CN107943097B - Aircraft control method and device and aircraft - Google Patents

Abstract

The present invention provides an aircraft control method, device and aircraft; wherein, the method includes: collecting initial flight state parameters output by a power system of the aircraft; generating an initial control signal according to the initial flight state parameters and a preset reference signal; Input the initial control signal to the power system of the aircraft, collect the current flight state parameters output by the power system, calculate the error signal between the current flight state parameter and the reference signal; generate the compensation control signal according to the error signal; according to the initial control signal and compensation The control signal determines the throttle valve and rudder deflection angle control signal of the aircraft to control the flight state of the aircraft. The present invention can restrain the influence of various uncertain factors generated in the flight process of the aircraft, and improve the tracking control performance of the controller to the flight state of the aircraft; at the same time, the present invention adopts a linear time-invariant control method, which is easy to implement, and Strong practicality.

Description

Aircraft control method, device and aircraft

technical field

The present invention relates to the technical field of aircraft, in particular to a control method, device and aircraft of an aircraft.

Background technique

High-speed aircraft are widely used in many fields because of their fast, efficient and reliable access to adjacent space. During flight, high-speed aircraft will be affected by nonlinearity and uncertainties including parameter uncertainty, nonlinear coupling, unstructured and external disturbances, especially when performing supersonic flight missions, these uncertainties will be serious Affects the tracking performance of the closed-loop control system in the aircraft. However, in the existing control methods, the influence of uncertainty is often ignored, or these influences are roughly estimated, resulting in a poor ability to suppress the uncertain factors of the aircraft, which in turn leads to poor tracking control performance.

SUMMARY OF THE INVENTION

In view of this, the purpose of the present invention is to provide an aircraft control method, device and aircraft, so as to suppress the influence of various uncertain factors during the flight of the aircraft and improve the tracking control performance of the controller on the flight state of the aircraft.

In a first aspect, an embodiment of the present invention provides a method for controlling an aircraft, and the method is applied to a controller of the aircraft; the method includes: collecting initial flight state parameters output by a power system of the aircraft; wherein the initial flight state parameters include flight speed and Flight altitude; generate the initial control signal according to the initial flight state parameters and the preset reference signal; input the initial control signal to the power system of the aircraft, collect the current flight state parameters output by the power system, and calculate the difference between the current flight state parameters and the reference signal. According to the error signal, the compensation control signal is generated; according to the initial control signal and the compensation control signal, the throttle valve and the rudder deflection angle control signal of the aircraft are determined to control the flight state of the aircraft.

In conjunction with the first aspect, an embodiment of the present invention provides a first possible implementation manner of the first aspect, wherein the above-mentioned step of generating an initial control signal according to the initial flight state parameters and a preset reference signal includes: Formula to calculate the initial control signal u _i,H :

u _i,H =K _i,H e _i ,i=1,2,;

Among them, i=1 represents the flight speed; i=2 represents the flight height; K _i,H suboptimal state feedback gain; e _i is the error signal.

In conjunction with the first possible implementation manner of the first aspect, the embodiment of the present invention provides a second possible implementation manner of the first aspect, wherein the above-mentioned suboptimal state feedback gain K _i,H is obtained by the following formula:

ω_n为自然角频率；

ρ为大气密度；v_trim为飞行器在巡航飞行阶段平衡状态时的速度；S为参考面积；

为平均气动弦长；C_Me为空气动力系数；I_yy为转动惯量；P_i为下述方程的对称正定解：in,

ω _n is the natural angular frequency;

ρ is the atmospheric density; v _trim is the speed of the aircraft in the equilibrium state during the cruise flight phase; S is the reference area;

is the average aerodynamic chord length; C _Me is the aerodynamic coefficient; I _yy is the moment of inertia; P _i is the symmetrical positive definite solution of the following equation:

a₂₁＝v_trim；a₂₂＝T_trim/m^N/V_trim；C_Tβ0和C_Tβ2为空气动力系数；m为飞机质量；ζ_n为阻尼比；T_trim为飞行器在巡航飞行阶段平衡状态时的推力；

和θ_i是权重参数；Q_i(i＝1,2)是对称正定矩阵，φ_i(i＝1,2)是设定的衰减因子；上标N表示参数为标称参数。

a ₂₁ =v _trim ; a ₂₂ =T _trim /m ^N /V _trim ; C _Tβ0 and C _Tβ2 are aerodynamic coefficients; m is the aircraft mass; _{ζ n} is the damping ratio; thrust;

and θ _i are weight parameters; Q _i (i=1,2) is a symmetric positive definite matrix, φ _i (i=1,2) is a set attenuation factor; the superscript N indicates that the parameters are nominal parameters.

With reference to the second possible implementation manner of the first aspect, the embodiment of the present invention provides a third possible implementation manner of the first aspect, wherein the above-mentioned step of generating a compensation control signal according to an error signal includes: by the following The formula to calculate the compensation control signal u _i,R (s):

f₁和f₂是设定的滤波参数，为正常数；y₂＝h-r_h，y₁＝v-r_v；v为飞行器的速度；h为飞行器的高度；r_v为飞行器在巡航飞行阶段的参考速度；r_h为飞行器在巡航飞行阶段的参考高度；G_i(s)为从u_i,R(s)至y_i的传递函数。Among them, s is the Laplacian operator; F ₁ (s)=f ₁ ³ /(s+f ₁ ) ³ ,

f ₁ and f ₂ are the set filter parameters, which are normal numbers; y ₂ =hr _h , y ₁ =vr _v ; v is the speed of the aircraft; h is the height of the aircraft; r _v is the reference of the aircraft in the cruise flight stage speed; _rh is the reference altitude of the aircraft in the cruise flight phase; G _i (s) is the transfer function from ui _,R (s) to _yi .

In conjunction with the third possible implementation manner of the first aspect, the embodiment of the present invention provides the fourth possible implementation manner of the first aspect, wherein the above-mentioned transfer function G _i (s) is obtained by the following formula:

G _i (s)=C _i (sI _i -A _i,H ) ^-1 B _i (i=1,2);

A_i,H＝A_i+B_iK_i,H。Among them, I _i (i=1,2) is the identity matrix;

A _i,H =A _i +B _i K _i,H .

In conjunction with the fourth possible implementation manner of the first aspect, the embodiment of the present invention provides the fifth possible implementation manner of the first aspect, wherein the above-mentioned determination of the throttle valve and the The control signal of the rudder deflection angle is used to control the flight state of the aircraft, including: calculating the final control signal of the aircraft u _i =ui _,H +u _i,R , (i=1,2); according to the final control signal Control the speed and/or altitude of the aircraft.

In a second aspect, an embodiment of the present invention provides a control device for an aircraft, the device is arranged on a controller of the aircraft; the device includes: a signal acquisition module for collecting initial flight state parameters output by a power system of the aircraft; wherein, the initial flight The state parameters include flight speed and flight height; the first signal generation module is used to generate the initial control signal according to the initial flight state parameters and the preset reference signal; the error signal calculation module is used to input the initial control signal to the power of the aircraft The system collects the current flight state parameters output by the power system, and calculates the error signal between the current flight state parameter and the reference signal; the second signal generation module is used to generate a compensation control signal according to the error signal; the control module is used to generate a compensation control signal according to the initial The control signal and the compensation control signal are used to determine the throttle valve and rudder deflection angle control signal of the aircraft to control the flight state of the aircraft.

In conjunction with the second aspect, the embodiment of the present invention provides a first possible implementation manner of the second aspect, wherein the above-mentioned first signal generation module is further configured to: calculate the initial control signal ui _,H by the following formula: u _i,H =K _i,H e _i ,i=1,2,wherein, i=1 represents the flight speed; i=2 represents the flight height; Ki _,H is the suboptimal state feedback gain; e _i is the error Signal.

With reference to the first possible implementation manner of the second aspect, the embodiment of the present invention provides a second possible implementation manner of the second aspect, wherein the above-mentioned first signal generation module is further configured to: a suboptimal state feedback gain K _i,H , obtained by the following formula:

ω_n为自然角频率；

ρ为大气密度；v_trim为飞行器在巡航飞行阶段平衡状态时的速度；S为参考面积；

为平均气动弦长；C_Me为空气动力系数；I_yy为转动惯量；P_i为下述方程的对称正定解：in,

ω _n is the natural angular frequency;

ρ is the atmospheric density; v _trim is the speed of the aircraft in the equilibrium state during the cruise flight phase; S is the reference area;

is the average aerodynamic chord length; C _Me is the aerodynamic coefficient; I _yy is the moment of inertia; P _i is the symmetrical positive definite solution of the following equation:

a₂₁＝v_trim；a₂₂＝T_trim/m^N/V_trim；C_Tβ0和C_Tβ2为空气动力系数；m为飞机质量；ζ_n为阻尼比；T_trim为飞行器在巡航飞行阶段平衡状态时的推力；

和θ_i是权重参数；Q_i(i＝1,2)是对称正定矩阵，φ_i(i＝1,2)是设定的衰减因子；上标N表示参数为标称参数。

a ₂₁ =v _trim ; a ₂₂ =T _trim /m ^N /V _trim ; C _Tβ0 and C _Tβ2 are aerodynamic coefficients; m is the aircraft mass; _{ζ n} is the damping ratio; thrust;

and θ _i are weight parameters; Q _i (i=1,2) is a symmetric positive definite matrix, φ _i (i=1,2) is a set attenuation factor; the superscript N indicates that the parameters are nominal parameters.

In a third aspect, an embodiment of the present invention provides an aircraft, the aircraft includes a processor and a sensor; the control device of the aircraft is provided in the processor.

The embodiments of the present invention have brought the following beneficial effects:

According to an aircraft control method, device and aircraft provided by the embodiments of the present invention, an initial control signal can be generated according to the initial flight state parameters output by the aircraft and a preset reference signal; after the initial control signal is input to the power system of the aircraft , collect the current flight state parameters output by the power system, and calculate the error signal between the current flight state parameter and the reference signal; then according to the error signal, the compensation control signal can be generated; then according to the initial control signal and the compensation control signal, the Determine the throttle valve and rudder deflection angle control signal of the aircraft, and then control the flight state of the aircraft; this method can suppress the influence of various uncertain factors generated by the aircraft during the flight process by compensating the control signal, and improve the controller's ability to The tracking control performance of the aircraft flight state.

Further, this method adopts a linear time-invariant control method, which is easy to implement and has strong practicability.

Additional features and advantages of the present invention will be set forth in the description which follows, or some may be inferred or unambiguously determined from the description, or may be learned by practicing the above-described techniques of the present invention.

In order to make the above-mentioned objects, features and advantages of the present invention more clearly understood, the preferred embodiments are exemplified below, and are described in detail as follows in conjunction with the accompanying drawings.

Description of drawings

In order to illustrate the specific embodiments of the present invention or the technical solutions in the prior art more clearly, the following briefly introduces the accompanying drawings that need to be used in the description of the specific embodiments or the prior art. Obviously, the accompanying drawings in the following description The drawings are some embodiments of the present invention. For those of ordinary skill in the art, other drawings can also be obtained based on these drawings without creative efforts.

1 is a flowchart of a method for controlling an aircraft according to an embodiment of the present invention;

2 is a schematic structural diagram of a flight controller according to an embodiment of the present invention;

3 is a schematic diagram showing the comparison of singular values of velocity channels of various controllers provided in an embodiment of the present invention;

FIG. 4 is a schematic diagram showing the comparison of singular values of height channels of various controllers provided in an embodiment of the present invention;

5 is a tracking response diagram of a speed channel and an altitude channel of an H _∞ controller to a reference signal provided by an embodiment of the present invention;

6 is a tracking response diagram of a speed channel and an altitude channel of a linear state feedback controller to a reference signal according to an embodiment of the present invention;

7 is a response diagram of the track angle, attack angle and roll angular velocity of the H _∞ controller and the linear state feedback controller provided by the embodiment of the present invention;

FIG. 8 is a schematic diagram showing the change of the input value of the aircraft power system by the H _∞ controller and the linear state feedback controller in the numerical simulation process provided by the embodiment of the present invention;

9 is a comparison diagram of the tracking errors of the H _∞ controller and the linear state feedback controller on the velocity channel and the altitude channel provided by an embodiment of the present invention;

FIG. 10 is a schematic structural diagram of a control device for an aircraft according to an embodiment of the present invention.

Detailed ways

In order to make the purposes, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are part of the embodiments of the present invention, not all of them. example. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

Considering the problem that the existing aircraft control methods have poor ability to suppress uncertain factors of the aircraft, resulting in poor tracking control performance, the embodiments of the present invention provide a control method, device and aircraft for an aircraft; the technology can be applied to In the flight control process of high-speed aircraft, unmanned aerial vehicles, etc.; the technology can be implemented by using relevant software or hardware, which will be described below through embodiments.

Referring to the flowchart of a control method of an aircraft shown in FIG. 1; the method is applied to the controller of the aircraft; the method includes the following steps:

Step S102, collecting initial flight state parameters output by the power system of the aircraft; wherein, the initial flight state parameters include flight speed and flight height;

Step S104, generating an initial control signal according to the initial flight state parameters and a preset reference signal;

Step S105, the initial control signal is input to the power system of the aircraft, the current flight state parameters output by the power system are collected, and the error signal between the current flight state parameter and the reference signal is calculated; The error signal of the flight height is processed separately, and then the flight speed and flight height of the UAV are tracked respectively.

Specifically, the above-mentioned initial control signal generates the speed channel control law and the altitude channel control law through the controller, and then is input to the aircraft; due to the influence of nonlinearity and uncertainty, there will be an error between the actual state of the aircraft flight and the reference signal; Based on this, the above step S105 calculates the error signal between the current flight state parameter and the reference signal, so as to compensate the initial control signal according to the error signal.

Specifically, the initial control signal ui _,H can be calculated by the following formula:

u _i,H =K _i,H e _i ,i=1,2,;

Among them, i=1 represents the flight speed; i=2 represents the flight height; K _i,H is the suboptimal state feedback gain; e _i is the error signal.

The suboptimal state feedback gain K _i,H can be obtained by the following formula:

ω_n为自然角频率；

ρ为大气密度；v_tr ⁱ _m为飞行器在巡航飞行阶段平衡状态时的速度；S为参考面积；

为平均气动弦长；C_Me为空气动力系数；I_yy为转动惯量；in,

ω _n is the natural angular frequency;

ρ is the atmospheric density; _vtri ^m is the _speed of the aircraft in the equilibrium state during the cruise flight stage; S is the reference area;

is the average aerodynamic chord length; C _Me is the aerodynamic coefficient; I _yy is the moment of inertia;

The above P _i is the symmetric positive definite solution of the following equation:

a₂₁＝v_trim；a₂₂＝T_trim/m^N/V_trim；C_Tβ0和C_Tβ2为空气动力系数；m为飞机质量；ζ_n为阻尼比；T_trim为飞行器在巡航飞行阶段平衡状态时的推力；

和θ_i是权重参数；Q_i(i＝1,2)是对称正定矩阵，φ_i(i＝1,2)是设定的衰减因子；上标N表示参数为标称参数。

a ₂₁ =v _trim ; a ₂₂ =T _trim /m ^N /V _trim ; C _Tβ0 and C _Tβ2 are aerodynamic coefficients; m is the aircraft mass; _{ζ n} is the damping ratio; thrust;

and θ _i are weight parameters; Q _i (i=1,2) is a symmetric positive definite matrix, φ _i (i=1,2) is a set attenuation factor; the superscript N indicates that the parameters are nominal parameters.

Step S106, generating a compensation control signal according to the error signal;

Specifically, the compensation control signal ui _,R (s) can be calculated by the following formula:

f₁和f₂是设定的滤波参数，为正常数；y₁＝v-r_v，y₂＝h-r_h；v为飞行器的速度；h为飞行器的高度；r_v为飞行器在巡航飞行阶段的参考速度；r_h为飞行器在巡航飞行阶段的参考高度；G_i(s)为从u_i,R(s)至y_i的传递函数。Among them, s is the Laplacian operator; F ₁ (s)=f ₁ ³ /(s+f ₁ ) ³ ,

f ₁ and f ₂ are the set filtering parameters, which are normal numbers; y ₁ =vr _v , y ₂ =hr _h ; v is the speed of the aircraft; h is the height of the aircraft; r _v is the reference of the aircraft in the cruise flight stage speed; _rh is the reference altitude of the aircraft in the cruise flight phase; G _i (s) is the transfer function from ui _,R (s) to _yi .

The above transfer function G _i (s) is obtained by the following formula:

G _i (s)=C _i (sI _i -A _i,H ) ^-1 B _i (i=1,2);

A_i,H＝A_i+B_iK_i,H。Among them, I _i (i=1,2) is the identity matrix;

A _i,H =A _i +B _i K _i,H .

Step S108: Determine the throttle valve and the rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal, so as to control the flight state of the aircraft.

This step S108 can be specifically implemented in the following manner:

Step (1), calculating the final control signal _ui =ui _,H +ui _,R of the aircraft, (i=1,2);

Step (2), control the flight speed and/or flight altitude of the aircraft according to the final control signal.

It can be seen from the above that i=1 represents the flight speed; i=2 represents the flight height; therefore, when i=1, the controller can control the flight speed of the aircraft according to u ₁ ; when i=2, it can control the aircraft according to u ₂ Of course, the controller can also generate u ₁ and u ₂ at the same time, so as to control the flight speed and flight height of the aircraft at the same time.

In a method for controlling an aircraft provided by an embodiment of the present invention, an initial control signal can be generated according to an initial flight state parameter output by the aircraft and a preset reference signal; after the initial control signal is input to the power system of the aircraft, the power system is collected Output the current flight state parameters, and calculate the error signal between the current flight state parameters and the reference signal; then according to the error signal, a compensation control signal can be generated; and then according to the initial control signal and the compensation control signal, the control signal of the aircraft can be determined. The control signal of flow valve and rudder deflection angle is used to control the flight state of the aircraft; this method can suppress the influence of various uncertain factors generated by the aircraft during flight by compensating the control signal, and improve the controller's ability to control the flight state of the aircraft. Track control performance.

Further, this method adopts a linear time-invariant control method, which is easy to implement and has strong practicability.

Embodiments of the present invention also provide another control method for an aircraft, which is implemented based on H _∞ theory and a robust compensation method; the method can suppress strong nonlinearity and include strong coupling, unstructured uncertainty and external disturbances, etc. Effects of equivalent disturbances on the designed closed-loop control system and generate robust linear control systems that do not satisfy matching conditions.

Specifically, the method first designs an H _∞ controller based on the H _∞ theory to achieve the expected tracking performance; since the H _∞ controller cannot suppress the influence of nonlinearity and uncertainty on the closed-loop control system in the entire frequency domain, so , and then design a robust compensator to suppress the effects of these equivalent disturbances. The combination of H _∞ controller and robust compensator constitutes a linear time-invariant controller, which can ensure that the velocity and altitude channel tracking errors of high-speed aircraft with nonlinearity and multiple uncertainties eventually converge to the neighborhood of a given pair near the origin. In addition, the method has linear time-invariant robustness, is easy to implement in practical applications, and has strong practicability.

In order to realize this method, the dynamic model of the high-speed aircraft needs to be established first. The longitudinal dynamics of the high-speed aircraft can be described by the following equations:

The model includes five state variables [v, γ, h, α, q], where v represents speed, γ represents flight path angle, h represents altitude, α represents angle of attack, q represents pitch rate; m represents aircraft mass , μ represents the gravitational constant, I _yy represents the moment of inertia, r=h+r _e , where r _e is the radius of the earth; d _i (i=V,γ,q,α,h) is the external time-varying atmospheric disturbance (Note: In practical applications, external disturbances mainly refer to the additional forces and moments introduced by the aircraft dynamics model. Therefore, d _h and d _α in formula (1) have no practical significance, and they are introduced into the stability analysis of the closed-loop control system for use) ; T, L, D and M _q represent thrust, lift, drag and pitching moment, respectively, satisfying the following formulas:

L=ρv ² SC _L /2,

T=ρv ² SC _T /2,

D=ρv ² SC _D /2,

分别代表大气密度，参考面积和平均气动弦长；C_L,C_T,C_D,C_Mα,C_Mδe分别代表推力系数，升力系数，阻力系数，攻角系数，偏航速率系数和俯仰速率系数，取决于攻角α和舵偏角δ_e，且满足以下：where ρ, S,

Represent air density, reference area and average aerodynamic chord length respectively; C _L , C _T , C _D , C _Mα , C _Mδe represent thrust coefficient, lift coefficient, drag coefficient, angle of attack coefficient, yaw rate coefficient and pitch rate coefficient, respectively , depends on the angle of attack α and the rudder deflection angle δ _e , and satisfies the following:

C _L =C _Lα α+Δ _L ,

C _D =C _Dα2 α ² +C _Dα α+C _D0 +Δ _D ,

C _Mα =C _Mα2 α ² +C _Mα α+C _α0 +Δ _Mα ,

C _Mδe =C _Me (δ _e -α)+Δ _δe ,

where β is the throttle valve opening, C _Lα , C _Tβ0 , C _Tβ1 , C _Tβ2 , C _Dα2 , C _Dα , C _D0 , C _Mα2 , C _Mα , C _α0 , C _Me , C _Mq2 , C _Mq , C _q0 is the aerodynamic coefficient, which is an unmodeled uncertainty; since it is difficult to determine the analytical relationship between the thrust isoforce and moment coefficients and the angle of attack or throttle opening, curve fitting techniques are used to describe the design controller the aerodynamic characteristics at the time, and then there is a mismatch _Δj between the real model and the control-oriented model. The high-speed aircraft engine dynamics model can be described by the following second-order system:

Among them, β _tsc , ω _n and ξ _n represent the throttle valve opening command, natural angular frequency and damping ratio, respectively; d _β represents the external disturbance acting on the engine. The elastic model of the longitudinal model of the high-speed aircraft can be constrained by an additional canard layout, and the dynamic d _β acting on the high-speed aircraft can be regarded as a bounded disturbance.

e_2,1＝y₂，e_2,2＝γ，e_2,3＝α，e_2,4＝q；这样由公式(1)-(3)表示的高速飞行器纵向模型可改写为：In this embodiment, the tracking of the reference speed _rv and the reference altitude _rh during the cruise flight phase of the high-speed aircraft can be realized by a robust feedback controller. Let the speed, altitude and thrust be v _trim , h _trim and T _trim , respectively, at the equilibrium state in the cruise flight phase. Define the two control inputs as u ₁ =β _tsc , u ₂ =δ _e ; define the output height error and velocity error as: y ₁ =vr _v , y ₂ =hr _h ; let e ₁ =[e _1,i ] _3×1 , e ₂ =[e _2,i ] _4×1 , where e _1,1 =y ₁ , e _1,2 =β,

e _2,1 =y ₂ , e _2,2 =γ, e _2,3 =α, e _2,4 =q; thus the longitudinal model of the high-speed aircraft represented by equations (1)-(3) can be rewritten as:

y _i =C _i e _i ,i=1,2,(5)

其中上标N表示参数为标称参数；Δ₁＝[Δ_1,i]_3×1和Δ₂＝[Δ_2,i]_4×1是等效扰动，包括参数不确定性，未建模不确定性，非线性，耦合，及外部干扰。a ₂₁ =v _trim , a ₂₂ =T _trim /m ^N /V _trim ,

where the superscript N indicates that the parameters are nominal; Δ ₁ =[Δ _1,i ] _3×1 and Δ ₂ =[Δ _2,i ] _4×1 are equivalent disturbances, including parameter uncertainties, not modeled Uncertainty, nonlinearity, coupling, and external disturbances.

After the dynamic model of the high-speed aircraft is established, a flight controller needs to be established to realize the above-mentioned control method of the aircraft; specifically, the standard H _∞ control theory and the robust compensation strategy can be combined to obtain the desired tracking effect and reduce the equivalent disturbance The effect of Δi ( _i =1,2). Referring to a schematic diagram of the structure of a flight controller shown in FIG. 2; the input _ui (i=1,2) of the flight controller includes the output values ui _,H (i=1,2) of the standard H _∞ control and based on The output values of robust compensation theory ui _,R (i=1,2), _ui (i=1,2) are expressed as follows:

u _i =u _i,H +u _i,R ,i=1,2. (6)

First, H _∞ controllers for altitude and velocity channels are designed using H _∞ control theory. Let _zi (i=1,2) denote the output performance and have the following form: _zi =C _i,e e _i +C _i,u u _i,H ,i=1,2;

The systematic error is:

计算得到，其中P_i(i＝1,2)是下列方程的对称正定解：where c _1i (i=1, 2, 3), c _2i (i=1, 2, 3, 4) and θ _i (i=1, 2) are weight parameters; the uncertain error system (i.e., formula (7) )) there is a suboptimal state feedback gain K _{i,H ,} which can be obtained by

Calculated, where P _i (i=1,2) is a symmetric positive definite solution to the equation:

Among them, Q _i (i=1,2) is a symmetric positive definite matrix, and φ _i (i=1,2) is a given attenuation factor;

From the above, the H _∞ state feedback control law can finally be given by the following equation:

u _i,H =K _i,H e _i ,i=1,2, (8)

Wherein, K _1,H =[k _1,i ] _1×3 , K _2,H =[k _2,i ] _1×4 .

Secondly, a robust compensator is designed to reduce the influence of uncertainty Δ _i (i=1,2) on the closed-loop control system;

From formulas (5) and (6), we can get:

y _i =C _i e _i ,i=1,2, (9)

where A _i,H =A _i +B _i K _i,H ; let G _i (s)(i=1,2) denote the transfer function of the altitude and velocity channels from input ui _,R to output _yi , and G _i (s)=C _i (sI _i -A _i,H ) ^-1 B _i (i=1,2), where I _i (i=1,2) represents an identity matrix. From formula (9), we can get:

y _i (s)=G _i (s)u _i,R (s)+C _i (sI _i -A _i,H ) ^-1 (e _i (0)+Δ _i (s)),i=1, 2. (10)

In order to reduce the influence of the equivalent disturbance Δ _i (i=1,2), the robust compensation input ui _,R (i=1,2) can be set as the following form:

可以完全抵消Δ_i的影响。公式(9)由于Δ_i(i＝1,2)涉及状态衍生，所以

无法直接实现，故需要引进低通鲁棒滤波器来去除这些衍生量的影响，低通滤波器具有如下形式：It is obvious from equation (11) that

The effect of _Δi can be completely canceled. Equation (9) Since Δ _i (i=1,2) involves state derivation, so

It cannot be directly realized, so it is necessary to introduce a low-pass robust filter to remove the influence of these derivatives. The low-pass filter has the following form:

F ₁ (s) and F ₂ (s) act on the velocity and altitude channels, respectively, where the filtering parameters f ₁ and f ₂ are undetermined constants. In this way, the robust compensation input ui _,R (i=1,2) can be redesigned as follows:

Since Δi ( _i =1,2) cannot be directly measured or obtained, the robust compensation controller can be obtained by substituting Equation (10) into Equation (13) as follows:

In order to prove the tracking performance of the above-mentioned control method of the aircraft, this embodiment theoretically proves the tracking control performance of the above-mentioned flight controller. controller) compared to using only the H _∞ controller (which may also be referred to as the standard H _∞ state feedback controller). Discuss with the following example: Calculate the singular values of the transfer functions of the velocity channel from Δ _1,3 to y ₁ and the height channel from Δ _2,4 to y ₂ for the two controllers, and compare them in the Bode plot. Among them, the parameters of the standard H _∞ state feedback controller are selected as follows: c ₁₁ =8×10 ⁵ ,c ₁₂ =10 ⁶ ,c ₁₃ =10,c ₂₁ =5,c ₂₂ =10 ⁴ ,c ₂₃ =78, c ₂₄ =16.2, θ ₁ =1, θ ₂ =0.05, φ ₁ =300, φ ₂ =3×10 ⁵ .

Refer to the schematic diagram of the singular value comparison of the velocity channel of various controllers shown in Figure 3, and the schematic diagram of the comparison of the singular value of the altitude channel of various controllers shown in Figure 4; it can be seen from Figure 3 and Figure 4 that in the low frequency band, The flight controller provided by this embodiment has a smaller singular value than the H _∞ controller, and after adding the robust filter shown in formula (12), if a larger filter parameter is selected, the singular value of the low frequency band can become more Small. This shows that the robust controller ui _,R (i=1,2) can limit the influence of the equivalent disturbances _Δ1,3 and _Δ2,4 on the control performance. The same method can be implemented in other input and output transfer functions of high-speed aircraft.

From equations (10)-(13), the closed-loop transfer matrix of the velocity and altitude channels can be obtained as follows:

y _i (s)=(1-F _i (s))C _i (sI _i -A _i,H ) ^-1 Δ _i (s),i=1,2,

If the robust filtering parameter takes a sufficiently large value, the robust filter becomes an all-pass filter, and the gain becomes 1, so that the influence of the equivalent disturbance Δ _i (i=1, 2) can be eliminated in the entire frequency domain inhibition.

Because Δ _i (i=1,2) cannot be considered only as external disturbance, it cannot be assumed to be bounded, and at the same time, the robust filtering parameter f _i (i=1,2) cannot be large enough in practical applications, Therefore, it is necessary to further discuss the robust stability of the control system designed under nonlinearity and uncertainty.

Because the high-speed aircraft model (ie, formula (9)) does not meet the matching conditions, the system variables are transformed as follows before discussing the system stability: Definition

x ₁ =[x _1,i ] _3×1 ,x ₂ =[x _2,i ] _4×1

且

where x _1,1 =e _1,1 , x _2,1 =e _2,1 ,

and

It is further rewritten as follows:

y ⁱ =C _i e _i ,i=1,2. (15)

The norm used in this example is as follows:

in,

The equivalent perturbation is assumed to have the following bounded norm:

Among them, λ _Δx4i , λ _Δx3i , λ _Δx2i , λ _Δx1i , λ _Δci are positive constants. The tracking performance of the closed-loop control system designed in this way can be described by the following theorem:

|y_i|≤ε(i＝1,2)；进一步如果x(0)＝0，则

|y_i|≤ε(i＝1,2)。Theorem 1: Under the above assumptions, for any given bounded initial state x(0) and any constant ε, positive constants f _min and T _min can be found, if f _i ≥ f _min ( i=1,2), then any state x is bounded, and

|y _i |≤ε(i=1,2); further if x(0)=0, then

|y _i |≤ε(i=1,2).

The proof of Theorem 1 is as follows:

From formula (15), it can be obtained as follows:

||x _i || _∞ ≤λ _xi(0) +δ _i ||Δ _i || _∞ ,i=1,2, (17)

δ_i＝||(sI_i-A_i,H)^-1(1-F_i)||₁。可以看出λ_xi(0)取决于初始条件，如果初始状态x(0)有界，则存在正常数λ_xi(0)满足

前面提到如果滤波参数选大的正值，则滤波器F_i(s)(i＝1,2)增益接近1，这样δ_i可以变得更小。参考刘昊等的文章，可以得到正常数f_min1和λ_δ，对任意f_i≥f_min1(i＝1,2)，||δ||_∞≤λ_δ/f成立。where _λxi(0) is a constant and satisfies:

δ _i =||(sI _i -A _{i ,H} ) ^-1 (1-Fi )|| ₁ . It can be seen that _λxi(0) depends on the initial conditions, if the initial state x(0) is bounded, then there is a constant _λxi(0) that satisfies

As mentioned above, if the filtering parameter selects a large positive value, the gain of the filter F _i (s) (i=1, 2) is close to 1, so that δ _i can become smaller. Referring to the article of Liu Hao et al., we can obtain the positive constants f _min1 and λ _δ , and for any f _i ≥ f _min1 (i=1,2), ||δ|| _∞ ≤λ _δ /f holds.

从公式(16)-(18)可以得到如下不等式：where, δ=max _i δ _i ; let λ _x(0) =max _i λ _xi(0) , λ _Δxj =max _i λ _Δxji (j=1,2,3,4), λ _Δc =max _i λ _Δci ,

From equations (16)-(18), the following inequalities can be obtained:

If f _i (i=1,2) satisfies:

f _i ≥λ _δ , i=1,2. (20)

Substituting equations (17) and (19) into (16), we get:

From equations (17) and (18), we get:

where _λxf is a positive constant and satisfies:

The first inequality of Equation (20) determines the region of attraction as follows:

{x:||x|| _∞ ≤ξ _max }, (23)

where ξ _max is the largest positive real root of the equation:

From the above expression, there exists a positive real number f _min2 , satisfying: when f _i ≥ f _min2 (i=1,2):

ξ _max >λ _x(0) ,

ξ _max ≥||x(0)|| _∞ . (24)

In this way, for any f _i ≥ f _min3 (i=1,2) satisfy:

Thus inequality (19) holds.

Finally, from (17), (18), and (22) we get:

则对任意给定有界初始状态x(0)，对任意给定的正实数ε，可以找到正实数f_min和T_min满足：f_min≥max{f_min1,f_min2,f_min3,f_min4}时，如果f_i≥f_min(i＝1,2)，则状态x是有界的，且

|y_i|≤ε(i＝1,2)；进一步如果x(0)＝0，则

|y_i|≤ε(i＝1,2)。where c _ik (i=1,2) is a vector with 1 in the kth row and 0 in the other rows. make

Then for any given bounded initial state x(0), for any given positive real number ε, it can be found that the positive real numbers f _min and T _min satisfy: f _min ≥max{f _min1 ,f _min2 ,f _min3 ,f _min4 }, if f _i ≥ f _min (i=1,2), then the state x is bounded, and

|y _i |≤ε(i=1,2); further if x(0)=0, then

|y _i |≤ε(i=1,2).

In order to prove the tracking performance of the above-mentioned control method of the aircraft, this embodiment performs numerical simulation verification on the tracking control performance of the above-mentioned flight controller. Specifically, the linear state feedback controller designed in this embodiment (that is, the above-mentioned flight controller) includes: The H _∞ state feedback controller (ie, equation (8)) and robust compensator (ie, equation (14)) are numerically simulated in the Matlab/Simulink environment to verify the effectiveness. The values of the nominal parameters are shown in Table 1 below:

Table 1

The parameters of the H _∞ state feedback controller are given before, and the robust filtering parameters are selected as: f ₁ =50, f ₂ =5. In the trim state, the speed and altitude channel of the high-speed aircraft follows the following two reference signals:

Among them, r _vic and r _hic are step reference input instructions, β _rv =0.6, β _rh =0.3. The aircraft climbed from 110,000 feet to 111,000 feet, and the speed climbed from 15,060 feet per second to 15,260 feet per second. The initial throttle valve opening command and the angle of attack are: β ₀ =0.1762, α ₀ =1.7905 degrees, respectively. In the simulation, all uncertainty parameters, aircraft parameters and aerodynamic coefficients were set to 50% of the nominal parameters. External disturbance d _i (i=v,γ,h,α,q,β) takes the value: d _v =6sin(0.2πt)+1,d _γ =0.002sin(0.1πt)-0.02,d _h =2sin (0.3πt)+0.4,d _α =0.05sin(0.2πt)−1,d _q =0.1sin(0.2πt)+0.05,d _β =0.01sin(0.3πt)+0.3.

The numerical simulation results are shown in Fig. 5 to Fig. 9; among them, Fig. 5 is the tracking response diagram of the velocity channel and the altitude channel of the H _∞ controller to the reference signal; Fig. 6 is the pair of the velocity channel and the altitude channel of the linear state feedback controller The tracking response diagram of the reference signal; Figure 7 is the response diagram of the track angle, attack angle and roll angular velocity of the H _∞ controller and the linear state feedback controller; Figure 8 is the H _∞ controller and the linear state during the numerical simulation process. Schematic diagram of the input value change of the feedback controller to the aircraft power system; Figure 9 is a comparison diagram of the tracking errors of the H _∞ controller and the linear state feedback controller for the velocity channel and the altitude channel.

It can be seen from the above-mentioned FIGS. 5 to 9 that, compared with the existing H _∞ controller, the flight controller (ie, the linear state feedback controller) provided in this embodiment does not ignore external disturbances, and considers that the external disturbances include time-invariant Partial and sinusoidal time-varying parts.

A method for controlling an aircraft provided by an embodiment of the present invention is implemented based on the flight controller, and has the following beneficial effects:

(1) In the embodiment of the present invention, a robust linear controller is designed for high-speed aircraft with strong nonlinearity and uncertainty. The design process of the controller completely takes into account all kinds of different problems that may be involved in the current high-speed aircraft. Deterministic, and can well suppress the impact of these uncertainties on high-speed aircraft

(2) In the embodiment of the present invention, a robust compensator is designed on the basis of the standard H _∞ controller. The combination of the two controllers can suppress the influence of nonlinearity and uncertainty on the closed-loop system in the entire frequency domain, and can affect the speed of the aircraft. The reference signals of the channel and height channel enable tracking, and the tracking error converges in the prior neighborhood of the origin.

(3) The embodiment of the present invention proves the proposed flight control method, and the method can be easily implemented in practical applications.

Corresponding to the above method embodiment, refer to the schematic structural diagram of a control device of an aircraft shown in FIG. 10; the device is arranged on the controller of the aircraft; the device includes:

The signal collection module 10 is used for collecting initial flight state parameters output by the power system of the aircraft; wherein, the initial flight state parameters include flight speed and flight altitude;

a first signal generation module 11, configured to generate an initial control signal according to the initial flight state parameters and a preset reference signal;

The error signal calculation module 12 is used to input the initial control signal to the power system of the aircraft, collect the current flight state parameters output by the power system, and calculate the error signal between the current flight state parameters and the reference signal;

The second signal generating module 13 is configured to generate a compensation control signal according to the error signal;

The control module 14 is configured to determine the throttle valve and the rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal, so as to control the flight state of the aircraft.

Further, the above-mentioned first signal generation module is also used to: calculate the initial control signal ui _,H by the following formula: ui _,H =K _i,H e _i ,i=1,2, wherein, i =1 represents the flight speed; i=2 represents the flight height; K _i,H is the suboptimal state feedback gain; e _i is the error signal.

Further, the above-mentioned first signal generation module is also used for: the suboptimal state feedback gain K _i,H is obtained by the following formula:

ω_n为自然角频率；

ρ为大气密度；v_tr ⁱ _m为飞行器在巡航飞行阶段平衡状态时的速度；S为参考面积；

为平均气动弦长；C_Me为空气动力系数；I_yy为转动惯量；P_i为下述方程的对称正定解：in,

ω _n is the natural angular frequency;

ρ is the atmospheric density; _vtri ^m is the _speed of the aircraft in the equilibrium state during the cruise flight stage; S is the reference area;

is the average aerodynamic chord length; C _Me is the aerodynamic coefficient; I _yy is the moment of inertia; P _i is the symmetrical positive definite solution of the following equation:

a₂₁＝v_trim；a₂₂＝T_trim/m^N/V_trim；C_Tβ0和C_Tβ2为空气动力系数；m为飞机质量；ζ_n为阻尼比；T_trim为飞行器在巡航飞行阶段平衡状态时的推力；

和θ_i是权重参数；Q_i(i＝1,2)是对称正定矩阵，φ_i(i＝1,2)是设定的衰减因子；上标N表示参数为标称参数。

a ₂₁ =v _trim ; a ₂₂ =T _trim /m ^N /V _trim ; C _Tβ0 and C _Tβ2 are aerodynamic coefficients; m is the aircraft mass; _{ζ n} is the damping ratio; thrust;

and θ _i are weight parameters; Q _i (i=1,2) is a symmetric positive definite matrix, φ _i (i=1,2) is a set attenuation factor; the superscript N indicates that the parameters are nominal parameters.

A control device for an aircraft provided by an embodiment of the present invention can generate an initial control signal according to the initial flight state parameters output by the aircraft and a preset reference signal; after the initial control signal is input to the power system of the aircraft, the power system is collected Output the current flight state parameters, and calculate the error signal between the current flight state parameters and the reference signal; then according to the error signal, the compensation control signal can be generated; and then according to the initial control signal and the compensation control signal, the control signal of the aircraft can be determined. The control signal of flow valve and rudder deflection angle is used to control the flight state of the aircraft; this method can suppress the influence of various uncertain factors generated by the aircraft during flight by compensating the control signal, and improve the controller's ability to control the flight state of the aircraft. Track control performance.

Further, this method adopts a linear time-invariant control method, which is easy to implement and has strong practicability.

An embodiment of the present invention also provides an aircraft, the aircraft includes a processor and a sensor; the control device of the aircraft is provided in the processor.

The aircraft provided by the embodiment of the present invention has the same technical features as the control method and device for the aircraft provided by the above-mentioned embodiments, so it can also solve the same technical problem and achieve the same technical effect.

The aircraft control method, device, and aircraft computer program product provided by the embodiments of the present invention include a computer-readable storage medium storing program codes, and the instructions included in the program codes can be used to execute the methods described in the foregoing method embodiments. For the specific implementation, reference may be made to the method embodiments, which will not be repeated here.

The functions, if implemented in the form of software functional units and sold or used as independent products, may be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention can be embodied in the form of a software product in essence, or the part that contributes to the prior art or the part of the technical solution. The computer software product is stored in a storage medium, including Several instructions are used to cause a computer device (which may be a personal computer, a server, or a network device, etc.) to execute all or part of the steps of the methods described in the various embodiments of the present invention. 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 .

Finally, it should be noted that the above-mentioned embodiments are only specific implementations of the present invention, and are used to illustrate the technical solutions of the present invention, but not to limit them. The protection scope of the present invention is not limited thereto, although referring to the foregoing The embodiment has been described in detail the present invention, those of ordinary skill in the art should understand: any person skilled in the art who is familiar with the technical field within the technical scope disclosed by the present invention can still modify the technical solutions described in the foregoing embodiments. Or can easily think of changes, or equivalently replace 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 invention, and should be covered in the present invention. within the scope of protection. Therefore, the protection scope of the present invention should be based on the protection scope of the claims.

Claims (3)

1. A control method for an aircraft, wherein the method is applied to a controller of the aircraft; the method comprises: Collect initial flight state parameters output by the power system of the aircraft; wherein, the initial flight state parameters include flight speed and flight altitude; generating an initial control signal according to the initial flight state parameters and a preset reference signal; inputting the initial control signal into the power system of the aircraft, collecting current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and the reference signal; generating a compensation control signal according to the error signal; determining a throttle valve and a rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal, so as to control the flight state of the aircraft; The step of generating an initial control signal according to the initial flight state parameters and the preset reference signal includes: The initial control signal ui _,H is calculated by the following formula: u _i,H =K _i,H e _i ,i=1,2,; Among them, i=1 represents the flight speed; i=2 represents the flight height; K _{i, H} is the suboptimal state feedback gain; e _i is the error signal; The suboptimal state feedback gain K _i,H is obtained by the following formula:

in,

ω _n is the natural angular frequency;

ρ is the atmospheric density; v _trim is the speed of the aircraft in the equilibrium state of the cruise flight stage; S is the reference area;

is the average aerodynamic chord length; C _Me is the aerodynamic coefficient; I _yy is the moment of inertia; _Pi is the symmetric positive definite solution of the following equation:

a ₂₁ =v _trim ; a ₂₂ =T _trim /m ^N /V _trim ; C _Tβ0 and C _Tβ2 are the aerodynamic coefficients; m is the aircraft mass; ζ _n is the damping ratio; T _trim is the balance of the aircraft in the cruise flight phase state of thrust;

c _1j , j=1, 2, 3, c _2j , j=1, 2, 3, 4 and θ _i are weight parameters; Q _i , i=1, 2 are symmetric positive definite matrices, φ _i , i=1, 2 is the set attenuation factor; the superscript N indicates that the parameter is a nominal parameter; The step of generating a compensation control signal according to the error signal includes: The compensation control signal ui _,R (s) is calculated by the following formula:

Among them, s is the Laplacian operator; F ₁ (s)=f ₁ ³ /(s+f ₁ ) ³ ,

f ₁ and f ₂ are set filtering parameters, which are positive numbers; y ₁ =vr _v , y ₂ =hr _h ; v is the speed of the aircraft; h is the height of the aircraft; r _v is the aircraft The reference speed in the cruise flight stage; _rh is the reference altitude of the aircraft in the cruise flight stage; G _i (s) is the transfer function from ui _,R (s) to _yi ; The transfer function G _i (s) is obtained by the following formula: G _i (s)=C _i (sI _i -A _i,H ) ^-1 B _i , i=1,2; Wherein, I _i , i=1, 2 is the identity matrix;

A _i,H =A _i +B _i K _i,H ; The step of determining the throttle valve and the rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal to control the flight state of the aircraft, including: Calculate the final control signal of the aircraft u _i =ui _,H +u _i,R , i=1,2; The flying speed and/or the flying height of the aircraft are controlled according to the final control signal. 2. A control device for an aircraft, wherein the device is arranged on a controller of the aircraft; the device comprises: a signal collection module, used for collecting initial flight state parameters output by the power system of the aircraft; wherein, the initial flight state parameters include flight speed and flight altitude; a first signal generation module, configured to generate an initial control signal according to the initial flight state parameters and a preset reference signal; An error signal calculation module, configured to input the initial control signal to the power system of the aircraft, collect the current flight state parameters output by the power system, and calculate the error between the current flight state parameters and the reference signal Signal; a second signal generating module, configured to generate a compensation control signal according to the error signal; a control module, configured to determine a throttle valve and a rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal, so as to control the flight state of the aircraft; The first signal generation module is also used for: The initial control signal ui _,H is calculated by the following formula: u _i,H =K _i,H e _i ,i=1,2,; Among them, i=1 represents the flight speed; i=2 represents the flight height; K _{i, H} is the suboptimal state feedback gain; e _i is the error signal; The first signal generation module is further configured to: obtain the suboptimal state feedback gain K _i,H by the following formula:

in,

ω _n is the natural angular frequency;

ρ is the atmospheric density; v _trim is the speed of the aircraft in the equilibrium state of the cruise flight stage; S is the reference area;

is the average aerodynamic chord length; C _Me is the aerodynamic coefficient; I _yy is the moment of inertia; _Pi is the symmetric positive definite solution of the following equation:

a ₂₁ =v _trim ; a ₂₂ =T _trim /m ^N /V _trim ; C _Tβ0 and C _Tβ2 are the aerodynamic coefficients; m is the aircraft mass; ζ _n is the damping ratio; T _trim is the balance of the aircraft in the cruise flight phase state of thrust;

c _1j , j=1, 2, 3, c _2j , j=1, 2, 3, 4 and θ _i are weight parameters; Q _i , i=1, 2 are symmetric positive definite matrices, φ _i , i=1, 2 is the set attenuation factor; the superscript N indicates that the parameter is a nominal parameter; The second signal generation module is also used for: The compensation control signal ui _,R (s) is calculated by the following formula:

Among them, s is the Laplacian operator; F ₁ (s)=f ₁ ³ /(s+f ₁ ) ³ ,

f ₁ and f ₂ are set filtering parameters, which are positive numbers; y ₁ =vr _v , y ₂ =hr _h ; v is the speed of the aircraft; h is the height of the aircraft; r _v is the aircraft The reference speed in the cruise flight stage; _rh is the reference altitude of the aircraft in the cruise flight stage; G _i (s) is the transfer function from ui _,R (s) to _yi ; The transfer function G _i (s) is obtained by the following formula: G _i (s)=C _i (sI _i -A _i,H ) ^-1 B _i , i=1,2; Wherein, I _i , i=1, 2 is the identity matrix;

A _i,H =A _i +B _i K _i,H ; The control module is also used for: Calculate the final control signal of the aircraft u _i =ui _,H +u _i,R , i=1,2; The flying speed and/or the flying height of the aircraft are controlled according to the final control signal. 3. An aircraft, characterized in that the aircraft comprises a processor and a sensor; the device of claim 2 is provided in the processor.

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