propeller tip does not generate shock wave loss, the rotation speed is further reduced, and the traditional multi-rotor fixed-pitch variable speed control is difficult to meet the needs of attitude control. This is because as the diameter of the blade increases, the moment of inertia increases and the control delay becomes serious.

At present, variable pitch control is used to solve this problem. The existing variable pitch blade structure is usually driven by a series of mechanical transmission structures such as hydraulic pull rods or electric steering gear pull rods. However, hydraulic pressure is not suitable for quickly adjusting the pitch, and electric steering gears The tie rod transmission structure is prone to failure during long-term use, has poor reliability, and is difficult to be maintenance-free.

technical problem

The present invention provides an electronically controlled variable pitch propeller, a multi-axis aircraft and a control method thereof to solve the technical problem of achieving attitude control without using a series of mechanical transmission structures such as servos and pull rods.

Technical solutions

The invention provides an electronically controlled variable pitch propeller, which includes a main shaft, variable pitch blades and a braking module, wherein:

The main shaft includes vertical rods and horizontal rods;

The braking module includes a coaxially arranged excitation coil, a braking eddy current disc and a braking gear, and the braking eddy current disc and braking gear are installed on the vertical pole;

The variable-pitch blade includes a variable-pitch bevel gear and a blade mounting clip connected to the variable-pitch bevel gear. The variable-pitch bevel gear is connected to the crossbar and meshes with the brake gear.

The present invention also provides a multi-axis aircraft, including a main body and at least two of the above-mentioned electronically controlled variable pitch propellers; the main body is provided with a control module, a wireless communication module and an attitude measurement module, and the electronically controlled variable pitch propellers and The control modules are connected; wherein:

The wireless communication module is used to receive control signals and send the control signals to the control module;

The attitude measurement module is used to monitor the actual attitude of the aircraft in real time to generate actual attitude data, and send the actual attitude data to the control module;

The control module is used to calculate a control quantity based on the control signal and the actual attitude data, and control the rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller accordingly.

The present invention also provides a control method for a multi-copter aircraft. The multi-copter aircraft includes at least two of the above-mentioned electronically controlled variable pitch propellers. The control method includes the steps:

Receive control signals and calculate target attitude data accordingly;

Obtain the actual attitude data and calculate the control amount combined with the target attitude data;

The rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller are controlled according to the control quantity.

Wherein, after controlling the rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller according to the control amount, it further includes:

Obtain attitude information of the variable pitch blades in the electronically controlled variable pitch propeller and adjust the current intensity of the excitation coil accordingly.

The present invention further provides a control method for a multi-copter aircraft. The multi-copter aircraft includes at least two of the above-mentioned electronically controlled variable pitch propellers. The control method includes the steps:

Output the thrust corresponding to each axis according to the target attitude and actual attitude of the multi-axis aircraft flight;

Obtain the thrust change amount according to the thrust force;

The excitation intensity of the excitation coil is changed according to the thrust variation.

Wherein, after changing the excitation intensity of the electromagnetic eddy current brake according to the thrust variation, the method further includes:

Obtain the attitude information of the blade and adjust the excitation intensity of the excitation coil accordingly.

beneficial effects

Without using a series of mechanical transmission structures such as steering gears and pull rods, the invention achieves variable pitch control by meshing the brake gear of the braking module with the variable pitch bevel gear of the variable pitch blade, thereby satisfying attitude control. needs.

Description of the drawings

The advantages and implementation manners of the present invention will be more obvious when the present invention is described in detail with reference to the accompanying drawings and examples. The contents shown in the accompanying drawings are only used for explanation of the present invention and do not constitute any limitation of the present invention. The limitations, in the attached figure:

Figure 1 is a schematic structural diagram of an electronically controlled variable pitch propeller according to an embodiment of the present invention.

Figure 2 is a schematic structural diagram of a spindle according to an embodiment of the present invention.

Figure 3 is a schematic structural diagram of an electronically controlled variable pitch propeller without blades installed according to an embodiment of the present invention.

Figure 4 is a schematic structural diagram of a propeller blade according to an embodiment of the present invention.

Figure 5 is a schematic structural diagram of a multi-copter aircraft according to an embodiment of the present invention.

FIG. 6 is a schematic flowchart of the control method of the multi-copter aircraft shown in FIG. 5 according to the present invention.

Figure 7 is a schematic structural diagram of a multi-copter aircraft according to an embodiment of the present invention.

FIG. 8 is a schematic flowchart of the control method of the multi-copter aircraft shown in FIG. 7 according to the present invention.

Reference number:

10. Main shaft; 11. Vertical rod; 12. Cross rod; 20. Variable pitch blade; 21. Variable pitch bevel gear; 22. Blade counterweight; 23. Blade; 24. Bushing; 25. Blade Installation clip; 30. Braking module; 31. Excitation coil; 32. Braking eddy current disc; 33. Braking gear; 40. Attitude measurement module; 41. Acceleration sensor; 42. Gyroscope; 43. Geomagnetic sensor; 50 , wireless communication module; 60. control module; 61. motor driver; 62. power driver; 63. main controller; 71. attitude detection module; 72. wireless communication module; 73. flight controller; 74. variable distance controller .

Embodiments of the invention

In the description of the present invention, unless otherwise stated, "plurality" means two or more; the terms "upper", "lower", "inner", "outer", "top", "bottom" The indicated orientations or positional relationships are based on the orientations or positional relationships shown in the drawings. They are only for the convenience of describing the present invention and simplifying the description. They are not intended to indicate or imply that the device or element referred to must have a specific orientation or in a specific manner. orientation construction and operation and therefore should not be construed as limitations of the invention. Furthermore, the terms "first," "second," etc. are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.

In the description of the present invention, it should be noted that, unless otherwise clearly stated and limited, the terms "connected" and "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection. Ground connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection 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.

As shown in Figures 1 to 4, the present invention provides an electronically controlled variable pitch propeller, which includes a main shaft 10, a variable pitch blade 20, and a brake module 30.

The main shaft 10 is composed of a vertical pole 11 and a horizontal bar 12 installed on the top of the vertical pole 11, and the vertical pole 11 and the horizontal bar 12 are perpendicular to each other. The crossbar 12 is used to connect with the variable pitch blade 20 . The vertical pole 11 is connected to a motor (not shown in the figure), and the vertical pole 11 can rotate at a preset speed driven by the motor.

The variable pitch blade 20 includes a blade mounting clip 25 and a variable pitch bevel gear 21 . Among them, one end face of the variable pitch bevel gear 21 is installed on the cross bar 12 through a sleeve or a bearing, and the blade mounting clip 25 is fixed on the other end face of the variable pitch bevel gear 21 through bolts. The variable pitch blade also includes a blade 23, which is installed on a blade mounting clip 25. The blade 23 can be installed on the blade mounting clip 25 through bolts. In one embodiment of the present application, the variable pitch bevel gear 21 is installed on the cross bar 12 using a bushing 24. At this time, one end surface of the bushing 24 is fixedly installed on the cross bar 12, and the other end surface is connected to the blade. The mounting clips 25 are fixedly connected together, and the variable pitch bevel gear 21 is arranged on the outer peripheral wall of the shaft sleeve 24 .

The variable pitch bevel gear 21 is composed of a bevel gear and a blade counterweight portion 22 connected to the bevel gear. The center line of the blade counterweight portion 22 is perpendicular to the center line of the bevel gear, and the center line of the blade counterweight portion 22 is perpendicular to the center line of the bevel gear. Parallel to the vertical pole. The weight of the blade counterweight portion 22 is determined based on the required deflection moment and operating speed of the variable pitch blade. In one embodiment of the present invention, the blade counterweight portion 22 is integrally formed with the bevel gear. If the weight of the blade counterweight portion 22 is to be replaced, the entire variable pitch bevel gear 21 needs to be replaced.

When the main shaft 10 rotates at a certain speed, the blade counterweight 22 will give the blade 23 a deflection moment due to the mismatch between the centripetal force and the structural tension line. It can be seen that the deflection moment is related to the rotation speed. When the rotation speed is constant, the deflection moment is a constant value, thereby causing the blade 23 to rotate in the direction of increasing pitch.

In this embodiment, there are two variable pitch blades 20 , which are symmetrically installed on the main shaft 10 . Of course, the variable pitch blades can also be 3 or 4 blades, which are selected according to the design of the aircraft.

The braking module 30 is composed of an excitation coil 31 , a braking eddy current disk 32 and a braking gear 33 . The braking eddy current disc 32 and the braking gear 33 are rigidly connected together, and the braking eddy current disc 32 and the braking gear 33 are coaxially mounted on the vertical pole 11 through bearings. The axial center line of the braking gear 33 and the axial center line of the variable pitch bevel gear 21 are perpendicular to each other. That is to say, the braking gear 33 and the variable pitch bevel gear 21 are in vertical meshing. In this embodiment, the brake gear 33 and the variable pitch bevel gear 21 are bevel gears meshing with each other. In another embodiment of the present invention, the braking gear 33 is a ring gear track, and the variable pitch bevel gear 21 is a gear meshing with the ring gear track. The present invention adopts eddy current electromagnetic braking, the main shaft 10 is rotating, and the excitation coil 31 is installed on the aircraft body.

When the excitation coil 31 is energized, the magnetic field generated on the coil moves relative to the braking gear 33 to generate an eddy current, which in turn generates a braking torque that prevents the rotation of the braking gear 33. This braking force causes the blade to rotate in the direction of decreasing pitch. What ultimately determines the direction of blade movement is the difference between the deflection moment generated by the counterweight and the moment generated by the eddy current brake disc 33 . The blade counterweight is determined based on the required deflection moment and operating speed of the blade. When the counterweights are consistent and the intensity of the excitation current is changed, the braking torque of the eddy current brake disc 33 can be changed to achieve the purpose of controlling the movement of the blades.

In another embodiment of the present invention, in order to more conveniently replace the weight of the counterweight block, the blade counterweight part 22 is composed of a counterweight mounting bracket and a counterweight block installed on the counterweight mounting bracket. At this time, one end of the counterweight mounting bracket is fixedly connected to the bevel gear, and the other end is provided with a counterweight mounting hole. The counterweight can be determined based on the required deflection moment and operating speed of the variable pitch blade. The counterweight block includes a first counterweight and a second counterweight connected by threads or snaps, and the first counterweight and/or the second counterweight are provided with mounting posts for passing through the counterweight installation holes, and the above-mentioned threads or The buckle can be arranged on the mounting post.

As shown in Figure 5, the present invention provides a multi-copter aircraft, including a main body and at least two electronically controlled variable pitch propellers provided by any embodiment of the present invention; for example, 3, 4, 6 or 8. Generally, 4 electronically controlled variable pitch propellers can be installed on the main body to form a 4-axis aircraft, or 8 electronically controlled variable pitch propellers can be installed on the main body to form an 8-axis aircraft. The present invention does not limit this.

The main body is provided with a control module 60, a wireless communication module 50 and an attitude measurement module 40, and the electronically controlled variable pitch propeller is connected to the control module 60; more specifically, the control module 60 includes multiple motors, a main controller 63 and a main controller 63. The motor driver 61 and the power driver 62 are connected to the device 63. The motor driver 61 is connected to the plurality of motors and is used to control the rotation speed of the motors. Each motor is connected to the main shaft of an electronically controlled variable pitch propeller. The power driver 62 is electrically connected to the excitation coil and is used to control the current intensity of the excitation coil. In addition, it also includes a power module (not shown in the figure), which is used to provide power for the control module 60, the wireless communication module 50, the attitude measurement module 40, etc.

The state of a multi-copter aircraft at a certain moment is described by six physical quantities, including three positions in three-dimensional coordinates and attitude quantities along three axes (called six degrees of freedom). As a detection device, the sensor can sense the measured information, and can transform the detected information into electrical signals or other required forms of information output according to certain rules to meet the needs of information transmission, processing, storage, Display, recording and control requirements. It is the primary link in realizing automatic detection and automatic control. Therefore, the attitude measurement module 40 is a device that provides various flight parameters for the flight control of the multi-axis aircraft, including an acceleration sensor 41 that measures the angular rate of the three-axis body, a gyroscope 42 that measures the three-axis acceleration of the body, a gyroscope 42 that measures the heading of the body, and Geomagnetic sensor 43 for attitude information, etc. During the measurement process, due to the temperature drift of the gyroscope 42 and the influence of the surrounding magnetic field on the geomagnetic sensor 43, the measured attitude information is inaccurate. Therefore, the acceleration sensor 41, the gyroscope 42 and the geomagnetic sensor 43 are combined to obtain accurate bias. Heading angle, roll angle, pitch angle information.

Among them, the acceleration sensor 41 is used to measure the inclination angle of the fuselage relative to the horizontal plane, and uses the earth's gravity to project the gravitational acceleration onto the X, Y, and Z axes to measure the posture of the object. The gyroscope 42 uses the invariance of the direction pointed by the rotation axis of the rotating object when it is not affected by external forces to measure the influence of external forces on the object. The geomagnetic sensor 43 is used to determine the direction, and it uses the geomagnetic field to determine the North Pole. The three-dimensional geomagnetic sensor can provide the heading angle, pitch angle and roll angle of the moving object by giving the geomagnetic force projection on the X-axis, Y-axis and Z-axis, so that the attitude of the object can be determined, which is actually the coordinates of the object. The orientation relationship between the coordinate system and the geographical coordinate system.

The wireless communication module 50 receives the control signal from the remote controller, and then transmits the control signal to the control module 60 . The attitude measurement module 40 uses a three-axis acceleration sensor 41 , a gyroscope 42 , and a geomagnetic sensor 43 to monitor the actual attitude of the aircraft in real time to generate actual attitude data, and transmits the actual attitude data of the aircraft to the control module 60 . After receiving the target attitude data and actual attitude data from the attitude measurement module 40 and the wireless communication module 50, the main controller 63 in the control module 60 completes a series of complex algorithms. The algorithm can be a Kalman filter algorithm, flight control algorithm, etc. PID algorithm, strapdown inertial navigation algorithm, etc., the present invention does not limit these. The main controller 63 calculates the control amount based on the attitude and position information of the aircraft. The control amount includes the rotational speed control amount and the variable pitch control amount. The motor driver 61 converts the rotational speed control amount into the corresponding PWM signal and then drives multiple motors through the drive circuit. The motor works to control the rotation speed of the main shaft in the electronically controlled variable pitch propeller; the power driver 62 converts the variable pitch control amount into a corresponding PWM signal and then controls the current intensity of the excitation coil through the drive circuit, thereby maintaining the stable flight of the multi-axis aircraft. Among them, the PWM pulse control method is to control the on and off of the inverter circuit switching device, so that the output end obtains a series of pulses with equal amplitude, and these pulses are used to replace the sine wave or the required waveform. That is, multiple pulses are generated in half a cycle of the output waveform, so that the equivalent voltage of each pulse is a sinusoidal waveform, and the obtained output is smooth and has few low-order harmonics. By modulating the width of each pulse according to certain rules, the output voltage of the inverter circuit can be changed, and the output frequency can also be changed.

During specific implementation, the motor driver 61 drives each motor to a specified rotation speed according to the instructions of the main controller 63, and feeds back the speed of the motor to the main controller 63 through the speed measurement feedback device. The main controller 63 adjusts the rotation speed control amount accordingly, using Closed-loop control is used to control the motor speed to the expected value, thereby achieving different flight states of the multi-axis aircraft.

In one embodiment of the present invention, the attitude measurement module further includes a pitch encoder. The pitch encoder can transmit the angular position of the blades in real time. The main controller 63 adjusts the pitch control amount according to the angular position of the blades, and The variable pitch control quantity is transmitted to the power driver 62, and the power driver 62 adjusts the current intensity of the excitation coil according to the adjusted variable pitch control quantity, and uses closed-loop control to control the angular position of the blade to a desired value.

As shown in Figure 6, the present invention also provides a control method for a multi-copter aircraft. The multi-copter aircraft is the multi-copter aircraft shown in Figure 5. The application object of the multi-copter aircraft control method provided by the present invention is a fixed-speed multi-rotor aircraft. Axis aircraft. For convenience of description, only the pitch control process of one axis is described. The control method includes steps:

100. Receive control signals and calculate target attitude data accordingly;

101. Obtain the actual attitude data and calculate the control amount combined with the target attitude data;

During specific implementation, a three-axis acceleration sensor, a gyroscope, and a geomagnetic sensor 43 can be used to monitor the actual attitude of the aircraft in real time and obtain actual attitude data. When calculating the control quantity in combination with the target attitude data, the Kalman filter algorithm, flight control PID algorithm, strapdown inertial navigation algorithm, etc. can be used, and the present invention does not limit this.

102. Control the rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller according to the control quantity.

In the present invention, the control quantity includes the rotation speed control quantity and the variable pitch control quantity. The rotation speed control quantity is converted into the corresponding PWM signal through the drive circuit, which can drive multiple motors to work and control the rotation speed of the main shaft in the electronically controlled variable pitch propeller; The variable pitch control amount is converted into the corresponding PWM signal and can be controlled by the drive circuit to control the current intensity of the excitation coil, thereby maintaining the stable flight of the multi-axis aircraft.

Preferably, after step 102, it further includes:

Obtain the attitude information of the variable pitch blades in the electronically controlled variable pitch propeller and adjust the current intensity of the excitation coil accordingly.

In specific implementation, the pitch encoder can be used to transmit the angular position of the blade in real time, adjust the variable pitch control amount according to the angular position of the blade, and adjust the current intensity of the excitation coil according to the adjusted variable pitch control amount, thereby using closed-loop control. Control the angular position of the blade to the expected value.

As shown in Figure 7, the present invention provides a multi-copter aircraft, including a main body and at least two electronically controlled variable pitch propellers provided by any embodiment of the present invention; for example, 3, 4, 6 or 8. Generally, four rotor mounting assemblies can be installed on the main body to form a four-rotor aircraft, or eight rotor mounting assemblies can be installed on the main body to form an eight-rotor aircraft. The present invention does not limit this.

The main body is provided with an attitude detection module 71, a wireless communication module 72, a flight controller 73, and a variable distance controller 74. The variable distance controller 74 is used to control the current intensity of the excitation coil 31. Although not shown in the figure, the flight controller is also connected to multiple motors, and the motors are connected to the pole 11. The rotation speed of the pole 11 can be controlled by the flight controller.

The so-called posture is used to describe the angle and positional relationship between a rigid body's own rigid body coordinate system and the reference coordinate system. Common description methods include Euler angles, that is, the three corners of pitch, roll, and yaw, and there are four more. Elements, directional cosine matrices and other description methods.

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

The gyroscope is an indirect angle measurement instrument. The output is the angular velocity of the carrier movement. The angular velocity needs to be integrated in the time domain to obtain the angle. The acceleration sensor is a device that measures the linear acceleration of the carrier. By measuring the acceleration caused by gravity, the inclination angle of the carrier relative to the horizontal plane can be calculated. The accelerometer is used to continuously correct the gyroscope, and the attitude data of the two can be fused to more accurately calculate the current attitude of the aircraft.

In the specific implementation, the data measured by the gyroscope and accelerometer can be used as input; the output is the inclination angle and angular velocity relative to the earth's coordinate system. The algorithm is detailed as follows:

Low-Pass Fllter: The original angular velocity data of the accelerometer is low-pass filtered, with the purpose of filtering short-term acceleration data fluctuations, filtering out burrs, and smoothing the data. In the implementation, the new measurement bits can be weighted with the previous calculated values, such as :

angle=(0.98)*angle+(0.02)*x-acc，

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

Numeric Integration: The output of the gyroscope is the angular velocity of the object's rotation. Through integration, a smooth angle can be obtained.

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

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

In one embodiment of the present invention, the attitude monitoring module further includes a pitch encoder. The pitch encoder can transmit the angular position of the blades in real time. The flight controller adjusts the current intensity of the excitation coil according to the angular position of the blades, using a closed-loop Control to control the angular position of the blade to the desired value.

The wireless communication module 72 is used to receive control instructions from the remote controller, and then obtain the target posture. After receiving the target attitude data and actual attitude data from the attitude monitoring module 71 and the wireless communication module 72, the flight controller 73 completes a series of complex algorithms and finally outputs the thrust corresponding to each axis. The variable pitch controller 44 is used to obtain the thrust change amount according to the thrust force and change the excitation intensity of the electromagnetic eddy current brake, that is, the current intensity of the excitation coil 31 accordingly. The amount of thrust change refers to whether the thrust corresponding to the target attitude needs to be increased, decreased, or maintained compared to the existing thrust.

As shown in Figure 8, the present invention also provides a control method for a multi-copter aircraft, which is used to control the flight of the multi-copter aircraft as shown in Figure 7. The control method includes the steps:

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

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

111. Obtain the thrust change amount according to the thrust force;

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

112. Change the excitation intensity of the electromagnetic eddy current brake according to the thrust change amount.

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

In this embodiment, it is not necessary to brake the actual propeller torque. Instead, the width of the PWM pulse of the excitation coil is controlled according to the thrust output of the flight controller to increase and decrease, and the thrust is maintained, thereby changing the electromagnetic intensity. braking torque to achieve the purpose of changing the propeller thrust.

Preferably, after step 112, it further includes:

Obtain the attitude information of the blade and adjust the excitation intensity of the electromagnetic eddy current brake accordingly.

In specific implementation, the pitch encoder can be used to transmit the angular position of the blades in real time, the thrust change amount can be adjusted according to the angular position of the blades, and the current intensity of the excitation coil can be adjusted according to the adjusted thrust change amount, thereby using closed-loop control to control the propellers. The angular position of the leaf is as expected.

The preferred embodiments of the present invention have been described above with reference to the accompanying drawings. Those skilled in the art can implement the present invention in various modifications without departing from the scope and essence of the present invention. For example, features shown or described as part of one embodiment can be used on another embodiment, to yield a still further embodiment. The above are only the best and feasible embodiments of the present invention, and do not limit the scope of rights of the present invention. All equivalent changes made by using the description and drawings of the present invention are included in the scope of rights of the present invention.

Industrial applicability

The electronically controlled variable pitch propeller, multi-axis aircraft and control method provided by the present invention achieve variable pitch control by meshing the brake gear of the braking module with the variable pitch bevel gear of the variable pitch blade, thereby meeting the needs of attitude control. , and has good protection, high reliability and simple maintenance. Therefore, it has industrial practicality.

Claims (16)

1. An electronically controlled variable pitch propeller, including a main shaft, variable pitch blades and a braking module, wherein:

   The main shaft includes vertical rods and horizontal rods;

   The braking module includes a coaxially arranged excitation coil, a braking eddy current disc and a braking gear, and the braking eddy current disc and braking gear are installed on the vertical pole;

   The variable-pitch blade includes a variable-pitch bevel gear and a blade mounting clip connected to the variable-pitch bevel gear. The variable-pitch bevel gear is connected to the crossbar and meshes with the brake gear.
2. The electronically controlled variable pitch propeller according to claim 1, wherein the variable pitch bevel gear includes a bevel gear and a blade counterweight connected to the bevel gear.
3. The electronically controlled variable pitch propeller according to claim 2, wherein the center line of the blade counterweight portion is perpendicular to the center line of the bevel gear.
4. The electronically controlled variable pitch propeller according to claim 2 or 3, wherein the blade counterweight portion includes a counterweight mounting bracket and a counterweight block installed on the counterweight mounting bracket.
5. The electronically controlled variable pitch propeller according to claim 4, wherein the counterweight mounting bracket is provided with a counterweight mounting hole.
6. The electronically controlled variable pitch propeller according to claim 4, wherein the counterweight is determined according to the required deflection torque and operating speed of the variable pitch blade.
7. The electronically controlled variable pitch propeller according to claim 1, wherein the brake gear and the variable pitch bevel gear are mutually meshed bevel gears.
8. The electronically controlled variable pitch propeller according to claim 1, wherein the brake gear is a ring gear track, and the variable pitch bevel gear is a gear meshing with the ring gear track.
9. The electronically controlled variable pitch propeller according to claim 1, wherein there are two variable pitch blades, which are symmetrically arranged at both ends of the crossbar.
10. The electronically controlled variable pitch propeller according to claim 1, wherein the variable pitch blade further includes a blade installed on the blade mounting clip.
11. A multi-axis aircraft, including a main body and at least two electronically controlled variable pitch propellers according to any one of claims 1 to 10; the main body is provided with a control module, a wireless communication module and an attitude measurement module, and the electronic The variable pitch propeller is connected to the control module; where:

    The wireless communication module is used to receive control signals and send the control signals to the control module;

    The attitude measurement module is used to monitor the actual attitude of the aircraft in real time to generate actual attitude data, and send the actual attitude data to the control module;

    The control module is used to calculate a control quantity based on the control signal and the actual attitude data, and control the rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller accordingly.
12. The multi-copter aircraft according to claim 11, wherein the control module includes a main controller, a motor driver, and a power driver.
13. A control method for a multi-copter aircraft, the multi-copter aircraft comprising at least two electronically controlled variable pitch propellers as claimed in any one of claims 1 to 10, the control method comprising the steps:

    Receive control signals and calculate target attitude data accordingly;

    Obtain the actual attitude data and calculate the control amount combined with the target attitude data;

    The rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller are controlled according to the control quantity.
14. The control method of a multi-copter aircraft according to claim 13, wherein after controlling the rotation speed of the main shaft and the current intensity of the excitation coil in the electronically controlled variable pitch propeller according to the control quantity, the method further includes:

    Obtain attitude information of the variable pitch blades in the electronically controlled variable pitch propeller and adjust the current intensity of the excitation coil accordingly.
15. A control method for a multi-copter aircraft, the multi-copter aircraft comprising at least two electronically controlled variable pitch propellers according to any one of claims 1 to 10, the control method comprising the steps:

    Output the thrust corresponding to each axis according to the target attitude and actual attitude of the multi-axis aircraft flight;

    Obtain the thrust change amount according to the thrust force;

    The excitation intensity of the excitation coil is changed according to the thrust variation.
16. The control method of a multi-copter aircraft according to claim 15, wherein after changing the excitation intensity of the electromagnetic eddy current brake according to the thrust variation, further comprising:

    Obtain the attitude information of the blade and adjust the excitation intensity of the excitation coil accordingly.