WO2022160618A1

WO2022160618A1 - Passenger drone flight posture control method and system, and passenger drone

Info

Publication number: WO2022160618A1
Application number: PCT/CN2021/107415
Authority: WO
Inventors: 周双久; 李�杰; 黄璐璐; 张书存; 邹露星
Original assignee: 广东汇天航空航天科技有限公司
Priority date: 2021-01-29
Filing date: 2021-07-20
Publication date: 2022-08-04
Also published as: CN112965505A

Abstract

A passenger drone flight posture control method and system, and a passenger drone. The passenger drone flight posture control method comprises: processing an inputted expected posture value of a drone via a tracking differentiator to produce a differential parameter value of the expected posture value (S101); processing an acquired current posture value of the drone via a state observation instrument to produce a posture feedback value of the current posture value (S102); performing an operation on the basis of the differential parameter value of the expected posture value and of the posture feedback value of the current posture value to produce a set error value (S103); processing the set error value via a nonlinear controller, and outputting a first control value for flight posture control to control the flight posture of the drone (S104). Improved autonomous anti-interference capability is provided and the flight stability of the passenger drone is increased.

Description

Manned aircraft flight attitude control method, system and manned aircraft

This application claims the priority of the Chinese patent application filed on January 29, 2021, with the application number of 202110134374.2 and the invention titled "Flight Attitude Control Method, System and Manned Aircraft for a Manned Aircraft", the entire contents of which are by reference Incorporated in this application.

technical field

The present application relates to the technical field of aircraft and control, and in particular, to a flight attitude control method and system for a manned aircraft, and a manned aircraft.

Background technique

At present, stability control is very important in aircraft technology, and various control algorithms are used in the stability control of the flight attitude of unmanned aerial vehicles or manned aerial vehicles. Commonly used controllers include PID (Proportion Integration Differentiation) controller, backstepping controller and so on. Among them, PID controller is widely used because of its simplicity and high efficiency.

In the traditional PID control method, the parameters involved mainly include three parameters: P, I, and D, where P is a proportional parameter, I is an integral parameter, and D is a differential parameter. The calculated value of the three parameters is calculated to obtain an output control value, which controls and adjusts the attitude error of reducing the flight attitude.

However, in the related art, during the flight of the aircraft, there are many disturbances, such as airflow disturbance, uneven center of gravity of the aircraft, inconsistent dynamic parameters, and other external or internal disturbances. Although the parameters of PID control are simple, its robustness is poor, there are limitations and bottlenecks in different degrees, and the anti-interference ability is poor. The integral link set due to the difference will make the control response of the system lag; the differential signal can only be approximated, which is easily polluted by noise and has poor control effect.

Therefore, the flight attitude control method of the related art needs to be further improved.

SUMMARY OF THE INVENTION

In order to overcome the problems existing in the related art, the present application provides a flight attitude control method and system for a manned aircraft, and a manned aircraft, which have better anti-jamming capability and can improve the flight stability of the manned aircraft.

A first aspect of the present application provides a method for controlling the flight attitude of a manned aircraft, including:

The input expected attitude value of the aircraft is processed by the tracking differentiator to obtain the differential parameter value of the expected attitude value;

The obtained current attitude value of the aircraft is processed by the state observer to obtain the attitude feedback value of the current attitude value;

According to the differential parameter value of the desired attitude value and the attitude feedback value of the current attitude value, a set error value is obtained;

The set error value is processed by a nonlinear controller, and a first control value for controlling the flight attitude is output to control the flight attitude of the aircraft.

In an implementation manner, after processing the set error value by the nonlinear controller and outputting the first control value for controlling the flight attitude, the method further includes:

The operation is performed according to the first control value and the attitude feedback value of the current attitude value, and a second control value for controlling the flying attitude is output to control the flying attitude of the aircraft.

In one embodiment, the inputted expected attitude value of the aircraft is processed by the tracking differentiator to obtain the differential parameter value of the expected attitude value, including:

Using the tracking differentiator and using the fastest control synthesis function, the input expected attitude value of the aircraft is processed according to the first preset algorithm, and the expected attitude angle and the differential value of the expected attitude angle are output.

In one embodiment, the state observer processes the acquired current attitude value of the aircraft to obtain the attitude feedback value of the current attitude value, including:

Using the state observer, the acquired current attitude value of the aircraft is processed according to the second preset algorithm, and the observed attitude value, the differential value of the observed attitude, and the current disturbance feedback value of the aircraft are obtained.

In an implementation manner, the calculation based on the differential parameter value of the desired attitude value and the attitude feedback value of the current attitude value to obtain a set error value includes: combining the desired attitude angle with the observed The attitude value is subtracted to obtain the attitude error value; the differential value of the desired attitude angle is subtracted from the observed attitude differential value to obtain the attitude differential error value;

The non-linear controller processes the set error value, and outputs a first control value for controlling the flight attitude to control the flight attitude of the aircraft, including:

Using a nonlinear controller, the attitude error value and the attitude differential error value are processed according to a third preset algorithm, and a first control value for controlling the flight attitude is output to control the flight attitude of the aircraft.

In an embodiment, the operation according to the first control value and the attitude feedback value of the current attitude value, and outputting the second control value for controlling the flight attitude to control the flight attitude of the aircraft, includes:

The first control value is subtracted from the disturbance feedback value of the current aircraft, and a second control value for controlling the flight attitude is output to control the flight attitude of the aircraft.

A second aspect of the present application provides a flight attitude control system, including:

a tracking differentiator, configured to process the input expected attitude value of the aircraft to obtain a differential parameter value of the expected attitude value;

a state observer, configured to process the acquired current attitude value of the aircraft to obtain an attitude feedback value of the current attitude value;

a nonlinear controller for processing the set error value and outputting a first control value for controlling the flight attitude to control the flight attitude of the aircraft, wherein the set error value is based on the differential parameter value of the desired attitude value Obtained by operation with the attitude feedback value of the current attitude value.

In one embodiment, the system further includes:

an adjustment module, configured to perform an operation according to the first control value output by the nonlinear controller and the attitude feedback value of the current attitude value obtained by the state observer, and output a second control value for controlling the flight attitude to control the aircraft flight attitude.

In one embodiment, the tracking differentiator uses the fastest control comprehensive function to process the input desired attitude value of the aircraft according to a first preset algorithm, and outputs the desired attitude angle and the differential value of the desired attitude angle;

The state observer processes the acquired current attitude value of the aircraft according to the second preset algorithm, and obtains the observation attitude value, the observation attitude differential value, and the disturbance feedback value of the current aircraft;

The nonlinear controller processes the attitude error value and the attitude differential error value according to a third preset algorithm, and outputs a first control value for controlling the flight attitude to control the flight attitude of the aircraft, wherein the desired attitude angle is compared with the flight attitude. The observed attitude values are subtracted to obtain an attitude error value, and the differential value of the desired attitude angle is subtracted from the observed attitude differential value to obtain an attitude differential error value.

In one embodiment, the adjustment module subtracts the first control value output by the nonlinear controller and the current disturbance feedback value of the aircraft obtained by the state observer, and outputs a second control value for controlling the flight attitude value to control the flight attitude of the aircraft.

A third aspect of the present application provides a manned aircraft, including the above-mentioned flight attitude control system.

A fourth aspect of the present application provides an electronic device, comprising:

processor; and

A memory having executable codes stored thereon, and when the executable codes are executed by the processor, causes the processor to execute the above-described method.

The technical solution provided by this application can include the following beneficial effects:

The flight attitude control method provided by the embodiment of the present application utilizes a tracking differentiator, a state observer and a nonlinear controller to perform comprehensive operations, and processes the input expected attitude value of the aircraft through the tracking differentiator to obtain the expected attitude value. The obtained differential parameter value of the aircraft is processed by the state observer to obtain the attitude feedback value of the current attitude value; according to the differential parameter value of the desired attitude value and the attitude feedback of the current attitude value The set error value is obtained by calculating the value; the set error value is processed by the nonlinear controller, and the first control value for controlling the flight attitude is output. In the related art, if there is a large deviation between the expected attitude value and the current actual attitude value, it is easy to generate overshoot (referring to the output exceeding its final steady state value) and overshoot (referring to exceeding the given value) using PID control. However, the solution of this application sets up a transition process through the processing of the tracking differentiator to make the input of the desired attitude value smoother. After a transition process, the desired attitude value can be realized under the premise of basically no overshoot The fast tracking and processing of the set error value combined with the state observer and the nonlinear controller can relatively stably output the control value for controlling the flight attitude, so that the control system has a better anti-interference ability and improves the manned aircraft. flight stability.

Further, in the method provided by the embodiment of the present application, the state observer can observe the unknown external disturbance in real time, feedback the disturbance component, that is, the disturbance feedback value, and finally subtract the disturbance component from the output part of the controller to eliminate the external disturbance. , to further enhance the stability of the flight attitude.

It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not limiting of the present application.

Description of drawings

The above and other objects, features and advantages of the present application will become more apparent from the more detailed description of the exemplary embodiments of the present application in conjunction with the accompanying drawings, wherein the same reference numerals generally represent the exemplary embodiments of the present application. same parts.

1 is a schematic flowchart of a method for controlling the flight attitude of a manned aircraft according to an embodiment of the present application;

2 is a schematic flowchart of another method for controlling the flight attitude of a manned aircraft according to an embodiment of the present application;

3 is a schematic flowchart of another method for controlling the flight attitude of a manned aircraft according to an embodiment of the present application;

4 is a schematic diagram of an application of the application control system shown in the embodiment of the present application to perform flight attitude control;

5 is a schematic structural diagram of a flight attitude control system shown in an embodiment of the present application;

FIG. 6 is a schematic structural diagram of an electronic device shown in an embodiment of the present application.

Detailed ways

Preferred embodiments of the present application will be described in more detail below with reference to the accompanying drawings. While preferred embodiments of the present application are shown in the drawings, it should be understood that the present application may be embodied in various forms and should not be limited by the embodiments set forth herein. Rather, these embodiments are provided so that this application will be thorough and complete, and will fully convey the scope of this application to those skilled in the art.

The terminology used in this application is for the purpose of describing particular embodiments only and is not intended to limit the application. As used in this application and the appended claims, the singular forms "a," "the," and "the" are intended to include the plural forms as well, unless the context clearly dictates otherwise. It will also be understood that the term "and/or" as used herein refers to and includes any and all possible combinations of one or more of the associated listed items.

It should be understood that although the terms "first", "second", "third", etc. may be used in this application to describe various information, such information should not be limited by these terms. These terms are only used to distinguish the same type of information from each other. For example, the first information may also be referred to as the second information, and similarly, the second information may also be referred to as the first information without departing from the scope of the present application. Thus, a feature defined as "first" or "second" may expressly or implicitly include one or more of that feature. In the description of the present application, "plurality" means two or more, unless otherwise expressly and specifically defined.

At present, the related art uses a PID-controlled flight attitude control method, and the control effect needs to be improved. In view of the above problems, the embodiments of the present application provide a method for controlling the flight attitude of a manned aircraft, which has better anti-jamming capability and can improve the flight stability of the manned aircraft. The technical solutions of the embodiments of the present application will be described in detail below with reference to the accompanying drawings.

FIG. 1 is a schematic flowchart of a method for controlling the flight attitude of a manned aircraft according to an embodiment of the present application.

Referring to Figure 1, the method includes:

In step S101, the input expected attitude value of the aircraft is processed by the tracking differentiator to obtain the differential parameter value of the expected attitude value.

Wherein, a tracking differentiator can be used and the fastest control comprehensive function can be used to process the input expected attitude value of the aircraft according to the first preset algorithm, and output the expected attitude angle and the differential value of the expected attitude angle.

In step S102, the acquired current attitude value of the aircraft is processed by the state observer to obtain the attitude feedback value of the current attitude value.

Wherein, a state observer may be used to process the acquired current attitude value of the aircraft according to the second preset algorithm to obtain the observed attitude value, the differential value of the observed attitude, and the disturbance feedback value of the current aircraft.

In step S103, an operation is performed according to the differential parameter value of the desired attitude value and the attitude feedback value of the current attitude value to obtain a set error value.

Among them, the expected attitude angle and the observed attitude value can be subtracted to obtain the attitude error value; the differential value of the expected attitude angle and the observed attitude differential value can be subtracted to obtain the attitude differential error value.

In step S104, the set error value is processed by the nonlinear controller, and the first control value for controlling the flight attitude is output to control the flight attitude of the aircraft.

Wherein, a nonlinear controller may be used to process the attitude error value and the attitude differential error value according to a third preset algorithm, and output a first control value for controlling the flight attitude to control the flight attitude of the aircraft.

It can be seen from this embodiment that in the related art, if there is a large deviation between the desired attitude value and the current actual attitude value, overshoot (meaning that the output exceeds its final steady-state value) and overshoot are easily generated by using PID control (referring to exceeding a given value), and the solution of this application sets up a transition process by tracking the processing of the differentiator to make the input of the desired attitude value smoother. Under the premise of achieving fast tracking of the desired attitude value, and combining the state observer and the nonlinear controller to process the set error value, the control value of the flight attitude control can be relatively stably output to control the flight attitude of the aircraft, so that the control The system has good anti-jamming capability and improves the flight stability of manned aircraft.

FIG. 2 is a schematic flowchart of another method for controlling the flight attitude of a manned aircraft according to an embodiment of the present application. Compared with Fig. 1, the process of Fig. 2 mainly subtracts the disturbance component from the output part of the nonlinear controller to further improve the stability.

Referring to Figure 2, the method includes:

In step S201, the input expected attitude value of the aircraft is processed by the tracking differentiator to obtain the differential parameter value of the expected attitude value.

In step S202, the acquired current attitude value of the aircraft is processed by the state observer to obtain the attitude feedback value of the current attitude value.

In step S203, an operation is performed according to the differential parameter value of the desired attitude value and the attitude feedback value of the current attitude value to obtain a set error value.

In step S204, the set error value is processed by the nonlinear controller, and the first control value for controlling the flight attitude is output.

Wherein, a nonlinear controller can be used to process the attitude error value and the attitude differential error value according to a third preset algorithm, and output the first control value for controlling the flight attitude.

In step S205, an operation is performed according to the first control value and the attitude feedback value of the current attitude value, and a second control value for controlling the flying attitude is output to control the flying attitude of the aircraft.

The first control value may be subtracted from the current disturbance feedback value of the aircraft, and the second control value for controlling the flight attitude may be output to control the flight attitude of the aircraft.

It can be seen from this embodiment that, in addition to making the input of the desired attitude value smoother through the processing of the tracking differentiator, the solution of the present application can realize fast tracking of the desired attitude value under the premise of basically no overshoot. The controller can observe the unknown external disturbance in real time, and feed back the disturbance component, that is, the disturbance feedback value. Finally, the disturbance component is subtracted from the output part of the nonlinear controller to eliminate the external disturbance and further enhance the stability of the flight attitude.

FIG. 3 is a schematic flowchart of another method for controlling the flight attitude of a manned aircraft according to an embodiment of the present application; FIG. 4 is a schematic diagram of an application of the application control system shown in the embodiment of the present application for flight attitude control. FIG. 3 introduces the technical solution of the present application in more detail with respect to FIG. 1 and FIG. 2 .

Referring to Figure 4, the control system includes: a tracking differentiator, a state observer and a nonlinear controller.

The function of the tracking differentiator is to reasonably extract the differential signal from the noise-contaminated signal. The state observer is a kind of dynamic system that obtains the estimated value of the state variable according to the measured value of the external variables (input variable and output variable) of the system, also known as the state reconstructor. A nonlinear controller refers to a system whose state and output variables cannot be described by a linear relationship under the influence of external conditions.

Assuming that the desired attitude value of the aircraft input to the tracking differentiator is V, the tracking differentiator outputs the desired attitude angle X ₁ and the differential value X ₂ of the desired attitude angle after processing; the state observer obtains the current attitude value of the aircraft y(k), observe the current attitude value y(k), and output the attitude feedback value of the current attitude value, including the observation attitude value Z ₁ , the observation attitude differential value Z ₂ , and the disturbance feedback value Z ₃ of the current aircraft. The observation attitude value Z ₁ and the observation attitude differential value Z ₂ can be used as state feedback values; the X ₁ and X ₂ output by the tracking differentiator are subtracted from the two state feedback values of the state observer to obtain two The error values are the attitude error value and the attitude differential error value, and are input into the nonlinear controller; the nonlinear controller processes according to the attitude error value and the attitude differential error value, and outputs the first control value u ₀ ; From the disturbance feedback value Z ₃ of the state observer, the final second control value u is obtained as a control quantity and output to the aircraft motor for flight control.

Referring to Figure 3, the method includes:

In step S301, the expected attitude value of the aircraft is input to the tracking differentiator, and the expected attitude value is processed by the tracking differentiator to obtain the desired attitude angle and the differential value of the desired attitude angle.

The tracking differentiator uses the fastest control comprehensive function to process the input expected attitude value of the aircraft according to the first preset algorithm, and outputs the expected attitude angle and the differential value of the expected attitude angle.

During the flight of the aircraft, the attitude refers to the current flight attitude of the aircraft. Attitude is actually a relationship between the aircraft coordinate system and the geographic coordinate system. In the aircraft coordinate system, the X axis is generally the direction of the aircraft wing, the Y axis is the direction of the nose, and the Z axis is perpendicular to the aircraft. This coordinate system changes with the attitude of the aircraft. The flight attitude can be represented by the attitude angle, which refers to the angle between the aircraft coordinate system and the geographic coordinate system, which can be represented by three angles: roll, pitch, and yaw. Among them, the roll angle: the angle between the plane of symmetry of the aircraft and the vertical plane passing through the longitudinal axis of the aircraft body, the right roll is positive; the pitch angle: the angle between the body axis and the ground plane (horizontal plane), the head of the aircraft is positive ;Yaw angle: the angle between the projection of the body axis on the horizontal plane and the ground axis, with the right deviation of the aircraft as positive.

The input of the tracking differentiator is the expected attitude value of the aircraft, and the expected attitude value may be derived from the pilot's manipulation input but is not limited thereto. When there is no pilot or no control, the expected value of the roll angle and pitch angle in the expected attitude value is 0, and the expected value of the yaw angle is the current heading angle (that is, the heading angle to keep the aircraft level and lock the current heading).

Tracking differentiators can be processed using the fastest control synthesis function fhan function. Assuming that the expected attitude value of the aircraft input to the tracking differentiator is V(k), the tracking differentiator outputs the expected attitude angle X ₁ and the differential value X ₂ of the expected attitude angle after processing according to the first preset algorithm.

The relevant formula of the first preset algorithm is as follows:

fh=fhan(X ₁ (k)-V(k),X ₂ (k),r,h)

X ₁ (k+1)=X ₁ (k)+h*X ₂ (k)

X ₂ (k+1)=X ₂ (k)+h*fh

In the above formula, X ₁ (k) represents the desired attitude angle at time k, X ₂ (k) represents the differential value of the desired attitude angle at time k, X ₁ (k+1) represents the desired attitude angle at time k+1, X ₂ (k+1) represents the differential value of the desired attitude angle at time k+1, V(k) represents the expected attitude value input at time k, r represents the speed factor that determines the tracking speed, h represents the time step, and fh represents Second derivative of the input data. The fhan function is called the fastest control synthesis function, and reference may be made to the calculation method of the related art for this function, which is not limited in this application.

In this application, the main purpose of setting the tracking differentiator for processing is to arrange a transition process, that is, to arrange the transition process of the closed-loop system, so that the desired angle value is no longer directly input into the control system by using the PID control algorithm, because if the The PID control algorithm is directly input into the control system, which is prone to overshoot (meaning that the output exceeds its final steady-state value) and overshoot (meaning exceeding a given value).

In step S302, the current attitude value of the aircraft is acquired by the state observer for processing, and the observed attitude value, the differential value of the observed attitude and the current disturbance feedback value of the aircraft are obtained.

The state observer processes the acquired current attitude value of the aircraft according to the second preset algorithm, and obtains the observation attitude value, the observation attitude differential value, and the disturbance feedback value of the current aircraft.

The state observer observes the current input attitude value y(k) of the aircraft, and estimates the system state value and disturbance component. Due to the existence of various disturbances such as external disturbances, there may be disturbance noise in the input current attitude value. After the operation is performed according to the second preset algorithm, three components Z ₁ , Z ₂ and Z ₃ can be obtained, that is, three attitudes Feedback value, where Z ₁ represents the observation attitude value observed by the observer (compared to the input attitude value, the observation attitude value filters out a part of the disturbance noise), Z ₂ represents the observed observation attitude differential value, and Z ₃ represents the external The disturbance component of is the disturbance feedback value of the current aircraft.

The relevant formula of the second preset algorithm is as follows:

ε=Z ₁ (k)-y(k)

Z ₁ (k+1)=Z ₁ (k)+h*[Z ₂ (k)-β ₀₁ *ε]

In the above formula, ε represents the error of the attitude value at time k, y(k) represents the current attitude value of the aircraft input at the current time k, and Z ₁ (k) represents the observed attitude value observed by the observer at time k (compared to the input Attitude value, this observation attitude value filters out part of the disturbance noise), Z ₂ (k) represents the observed attitude differential value observed at time k, Z ₃ (k) represents the external disturbance component at time k, that is, the disturbance feedback value of the current aircraft, Z ₁ (k+1) represents the observation attitude value observed by the observer at time k+1, Z ₂ (k+1) represents the differential value of the observation attitude observed at time k+1, and Z ₃ (k+1) represents k The external disturbance component at time +1 is the disturbance feedback value of the current aircraft, h represents the time step, β ₀₁ represents the angle gain coefficient, β ₀₂ represents the angular velocity gain coefficient, β ₀₃ represents the angular acceleration gain coefficient, and δ represents the filtering of the fal filter function. factor, b ₀ is a compensation factor, u represents the last control amount (control value), and fal is a nonlinear filter function. Among them, the filtering factor and the compensation factor can be empirical values according to actual needs.

In step S303, the attitude error value is calculated according to the expected attitude angle and the observed attitude value; the attitude differential error value is calculated according to the differential value of the expected attitude angle and the observed attitude differential value.

The desired attitude angle X ₁ output by the tracking differentiator and the differential value X ₂ of the desired attitude angle are subtracted from the two state feedback values Z ₁ and Z ₂ of the state observer to obtain two error values e ₁ and e ₂ , into the nonlinear controller.

The formula for the nonlinear controller to calculate the attitude error value is as follows:

e ₁ =X ₁ -Z ₁

e ₂ =X ₂ -Z ₂

In step S304, the attitude error value and the attitude differential error value are processed by the nonlinear controller, and the first control value for controlling the flight attitude is output.

The nonlinear controller processes the attitude error value e ₁ and the attitude differential error value e ₂ according to a third preset algorithm, and outputs a first control value for controlling the flight attitude.

After the attitude error value is input to the nonlinear controller, the nonlinear controller performs related operations, and the output control value is the first control value u ₀ . The relevant formula of the third preset algorithm is as follows:

u ₀ =β ₁ *fal(e ₁ ,a ₁ ,δ)+β ₂ *fal(e ₂ ,a ₂ ,δ)

In the above formula, β ₁ and β ₂ represent the angle feedback gain and angular velocity feedback gain respectively, a ₁ and a ₂ are the tracking factors of the two fal functions, δ represents the filter factor of the fal filter function, and fal is a nonlinear filter function , e ₁ is the attitude error value, and e ₂ is the attitude differential error value. The filtering algorithm can usually use Kalman filtering or complementary filtering, etc. Oscillation during filtering can be prevented by increasing the filter factor.

In step S305, an operation is performed according to the first control value and the current disturbance feedback value of the aircraft, and a second control value for controlling the flight attitude is output to the aircraft motor.

Wherein, the first control value can be subtracted from the disturbance feedback value of the current aircraft, and the second control value that controls the flight attitude can be output and transmitted to the aircraft motor to control the flight attitude of the aircraft, for example, to the motor controller in the aircraft motor, so that The motor controller may control the rotational speed of the motor according to the second control value, thereby controlling the flight attitude of the aircraft.

The output first control value u ₀ is then subtracted from the disturbance feedback value Z ₃ of the state observer to obtain the final control value, that is, the second control value u. The first control value u ₀ can be divided by the compensation factor b ₀ after subtracting the disturbance feedback value Z ₃ of the state observer to obtain the final control value, that is, the second control value u.

The relevant formula is as follows:

In the above formula, b ₀ represents the compensation factor, and the compensation factor can be an empirical value according to actual needs.

Finally, the control quantity, that is, the second control value u, is finally output to the motor controller of the aircraft, and the motor controller controls the rotation speed of the motor according to the second control value u, thereby controlling the flight attitude of the aircraft, and the manned aircraft can be realized. smooth control of the flight attitude.

To sum up, the solution of the present application has the following beneficial effects:

1) In the related art, if there is a large deviation between the expected attitude value and the current actual attitude value, it is easy to generate overshoot (referring to the output exceeding its final steady state value) and overshoot (referring to exceeding the given value) using PID control. ) problem, and the solution of this application sets up a transition process by tracking the processing of the differentiator to make the input of the desired attitude value smoother. The fast tracking of the attitude value, combined with the state observer and the nonlinear controller to process the set error value, can relatively stably output the control value of the flight attitude control, so that the control system has a better anti-interference ability and improves the load. Human aircraft flight stability.

2) In the solution of this application, an unknown external disturbance can be observed in real time through the state observer, and the disturbance component, that is, the disturbance feedback value, can be fed back. Finally, the disturbance component is subtracted from the output part of the controller to eliminate the external disturbance and further enhance the flight. Postural stability.

It can be seen that the scheme of this application is different from the control methods of manned aircraft commonly used in the market. The control scheme proposed in this application can effectively remove the interference of external factors on the aircraft, and can also guarantee the Aircraft such as manned aircraft fly safely and stably.

The above describes in detail the flight attitude control method of the manned aircraft of the present application. Correspondingly, the present application also provides a flight attitude control system, a manned aircraft and related equipment.

FIG. 5 is a schematic structural diagram of a flight attitude control system shown in an embodiment of the present application.

Referring to FIG. 5 , the flight attitude control system 50 provided by the present application includes: a tracking differentiator 51 , a state observer 52 , and a nonlinear controller 53 .

The tracking differentiator 51 is used for processing the input expected attitude value of the aircraft to obtain the differential parameter value of the expected attitude value.

The state observer 52 is configured to process the acquired current attitude value of the aircraft to obtain an attitude feedback value of the current attitude value.

The nonlinear controller 53 is used to process the set error value and output the first control value for controlling the flight attitude to control the flight attitude of the aircraft, wherein the set error value is based on the differential parameter of the desired attitude value obtained by the tracking differentiator 51 The value is obtained by operation with the attitude feedback value of the current attitude value obtained by the state observer 52 .

In one embodiment, the flight attitude control system 50 may further include: an adjustment module 54 .

The adjustment module 54 is used for calculating according to the first control value output by the nonlinear controller 53 and the attitude feedback value of the current attitude value obtained by the state observer 52, and outputs the second control value for controlling the flight attitude to control the flight attitude of the aircraft .

In one embodiment, the tracking differentiator 51 uses the fastest control synthesis function to process the input expected attitude value of the aircraft according to the first preset algorithm, and outputs the expected attitude angle and the differential value of the expected attitude angle.

The relevant formula of the first preset algorithm is as follows:

fh=fhan(X ₁ (k)-V(k),X ₂ (k),r,h)

X ₁ (k+1)=X ₁ (k)+h*X ₂ (k)

X ₂ (k+1)=X ₂ (k)+h*fh

The state observer 52 processes the acquired current attitude value of the aircraft according to the second preset algorithm, and obtains the observed attitude value, the differential value of the observed attitude, and the disturbance feedback value of the current aircraft.

The relevant formula of the second preset algorithm is as follows:

ε=Z ₁ (k)-y(k)

Z ₁ (k+1)=Z ₁ (k)+h*[Z ₂ (k)-β ₀₁ *ε]

In the above formula, ε represents the error of the attitude value at time k, y(k) represents the current attitude value of the aircraft input at the current time k, and Z ₁ (k) represents the observed attitude value observed by the observer at time k (compared to the input Attitude value, this observation attitude value filters out part of the disturbance noise), Z ₂ (k) represents the observed attitude differential value observed at time k, Z ₃ (k) represents the external disturbance component at time k, that is, the disturbance feedback value of the current aircraft, Z ₁ (k+1) represents the observation attitude value observed by the observer at time k+1, Z ₂ (k+1) represents the differential value of the observation attitude observed at time k+1, and Z ₃ (k+1) represents k The external disturbance component at time +1 is the disturbance feedback value of the current aircraft, h represents the time step, β ₀₁ represents the angle gain coefficient, β ₀₂ represents the angular velocity gain coefficient, β ₀₃ represents the angular acceleration gain coefficient, and δ represents the filtering of the fal filter function. factor, b ₀ is a compensation factor, u represents the last control amount (control value), and fal is a nonlinear filter function.

The nonlinear controller 53 processes the attitude error value and the attitude differential error value according to a third preset algorithm, and outputs a first control value for controlling the flight attitude to control the flight attitude of the aircraft, wherein the desired attitude angle is subtracted from the observed attitude value. , get the attitude error value, and subtract the differential value of the desired attitude angle from the observed attitude differential value to obtain the attitude differential error value.

e ₁ =X ₁ -Z ₁

e ₂ =X ₂ -Z ₂

The relevant formula of the third preset algorithm is as follows:

u ₀ =β ₁ *fal(e ₁ ,a ₁ ,δ)+β ₂ *fal(e ₂ ,a ₂ ,δ)

In the above formula, β ₁ and β ₂ represent the angle feedback gain and angular velocity feedback gain respectively, a ₁ and a ₂ are the tracking factors of the two fal functions, δ represents the filter factor of the fal filter function, and fal is a nonlinear filter function . The filtering algorithm can usually use Kalman filtering or complementary filtering, etc. Oscillation during filtering can be prevented by increasing the filter factor.

In one embodiment, the adjustment module 54 subtracts the first control value output by the nonlinear controller and the current disturbance feedback value of the aircraft obtained by the state observer, and outputs the second control value for controlling the flight attitude to control the aircraft's flight attitude.

The relevant formula is as follows:

The present application also provides a manned aircraft, including the above-mentioned flight attitude control system 50 . The structure of the flight attitude control system 50 can be referred to the description in FIG. 5 , which will not be repeated here.

Regarding the apparatus in the above-mentioned embodiment, the specific manner in which each module performs the operation has been described in detail in the embodiment of the method, and will not be described in detail here.

FIG. 6 is a schematic structural diagram of an electronic device shown in an embodiment of the present application. The electronic device may be, for example, a control system device or the like.

Referring to FIG. 6 , an electronic device 600 includes a memory 610 and a processor 620 .

The processor 620 may be a central processing unit (Central Processing Unit, CPU), or other general-purpose processors, digital signal processors (Digital Signal Processors, DSPs), application specific integrated circuits (Application Specific Integrated Circuits, ASICs), field-available processors. Field-Programmable Gate Array (FPGA) or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, etc. A general purpose processor may be a microprocessor or the processor may be any conventional processor or the like.

Memory 610 may include various types of storage units, such as system memory, read only memory (ROM), and persistent storage. The ROM may store static data or instructions required by the processor 620 or other modules of the computer. Persistent storage devices may be readable and writable storage devices. Permanent storage may be a non-volatile storage device that does not lose stored instructions and data even if the computer is powered off. In some embodiments, persistent storage devices employ mass storage devices (eg, magnetic or optical disks, flash memory) as persistent storage devices. In other embodiments, persistent storage may be a removable storage device (eg, a floppy disk, an optical drive). System memory can be a readable and writable storage device or a volatile readable and writable storage device, such as dynamic random access memory. System memory can store some or all of the instructions and data that the processor needs at runtime. Additionally, memory 610 may include any combination of computer-readable storage media, including various types of semiconductor memory chips (DRAM, SRAM, SDRAM, flash memory, programmable read-only memory), and magnetic and/or optical disks may also be employed. In some embodiments, memory 610 may include a removable storage device that is readable and/or writable, such as a compact disc (CD), a read-only digital versatile disc (eg, DVD-ROM, dual-layer DVD-ROM), Read-only Blu-ray Discs, Ultra-Density Discs, Flash Cards (eg SD Cards, Min SD Cards, Micro-SD Cards, etc.), Magnetic Floppy Disks, etc. Computer-readable storage media do not contain carrier waves and transient electronic signals transmitted over wireless or wireline.

Executable code is stored on the memory 610, and when the executable code is processed by the processor 620, the processor 620 can be caused to perform some or all of the above-mentioned methods.

The solution of the present application has been described in detail above with reference to the accompanying drawings. In the above-mentioned embodiments, the description of each embodiment has its own emphasis. For parts that are not described in detail in a certain embodiment, reference may be made to the relevant descriptions of other embodiments. Those skilled in the art should also know that the actions and modules involved in the description are not necessarily required by the present application. In addition, it can be understood that the steps in the method of the embodiment of the present application may be sequentially adjusted, combined and deleted according to actual needs, and the modules in the device of the embodiment of the present application may be combined, divided and deleted according to actual needs.

Furthermore, the method according to the present application can also be implemented as a computer program or computer program product comprising computer program code instructions for performing some or all of the steps in the above method of the present application.

Alternatively, the present application can also be implemented as a non-transitory machine-readable storage medium (or computer-readable storage medium, or machine-readable storage medium) on which executable codes (or computer programs, or computer instruction codes are stored) ), when the executable code (or computer program, or computer instruction code) is executed by the processor of the electronic device (or electronic device, server, etc.), the processor is caused to execute each step of the above-mentioned method according to the present application part or all of it.

Those skilled in the art will also appreciate that the various exemplary logical blocks, modules, circuits, and algorithm steps described in connection with the applications herein may be implemented as electronic hardware, computer software, or combinations of both.

The flowchart and block diagrams in the figures illustrate the architecture, functionality, and operation of possible implementations of systems and methods according to various embodiments of the present application. In this regard, each block in the flowchart or block diagrams may represent a module, segment, or portion of code that contains one or more functions for implementing the specified logical function(s) executable instructions. It should also be noted that, in some alternative implementations, the functions noted in the block may occur out of the order noted in the figures. For example, two blocks in succession may, in fact, be executed substantially concurrently, or the blocks may sometimes be executed in the reverse order, depending upon the functionality involved. It is also noted that each block of the block diagrams and/or flowchart illustrations, and combinations of blocks in the block diagrams and/or flowchart illustrations, can be implemented in dedicated hardware-based systems that perform the specified functions or operations , or can be implemented in a combination of dedicated hardware and computer instructions.

Various embodiments of the present application have been described above, and the foregoing descriptions are exemplary, not exhaustive, and not limiting of the disclosed embodiments. Numerous modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the described embodiments. The terminology used herein was chosen to best explain the principles of the various embodiments, the practical application or improvement over the technology in the marketplace, or to enable others of ordinary skill in the art to understand the various embodiments disclosed herein.

Claims

A method for controlling the flight attitude of a manned aircraft, comprising:

The input expected attitude value of the aircraft is processed by the tracking differentiator to obtain the differential parameter value of the expected attitude value;

The obtained current attitude value of the aircraft is processed by the state observer to obtain the attitude feedback value of the current attitude value;

According to the differential parameter value of the desired attitude value and the attitude feedback value of the current attitude value, a set error value is obtained;

The set error value is processed by a nonlinear controller, and a first control value for controlling the flight attitude is output to control the flight attitude of the aircraft.
The method according to claim 1, characterized in that, after the processing of the set error value by the nonlinear controller and outputting the first control value for controlling the flight attitude, the method further comprises:

The operation is performed according to the first control value and the attitude feedback value of the current attitude value, and a second control value for controlling the flying attitude is output to control the flying attitude of the aircraft.
The method according to claim 1 or 2, characterized in that, processing the input expected attitude value of the aircraft through a tracking differentiator to obtain a differential parameter value of the expected attitude value, comprising:

Using the tracking differentiator and using the fastest control synthesis function, the input expected attitude value of the aircraft is processed according to the first preset algorithm, and the expected attitude angle and the differential value of the expected attitude angle are output.
The method according to claim 3, wherein the process of processing the acquired current attitude value of the aircraft through the state observer to obtain the attitude feedback value of the current attitude value comprises:

Using the state observer, the acquired current attitude value of the aircraft is processed according to the second preset algorithm, and the observed attitude value, the differential value of the observed attitude, and the current disturbance feedback value of the aircraft are obtained.
The method according to claim 4, wherein:

The calculating according to the differential parameter value of the desired attitude value and the attitude feedback value of the current attitude value to obtain a set error value includes: subtracting the desired attitude angle and the observed attitude value to obtain the attitude Error value; subtract the differential value of the desired attitude angle from the differential value of the observed attitude to obtain the attitude differential error value;

The non-linear controller processes the set error value, and outputs a first control value for controlling the flight attitude to control the flight attitude of the aircraft, including:

Using a nonlinear controller, the attitude error value and the attitude differential error value are processed according to a third preset algorithm, and a first control value for controlling the flight attitude is output to control the flight attitude of the aircraft.
The method according to claim 4, wherein the calculation is performed according to the attitude feedback value of the first control value and the current attitude value, and a second control value for controlling the flight attitude is output to control the flight attitude of the aircraft. flight attitude, including:

The first control value is subtracted from the disturbance feedback value of the current aircraft, and a second control value for controlling the flight attitude is output to control the flight attitude of the aircraft.
A flight attitude control system, comprising:

a tracking differentiator, configured to process the input expected attitude value of the aircraft to obtain a differential parameter value of the expected attitude value;

a state observer, configured to process the acquired current attitude value of the aircraft to obtain an attitude feedback value of the current attitude value;

a non-linear controller for processing the set error value and outputting a first control value for controlling the flight attitude to control the flight attitude of the aircraft, wherein the set error value is based on the desired attitude obtained by the tracking differentiator The differential parameter value of the value is obtained by operation with the attitude feedback value of the current attitude value obtained by the state observer.
The system of claim 7, wherein the system further comprises:

an adjustment module, configured to perform an operation according to the first control value output by the nonlinear controller and the attitude feedback value of the current attitude value obtained by the state observer, and output a second control value for controlling the flight attitude to control the aircraft flight attitude.
The system of claim 8, wherein:

The tracking differentiator uses the fastest control comprehensive function to process the input desired attitude value of the aircraft according to the first preset algorithm, and outputs the desired attitude angle and the differential value of the desired attitude angle;

The state observer processes the acquired current attitude value of the aircraft according to the second preset algorithm, and obtains the observation attitude value, the observation attitude differential value, and the disturbance feedback value of the current aircraft;

The nonlinear controller processes the attitude error value and the attitude differential error value according to a third preset algorithm, and outputs a first control value for controlling the flight attitude to control the flight attitude of the aircraft, wherein the desired attitude angle is compared with the flight attitude. The observed attitude values are subtracted to obtain an attitude error value, and the differential value of the desired attitude angle is subtracted from the observed attitude differential value to obtain an attitude differential error value.
The system of claim 9, wherein:

The adjustment module subtracts the first control value output by the nonlinear controller from the disturbance feedback value of the current aircraft obtained by the state observer, and outputs a second control value for controlling the flight attitude to control the aircraft's flight attitude.
A manned aircraft, characterized by comprising the flight attitude control system according to any one of claims 7-10.
An electronic device, comprising:

processor; and

A memory having executable code stored thereon which, when executed by the processor, causes the processor to perform the method of any one of claims 1 to 6.