CN112027117B - Aircraft sideslip and roll composite turning control method based on attitude measurement - Google Patents
Aircraft sideslip and roll composite turning control method based on attitude measurement Download PDFInfo
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Abstract
The invention relates to an aircraft sideslip and roll composite turning control method based on attitude measurement. A YIN600-R inertial integrated navigation system is arranged to measure a yaw angle, a roll angle and a lateral acceleration, and then a lateral speed and a position signal are obtained through integration. And then designing a nonlinear low-pass integral corrector to obtain a position error integral signal, designing a nonlinear low-pass filter corrector to obtain a speed filter signal, symmetrically designing and combining to obtain expected signals of a yaw angle and a roll angle, then obtaining final control signals of a yaw channel and a roll channel through the error of an attitude angle, high-pass differential correction and nonlinear integration, and realizing the cooperative integrated lateral turning control of the aircraft by the simultaneous action of the two channels. The method has the advantages that the integrated control of the sideslip and the rolling ensures that the aircraft has good rapidity in lateral turning, and meanwhile, the design of symmetry ensures that the aircraft has good stability.
Description
Technical Field
The invention relates to the field of aircraft stabilization and turning control, in particular to an aircraft sideslip rolling composite turning control method based on attitude measurement.
Background
The main current methods for the lateral mass center motion of an aircraft include sideslip turning and bank turning, wherein the bank turning is also called rolling turning, and the change of the lift direction in the rolling process of the aircraft is utilized to provide centripetal force, so that the turning process of the aircraft is faster than that of the traditional sideslip turning. However, the traditional sideslip turning idea is to basically make the rolling channel still and provide turning power by utilizing the sideslip angle; banked turns to eliminate coupling of the yaw path are typically in a steady state in the yaw path. Therefore, in order to ensure the stability of the system, the two methods completely adopt the design concept of two-channel flying away and do not utilize the nonlinear coupling effect between yaw and sideslip. Based on the background reasons, the aircraft turning control method which adopts attitude measurement and attitude stabilization as the foundation stone and realizes the cooperative integration of sideslip and roll through the coordinated control of the yaw angle and the roll angle realizes good rapidity of the turning process and has high engineering application value.
It is to be noted that the information invented in the above background section is only for enhancing 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.
Disclosure of Invention
The invention aims to provide an aircraft sideslip and roll composite turning control method based on attitude measurement, and further solves the problem that the yaw and roll channels of the aircraft cannot be matched and controlled in a coordinated mode due to the limitations and defects of the related technology at least to a certain extent.
According to one aspect of the invention, an aircraft sideslip and roll composite turning control method based on attitude measurement is provided, and comprises the following steps:
step S10, installing a YIN600-R inertial integrated navigation system on an aircraft, and measuring the lateral acceleration, the yaw angle and the roll angle of the aircraft;
step S20, according to a lateral acceleration signal measured by a YIN600-R inertial integrated navigation system, performing twice integration to respectively obtain a lateral speed signal and a lateral position signal, and comparing the lateral speed signal and the lateral position signal with a lateral position instruction signal to obtain a lateral position error signal;
step S30, designing a nonlinear low-pass integral corrector according to the lateral position error signal and carrying out linear integration to obtain a lateral position filtering integral signal;
step S40, designing a nonlinear low-pass filtering corrector according to the lateral velocity signal to obtain a filtering velocity signal and provide a damping signal for a system;
s50, carrying out linear combination and superposition according to the lateral position error signal, the lateral position filtering integral signal and the filtering speed signal, and respectively designing a yaw angle expected signal and a roll angle expected signal;
step S60, comparing a yaw angle signal measured by the YIN600-R inertial integrated navigation system with the yaw angle expected signal to obtain a yaw angle error signal, designing a nonlinear high-pass differential corrector to obtain a yaw angle error filtering differential signal, and superposing an error nonlinear integral signal to form a yaw channel control signal;
and S70, comparing a rolling angle signal obtained by measuring according to the YIN600-R inertial integrated navigation system with the expected rolling angle signal to obtain a rolling angle error signal, designing a nonlinear high-pass differential corrector to obtain a rolling angle error filtering differential signal, superposing an error nonlinear integral signal to form a rolling channel control signal, and controlling the rolling channel control signal and a yaw channel simultaneously to realize sideslip rolling cooperative turning of the aircraft.
In an exemplary embodiment of the invention, a YIN600-R inertial integrated navigation system is installed on an aircraft to measure the yaw angle, the roll angle and the lateral acceleration of the aircraft, then the yaw angle of the aircraft is measured, the two times of integration are carried out to respectively obtain a lateral speed signal and a lateral position signal, and the lateral speed signal and the lateral position signal are compared with a lateral position command signal to obtain a lateral position error signal;
v z ＝∫a z dt；
z＝∫v z dt；
e z ＝z-z d ；
wherein a is z For measuring the lateral acceleration of the aircraft using the YIN600-R inertial integrated navigation system, a z (n) represents the lateral acceleration data at time T = n x Δ T, where n =1,2,3 \ 8230, Δ T is the data sampling period. v. of z For lateral velocity signals, dt represents the integration of the time signal. z is a lateral position signal, z d Setting a lateral desired position signal for a lateral mission of the aircraft, e z Is a lateral position error signal.
In an exemplary embodiment of the present invention, designing a non-linear low-pass integral corrector according to the lateral position error signal and performing linear integration to obtain a lateral position filtering integral signal includes:
e z1 (n+1)＝e z1 (n)+D a *T a ；
s 1 ＝∫e z1 dt；
wherein e z For said lateral position error signal, e z1 For low-pass correction of the signal, T a Time interval, k, representing data 1 、k 2 And epsilon 1 The detailed design of the parameter is described in the following examples. dt represents the integral of the time signal, s 1 The integrated signal is filtered for the final lateral position.
In an exemplary embodiment of the present invention, designing a non-linear low-pass filter corrector according to the lateral velocity signal, and obtaining a filtered velocity signal includes:
v z1 (n+1)＝v z1 (n)+D b *T a ；
wherein v is z For said lateral velocity signal, v z1 For filtering the speed signal, k 3 、k 4 And e 1 The detailed design of the parameter is described in the following examples.
In an exemplary embodiment of the present invention, the linearly combining and superimposing according to the lateral position error signal, the lateral position filtered integrated signal and the filtered velocity signal, and the separately designing the yaw angle desired signal and the roll angle desired signal comprises:
wherein e z For said lateral position error signal, s 1 Filtering the integrated signal for lateral position v z1 For filtering velocity signals, # d For yaw angle desired signal, gamma d As roll angle desired signal, c 11 、c 12 、c 13 、 c 14 、ε 11 、c 21 、c 22 、c 23 、c 24 、ε 21 The detailed design of the parameter is described in the following examples.
In an exemplary embodiment of the present invention, comparing a yaw angle signal measured by a YIN600-R inertial integrated navigation system with the desired yaw angle signal to obtain a yaw angle error signal, designing a non-linear high-pass differential corrector to obtain a filtered differential signal of the yaw angle error, and superimposing an error non-linear integral signal to form a yaw channel control signal includes:
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 a yaw angle measurement signal, e ψ Is the yaw angle error signal, e ψ The detailed design of the constant value parameter is described in the examples, and dt represents the integral of the time signal. D ψ As a yaw angle errorOf the differential correction signal, T a1 Time intervals for data sampling, d 1 、d 2 、d 3 、d 4 、ε k1 The parameters are normal parameters, and the detailed selection is implemented later. u. of p For yaw channel control signals,/ 1 ,l 2 ,l 3 ,l 4 ,ε l The detailed design of the control parameter is described in the following examples.
In an exemplary embodiment of the present invention, comparing a roll angle signal measured by a YIN600-R inertial integrated navigation system with the desired roll angle signal to obtain a roll angle error signal, designing a non-linear high-pass differential corrector to obtain a roll angle error filtering differential signal, and then superimposing an error non-linear integral signal to form a roll channel control signal, the method includes:
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 ；
wherein gamma is a roll angle measurement signal, e γ Is the roll angle error signal, ε γ Is a constant parameter, the detailed design of which is described in the examples hereinafter, dt representing the integral over time. s is 3 As a non-linear integral signal of the roll angle error, D γ For differential correction of roll angle errors, T a1 Time interval for data samplingPartition, d 5 、d 6 、d 7 、d 8 、ε k2 The parameters are commonly used and are selected in detail later. u. of g For the roll channel control signal,/ 5 ,l 6 ,l 7 ,l 8 ,ε g The detailed design of the control parameter is described in the following examples.
Finally, the obtained yaw channel control quantity u p And the tracking angle is transmitted to a yaw rudder system to realize the tracking of the yaw angle expected value. The obtained controlled quantity u of the rolling channel g And the expected value of the roll angle is tracked by transmitting the expected value to a roll rudder system. Therefore, under the combined action of the rolling channel and the yawing channel, the sideslip and rolling coordinated turning of the aircraft is realized.
Advantageous effects
The invention provides an aircraft sideslip and roll composite turning control method based on attitude measurement. The aircraft can carry out sideslip movement while rolling, so that the turning rapidity is effectively improved, and the maneuverability of the aircraft is enhanced. Especially, by the design of the symmetrical yaw angle and roll angle expected signals, the stability of the whole aircraft in the turning process is ensured by 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, as claimed.
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. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 is a flow chart of an aircraft sideslip roll composite turning control method based on attitude measurement provided by the invention;
FIG. 2 is a diagram of a YIN600-R inertial integrated navigation system embodying the method of the present invention;
FIG. 3 is a graph of the lateral acceleration of an aircraft (in meters per second squared) using a method provided by an embodiment of the present invention;
FIG. 4 is a plot of aircraft yaw angle (in degrees) according to a method provided by an embodiment of the present invention;
FIG. 5 is a graph of aircraft roll angle (in degrees) for a method provided by an embodiment of the invention;
FIG. 6 is a graph of the lateral velocity of an aircraft (in meters per second) using a method provided by an embodiment of the present invention;
FIG. 7 is a plot of aircraft lateral position (in meters) for a method provided by an embodiment of the present invention;
FIG. 8 is a plot of aircraft lateral position error (in meters) for a method provided by an embodiment of the present invention;
FIG. 9 is a plot (in units) of the integrated filtered aircraft lateral position signal in accordance with a method provided by an embodiment of the present invention;
FIG. 10 is a plot (in units) of an aircraft filtered velocity signal for a method provided by an embodiment of the present invention;
FIG. 11 is a plot of yaw angle desired signal (in degrees) for a method provided by an embodiment of the present invention;
FIG. 12 is a roll angle desired signal curve (in degrees) for a method provided by an embodiment of the present invention;
FIG. 13 is a yaw path control signal plot (without units) of a method provided by an embodiment of the present invention;
fig. 14 is a roll channel control signal plot (without units) for a method provided by an embodiment of the present invention.
FIG. 15 is a graph of a sideslip angle signal (in degrees) in accordance with a method provided by an embodiment of the present invention.
Detailed Description
Example embodiments will now be described more fully with reference to the accompanying drawings. Example embodiments may, however, be embodied in many different 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 will 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 to provide a thorough understanding of embodiments of the invention. One skilled in the relevant art will recognize, however, that the invention can be practiced without one or more of the specific details, or with other methods, components, devices, steps, and so forth. 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 composite turning control method based on attitude measurement. And simultaneously, the symmetrical attitude command signal and the attitude stabilizing system are designed, the stability of the two channels is ensured, and the rapidity and the stability of the method can be improved and improved.
The method for controlling the aircraft sideslip and roll composite turning based on the attitude measurement according to the invention will be further explained and explained with reference to the attached drawings. Referring to fig. 1, the method for controlling the aircraft sideslip roll composite turning based on the attitude measurement comprises the following steps:
and S10, installing a YIN600-R inertial integrated navigation system on the aircraft, and measuring the lateral acceleration, the yaw angle and the roll angle of the aircraft.
Specifically, a YIN600-R inertial integrated navigation system is installed on an aircraft, a physical picture of the system is shown in fig. 2, and performance indexes of the system are as follows: weight 150g, size 59 x 45 x 23.8mm, acceleration measurement range-16 g to 16g, attitude accuracy 0.1 degree, bandwidth 80Hz, power consumption 1150mW, output rate 200Hz, speed accuracy 0.05 meters per second, and positioning accuracy 0.1 meter. The output mode is that RS232, RS485 and RS422 level interfaces are selectable in standard, and the working temperature is-40- +85 ℃.
Then, the aircraft yaw angle is measured by YIN600-R inertial integrated navigation system, and psi (n) is calculated and represents the data of the yaw angle at the time T = n × Δ T, wherein n =1,2,3 \8230;, Δ T is the data sampling period, and the detailed design thereof is shown in the examples later. And simultaneously measuring the lateral acceleration of the aircraft by adopting the inertial navigation system, and counting as az, wherein az (n) represents data of the lateral acceleration at the time T = n × Δ T, wherein n =1,2,3 \8230, Δ T is a data sampling period, and the detailed design of the system can be selected to be the same as that of yaw angle measurement.
Finally, the YIN600-R inertial integrated navigation system is adopted to measure the roll angle of the aircraft, and the gamma (n) represents the data of the roll angle at the time T = n × Δ T, wherein n =1,2,3 \8230, Δ T is the data sampling period, and the detailed design thereof is shown in the embodiment of the later case.
Step S20, according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system, performing twice integration to respectively obtain a lateral speed signal and a lateral position signal, and comparing the lateral speed signal and the lateral position signal with a lateral position instruction signal to obtain a lateral position error signal;
specifically, firstly, according to the lateral acceleration measurement signal a z Integrating to obtain lateral speed signal, and counting as v z The integration method is as follows:
v z ＝∫a z dt；
where dt represents the integration of the time signal.
Thirdly, a lateral velocity measurement signal a is measured z Linear integration is performed to obtain a lateral position signal, which is measured as zThe integration is as follows:
z＝∫v z dt；
where dt represents the integration of the time signal.
Finally, a lateral desired position signal is set according to the lateral mission of the aircraft, denoted as z d . Then comparing it with the said lateral position signal to obtain lateral position error signal, and recording it as e z The comparison is as follows:
e z ＝z-z d ；
and S30, designing a nonlinear low-pass integral corrector according to the lateral position error signal and carrying out linear integration to obtain a lateral position filtering integral signal.
In particular, first the lateral position error signal e is addressed z The following nonlinear low-pass correction is performed to obtain a low-pass correction signal counted as e z1 The calculation method is as follows:
e z1 (n+1)＝e z1 (n)+D a *T a ；
wherein T is a Time interval, k, representing data 1 、k 2 And e 1 The detailed design of the parameter is described in the following examples.
Secondly, the low-pass correction signal is integrated to obtain a final lateral position filtering integral signal which is recorded as s 1 The calculation method is as follows:
s 1 ＝∫e z1 dt
where dt represents the integral of the time signal.
Step S40, designing a nonlinear low-pass filtering corrector according to the lateral velocity signal to obtain a filtering velocity signal and provide a damping signal for a system;
in particular, for said lateral velocity signal v z Carrying out nonlinear low-pass filtering correction to obtain filtering speed signalNumber, denoted by v z1 The calculation method is as follows:
v z1 (n+1)＝v z1 (n)+D b *T a ；
wherein k is 3 、k 2 And epsilon 1 The detailed design of the parameter is described in the following examples.
And S50, carrying out linear combination and superposition according to the lateral position error signal, the lateral position filtering integral signal and the filtering speed signal, and respectively designing a yaw angle expected signal and a roll angle expected signal.
Specifically, first, the error signal e is determined according to the lateral position z Filtering the integrated signal s in the lateral position 1 And the filtered velocity signal v z1 Linearly combining and superposing to design yaw angle expected signal written as psi d The calculation method is as follows:
wherein c is 11 、c 12 、c 13 、c 14 、ε 11 The detailed design of the parameter is described in the following examples.
Secondly, adopting a symmetrical mode according to the lateral position error signal e z Filtering the integrated signal s in the lateral direction 1 And the filtered velocity signal v z1 Linear combination and superposition are carried out, and a roll angle expected signal is designed and recorded as gamma d The calculation method is as follows:
wherein c is 21 、c 22 、c 23 、c 24 、ε 21 Is a constant parameter, which is detailedDetailed design is described in the following example implementation.
And S60, comparing a yaw angle signal measured by the YIN600-R inertial integrated navigation system with the yaw angle expected signal to obtain a yaw angle error signal, designing a nonlinear high-pass differential corrector to obtain a yaw angle error filtering differential signal, and superposing an error nonlinear integral signal to form a yaw channel control signal.
Specifically, the yaw angle measurement signal is compared with the yaw angle expected signal to obtain a yaw angle error signal, which is denoted as e ψ The comparison is as follows:
e ψ ＝ψ-ψ d ；
secondly, according to the yaw angle error signal, carrying out nonlinear linear integration to obtain a yaw angle error nonlinear integral signal which is recorded as s 2 The integration method is as follows:
wherein epsilon ψ The detailed design of the constant value parameter is described in the examples, and dt represents the integral of the time signal.
Then, according to the yaw angle error signal, constructing a nonlinear high-pass differential corrector to obtain a differential correction signal of the yaw angle error, and calculating the differential correction signal as D ψ The 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 ；
wherein T is a1 Time interval for data sampling, d 1 、d 2 、d 3 、d 4 、ε k1 Is a constant valueThe parameters, the detailed selection of which is made later.
Finally, for said aircraft yaw angle error signal e ψ And a non-linear integral signal s of the yaw angle error 2 Differential correction signal D for yaw angle error ψ Linear combination is carried out to obtain the final yaw channel control signal which is recorded as u p The calculation method is as follows:
wherein l 1 ,l 2 ,l 3 ,l 4 ,ε l The detailed design of the control parameter is described in the following examples.
Finally, the obtained yaw channel control quantity u p And the tracking angle is transmitted to a yaw rudder system to realize the tracking of the yaw angle expected value.
And S70, comparing a roll angle signal obtained by measuring according to the YIN600-R inertia combined navigation system with the roll angle expected signal to obtain a roll angle error signal, designing a nonlinear high-pass differential corrector to obtain a roll angle error filtering differential signal, superposing an error nonlinear integral signal to form a roll channel control signal, and controlling the roll channel control signal and a yaw channel simultaneously to realize sideslip and roll cooperative turning of the aircraft.
Specifically, the roll angle measurement signal is compared with the roll angle expected signal to obtain a roll angle error signal, which is recorded as e γ The comparison is as follows:
e γ ＝γ-γ d ；
secondly, according to the rolling angle error signal, carrying out nonlinear linear integration to obtain a rolling angle error nonlinear integral signal which is recorded as s 3 The integration method is as follows:
wherein epsilon γ Is a constant value parameterDetailed design thereof is described in the examples hereinafter, dt represents the integral of the time signal. Then, constructing a nonlinear high-pass differential corrector according to the rolling angle error signal to obtain a differential correction signal of the rolling angle error, and calculating the differential correction signal as D γ The 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 ；
wherein T is a1 Time intervals for data sampling, d 5 、d 6 、d 7 、d 8 、ε k2 The parameters are normal parameters, and the detailed selection is implemented later.
Finally, aiming at the aircraft roll angle error signal e γ And a roll angle error nonlinear integral signal s 3 Differential correction signal D of roll angle error γ Linear combination is carried out to obtain the final control signal of the rolling channel, which is recorded as u g The calculation method is as follows:
wherein l 5 ,l 6 ,l 7 ,l 8 ,ε g The detailed design of the control parameters is described in the following embodiments.
Finally, the obtained controlled quantity u of the rolling channel g And the tracking angle is transmitted to a rolling rudder system to realize the tracking of the expected value of the rolling angle. Therefore, under the combined action of the rolling channel and the yawing channel, the sideslip and rolling coordinated turning of the aircraft is realized.
Case implementation and simulation experiment result analysis
In step S10, a YIN600-R inertial integrated navigation system is installed on the aircraft, the lateral acceleration of the aircraft is measured as shown in FIG. 3, the yaw angle is measured as shown in FIG. 4, and the roll angle is measured as shown in FIG. 5.
In step S20, two times of integration are performed according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system to obtain a lateral velocity signal and a lateral position signal as shown in fig. 6 and 7, respectively, and a lateral position error signal as shown in fig. 8.
In step S30, T is selected a ＝0.001，k 1 ＝10、k 2 =5 and ε 1 And =0.2, and the lateral position filtered integrated signal is obtained as shown in fig. 9.
In step S40, k is selected 3 ＝10、k 4 =6 and ε 1 =0.5, the filtered speed signal is obtained as shown in fig. 10.
In step S50, c is selected 11 ＝0.1、c 12 ＝0.05、c 13 ＝0.1、c 14 ＝3、ε 11 =0.8, get yaw angle desired signal as shown in fig. 11, select c 21 ＝0.2、c 22 ＝0.1、c 23 ＝0.2、c 24 ＝8、ε 21 =0.4; the roll angle desired signal is shown in fig. 12.
In step S60, d is selected 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, resulting yaw channel control signal is shown in fig. 13.
In step S70, d is selected 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, and the roll channel control signal is obtained as shown in fig. 14. The yaw channel and the yaw channel are simultaneously controlled to realize the sideslip and roll cooperative turning of the aircraft, and the sideslip angle during the turning process is shown in figure 15.
As can be seen from fig. 4, the yaw angle reaches 6 degrees at maximum, as can be seen from fig. 5, the roll angle reaches 20 degrees at maximum, and as can be seen from fig. 6, the lateral velocity reaches 14 meters per second at maximum, as can be seen from fig. 7, the aircraft completes a large-maneuvering turn in about 5 s. As can be seen from fig. 15, the sideslip angle reaches 4.2 degrees at most, and as can be seen from fig. 13 and 14, the yaw channel control signal does not exceed 8 degrees, and the roll channel control signal does not exceed 6 degrees, which means that both the yaw rudder and the roll rudder do not exceed 8 degrees, thereby meeting the restriction requirements of engineering application. The case shows that the turning strategy of rolling and sideslip integration enables the turning process of the aircraft to be very rapid and coupling to be more beneficial to finishing the turning action, so that the invention has good engineering practical value.
Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the invention following, in general, the principles of the invention and including such departures from the present disclosure as come within known or customary practice within the art to which the invention pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.
Claims (6)
1. An aircraft sideslip and roll composite turning control method based on attitude measurement is characterized by comprising the following steps:
s10, installing a YIN600-R inertial integrated navigation system on the aircraft, and measuring the lateral acceleration, the yaw angle and the roll angle of the aircraft;
step S20, according to the lateral acceleration signal measured by the YIN600-R inertial integrated navigation system, performing twice integration to respectively obtain a lateral speed signal and a lateral position signal, and comparing the lateral speed signal and the lateral position signal with a lateral position instruction signal to obtain a lateral position error signal;
step S30, designing a nonlinear low-pass integral corrector according to the lateral position error signal and carrying out linear integration to obtain a lateral position filtering integral signal;
step S40, designing a nonlinear low-pass filtering corrector according to the lateral velocity signal to obtain a filtering velocity signal and provide a damping signal for a system;
s50, carrying out linear combination and superposition according to the lateral position error signal, the lateral position filtering integral signal and the filtering speed signal, and respectively designing a yaw angle expected signal and a roll angle expected signal;
step S60, comparing a yaw angle signal obtained by measurement of a YIN600-R inertia integrated navigation system with the yaw angle expected signal to obtain a yaw angle error signal, designing a nonlinear high-pass differential corrector to obtain a yaw angle error filtering differential signal, and superposing an error nonlinear integral signal to form a yaw channel control signal;
and S70, comparing a roll angle signal obtained by measuring according to the YIN600-R inertia combined navigation system with the roll angle expected signal to obtain a roll angle error signal, designing a nonlinear high-pass differential corrector to obtain a roll angle error filtering differential signal, superposing an error nonlinear integral signal to form a roll channel control signal, and controlling the roll channel control signal and a yaw channel simultaneously to realize sideslip and roll cooperative turning of the aircraft.
2. The method of claim 1, wherein the designing a non-linear low-pass integral corrector and performing linear integration according to the lateral position error signal to obtain a lateral position filtering integral signal comprises:
v z ＝∫a z dt；
z＝∫v z dt；
e z ＝z-z d ；
e z1 (n+1)＝e z1 (n)+D a *T a ；
s 1 ＝∫e z1 dt；
wherein a is z For measuring the lateral acceleration of the aircraft using the YIN600-R inertial integrated navigation system, a z (n) data representing lateral acceleration at time T = n x Δ T, where n =1,2,3 \ 8230, Δ T being the data sampling period; v. of z For lateral velocity signals, dt represents the integration of the time signal; z is a lateral position signal, z d Setting a lateral expected position signal according to a lateral task of the aircraft, wherein ez is a lateral position error signal; e.g. of the type z For said lateral position error signal, e z1 For low-pass correction of the signal, T a Time interval, k, representing data 1 、k 2 And epsilon 1 Is a constant parameter; dt represents the integral of the time signal, s 1 The integrated signal is filtered for the final lateral position.
3. The method for controlling the aircraft side-slip and roll composite turn based on the attitude measurement according to claim 2, wherein the designing of the nonlinear low-pass filter corrector according to the lateral velocity signal to obtain the filtered velocity signal comprises:
v z1 (n+1)＝v z1 (n)+D b *T a ；
wherein v is z For said lateral velocity signal, v z1 For filtering the speed signal, k 3 、k 4 And e 2 Is a constant parameter.
4. The method of claim 3, wherein the step of designing the yaw angle desired signal and the roll angle desired signal separately comprises the steps of, according to the lateral position error signal, the lateral position filtered integrated signal and the filtered velocity signal, linearly combining and superimposing the signals:
wherein e z For said lateral position error signal, s 1 Filtering the integrated signal for lateral position v z1 For filtering velocity signals,/ d For yaw angle desired signal, gamma d As roll angle desired signal, c 11 、c 12 、c 13 、c 14 、ε 11 、c 21 、c 22 、c 23 、c 24 、ε 21 Is a constant parameter.
5. The method of claim 4, wherein the step of comparing a yaw angle signal measured by a YIN600-R inertial integrated navigation system with the desired yaw angle signal to obtain a yaw angle error signal, the step of designing a nonlinear high-pass differential corrector to obtain a filtered differential yaw angle signal, and the step of superimposing a nonlinear integral error signal to form a yaw channel control signal comprises the steps of:
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 psi is the yaw angle measurement signal, gamma d For roll angle desired signal, # d For yaw angle desired signal, e ψ For yaw angle error signal, e ψ Dt represents the integral of the time signal, being a constant parameter; d ψ For differential correction of yaw angle errors, s 2 For non-linearly integrated signals of yaw angle error, T a1 Time intervals for data sampling, d 1 、d 2 、d 3 、d 4 、ε k1 Is a constant parameter; u. of p For yaw channel control signals,/ 1 ,l 2 ,l 3 ,l 4 ,ε l The parameter is controlled to be constant.
6. The method of claim 5, wherein the step of obtaining a roll angle error signal by comparing a roll angle signal measured by a YIN600-R inertial integrated navigation system with the roll angle desired signal, the step of designing a non-linear high-pass differential corrector to obtain a roll angle error filtering differential signal, and the step of superposing an error non-linear integral signal to form a roll channel control signal comprises:
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 ；
wherein gamma is a roll angle measurement signal, e γ Is the roll angle error signal, ε γ Dt represents the integral of the time signal, being a constant parameter; s 3 As a non-linear integral signal of the roll angle error, D γ For differential correction of roll angle errors, T a1 Time intervals for data sampling, d 5 、d 6 、d 7 、d 8 、ε k2 Is a constant parameter; u. of g For the roll channel control signal,/ 5 ,l 6 ,l 7 ,l 8 ,ε g Constant control parameters; finally, the obtained yaw channel control signal u p The yaw angle expected value is transmitted to a yaw rudder system to realize the tracking of the yaw angle expected value; the obtained control quantity u of the rolling channel g The tracking angle is transmitted to a rolling rudder system to realize the tracking of the expected value of the rolling angle; therefore, under the combined action of the rolling channel and the yawing channel, the sideslip and rolling coordinated turning of the aircraft is realized.
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