CN108196561A - The robust wind disturbance resistance position control method and device of a kind of unmanned vehicle - Google Patents

Abstract

The present application provides a method and device for anti-wind interference position control of an unmanned aerial vehicle, including: based on the set pitch reference signal and the acquired pitch state information output by the aircraft, respectively determine the pitch nominal control signal and the pitch robust compensation signal , according to the pitch nominal control signal and the pitch robust compensation signal, determine the first pitch control input signal of the aircraft; based on the set yaw reference signal and the obtained yaw state information and roll state information output by the aircraft, the yaw A navigation mark control signal, and a first yaw robust compensation signal determined according to the yaw reference signal and the yaw state information, and a first yaw robust compensation signal determined according to the yaw nominal control signal and the first yaw robust compensation signal Determine the yaw control input signal of the aircraft; control the flight attitude of the aircraft according to the first pitch control input signal and the yaw control input signal.

Description

Robust anti-wind disturbance position control method and device for unmanned aerial vehicle

technical field

The present application relates to the technical field of aircraft control, in particular, to a method and device for robust anti-wind disturbance position control of an unmanned aircraft.

Background technique

Unmanned aerial vehicles have been widely used in civil and military fields due to their maneuverability, flexibility and versatility. However, the tracking performance of UAVs will be affected by adverse weather conditions such as persistent wind gusts and atmospheric turbulence, and a robust control system is required to suppress these disturbances. Many previous experiments have only studied the stability of UAVs under constant constant wind gusts when hovering or near-hovering, so in the case of time-varying gust disturbances and large maneuvers, the Achieving better tracking performance for aerial vehicles remains challenging.

Contents of the invention

In view of this, the purpose of this application is to provide a robust anti-wind disturbance position control method and device for an unmanned aerial vehicle, which is used to solve the instability problem caused by the influence of wind during the flight of the aircraft in the prior art.

In the first aspect, the embodiment of the present application provides a robust anti-wind disturbance position control method for an unmanned aerial vehicle, the method comprising:

Based on the set pitch reference signal and the acquired pitch state information output by the aircraft, respectively determine the pitch nominal control signal and the pitch robust compensation signal, and determine the first pitch of the aircraft according to the pitch nominal control signal and the pitch robust compensation signal control input signal;

The yaw nominal control signal determined based on the set yaw reference signal and the acquired yaw state information and roll state information output by the aircraft, and the first determined according to the yaw reference signal and the yaw state information. a yaw robust compensation signal, determining a yaw control input signal of the aircraft according to the yaw nominal control signal and the first yaw robust compensation signal;

The flight attitude of the aircraft is controlled according to the first pitch control input signal and the yaw control input signal.

Optionally, based on the set pitch reference signal and the acquired pitch state information output by the aircraft, the pitch nominal control signal is determined, including:

determining a pitch robust compensation state quantity according to the pitch reference signal and preset pitch parameters of the aircraft;

The pitch nominal control signal is determined according to the pitch robust compensation state quantity, the pitch state information, and the pitch nominal control parameter.

Optionally, based on the set pitch reference signal and the acquired pitch state information output by the aircraft, the pitch robust compensation signal is determined, including:

calculating a difference between the value of the pitch state information and the value of the pitch reference signal, and using the difference as a pitch angle error;

A pitch robust compensation signal is determined according to the pitch angle error, the recorded second pitch control input signal and the pitch robust compensation state quantity.

Optionally, the yaw nominal control signal determined based on the set yaw reference signal and the acquired yaw state information and roll state information output by the aircraft includes:

calculating the difference between the value of the yaw state information and the value of the yaw reference signal and using the difference as a yaw angle error;

The yaw nominal control signal is determined based on the yaw angle error, and the roll state information and the yaw state information.

Optionally, the first yaw robust compensation signal determined according to the yaw reference signal and the acquired yaw state information output by the aircraft includes:

determining a yaw state quantity according to the yaw reference signal and the yaw state information;

The first yaw robust compensation signal is determined according to the yaw state quantity and the recorded second yaw robust compensation signal.

In the second aspect, the embodiment of the present application provides a robust anti-wind disturbance position control device for an unmanned aerial vehicle, which includes:

The first processing module is used to respectively determine the pitch nominal control signal and the pitch robust compensation signal based on the set pitch reference signal and the acquired pitch state information output by the aircraft, according to the pitch nominal control signal and the pitch robust compensation signal , to determine the first pitch control input signal of the aircraft;

The second processing module is based on the yaw nominal control signal determined based on the set yaw reference signal and the acquired yaw state information and roll state information output by the aircraft, and according to the yaw reference signal and the yaw state A first yaw robust compensation signal determined by the information, determining an aircraft yaw control input signal according to the yaw nominal control signal and the first yaw robust compensation signal;

A third processing module, configured to control the flight attitude of the aircraft according to the first pitch control input signal and the yaw control input signal.

Optionally, the first processing module is specifically configured to:

determining a pitch robust compensation state quantity according to the pitch reference signal and preset pitch parameters of the aircraft;

The pitch nominal control signal is determined according to the pitch robust compensation state quantity, the pitch state information, and the pitch nominal control parameter.

Optionally, the first processing module is specifically configured to:

calculating a difference between the value of the pitch state information and the value of the pitch reference signal, and using the difference as a pitch angle error;

A pitch robust compensation signal is determined according to the pitch angle error, the recorded second pitch control input signal and the pitch robust compensation state quantity.

Optionally, the second processing module is specifically configured to:

calculating the difference between the value of the yaw state information and the value of the yaw reference signal and using the difference as a yaw angle error;

The yaw nominal control signal is determined based on the yaw angle error, and the roll state information and the yaw state information.

Optionally, the second processing module is specifically configured to:

determining a yaw state quantity according to the yaw reference signal and the yaw state information;

The first yaw robust compensation signal is determined according to the yaw state quantity and the recorded second yaw robust compensation signal.

A method and device for robust anti-wind disturbance position control of an unmanned aerial vehicle provided in the embodiments of the present application, based on the set pitch reference signal and the acquired pitch state information output by the aircraft, respectively determine the pitch nominal control signal and pitch robust Rod compensation signal, according to the pitch nominal control signal and the pitch robust compensation signal, determine the first pitch control input signal of the aircraft; based on the set yaw reference signal and the acquired aircraft output yaw state information and roll state information The determined yaw nominal control signal, and the yaw robust compensation signal determined according to the yaw state quantity, the first yaw control of the aircraft is determined according to the yaw nominal control signal and the yaw robust compensation signal Input signal: According to the first pitch control input signal and the first yaw control input signal, the flight attitude of the aircraft is controlled, which is easy to implement in engineering applications, and it is easy to obtain ideal dynamic and steady-state tracking performance of the aircraft.

In order to make the above-mentioned purpose, features and advantages of the present application more comprehensible, preferred embodiments will be described in detail below together with the accompanying drawings.

Description of drawings

In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following will briefly introduce the accompanying drawings used in the embodiments. It should be understood that the following drawings only show some embodiments of the present application, so It should be regarded as a limitation on the scope, and those skilled in the art can also obtain other related drawings based on these drawings without creative work.

Fig. 1 is a schematic structural view of an unmanned aerial vehicle involved in an embodiment of the present application;

FIG. 2 is a schematic flow diagram of a robust anti-wind disturbance position control method for an unmanned aerial vehicle provided in an embodiment of the present application;

FIG. 3 is a schematic diagram of signal control flow in an aircraft provided by an embodiment of the present application;

FIG. 4 is a schematic structural diagram of a robust anti-wind disturbance position control device for an unmanned aerial vehicle provided in an embodiment of the present application;

5A-B are respectively a response effect diagram of a pitch angle and a response effect diagram of a yaw angle in a static feedback control provided in Embodiment 1 of the present application;

6A-B are respectively a response effect diagram of a pitch angle and a response effect diagram of a yaw angle in a robust control provided in Embodiment 1 of the present application;

7A-C are respectively a response effect diagram, a yaw angle response effect diagram, and a roll angle response effect diagram in a static feedback control provided in Embodiment 2 of the present application;

8A-C are respectively a response effect diagram of a pitch angle, a response effect diagram of a yaw angle, and a response effect diagram of a roll angle in a robust control provided in Embodiment 2 of the present application;

9A-C are respectively a response effect diagram, a yaw angle response effect diagram, and a roll angle response effect diagram in a static feedback control provided in Embodiment 3 of the present application;

10A-C are respectively a response effect diagram of a pitch angle, a response effect diagram of a yaw angle, and a response effect diagram of a roll angle in a robust control provided in Embodiment 3 of the present application.

Detailed ways

In order to make the purpose, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only It is a part of the embodiments of this application, not all of them. The components of the embodiments of the application generally described and illustrated in the figures herein may be arranged and designed in a variety of different configurations. Accordingly, the following detailed description of the embodiments of the application provided in the accompanying drawings is not intended to limit the scope of the claimed application, but merely represents selected embodiments of the application. Based on the embodiments of the present application, all other embodiments obtained by those skilled in the art without making creative efforts belong to the scope of protection of the present application.

As shown in Fig. 1, the signal in the unmanned aerial vehicle control method disclosed in the present application is based on the following model, and the dynamic model of this aircraft can be described as:

Among them, θ(t) is the pitch angle output by the aircraft, φ(t) is the roll angle output by the aircraft, ψ(t) is the yaw angle output by the aircraft, V _f (t) and V _b (t) are respectively applied The voltage on the front and rear motors of the aircraft, w _i (t) (i=θ, φ, ψ) is the additional force generated by the external gust acting on the aircraft, V _op is a normal constant, a _ij (i = θ, φ, ψ; j=1,2) and b _i (i=θ,φ) are aircraft parameters, respectively. Aircraft parameters can be divided into predetermined nominal parameters (indicated by superscript N) and uncertain parameters (indicated by Δ).

where b _ψ = a _ψ2 V _op ,

Define the aircraft pitch control input signal u _θ (t) = V _f (t) + V _b (t) and yaw control input signal u _ψ (t) = V _f (t) - V _b (t), then, The aircraft model of formula group (1) can be rewritten as:

Among them, q _i (t) (i=θ, φ, ψ1) is called the equivalent disturbance, including parameter change, nonlinearity, coupling and external wind disturbance, etc. The equivalent disturbance is expressed by the following formula:

Among them, u _θ (t) is the first pitch control input signal;

u _ψ (t) is the first yaw control input signal;

w _θ (t) The additional force generated by the external gust acting on the pitch channel of the aircraft;

w _ψ (t) The additional force generated by the external gust acting on the yaw channel of the aircraft;

w _φ (t) The additional force generated by the external gust acting on the roll channel of the aircraft;

Differentiating the two sides of the yaw angle equation in formula group (2) can get:

Among them, ψ ⁽³⁾ (t) is the third order differential of yaw angle; is the first order differential of the roll angle; is the first order differential of the yaw equivalent disturbance;

Differentiate formula (4) further to get:

Among them, ψ ⁽⁴⁾ (t) is the fourth order differential of yaw angle.

Substituting formula group (2) and formula (4) into formula (5) can get:

where q _ψ (t)

Among them, q _φ (t) is the rolling equivalent disturbance signal;

Define the yaw state quantity:

where x _θ1 (t) = θ(t) - θ _r (t)

x _ψ1 (t) = ψ(t) - ψ _r (t)

Among them, x _θ1 (t) is pitch angle error;

θ _r (t) is the pitch angle reference signal;

ψ _r (t) is the yaw angle reference signal;

Indicates that x _ψ5 (t) is obtained by integrating the yaw angle error x _ψ1 (t);

Adding x _θ3 (t) and x _ψ5 (t) is used to design the static feedback nominal controller, the purpose is to improve the steady-state performance of the nominal closed-loop system. define parameters and Then use the following formula to express the error system in state space form

where the matrix:

B _ψ ＝[0 0 0 1 0] ^T

Robust compensator state quantity:

Among them, r _θ (t) is the pitch robust compensation state quantity;

r _ψ (t) is the yaw robust compensation state quantity; the embodiment of the present application provides a robust anti-wind disturbance position control method for an unmanned aerial vehicle, as shown in Figure 2, the method includes the following steps:

S201. Based on the set pitch reference signal and the acquired pitch state information output by the aircraft, respectively determine the pitch nominal control signal and the pitch robust compensation signal, and determine the first pitch of the aircraft according to the pitch nominal control signal and the pitch robust compensation signal. a pitch control input signal;

Optionally, based on the set pitch reference signal and the acquired pitch state information output by the aircraft, the pitch nominal control signal is determined, including:

determining a pitch robust compensation state quantity according to the pitch reference signal and preset pitch parameters of the aircraft;

The pitch nominal control signal is determined according to the pitch robust compensation state quantity, the pitch state information, and the pitch nominal control parameter.

Optionally, based on the set pitch reference signal and the acquired pitch state information output by the aircraft, the pitch robust compensation signal is determined, including:

calculating a difference between the value of the pitch state information and the value of the pitch reference signal, and using the difference as a pitch angle error;

A pitch robust compensation signal is determined according to the pitch angle error, the recorded second pitch control input signal and the pitch robust compensation state quantity.

As shown in Fig. 3, the first pitch control input signal u _θ (t) has two parts: pitch nominal control input and pitch robust compensation signal v _θ (t) based on robust compensation technology. Therefore, the first pitch control input signal is as follows:

Among them, u _θ (t) is the first pitch control input signal, is the pitch nominal control signal, and v _θ (t) is the pitch robust compensation signal.

Neglecting the pitch equivalent disturbance q _θ (t), the linear static feedback control law is designed as follows:

Among them, K _θ is the pitch controller parameter, X _θ (t) is the pitch error state quantity, r _θ (t) is the pitch robust compensator state quantity, is the first pitch nominal parameter of the aircraft.

X _θ (t)＝[x _θ1 (t) x _θ2 (t) x _θ3 (t)] ^T

x _θ1 (t)=θ(t) _-θr (t)

Among them, θ(t) is the pitch state information, θ _r (t) is the pitch reference signal;

x _θ1 (t) is the first pitch error state quantity;

x _θ2 (t) is the first order differential of the first pitch error state quantity.

It is necessary to select an appropriate pitch controller parameter K _θ =[k _θj ] _1×3 so that the nominal controller system matrix A _θH =A _θ +B _θ K _θ is a Hurwitz matrix. Among them, A _θ is the nominal controller system matrix, and B _θ is the aircraft output matrix.

Among them, A _θH is the pitch nominal controller system matrix; X _θ (t) is the pitch error state quantity; is the first-order differential of the pitch error state quantity; B _θ is the output matrix of the aircraft; is the first pitch nominal parameter of the aircraft; v _θ (t) is the recorded second pitch robust compensation signal; q _θ (t) is the pitch equivalent interference signal;

Among them, r _θ (t) is the pitch robust compensation state quantity;

is the first control parameter of the aircraft;

is the second control parameter of the aircraft;

Laplace transform formula: s is the Laplace operator, and x(s) is the representation of the time function x(t) in the complex frequency domain. The variable transformations below are all transformed according to this formula.

In order to suppress the influence of uncertainty, the pitch robust compensation function F _θ (s) is added, and the resulting pitch robust compensation signal v _θ (s) is expressed by the following formula:

Among them, q _θ (s) is the equivalent disturbance parameter in pitch frequency domain;

Among them, s is the Laplacian operator, and,

Among them, f _θ1 is the first pitching robust compensation parameter, and f _θ2 is the second pitching robust compensator parameter;

Its parameters f _θ1 >0, f _θ2 >0. From equations (14)-(16), the state description of the robust compensator for the pitch channel can be obtained:

where x _θ1 (t) = θ(t) - θ _r (t)

Among them, v _θ (t) is the pitch robust compensation signal, z _1θ (t) is the state quantity of the first pitch robust compensator, z _2θ (t) is the state quantity of the second pitch robust compensator; f _θ1 is the state quantity of the first pitch robust compensator A pitching robust compensator parameter, f _θ2 is the second pitching robust compensator parameter; is the first pitch nominal parameter of the aircraft; u _θ0 (t) is the recorded second pitch control input signal; x _θ1 (t) is the pitch angle error, θ(t) is the pitch information output by the aircraft, θ _r (t ) is the set pitch reference information, u _θ0 (t) is the recorded second pitch control input signal.

S202, the yaw nominal control signal determined based on the set yaw reference signal and the acquired yaw state information and roll state information output by the aircraft, and the yaw nominal control signal determined according to the yaw reference signal and the yaw state information a first yaw robust compensation signal, determining a yaw control input signal of the aircraft according to the yaw nominal control signal and the first yaw robust compensation signal;

Optionally, the yaw nominal control signal determined based on the set yaw reference signal and the obtained yaw state information and roll state information output by the aircraft includes:

calculating the difference between the value of the yaw state information and the value of the yaw reference signal and using the difference as a yaw angle error;

The yaw nominal control signal is determined based on the yaw angle error, and the roll state information and the yaw state information.

Optionally, the first yaw robust compensation signal determined according to the yaw reference signal and the acquired yaw state information output by the aircraft includes:

determining a yaw state quantity according to the yaw reference signal and the yaw state information;

The first yaw robust compensation signal is determined according to the yaw state quantity and the recorded second yaw robust compensation signal.

The first yaw control input signal u _ψ (t) has two parts: yaw nominal control input and the yaw robust compensation signal v _ψ (t) based on the robust compensation technique. Therefore, the first yaw control input signal is as follows:

Among them, u _ψ (t) is the first yaw control input signal, is the nominal control signal for yaw, and v _ψ (t) is the robust compensation signal for yaw.

Neglecting the uncertainty term q _ψ (t), then design a linear static feedback control law as follows

Among them, K _ψ is the parameter of yaw controller, X _ψ (t) is the state quantity of yaw error, r _ψ (t) is the state of yaw robust compensator, is the nominal parameter of aircraft yaw.

It is necessary to select the appropriate yaw controller parameter K _ψ = [k _ψj ] _1×5 so that the matrix A _ψH = A _ψ + B _ψ K _ψ is a Hurwitz matrix.

Among them, A _ψH is the yaw nominal controller system matrix;

X _ψ (t) is the yaw error state quantity matrix;

B _ψ is the aircraft output matrix;

is the nominal parameter of aircraft yaw;

v _ψ (t) is the recorded second yaw robust compensation signal;

q _ψ (t) is the yaw equivalent interference signal;

is the first order differential of the yaw error state quantity;

and ψ ⁽³⁾ (t) cannot be obtained by measurement, so redesign the yaw nominal control input

Among them, k _ψ1 is the first yaw control parameter, k _ψ2 is the second yaw control parameter, k _ψ3 is the third yaw control parameter, k _ψ4 is the fourth yaw control parameter, and k _ψ5 is the fifth yaw control parameter control parameters; r _ψ (t) is the state of the yaw robust compensator; x _ψ1 (t) is the first yaw state quantity, x _ψ2 (t) is the second yaw state quantity, x _ψ5 (t) is the second yaw state quantity Three yaw state quantities; is the first yaw nominal parameter of the aircraft, is the second yaw nominal parameter of the aircraft, is the third yaw nominal parameter of the aircraft; is the first order differential of the roll angle; is the first order differential of the yaw angle, is the second order differential of the yaw angle reference signal, is the third order differential of the yaw angle reference signal.

Laplace transform formula: s is the Laplace operator, and x(s) is the representation of the time function x(t) in the complex frequency domain. The variable transformations below are all transformed according to this formula.

In order to suppress the influence of uncertainty, the pitch robust compensation function F _θ (s) is added, and the resulting pitch robust compensation signal v _θ (s) is expressed by the following formula:

Among them, q _ψ (s) is the equivalent disturbance signal in the yaw frequency domain;

Among them, s is the Laplacian operator, and,

Among them, f _ψ1 is the first yaw robust compensator parameter, f _ψ2 is the second yaw robust compensator parameter

The parameter f _ψj of the yaw robust compensator >0 (j=1,2), from equations (14)-(16), the state space description of the robust compensator for the pitch angle channel can be obtained.

because and ψ ⁽³⁾ (t) cannot be measured, define the function d _ψH (s) = det(sI-A _ψH ) and the state y _ψ = x _ψ1 + x _ψ5 . The equivalent interference signal can be obtained from formula (20):

Among them, y _ψ (s) is the frequency domain yaw state quantity of the aircraft; v _ψ (s) is the yaw robust frequency domain compensation signal.

From formulas (24) and (25):

Among them, v _ψ (s) is the yaw robust frequency domain compensation signal; y _ψ (s) is the yaw frequency domain state quantity of the aircraft; s is the Laplacian operator; d _ψH (s) is the yaw transfer function; v _ψ0 (s) is the recorded second yaw robust frequency domain compensation signal.

v _ψ (s) is converted to v _ψ (t) by the Laplace inverse transformation formula.

Among them, y _ψ (s) can be obtained by processing the yaw state quantity y _ψ (t) through the Laplace transformation formula.

y _ψ (t) = x _ψ1 (t) + x _ψ5 (t)

x _ψ1 (t) = ψ(t) - ψ _r (t)

Among them, ψ(t) is the yaw state information output by the aircraft, and ψ _r (t) is the yaw reference signal.

The robust compensator turns on only when disturbances are present. A robust control part is introduced to suppress the influence of external wind disturbance, which ensures the asymptotic stability of the closed-loop system when there is uncertainty.

S203. Control the flight attitude of the aircraft according to the first pitch control input signal and the yaw control input signal.

Wherein, the flight attitude of the unmanned aerial vehicle is periodically controlled through the first pitch control input signal and the yaw control input signal, and the above-mentioned second yaw control input signal, the second pitch control input signal, and the second pitch robust compensation signal are all Signal recorded for the previous period adjacent to the current period.

In Example 1, the robust control method is compared with the static feedback control method in the hovering state. The aircraft is subjected to a gust of wind with a fixed velocity of 4.2 m/s. The purpose of the experiment is to keep the pitch angle and yaw angle of the aircraft stable at 0 degrees. The experimental results of the static feedback controller and the robust controller are shown in Fig. 5A-B and Fig. 6A-B respectively. It can be seen that both methods keep both angles stable in the presence of gusts, despite differences in tracking error.

In Example 2, large maneuvers were performed under a wind gust with a wind speed of 4.2 m/s. The aircraft needs to track two large-angle reference signals, and the two channels are coupled. The reference signal is given by:

in, and are the reference signals of the pitch angle and the yaw angle respectively, the reference signal is a square wave signal with a period of 60s, and its amplitudes are 20 degrees and 60 degrees respectively. When performing large maneuvers, it can be seen that the yaw and pitch channels are affected by external forces generated by gusts of wind. Comparing the tracking performance of the static feedback controller with the robust controller, the experimental results are shown in Fig. 7A-C and Fig. 8A-C, respectively. It can be seen that the tracking response of the static feedback controller is greatly affected by the external wind disturbance, while the robust method has better dynamic performance. It can be seen from Fig. 7A-C and the third sub-figure of Fig. 8A-C that the roll angle has not reached the mechanical limit.

In embodiment 3, the three-degree-of-freedom aircraft is required to achieve large maneuvering position tracking under unknown time-varying wind interference. The desired locus is the same as in Example 2. Figures 9A-C and Figures 10A-C depict the tracking response with a static feedback controller and a robust controller, respectively. The corresponding external wind speeds are shown in the third subgraphs of Figures 9A-C and Figures 10A-C, respectively. It can be seen from the figure that the robust control method has better dynamic and steady-state performance than the static feedback control method, and realizes large-angle tracking control.

The embodiment of the present application provides an anti-wind disturbance position control device for an unmanned aerial vehicle, as shown in Figure 4, the device includes:

The first processing module 41 is used to determine the pitch nominal control signal and the pitch robust compensation signal respectively based on the set pitch reference signal and the acquired pitch state information output by the aircraft, according to the pitch nominal control signal and the pitch robust compensation signal signal, to determine the first pitch control input signal of the aircraft;

The second processing module 42 is based on the yaw nominal control signal determined based on the set yaw reference signal and the acquired yaw state information and roll state information output by the aircraft, and according to the yaw reference signal and the yaw The first yaw robust compensation signal determined by the state information, the yaw control input signal of the aircraft is determined according to the yaw nominal control signal and the first yaw robust compensation signal;

The third processing module 43 is configured to control the flight attitude of the aircraft according to the first pitch control input signal and the yaw control input signal.

Optionally, the first processing module 41 is specifically configured to:

determining a pitch robust compensation state quantity according to the pitch reference signal and preset pitch parameters of the aircraft;

The pitch nominal control signal is determined according to the pitch robust compensation state quantity, the pitch state information, and the pitch nominal control parameter.

Optionally, the first processing module 41 is specifically configured to:

calculating a difference between the value of the pitch state information and the value of the pitch reference signal, and using the difference as a pitch angle error;

A pitch robust compensation signal is determined according to the pitch angle error, the recorded second pitch control input signal and the pitch robust compensation state quantity.

Optionally, the second processing module 42 is specifically configured to:

calculating the difference between the value of the yaw state information and the value of the yaw reference signal and using the difference as a yaw angle error;

The yaw nominal control signal is determined based on the yaw angle error, and the roll state information and the yaw state information.

Optionally, the second processing module 42 is specifically configured to:

determining a yaw state quantity according to the yaw reference signal and the yaw state information;

The first yaw robust compensation signal is determined according to the yaw state quantity and the recorded second yaw robust compensation signal.

In the embodiments provided in this application, it should be understood that the disclosed devices and methods may be implemented in other ways. The device embodiments described above are only illustrative. For example, the division of the units is only a logical function division. In actual implementation, there may be other division methods. For example, multiple units or components can be combined or May be integrated into another system, or some features may be ignored, or not implemented. In another point, the mutual coupling or direct coupling or communication connection shown or discussed may be through some communication interfaces, and the indirect coupling or communication connection of devices or units may be in electrical, mechanical or other forms.

The units described as separate components may or may not be physically separated, and the components shown as units may or may not be physical units, that is, they may be located in one place, or may be distributed to multiple network units. Part or all of the units can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

In addition, each functional unit in the embodiments provided by the present application may be integrated into one processing unit, each unit may exist separately physically, or two or more units may be integrated into one unit.

If the functions described above are realized in the form of software function units and sold or used as independent products, they can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present application is essentially or the part that contributes to the prior art or the part of the technical solution can be embodied in the form of a software product, and the computer software product is stored in a storage medium, including Several instructions are used to make a computer device (which may be a personal computer, a server, or a network device, etc.) execute all or part of the steps of the methods described in the various embodiments of the present application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM, Read-Only Memory), random access memory (RAM, Random Access Memory), magnetic disk or optical disk and other media that can store program codes. .

It should be noted that like numerals and letters denote similar items in the following drawings, therefore, once an item is defined in one drawing, it does not require further definition and explanation in subsequent drawings, In addition, the terms "first", "second", "third", etc. are only used for distinguishing descriptions, and should not be construed as indicating or implying relative importance.

Finally, it should be noted that: the above-described embodiments are only specific implementations of the application, used to illustrate the technical solutions of the application, rather than limiting it, and the scope of protection of the application is not limited thereto, although referring to the aforementioned The embodiment has described this application in detail, and those of ordinary skill in the art should understand that any person familiar with this technical field can still modify the technical solutions described in the foregoing embodiments within the technical scope disclosed in this application Changes can be easily imagined, or equivalent replacements can be made to some of the technical features; and these modifications, changes or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present application. All should be covered within the scope of protection of this application. Therefore, the protection scope of the present application should be based on the protection scope of the claims.

Claims (10)

1. A robust wind disturbance resistant position control method of an unmanned aerial vehicle is characterized by comprising the following steps:

respectively determining a pitching nominal control signal and a pitching robust compensation signal based on a set pitching reference signal and the obtained pitching state information output by the aircraft, and determining a first pitching control input signal of the aircraft according to the pitching nominal control signal and the pitching robust compensation signal;

determining a yaw nominal control signal based on a set yaw reference signal and the obtained yaw state information and roll state information output by the aircraft, determining a first yaw robust compensation signal according to the yaw reference signal and the obtained yaw state information, and determining a yaw control input signal of the aircraft according to the yaw nominal control signal and the first yaw robust compensation signal;

and controlling the flight attitude of the aircraft according to the first pitching control input signal and the yawing control input signal.

2. The method of claim 1, wherein determining a pitch nominal control signal based on the set pitch reference signal and the acquired pitch status information of the aircraft output comprises:

determining a pitching robust compensation state quantity according to the pitching reference signal and a preset pitching parameter of the aircraft;

and determining the pitching nominal control signal according to the pitching robust compensation state quantity, the pitching state information and the pitching nominal control parameter.

3. The method of claim 1, wherein determining a pitch robust compensation signal based on the set pitch reference signal and the acquired pitch state information of the aircraft output comprises:

calculating a difference value between the value of the pitching state information and the value of the pitching reference signal, and taking the difference value as a pitching angle error;

and determining a pitching robust compensation signal according to the pitching angle error, the recorded second pitching control input signal and the recorded pitching robust compensation state quantity.

4. The method of claim 1, wherein the yaw nominal control signal determined based on the set yaw reference signal and the obtained yaw state information and roll state information of the aircraft output comprises:

calculating a difference value between the value of the yaw state information and the value of the yaw reference signal, and taking the difference value as a yaw angle error;

and determining the yaw nominal control signal based on the yaw angle error, the roll state information and the yaw state information.

5. The method of claim 1, wherein determining a first yaw robust compensation signal based on the yaw reference signal and the obtained yaw state information of the aircraft output comprises:

determining a yaw state quantity according to the yaw reference signal and the yaw state information;

and determining the first yaw robust compensation signal according to the yaw state quantity and the recorded second yaw robust compensation signal.

6. A robust wind disturbance rejection position control apparatus for an unmanned aerial vehicle, the apparatus comprising:

the first processing module is used for respectively determining a pitching nominal control signal and a pitching robust compensation signal based on a set pitching reference signal and the obtained pitching state information output by the aircraft, and determining a first pitching control input signal of the aircraft according to the pitching nominal control signal and the pitching robust compensation signal;

the second processing module is used for determining a yaw nominal control signal based on a set yaw reference signal, the obtained yaw state information and the obtained roll state information output by the aircraft, determining a first yaw robust compensation signal according to the yaw reference signal and the yaw state information, and determining a yaw control input signal of the aircraft according to the yaw nominal control signal and the first yaw robust compensation signal;

and the third processing module is used for controlling the flight attitude of the aircraft according to the first pitching control input signal and the yawing control input signal.

7. The apparatus of claim 6, wherein the first processing module is specifically configured to:

determining a pitching robust compensation state quantity according to the pitching reference signal and a preset pitching parameter of the aircraft;

and determining the pitching nominal control signal according to the pitching robust compensation state quantity, the pitching state information and the pitching nominal control parameter.

8. The apparatus of claim 6, wherein the first processing module is specifically configured to:

calculating a difference value between the value of the pitching state information and the value of the pitching reference signal, and taking the difference value as a pitching angle error;

and determining a pitching robust compensation signal according to the pitching angle error, the recorded second pitching control input signal and the recorded pitching robust compensation state quantity.

9. The apparatus of claim 6, wherein the second processing module is specifically configured to:

calculating a difference value between the value of the yaw state information and the value of the yaw reference signal, and taking the difference value as a yaw angle error;

and determining the yaw nominal control signal based on the yaw angle error, the roll state information and the yaw state information.

10. The apparatus of claim 6, wherein the second processing module is specifically configured to:

determining a yaw state quantity according to the yaw reference signal and the yaw state information;

and determining the first yaw robust compensation signal according to the yaw state quantity and the recorded second yaw robust compensation signal.