CN112099348A - Collision angle control guidance method based on observer and global sliding mode - Google Patents

Collision angle control guidance method based on observer and global sliding mode Download PDF

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CN112099348A
CN112099348A CN202010837398.XA CN202010837398A CN112099348A CN 112099348 A CN112099348 A CN 112099348A CN 202010837398 A CN202010837398 A CN 202010837398A CN 112099348 A CN112099348 A CN 112099348A
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missile
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sliding mode
guidance
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易文俊
张文广
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Nanjing University of Science and Technology
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Abstract

The invention discloses an impact angle control guidance method based on an observer and a global sliding mode. The method comprises the following steps: firstly, establishing a bullet relative kinematics model; then acquiring the motion parameters of the projectile body and the target through a projectile-borne sensor; estimating target acceleration information; then designing a guidance law; finally, inputting a guidance instruction into an actuator to control the missile to fly; and repeating the step motion parameter measurement and the guidance control until the guidance is finished. The invention has the characteristic of limited time convergence and is suitable for static targets and maneuvering targets.

Description

Collision angle control guidance method based on observer and global sliding mode
Technical Field
The invention relates to the technical field of missile guidance, in particular to a collision angle control guidance method based on an observer and a global sliding mode.
Background
In order to achieve accurate strike, a designed guidance law is required to ensure a small miss distance. However, in some special application scenarios, such as anti-tank missiles and anti-ship missiles, a designed guidance law can be required to control the collision angle so as to enhance the damage effect. Scholars at home and abroad propose solutions to the multi-constraint guidance law problem. For example, Dou L. (Dou, L., Dou, J.: The design of optimal guidance law with multi-constraints using block pulse functions. Aerospace Science and Technology 23(1), 201-. Lee C.H. et al (Lee, C.H., Shin, H.S., Tsourdos, A.: A New Command Shaping navigation Law using Lagrange Multiplier. IFAC paperOnLine 50(1), 15185-. Liu D.W. et al (Liu, D.W., Xia, Q., Du, Y.: Miss Distance and Impact Angle Error of trajectory Shaping guide Law Based on Statistical linear linearity addition method 11 (13)), 6370-6377(2011)) propose an Impact Angle control Guidance Law Based on a linear quadratic optimization control theory, and adopt a Statistical linearity adaptation Guidance Law to study the characteristics of the proposed Guidance Law. Erer K.S. et al (er K.S.: index Impact-Angle-Control acquisition Targets Using binary Pure presentation navigation. journal of Guidance controls and Dynamics,35 (2)), 700-. However, the above methods all need to acquire the remaining flight time of the missile, and the remaining flight time of the missile is usually acquired by an estimation method, so that estimation errors are inevitably introduced, and the missile cannot complete a precise hitting task; on the other hand, none of the above methods can guarantee the finite time convergence characteristic.
Disclosure of Invention
The invention aims to provide a collision angle control guidance method based on an observer and a global sliding mode and having a finite time convergence characteristic, so that a missile can complete a task of accurately striking.
The technical solution for realizing the purpose of the invention is as follows: an impact angle control guidance method based on an observer and a global sliding mode comprises the following steps:
step 1, establishing a bullet relative kinematics model;
step 2, acquiring the motion parameters of a projectile body and a target through a projectile-borne sensor;
step 3, estimating target acceleration information;
step 4, designing a guidance law;
step 5, inputting instructions to an actuator to control the missile to fly;
and 6, repeating the steps 2 to 5 until the guidance is finished.
Further, the establishment of the bullet relative kinematics model in the step 1 specifically includes the following steps:
the relative kinematics model of the missile-target in the two-dimensional plane is established as follows:
Figure BDA0002640204470000021
wherein, the lower corner marks T and M represent the target and the missile; a represents acceleration, V represents velocity, θ represents line of sight angle, λ and the direction of flight angles of the missile and target, respectively, relative to the line of sight LOS; γ represents a flight path angle;
Figure BDA0002640204470000022
and
Figure BDA0002640204470000023
respectively representing the acceleration components of the missile and the target perpendicular to the LOS; vθAnd VRRepresenting relative velocity perpendicular to and along the LOS, respectively; the calculation is made according to the following formula:
Figure BDA0002640204470000024
Vθ＝VMsin-VTsinλ,VR＝VMcos-VTcosλ (3)。
further, the step 2 of obtaining the motion parameters of the projectile body and the target through the projectile-loaded sensor specifically includes:
the projectile motion parameters obtained by the projectile loading sensor are as follows: vMλ; the target motion parameters obtained by the missile-borne sensor are as follows: vT，。
Further, the estimating of the target acceleration information in step 3 specifically includes:
observer dynamics are represented by:
Figure BDA0002640204470000025
Figure BDA0002640204470000026
Figure BDA0002640204470000027
Figure BDA0002640204470000028
wherein h is0、h1、h2、z0、z1、z2Intermediate variables, with no specific meaning; sgn () is a sign function; k>0 is a design constant;
the dynamic state of the observer is solved in real time to obtain the estimation of the target acceleration
Figure BDA0002640204470000031
Comprises the following steps:
Figure BDA0002640204470000032
further, designing a guidance law in the step 4 specifically includes:
defining angular error variable
Figure BDA0002640204470000033
Is composed of
Figure BDA0002640204470000034
To pair
Figure BDA0002640204470000035
Derivative to obtain
Figure BDA0002640204470000036
Design the sliding mode variable s as
Figure BDA0002640204470000037
Wherein, mu1>0 and mu2>0 is a constant, t represents time;
acceleration of missile of
Figure BDA0002640204470000038
Wherein g >0 is a constant;
it is known that
Figure BDA0002640204470000039
Require that
Figure BDA00026402044700000310
A is an intermediate variable, with no specific meaning, calculated according to the following formula:
Figure BDA00026402044700000311
compared with the prior art, the invention has the following remarkable advantages: (1) the guidance method is based on a global sliding mode control method and a fast observer, and has the characteristic of finite time convergence; (2) by adopting a sliding mode control technology, the control of a large-range impact angle can be realized; (3) the method is suitable for static targets and maneuvering targets and has a wide application range.
Drawings
FIG. 1 is a schematic flow chart diagram of an impact angle control guidance method based on a global sliding mode.
FIG. 2 is a diagram of the missile flight trajectory during the simulation.
FIG. 3 is a global sliding mode variable trajectory diagram in the simulation process.
FIG. 4 is a view angle trace diagram during simulation.
FIG. 5 is a diagram of acceleration trajectory during simulation.
Detailed Description
The invention discloses an impact angle control guidance method based on an observer and a global sliding mode, which comprises the following steps of:
step 1, establishing a bullet relative kinematics model;
step 2, acquiring the motion parameters of a projectile body and a target through a projectile-borne sensor;
step 3, estimating target acceleration information;
step 4, designing a guidance law;
step 5, inputting instructions to an actuator to control the missile to fly;
and 6, repeating the steps 2 to 5 until the guidance is finished.
Further, the establishment of the bullet relative kinematics model in the step 1 specifically includes the following steps:
the relative kinematics model of the missile-target in the two-dimensional plane is established as follows:
Figure BDA0002640204470000041
wherein, the lower corner marks T and M represent the target and the missile; a represents acceleration, V represents velocity, θ represents line of sight angle, λ and the direction of flight angles of the missile and target, respectively, relative to the line of sight LOS; γ represents a flight path angle;
Figure BDA0002640204470000042
and
Figure BDA0002640204470000043
respectively representing the acceleration components of the missile and the target perpendicular to the LOS; vθAnd VRRepresenting relative velocity perpendicular to and along the LOS, respectively; the calculation is made according to the following formula:
Figure BDA0002640204470000044
Vθ＝VMsin-VTsinλ,VR＝VMcos-VTcosλ (3)。
further, the step 2 of obtaining the motion parameters of the projectile body and the target through the projectile-loaded sensor specifically includes:
the projectile motion parameters obtained by the projectile loading sensor are as follows: vMλ; the target motion parameters obtained by the missile-borne sensor are as follows: vT，。
Further, the estimating of the target acceleration information in step 3 specifically includes:
observer dynamics are represented by:
Figure BDA0002640204470000045
Figure BDA0002640204470000046
Figure BDA0002640204470000047
Figure BDA0002640204470000048
wherein h is0、h1、h2、z0、z1、z2Intermediate variables, with no specific meaning; sgn () is a sign function; k>0 is a design constant;
the dynamic state of the observer is solved in real time to obtain the estimation of the target acceleration
Figure BDA0002640204470000051
Comprises the following steps:
Figure BDA0002640204470000052
further, designing a guidance law in the step 4 specifically includes:
defining angular error variable
Figure BDA0002640204470000053
Is composed of
Figure BDA0002640204470000054
To pair
Figure BDA0002640204470000055
Derivative to obtain
Figure BDA0002640204470000056
Design the sliding mode variable s as
Figure BDA0002640204470000057
Wherein, mu1>0 and mu2>0 is a constant, t represents time;
acceleration of missile of
Figure BDA0002640204470000058
Wherein g >0 is a constant;
it is known that
Figure BDA0002640204470000059
Require that
Figure BDA00026402044700000510
A is an intermediate variable, with no specific meaning, calculated according to the following formula:
Figure BDA00026402044700000511
the invention is described in further detail below with reference to the figures and the embodiments.
Example 1
With reference to fig. 1, the invention provides an impact angle control guidance method based on an observer and a global sliding mode, which comprises the following steps:
step 1, establishing a bullet relative kinematics model, which comprises the following specific steps:
the relative kinematics model of the missile-target in the two-dimensional plane can be established as follows:
Figure BDA00026402044700000512
wherein, the lower corner marks T and M represent the target and the missile; a represents acceleration, V represents velocity, θ represents line of sight angle, λ and the direction of flight angles of the missile and target, respectively, relative to the line of sight (LOS); γ represents the flight path angle.
Figure BDA0002640204470000061
And
Figure BDA0002640204470000062
respectively representing the acceleration components of the missile and the target perpendicular to the LOS; vθAnd VRIndicating the relative velocity perpendicular to and along the LOS, respectively. They can be calculated according to the following formula:
Figure BDA0002640204470000063
Vθ＝VMsin-VTsinλ,VR＝VMcos-VTcosλ (3)
step 2, acquiring the motion parameters of the projectile body and the target through the missile-borne sensor, specifically:
the projectile motion parameters obtained by the projectile loading sensor are as follows: vMλ; the target motion parameters obtained by the missile-borne sensor are as follows: vT，。
Step 3, estimating target acceleration information, specifically:
observer dynamics can be represented by:
Figure BDA0002640204470000064
Figure BDA0002640204470000065
Figure BDA0002640204470000066
Figure BDA0002640204470000067
wherein h is0、h1、h2、z0、z1、z2Intermediate variables, with no specific meaning; sgn () is a sign function; k>0 is a design constant.
The observer dynamics is solved in real time, and the estimation of the target acceleration can be obtained
Figure BDA0002640204470000068
Comprises the following steps:
Figure BDA0002640204470000069
step 4, designing a guidance law, which specifically comprises the following steps:
defining angular error variable
Figure BDA00026402044700000610
Is composed of
Figure BDA00026402044700000611
To pair
Figure BDA00026402044700000612
Derivative to obtain
Figure BDA00026402044700000613
Design the sliding mode variable s as
Figure BDA00026402044700000614
Wherein, mu1>0 and mu2>0 is a design constant; t represents time.
The missile acceleration can be designed as
Figure BDA0002640204470000071
Wherein g is a design parameter. It is known that
Figure BDA0002640204470000072
Require that
Figure BDA0002640204470000073
A is an intermediate variable, with no specific meaning, and can be calculated according to the following formula:
Figure BDA0002640204470000074
step 5, inputting instructions to an actuator to control the flight;
and 6, repeating the steps 2 to 6 until the guidance is finished.
By means of a handleA collision angle control guidance method based on an observer and a global sliding mode is provided, and a guidance simulation experiment is designed. The simulation parameters are designed as follows: vM＝900,VT＝1000,0＝180°,λ0＝180°,μ1＝0.6；μ2＝0.8,τT＝0.1s,θ0＝0°，R0In the simulation, assume the target is at aTManeuver was performed at 50sin (0.7t + π). The initial conditions were set as: target coordinates: (0,0) m; missile coordinates are as follows: (15000,0) m.
Fig. 2 is a diagram of the flight trajectory of a missile during simulation, and it can be seen from fig. 2 that the proposed method results in three significantly different flight trajectories in order to obtain a desired terminal angle.
Fig. 3 is a trace diagram of a global sliding mode variable in a simulation process, and it can be seen from fig. 3 that an initial value of the sliding mode variable is always 0 and is kept near a value of 0 in the simulation process.
Fig. 4 is a view of the trajectory of the viewing angle in the simulation process, and it can be seen from fig. 4 that the missile can finally obtain the desired viewing angle.
Fig. 5 is a diagram of acceleration traces in the simulation process, and it can be seen from fig. 5 that the acceleration traces are smoother.
In conclusion, the guidance method is based on a global sliding mode control method and a fast observer, and has the characteristic of finite time convergence; by adopting a sliding mode control technology, the control of a large-range impact angle can be realized; the method is suitable for static targets and maneuvering targets and has a wide application range.

Claims (5)

1. An impact angle control guidance method based on an observer and a global sliding mode is characterized by comprising the following steps:
step 1, establishing a bullet relative kinematics model;
step 2, acquiring the motion parameters of a projectile body and a target through a projectile-borne sensor;
step 3, estimating target acceleration information;
step 4, designing a guidance law;
step 5, inputting instructions to an actuator to control the missile to fly;
and 6, repeating the steps 2 to 5 until the guidance is finished.
2. The method for controlling and guiding the collision angle based on the observer and the global sliding mode according to claim 1, wherein the establishment of the missile-target relative kinematics model in the step 1 specifically comprises the following steps:
the relative kinematics model of the missile-target in the two-dimensional plane is established as follows:
Figure FDA0002640204460000011
wherein, the lower corner marks T and M represent the target and the missile; a represents acceleration, V represents velocity, θ represents line of sight angle, λ and the direction of flight angles of the missile and target, respectively, relative to the line of sight LOS; γ represents a flight path angle;
Figure FDA0002640204460000013
and
Figure FDA0002640204460000014
respectively representing the acceleration components of the missile and the target perpendicular to the LOS; vθAnd VRRepresenting relative velocity perpendicular to and along the LOS, respectively; the calculation is made according to the following formula:
Figure FDA0002640204460000012
Vθ＝VMsin-VTsinλ,VR＝VMcos-VTcosλ (3)。
3. the collision angle control guidance method based on the observer and the global sliding mode according to claim 1, wherein the step 2 of obtaining the motion parameters of the projectile body and the target through the missile-borne sensor specifically comprises the following steps:
the projectile motion parameters obtained by the projectile loading sensor are as follows: vMλ; the target motion parameters obtained by the missile-borne sensor are as follows: vT，。
4. The method for controlling and guiding an impact angle based on an observer and a global sliding mode according to claim 1, wherein the estimating of the target acceleration information in step 3 is specifically:
observer dynamics are represented by:
Figure FDA0002640204460000021
Figure FDA0002640204460000022
Figure FDA0002640204460000023
Figure FDA0002640204460000024
wherein h is0、h1、h2、z0、z1、z2Intermediate variables, with no specific meaning; sgn () is a sign function; k>0 is a design constant;
the dynamic state of the observer is solved in real time to obtain the estimation of the target acceleration
Figure FDA00026402044600000211
Comprises the following steps:
Figure FDA0002640204460000025
5. the method for controlling and guiding an impact angle based on an observer and a global sliding mode according to claim 1, wherein the designing of the guidance law in the step 4 specifically comprises:
defining angular error variable
Figure FDA00026402044600000214
Is composed of
Figure FDA0002640204460000026
To pair
Figure FDA00026402044600000215
Derivative to obtain
Figure FDA0002640204460000027
Design the sliding mode variable s as
Figure FDA0002640204460000028
Wherein, mu1>0 and mu2>0 is a constant, t represents time;
acceleration of missile of
Figure FDA0002640204460000029
Wherein g >0 is a constant;
it is known that
Figure FDA00026402044600000212
Require that
Figure FDA00026402044600000213
A is an intermediate variable, with no specific meaning, calculated according to the following formula:
Figure FDA00026402044600000210
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Publication number Priority date Publication date Assignee Title
CN113359819A (en) * 2021-05-27 2021-09-07 北京航空航天大学 Optimal guidance law with collision angle constraint and acceleration limitation
CN113589839A (en) * 2021-07-09 2021-11-02 山东思达特测控设备有限公司 Unmanned aerial vehicle automatic collision avoidance method based on rapid finite time convergence sliding mode guidance
CN113589840A (en) * 2021-07-09 2021-11-02 山东思达特测控设备有限公司 Unmanned aerial vehicle automatic collision avoidance method based on finite time convergence sliding mode guidance

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CN110017729A (en) * 2019-04-18 2019-07-16 西安交通大学 A kind of more guided missile time coordination method of guidance with impingement angle constraint
CN110220416A (en) * 2019-05-15 2019-09-10 南京理工大学 A kind of adaptive quickly path tracking method of guidance

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CN113359819A (en) * 2021-05-27 2021-09-07 北京航空航天大学 Optimal guidance law with collision angle constraint and acceleration limitation
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CN113589839A (en) * 2021-07-09 2021-11-02 山东思达特测控设备有限公司 Unmanned aerial vehicle automatic collision avoidance method based on rapid finite time convergence sliding mode guidance
CN113589840A (en) * 2021-07-09 2021-11-02 山东思达特测控设备有限公司 Unmanned aerial vehicle automatic collision avoidance method based on finite time convergence sliding mode guidance

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