CN108646554B - Aircraft rapid anti-interference longitudinal guidance method based on designated performance - Google Patents

Aircraft rapid anti-interference longitudinal guidance method based on designated performance Download PDF

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CN108646554B
CN108646554B CN201810419551.XA CN201810419551A CN108646554B CN 108646554 B CN108646554 B CN 108646554B CN 201810419551 A CN201810419551 A CN 201810419551A CN 108646554 B CN108646554 B CN 108646554B
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sliding mode
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乔建忠
张丹瑶
郭雷
朱玉凯
徐健伟
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Beihang University
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    • G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
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Abstract

The invention relates to an aircraft rapid anti-interference longitudinal guidance method based on designated performance, which is characterized in that a longitudinal aircraft dynamics model containing uncertain aerodynamic parameters is established, wherein the aerodynamic parameters comprise a lift coefficient and a resistance coefficient; performing appointed performance conversion through the designed appointed performance function and the conversion function to obtain a converted model; designing a sliding mode disturbance observer to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft to obtain a disturbance estimated value; designing a sliding mode control law to meet the requirement of a quick control task; and designing a composite sliding mode controller to finish the rapid anti-interference longitudinal interference guidance method of the aircraft based on the designated performance. The invention can specify three performances of convergence rate, overshoot and steady-state error of the control system, has the characteristics of rapidity and high precision, is suitable for specified performance rapid anti-interference guidance systems of various flight systems and other high-altitude unmanned aerial vehicles, and can also solve the problem of aircraft faults such as rapid fault tolerance.

Description

Aircraft rapid anti-interference longitudinal guidance method based on designated performance
Technical Field
The invention relates to an aircraft rapid anti-interference longitudinal guidance method based on designated performance, which solves the problem of rapid high-precision anti-interference guidance of an aircraft with uncertain pneumatic parameters.
Background
The aircrafts such as unmanned planes, hypersonic aircrafts, missiles and the like are used as important weapons in the military field, have the important function in the fields of high-altitude pursuit, unmanned investigation, global striking and the like, have the advantages of rapidity, high maneuverability, accuracy and the like, and are widely applied to multiple fields such as military use, civil use and the like. During the mission of the aircraft, the reentry guidance of the tip is particularly important, and is one of the key technologies of the mission, and the requirements of rapidity and accuracy need to be met. For example, if the weapons such as missiles and the like cannot meet certain rapidity indexes, the weapons are easy to be intercepted and destroyed by an interception system, so that the task fails, and thus the rapidity is the basic requirement of the reentry process. However, the flight span of the reentry process of the aircraft is large, the aerodynamic heat generated at high speed causes the elastic deformation of the aircraft, and the flight environment changes rapidly and complexly, so that the aerodynamic parameters of the aircraft are unstable, and great uncertainty exists. The uncertainty of the pneumatic parameters directly influences the accuracy of the reentry process of the aircraft, and also seriously influences the basic condition of rapidity, and the uncertainty needs to be subjected to anti-interference control to improve the performance of the reentry guidance system. Therefore, by combining the prior art, the design of the rapid anti-interference longitudinal guidance method of the aircraft with the designated performance is important, and the method has a wide application prospect.
At present, the research on the control of the assigned performance by domestic scholars is less, the literature 'cooperative control of a class of nonlinear multi-agent systems meeting the assigned performance' is provided, a leader-follower cooperative control problem meeting the assigned performance is researched for a class of uncertain nonlinear multi-agent systems, and an adaptive fuzzy cooperative control algorithm is provided based on a dynamic surface control technology and an assigned performance control technology, but the method does not consider the problem of anti-interference control and ignores the uncertainty of the system. In order to solve the problem of guidance of reentry of aircrafts, scholars at home and abroad make a great deal of research. Patent number 201610366190.8 proposes an air-to-air missile guidance method based on a sliding mode variable structure, which improves estimation of a target motion state in the initial guidance stage and improves missile hit precision in the final guidance stage. The article 'hypersonic aircraft longitudinal plane gliding flight guidance control method' utilizes a dynamic surface control method, a terminal sliding mode control method and a second-order sliding mode control method to complete the design of a height control system in the hypersonic aircraft longitudinal plane. The above patents and articles all use sliding mode control methods to enable the system to be converged within a limited time, but do not consider the problem of interference resistance under a complex environment, and for the problem, the following research improves the aircraft height sliding mode control system. Patent application No. 201610306205.1 proposes an anti-interference composite online guidance method for an atmospheric admission segment of a mars lander, but an observer used by the method does not have limited time convergence capability and cannot meet the requirement of rapid guidance. In the article, "hypersonic aircraft Terminal sliding mode control based on disturbance observer", a nonlinear disturbance observer is designed for enhancing the robustness of a controller on the basis of sliding mode control, model uncertainty items are subjected to self-adaptive estimation and compensation, but the anti-disturbance sliding mode control method cannot perform specified performance control on the height.
In conclusion, the conventional method lacks the rapid and high-precision control capability with specified performance under the condition that the pneumatic parameters are uncertain, and needs to overcome the rapid anti-interference longitudinal guidance method of the aircraft based on the specified performance.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: aiming at the problem of rapid anti-interference altitude control of an aircraft with uncertain pneumatic parameters, the defects of the prior art are overcome, and the rapid anti-interference longitudinal guidance method of the aircraft based on the designated performance is provided, so that rapid altitude control of the aircraft based on the designated performance and rapid estimation and compensation of uncertainty of the pneumatic parameters are realized, designated performance constraint is carried out on the altitude control process of the aircraft, and rapidity, accuracy and anti-interference capability of the altitude control process of the aircraft are improved.
The invention and the technical solution are as follows: a rapid anti-interference longitudinal guidance method of an aircraft based on designated performance is characterized in that a dynamic model containing uncertain pneumatic parameters is established, designated performance functions and conversion functions are designed to carry out designated performance conversion on a system, and a composite rapid anti-interference height controller is designed by utilizing a method of combining a sliding mode controller and a sliding mode interference observer on the basis of conversion, and the specific implementation steps are as follows:
the method comprises the following steps of firstly, establishing a longitudinal aircraft dynamic model with uncertain aerodynamic parameters, wherein the aerodynamic parameters comprise a lift coefficient and a drag coefficient:
Figure BDA0001650318110000021
Figure BDA0001650318110000022
Figure BDA0001650318110000023
wherein, the distance from the geocenter to the mass center of the aircraft is r, the relative earth speed of the aircraft is V, and the track dip angle is gamma;
Figure BDA0001650318110000031
are respectively one of r, V and gammaA first derivative; sigma is the aircraft roll angle, g is the gravitational acceleration, d1、d2The method is characterized in that equivalent interference with uncertain aerodynamic parameters is represented, L and D respectively represent lift acceleration and drag acceleration, and the expression form is as follows:
Figure BDA0001650318110000032
Figure BDA0001650318110000033
where ρ is the atmospheric density, S is the reference area of the aircraft, m is the mass of the aircraft, CLAnd CDRespectively, the lift coefficient and the drag coefficient of the whole body. The lift coefficient and drag coefficient are modeled as follows:
CL＝CL1α2+CL2α+CL3Ma+CL4
CD＝CD1α2+CD2α+CD3Ma+CD4
wherein M isaMach number, α angle of attack, CL1、CL2、CL3、CL4The coefficient is a second-order attack angle coefficient, a first-order attack angle coefficient, a Mach number coefficient and a constant coefficient of the lift coefficient; cD1、CD2、CD3、CD4The control variables are selected as an aircraft roll angle sigma and an attack angle α.
And secondly, performing specified performance conversion through the designed specified performance function and the conversion function according to the longitudinal dynamics model in the first step to obtain a converted model:
designing a specified performance function
Figure BDA0001650318110000034
The following were used:
Figure BDA0001650318110000035
wherein, t is the time,
Figure BDA0001650318110000036
a > 0 is a performance function parameter.
The transfer function Z is designed as follows:
Figure BDA0001650318110000037
the first time derivative is found for the conversion function Z:
Figure BDA0001650318110000038
wherein the content of the first and second substances,
Figure BDA0001650318110000039
as the first time derivative of the transfer function Z,
Figure BDA00016503181100000310
is the first time derivative of the aircraft's distance r from the geocentric,
Figure BDA00016503181100000311
for specifying a function of performance
Figure BDA00016503181100000312
First time derivative of, τzFor the parameters of the transfer function, expressed as
Figure BDA0001650318110000041
The second time derivative continues to be found for the conversion function Z:
Figure BDA0001650318110000042
wherein the content of the first and second substances,
Figure BDA0001650318110000043
being the second time derivative of the transfer function Z,
Figure BDA0001650318110000044
for specifying a function of performance
Figure BDA0001650318110000045
The second-order time derivative of (a),
Figure BDA0001650318110000046
is tauzThe first time derivative of (a). The above model is simplified into the following expression:
Figure BDA0001650318110000047
wherein the content of the first and second substances,
Figure BDA0001650318110000048
for the non-linear term of the transformed model, uz＝τz(-Dsin γ + Lcos γ cos σ) is the equivalent control input for the transformed model, dz＝τz(sinγd1+cosγVd2) Is the equivalent interference of the transformed model.
Thirdly, designing a sliding mode disturbance observer according to the model converted in the second step to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft, and obtaining a disturbance estimation value:
the disturbance observer is designed as follows:
Figure BDA0001650318110000049
Figure BDA00016503181100000410
Figure BDA00016503181100000411
Figure BDA00016503181100000412
wherein z is0In the case of the intermediate variables of the state,
Figure BDA00016503181100000413
is z0First derivative of v0For the intermediate variables of the function, the intermediate variables,
Figure BDA00016503181100000414
is v is0The first derivative of (a) is,
Figure BDA00016503181100000415
as an unknown equivalent interference dzIs determined by the estimated value of (c),
Figure BDA00016503181100000416
is composed of
Figure BDA00016503181100000417
The first derivative of (a) is,
Figure BDA00016503181100000418
for first derivative of unknown equivalent interference
Figure BDA00016503181100000419
Is determined by the estimated value of (c),
Figure BDA00016503181100000420
is composed of
Figure BDA00016503181100000421
First derivative of, λ0、λ1、λ2Is the observer gain and is a positive number. sign (·) denotes the derivation of a sign function.
Fourthly, designing a sliding mode control law to meet the requirement of a quick control task:
the sliding mode control law is designed as follows:
Figure BDA00016503181100000422
Figure BDA00016503181100000423
wherein the content of the first and second substances,
Figure BDA0001650318110000051
is a sliding mode controller, b is a state coefficient more than 0, 1 is more than tau and less than 2 is a symbol state coefficient, k1More than 0 is the sliding mode surface coefficient, k2More than 0 is the coefficient of the symbol sliding mode surface, more than 0 mu and less than 1 is the order value of the sliding mode, and s is the sliding mode surface.
Fifthly, designing a composite sliding mode controller by using the interference estimation value in the third step and the sliding mode control law in the fourth step, and completing the rapid anti-interference longitudinal interference guidance method of the aircraft based on the specified performance, wherein the method comprises the following steps:
designing a composite proportional pilot controller:
Figure BDA0001650318110000052
wherein u iseIs a sliding-mode controller, which is provided with a sliding-mode controller,
Figure BDA0001650318110000053
as an unknown equivalent interference dzAn estimate of (d).
Compared with the prior art, the invention has the advantages that: the invention relates to an aircraft rapid anti-interference longitudinal guidance method based on specified performance, aiming at the defect that the existing method lacks the rapid and high-precision control capability with the specified performance under the condition of uncertain pneumatic parameters, firstly, establishing a longitudinal aircraft dynamic model with uncertain pneumatic parameters, wherein the pneumatic parameters comprise a lift coefficient and a drag coefficient; secondly, performing specified performance conversion through the designed specified performance function and the conversion function according to the longitudinal dynamics model in the first step to obtain a converted model; then, designing a sliding mode disturbance observer according to the model converted in the second step to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft to obtain a disturbance estimated value; then designing a sliding mode control law to finish the requirement of a quick control task; and finally, designing a composite sliding mode controller by using the interference estimation value in the third step and the sliding mode control law in the fourth step, and finishing the rapid anti-interference longitudinal interference guidance method of the aircraft based on the specified performance. The invention adopts the rapid anti-interference guidance method based on the combination of the sliding mode interference observer with the sliding mode controller with the specified performance, can specify the three performances of the convergence rate, the overshoot and the steady-state error of the control system, has the characteristics of rapidity and high precision, is suitable for the specified performance rapid anti-interference guidance systems of various flight systems and other high-altitude unmanned aerial vehicles, can solve the aircraft fault problems of rapid fault tolerance and the like, and meets the requirements of the rapidity, the high precision and the like of an aircraft height control system.
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FIG. 1 is a design flow chart of the rapid anti-interference longitudinal guidance method of the aircraft based on the designated performance.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings and examples.
As shown in FIG. 1, the invention relates to a rapid anti-interference longitudinal guidance method for an aircraft based on specified performance. The method comprises the steps of firstly, establishing a longitudinal aircraft dynamic model with uncertain aerodynamic parameters, wherein the aerodynamic parameters comprise a lift coefficient and a drag coefficient; secondly, performing specified performance conversion through the designed specified performance function and the conversion function according to the longitudinal dynamics model in the first step to obtain a converted model; thirdly, designing a sliding mode disturbance observer according to the model converted in the second step to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft to obtain a disturbance estimated value; fourthly, designing a sliding mode control law to meet the requirement of a quick control task; and fifthly, designing a composite sliding mode controller by using the interference estimation value in the third step and the sliding mode control law in the fourth step, and finishing the rapid anti-interference longitudinal interference guidance method of the aircraft based on the specified performance. The invention adopts the rapid anti-interference guidance method based on the combination of the sliding mode interference observer with the sliding mode controller with specified performance, can specify the three performances of convergence rate, overshoot and steady-state error of the control system, has the characteristics of rapidity and high precision, is suitable for the specified performance rapid anti-interference guidance systems of various flight systems and other high-altitude unmanned aerial vehicles, and can also solve the problems of rapid fault tolerance and other aircraft faults.
The specific implementation steps are as follows:
the method comprises the following steps of firstly, establishing a longitudinal aircraft dynamic model with uncertain aerodynamic parameters, wherein the aerodynamic parameters comprise a lift coefficient and a drag coefficient:
Figure BDA0001650318110000061
Figure BDA0001650318110000062
Figure BDA0001650318110000063
wherein the initial value of the distance from the geocenter to the mass center of the aircraft is 30480km, the relative earth speed of the aircraft is 3352.8m/s, the track inclination angle is-0.785 rad. (ii) a
Figure BDA0001650318110000064
First derivatives of r, V, and γ, respectively; sigma is the aircraft roll angle, g is the gravitational acceleration, and the value is 9.8m/s2，d1、d2The method is characterized in that equivalent interference with uncertain aerodynamic parameters is represented, L and D respectively represent lift acceleration and drag acceleration, and the expression form is as follows:
Figure BDA0001650318110000065
Figure BDA0001650318110000066
wherein rho is the atmospheric density and takes the value of 1.225kg/m3S is the reference area of the aircraft, and the value is 149.4m2M is the mass of the aircraft, and the value is 35828kg, CLAnd CDRespectively, the lift coefficient and the drag coefficient of the whole body. The lift coefficient and drag coefficient are modeled as follows:
CL＝-0.000522α2+0.03506α-0.04857Ma+0.1577
CD＝0.0001432α2+0.00558α-0.01048Ma+0.2204
wherein M isaAt mach number, the initial value is 11Ma, α is the angle of attack, and the control variables are selected as the aircraft roll angle σ and the angle of attack α.
And secondly, performing specified performance conversion through the designed specified performance function and the conversion function according to the longitudinal dynamics model in the first step to obtain a converted model:
designing a specified performance function
Figure BDA0001650318110000071
The following were used:
Figure BDA0001650318110000072
wherein, t is the time,
Figure BDA0001650318110000073
and a is 3.2 as a performance function parameter.
The transfer function Z is designed as follows:
Figure BDA0001650318110000074
the first time derivative is found for the conversion function Z:
Figure BDA0001650318110000075
wherein the content of the first and second substances,
Figure BDA0001650318110000076
as the first time derivative of the transfer function Z,
Figure BDA0001650318110000077
is the first time derivative of the aircraft's distance r from the geocentric,
Figure BDA0001650318110000078
for specifying a function of performance
Figure BDA0001650318110000079
First time derivative of, τzFor the parameters of the transfer function, expressed as
Figure BDA00016503181100000710
The second time derivative continues to be found for the conversion function Z:
Figure BDA00016503181100000711
wherein the content of the first and second substances,
Figure BDA00016503181100000712
being the second time derivative of the transfer function Z,
Figure BDA00016503181100000713
for specifying a function of performance
Figure BDA00016503181100000714
The second-order time derivative of (a),
Figure BDA00016503181100000715
is tauzThe first time derivative of (a). The above model is simplified into the following expression:
Figure BDA00016503181100000716
wherein the content of the first and second substances,
Figure BDA00016503181100000717
for the non-linear term of the transformed model, uz＝τz(-Dsinγ+Lcosγcos σ) as an equivalent control input to the transformed model, dz＝τz(sinγd1+cosγVd2) Is the equivalent interference of the transformed model.
Thirdly, designing a sliding mode disturbance observer according to the model converted in the second step to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft, and obtaining a disturbance estimation value:
the disturbance observer is designed as follows:
Figure BDA0001650318110000081
Figure BDA0001650318110000082
Figure BDA0001650318110000083
Figure BDA0001650318110000084
wherein z is0In the case of the intermediate variables of the state,
Figure BDA0001650318110000085
is z0First derivative of v0For the intermediate variables of the function, the intermediate variables,
Figure BDA0001650318110000086
is v is0The first derivative of (a) is,
Figure BDA0001650318110000087
as an unknown equivalent interference dzIs determined by the estimated value of (c),
Figure BDA0001650318110000088
is composed of
Figure BDA0001650318110000089
The first derivative of (a) is,
Figure BDA00016503181100000810
for first derivative of unknown equivalent interference
Figure BDA00016503181100000811
Is determined by the estimated value of (c),
Figure BDA00016503181100000812
is composed of
Figure BDA00016503181100000813
First derivative of, λ0、λ1、λ2For observer gain, 2, 1.5, 1.1 can be taken, respectively. sign (·) denotes the derivation of a sign function.
Fourthly, designing a sliding mode control law to meet the requirement of a quick control task:
the sliding mode control law is designed as follows:
Figure BDA00016503181100000814
Figure BDA00016503181100000815
wherein the content of the first and second substances,
Figure BDA00016503181100000816
is a sliding mode controller, b is a state coefficient with the value of 1 when being more than 0, tau is more than 1 and less than 2, is a symbol state coefficient with the value of 1.5, and k is1The coefficient of sliding mode surface is more than 0, the value is 2, k2The coefficient of the symbol sliding mode surface is more than 0, the value is 1.3, the value of mu is more than 0 and less than 1 is the sliding mode order value, the value is 0.7, and s is the sliding mode surface.
Fifthly, designing a composite sliding mode controller by using the interference estimation value in the third step and the sliding mode control law in the fourth step, and completing the rapid anti-interference longitudinal interference guidance method of the aircraft based on the specified performance, wherein the method comprises the following steps:
designing a composite proportional pilot controller:
Figure BDA00016503181100000817
wherein u iseIs a sliding-mode controller, which is provided with a sliding-mode controller,
Figure BDA00016503181100000818
as an unknown equivalent interference dzIs estimated value of
The method of the invention is adopted for longitudinal guidance, the performance of height control can be limited within the range of a performance function, the convergence rate, overshoot and steady-state error of the system are ensured to be limited values, the upper limit of the convergence rule can be limited to 0.01m/s according to the performance function parameters in the specific implementation steps, the overshoot is controlled to be below 0m, and the steady-state error is +/-5 multiplied by 10-4And m is selected. Meanwhile, compared with a controller with interference-free estimation and compensation, the control effect can reduce the guidance time by 10-20%, and the interference estimation error is stable within 1s, thereby meeting the requirements of stability and rapidity.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.

Claims (2)

1. An aircraft rapid anti-interference longitudinal guidance method based on designated performance is characterized by comprising the following steps:
the method comprises the steps of firstly, establishing a longitudinal aircraft dynamic model with uncertain aerodynamic parameters, wherein the aerodynamic parameters comprise a lift coefficient and a drag coefficient;
secondly, performing specified performance conversion through a designed specified performance function and a designed conversion function according to the longitudinal aircraft dynamics model in the first step to obtain a converted model;
thirdly, designing a sliding mode disturbance observer to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft according to the model converted in the second step to obtain a disturbance estimation value;
fourthly, designing a sliding mode control law for finishing the requirement of the rapid control task;
fifthly, designing a composite sliding mode controller by using the interference estimation value in the third step and the sliding mode control law in the fourth step, and completing the rapid anti-interference longitudinal guidance method of the aircraft based on specified performance;
in the first step, a longitudinal aircraft dynamic model with uncertain aerodynamic parameters is established, wherein the aerodynamic parameters comprise a lift coefficient and a drag coefficient, and the method comprises the following specific steps:
Figure FDA0002352867330000011
Figure FDA0002352867330000012
Figure FDA0002352867330000013
wherein, the distance from the geocenter to the mass center of the aircraft is r, the relative earth speed of the aircraft is V, and the track dip angle is gamma;
Figure FDA0002352867330000014
first derivatives of r, V, and γ, respectively; sigma is the aircraft roll angle, g is the gravitational acceleration, d1、d2The method is characterized in that equivalent interference with uncertain aerodynamic parameters is represented, L and D respectively represent lift acceleration and drag acceleration, and the expression form is as follows:
Figure FDA0002352867330000015
Figure FDA0002352867330000016
where ρ is the atmospheric density, S is the reference area of the aircraft, m is the mass of the aircraft, CLAnd CDThe lift coefficient and the drag coefficient are respectively integral, and the lift coefficient and the drag coefficient are modeled as follows:
CL＝CL1α2+CL2α+CL3Ma+CL4
CD＝CD1α2+CD2α+CD3Ma+CD4
wherein M isaMach number, α angle of attack, CL1、CL2、CL3、CL4The coefficient is a second-order attack angle coefficient, a first-order attack angle coefficient, a Mach number coefficient and a constant coefficient of the lift coefficient; cD1、CD2、CD3、CD4The control quantity is selected as an aircraft roll angle sigma and an attack angle α;
in the second step, the first step is carried out,
designing a specified performance function
Figure FDA0002352867330000021
The following were used:
Figure FDA0002352867330000022
wherein, t is the time,
Figure FDA0002352867330000023
a is more than 0 and is a performance function parameter;
the transfer function Z is designed as follows:
Figure FDA0002352867330000024
the first time derivative is found for the conversion function Z:
Figure FDA0002352867330000025
wherein the content of the first and second substances,
Figure FDA0002352867330000026
as the first time derivative of the transfer function Z,
Figure FDA0002352867330000027
is the first time derivative of the centroid to aircraft centroid distance r,
Figure FDA0002352867330000028
for specifying a function of performance
Figure FDA0002352867330000029
First time derivative of, τzFor the parameters of the transfer function, expressed as
Figure FDA00023528673300000210
The second time derivative continues to be found for the conversion function Z:
Figure FDA00023528673300000211
wherein the content of the first and second substances,
Figure FDA00023528673300000212
being the second time derivative of the transfer function Z,
Figure FDA00023528673300000213
for specifying a function of performance
Figure FDA00023528673300000214
The second-order time derivative of (a),
Figure FDA00023528673300000215
is tauzThe above model is simplified to the following expression:
Figure FDA00023528673300000216
wherein the content of the first and second substances,
Figure FDA00023528673300000217
for the non-linear term of the transformed model, uz＝τz(-D sin γ + L cos γ cos σ) is the equivalent control input to the transformed model, Dz＝τz(sinγd1+cosγVd2) Is the equivalent interference of the transformed model;
in the third step, a sliding mode disturbance observer is designed according to the model converted in the second step to quickly estimate the uncertainty of the aerodynamic parameters of the aircraft, so that a disturbance estimation value is obtained:
the disturbance observer is designed as follows:
Figure FDA0002352867330000031
Figure FDA0002352867330000032
Figure FDA0002352867330000033
Figure FDA0002352867330000034
wherein z is0In the case of the intermediate variables of the state,
Figure FDA00023528673300000319
is z0First derivative of v0For the intermediate variables of the function, the intermediate variables,
Figure FDA0002352867330000035
is v is0The first derivative of (a) is,
Figure FDA0002352867330000036
as an unknown equivalent interference dzIs determined by the estimated value of (c),
Figure FDA0002352867330000037
is composed of
Figure FDA0002352867330000038
The first derivative of (a) is,
Figure FDA0002352867330000039
for first derivative of unknown equivalent interference
Figure FDA00023528673300000310
Is determined by the estimated value of (c),
Figure FDA00023528673300000311
is composed of
Figure FDA00023528673300000312
First derivative of, λ0、λ1、λ2For observer gain and positive number, sign (·) represents solving a sign function;
in the fourth step, a sliding mode control law is designed as follows:
Figure FDA00023528673300000313
Figure FDA00023528673300000314
wherein the content of the first and second substances,
Figure FDA00023528673300000315
is a sliding mode controller, b is a state coefficient more than 0, 1 is more than tau and less than 2 is a symbol state coefficient, k1More than 0 is the sliding mode surface coefficient, k2More than 0 is the coefficient of the symbol sliding mode surface, more than 0 mu and less than 1 is the order value of the sliding mode, and s is the sliding mode surface.
2. The rapid anti-interference longitudinal guidance method for the aircraft based on the designated performance as claimed in claim 1, characterized in that: and the fifth step, designing a composite sliding mode controller:
Figure FDA00023528673300000316
wherein the content of the first and second substances,
Figure FDA00023528673300000317
is a sliding-mode controller, which is provided with a sliding-mode controller,
Figure FDA00023528673300000318
as an unknown equivalent interference dzAn estimate of (d).
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