CN112027117B - A control method for side-slip and roll compound turns of aircraft based on attitude measurement - Google Patents

Abstract

The invention relates to an attitude measurement-based aircraft sideslip-roll compound turning control method. It installs YIN600-R inertial integrated navigation system to measure yaw angle, roll angle and lateral acceleration, and then obtains lateral velocity and position signals through integration. Then design a nonlinear low-pass integral corrector to obtain the position error integral signal, and then design a nonlinear low-pass filter corrector to obtain the rate filter signal, and then through symmetrical design and combination, get the expected signal of yaw angle and roll angle, and then pass the attitude Angle error, high-pass differential correction and nonlinear integration to obtain the final yaw and roll channel control signals, and the two channels act simultaneously to realize the cooperative and integrated lateral turning control of the aircraft. The advantage of this method is that the integrated control of sideslip and roll makes the lateral turning of the aircraft very fast, and at the same time, the symmetrical design makes it very stable.

Description

A control method for side-slip and roll compound turns of aircraft based on attitude measurement

technical field

The invention relates to the field of aircraft stability and turning control, in particular to an attitude measurement-based aircraft sideslip and roll compound turning control method.

Background technique

The current mainstream methods for the lateral center of mass movement of the aircraft are sideslip turn and bank turn. Among them, the bank turn is also called roll turn, which uses the change of lift direction during the rolling process of the aircraft to provide centripetal force, so that the turning process of the aircraft is faster than the traditional sideways turn. Slippery turns are faster. However, the idea of the traditional sideslip turn is to keep the roll channel basically stationary, and use the sideslip angle to provide the turning power; in order to eliminate the coupling of the yaw channel, the banked turn is generally in a stable state in the yaw channel. Therefore, in order to ensure the stability of the system, these two methods fully adopt the two-channel fly-off design idea, and do not use the nonlinear coupling effect between yaw and sideslip. We know that the moving object has a coupled physical nature when it rolls and slides sideways. Based on the above background reasons, we propose a method that uses attitude measurement and attitude stabilization as the cornerstone, and at the same time coordinates the yaw angle and roll angle. Control, realizes the side-slip and roll cooperative integration of the aircraft turning control method, which makes the turning process very fast, and also makes this method have high engineering application value.

It should be noted that the information disclosed in the above background technology section is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those of ordinary skill in the art.

Contents of the invention

The purpose of the present invention is to provide a method for controlling aircraft side-slip and roll compound turns based on attitude measurement, and then at least to a certain extent overcome the uncoordinated matching control of aircraft turning, yaw, and roll channels caused by the limitations and defects of related technologies The problem.

According to one aspect of the present invention, there is provided a method for controlling an aircraft sideslip and roll compound turn based on attitude measurement, comprising the following steps:

Step S10, install the YIN600-R inertial integrated navigation system on the aircraft, and measure the lateral acceleration, yaw angle and roll angle of the aircraft;

Step S20, according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system, perform two integrations to obtain the lateral speed signal and the lateral position signal respectively, and compare them with the lateral position command signal to obtain the lateral position error Signal;

Step S30, according to the lateral position error signal, design a nonlinear low-pass integral corrector and perform linear integration to obtain a lateral position filtered integral signal;

Step S40, according to the lateral speed signal, design a nonlinear low-pass filter corrector, obtain the filtered speed signal, and provide a damping signal for the system;

Step S50, performing linear combination and superposition according to the lateral position error signal, the lateral position filtered integral signal and the filtered speed signal, respectively designing a desired yaw angle signal and a desired roll angle signal;

Step S60, compare the yaw angle signal measured by the YIN600-R inertial integrated navigation system with the desired yaw angle signal to obtain the yaw angle error signal, and then design a nonlinear high-pass differential corrector to obtain the yaw angle The error filter differential signal is superimposed on the error nonlinear integral signal to form the yaw channel control signal;

Step S70, comparing the roll angle signal measured by the YIN600-R inertial integrated navigation system with the roll angle expected signal to obtain a roll angle error signal, and then designing a nonlinear high-pass differential corrector to obtain a roll angle error filtered differential signal , and then superimpose the error nonlinear integral signal to form the roll channel control signal, which is controlled simultaneously with the yaw channel to realize the side-slip roll cooperative turning of the aircraft.

In an exemplary embodiment of the present invention, the YIN600-R inertial integrated navigation system is installed on the aircraft to measure the yaw angle, roll angle and lateral acceleration of the aircraft, and then the yaw angle of the aircraft is measured and integrated twice to obtain The lateral speed signal and the lateral position signal are compared with the lateral position command signal to obtain the lateral position error signal including;

v _z = ∫ a _z dt;

z=∫v _z dt;

e _z = zz ^d ;

Among them, a _z is the lateral acceleration of the aircraft measured by the YIN600-R inertial integrated navigation system, and a _z (n) represents the data of the lateral acceleration at time t=n*ΔT, where n=1,2,3 ..., ΔT is the data sampling period. v _z is the lateral speed signal, and dt means integrating the time signal. z is the lateral position signal, z ^d is the lateral desired position signal set according to the lateral task of the aircraft, e _z is the lateral position error signal.

In an exemplary embodiment of the present invention, according to the lateral position error signal, a nonlinear low-pass integral corrector is designed and linearly integrated to obtain the lateral position filtered integral signal including:

e _z1 (n+1)=e _z1 (n)+D _a *T _a ;

s ₁ =∫e _z1 dt;

Where e _z is the above-mentioned lateral position error signal, e _z1 is the low-pass correction signal, T _a represents the time interval of the data, k ₁ , k ₂ and ε ₁ are constant parameters, and the detailed design is shown in the following case implementation . dt represents the integral of the time signal, and s ₁ is the final lateral position filtered integral signal.

In an exemplary embodiment of the present invention, according to the lateral velocity signal, a non-linear low-pass filter corrector is designed to obtain the filtered velocity signal including:

v _z1 (n+1)=v _z1 (n)+D _b *T _a ;

Among them, v _z is the lateral speed signal mentioned above, v _z1 is the filtered speed signal, k ₃ , k ₄ and ε ₁ are constant parameters, and the detailed design is shown in the case implementation later.

In an exemplary embodiment of the present invention, linear combination and superposition are performed according to the lateral position error signal, the lateral position filtered integral signal and the filtered speed signal, and the expected yaw angle signal and the expected roll angle signal are respectively designed include:

Where e _z is the lateral position error signal, s ₁ is the filtered integral signal of the lateral position, v _z1 is the filtered velocity signal, ψ ^d is the expected signal of yaw angle, γ ^d is the expected signal of roll angle, c ₁₁ , c ₁₂ , c ₁₃ , c ₁₄ , ε ₁₁ , c ₂₁ , c ₂₂ , c ₂₃ , c ₂₄ , and ε ₂₁ are constant value parameters, and their detailed design is shown in the following case implementation.

In an exemplary embodiment of the present invention, the yaw angle signal measured according to the YIN600-R inertial integrated navigation system is compared with the described yaw angle expected signal to obtain the yaw angle error signal, and then the nonlinear high-pass is designed The differential corrector obtains the yaw angle error filtered differential signal, and then superimposes the error nonlinear integral signal to form a yaw channel control signal including:

e _ψ = ψ-ψ ^d ;

D ₁ =d ₁ (e _ψ (n+1)-e _ψ (n))+d ₂ sin(e _ψ (n+1)-e _ψ (n));

D ₂ (n+1)=D ₂ (n)+D _ψ *T _a1 ;

Among them, ψ is the yaw angle measurement signal, e _ψ is the yaw angle error signal, ε _ψ is a constant value parameter, and its detailed design is shown in the following case implementation, and dt means the integral of the time signal. D _ψ is the differential correction signal of yaw angle error, T _a1 is the time interval of data sampling, and d ₁ , d ₂ , d ₃ , d ₄ , and ε _k1 are constant parameters, and the detailed selection will be implemented later. u _p is the control signal of the yaw channel, l ₁ , l ₂ , l ₃ , l ₄ , and ε _l are the constant value control parameters, and the detailed design is shown in the following case implementation.

In an exemplary embodiment of the present invention, the roll angle signal measured by the YIN600-R inertial integrated navigation system is compared with the roll angle expected signal to obtain the roll angle error signal, and then the nonlinear high-pass differential corrector is designed , to obtain the roll angle error filtered differential signal, and then superimpose the error nonlinear integral signal to form the roll channel control signal including:

e _γ = γ-γ ^d ;

D ₃ =d ₅ (e _γ (n+1)-e _γ (n))+d ₆ sin(e _γ (n+1)-e _γ (n));

D ₄ (n+1)=D ₄ (n)+D _γ *T _a1 ;

Among them, γ is the roll angle measurement signal, e _γ is the roll angle error signal, ε _γ is a constant value parameter, and its detailed design is shown in the following case implementation, and dt represents the integral of the time signal. s ₃ is the nonlinear integral signal of roll angle error, D _γ is the differential correction signal of roll angle error, T _a1 is the time interval of data sampling, d ₅ , d ₆ , d ₇ , d ₈ , ε _k2 are constant parameters, For detailed selection, see the implementation below. u _g is the control signal of the rolling channel, l ₅ , l ₆ , l ₇ , l ₈ , and ε _g are the constant value control parameters, and its detailed design is shown in the following case implementation.

Finally, the obtained yaw channel control value u _p is sent to the yaw rudder system to realize the tracking of the expected value of yaw angle. The obtained rolling channel control value _ug is sent to the rolling rudder system to realize the tracking of the expected value of the rolling angle. Therefore, under the joint action of the roll channel and the yaw channel, the sideslip and roll coordinated turn of the aircraft can be realized.

Beneficial effect

The present invention provides an attitude measurement-based aircraft side-slip and roll composite turn control method, which has the advantage of realizing the coordinated control of sideslip and roll turns, so that the coupling effect of the yaw channel and the roll channel is obtained. effective use. It also enables the aircraft to perform side-slip motion while side-rolling, so that the rapidity of turning is effectively improved, and the maneuverability of the aircraft is also enhanced. Especially through the design of symmetrical yaw angle and roll angle expected signals, the stability of the entire aircraft turning process is guaranteed through attitude measurement and attitude stabilization, so that the whole method has high engineering application value.

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

Description of drawings

The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and together with the description serve to explain the principles of the invention. Apparently, the drawings in the following description are only some embodiments of the present invention, and those skilled in the art can obtain other drawings according to these drawings without creative efforts.

Fig. 1 is the flow chart of a kind of aircraft sideslip and roll compound turning control method based on attitude measurement provided by the present invention;

Fig. 2 is the physical figure of the YIN600-R inertial integrated navigation system of the method provided by the embodiment of the present invention;

Fig. 3 is the aircraft lateral acceleration curve (unit: meter per second square) of the method provided by the embodiment of the present invention;

Fig. 4 is the aircraft yaw angle curve (unit: degree) of the method provided by the embodiment of the present invention;

Fig. 5 is the aircraft roll angle curve (unit: degree) of the method provided by the embodiment of the present invention;

Fig. 6 is the aircraft lateral velocity curve (unit: meter per second) of the method provided by the embodiment of the present invention;

Fig. 7 is the aircraft lateral position curve (unit: meter) of the method provided by the embodiment of the present invention;

Fig. 8 is the aircraft lateral position error curve (unit: meter) of the method provided by the embodiment of the present invention;

Fig. 9 is the aircraft lateral position filter integral signal curve (unitless) of method provided by the embodiment of the present invention;

Fig. 10 is the aircraft filtering speed signal curve (no unit) of the method provided by the embodiment of the present invention;

Fig. 11 is the yaw angle expected signal curve (unit: degree) of the method provided by the embodiment of the present invention;

Fig. 12 is the roll angle expected signal curve (unit: degree) of the method provided by the embodiment of the present invention;

Fig. 13 is the yaw channel control signal curve (no unit) of the method provided by the embodiment of the present invention;

Fig. 14 is a rolling channel control signal curve (without unit) of the method provided by the embodiment of the present invention.

Fig. 15 is a sideslip angle signal curve (unit: degree) of the method provided by the embodiment of the present invention.

Detailed ways

Example embodiments will now be described more fully with reference to the accompanying drawings. Example embodiments may, however, be embodied in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete and fully convey the concept of example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided in order to give a thorough understanding of embodiments of the invention. However, those skilled in the art will appreciate that the technical solution of the present invention may be practiced without one or more of the specific details, or other methods, components, devices, steps, etc. may be adopted. In other instances, well-known technical solutions have not been shown or described in detail to avoid obscuring aspects of the invention.

The invention provides an aircraft sideslip and roll compound turning control method based on attitude measurement, which measures the yaw angle and the roll angle of the aircraft, and simultaneously stabilizes and tracks the yaw angle and the roll angle. Then, the lateral error is obtained by measuring the lateral acceleration of the aircraft and comparing the integral with the lateral center of mass control task command signal. At the same time, the expected signal of the yaw angle and pitch angle is obtained through velocity filtering and error integration. The turning channel tracks the desired attitude angle and realizes the lateral center of mass control, so that the two channels of roll and yaw can complete coordination and cooperation, greatly improving the speed of turning, and at the same time, the design of symmetrical attitude command signal and attitude stabilization system , and the stability of the two channels is guaranteed, so that the rapidity and stability of the method can be improved and enhanced.

Next, an attitude measurement-based aircraft side-slip-roll compound turn control method of the present invention will be further explained and described in conjunction with the accompanying drawings. Referring to Fig. 1, this a kind of attitude measurement-based aircraft sideslip and roll compound turning control method includes the following steps:

Step S10, install the YIN600-R inertial integrated navigation system on the aircraft, and measure the lateral acceleration, yaw angle and roll angle of the aircraft.

Specifically, first install the YIN600-R inertial integrated navigation system on the aircraft. Its physical picture is shown in Figure 2. Its performance indicators are as follows: weight 150g, size 59*45*23.8mm, acceleration measurement range -16g to 16g, attitude The accuracy is 0.1 degrees, the bandwidth is 80Hz, the power consumption is 1150mW, the output rate is 200Hz, the speed accuracy is 0.05 meters per second, and the positioning accuracy is 0.1 meters. The output mode is RS232, RS485, RS422 level interface standard optional, working temperature -40～+85℃.

Secondly, the YIN600-R inertial integrated navigation system is used to measure the yaw angle of the aircraft, which is counted as ψ, ψ(n) represents the data of the yaw angle at time t=n*ΔT, where n=1,2, 3..., ΔT is the data sampling period, and its detailed design is shown in the following case implementation. Adopt this inertial navigation system to measure the lateral acceleration of aircraft simultaneously, count as az, az (n) represents the data of lateral acceleration at time t=n*ΔT moment, wherein n=1,2,3..., ΔT is data The sampling period, its detailed design can be chosen to be the same as that of the yaw angle measurement.

Finally, use the YIN600-R inertial integrated navigation system to measure the roll angle of the aircraft, which is counted as γ, γ(n) represents the data of the roll angle at time t=n*ΔT, where n=1,2,3... , ΔT is the data sampling period, and its detailed design is shown in the following case implementation.

Step S20, according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system, perform two integrations to obtain the lateral speed signal and the lateral position signal respectively, and compare them with the lateral position command signal to obtain the lateral position error Signal;

Specifically, firstly, according to the lateral acceleration measurement signal a _z , integration is performed to obtain a lateral velocity signal, which is counted as v _z , and the integration method is as follows:

v _z = ∫ a _z dt;

where dt means integrating the time signal.

Again, linearly integrate the lateral speed measurement signal a _z to obtain the lateral position signal, which is counted as z, and the integration method is as follows:

z=∫v _z dt;

where dt means integrating the time signal.

Finally, set the lateral desired position signal according to the lateral task of the aircraft, denoted as z ^d . Then compare it with the said lateral position signal to get the lateral position error signal, denoted as e _z , the comparison method is as follows:

e _z = zz ^d ;

Step S30, according to the lateral position error signal, design a nonlinear low-pass integral corrector and perform linear integration to obtain a lateral position filtered integral signal.

Specifically, first, the following nonlinear low-pass correction is performed on the lateral position error signal e _z , and the obtained low-pass correction signal is counted as e _z1 , and its calculation method is as follows:

e _z1 (n+1)=e _z1 (n)+D _a *T _a ;

Among them, T _a represents the time interval of the data, and k ₁ , k ₂ and ε ₁ are constant parameters. For the detailed design, please refer to the case implementation later.

Secondly, the above-mentioned low-pass correction signal is integrated to obtain the final lateral position filtered integral signal, denoted as s ₁ , and its calculation method is as follows:

s ₁ =∫e _z1 dt

where dt represents the integration of the time signal.

Step S40, according to the lateral speed signal, design a nonlinear low-pass filter corrector, obtain the filtered speed signal, and provide a damping signal for the system;

Specifically, the following nonlinear low-pass filter correction is performed on the lateral velocity signal v _z to obtain the filtered velocity signal, denoted as v _z1 , and its calculation method is as follows:

v _z1 (n+1)=v _z1 (n)+D _b *T _a ;

Among them, k ₃ , k ₂ and ε ₁ are constant parameters, and their detailed design is shown in the following case implementation.

Step S50, performing linear combination and superposition according to the lateral position error signal, the lateral position filtered integral signal and the filtered speed signal, respectively designing a desired yaw angle signal and a desired roll angle signal.

Specifically, firstly, according to the lateral position error signal e _z , the lateral position filtered integral signal s ₁ and the filtered speed signal v _z1 are linearly combined and superimposed to design the expected yaw angle signal, denoted by ψ ^d , where It is calculated as follows:

Among them, c ₁₁ , c ₁₂ , c ₁₃ , c ₁₄ , and ε ₁₁ are constant value parameters, and their detailed design is shown in the following case implementation.

Secondly, in a symmetrical manner, according to the lateral position error signal e _z , the lateral position filtered integral signal s ₁ and the filtered speed signal v _z1 are linearly combined and superimposed to design the expected roll angle signal, denoted as γ ^d , which is calculated as follows:

Among them, c ₂₁ , c ₂₂ , c ₂₃ , c ₂₄ , and ε ₂₁ are constant value parameters, and their detailed design is shown in the following case implementation.

Step S60, compare the yaw angle signal measured by the YIN600-R inertial integrated navigation system with the desired yaw angle signal to obtain the yaw angle error signal, and then design a nonlinear high-pass differential corrector to obtain the yaw angle The error filter differential signal is superimposed on the error nonlinear integral signal to form the yaw channel control signal.

Specifically, firstly, the yaw angle measurement signal is compared with the yaw angle expected signal to obtain a yaw angle error signal, denoted as e _ψ , and the comparison method is as follows:

e _ψ = ψ-ψ ^d ;

Secondly, according to the yaw angle error signal, the nonlinear linear integration is performed to obtain the yaw angle error nonlinear integral signal, denoted as s ₂ , and the integration method is as follows:

Among them, ε _ψ is a constant value parameter, and its detailed design is shown in the following case implementation, and dt represents the integral of the time signal.

Then, according to the yaw angle error signal, construct the nonlinear high-pass differential corrector as follows, and obtain the differential correction signal of the yaw angle error, which is counted as D _ψ , and its calculation method is as follows:

D ₁ =d ₁ (e _ψ (n+1)-e _ψ (n))+d ₂ sin(e _ψ (n+1)-e _ψ (n));

D ₂ (n+1)=D ₂ (n)+D _ψ *T _a1 ;

Among them, T _a1 is the time interval of data sampling, and d ₁ , d ₂ , d ₃ , d ₄ , and ε _k1 are constant value parameters, and their detailed selection is described later in the implementation.

Finally, linearly combine the yaw angle error signal e _ψ of the aircraft, the nonlinear integral signal s ₂ of the yaw angle error, and the differential correction signal D _ψ of the yaw angle error to obtain the final yaw channel control signal. As u _p , its calculation method is as follows:

Among them, l ₁ , l ₂ , l ₃ , l ₄ , and ε _l are constant value control parameters, and their detailed design is shown in the following case implementation.

Finally, the obtained yaw channel control value u _p is sent to the yaw rudder system to realize the tracking of the expected value of yaw angle.

Step S70, comparing the roll angle signal measured by the YIN600-R inertial integrated navigation system with the roll angle expected signal to obtain a roll angle error signal, and then designing a nonlinear high-pass differential corrector to obtain a roll angle error filtered differential signal , and then superimpose the error nonlinear integral signal to form the roll channel control signal, which is controlled simultaneously with the yaw channel to realize the side-slip roll cooperative turning of the aircraft.

Specifically, firstly, the roll angle measurement signal is compared with the roll angle expected signal to obtain the roll angle error signal, denoted as e _γ , and the comparison method is as follows:

e _γ = γ-γ ^d ;

Secondly, according to the roll angle error signal, the nonlinear linear integration is performed to obtain the roll angle error nonlinear integral signal, denoted as s ₃ , and the integration method is as follows:

Among them, ε _γ is a constant value parameter, and its detailed design is shown in the following case implementation, and dt represents the integral of the time signal. Then, construct the nonlinear high-pass differential corrector as follows according to the roll angle error signal, and obtain the differential correction signal of the roll angle error, which is counted as D _γ , and its calculation method is as follows:

D ₃ =d ₅ (e _γ (n+1)-e _γ (n))+d ₆ sin(e _γ (n+1)-e _γ (n));

D ₄ (n+1)=D ₄ (n)+D _γ *T _a1 ;

Among them, T _a1 is the time interval of data sampling, and d ₅ , d ₆ , d ₇ , d ₈ , and ε _k2 are constant value parameters, and their detailed selection is described later in the implementation.

Finally, linearly combine the aircraft roll angle error signal e _γ , roll angle error nonlinear integral signal s ₃ , and roll angle error differential correction signal D _γ to obtain the final roll channel control signal, denoted as u _g , which is calculated as follows:

Among them, l ₅ , l ₆ , l ₇ , l ₈ , and ε _g are constant control parameters, and their detailed design is shown in the following case implementation.

Finally, the obtained roll channel control value _ug is sent to the roll rudder system to track the expected value of the roll angle. Therefore, under the joint action of the roll channel and the yaw channel, the sideslip and roll coordinated turn of the aircraft can be realized.

Case implementation and simulation experiment results analysis

In step S10, install the YIN600-R inertial integrated navigation system on the aircraft, measure the lateral acceleration of the aircraft as shown in Figure 3, measure the yaw angle as shown in Figure 4, and measure the roll angle as shown in Figure 5.

In step S20, according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system, two integrations are performed to obtain the lateral velocity signal and the lateral position signal respectively as shown in Figure 6 and Figure 7, and the lateral position error is obtained The signal is shown in Figure 8.

In step S30, T _a =0.001, k ₁ =10, k ₂ =5 and ε ₁ =0.2 are selected to obtain the filtered integral signal of the lateral position as shown in FIG. 9 .

In step S40, k ₃ =10, k ₄ =6 and ε ₁ =0.5 are selected to obtain the filtered velocity signal as shown in FIG. 10 .

In step S50, select c ₁₁ =0.1, c ₁₂ =0.05, c ₁₃ =0.1, c ₁₄ =3, ε ₁₁ =0.8 to obtain the desired yaw angle signal as shown in Figure 11, select c ₂₁ =0.2, c ₂₂ =0.1, c ₂₃ =0.2, c ₂₄ =8, ε ₂₁ =0.4; the desired roll angle signal is shown in FIG. 12 .

In step S60, select d ₁ =1000, d ₂ =500, d ₃ =20, d ₄ =8, ε _k1 =2, T _a1 =0.001, ε _ψ =0.7, l ₁ =2, l ₂ =0.2 , l ₃ =0.4, l ₄ =5, ε _l =4, the yaw channel control signal is obtained as shown in Figure 13 .

In step S70, select d ₅ =1000, d ₆ =700, d ₇ =10, d ₈ =8, ε _k2 =4, T _a1 =0.001, ε _γ =.07, l ₅ =2, l ₆ = 0.1, l ₇ =0.3, l ₈ =7, ε _g =6, the control signal of the rolling channel is obtained as shown in FIG. 14 . It is controlled simultaneously with the yaw channel to realize the sideslip and roll cooperative turning of the aircraft. The sideslip angle during the turning process is shown in Figure 15.

It can be seen from Figure 4 that the maximum yaw angle reaches 6 degrees, it can be seen from Figure 5 that the maximum roll angle reaches 20 degrees, and it can be seen from Figure 6 that the maximum lateral speed reaches 14 meters per second. It can be seen from Fig. 7 that the aircraft completes the large maneuvering turn in about 5s. It can be seen from Figure 15 that the sideslip angle reaches a maximum of 4.2 degrees, and it can be seen from Figures 13 and 14 that the yaw channel control signal does not exceed 8 degrees, and the roll channel control signal does not exceed 6 degrees, which means Both the yaw rudder and the roll rudder are not more than 8 degrees, so as to meet the limit requirements of engineering applications. It can be seen from the above cases that the integrated turning strategy of roll and sideslip makes the turning process of the aircraft very fast and the coupling is more conducive to the completion of the turning action, so that the present invention has good engineering practical value.

Other embodiments of the invention will be readily apparent to those skilled in the art from consideration of the specification and practice of such an invention. This application is intended to cover any modification, use or adaptation of the present invention. These modifications, uses or adaptations follow the general principles of the present invention and include common knowledge or conventional technical means in the technical field not specified in the present invention . The specification and examples are to be considered exemplary only, with the true scope and spirit of the invention indicated by the appended claims.

Claims (6)

1. an aircraft sideslip and roll compound turning control method based on attitude measurement, is characterized in that, comprises the following steps: Step S10, install the YIN600-R inertial integrated navigation system on the aircraft, and measure the lateral acceleration, yaw angle and roll angle of the aircraft; Step S20, according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system, perform two integrations to obtain the lateral speed signal and the lateral position signal respectively, and compare them with the lateral position command signal to obtain the lateral position error Signal; Step S30, according to the lateral position error signal, design a nonlinear low-pass integral corrector and perform linear integration to obtain a lateral position filtered integral signal; Step S40, according to the lateral speed signal, design a nonlinear low-pass filter corrector, obtain the filtered speed signal, and provide a damping signal for the system; Step S50, performing linear combination and superposition according to the lateral position error signal, the lateral position filtered integral signal and the filtered speed signal, respectively designing a desired yaw angle signal and a desired roll angle signal; Step S60, compare the yaw angle signal measured by the YIN600-R inertial integrated navigation system with the desired yaw angle signal to obtain the yaw angle error signal, and then design a nonlinear high-pass differential corrector to obtain the yaw angle The error filter differential signal is superimposed on the error nonlinear integral signal to form the yaw channel control signal; Step S70, comparing the roll angle signal measured by the YIN600-R inertial integrated navigation system with the roll angle expected signal to obtain a roll angle error signal, and then designing a nonlinear high-pass differential corrector to obtain a roll angle error filtered differential signal , and then superimpose the error nonlinear integral signal to form the roll channel control signal, which is controlled simultaneously with the yaw channel to realize the side-slip roll cooperative turning of the aircraft. 2. a kind of aircraft sideslip and roll compound turning control method based on attitude measurement according to claim 1, according to described lateral position error signal, design nonlinear low-pass integral corrector and carry out linear integration, obtain side Filtering the integrated signal to position involves: v _z = ∫ a _z dt; z=∫v _z dt; e _z = zz ^d ;

e _z1 (n+1)=e _z1 (n)+D _a *T _a ; s ₁ =∫e _z1 dt; Among them, a _z is the lateral acceleration of the aircraft measured by the YIN600-R inertial integrated navigation system, and a _z (n) represents the data of the lateral acceleration at time t=n*ΔT, where n=1,2,3 …, ΔT is the data sampling period; v _z is the lateral velocity signal, dt represents the integration of the time signal; z is the lateral position signal, z ^d is the lateral desired position signal set according to the lateral task of the aircraft, and ez is lateral position error signal; e _z is the lateral position error signal described above, e _z1 is the low-pass correction signal, T _a represents the time interval of data, k ₁ , k ₂ and ε ₁ are constant parameters; dt represents the Integral of the time signal, s ₁ is the final lateral position filtered integral signal. 3. a kind of aircraft sideslip and roll compound turning control method based on attitude measurement according to claim 2, according to described lateral velocity signal, design non-linear low-pass filter corrector, obtain filter velocity signal and comprise:

v _z1 (n+1)=v _z1 (n)+D _b *T _a ; Where v _z is the lateral speed signal, v _z1 is the filtered speed signal, and k ₃ , k ₄ and ε ₂ are constant parameters. 4. a kind of aircraft sideslip and roll compound turning control method based on attitude measurement according to claim 3, according to described lateral position error signal, lateral position filter integration signal and filter speed signal, carry out linear combination and Superposition, respectively design the expected signal of yaw angle and the expected signal of roll angle include:

Where e _z is the lateral position error signal, s ₁ is the filtered integral signal of the lateral position, v _z1 is the filtered velocity signal, ψ ^d is the expected signal of yaw angle, γ ^d is the expected signal of roll angle, c ₁₁ , c ₁₂ , c ₁₃ , c ₁₄ , ε ₁₁ , c ₂₁ , c ₂₂ , c ₂₃ , c ₂₄ , ε ₂₁ are constant value parameters. 5. a kind of aircraft sideslip and roll compound turning control method based on attitude measurement according to claim 4, according to the yaw angle signal that YIN600-R inertial integrated navigation system measures and obtains and described yaw angle expectation signal carries out By comparison, the yaw angle error signal is obtained, and then a nonlinear high-pass differential corrector is designed to obtain the yaw angle error filtered differential signal, and then the error nonlinear integral signal is superimposed to form a yaw channel control signal including: e _ψ = ψ-ψ ^d ;

D ₁ =d ₁ (e _ψ (n+1)-e _ψ (n))+d ₂ sin(e _ψ (n+1)-e _ψ (n));

D ₂ (n+1)=D ₂ (n)+D _ψ *T _a1 ;

Where ψ is the measurement signal of yaw angle, γ ^d is the expected signal of roll angle, ψ ^d is the expected signal of yaw angle, e _ψ is the error signal of yaw angle, ε _ψ is a constant value parameter, and dt represents the integral of the time signal; D _ψ is the differential correction signal of yaw angle error, s ₂ is the nonlinear integral signal of yaw angle error, T _a1 is the time interval of data sampling, d ₁ , d ₂ , d ₃ , d ₄ , and ε _k1 are constant values parameters; u _p is the yaw channel control signal, l ₁ , l ₂ , l ₃ , l ₄ , ε _l are constant control parameters. 6. a kind of aircraft sideslip and roll compound turning control method based on attitude measurement according to claim 5, according to the roll angle signal that YIN600-R inertial integrated navigation system measures and obtains, compare with described roll angle expectation signal, The roll angle error signal is obtained, and then a nonlinear high-pass differential corrector is designed to obtain the roll angle error filtered differential signal, and then the error nonlinear integral signal is superimposed to form a roll channel control signal including: e _γ = γ-γ ^d ;

D ₃ =d ₅ (e _γ (n+1)-e _γ (n))+d ₆ sin(e _γ (n+1)-e _γ (n));

D ₄ (n+1)=D ₄ (n)+D _γ *T _a1 ;

Where γ is the roll angle measurement signal, e _γ is the roll angle error signal, ε _γ is a constant parameter, dt is the integral of the time signal; s ₃ is the roll angle error nonlinear integral signal, D _γ is the roll angle error differential Correction signal, T _a1 is the time interval of data sampling, d ₅ , d ₆ , d ₇ , d ₈ , ε _k2 are constant parameters; u _g is the roll channel control signal, l ₅ ,l ₆ ,l ₇ ,l ₈ , ε _g is a constant control parameter; finally, the obtained yaw channel control signal u _p is sent to the yaw rudder system to track the expected value of the yaw angle; the obtained roll channel control value u _g is sent to The roll rudder system is used to track the expected value of the roll angle; thus, under the joint action of the roll channel and the yaw channel, the aircraft sideslip and roll coordinated turn is realized.

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