Aircraft attitude tracking control method, system and computer equipment

Abstract

The invention provides an aircraft attitude tracking control method, system and computer equipment, which belongs to the field of aircraft attitude control, comprising: obtaining aircraft operating parameters and establishing a kinematic model of the attitude tracking error of the aircraft. A specified time performance constraint function and attitude error boundary based on the hyperbolic cosecant composite function are then constructed. A constrained attitude error conversion function and its inverse function are utilized to obtain an unconstrained attitude tracking conversion error. Virtual control instructions are solved in kinematics using a sigmoid tracking differentiator to obtain a numerical derivative. The control moment in the dynamics is solved and the complete form of the attitude tracking algorithm is given for the aircraft attitude tracking system. This method provides a computationally efficient means of controlling aircraft to ensure that error convergence is achieved within a preset time.

Description

CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to Chinese Patent Application No. 202311664381.9, filed on Dec. 6, 2023 before the China National Intellectual Property Administration, the disclosure of which is incorporated herein by reference in entirety.

TECHNICAL FIELD

The present invention belongs to the technical field of attitude control of rigid body aircraft, and specifically relates to an aircraft attitude tracking control method, system and computer equipment.

BACKGROUND

With the advancement of technology and the continuous growth of mission requirements, aircraft and space missions have become increasingly complex and diversified, and mission types have gradually evolved from traditional single-aircraft missions such as earth observation and communication relay to emerging multi-aircraft cluster missions such as rendezvous, docking and formation flight. For all kinds of aircraft, effective attitude control is related to various flight processes such as aircraft take-off, landing and flight direction, and is an important foundation to ensure the smooth execution of various aerospace missions. Due to the relatively close relative distance between the aircraft in a swarm mission, for the sake of the mission time window and safety collision avoidance, it is necessary to pay attention not only to the final steady-state goal of attitude control, but also to transient performance characteristics such as convergence speed, overshoot, and trajectory boundary during the motion processes. Although the transient performance of the system can be adjusted by changing the parameters of traditional control methods, transient performance is not considered in the control design process. Thus, there is no theoretical framework to guarantee suitable transient performance, making it difficult to meet the requirements for the safety and reliability of attitude control in swarm aerospace missions.

Preset performance control is a convenient, effective and widely used method to achieve a priori constraints on transient performance such as overshoot and convergence time. However, there are two main problems in current implementations of preset performance attitude control of aircraft: (1) the traditional exponential performance constraint function has limitations of not enabling preset convergence time, requiring initial error information for the system, and not allowing fine adjustment of convergence speed. These functions have the characteristics of fast convergence speed in the initial stage, and large systematic error in the initial stage. Thus, it is easy to cause phenomena such as surge of initial control quantity and jitter, which then cause performance degradation or even instability of the closed-loop system. (2) Most of the existing control algorithm designs focus is on theoretical derivation results, and simulation tests are carried out on high-performance computers with abundant computing resources. However, available computing resources in actual airborne flight control systems are very limited, and most of the attitude control algorithms do not take into account the need for lightweight processing of calculations. Thus it is easy for a huge disparity between simulation test performance and practical application effects to emerge.

SUMMARY

In order to solve the problem of high-performance, robust practical aircraft attitude tracking control under the condition of limited computing resources, the present invention provides an aircraft attitude tracking control method, system and computer equipment, which is based on the composite hyperbolic cosecant composite function to carry out the design of an attitude tracking controller with preset performance at specified time, which can simultaneously solve the problems in existing methods of infinite theoretical convergence time, dependence on initial state information, and difficulty in adjusting convergence speed. Based on a sigmoid tracking differentiator, the method obtains the derivative information of the virtual control instruction in the backstep design, which can solve the problems of complex controller form and the sharp increase of online computation caused by the analytical derivation of existing methods.

In order to achieve the above purpose, the invention provides the following technical scheme:

An aircraft attitude tracking control method comprises the following steps:

• • Obtain the operating parameters of the aircraft;
  • Based on the operation parameters and quaternion description of the aircraft, the kinematic model and dynamic model of the attitude tracking error of the aircraft were established;
  • Based on the hyperbolic cosecant composite function, the performance constraint function of the specified time convergence is constructed, and the attitude tracking error constraint boundary is set based on the performance constraint function.

Construct a constrained attitude error conversion function based on the attitude tracking error constraint boundary, and solve the inverse function of the constrained attitude error conversion function to obtain the dynamic equation of the unconstrained attitude tracking conversion error.

According to the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error, establish a dynamic equation of the unconstrained conversion.

Solve the virtual control instructions in the kinematic model based on the dynamic equation and the backstepping method of the unconstrained transformation error, and construct a Sigmoid tracking differentiator to obtain the numerical derivative of the virtual control instructions;

Wherein the control torque is designed based on the virtual control instructions and numerical derivatives, the actual control input is solved by the backstepping method in the dynamic model, and the control torque and the actual control input are used in the aircraft attitude tracking system to realize the aircraft attitude tracking control.

Preferably, the specific construction methods of the kinematic model and dynamic model of the aircraft attitude tracking error are as follows:

F_I denotes the earth's inertial system, F_B is the aircraft system, F_D is the desired motion coordinate system, and the motion of the aircraft unit is expressed as F_B relative to F₁ motion. The single body attitude kinematics model and dynamics model based on quaternion description are:

$\begin{array}{cc}{\dot{q}}_B=\frac{1}{2}{q}_B{\omega }_B^B& \left(1\right)\end{array}$ $\begin{array}{cc}{J}_B{\dot{\omega }}_B^B=-{\omega }_B^B\times J_B{\omega }_B^B+{\tau }^B& \left(2\right)\end{array}$

• • wherein, q_B and ω_B^B are the attitude quaternion and angular velocity of the aircraft, {dot over (q)}_B and {dot over (ω)}_B^B are the derivatives of the attitude quaternion and angular velocity of the aircraft, respectively, J_B is the moment of inertia of the aircraft and τ^B is the control input torque;

Desired motion is described as F_D relative to F_I motion that satisfies:

$\begin{array}{cc}{\dot{q}}_D=\frac{1}{2}{q}_D{\omega }_D^D& \left(3\right)\end{array}$

• • where q_D and ω_D^D are the desired attitude and the desired angular velocity, respectively;

The tracking motion of the aircraft is described as the motion of F_B relative to F_D. Based on formulas (1)-(3), the kinematic model and dynamic model of the aircraft attitude tracking error are established as follows:

$\begin{array}{cc}{\dot{q}}_e=\frac{1}{2}{q}_e{\omega }_e^B& \left(4\right)\end{array}$ $\begin{array}{cc}{J}_B{\dot{\omega }}_e^B=-{\omega }_B^B\times J_B{\omega }_B^B+J_B\left(C_e{\dot{\omega }}_D^D-{\omega }_e^B\times C_e{\omega }_D^D\right)+{\tau }^B& \left(5\right)\end{array}$

• • wherein, q_e=[q_e0, q_e^T]^T is the attitude tracking error quaternion, q_e0 is the scalar part of the quaternion, q_e=[q_e1, q_e2, q_e3]^T is the vector part of the quaternion, ω_e^B is the angular velocity tracking error, J_B is the moment of inertia of the aircraft, ω_B^B is the angular velocity of the aircraft, WB and co are the expected angular velocity and angular acceleration respectively, C_e is the rotation matrix, and τ^B is the control torque.

Preferably, the performance constraint function of the specified time convergence is constructed based on the hyperbolic cosecant composite function, and the attitude tracking error constraint boundary is set based on the performance constraint function, specifically:

Based on the hyperbolic cosecant composite function, a performance constraint function for the convergence of the specified time is constructed ρ_i(t):

$\begin{array}{cc}{\rho }_i\left(t\right)=\{\begin{array}{cc}\mathrm{csch}\left({\rho }_{i0}+\frac{r_{i}t}{T_i-t}\right)+{\rho }_{i\infty },& 0\le t<T_i\\ {\rho }_{i\infty }& t<T_i\end{array},& \left(6\right)\end{array}$ $i=1,2,3$

• • wherein, t is the time variable, csch(⋅) is the hyperbolic cosecant composite function, T_i is the specified convergence time, ρ_i0 is used to adjust the initial point of the boundary function, r_i is used to adjust the convergence speed of the function, and ρ_i∞ is the steady-state upper bound of the error;

Based on the performance constraint function, the attitude tracking error constraint boundary is designed as:

$\begin{array}{cc}-\underset{\_}{{\delta }_i}{\rho }_i\left(t\right)<q_{\mathrm{ei}}<\stackrel{\_}{{\delta }_i}{\rho }_i\left(t\right),& \left(7\right)\end{array}$ $i=1,2,3.$

• • where δ_i and δ_i are the overshoot limit parameters, and δ_i, δ_i∈(0,1], q_ei are the corresponding components in the quaternion of the attitude tracking error.

Preferably, the constrained attitude error conversion function is constructed based on the attitude tracking error constraint boundary, and the inverse function of the constrained attitude error conversion function is solved to obtain a dynamic equation for the unconstrained attitude tracking conversion error, specifically:

Based on the boundary constraint of the attitude tracking error, the constrained attitude error conversion function is constructed, which is used to convert the constrained attitude tracking error into an unconstrained form, and the constrained attitude error conversion function S(⋅) is designed as follows

$\begin{array}{cc}S\left({\epsilon }_i\right)=\frac{\overline{{\delta }_i}{e}^{{\epsilon }_i}-\underset{¯}{{\delta }_i}{e}^{-{\epsilon }_i}}{e^{{\epsilon }_i}+e^{-{\epsilon }_i}},& \left(8\right)\end{array}$ $i=1,2,3$

• • wherein ε_i is the unconstrained transformation error, and the constrained actual attitude error q_e and the unconstrained transformation error ε_i satisfy:

$\begin{array}{cc}{q}_{\mathrm{ei}}={\rho }_{i}S\left({\epsilon }_i\right),& \left(9\right)\end{array}$ $i=1,2,3$

• • based on equation (9), the inverse function of the constrained attitude error conversion function S⁻¹(⋅) is solved to give an explicit expression of the transformation error:

$\begin{array}{cc}{\epsilon }_i=S^{-1}\left(z_i\right)=\frac{1}{2}\ln \left(\frac{z_i+\underset{\_}{{\delta }_i}}{\stackrel{\_}{{\delta }_i}-z_i}\right),& \left(10\right)\end{array}$ $z_i=q_{\mathrm{ei}}\left(t\right)/{\rho }_i\left(t\right),$ $i=1,2,3$

• • where ε_i is the conversion error, and z_i is the normalized conversion error.

Preferably, the dynamic equation of the unconstrained conversion error is established according to the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error, specifically:

$\begin{array}{cc}\dot{\epsilon }={\mathrm{RQ}}_e{\omega }_e+v& \left(11\right)\end{array}$

• • where ε=[ε₁, ε₂, ε₃]^T is the transformation error vector, {dot over (ε)} represents the derivative of ε; R=diag[r₁, r₂, r₃], is a third-order diagonal matrix;

$r_i=\frac{\partial S^{-1}}{\partial z_i}·\frac{1}{{\rho }_i},$ $i=1,2,3,$ $Q_e=\frac{q_{e0}{I}_3+q_e^{\times }}{2},$ $v={\left[v_1,v_2,v_3\right]}^T,,$ $v_i=-r_i\frac{q_{\mathrm{ei}}{\stackrel{.}{\rho }}_i}{{\rho }_i},$ $i=1,2,3;$

S⁻¹ is the inverse of the error conversion function, r_i is an element in the diagonal matrix, Q_e is the attitude multiplication matrix, q_e0 is the scalar part of the quaternion, I₃ is the third-order identity matrix, q_e^× is the attitude cross product matrix, ν is the auxiliary vector, and ν_i is an element in the auxiliary vector.

Preferably, the virtual control instructions of the kinematic loop are solved based on the dynamic equation and the backstepping method of the unconstrained conversion error, and a sigmoid tracking differentiator is constructed to obtain the numerical derivative of the virtual control instructions, specifically:

Based on the dynamic equation of the unconstrained transformation error, the virtual control instructions in the kinematic model are designed as:

$\begin{array}{cc}{\omega }_{\mathrm{ec}}=-Q_e^{-1}{R}^{-1}\left(k_p\epsilon +v\right)& \left(12\right)\end{array}$

• • wherein k_p>0 is the control parameter, and the −1 in the upper right corner of the parameter represents the inversion.

To build a tracking differentiator based on sigmoid function:

$\{\begin{array}{c}{\stackrel{.}{x}}_1=x_2\\ {\stackrel{.}{x}}_2=-r^2\mathrm{sig}\left(x_1-w;a_1,b_1\right)-\mathrm{sig}\left(x_2/r;a_2{b}_2\right)\end{array},$ $\mathrm{sig}\left(x;a,b\right)=a\left[{\left(1+e^{-\mathrm{bx}}\right)}^{-1}-0.5\right]$

• • where sig(⋅) is the sigmoid function, a₁, b₁, a₂, b₂>0 is the design parameter, r>0 is the acceleration factor, w is the reference signal input, and x₁ and x₂ are the two states of the tracking differentiator.

The sigmoid tracking differentiator is used to find the derivative of the virtual control command:

$\begin{array}{cc}\{\begin{array}{c}{\stackrel{.}{\omega }}_{\mathrm{ecfi}}=x_{\omega i}\\ {\stackrel{.}{x}}_{\omega i}=r_i^2[\mathrm{sig}\left({\omega }_{\mathrm{ecfi}}-{\omega }_{\mathrm{eci}};a_{i1},b_{i1}\right)+\mathrm{sig}\left(x_{\omega i}/r_i;a_{i2}{b}_{i2}\right)\end{array},& \left(14\right)\end{array}$ $i=1,2,3$

• • wherein r_i, a_i1, a_i2, b_i1, b_i2>0, ω_ecf=[ω_ec1, ω_ec2, ω_ec3]^T and {dot over (ω)}_ecf=[x_ω1, x_ω2, x_ω3]^T are the filtered signal and the filtered derivative of the virtual control command, respectively, and x_ωi is a state of the tracking differentiator.

Preferably, the control torque is designed based on the virtual control instruction and the numerical derivative, the actual control input quantity is solved by using the backstepping method in the dynamic model, and the control torque and the actual control input quantity are used in the aircraft attitude tracking system to realize the aircraft attitude tracking control, specifically:

Based on the virtual control instruction and its filtering signal, the angular velocity loop tracking error is defined as z_ω=ω_e−ω_ecf, and its dynamic equation is:

$\begin{array}{cc}{\dot{z}}_w=J_B^{-1}(-{\omega }_B^B\times J_B{\omega }_B^B-J_B\left(C_e{\dot{\omega }}_D^D-{\omega }_e^B\times C_e{\omega }_D^D\right)+{\tau }^B& \left(16\right)\end{array}$

According to equation (16) and the principle of backstepping method, the control torque input τ^B is designed as:

$\begin{array}{cc}{\tau }^B=-\Xi +J_B{\dot{\omega }}_{ecf}-k_d{z}_{\omega }-Q_e^T{R}^T\epsilon & \left(17\right)\end{array}$

• • wherein k_d>0 is the control parameter, and Ξ=−ω_B^B×J_Bω_B^B−J_B(C_e{dot over (ω)}_D^D=ω_e^B×C_eω_D^D) is the auxiliary variable;

Substituting the control torque input τ^B into the dynamical model, the closed-loop dynamics equation is as follows:

$\begin{array}{cc}{\dot{z}}_{\omega }=-k_d{z}_{\omega }-Q_e^T{R}^T\epsilon & \left(18\right)\end{array}$

The candidate Lyapunov equation for defining the overall closed-loop system is:

$\begin{array}{cc}V=\frac{1}{2}{\epsilon }^T\epsilon +\frac{1}{2}{z}_{\omega }^T{z}_{\omega }& \left(19\right)\end{array}$

The derivative V along the trajectory of the closed-loop system can be obtained {dot over (V)}=k_pε²−k_dz_ω²+ε^TRQ_es.

Preferably, the aircraft attitude tracking controller is:

$\begin{array}{cc}\{\begin{array}{c}{\omega }_{ec}=-2Q_e^{-1}{R}^{-1}\left(k_p\epsilon +v\right)\\ \{\begin{array}{c}{\dot{\omega }}_{\mathrm{ecfi}}=x_{\omega i},i=1,2,3\\ {\stackrel{.}{x}}_{\omega i}=r_i^2[\mathrm{sig}\left({\omega }_{\mathrm{ecfi}}-{\omega }_{\mathrm{eci}};a_{i1},b_{i1}\right)+\mathrm{sig}\left(x_{\omega i}/r_i;a_{i2}{b}_{i2}\right)\end{array}\\ {\tau }^B=-\Xi +J_B{\dot{\omega }}_{\mathrm{ecf}}-k_d{z}_{\omega }-Q_e^T{R}^T\epsilon \end{array}& \left(20\right)\end{array}$

• • wherein, x_ωi is a state of the tracking differentiator, k_d is the control parameter and Ξ is the auxiliary variable.

The invention also provides an aircraft attitude tracking control system, comprising:

• • The data acquisition module is used to obtain the operating parameters of the aircraft;
  • The model building block is used to establish the kinematic model and dynamic model of the aircraft attitude tracking error based on the aircraft operating parameters and quaternion description methods;
  • The constraint boundary setting module is used to construct the performance constraint function of the specified time convergence based on the hyperbolic cosecant composite function, and to set the constraint boundary of the attitude tracking error based on the performance constraint function;
  • The first dynamic equation is used to construct a constrained attitude error conversion function based on the attitude tracking error constraint boundary, and to solve the inverse function of the constrained attitude error conversion function to obtain a dynamic equation of the unconstrained attitude tracking conversion error;
  • The second dynamic equation is used to construct a model based on the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error.

The derivative solving module is used for solving the virtual control instructions in the kinematic model based on the dynamic equation and the backstepping method of the unconstrained transformation error, and constructing a Sigmoid tracking differentiator to obtain the numerical derivative of the virtual control instructions;

The tracking module is used for designing the control moment based on the virtual control instruction and the numerical derivative, using the backstepping method to solve the actual control input quantity in the dynamic model, and using the control torque and the actual control input quantity for the aircraft attitude tracking system to realize the aircraft attitude tracking control.

The invention also provides a computer device, comprising a memory, a processor and a computer program stored on the memory, wherein the processor executes the computer program to realize any of the steps described in the aircraft attitude tracking control method.

The aircraft attitude tracking control method provided by the present invention has the following beneficial effects:

• • The attitude tracking algorithm in the present invention is designed based on the preset performance method, and the performance constraints are applied in advance in the controller design process, so as to theoretically ensure the constraint and adjustability of the transient and steady-state performance of the closed-loop system; The performance constraint function based on the composite hyperbolic cosecant is constructed, and the prior constraints of overshoot and steady-state error, the pre-specified convergence time and the fine adjustment of the convergence speed are realized at the same time, so as to ensure the strict controllability and smooth reliability of the whole tracking process. The Sigmoid tracking differentiator is used to derive the virtual control instructions in the backstep design, so as to realize the computational lightweight of the control algorithm and ensure its practical application performance in the airborne flight control system.

BRIEF DESCRIPTION OF THE DRAWINGS

In order to more clearly illustrate the embodiment of the present invention and its design scheme, the accompanying drawings required for the present embodiment will be briefly described below. The drawings described below are only partial embodiments of the present invention, and for those of ordinary skill in the art, other drawings may be obtained from these drawings without creative effort.

FIG. 1 is the flow chart of the aircraft attitude tracking control method of embodiment 1 of the present invention.

FIG. 2 is a schematic diagram of three coordinate systems used in the attitude tracking motion modeling in the present invention.

FIG. 3 is the closed-loop structure diagram of the lightweight specified time preset performance attitude tracking control system in the present invention.

DETAILED DESCRIPTION OF EMBODIMENTS

In order to enable those skilled in the art to better understand the technical scheme of the present invention and to be able to implement it, the present invention is described in detail below in conjunction with the accompanying drawings and specific embodiments. The following embodiments are only used to more clearly illustrate the technical scheme of the present invention and cannot be used to limit the scope of protection of the present invention.

Embodiment 1

The present invention provides an aircraft attitude tracking control method, as shown in FIG. 1 to FIG. 3, and comprises the following steps:

Step 1: Firstly, the aircraft operating parameters are obtained, and the kinematic model and dynamic model of the aircraft attitude tracking error based on the aircraft operating parameters and quaternion description are established.

In order to establish the aircraft attitude motion model (hereinafter referred to as the kinematic model), three coordinates are defined as shown in FIG. 2, in which the earth's inertial frame is denoted as F₁, the aircraft system is denoted as, F_B and the expected motion coordinate system is denoted as F_D. The motion of the aircraft unit is expressed as F_B. F₁ represents relative motion, and the kinematic model and dynamic model of the attitude of the vehicle based on the quaternion description are as follows

$\begin{array}{cc}{\dot{q}}_B=\frac{1}{2}{q}_B{\omega }_B^B& \left(1\right)\end{array}$ $\begin{array}{cc}{J}_B{\dot{\omega }}_B^B=-{\omega }_B^B\times J_B{\omega }_B^B+{\tau }^B& \left(2\right)\end{array}$

• • wherein, q_B and ω_B^B are the attitude quaternion and angular velocity of the aircraft, {dot over (q)}_B and ω_B^B are the derivatives of the attitude quaternion and angular velocity of the aircraft, respectively, J_B is the moment of inertia of the aircraft and τ^B is the control input torque.

Desired motion is described as F_D relative to F₁ motion that satisfies:

$\begin{array}{cc}{\dot{q}}_D=\frac{1}{2}{q}_D{\omega }_D^D& \left(3\right)\end{array}$

• • where q_D and ω_D^D are the desired attitude and the desired angular velocity, respectively;

The tracking motion of the aircraft is described as the motion of F_B relative to F_D. Based on formulas (1)-(3), the kinematic model and dynamic model of the aircraft attitude tracking error are established as follows:

$\begin{array}{cc}{\dot{q}}_e=\frac{1}{2}{q}_e{\omega }_e^B& \left(4\right)\end{array}$ $\begin{array}{cc}{J}_B{\dot{\omega }}_e^B=-{\omega }_B^B\times J_B{\omega }_B^B+J_B\left(C_e{\dot{\omega }}_D^D-{\omega }_e^B\times C_e{\omega }_D^D\right)+{\tau }^B& \left(5\right)\end{array}$

• • wherein, q_e=q_D*q_B is the attitude tracking error quaternion, q_e0 is the scalar part of the quaternion, q_e=[q_e1, q_e2, q_e3]^T is the vector part of the quaternion, ω_e^B=ω_B^B−C_eω_D^D is the angular velocity tracking error, C_e is the rotation matrix from F_D to F_B, ω_D^D is the desired angular velocity, {dot over (ω)}_D^D is the is the desired attitude angular acceleration, J_B is the moment of inertia of the aircraft, ω_B^B is the angular velocity of the aircraft, and τ^B is the control torque.

Step 2: Construct a specified time performance constraint function based on a hyperbolic cosecant composite function and a posture tracking error constraint boundary.

In order to achieve the goal of fine-tuning the convergence speed and pre-specifying the convergence time, a performance constraint function based on a piecewise composite hyperbolic cosecant composite function is constructed:

$\begin{array}{cc}{\rho }_i\left(t\right)=\{\begin{array}{cc}\mathrm{csch}\left({\rho }_{i0}+\frac{r_{i}t}{T_i-t}\right)+{\rho }_{i\infty },& 0\le t<T_i\\ {\rho }_{i\infty }& t<T_i\end{array},i=1,2,3& \left(6\right)\end{array}$

• • wherein, t is the time variable, csch(⋅) is the hyperbolic cosecant composite function, T_i is the specified convergence time, ρ_i0 is used to adjust the initial value of the performance constraint function; with the help of the properties of the hyperbolic cosecant composite function,

$\underset{{\rho }_{i0\to 0}}{\lim }csch\left(0\right)=+\infty$

we can design a sufficiently small ρ_i0 to make the initial value of the performance constraint function sufficiently large, thereby ensuring that the initial conditions are always met and eliminating the dependence on the initial state of the system.

Based on the performance constraint function, the attitude tracking error constraint boundary is designed as:

$\begin{array}{cc}-\underset{\_}{{\delta }_i}{\rho }_i\left(t\right)<q_{ei}<\overline{{\delta }_i}{\rho }_i\left(t\right),i=1,2,3.& \left(7\right)\end{array}$

• • where δ_i and δ_i are the overshoot limit parameters, and δ_i, δ_i∈(0,1], which is used to adjust the maximum boundary of the attitude tracking error, is the overshoot of the adjustment convergence process, and q_ei is the corresponding component in the quaternion of the attitude tracking error.

Step 3: Construct the constrained attitude error conversion function and its inverse function to obtain the dynamic equation of the unconstrained attitude tracking conversion error.

Based on the boundary constraint of the attitude tracking error constraint set in the second step, a constrained attitude error conversion function is constructed to convert the constrained attitude tracking error into an unconstrained form, and the constrained attitude error conversion function S(⋅) is designed as follows

$\begin{array}{cc}S\left({\epsilon }_i\right)=\frac{\overline{{\delta }_i}{e}^{{\epsilon }_i}-\underset{\_}{{\delta }_i}{e}^{-{\epsilon }_i}}{e^{{\epsilon }_i}+e^{-{\epsilon }_i}},i=1,2,3& \left(8\right)\end{array}$

• • wherein ε_i is the unconstrained transformation error, and the constrained actual attitude error q_ei and the unconstrained transformation error ε_i satisfy:

$\begin{array}{cc}{q}_{ei}={\rho }_{i}S\left({\epsilon }_i\right),i=1,2,3& \left(9\right)\end{array}$

Based on equation (9), the inverse function of the constrained attitude error conversion function S⁻¹(⋅) is solved to give the explicit expression of the transformation error:

$\begin{array}{cc}{\epsilon }_i=S^{-1}\left(z_i\right)=\frac{1}{2}\ln \left(\frac{z_i+\underset{\_}{{\delta }_i}}{\stackrel{\_}{{\delta }_i}-z_i}\right),z_j=q_{ei}\left(t\right)/{\rho }_i\left(t\right),i=1,2,3& 10\end{array}$

• • where ε_i is the conversion error, and z_i is the normalized conversion error.

Combined with the kinematic model of the attitude tracking error of the aircraft established in the first step, the dynamic equation of the unconstrained conversion error is established:

$\begin{array}{cc}\dot{\epsilon }=R{Q}_e{\omega }_e+v& \left(11\right)\end{array}$

• • where ε=[ε₁, ε₂, ε₃]^T is the transformation error vector, E represents the derivative of ε; R=diag[r₁, r₂, r₃], is a third-order diagonal matrix;

$r_i=\frac{\partial s^{-1}}{\partial z_i}·\frac{1}{{\rho }_i},$ $i=1,2,3,Q_e=\frac{q_{eo}{I}_3+q_e^{\times }}{2},v={\left[v_1,v_2,v_3\right]}^T,v_i=-r_i\frac{q_{ei}{\stackrel{.}{\rho }}_i}{{\rho }_i},i=1,2,3;$

S⁻¹ is the inverse of the error conversion function, r_i is an element in the diagonal matrix, Q_e is the attitude multiplication matrix, q_e0 is the scalar part of the quaternion, I₃ is the third-order identity matrix, q_e^× is the attitude cross product matrix, ν is the auxiliary vector, and ν_i is an element in the auxiliary vector.

Based on the dynamic equation of the unconstrained conversion error given in the third step, the virtual control instructions in the kinematic model are designed as follows:

$\begin{array}{cc}{\omega }_{ec}=-Q_e^{-1}{R}^{-1}\left(k_p\epsilon +v\right)& \left(12\right)\end{array}$

• • wherein k_p>0 is a control parameter.

In order to obtain the derivatives of virtual control instructions and reduce the burden of online computation, a tracking differentiator based on Sigmoid function is constructed:

$\begin{array}{cc}\{\begin{array}{c}{\dot{x}}_1=x_2\\ \begin{array}{c}{\dot{x}}_2=-r^{2}sig(x_1-w;a_1,\\ b_1)-(\mathrm{sig}\left(x_2/r;a_2{b}_2\right)\end{array}\end{array},\mathrm{sig}\left(x;a,b\right)=a\left[{\left(1+e^{-bx}\right)}^{-1}-0.5\right].& \left(13\right)\end{array}$

• • where sig(⋅) is the sigmoid function, a₁, b₁, a₂, b₂>0 is the design parameter, r>0 is the acceleration factor, w is the reference signal input, and x₁ and x₂ are the two states of the tracking differentiator. In order to ensure the best performance of the tracking differentiator, the following parameter adjustment rules are followed: first, gradually increase r to obtain the desired tracking accuracy; secondly, gradually increase a₁ and b₁ to ensure stability and speed with no overshoot; finally, gradually reduce a₂ and b₂ to obtain optimal dynamic performance. Under the action of this tracking differentiator, there are:

$\underset{t\to \infty }{\lim }{x}_1=w,\underset{t\to \infty }{\lim }{x}_2=\dot{w},$

i.e., x₁ asymptotic convergence to the reference signal and w, x₂ asymptotically converge to the reference signal derivative {dot over (w)}.

The sigmoid tracking differentiator is used to obtain the derivative signal of the virtual control instruction in the following form:

$\begin{array}{cc}\{\begin{array}{c}{\stackrel{.}{\omega }}_{\mathrm{ecfi}}=x_{\omega i}\\ {\stackrel{.}{x}}_{\omega i}=r_i^2[\mathrm{sig}\left({\omega }_{\mathrm{ecfi}}-{\omega }_{\mathrm{eci}};a_{i1},b_{i1}\right)+\mathrm{sig}\left(x_{\omega i}/r_i;a_{i2}{b}_{i2}\right)\end{array},i=1,2,3& \left(14\right)\end{array}$

• • wherein r_i, a_i1, a_i2, b_i1, b_i2>0, ω_ecf=[ω_ec1, ω_ec2, ω_ec3]^T and {dot over (ω)}_ecf=[x_ω1, x_ω2, x_ω3]^T are the filtered signal and the filtered derivative of the virtual control command, respectively, and x_ωi is a state of the tracking differentiator.

Where s=ω_ecf−ω_ec is denoted as the filter estimation error and z_ω=ω_e−ω_ecf is the angular velocity tracking error, then the dynamic equation that constrains the conversion error can be written as:

$\begin{array}{cc}\dot{\epsilon }=-k_p\epsilon +R{Q}_{e}s+R{Q}_e{z}_{\omega }& \left(15\right)\end{array}$

Step 5: Solve the control moment in the dynamic model and give the complete form of the attitude tracking algorithm for use in the flight control system.

Based on the virtual control command and its filtering information given in the fourth step, the angular velocity tracking error is defined as z_ω=ω_e−ω_ecf, and its dynamic equation is as follows:

$\begin{array}{cc}{\dot{z}}_w=J_B^{-1}(-{\omega }_B^B\times J_B{\omega }_B^B-J_B\left(C_e{\dot{\omega }}_D^D-{\omega }_e^B\times C_e{\omega }_D^D\right)+{\tau }^B& \left(16\right)\end{array}$

According to the above derivation and the principle of backstep, the control torque input τ^B is designed as:

$\begin{array}{cc}{\tau }^B=-\Xi +J_B{\dot{\omega }}_{\mathrm{ecf}}-k_d{z}_{\omega }-Q_e^T{R}^T\epsilon & \left(17\right)\end{array}$

• • wherein k_d>0 is the control parameter, and Ξ=−ω_B^B×J_Bω_B^B−J_B(C_e{dot over (ω)}_D^D−ω_e^B×C_eω_D^D) is the auxiliary variable.

Substituting the control torque input τ^B into the dynamics model yields the closed-loop dynamics equation as follows:

$\begin{array}{cc}{\dot{z}}_{\omega }=-k_d{z}_{\omega }-Q_e^T{R}^T\epsilon & \left(18\right)\end{array}$

The candidate Lyapunov equation for defining the overall closed-loop system is:

$\begin{array}{cc}V=\frac{1}{2}{\epsilon }^T\epsilon +\frac{1}{2}{z}_{\omega }^T{z}_{\omega }& \left(19\right)\end{array}$

The derivative V along the trajectory of the closed-loop system can be obtained {dot over (V)}=k_pε²−k_dz_ω²+ε^TRQ_es. According to the nature of the sigmoid tracking differentiator,

$\underset{t\to \infty }{\lim }s=0,$

so ε and z_ω are ultimately bounded. The attitude tracking error of the aircraft converges at a specified time and satisfies the requirements of the performance constraint function according to the backstepping method and the basic principle of preset performance control.

Finally, the complete form of the proposed lightweight aircraft attitude tracking controller with preset performance at specified time is given:

$\begin{array}{cc}\{\begin{array}{c}{\omega }_{ec}=-2Q_e^{-1}{R}^{-1}\left(k_p\epsilon +v\right)\\ \{\begin{array}{c}{\dot{\omega }}_{\mathrm{ecfi}}=x_{\mathrm{\omega i}},i=1,2,3\\ {\stackrel{.}{x}}_{\omega i}=r_i^2[sig\left({\omega }_{\mathrm{ecfi}}-{\omega }_{eci};a_{i1},b_{i1}\right)+sig\left(x_{\omega i}/r_i;a_{i2}{b}_{i2}\right)\end{array}\\ {\tau }^B=-\Xi +J_B{\dot{\omega }}_{\mathrm{ecf}}-k_d{z}_{\omega }-Q_e^T{R}^T\epsilon \end{array}& \left(20\right)\end{array}$

• • wherein, x_ωi is a state of the tracking differentiator, k_d is the control parameter and Ξ is the auxiliary variable.

The aircraft attitude tracking control method in the present invention consists of a virtual control law (7), tracking differentiator (10), and the actual control torque (12). It is used in the aircraft motion control system to achieve high-performance attitude tracking control under the limitation of computing resources.

Based on the same invention concept, the invention also provides an aircraft attitude tracking control system, which comprises a data acquisition module, a model building module, a constraint boundary setting module, a first dynamic equation construction model, a second dynamic equation construction model, a derivative solving module and a tracking module.

Specifically, the data acquisition module is used to obtain the operating parameters of the aircraft. The model building module is used to establish the kinematic model and dynamic model of the aircraft attitude tracking error based on the aircraft operating parameters and quaternion description. The constraint boundary setting module is used to construct the performance constraint function of the specified time convergence based on the hyperbolic cosecant composite function, and set the constraint boundary of the attitude tracking error based on the performance constraint function. the first dynamic equation construction model is used to construct a constrained attitude error conversion function based on the constraint boundary of the attitude tracking error, and to solve the inverse function of the constrained attitude error conversion function to obtain a dynamic equation of the unconstrained attitude tracking conversion error. The second dynamic equation construction model is used to establish the dynamic equation of unconstrained conversion error according to the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error. the derivative solving module is used for solving the virtual control instructions in the kinematic model based on the dynamic equation and the backstepping method of the unconstrained transformation error, and constructs a Sigmoid tracking differentiator to obtain numerical derivatives for the virtual control instructions. The tracking module is used to design the control moment based on the virtual control instruction and the numerical derivative, uses the backstepping method to solve the actual control input quantity in the dynamic model, and applies the control torque and the actual control input quantity to the aircraft attitude tracking system to realize the aircraft attitude tracking control.

Each module in the aircraft attitude tracking control system can be realized in whole or in part through software, hardware and a combination thereof. Each module can be embedded in or independent of the processor in the computer equipment in the form of hardware, or can be stored in the memory in the computer device in the form of software, so that the processor can be called to perform the corresponding operations of each of the above modules.

At the same time, the invention also provides a computer device, comprising a memory, a processor and a computer program stored on the memory, and the processor executes the steps in the embodiment of the aircraft attitude tracking control method for implementing the computer program. The specific implementation method can be found in the method embodiment, which will not be repeated here.

Those skilled in the art should understand that embodiments of the present invention may be provided as a process, system, or computer program product. Therefore, the present invention may take the form of a complete hardware embodiment, a complete software embodiment, or a combination of software and hardware embodiments. Moreover, the invention may take the form of a computer program product implemented on one or more computer-available storage media (including, but not limited to, disk memory, CD-ROM, optical memory, etc.) containing computer-available program code.

The present invention is described with reference to a flow diagram and/or block diagram of a method, device (system) and computer program product according to an embodiment of the present invention. It should be understood that each process and/or box in the flowchart and/or block diagram, as well as the combination of the process and/or box in the flowchart and/or block diagram, can be implemented by computer program instructions. These computer program instructions may be provided to the processor of a general-purpose computer, a specialized computer, an embedded processing machine, or other programmable data processing device to produce a machine such that the instructions executed by the processor of the computer or other programmable data processing device produce a device for implementing the functions specified in one or more processes and/or block diagrams, one or more boxes.

These computer program instructions may also be stored in computer-readable memory capable of directing a computer or other programmable data-processing device to work in a particular manner such that the instructions stored in the computer-readable memory result in a manufactured product comprising a directive device that implements the functions specified in a flowchart process or processes and/or block diagram boxes or boxes.

These computer program instructions may also be loaded onto a computer or other programmable data processing device such that a series of operational steps are performed on the computer or other programmable device to produce computer-implemented processing, so that the instructions executed on the computer or other programmable device provide steps for implementing the function specified in a flowchart process or processes and/or block diagram boxes or boxes.

It should be noted that the specific embodiments described above may enable those skilled in the art to understand the invention more comprehensively, but do not limit the invention in any way. Therefore, although the present specification and embodiments have described the invention in detail, those skilled in the art should understand that the present invention may still be modified or equivalent replaced; And all technical solutions and improvements that do not deviate from the spirit and scope of the present invention are covered by the scope of protection of the patent for the present invention. Any drawing mark in the claims shall not be deemed to be a limitation of the claims to which it relates.

Claims

1. A method for aircraft attitude tracking control, characterized by comprising the following steps: acquire aircraft operating parameters;establish a kinematic model and a dynamic model of aircraft attitude tracking error based on aircraft operating parameters and quaternion description;construct a performance constraint function that converges at a specified time based on a hyperbolic cosecant composite function, and set an attitude tracking error constraint boundary based on the performance constraint function;construct a constrained attitude error conversion function based on the attitude tracking error constraint boundary, and solve the inverse function of the constrained attitude error conversion function to obtain a dynamic equation of an unconstrained attitude tracking conversion error;according to the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error, establish a dynamic equation of the unconstrained conversion error;solve the virtual control instructions in the kinematic model based on the dynamic equation of the unconstrained conversion error and the backstepping method, and construct a sigmoid tracking differentiator to obtain the numerical derivative of the virtual control instructions;design a control torque based on the virtual control instructions and the numerical derivative, use the backstepping method to solve the actual control input in the dynamic model, and use the control torque and the actual control input in the aircraft attitude tracking system to realize aircraft attitude tracking control.

2. The aircraft attitude tracking control method of claim 1, wherein the specific construction method of the kinematic model and the dynamic model of the aircraft attitude tracking error is as follows: taking F1 as the earth's inertial system, FB as the aircraft system, and FD as the expected motion coordinate system, the motion of the aircraft unit is expressed as the motion of FB relative to F1 and the single body kinematic model and dynamic model based on quaternion description are, respectively:

3. The aircraft attitude tracking control method of claim 2, wherein the performance constraint function of the specified time convergence is constructed based on the hyperbolic cosecant composite function, and the attitude tracking error constraint boundary is set based on the performance constraint function, specifically: based on the hyperbolic cosecant composite function, a performance constraint function for the convergence at the specified time is constructed ρi(t):

4. The aircraft attitude tracking control method of claim 3, wherein the constrained attitude error conversion function is constructed based on the attitude tracking error constraint boundary, and the inverse function of the constrained attitude error conversion function is solved to obtain a dynamic equation of the unconstrained attitude tracking conversion error, specifically: based on the constraint boundary of the attitude tracking error, the constrained attitude error conversion function is constructed, which is used to convert the constrained attitude tracking error into an unconstrained form, and the constrained attitude error conversion function S(⋅) is designed as follows

5. The aircraft attitude tracking control method of claim 4, wherein the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error are established, specifically:

6. The aircraft attitude tracking control method of claim 5, wherein the virtual control instruction of the kinematic loop is solved based on the dynamic equation and the backstepping method of the unconstrained transformation error, and the sigmoid tracking differential is constructed to obtain the numerical derivative of the virtual control instruction, specifically: based on the dynamic equation of the unconstrained transformation error, the virtual control instructions in the kinematic model are designed as:

7. The aircraft attitude tracking control method of claim 1, wherein the control moment is designed based on the virtual control instruction and the numerical derivative, the actual control input quantity is solved by using the backstepping method in the dynamic model, and the control moment and the actual control input quantity are used in the aircraft attitude tracking system to realize the aircraft attitude tracking control, specifically: based on the virtual control instruction and its filtering signal, the angular velocity loop tracking error is defined as zω=ωe−ωecf, and its dynamic equation is:

8. The aircraft attitude tracking control method of claim 1, wherein the aircraft attitude tracking controller is:

9. An aircraft attitude tracking control system, comprising: a data acquisition module configured to obtain the operating parameters of the aircraft;wherein, a model building block is used to establish the kinematic model and dynamic model of the aircraft attitude tracking error based on the aircraft operating parameters and quaternion description methods;a constraint boundary setting module is used to construct a performance constraint function of a specified time convergence based on a hyperbolic cosecant composite function, and to set the constraint boundary of the attitude tracking error based on the performance constraint function;a first dynamic equation is used to construct a constrained attitude error conversion function based on the attitude tracking error constraint boundary, and to solve the inverse function of the constrained attitude error conversion function to obtain a dynamic equation of the unconstrained attitude tracking conversion error;the second dynamic equation is used to construct a model based on the dynamic equation of the unconstrained attitude tracking conversion error and the kinematic model of the aircraft attitude tracking error;wherein, the derivative solving module is used for solving the virtual control instructions in the kinematic model based on the dynamic equation and the backstepping method of the unconstrained transformation error, and constructing a Sigmoid tracking differentiator to obtain the numerical derivative of the virtual control instructions;the tracking module is used for designing the control moment based on the virtual control instruction and the numerical derivative, using the backstepping method to solve the actual control input quantity in the dynamic model, and using the control torque and the actual control input quantity for the aircraft attitude tracking system to realize the aircraft attitude tracking control.

10. A computer device comprising a non-transient memory, a processor and a computer program stored on the memory, wherein the processor executes the steps of a computer program configured to realize the method of claim 1.