CN112027117A - 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 PDF

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CN112027117A
CN112027117A CN202010948664.6A CN202010948664A CN112027117A CN 112027117 A CN112027117 A CN 112027117A CN 202010948664 A CN202010948664 A CN 202010948664A CN 112027117 A CN112027117 A CN 112027117A
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roll
lateral
angle
yaw
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CN112027117B (en
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雷军委
李恒
王瑞奇
王玲玲
李辉
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Naval Aeronautical University
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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. 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, then obtaining an expected signal of a yaw angle and a roll angle through symmetrical design and combination, then obtaining a final control signal 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 through 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

Aircraft sideslip and roll composite turning control method based on attitude measurement
Technical Field
The invention relates to the field of aircraft stabilization and turning control, in particular to an aircraft sideslip and roll 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 that a rolling channel is basically fixed, and the power for turning is provided by utilizing a 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 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 the lateral position instruction signal to obtain a lateral position error signal;
step S30, designing a non-linear 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 the system;
step S50, according to the lateral position error signal, the lateral position filtering integral signal and the filtering speed signal, linear combination and superposition are carried out, and a yaw angle expected signal and a roll angle expected signal are respectively designed;
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 integrated 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 the 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.
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 velocity signal and a lateral position signal, and the lateral velocity signal and the lateral position signal are compared with a lateral position command signal to obtain a lateral position error signal;
vz＝∫azdt；
z＝∫vzdt；
ez＝z-zd；
wherein a iszFor measuring the lateral acceleration of the aircraft using the YIN600-R inertial integrated navigation system, az(n) represents the lateral acceleration data at time T ═ n × Δ T, where n is 1,2,3 … and Δ T is the data sampling period. v. ofzFor lateral velocity signals, dt represents the integration of the time signal. z is a lateral position signal, zdSetting a lateral desired position signal for a lateral mission of the aircraft, ezIs 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:
Figure BDA0002676159120000031
ez1(n+1)＝ez1(n)+Da*Ta；
s1＝∫ez1dt；
wherein ezFor said lateral position error signal, ez1For low-pass correction of the signal, TaTime interval, k, representing data1、k2And1the detailed design of the parameter is described in the following examples. dt represents the integral of the time signal, s1The integrated signal is filtered for the final lateral position.
In an exemplary embodiment of the present invention, designing a non-linear low-pass filtering corrector according to the lateral velocity signal, and obtaining a filtered velocity signal includes:
Figure BDA0002676159120000032
vz1(n+1)＝vz1(n)+Db*Ta；
wherein v iszFor said lateral velocity signal, vz1For filtering the speed signal, k3、k2And1the detailed design of the parameter is described in the following examples.
In an exemplary embodiment of the present invention, the linear combination and superposition according to the lateral position error signal, the lateral position filtered integrated signal and the filtered velocity signal, and the designing the yaw angle desired signal and the roll angle desired signal respectively comprises:
Figure BDA0002676159120000041
Figure BDA0002676159120000042
wherein ezFor said lateral position error signal, s1Filtering the integrated signal for lateral position vz1For filtering velocity signals, #dFor yaw angle desired signal, gammadAs roll angle desired signal, c11、c12、c13、c14、11、c21、c22、c23、c24、21The 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:
eψ＝ψ-ψd；
Figure BDA0002676159120000043
D1＝d1(eψ(n+1)-eψ(n))+d2 sin(eψ(n+1)-eψ(n))；
Figure BDA0002676159120000044
D2(n+1)＝D2(n)+Dψ*Ta1；
Figure BDA0002676159120000045
where psi is the yaw angle measurement signal, eψIn order to be a yaw angle error signal,ψis a constant parameter, the detailed design of which is described in the examples hereinafter, dt representing the integral over time. DψFor differential correction of yaw angle errors, Ta1Time intervals for data sampling, d1、d2、d3、d4、k1The parameters are commonly used and are selected in detail later. u. ofpFor yaw channel control signals,/1,l2,l3,l4,lThe 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:
eγ＝γ-γd；
Figure BDA0002676159120000051
D3＝d5(eγ(n+1)-eγ(n))+d6 sin(eγ(n+1)-eγ(n))；
Figure BDA0002676159120000052
D4(n+1)＝D4(n)+Dγ*Ta1；
Figure BDA0002676159120000053
where gamma is the roll angle measurement signal, eγIn order to provide a roll angle error signal,γis a constant parameter, the detailed design of which is described in the examples hereinafter, dt representing the integral over time. s3As a non-linear integral signal of the roll angle error, DγFor differential correction of roll angle errors, Ta1Time intervals for data sampling, d5、d6、d7、d8、k2The parameters are commonly used and are selected in detail later. u. ofgFor the roll channel control signal,/5,l6,l7,l8,gThe detailed design of the control parameter is described in the following examples.
Finally, the obtained yaw channel control quantity uhAnd the tracking angle 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 channelgAnd 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.
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. Particularly, 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 a method for controlling an aircraft sideslip and roll composite turn based on attitude measurement according to the present 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 plot of the lateral acceleration of the aircraft (in meters per second) for 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 plot of aircraft lateral velocity (in meters per second) for 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 (without units) of the filtered integrated signal for the lateral position of the aircraft 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 in accordance with 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 may 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 and roll composite turning based on the attitude measurement comprises the following steps:
and step 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: 150g of weight, 59 × 45 × 23.8mm of size, acceleration measurement range of-16 g to 16g, attitude accuracy of 0.1 degree, bandwidth of 80Hz, power consumption of 1150mW, output rate of 200Hz, speed accuracy of 0.05 meter per second and positioning accuracy of 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 ℃.
Next, the YIN600-R integrated inertial navigation system is used to measure the yaw angle of the aircraft, and ψ (n) is calculated, where n is 1,2,3 … and Δ T is the data sampling period, and the detailed design thereof is described in the following example. Simultaneously, the inertial navigation system is adopted to measure the lateral acceleration of the aircraft and is counted as az，azAnd (n) represents the data of the lateral acceleration at the time T, n, Δ T, wherein n is 1,2,3 …, Δ T is the data sampling period, and the detailed design of the data can be selected to be the same as the yaw angle measurement.
Finally, the YIN600-R integrated inertial navigation system is used to measure the roll angle of the aircraft, and γ (n) represents the data of the roll angle at the time T n Δ T, where n is 1,2,3 … and Δ T is the data sampling period.
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 the lateral position instruction signal to obtain a lateral position error signal;
specifically, firstly, according to the lateral acceleration measurement signal azIntegrating to obtain lateral speed signal, and counting as vzThe integration method is as follows:
vz＝∫azdt；
where dt represents the integration of the time signal.
Thirdly, a lateral velocity measurement signal a is measuredzAnd performing linear integration to obtain a lateral position signal, and calculating the lateral position signal as z, wherein the integration mode is as follows:
z＝∫vzdt；
where dt represents the integration of the time signal.
Finally, a lateral desired position signal, denoted z, is set according to the lateral mission of the aircraftd. Then comparing with the lateral position signal to obtain a lateral position error signal, and recording the lateral position error signal as ezThe comparison is as follows:
ez＝z-zd；
and 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.
In particular, first the lateral position error signal e is addressedzThe following nonlinear low-pass correction is performed to obtain a low-pass correction signal counted as ez1The calculation method is as follows:
Figure BDA0002676159120000091
ez1(n+1)＝ez1(n)+Da*Ta；
wherein T isaTime interval, k, representing data1、k2And1the 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 s1The calculation method is as follows:
s1＝∫ez1dt
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 the system;
in particular, for said lateral velocity signal vzThe filtered velocity signal, denoted v, is obtained by nonlinear low-pass filtering correctionz1The calculation method is as follows:
Figure BDA0002676159120000101
vz1(n+1)＝vz1(n)+Db*Ta；
wherein k is3、k2And1the detailed design of the parameter is described in the following examples.
And step S50, performing 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, the error signal e is first determined according to the lateral positionzFiltering the integrated signal s in the lateral position1And the filtered velocity signal vz1Linearly combining and superposing to design yaw angle expected signal written as psidThe calculation method is as follows:
Figure BDA0002676159120000102
wherein c is11、c12、c13、c14、11The detailed design of the parameter is described in the following examples.
Secondly, a symmetrical mode is adopted, and the error signal e of the lateral position is obtainedzFiltering the integrated signal s in the lateral position1And the filtered velocity signal vz1Linear combination and superposition are carried out, and a roll angle expected signal is designed and recorded as gammadThe calculation method is as follows:
Figure BDA0002676159120000103
wherein c is21、c22、c23、c24、21The detailed design of the parameter is described in the following examples.
And step S60, comparing a yaw angle signal obtained by measurement of 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 eψThe comparison is as follows:
eψ＝ψ-ψd；
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 s2The integration method is as follows:
Figure BDA0002676159120000111
whereinψIs a constant parameter, the detailed design of which is shown belowIn the context of the examples, dt represents the integral of the time signal.
Then, a nonlinear high-pass differential corrector is constructed according to the yaw angle error signal to obtain a differential correction signal of the yaw angle error, and the differential correction signal is counted as DψThe calculation method is as follows:
D1＝d1(eψ(n+1)-eψ(n))+d2 sin(eψ(n+1)-eψ(n))；
Figure BDA0002676159120000112
D2(n+1)＝D2(n)+Dψ*Ta1；
wherein T isa1Time intervals for data sampling, d1、d2、d3、d4、k1The parameters are commonly used and are selected in detail later.
Finally, for said aircraft yaw angle error signal eψAnd a non-linear integral signal s of the yaw angle error2Differential correction signal D for yaw angle errorψLinear combination is carried out to obtain the final yaw channel control signal which is recorded as upThe calculation method is as follows:
Figure BDA0002676159120000113
wherein l1,l2,l3,l4,lThe detailed design of the control parameter is described in the following examples.
Finally, the obtained yaw channel control quantity uhAnd 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 integrated 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 the 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 eγThe comparison is as follows:
eγ＝γ-γd；
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 s3The integration method is as follows:
Figure BDA0002676159120000121
whereinγIs a constant parameter, the detailed design of which is described in the examples hereinafter, dt representing the integral over time. Then, a nonlinear high-pass differential corrector is constructed according to the rolling angle error signal to obtain a differential correction signal of the rolling angle error, and the differential correction signal is counted as DγThe calculation method is as follows:
D3＝d5(eγ(n+1)-eγ(n))+d6 sin(eγ(n+1)-eγ(n))；
Figure BDA0002676159120000122
D4(n+1)＝D4(n)+Dγ*Ta1；
wherein T isa1Time intervals for data sampling, d5、d6、d7、d8、k2The parameters are commonly used and are selected in detail later.
Finally, aiming at the aircraft roll angle error signal eγAnd a roll angle error nonlinear integral signal s3Differential correction signal D of roll angle errorγAdvancing lineThe final rolling channel control signal is obtained by sexual combination and is recorded as ugThe calculation method is as follows:
Figure BDA0002676159120000123
wherein l5,l6,l7,l8,gThe detailed design of the control parameter is described in the following examples.
Finally, the obtained rolling channel control quantity ugAnd 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, the YIN600-R integrated inertial 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 to obtain a lateral velocity signal and a lateral position signal respectively as shown in fig. 6 and 7 and a lateral position error signal as shown in fig. 8 according to the lateral acceleration signal measured by YIN600-R inertial integrated navigation system.
In step S30, T is selecteda＝0.001，k1＝10、k2Not equal to 5 and1the integrated signal of the lateral position filter is obtained as shown in fig. 9 at 0.2.
In step S40, k is selected3＝10、k 46 and1the filtered speed signal is obtained as shown in fig. 10, at 0.5.
In step S50, c is selected11＝0.1、c12＝0.05、c13＝0.1、c14＝3、11When the yaw angle desired signal is obtained, 0.8, c is selected as shown in fig. 1121＝0.2、c22＝0.1、c23＝0.2、c24＝8、210.4; roll angle desired signal is shown in FIG. 12。
In step S60, d is selected1＝1000、d2＝500、d3＝20、d4＝8、k1＝2、Ta1＝0.001，ψ＝0.7、l1＝2,l2＝0.2,l3＝0.4,l4＝5,lThe resulting yaw channel control signal is shown in fig. 13, 4.
In step S70, d is selected5＝1000、d6＝700、d7＝10、d8＝8、k2＝4、Ta1＝0.001、γ＝.07，l5＝2,l6＝0.1,l7＝0.3,l8＝7,gThe roll channel control signal is obtained as shown in fig. 14, 6. 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:
step 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 the lateral position instruction signal to obtain a lateral position error signal;
step S30, designing a non-linear 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 the system;
step S50, according to the lateral position error signal, the lateral position filtering integral signal and the filtering speed signal, linear combination and superposition are carried out, and a yaw angle expected signal and a roll angle expected signal are respectively designed;
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 integrated 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 the 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:
vz＝∫azdt；
z＝∫vzdt；
ez＝z-zd；
Figure RE-FDA0002728542930000021
ez1(n+1)＝ez1(n)+Da*Ta；
s1＝∫ez1dt；
wherein a iszFor measuring the lateral acceleration of the aircraft using the YIN600-R inertial integrated navigation system, az(n) represents the lateral acceleration data at time T ═ n × Δ T, where n is 1,2,3 … and Δ T is the data sampling period. v. ofzFor lateral velocity signals, dt represents the integration of the time signal. z is a lateral position signal, zdSetting a lateral desired position signal for a lateral mission of the aircraft, ezIs a lateral position error signal. e.g. of the typezFor said lateral position error signal, ez1For low-pass correction of the signal, TaTime interval, k, representing data1、k2And1the detailed design of the parameter is described in the following examples. dt represents the integral of the time signal, s1The integrated signal is filtered for the final lateral position.
3. The method of claim 1, wherein the step of designing a non-linear low-pass filter corrector according to the lateral velocity signal to obtain a filtered velocity signal comprises:
Figure RE-FDA0002728542930000022
vz1(n+1)＝vz1(n)+Db*Ta；
wherein v iszFor said lateral velocity signal, vz1For filtering the speed signal, k3、k2And1is a constant parameter.
4. The method of claim 1, wherein the step of designing the yaw angle desired signal and the roll angle desired signal respectively comprises the steps of performing linear combination and superposition on the lateral position error signal, the lateral position filtered integrated signal and the filtered velocity signal according to the attitude measurement, and the step of designing the yaw angle desired signal and the roll angle desired signal respectively comprises the steps of:
Figure RE-FDA0002728542930000031
Figure RE-FDA0002728542930000032
wherein ezFor said lateral position error signal, s1Filtering the integrated signal for lateral position vz1For filtering velocity signals, #dFor yaw angle desired signal, gammadAs roll angle desired signal, c11、c12、c13、c14、11、c21、c22、c23、c24、21Is a constant parameter.
5. The method of claim 1, wherein 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 yaw angle signal, and superimposing an error non-linear integral signal to form a yaw channel control signal comprises:
eψ＝ψ-ψd；
Figure RE-FDA0002728542930000033
D1＝d1(eψ(n+1)-eψ(n))+d2 sin(eψ(n+1)-eψ(n))；
Figure RE-FDA0002728542930000034
D2(n+1)＝D2(n)+Dψ*Ta1；
Figure RE-FDA0002728542930000035
where psi is the yaw angle measurement signal, eψIn order to be a yaw angle error signal,ψdt represents the integral of the time signal, a constant parameter. DψFor differential correction of yaw angle errors, Ta1Time intervals for data sampling, d1、d2、d3、d4、k1Is a constant parameter. u. ofpFor yaw channel control signals,/1,l2,l3,l4,lThe parameter is controlled to be constant.
6. The method of claim 1, 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 filtered differential signal, and the step of superimposing an error non-linear integral signal to form a roll channel control signal comprises:
eγ＝γ-γd；
Figure RE-FDA0002728542930000041
D3＝d5(eγ(n+1)-eγ(n))+d6 sin(eγ(n+1)-eγ(n))；
Figure RE-FDA0002728542930000042
D4(n+1)＝D4(n)+Dγ*Ta1；
Figure RE-FDA0002728542930000043
where gamma is the roll angle measurement signal, eγIn order to provide a roll angle error signal,γdt represents the integral of the time signal, a constant parameter. s3As a non-linear integral signal of the roll angle error, DγFor differential correction of roll angle errors, Ta1Time intervals for data sampling, d5、d6、d7、d8、k2Is a constant parameter. u. ofgFor the roll channel control signal,/5,l6,l7,l8,gThe parameter is controlled to be constant. Finally, the obtained yaw channel control quantity uhAnd the tracking angle 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 channelgAnd 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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