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CN114137999A - Method and device for controlling transverse and lateral channels of fixed-wing aircraft - Google Patents

Method and device for controlling transverse and lateral channels of fixed-wing aircraft Download PDF

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Publication number
CN114137999A
CN114137999A CN202110792449.6A CN202110792449A CN114137999A CN 114137999 A CN114137999 A CN 114137999A CN 202110792449 A CN202110792449 A CN 202110792449A CN 114137999 A CN114137999 A CN 114137999A
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yaw
roll
channel
angle
wing aircraft
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张一癸
李其骏
蒋自成
唐国金
罗林燕
周哲
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Hunan Kunpeng Zhihui Unmanned Plane Technology Co ltd
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Hunan Kunpeng Zhihui Unmanned Plane Technology Co ltd
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    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft

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Abstract

Aiming at the problem of poor disturbance resistance in the control of the transverse and lateral channels of the existing fixed wing aircraft, the invention provides a method and a device for controlling the transverse and lateral channels of the fixed wing aircraft, according to a target track, establishing a fixed wing aircraft lateral navigation control law to obtain a target roll angle and a target yaw angle required by tracking the target track, then establishing a fixed wing aircraft lateral channel expansion state model containing unmodeled dynamic and external interference, selecting a proper tracking differentiator to carry out smooth processing on the target roll angle and the target yaw angle, designing a PID (proportion integration differentiation) controller by using the processed signals, designing an observer by using the control quantity after delay processing, and estimating the expansion state as the total disturbance of the system in real time, and disturbance compensation is carried out on the control quantity output by the controller to obtain the final control quantity, so that the control precision of the control system is improved, and the disturbance resistance capability is enhanced.

Description

Method and device for controlling transverse and lateral channels of fixed-wing aircraft
Technical Field
The invention belongs to the technical field of aircraft control, and particularly relates to a method and a device for controlling a transverse and lateral channel of a fixed-wing aircraft.
Background
The fixed-wing aircraft transverse and lateral channel control mainly means that according to a target waypoint (set) or a target path, a navigation control law gives out an expected roll angle and a desired yaw angle, and then the aircraft flies towards the target waypoint or along the target path by controlling the roll angle and the yaw angle.
When the fixed-wing aircraft executes a flight task, a target flight path is required to be accurately tracked, the motion of a transverse lateral channel of the fixed-wing aircraft has strong coupling, and various external interferences, such as wind interference, exist, so that the aircraft is difficult to accurately fly according to a preset track.
Disclosure of Invention
The invention provides a fixed wing aircraft transverse and lateral channel control method and device, aiming at the problems of low precision and poor interference rejection capability in the existing fixed wing aircraft transverse and lateral channel control. The invention can effectively improve the control precision and the disturbance resistance of the whole system and improve the robustness.
In order to solve the technical problems, the technical scheme provided by the invention is as follows:
the fixed-wing aircraft transverse and lateral channel control method comprises the following steps:
firstly, designing a lateral navigation control law according to a target track, and outputting the current target roll angle of the aircraft
Figure BDA0003161522280000011
And target yaw angle psit；
Decoupling the transverse and lateral channel model of the fixed-wing aircraft, and expressing unmodeled dynamics, parameter uncertainty and external disturbance by using an expansion state to establish an expansion state model of the transverse and lateral channel of the fixed-wing aircraft containing the total disturbance;
thirdly, the target roll angle is adjusted
Figure BDA0003161522280000012
And target yaw angle psitAdopting the steepest control synthesis function fhan to construct a tracking differentiator, carrying out smooth filtering processing on the signal, outputting a reference signal and a differential signal thereof, and converting the reference signal and the differential signal into a reference signalRoll angular velocity prefAnd a reference yaw rate rref；
The fourth step, with reference roll angle
Figure BDA0003161522280000021
Reference roll angular velocity prefTrue roll angle
Figure BDA0003161522280000022
The actual roll angular speed p is input, and a double closed-loop series PID controller of a roll channel is designed, wherein the angle control is an outer loop, and the angular speed control is an inner loop;
a fifth step of referencing the yaw angle psirefReference yaw rate rrefThe actual yaw angle psi and the actual yaw angle speed r are used as input, and a double closed-loop series PID controller of a yaw channel is designed;
and sixthly, designing second-order extended state observers of a roll channel and a yaw channel respectively, estimating total disturbance in the extended state model of the fixed wing aircraft lateral channel established in the second step by using the fed-back actual control quantity and the control quantity after delay processing of the actuator, and performing disturbance compensation on the output of the controller to obtain the final control quantity.
The invention provides computer equipment which comprises a memory and a processor, wherein the memory stores a computer program, and the processor realizes the steps of a fixed-wing aircraft transverse and lateral channel control method when executing the computer program.
The invention provides a storage medium having stored thereon a computer program which, when executed by a processor, carries out the steps of a fixed-wing aircraft lateral passage control method.
Compared with the prior art, the invention has the advantages that:
the invention relates to a fixed wing aircraft transverse channel control method based on an active disturbance rejection control technology, which comprises the steps of establishing a fixed wing aircraft transverse channel expansion state model containing unmodeled dynamic and external disturbance, performing smooth processing on a target instruction signal by using a tracking differentiator, designing a PID (proportion integration differentiation) controller by using the processed signal, designing an observer by using a control quantity after delay processing, estimating the expansion state as the total disturbance of a system in real time, performing disturbance compensation on the control quantity output by the controller to obtain a final control quantity, and enhancing the disturbance rejection capability of the control system.
The invention completes the control of the transverse and lateral channels of the fixed-wing unmanned aerial vehicle based on the active disturbance rejection control technology, realizes good dynamic performance of the controller without steady-state error under the condition of large disturbance, and improves the robustness and the control precision of the control system.
Drawings
FIG. 1 is a flow chart of an embodiment of the present invention;
FIG. 2 is a block diagram of a fixed wing aircraft lateral channel controller in accordance with an embodiment of the present invention;
FIG. 3 is a schematic diagram of a fixed-wing aircraft lateral navigation control law in accordance with an embodiment of the present invention;
fig. 4 is a graph of a sinusoidal signal tracked by a fastest control synthesis function with fhan configuration under a certain noise condition according to an embodiment of the present invention, where a fast factor is 4 and a filtering factor is 0.005;
fig. 5 is a graph of a sinusoidal signal tracked by a fastest control synthesis function with fhan configuration under a certain noise condition according to an embodiment of the present invention, where a fast factor is 2 and a filtering factor is 0.01;
fig. 6 is a schematic structural diagram of a dual closed-loop series PID controller according to an embodiment of the present invention.
Detailed Description
The invention is further described below with reference to the figures and the specific embodiments of the description.
Referring to fig. 1 and 2, a method for controlling a lateral channel of a fixed-wing aircraft provided in an embodiment of the present invention includes:
firstly, designing a lateral navigation control law according to a target track, and outputting the current target roll angle of the aircraft
Figure BDA0003161522280000031
And target yaw angle psit。
Referring to FIG. 3, the trajectory is shown as the target trajectory that the aircraft is expected to track. Selecting a reference point P on the target track, wherein the distance between the reference point P and the current position O point of the aircraft is L1Ground speed of the aircraft VnAnd vector
Figure BDA0003161522280000032
The included angle therebetween is eta.
Lateral acceleration a required to drive the aircraft to the reference point P along the circular dotted-line trajectorysComprises the following steps:
Figure BDA0003161522280000033
wherein, R is the radius of the circular broken line track.
Distance L1The included angle eta and the radius R have the following geometrical relationship:
L1＝2R sin η
namely, it is
Figure BDA0003161522280000034
Substituting the above formula into one can obtain:
Figure BDA0003161522280000041
assuming that the aircraft performs horizontal linear fixed-height uniform-speed flight, the lift force F generated by the wings is generated at the momentlWhich counteracts the aircraft gravity mg. When the aircraft turns, the rolling angle is generated
Figure BDA0003161522280000042
Lifting force FlWill provide a lateral component to cause the aircraft to generate a lateral acceleration, lateral acceleration asAngle of transverse rolling
Figure BDA0003161522280000043
The relationship of (1) is:
Figure BDA0003161522280000044
where g is the acceleration of gravity.
The simplification can be obtained:
Figure BDA0003161522280000045
further, the target roll angle can be obtained
Figure BDA0003161522280000046
The expression of (a) is:
Figure BDA0003161522280000047
target yaw angle psitThe acquisition of (a) provides for pointing the aircraft nose direction towards the reference point P, i.e.:
Figure BDA0003161522280000048
wherein,
Figure BDA0003161522280000049
representing the vector from point O to point P,
Figure BDA00031615222800000410
vector of fingers
Figure BDA00031615222800000411
And the included angle between the north-east plane coordinate system and the true north direction (the north is true when the east is deviated from the north).
Therefore, in the first step in an embodiment of the present invention, a fixed-wing aircraft lateral navigation control law is constructed according to a target track, and a target roll angle is obtained
Figure BDA00031615222800000412
And target yaw angle psitThe formula of (1) is:
Figure BDA00031615222800000413
wherein, VnIs the ground speed of the aircraft, g is the acceleration of gravity, L1The distance between the current position O point of the aircraft and a reference point P is shown, and eta is the ground speed V of the aircraftnAnd vector
Figure BDA00031615222800000414
The included angle between the two parts is included,
Figure BDA00031615222800000415
representing the vector from point O to point P,
Figure BDA00031615222800000416
vector of fingers
Figure BDA00031615222800000417
And the included angle between the north-east plane coordinate system and the true north direction (the north is true when the east is deviated from the north).
And secondly, decoupling the fixed wing aircraft transverse lateral channel model, and representing the unmodeled dynamics, the parameter uncertainty and the external disturbance by using the expansion state to establish the fixed wing aircraft transverse lateral channel expansion state model containing the total disturbance.
According to a kinetic equation of the aircraft rotating around the center of mass, a fixed-wing aircraft lateral dynamics model is established as follows:
Figure BDA0003161522280000051
wherein p, q, r respectively represent roll angular velocity, pitch angular velocity and yaw angular velocity,
Figure BDA0003161522280000052
is the differential of the roll angular velocity,
Figure BDA0003161522280000053
is the differential of yaw angular velocity, i is the roll moment, n is the yaw moment, Γ1,Γ2,Γ3,Γ4,Γ5,Γ6Is a coefficient related to the moment of inertia.
According to the stress analysis, a fixed wing aircraft lateral moment model is established as follows:
Figure BDA0003161522280000054
wherein l is roll moment, n is yaw moment, ρ is air density, and VaIs airspeed, S is wing area, b is wing span, beta is sideslip angle, p is roll angular velocity, r is yaw angular velocity, deltaaAmount of aileron rudder, deltarAs the amount of the rudder,
Figure BDA0003161522280000055
are all pneumatic coefficients.
Substituting the moment model into a dynamic model, and simplifying to obtain a fixed wing aircraft transverse lateral channel coupling model in an expansion state:
Figure BDA0003161522280000056
wherein f is1(p,r,w1) For the total disturbance, w, in the roll pass including unmodeled dynamics of the system, uncertainty of the parameters, and external disturbances1External disturbance of the rolling path, f2(p,r,w2) For the sum disturbance, w, including unmodeled dynamics of the system, uncertainty of the parameters, and external disturbances in the yaw path2The external disturbance of the yaw channel is shown, and A, B, C and D are model coefficients.
Order to
Figure BDA0003161522280000057
And P is reversible.
Then the fixed-wing aircraft lateral passage expansion state coupling model can be rewritten as:
Figure BDA0003161522280000061
defining a virtual control quantity U:
Figure BDA0003161522280000062
wherein u is1And u2Virtual control quantities of the roll channel and the yaw channel are provided respectively.
Knowing the virtual control quantity U, the actual aileron rudder quantity and rudder quantity can be derived by inverting P, i.e.:
Figure BDA0003161522280000063
therefore, the decoupling model of the lateral channel expansion state of the fixed-wing aircraft can be represented as follows:
Figure BDA0003161522280000064
further, considering the delay effect of the actuator, the expansion state model can be expressed as:
Figure BDA0003161522280000065
wherein u is1(t-τ1) Delay tau for rolling path1Virtual control quantity after time, u2(t-τ2) Delaying tau for the yaw path2The virtual control amount after the time.
Thirdly, the target roll angle is adjusted
Figure BDA0003161522280000066
And target yaw angle psitAdopting the steepest control comprehensive function fhan to construct a tracking differentiator, carrying out smooth filtering processing on the signal, outputting a reference signal and a differential signal thereof, and converting the reference signal and the differential signal into a reference roll angular velocity prefAnd a reference yaw rate rref。
In the third step in the embodiment of the present invention, a steepest control synthesis function fhan is used to construct a rolling channel tracking differentiator, as follows:
Figure BDA0003161522280000067
wherein,
Figure BDA0003161522280000068
in order to input the target roll angle,
Figure BDA0003161522280000069
for reference to the roll angle,
Figure BDA00031615222800000610
for reference to the differential signal of the roll angle,
Figure BDA00031615222800000611
the fast factor of the differentiator is tracked for the roll path,
Figure BDA00031615222800000612
and f, a filter factor of a rolling channel tracking differentiator, u is an intermediate variable, and fhan is a steepest control comprehensive function.
Referring to FIG. 4, a fast factor is used in one embodiment of the present invention
Figure BDA00031615222800000613
Is 4, a filter factor
Figure BDA00031615222800000614
The sinusoidal signal under certain noise conditions is tracked for a fhan function of 0.005.
Similarly, in the third step in the embodiment of the present invention, the steepest control synthesis function fhan is used to construct the yaw channel tracking differentiator, as follows:
Figure BDA0003161522280000071
wherein psitFor an input target yaw angle, psirefIn order to refer to the yaw angle,
Figure BDA0003161522280000072
as a differential signal with reference to yaw angle, rψFast factor, h, for tracking differentiators for yaw pathψAnd f, tracking a filter factor of a differentiator for a yaw channel, wherein u is an intermediate variable, and fhan is a steepest control comprehensive function.
Referring to FIG. 5, in one embodiment of the present invention, a fast factor r is usedψIs 2, the filter factor hψThe sinusoidal signal under certain noise conditions is tracked for a fhan function of 0.01.
Output reference roll angle of roll channel tracking differentiator
Figure BDA0003161522280000073
And its differential signal
Figure BDA0003161522280000074
Yaw channel tracking differentiator outputs a reference yaw angle psirefAnd its differential signal
Figure BDA0003161522280000075
Convert it into a reference roll angular velocity prefAnd a reference yaw rate rrefThe formula is as follows:
Figure BDA0003161522280000076
wherein,
Figure BDA0003161522280000077
for reference to the roll angle, θrefFor reference to the pitch angle.
The fourth step, with reference roll angle
Figure BDA0003161522280000078
Reference roll angular velocity prefTrue roll angle
Figure BDA0003161522280000079
The actual roll angular speed p is input, and a double closed-loop series PID controller of a roll channel is designed, wherein the angle control is an outer loop, and the angular speed control is an inner loop.
Referring to fig. 6, the expression formula of the double closed-loop series PID controller of the roll channel designed in the fourth step of the embodiment of the present invention is as follows:
Figure BDA00031615222800000710
wherein,
Figure BDA00031615222800000711
error of roll angle, epIn order to determine the error of the roll angular velocity,
Figure BDA00031615222800000712
is the output of a double closed-loop series PID controller of a rolling channel,
Figure BDA0003161522280000081
Kp,p，Ki,p，Kd,pis the PID control parameter of the double closed-loop series PID controller of the rolling channel.
A fifth step of referencing the yaw angle psirefReference yaw rate rrefAnd the actual yaw angle psi and the actual yaw angle speed r are used as input, and a double closed-loop series PID controller of a yaw channel is designed.
Referring to fig. 6, the expression formula of the dual closed-loop series PID controller of the yaw channel designed in the fifth step of the embodiment of the present invention is as follows:
Figure BDA0003161522280000082
wherein e isψAs yaw angle error, erIn order to be able to determine the yaw rate error,
Figure BDA0003161522280000083
double closed-loop series PID controller output for yaw channel, Kp,ψ，Kp,r，Ki,r，Kd,rThe PID control parameters of the double closed-loop series PID controller of the yaw channel.
And sixthly, designing second-order extended state observers of a roll channel and a yaw channel respectively, estimating total disturbance in the extended state model of the fixed wing aircraft lateral channel established in the second step by using the fed-back actual control quantity and the control quantity after delay processing of the actuator, and performing disturbance compensation on the output of the controller to obtain the final control quantity.
The second-order extended state observer of the roll channel designed in the sixth step of the embodiment of the present invention is as follows:
Figure BDA0003161522280000084
where p is the actual roll angular velocity fed back, pesoFor the estimation of the roll angular velocity by the extended state observer, e1Representing the estimation error of roll angular velocity, Z1Representing the sum f of unmodeled dynamics, parametric perturbations and external disturbances of the extended state observer on the roll channel1(p,r,w1) Estimated value of u1(t-tau) is the rolling channel virtual control quantity delay tau1Value after time, betap1And betap2For the extended state observer parameters of the roll channel, fal is a continuous power function with a linear section near the origin, and delta is the interval length of the linear section, and is taken as 0.05.
The second-order extended state observer of the yaw channel designed in the sixth step of the embodiment of the present invention is as follows:
Figure BDA0003161522280000091
where r is the actual yaw rate of the feedback, resoFor the extended state observer estimation of yaw rate, e2Representing the yaw rate estimation error, Z2Representing the sum f of unmodeled dynamics, parametric perturbations and external disturbances of the extended state observer on the yaw path2(p,r,w2) Estimated value of u2(t-tau) is the virtual control quantity delay tau of the yaw channel2Value after time, betar1And betar2Is the extended state observer parameter of the yaw channel.
fal is a continuous power function with a linear segment near the origin, and is expressed as follows
Figure BDA0003161522280000092
Where δ is the interval length of the linear segment. In the invention, the parameters delta of the fal functions of the extended state observers of the roll channel and the yaw channel are 0.05.
Obtaining the total disturbance f through a second-order extended state observer of a roll channel and a yaw channel1(p,r,w1) And f2(p,r,w2) Is estimated by1And Z2Then, the output of the controller is compensated, including
Figure BDA0003161522280000093
Wherein,
Figure BDA0003161522280000094
and
Figure BDA0003161522280000095
double closed-loop series connection of rolling channel and yawing channel respectivelyType PID controller output, u1And u2And the virtual control quantities are respectively the virtual control quantities of the roll channel and the yaw channel after disturbance compensation.
Then the virtual control quantity is converted into an actual aileron rudder quantity delta through the following formulaaAnd rudder amount δrAnd (3) outputting:
Figure BDA0003161522280000096
an embodiment of the present invention provides a fixed-wing aircraft lateral passage control device, including:
a first module for designing a lateral navigation control law according to the target track and outputting the current target roll angle of the aircraft
Figure BDA0003161522280000101
And target yaw angle psit；
The second module is used for establishing a fixed wing aircraft transverse lateral channel expansion state decoupling model containing total disturbance;
a third module for inputting a target roll angle
Figure BDA0003161522280000102
And target yaw angle psitAdopting the steepest control comprehensive function fhan to construct a tracking differentiator, carrying out smooth filtering processing on the signal, outputting a reference signal and a differential signal thereof, and converting the reference signal and the differential signal into a reference roll angular velocity prefAnd a reference yaw rate rref
A fourth module for referencing roll angle
Figure BDA0003161522280000103
Reference roll angular velocity prefTrue roll angle
Figure BDA0003161522280000104
The actual roll angular speed p is input, and a double closed-loop series PID controller of a roll channel is designed;
a fifth module for calculating a reference yaw angle psirefReference yaw rate rrefThe actual yaw angle psi and the actual yaw angle speed r are used as input, and a double closed-loop series PID controller of a yaw channel is designed;
and the sixth module is used for respectively designing a second-order expansion state observer of a roll channel and a yaw channel, estimating total disturbance in the fixed wing aircraft lateral channel expansion state model established in the second step by using the fed-back actual control quantity and the control quantity after delay processing of the actuator, and performing disturbance compensation on the output of the controller to obtain final control quantity.
The technical features of the above embodiments can be arbitrarily combined, and for the sake of brevity, all possible combinations of the technical features in the above embodiments are not described, but should be considered as the scope of the present specification as long as there is no contradiction between the combinations of the technical features.
The above-mentioned embodiments only express several embodiments of the present application, and the description thereof is more specific and detailed, but not construed as limiting the scope of the invention. It should be noted that, for a person skilled in the art, several variations and modifications can be made without departing from the concept of the present application, which falls within the scope of protection of the present application. Therefore, the protection scope of the present patent shall be subject to the appended claims.

Claims (10)

1. The fixed wing aircraft transverse lateral channel control method is characterized by comprising the following steps:
firstly, designing a lateral navigation control law according to a target track, and outputting the current target roll angle of the aircraft
Figure FDA0003161522270000011
And target yaw angle psit；
Decoupling the transverse and lateral channel model of the fixed-wing aircraft, and expressing unmodeled dynamics, parameter uncertainty and external disturbance by using an expansion state to establish an expansion state model of the transverse and lateral channel of the fixed-wing aircraft containing the total disturbance;
thirdly, the target roll angle is adjusted
Figure FDA0003161522270000012
And target yaw angle psitAdopting the steepest control comprehensive function fhan to construct a tracking differentiator, carrying out smooth filtering processing on the signal, outputting a reference signal and a differential signal thereof, and converting the reference signal and the differential signal into a reference roll angular velocity prefAnd a reference yaw rate rref；
The fourth step, with reference roll angle
Figure FDA0003161522270000013
Reference roll angular velocity prefTrue roll angle
Figure FDA0003161522270000014
The actual roll angular speed p is input, and a double closed-loop series PID controller of a roll channel is designed, wherein the angle control is an outer loop, and the angular speed control is an inner loop;
a fifth step of referencing the yaw angle psirefReference yaw rate rrefThe actual yaw angle psi and the actual yaw angle speed r are used as input, and a double closed-loop series PID controller of a yaw channel is designed;
and sixthly, designing second-order extended state observers of a roll channel and a yaw channel respectively, estimating total disturbance in the extended state model of the fixed wing aircraft lateral channel established in the second step by using the fed-back actual control quantity and the control quantity after delay processing of the actuator, and performing disturbance compensation on the output of the controller to obtain the final control quantity.
2. The fixed-wing aircraft lateral passage control method according to claim 1, wherein in the second step, the fixed-wing aircraft lateral passage expansion state decoupling model is:
Figure FDA0003161522270000015
wherein p is the roll angular velocity, r is the yaw angular velocity,
Figure FDA0003161522270000016
is a differential signal of the roll angular velocity,
Figure FDA0003161522270000017
as a differential signal of yaw rate, f1(p，r，w1) For the total disturbance, w, in the roll pass including unmodeled dynamics of the system, uncertainty of the parameters, and external disturbances1External disturbance of the rolling path, f2(p，r，w2) For the sum disturbance, w, including unmodeled dynamics of the system, uncertainty of the parameters, and external disturbances in the yaw path2For external disturbances of the yaw channel, u1(t-τ1) Delaying tau for a virtual control quantity for a rolling channel1Output value after time, u2(t-τ2) Delaying tau for a virtual control quantity for a yaw channel2The output value after the time.
3. The fixed-wing aircraft lateral passage control method according to claim 2, characterized in that in the third step, a roll and yaw tracking differentiator is constructed using a steepest control synthesis function fhan.
The roll tracking differentiator constructed with fhan is as follows:
Figure FDA0003161522270000021
wherein,
Figure FDA0003161522270000022
in order to input the target roll angle,
Figure FDA0003161522270000023
for reference purposesThe transverse rolling angle is formed by the transverse rolling angle,
Figure FDA0003161522270000024
for reference to the differential signal of the roll angle,
Figure FDA0003161522270000025
the fast factor of the differentiator is tracked for the roll path,
Figure FDA0003161522270000026
the filter factor of the differentiator is tracked for the roll path, and u is an intermediate variable.
The yaw tracking differentiator constructed with fhan is as follows:
Figure FDA0003161522270000027
wherein psitFor an input target yaw angle, psirefIn order to refer to the yaw angle,
Figure FDA0003161522270000028
as a differential signal with reference to yaw angle, rψFast factor, h, for tracking differentiators for yaw pathψThe filter factor of the differentiator is tracked for the yaw channel, and u is the intermediate variable.
4. The fixed-wing aircraft lateral passage control method according to claim 3, wherein in the third step, the roll passage tracking differentiator outputs the reference roll angle
Figure FDA0003161522270000029
And its differential signal
Figure FDA00031615222700000210
Yaw channel tracking differentiator outputs a reference yaw angle psirefAnd its differential signal
Figure FDA00031615222700000211
Conversion into a reference roll angular velocity prefAnd a reference yaw rate rrefThe formula is as follows:
Figure FDA00031615222700000212
wherein,
Figure FDA00031615222700000213
for reference to the roll angle, θrefFor reference to the pitch angle.
5. The fixed-wing aircraft lateral passage control method according to any one of claims 1 to 4, wherein the expression of the double closed-loop series PID controller of the roll passage in the fourth step is as follows:
Figure FDA0003161522270000031
wherein,
Figure FDA0003161522270000032
error of roll angle, epIn order to determine the error of the roll angular velocity,
Figure FDA0003161522270000033
is the output of a double closed-loop series PID controller of a rolling channel,
Figure FDA0003161522270000034
Kp，p，Ki，p，Kd，pis the PID control parameter of the double closed-loop series PID controller of the rolling channel.
6. The fixed-wing aircraft lateral channel control method according to claim 5, wherein the expression of the double closed-loop series PID controller of the yaw channel in the fifth step is as follows:
Figure FDA0003161522270000035
wherein e isψAs yaw angle error, erIn order to be able to determine the yaw rate error,
Figure FDA0003161522270000036
double closed-loop series PID controller output for yaw channel, Kp，ψ，Kp，r，Ki，r，Kd，rThe PID control parameters of the double closed-loop series PID controller of the yaw channel.
7. The fixed-wing aircraft lateral passage control method according to claim 1, 2, 3, 4 or 6, wherein the second-order extended state observer of the roll passage in the sixth step is as follows:
Figure FDA0003161522270000037
where p is the actual roll angular velocity fed back, pesoFor the estimation of the roll angular velocity by the extended state observer, e1Representing the estimation error of roll angular velocity, Z1Representing the sum f of unmodeled dynamics, parametric perturbations and external disturbances of the extended state observer on the roll channel1(p，r，w1) Estimated value of u1(t-tau) is the rolling channel virtual control quantity delay tau1Value after time, betap1And betap2For the extended state observer parameters of the roll channel, fal is a continuous power function with a linear section near the origin, and delta is the interval length of the linear section, and is taken as 0.05.
8. The fixed-wing aircraft lateral passage control method according to claim 7, wherein the second-order extended state observer of the yaw passage in the sixth step is as follows:
Figure FDA0003161522270000041
where r is the actual yaw rate of the feedback, resoFor the extended state observer estimation of yaw rate, e2Representing the yaw rate estimation error, Z2Representing the sum f of unmodeled dynamics, parametric perturbations and external disturbances of the extended state observer on the yaw path2(p，r，w2) Estimated value of u2(t-tau) is the virtual control quantity delay tau of the yaw channel2Value after time, betar1And betar2Is the extended state observer parameter of the yaw channel.
9. The fixed-wing aircraft lateral passage control method according to claim 8, wherein in the sixth step, the final control amount is obtained by:
firstly, the controller output is compensated by using an extended state observer to estimate the disturbance value, namely:
Figure FDA0003161522270000042
wherein,
Figure FDA0003161522270000043
and
Figure FDA0003161522270000044
double closed-loop series PID controller output u, of roll and yaw channels, respectively1And u2And the virtual control quantities are respectively the virtual control quantities of the roll channel and the yaw channel after disturbance compensation.
Then the virtual control quantity is converted into an actual aileron rudder quantity delta through the following formulaaAnd rudder amount δrAnd (3) outputting:
Figure FDA0003161522270000045
wherein P is an invertible matrix.
10. Horizontal lateral channel controlling means of fixed wing aircraft, its characterized in that includes:
a first module for designing a lateral navigation control law according to the target track and outputting the current target roll angle of the aircraft
Figure FDA0003161522270000046
And target yaw angle psit；
The second module is used for establishing a fixed wing aircraft transverse lateral channel expansion state decoupling model containing total disturbance;
a third module for inputting a target roll angle
Figure FDA0003161522270000047
And target yaw angle psitAdopting the steepest control comprehensive function fhan to construct a tracking differentiator, carrying out smooth filtering processing on the signal, outputting a reference signal and a differential signal thereof, and converting the reference signal and the differential signal into a reference roll angular velocity prefAnd a reference yaw rate rref
A fourth module for referencing roll angle
Figure FDA0003161522270000048
Reference roll angular velocity prefTrue roll angle
Figure FDA0003161522270000049
The actual roll angular speed p is input, and a double closed-loop series PID controller of a roll channel is designed;
a fifth module for calculating a reference yaw angle psirefReference yaw rate rrefDesigning a yaw channel by taking an actual yaw angle psi and an actual yaw angular speed r as inputThe double closed-loop series PID controller;
and the sixth module is used for respectively designing a second-order expansion state observer of a roll channel and a yaw channel, estimating total disturbance in the fixed wing aircraft lateral channel expansion state model established in the second step by using the fed-back actual control quantity and the control quantity after delay processing of the actuator, and performing disturbance compensation on the output of the controller to obtain final control quantity.
CN202110792449.6A 2021-07-14 2021-07-14 Method and device for controlling transverse and lateral channels of fixed-wing aircraft Withdrawn CN114137999A (en)

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115509249A (en) * 2022-10-24 2022-12-23 北京理工大学 Attitude control method of four-rotor aircraft in Mars gust strong interference environment

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107491081A (en) * 2017-07-12 2017-12-19 西北工业大学 A kind of anti-interference four rotor wing unmanned aerial vehicles attitude control method
CN107688295A (en) * 2017-08-29 2018-02-13 浙江工业大学 A kind of quadrotor finite time self-adaptation control method based on fast terminal sliding formwork
CN107943094A (en) * 2017-12-27 2018-04-20 上海应用技术大学 The sliding-mode control and its controller of a kind of quadrotor
CN112987559A (en) * 2021-04-06 2021-06-18 广东英诺威盛科技有限公司 Method and device for controlling longitudinal channel of fixed-wing aircraft
CN113110593A (en) * 2021-05-06 2021-07-13 西北工业大学 Flight formation cooperative self-adaptive control method based on virtual structure and estimation information transmission

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107491081A (en) * 2017-07-12 2017-12-19 西北工业大学 A kind of anti-interference four rotor wing unmanned aerial vehicles attitude control method
CN107688295A (en) * 2017-08-29 2018-02-13 浙江工业大学 A kind of quadrotor finite time self-adaptation control method based on fast terminal sliding formwork
CN107943094A (en) * 2017-12-27 2018-04-20 上海应用技术大学 The sliding-mode control and its controller of a kind of quadrotor
CN112987559A (en) * 2021-04-06 2021-06-18 广东英诺威盛科技有限公司 Method and device for controlling longitudinal channel of fixed-wing aircraft
CN113110593A (en) * 2021-05-06 2021-07-13 西北工业大学 Flight formation cooperative self-adaptive control method based on virtual structure and estimation information transmission

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115509249A (en) * 2022-10-24 2022-12-23 北京理工大学 Attitude control method of four-rotor aircraft in Mars gust strong interference environment

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