Simple aircraft device based on wing flapping trajectory attitude stabilization algorithm

Abstract

The application discloses a kind of micro air vehicle devices based on wing flapping trajectory attitude stabilization algorithm.The overall scheme is composed of aircraft overall structure, hardware circuit controller module, wing flapping trajectory attitude moment generation method, attitude stabilization control algorithm.Aircraft mechanical flapping structure is composed of two groups of simple gear-spring-wing rod structure, which can realize the efficient and stable flapping of two wings;In the microcontroller, the IMU attitude data is read to implement the cascade PID control algorithm, through the torque-voltage linear mapper and MCU voltage modulator, the PWM waveform is input to the motor drive circuit, so as to control the high-speed forward and reverse rotation of the motor, and finally drive the wing flapping regularly through the flapping mechanism, the wing flapping regularly generates the attitude control moment of the aircraft, and the stable flight of the aircraft is realized.The cascade PID scheme is adopted in the attitude stabilization PID, so that the attitude control of the aircraft will not overshoot, and good stability is achieved.

Description

Technical Field

[0001] The present invention relates to the technical field related to attitude control and motor drive of a micro flapping-wing aircraft, and more particularly to a simple micro aircraft device driven by dual motors based on an algorithm for stabilizing the attitude of a wing flapping trajectory. Background Art

[0002] Micro air vehicles (MAVs), due to their small size, portability, excellent maneuverability, and stealth, are well-suited for missions in confined spaces, such as disaster rescue and military reconnaissance. Due to their demand for specialized applications, MAVs have become a growing research topic among scientists and engineers both domestically and internationally in recent years.

[0003] Aircraft can be broadly categorized based on their flight principles and structures: flapping-wing aircraft, fixed-wing aircraft, and rotary-wing aircraft. Each of these three types of aircraft has distinct advantages. The primary advantage of fixed-wing aircraft is their ability to achieve high-speed flight, but their disadvantage is difficulty achieving hovering and vertical takeoff and landing. Rotary-wing aircraft have the disadvantage of difficulty balancing the torque generated by their own rotors and low aerodynamic efficiency. Micro flapping-wing aircraft, emulating the flight patterns of natural insects and other organisms, align with biomimetic principles, enabling vertical takeoff and hover, offering highly flexible flight modes and excellent aerodynamic efficiency. Therefore, research into the flight structure and flight control of micro flapping-wing aircraft is a promising research area. However, current research on micro flapping-wing aircraft mostly utilizes tail-fin attitude control devices to maintain attitude stability, which undoubtedly increases the complexity of the flapping-wing aircraft system. The development of new, simplified micro flapping-wing aircraft systems and attitude balancing solutions is a current research hotspot. Summary of the Invention

[0004] To address the aforementioned issues, the present invention discloses a micro-aircraft device designed based on a wing flapping trajectory attitude stabilization algorithm. The device utilizes two motors driving two flexible wings to flap through a transmission mechanism. A main control unit (MCU) controls the motor input voltage via a motor drive circuit, thereby controlling the wings to produce regular flapping. According to the quasi-steady-state aerodynamics theoretical model, the interaction between the wing flapping trajectory and air resistance generates a torque that controls the prototype itself. By properly modulating the motor input voltage, the aircraft's attitude stability can be relatively well controlled. This flapping mechanism is simple, energy-efficient, and provides stable and reliable flapping. Furthermore, the attitude control stabilization algorithm is relatively novel, eliminating the need for additional attitude control mechanisms and reducing system complexity.

[0005] A simple aircraft device based on wing flapping trajectory attitude stabilization algorithm, which is composed of aircraft overall structure, hardware circuit controller module, wing flapping trajectory attitude moment generation method and attitude stabilization control algorithm. The process is that the given microcontroller desired attitude data is outputted to the motor control signal to the motor drive circuit through the attitude stabilization control algorithm, the motor rotates to drive the transmission mechanism to move to drive the wing flapping, thereby controlling the regularity trajectory of the aircraft on both sides of the wing to control the stability of the aircraft.

[0006] The whole structure of the aircraft is composed of driving gear one, driving gear two, driven gear one, driven gear two, shaft one, shaft two, spring sleeve one, spring sleeve two, wing rod one, wing rod two, wing one, wing two, upper bottom sleeve one, upper bottom sleeve two, lower bottom sleeve one, lower bottom sleeve two, upper bottom shaft, lower bottom shaft, bottom sleeve support shaft one, bottom sleeve support shaft two, left bearing, right bearing, battery one, battery two, battery support shaft, base, upper support sleeve, lower support sleeve, spring one, spring two, motor one, motor two, motor upper sleeve, motor lower sleeve, motor support shaft, IMU circuit board, microcontroller circuit board, motor drive circuit board and voltage conversion circuit board. The base and the battery support shaft are fixed axially and centrally, the battery one and the battery two are placed symmetrically in front and back on the battery support shaft in series to provide power for the system. The voltage conversion circuit board, the microcontroller circuit board and the motor drive circuit board are arranged on the same circuit board, which respectively provide the required voltage for the circuit, output control signals and drive the motor to rotate. The circuit board is fixed in the hole of the base and the lower support sleeve respectively. The lower bottom sleeve one and the lower bottom sleeve two are connected by the lower bottom shaft, the lower bottom shaft is connected with the lower support sleeve through the cylindrical through hole, the left bearing and the right bearing are fixed with the lower bottom sleeve one and the lower bottom sleeve two, the upper bottom sleeve one and the upper bottom sleeve two through the through hole, the bottom sleeve support shaft one and the bottom sleeve support shaft two are fixed with the left bearing and the right bearing through interference fit, the motor support shaft is fixed with the lower support sleeve and the upper support sleeve through the through hole, the motor one and the motor two are fixed symmetrically on the lower support sleeve and fixed with the motor lower sleeve and the motor upper sleeve through the through hole, the IMU circuit board is fixed between the motor upper sleeve and the motor lower sleeve to measure the attitude data of the prototype in real time. The spring one and the spring two are coaxially matched with the bottom sleeve support shaft one and the bottom sleeve support shaft two respectively, and are hinged with the motor upper sleeve and the spring sleeve one and the spring sleeve two respectively. The spring one and the spring two can store the output electric energy of the motor one and the motor two. The spring sleeve one and the spring sleeve two are engaged with the driven gear one and the driven gear two through the inner and outer teeth, the driven gear one and the driven gear two are engaged with the driving gear one and the driving gear two respectively, the driving gear one and the driving gear two are fixed on the output shaft of the motor one and the output shaft of the motor two respectively, the shaft one and the shaft two are coaxially fixed with the driven gear one and the driven gear two respectively, and are fixed with the wing rod one and the wing rod two through the through hole respectively. The shaft and the driven gear can transmit the flapping torque output by the motor. The upper bottom sleeve one and the upper bottom sleeve two are fixed on the surface of the shaft one and the shaft two respectively, and are connected by the upper bottom shaft, the upper bottom shaft is fixed with the upper support sleeve through the through hole, the wing one and the wing two are axially fixed with the wing rod one and the wing rod two, the bottom sleeve support shaft one and the bottom sleeve support shaft two.

[0007] The wing flapping trajectory attitude control method is composed of wing flapping trajectory amplitude control, wing trajectory flapping bias control and wing flapping trajectory frequency control. The wings on both sides of the aircraft are respectively driven by two flapping mechanisms. When the flapping amplitudes of the wings on both sides are the same and the lift generated by the wings is balanced with the gravity, it is the calibration state of the aircraft. For wing flapping trajectory amplitude control, in the calibration state, the flapping amplitude of the left wing (right wing) is larger than that of the right wing (left wing), the lift generated by the wings on both sides is unbalanced, and the positive (negative) roll control moment of the aircraft is generated. For wing trajectory flapping bias, in the calibration state, the flapping amplitude of the left wing (right wing) generates a same direction bias, the lift generated by the aircraft in front and back is unbalanced, and the positive (negative) pitch control moment of the aircraft is generated. For wing flapping trajectory frequency control, in the calibration state, the left wing flaps up fast (flaps down slow), and the right wing flaps up slow (flaps down fast), the positive rotation moment of the aircraft in the vertical direction of the prototype is generated, the left wing flaps up slow (flaps down fast), and the right wing flaps up fast (flaps down slow), the negative rotation moment of the aircraft in the vertical direction of the prototype is generated, and the above rotation moments are the yaw control moments of the aircraft.

[0008] The hardware circuit controller module is composed of a battery, a microcontroller module, a voltage conversion circuit module, a motor driving circuit module and a motor. The battery is connected with the voltage conversion circuit to provide system voltage input; the voltage conversion circuit is connected with the microcontroller module, the motor driving circuit module and the motor respectively to provide converted input voltage; the microcontroller module is connected with the motor driving circuit module to output PWM wave and enable signal to the motor driving circuit module; the motor driving circuit module is connected with the motor to output MOSFET opening and closing signal and control the rotation of the motor. The battery is composed of 3.7V lithium batteries in series; the microcontroller is STM32F303CCT6 type MCU, the timer PWM output channel and general GPIO output of the MCU are connected with the motor driving circuit module, the motor driving module can be enabled and the positive and negative rotation of the motor can be controlled by controlling the high and low level of the PWM; the voltage conversion circuit module is composed of a microcontroller voltage conversion circuit, a motor voltage conversion circuit and a motor hall voltage conversion circuit, the voltage conversion circuit converts the battery voltage into corresponding power supply voltage for the microcontroller, the motor and the motor hall component through linear voltage stabilizing chips TPS76033, TPS61040 and TPS76050; the motor driving module includes motor driving chip L6235Q and its peripheral circuit, the driving circuit has motor hall signal decoding unit, output current limiting circuit and motor direction control function, by accessing the motor hall component and decoding the signal, the next three-phase full-bridge MOSFET path can be determined, by changing the resistance and voltage of the output current limiting circuit, the motor load current size can be limited; the motor is a miniature brushless DC motor with evenly distributed hall components, the next MOSFET opening and closing selection can be realized through the hall decoding function of the motor driving circuit.

[0009] The attitude stabilization control algorithm is composed of IMU data measurement and attitude fusion algorithm, cascade PID control algorithm, torque-voltage linear mapper algorithm and MCU voltage modulation algorithm. The IMU data measurement and attitude fusion algorithm is connected with the cascade PID control algorithm, detects the actual three-axis angular velocity of the aircraft and the actual three-axis angle after attitude fusion, and provides the input signals of the outer loop and the inner loop of the cascade PID algorithm; the cascade PID algorithm is connected with the torque-voltage linear mapper algorithm, and provides the required control torque input; the torque-voltage linear mapper algorithm is connected with the MCU modulation algorithm, and determines the voltage modulation mode according to the input torque type; the MCU modulation algorithm outputs the motor control signal to drive the wing flapping to finally change the attitude of the aircraft.The IMU data measurement and attitude fusion use the MPU9250 sensor. When the built-in sensing unit updates data, the 16-bit accelerometer data register, the 16-bit gyroscope data register, and the 14-bit magnetometer data register are updated. The gyroscope data register is read into the digital filter for filtering, and the actual three-axis angular velocity of the aircraft is output. The data of the three registers are read and respectively subjected to digital filtering. The DMP digital motion processor is used, the three-axis attitude fusion is realized through the quaternion algorithm, and the actual three-axis angle of the aircraft is output. The cascade PID control algorithm is composed of an outer loop PID controller and an inner loop PID controller. The input of the outer loop PID controller is the difference between the expected three-axis attitude angle and the actual data fusion three-axis attitude angle. The output of the outer loop PID controller is the expected three-axis angular velocity. The input of the inner loop PID controller is the difference between the expected three-axis angular velocity and the actual measured three-axis angular velocity of the attitude. The output of the inner loop PID controller is the actual required three-axis control moment of the aircraft. The moment-voltage linear mapper is composed of a roll moment (Roll)-voltage amplitude difference linear mapper, a pitch moment (Pitch)-voltage bias linear mapper, and a yaw moment (Yaw)-anti-symmetric frequency linear mapper. The corresponding voltage modulation parameters can be obtained according to the size of the attitude control moment. For the roll moment (Roll)-voltage amplitude difference linear mapper, the size of the roll control moment output by the cascade PID is input into the roll moment (Roll)-voltage amplitude difference linear mapper, and the voltage amplitude modulation parameter is output to the voltage modulation input signal. For the pitch moment (Pitch)-voltage bias linear mapper, the size of the pitch control moment output by the cascade PID is input into the pitch moment (Pitch)-voltage bias linear mapper, and the voltage bias modulation parameter is output to the voltage modulation input signal. For the yaw moment (Yaw)-anti-symmetric frequency linear mapper, the size of the yaw control moment output by the cascade PID is input into the yaw moment (Yaw)-anti-symmetric frequency linear mapper, and the voltage frequency modulation parameter is output to the voltage modulation input signal. The MCU voltage modulation algorithm is composed of voltage amplitude modulation, voltage bias modulation, and voltage frequency modulation. According to the output voltage parameters of the above moment-voltage linear mapper, the sizes of the three voltage parameters (amplitude parameter, bias parameter, and frequency parameter) are determined. The DSP unit and the FPU unit are used to accelerate the calculation of the trigonometric function. The scale factor, bias, and frequency coefficient of the trigonometric function are changed, respectively. The duty cycle of the PWM is modified through the timer interrupt program, and the expected control signal of the voltage is output to the motor driving circuit. Finally, the desired trajectory is generated by the motor, the flapping mechanism, and the wing.

[0010] The attitude control mechanism of the micro air vehicle of the application is a mechanism that generates an attitude control moment of the micro air vehicle by controlling the amplitude difference or speed difference of the flapping of the wings on both sides of the micro air vehicle, and controls the size and direction of the control moment generated by the wings according to the size of the amplitude difference or speed difference.

[0011] The micro air vehicle device drives the flapping of the two wings through a simple and efficient flapping mechanism, outputs a PWM waveform to a motor driving circuit through an STM32 single-chip microcomputer, controls the high-speed forward and reverse rotation of a micro brushless direct current motor through the driving circuit, prevents the problems of stator overheating and the meshing disengagement of the driving gear and the driven gear caused by the four-quadrant rotation of the motor through the energy recovery and damping effect of the torsion spring, and realizes the stable and efficient reciprocating flapping of the wings.

[0012] Advantages of the application:

[0013] 1. In terms of wing flapping mechanical structure, a mechanical structure of motor-gear-spring-coupling-wing is used, which is small in size, light in weight, clear and simple in structure, has no tail wing and additional complex structure for controlling the attitude of the micro flapping-wing aircraft.

[0014] 2. From the attitude detection of the micro flapping-wing aircraft to the expected attitude, a cascade PID algorithm is used for control, which can improve the overall stability of the system and prevent overshoot in flapping adjustment.

[0015] 3. The MPU9250 attitude sensor has a DMP digital motion processor, which can perform additional data processing, and has an external interrupt that wakes up the main control MCU while updating data, effectively reducing the burden of the MCU.

[0016] 4. The FPU and DSP in the MCU can accelerate floating point and trigonometric function operations, greatly improving the operation speed of the interrupt program and making it possible to calculate the pulse comparison value in real time.

[0017] 5、Based on the quasi-steady aerodynamic model, the wing flapping amplitude, bias and frequency can be mapped to the attitude control torque. From the perspective of voltage input control, the timer output PWM method is adopted, which not only meets the amplitude, bias and frequency modulation of the input voltage, but also has strong real-time and rapidity. The control method is simple and stable, and has good control accuracy. BRIEF DESCRIPTION OF DRAWINGS

[0018] Figure 1 is a structure of a micro air vehicle device;

[0019] Figure 2 is a structure of a micro air vehicle device

[0020] Figure 3 is a wing flapping trajectory state and generated torque schematic diagram;

[0021] Figure 4 is a hardware circuit controller module schematic diagram;

[0022] Figure 5 is an attitude stability control algorithm schematic diagram. DETAILED DESCRIPTION

[0023] The present application will be further illustrated below in conjunction with the drawings and specific embodiments, and it should be understood that the following specific embodiments are only used to illustrate the present application and not to limit the scope of the present application. It should be noted that the words "front", "back", "left", "right", "up" and "down" used in the following description refer to the directions in the drawings, and the words "in" and "out" refer to the directions towards or away from the geometric center of a particular component.

[0024] Figure 1 and Figure 2It is respectively the main view and the plan view of the micro air vehicle device structure of the application, a double motor driven micro air vehicle device based on wing flap trajectory attitude stability algorithm. The left and right whole adopts symmetry structure design, the whole structure of the aircraft is composed of the whole structure of the aircraft, the driving gear one, the driving gear two, the driven gear one, the driven gear two, the shaft one, the shaft two, the spring sleeve one, the spring sleeve two, the wing rod one, the wing rod two, the wing one, the wing two, the upper bottom sleeve one, the upper bottom sleeve two, the lower bottom sleeve one, the lower bottom sleeve two, the upper bottom shaft, the lower bottom shaft, the bottom sleeve support shaft one, the bottom sleeve support shaft two, the left bearing, the right bearing, the battery one, the battery two, the battery support shaft, the base, the upper support sleeve, the lower support sleeve, the spring one, the spring two, the motor one, the motor two, the motor upper sleeve, the motor lower sleeve, the motor support shaft, the IMU circuit board, the microcontroller circuit board, the motor drive circuit board and the voltage conversion circuit board; the base and the battery support shaft keep axial center fixed, the battery one and the battery two are placed in front of and behind the battery support shaft in series, and provide energy for the system; the voltage conversion circuit board, the microcontroller circuit board and the motor drive circuit board are arranged on the same circuit board, and the functions are respectively to provide the required voltage of the circuit, output control signal and drive motor rotation, and the circuit board is fixed in the hole of the base and the lower support sleeve respectively; the lower bottom sleeve one and the lower bottom sleeve two are connected through the lower bottom shaft, the lower bottom shaft is connected with the lower support sleeve through the cylindrical through hole, the left bearing and the right bearing are fixed with the lower bottom sleeve one and the lower bottom sleeve two, the upper bottom sleeve one and the upper bottom sleeve two through the through hole respectively, the bottom sleeve support shaft one and the bottom sleeve support shaft two are fixed with the left bearing and the right bearing through the interference fit of the inner diameter, the motor support shaft is fixed with the lower support sleeve and the upper support sleeve through the through hole, the motor one and the motor two are fixed on the lower support sleeve symmetrically, and are fixed with the motor lower sleeve and the motor upper sleeve through the through hole, the IMU circuit board is fixed between the motor upper sleeve and the motor lower sleeve, and the attitude data of the prototype is measured in real time; the spring one and the spring two are coaxially matched with the bottom sleeve support shaft one and the bottom sleeve support shaft two respectively, and are hinged with the motor upper sleeve and the spring sleeve one and the spring sleeve two respectively, and the spring one and the spring two can store the output electric energy of the motor one and the motor two; the spring sleeve one and the spring sleeve two are engaged with the driven gear one and the driven gear two through the inner and outer teeth, the driven gear one and the driven gear two are engaged with the driving gear one and the driving gear two respectively, the driving gear one and the driving gear two are fixed on the motor one output shaft and the motor two output shaft respectively, the shaft one and the shaft two are coaxially fixed with the driven gear one and the driven gear two respectively, and are fixed with the wing rod one and the wing rod two through the through hole respectively, the shaft and the driven gear can transmit the flapping torque output by the motor; the upper bottom sleeve one and the upper bottom sleeve two are fixed on the surface of the shaft one and the shaft two respectively, and are connected through the upper bottom shaft, the upper bottom shaft is fixed with the upper support sleeve through the through hole, the wing one and the wing two are axially fixed with the wing rod one and the wing rod two, the bottom sleeve support shaft one and the bottom sleeve support shaft two.

[0025] According to the aircraft device structure in the above figure, the brushless DC motor angle output sine waveform can be controlled to reasonably control the wing flapping amplitude value. According to the system structure, the following equations can be obtained:

[0026] (1)

[0028] T e =K a I (2)

[0029]

[0030] Since the micro brushless DC motor has small inductance, its circuit equation is shown in equation (1), wherein u is the circuit input voltage, I is the circuit current, R is the motor inter-phase resistance, Ka is the motor torque and speed constant (under certain conditions, the torque constant is equal to the speed constant), is the motor angle, is the wing flapping angle, wherein n is the gear transmission ratio, equation (2) is the relationship between the current and the motor output electromagnetic torque, equation (3) is the motor torque equation, wherein Jm is the motor shaft moment of inertia, T l is the system air damping and friction damping load, Ks is the torsional spring stiffness coefficient. T l The calculation of the value is related to the wing area and the wing flapping speed, according to the quasi-steady aerodynamics model, generally wherein σ is a constant related to the wing size, then through the above said, the transfer function equation is simplified as follows:

[0031]

[0032] The above equation (4) is the general transfer equation of the flapping angle, and the system is nonlinear according to the equation. Through local linearization, a linearized system at a certain fixed frequency can be realized, and the linearized system is a traditional second-order system. Through the modulation of the input voltage u of the sine signal, the system output is a sine signal, and through the amplitude, bias and frequency modulation of the input signal, the expected flapping trajectory output can be realized.

[0033] Figure 3The figure shows the three-axis attitude control torque diagram generated by the wing flapping trajectory state. The wing flapping trajectory attitude control method consists of three parts: wing flapping trajectory amplitude control, wing flapping trajectory bias control, and wing flapping trajectory frequency control. The wings on both sides of the aircraft are driven by two flapping mechanisms. When the flapping amplitudes of the wings on both sides are the same and the lift generated by them is balanced with the gravity, it is the calibration state of the aircraft. For wing flapping trajectory amplitude control, in the calibration state, the flapping amplitude of the left wing (right wing) is larger than that of the right wing (left wing), and the lift generated by the wings on both sides is unbalanced, generating positive (negative) roll control torque of the aircraft. For wing flapping trajectory bias, in the calibration state, the left wing (right wing) flapping amplitude generates a same direction bias, and the lift generated by the aircraft in the front and back is unbalanced, generating positive (negative) pitch control torque of the aircraft. For wing flapping trajectory frequency control, in the calibration state, the left wing flaps up fast (flaps down slow), and the right wing flaps up slow (flaps down fast), generating positive rotation torque of the aircraft in the vertical direction of the prototype, and the left wing flaps up slow (flaps down fast), and the right wing flaps up fast (flaps down slow), generating negative rotation torque of the aircraft in the vertical direction of the prototype. The above rotation torque is the yaw control torque of the aircraft. For Figure 3 (a) The figure shows that the flapping amplitudes of the wings on both sides are equal, at which the lift of the aircraft is balanced, and it is the calibration state. For Figure 3 (b) The figure shows that the flapping amplitude of the left wing is significantly larger than that of the right wing, and the flapping trajectories of the left and right wings are symmetrical. According to the principle of quasi-steady aerodynamics, the flapping trajectories of the left and right wings are symmetrical, the lift generated by the left and right wings is balanced, and the flapping amplitude of the left wing is larger than that of the right wing, which will generate a larger range of lift on the side with larger amplitude. This will cause the prototype to generate a roll torque (roll) from the top of the left wing to the top of the right wing as shown in the figure. The size of the roll torque is linearly related to the difference in flapping amplitude of the left and right wings. For Figure 3 (c) The figure shows that the flapping amplitudes of the left and right wings are basically the same, but they have an additional amplitude bias. From the perspective of wing lift mechanism and quasi-steady aerodynamics model, the size of the wing lift is positively related to the size of the amplitude. The left and right wings are symmetrical, and the lift generated by the two sides will be balanced. The up and down flapping is asymmetrical, which will generate an up and down unbalanced torque, such as the pitch torque (pitch) shown in the figure. For Figure 3 (d) The right side of the figure shows that the flapping amplitudes are symmetrical up and down, and the flapping amplitudes of the left and right wings are symmetrical, so the roll torque and pitch torque are balanced. However, during the back and forth flapping of the wings, the period is constant, the up flapping speed is faster, and the down flapping speed is slower, which is called period time invariance and half period time complementary phenomenon. The size of the air resistance experienced by the wings is determined by the flapping speed. The faster the flapping speed, the faster the air resistance experienced by the wings. Figure 3(d)As shown, the left wing beats faster, the right wing beats slower, and the air resistance on the left and right wings is not equal, so that the micro flapping-wing aircraft generates a moment that causes the prototype to deviate. The specific deviation moment and direction are determined by the up-and-down beating speed of the two wings.

[0034] Figure 4 The hardware circuit controller module of the micro aircraft device is shown in the figure. The hardware circuit controller module is composed of a battery, a microcontroller module, a voltage conversion circuit module, a motor driving circuit module and a motor. The battery is connected with the voltage conversion circuit to provide system voltage input; the voltage conversion circuit is connected with the microcontroller module, the motor driving circuit module and the motor respectively to provide converted input voltage; the microcontroller module is connected with the motor driving circuit module to output PWM wave and enable signal to the motor driving circuit module; the motor driving circuit module is connected with the motor to output MOSFET opening and closing signal to control the rotation of the motor. The battery is composed of 3.7V lithium batteries in series; the microcontroller is an STM32F303CCT6 type MCU, the timer PWM output channel and the general GPIO output of the MCU are connected with the motor driving circuit module, the motor driving module can be enabled and the forward and reverse motion of the motor can be controlled by controlling the high and low level of the PWM; the voltage conversion circuit module is composed of a microcontroller voltage conversion circuit, a motor voltage conversion circuit and a motor hall voltage conversion circuit, the voltage conversion circuit converts the battery voltage into corresponding supply voltage for the microcontroller, the motor and the motor hall component through linear voltage stabilizing chips TPS76033, TPS61040 and TPS76050; the motor driving module includes a motor driving chip L6235Q and its peripheral circuit, the driving circuit has a motor hall signal decoding unit, an output current limiting circuit and a motor direction control function, by connecting the motor hall component and decoding the signal, the next three-phase full-bridge MOSFET path can be determined, by changing the resistance and voltage of the output current limiting circuit, the motor load current size can be limited; the motor is a micro brushless DC motor with evenly distributed hall components, the next MOSFET opening and closing selection can be realized through the hall decoding function of the motor driving circuit.

[0035] Figure 5The attitude stabilizing control algorithm of the micro air vehicle device is shown in the figure. The attitude stabilizing control algorithm is composed of an IMU data measurement and attitude fusion algorithm, a cascade PID control algorithm, a torque-voltage linear mapper algorithm and an MCU voltage modulation algorithm. The IMU data measurement and attitude fusion algorithm is connected with the cascade PID control algorithm, the actual three-axis angular velocity of the aircraft and the actual three-axis angle after attitude fusion are detected, and the input signals of the outer loop and the inner loop of the cascade PID algorithm are provided; the cascade PID algorithm is connected with the torque-voltage linear mapper algorithm, the required control torque input is provided; the torque-voltage linear mapper algorithm is connected with the MCU modulation algorithm, the voltage modulation mode is determined according to the input torque type; the MCU modulation algorithm outputs the motor control signal to drive the wing flapping to finally change the attitude of the aircraft.The IMU data measurement and attitude fusion use the MPU9250 sensor. When the built-in sensing unit updates data, the 16-bit accelerometer data register, the 16-bit gyroscope data register and the 14-bit magnetometer data register are updated. The gyroscope data register is read into the digital filter for filtering, and the actual three-axis angular velocity of the aircraft is output. The data of the three registers are read and respectively filtered, the DMP digital motion processor is used, the three-axis attitude fusion is realized through the quaternion algorithm, and the actual three-axis angle of the aircraft is output. The cascade PID control algorithm is composed of an outer loop PID controller and an inner loop PID controller. The input of the outer loop PID controller is the difference between the expected three-axis attitude angle and the actual data fusion three-axis attitude angle. The output of the outer loop PID controller is the expected three-axis angular velocity. The input of the inner loop PID controller is the difference between the expected three-axis angular velocity and the actual measured three-axis angular velocity of the attitude. The output of the inner loop PID controller is the actual required three-axis control moment of the aircraft. The moment-voltage linear mapper is composed of a roll moment (Roll)-voltage amplitude difference linear mapper, a pitch moment (Pitch)-voltage bias linear mapper and a yaw moment (Yaw)-anti-symmetry frequency linear mapper, which can obtain the corresponding voltage modulation parameters according to the size of the attitude control moment. For the roll moment (Roll)-voltage amplitude difference linear mapper, the size of the roll control moment output by the cascade PID is input into the roll moment (Roll)-voltage amplitude difference linear mapper, and the voltage amplitude modulation parameter is output to the voltage modulation input signal. For the pitch moment (Pitch)-voltage bias linear mapper, the size of the pitch control moment output by the cascade PID is input into the pitch moment (Pitch)-voltage bias linear mapper, and the voltage bias modulation parameter is output to the voltage modulation input signal. For the yaw moment (Yaw)-anti-symmetry frequency linear mapper, the size of the yaw control moment output by the cascade PID is input into the yaw moment (Yaw)-anti-symmetry frequency linear mapper, and the voltage frequency modulation parameter is output to the voltage modulation input signal. The MCU voltage modulation algorithm is composed of voltage amplitude modulation, voltage bias modulation and voltage frequency modulation. According to the output voltage parameters of the above moment-voltage linear mapper, the sizes of the three voltage parameters (amplitude parameter, bias parameter and frequency parameter) are determined. The DSP unit and the FPU unit are used to accelerate the calculation of the trigonometric function, the scale factor, the bias and the frequency coefficient of the trigonometric function are changed respectively, the duty cycle of the PWM is modified through the timer interrupt program, and the voltage expected control signal is output to the motor driving circuit. Finally, the desired trajectory is generated by the motor, the flapping mechanism and the wing. The PWM modulation adopts the real-time pulse comparison value calculation mode, and the fast trigonometric function API in the DSP library is used, and the formula is as follows.

[0036]

[0037] This formula is the interrupt program calculation formula of the modulated SPWM waveform, value is the pulse comparison value, V1 is the timer PWM overflow value, index is the SPWM sampling number value, and pluse is the position mark of the current sampling point. The sine wave period is determined by the timer overflow value and the sampling number. Among them, for amplitude parameter modulation, according to the area impulse equivalent principle, when the duty cycle reaches 1, the amplitude is the power supply voltage input value. By adjusting the coefficient value of ε, when ε is 1, the duty cycle reaches the maximum value of 1, and the motor input signal amplitude is the power supply value. The value of ε is always between 0 and 1. For bias parameter modulation, it is determined by the δ parameter. The range of δ value is to (Generally, avoid the case where the duty cycle is 1, at this time the bias voltage exists modulation space), in the bias calibration case, the value is always 0, and the maximum adjustable range is half of the power supply amplitude. Exceeding this range will cause over-regulation, that is, due to the fixed nature of the power supply voltage, SPWM waveform cannot be generated. For frequency parameter modulation, it is realized by controlling the index size. When the index sampling point is half of the single cycle sampling point, the speed of the wing up and down is equal. When the index sampling point is less than half, the speed is slower in the half cycle, and the speed is faster in the other half cycle. The principle is that in the case of constant PWM cycle time, the number of 0-π sampling points and 0-2π sampling points is controlled, the total sampling point number is unchanged, and the sum of the two sampling point numbers is equal to the total sampling point number. Thus, the speed of the wing up and down is not equal, and the yaw attitude control moment is generated. By calling the above three modulation methods in the timer interrupt service function, the required SPWM waveform voltage output can be accurately controlled. After filtering by motor inductance, an approximate sine wave voltage is output, and the desired voltage input is achieved.

[0038] The technical means disclosed in the present application scheme is not limited to the technical means disclosed in the above embodiments, but also includes technical solutions composed of any combination of the above technical features.

Claims

1. A simple aircraft control method based on wing flapping trajectory attitude stabilization algorithm, characterized in that: The application relates to a wing flapping trajectory attitude moment generation method and an attitude stabilization control algorithm, and the process is as follows: through given microcontroller desired attitude data, a motor control signal is output to a motor driving circuit through the attitude stabilization control algorithm, the motor rotates to drive the transmission mechanism to move so as to drive the wing flapping, thereby controlling the regularity trajectory of the wings on both sides of the aircraft to control the stability of the aircraft; wherein the wing flapping trajectory attitude moment generation method is composed of wing flapping trajectory amplitude control, wing trajectory flapping bias control and wing flapping trajectory frequency control; the wings on both sides of the aircraft are driven by two flapping mechanisms respectively, when the flapping amplitudes of the wings on both sides are the same and the generated lift force is balanced with the gravity, the aircraft is in a calibration state; for the wing flapping trajectory amplitude control, under the calibration state, the flapping amplitude of the left wing (right wing) is larger than that of the right wing (left wing), the lift forces generated by the wings on the left and right sides are unbalanced, and a positive (negative) roll control moment of the aircraft is generated; for the wing trajectory flapping bias, under the calibration state, the flapping amplitudes of the left wing (right wing) generate a same direction bias, the lift forces generated by the aircraft in the front and back directions are unbalanced, and a positive (negative) pitch control moment of the aircraft is generated; for the wing flapping trajectory frequency control, under the calibration state, the left wing flaps up fast (flaps down slow), the right wing flaps up slow (flaps down fast), a positive rotation moment of the aircraft in the vertical direction of the prototype is generated, the left wing flaps up slow (flaps down fast), the right wing flaps up fast (flaps down slow), a negative rotation moment of the aircraft in the vertical direction of the prototype is generated, and the rotation moment is a yaw control moment of the aircraft; the attitude stabilization control algorithm is composed of an IMU data measurement and attitude fusion algorithm, a cascade PID control algorithm, a moment-voltage linear mapper algorithm and a MCU voltage modulation algorithm; the IMU data measurement and attitude fusion algorithm is connected with the cascade PID control algorithm, actual three-axis angular velocities and actual three-axis angles after attitude fusion of the aircraft are detected, and input signals of the outer loop and the inner loop of the cascade PID algorithm are provided; the cascade PID algorithm is connected with the moment-voltage linear mapper algorithm, and required control moment input is provided; the moment-voltage linear mapper algorithm is connected with the MCU voltage modulation algorithm, voltage modulation modes are determined according to the input moment types; the MCU voltage modulation algorithm outputs a motor control signal to drive the wing flapping and finally change the attitude of the aircraft; wherein the IMU data measurement and attitude fusion use an MPU9250 sensor, when the built-in sensing unit updates data, 16-bit accelerometer data registers, 16-bit gyroscope data registers and 14-bit magnetometer data registers are updated, gyroscope data registers are read to a digital filter for filtering, actual three-axis angular velocities of the aircraft are finally output, data of the three registers are read and respectively subjected to digital filtering, a DMP digital motion processor is used, three-axis attitude fusion is realized through a quaternion algorithm, and actual three-axis angles of the aircraft are output.The cascade PID control algorithm is composed of an outer loop PID controller and an inner loop PID controller, the input of the outer loop PID controller is the difference between the expected three-axis attitude angle and the actual three-axis attitude angle, the output of the outer loop PID controller is the expected three-axis angular velocity, the input of the inner loop PID controller is the difference between the expected three-axis angular velocity and the actual three-axis angular velocity, and the output of the inner loop PID controller is the actual required three-axis control moment of the aircraft; the moment-voltage linear mapper is composed of a roll moment (Roll)-voltage amplitude difference linear mapper, a pitch moment (Pitch)-voltage bias linear mapper and a yaw moment (Yaw)-anti-symmetry frequency linear mapper, which can obtain corresponding voltage modulation parameters according to the size of the attitude control moment; for the roll moment (Roll)-voltage amplitude difference linear mapper, the cascade PID outputs the size of the roll control moment, which is input into the voltage modulation input signal after the roll moment (Roll)-voltage amplitude difference linear mapper; for the pitch moment (Pitch)-voltage bias linear mapper, the cascade PID outputs the size of the pitch control moment, which is input into the voltage modulation input signal after the pitch moment (Pitch)-voltage bias linear mapper; for the yaw moment (Yaw)-anti-symmetry frequency linear mapper, the cascade PID outputs the size of the yaw control moment, which is input into the voltage modulation input signal after the yaw moment (Yaw)-anti-symmetry frequency linear mapper; the MCU voltage modulation algorithm is composed of voltage amplitude modulation, voltage bias modulation and voltage frequency modulation, according to the output voltage parameters of the above moment-voltage linear mapper, the sizes of three voltage parameters: amplitude parameter, bias parameter and frequency parameter are determined, the calculation of the trigonometric function is accelerated by using the DSP unit and the FPU unit, the proportional coefficient, the bias and the frequency coefficient of the trigonometric function are changed respectively, the duty cycle of the PWM is modified through the timer interrupt program, the voltage expected control signal is output to the motor driving circuit, and finally the expected trajectory is generated by the motor, the flapping mechanism and the wing.

2. A simple aircraft control method based on wing flapping trajectory attitude stabilization algorithm according to claim 1, characterized in that: The MCU voltage modulation algorithm is composed of amplitude modulation, bias modulation and frequency modulation algorithm; the MCU timer and the function of fast trigonometric function calculation of the DSP are used to calculate the next PWM duty cycle in real time, and the amplitude, bias and frequency of the sine wave can be adjusted through the duty cycle size of the PWM and the number of sampling points, so as to achieve the expected voltage modulation output.

3. The simple aircraft control method based on wing flapping trajectory attitude stabilization algorithm of claim 1, wherein: The attitude stability control adopts a cascade PID control algorithm; the cascade PID control algorithm is composed of an outer loop PID algorithm and an inner loop PID algorithm; the input of the outer loop PID link is the error between the expected three-axis Euler angle of the aircraft and the actual aircraft attitude detection Euler angle, the output of the outer loop PID is the expected angular velocity value of the inner loop, the input of the inner loop PID link is the error between the expected output value of the outer loop and the aircraft attitude detection angular velocity, and the output of the inner loop PID link is the expected control torque link of the aircraft; through the cascade PID control algorithm, the attitude control stability of the micro air vehicle system can be realized, and the attitude overshoot in the system regulation process can be effectively prevented.

4. A simple aircraft device using a simple aircraft control method based on a wing flapping trajectory attitude stabilization algorithm as claimed in claim 1, comprising an aircraft overall structure and a hardware circuit controller module: characterized by: The whole structure of the aircraft is composed of a driving gear one, a driving gear two, a driven gear one, a driven gear two, a shaft one, a shaft two, a spring sleeve one, a spring sleeve two, a wing rod one, a wing rod two, a wing one, a wing two, an upper bottom sleeve one, an upper bottom sleeve two, a lower bottom sleeve one, a lower bottom sleeve two, an upper bottom shaft, a lower bottom shaft, a bottom sleeve support shaft one, a bottom sleeve support shaft two, a left bearing, a right bearing, a battery one, a battery two, a battery support shaft, a base, an upper support sleeve, a lower support sleeve, a spring one, a spring two, a motor one, a motor two, a motor upper sleeve, a motor lower sleeve, a motor support shaft, an IMU circuit board, a microcontroller circuit board, a motor driving circuit board and a voltage conversion circuit board; the base and the battery support shaft are fixed axially and centrally, the battery one and the battery two are placed symmetrically in front and back on the battery support shaft in series to provide energy for the system; the voltage conversion circuit board, the microcontroller circuit board and the motor driving circuit board are arranged on the same circuit board, which respectively provide the required voltage for the circuit, output control signals and drive the motor to rotate, and are fixed in the holes of the base and the lower support sleeve respectively; the lower bottom sleeve one and the lower bottom sleeve two are connected through the lower bottom shaft, the lower bottom shaft is connected with the lower support sleeve through a cylindrical through hole, the left bearing and the right bearing are fixed with the lower bottom sleeve one and the lower bottom sleeve two, the upper bottom sleeve one and the upper bottom sleeve two through through holes, the bottom sleeve support shaft one and the bottom sleeve support shaft two are fixed with the left bearing and the right bearing through interference fit, the motor support shaft is fixed with the lower support sleeve and the upper support sleeve through a through hole, the motor one and the motor two are fixed symmetrically on the lower support sleeve and fixed with the motor lower sleeve and the motor upper sleeve through a through hole, the IMU circuit board is fixed between the motor upper sleeve and the motor lower sleeve to measure the attitude data of the prototype in real time; the spring one and the spring two are coaxially matched with the bottom sleeve support shaft one and the bottom sleeve support shaft two respectively and are hinged with the motor upper sleeve and the spring sleeve one and the spring sleeve two respectively, the spring one and the spring two can store the output electric energy of the motor one and the motor two; the spring sleeve one and the spring sleeve two are engaged with the driven gear one and the driven gear two through inner and outer teeth, the driven gear one and the driven gear two are engaged with the driving gear one and the driving gear two respectively, the driving gear one and the driving gear two are fixed on the output shaft of the motor one and the output shaft of the motor two respectively, the shaft one and the shaft two are coaxially fixed with the driven gear one and the driven gear two respectively and are fixed with the wing rod one and the wing rod two through a through hole, the shaft and the driven gear can transmit the flapping torque output by the motor; the upper bottom sleeve one and the upper bottom sleeve two are fixed on the surface of the shaft one and the shaft two respectively and are connected through the upper bottom shaft, the upper bottom shaft is fixed with the upper support sleeve through a through hole, the wing one and the wing two are axially fixed with the wing rod one and the wing rod two, the bottom sleeve support shaft one and the bottom sleeve support shaft two.

5. A simple aircraft device according to claim 4, characterized in that: The hardware circuit controller module is composed of a battery, a microcontroller module, a voltage conversion circuit module, a motor driving circuit module and a motor; the battery is connected with the voltage conversion circuit to provide system voltage input; the voltage conversion circuit is connected with the microcontroller module, the motor driving circuit module and the motor respectively to provide converted input voltage; the microcontroller module is connected with the motor driving circuit module to output PWM wave and an enable signal to the motor driving circuit module; the motor driving circuit module is connected with the motor to output MOSFET on-off signal and control the motor rotation; the microcontroller is an STM32F303CCT6 type MCU, and the timer PWM output channel and the general GPIO output of the MCU are connected with the motor driving circuit module; the voltage conversion circuit module is composed of a microcontroller voltage conversion circuit, a motor voltage conversion circuit and a motor Hall voltage conversion circuit, the voltage conversion circuit converts the battery voltage into corresponding supply voltage for the microcontroller, the motor and the motor Hall component through a linear voltage stabilizing chip; the motor driving module is composed of a motor driving chip and its peripheral circuit, the driving circuit has a motor Hall signal decoding unit, an output current limiting circuit and a motor direction control function, by accessing the motor Hall component and decoding the signal, the next three-phase full-bridge MOSFET path can be determined, by changing the resistance and voltage of the output current limiting circuit, the motor load current size can be limited; the motor is a miniature brushless DC motor, has evenly distributed Hall components, and the next MOSFET on-off selection can be realized through the Hall decoding function of the motor driving circuit.

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