CN112033234B - Multi-missile cooperative guidance law design method - Google Patents

Multi-missile cooperative guidance law design method Download PDF

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CN112033234B
CN112033234B CN202010825332.9A CN202010825332A CN112033234B CN 112033234 B CN112033234 B CN 112033234B CN 202010825332 A CN202010825332 A CN 202010825332A CN 112033234 B CN112033234 B CN 112033234B
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CN112033234A (en
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赵启伦
李文
王晓东
宋勋
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Beijing Institute of Electronic System Engineering
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/01—Arrangements thereon for guidance or control
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/34—Direction control systems for self-propelled missiles based on predetermined target position data
    • G—PHYSICS
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Abstract

The invention discloses a method for designing a multi-missile cooperative guidance law, which divides the flight process of each missile into a primary guidance section, a middle guidance section and a final guidance section.

Description

Multi-missile cooperative guidance law design method
Technical Field
The invention relates to the field of missile launching. More particularly, the method relates to a composite guidance strategy comprising initial guidance, intermediate guidance and final guidance, and a multi-missile cooperative guidance law design method under the constraint of a falling angle and the constraint of attack time.
Background
In order to improve the penetration capability of the aircraft, the document [ 1 ] Jeon I S, Lee J I, Tahk M J.Impact-time-Control guiding law for anti-ship missiles [ J ]. IEEE Transactions on Control Systems Technology, 2006, 14(2): 260-.
Literature [ 2 ] Zhao J, Zhou S Y, Zhou r.distributed time-constrained using Nonlinear model predictive control [ J ]. Nonlinear Dynamics, 2016; 84(3): 1399-.
The document [ 4 ] Wang Y J, Dong S, Ou L L, et al, cooperative Control of multi-mission systems [ J ]. IET Control the Theory of the invention and Applications, 2014,9(3): 441) 446. by means of the idea of feedback linearization, the cooperative guidance problem is transformed into the nonlinear system consistency problem, and the cooperative attack is realized by coordinating the projectile distance and the lead angle of a plurality of aircrafts.
Document [ 5 ] Hou D L, Wang Q, Sun X J, et al, finish-time cooperative identification law for multiple participation with access establishment constraints [ J ]. IET Control Theory & Applications, 2015. the problem of finite consistency of attack times in the case of acceleration command anti-saturation was investigated.
The document [ 6 ] Wang X L, Zhang Y A, Wu H L.distributed cooperative guidance of multiple anti-ship hybrids with arbitrary impact and geometry [ J ]. Aerospace Science and Technology, 2015.
The documents [ 7 ] Sun X J, Zhou R, Hou D L, et al, consensus of leader-followers system of multi-missile with time-delays and switching topologies [ J ]. Optik-International Journal for Light and Electron Optics, 2014, 125(3): 1202-.
The guidance law in the literature [ 1-8 ] can guide multiple aircrafts to achieve 'simultaneous shooting landing' attack, but only coordination and control on hit time are considered, in practical application, in order to obtain the optimal relative information measurement effect of the bullets, detection equipment such as an optical imaging seeker and a radio frequency seeker has special requirements on the intersection angle of the bullets, and therefore the falling angle needs to be controlled; document [ 9 ] wangxiang, hongxin, linhain a method of controlling multi-bullet coordinated attack time and angle [ J ]. ballistics bullets.2012, 24 (2): 1-5, and [ 10 ] Zhao Qilun, Cheng and Li Qingdong, etc. super weapons and conventional missile cooperative attack strategy feasible domain research [ J ] aviation bulletin 2015, 36(7): 2291-.
In an actual application scene, because an aircraft fights in a three-dimensional space, the problem of aircraft guidance in the three-dimensional space needs to be discussed, in addition, acceleration instructions are generated in the whole process of missile flight according to different missile-target relative information measurement means, guidance laws of an initial guidance section, a middle guidance section and a final guidance section need to be designed respectively, so that the cost of a single aircraft is reduced as far as possible while the final attack precision is met, and the cost effectiveness ratio is improved.
Disclosure of Invention
In order to solve at least one of the technical problems proposed in the background art, the invention provides a design method of a multi-missile cooperative guidance law, which divides the flight process of each missile into a primary guidance section, a middle guidance section and a final guidance section, and comprises the following steps:
s1, establishing a space rectangular coordinate system according to the initial positions of the target and each missile, and obtaining the initial value (x) of the target in the space rectangular coordinate system T ,y T ,z T ) Τ And initial value (x) of each missile i ,y i ,z i ) Τ Wherein the subscript i represents each missile number;
s2, calculating expected trajectory deflection angle of each guided missile primary guiding section based on the initial value of the target and the initial value of each guided missile
Figure BDA0002635954700000021
And desired ballistic inclination
Figure BDA0002635954700000022
S3, judging whether each missile enters a final guidance segment at the current moment or not based on the working state of each missile guidance head, and when each missile is in the final guidance segment, determining the trajectory deflection angle psi of each missile at the current moment based on i And ballistic inclination angle theta i Calculating the acceleration instruction of each missile under the final guidance section, and when each missile is not in the final guidance section, each missile is in the intermediate guidance section or the primary guidance section;
s4, based on the trajectory deflection angle psi of each missile at the current moment i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure BDA0002635954700000023
And desired ballistic inclination
Figure BDA0002635954700000024
Judging whether each missile is in a primary guidance section or a middle guidance section at the current moment, calculating the acceleration instruction of each missile under the primary guidance section when each missile is in the primary guidance section, and calculating the acceleration instruction of each missile under the middle guidance section when each missile is in the middle guidance section;
and S5, controlling each missile to hit the target at the same time according to the acceleration instruction of each missile at the current moment.
In some possible implementations, the desired trajectory deviation angle of the primary guidance segment of each missile in S2
Figure BDA0002635954700000025
Obtained by the following formula:
Figure BDA0002635954700000026
expected ballistic inclination angle of primary guiding section of each missile in S2
Figure BDA0002635954700000027
Obtained by the following formula:
Figure BDA0002635954700000031
wherein, c i A fall angle scaling factor is desired for initial guidance.
In some possible implementations, the determining, in S3, whether each missile enters the terminal segment at the current time based on the operating state of each missile seeker includes:
each missile seeker feeds back a preset signal, and each missile enters a final guidance section.
In some possible implementations, the ballistic deflection angle ψ of each missile at the current time in S3 and S4 i Obtained by the following formula:
ψ i ＝atan2(-V z,i /V x,i )；
the trajectory inclination angle theta of each missile at the current moment in the S3 and S4 i Obtained by the following formula:
Figure BDA0002635954700000032
wherein, V x,i ,V y,i ,V z,i And the velocity components of the missile in the space rectangular coordinate system are respectively.
In some possible implementations, the calculating the acceleration command of each missile under the final guidance segment in S3 includes:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure BDA0002635954700000033
the acceleration instruction of the missile on the vertical plane is as follows:
Figure BDA0002635954700000034
wherein k is P ∈[3,6]Is the ratio of the navigation to the navigation,
Figure BDA0002635954700000035
is the rate of approach of the projectile to the eye,
Figure BDA0002635954700000036
is the line-of-sight azimuth rate,
Figure BDA0002635954700000037
is the angular velocity of the line of sight, q e,i Is the elevation angle of the line of sight, k A Greater than 0 is the proportional coefficient of the falling angle constraint term of the vertical plane, t go,i The remaining flight time of the ith missile is shown,
Figure BDA0002635954700000038
is the desired fall angle, g is the gravitational constant.
In some possible implementations, the trajectory deviation angle ψ of each missile based on the current time in S4 i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure BDA0002635954700000039
And desired ballistic inclination
Figure BDA00026359547000000310
Judging whether each missile at the current moment is in the primary guidance section or the intermediate guidance section comprises the following steps:
ballistic deflection angle psi of each missile at current moment i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure BDA00026359547000000315
And desired ballistic inclination
Figure BDA00026359547000000314
When the requirement of the following formula is met, each guided missile at the current moment is judged to be in the middle brake guide section,
Figure BDA00026359547000000312
Figure BDA00026359547000000313
wherein, delta b Error range allowed for ballistic declination, δ e Allowable error range for ballistic inclination；
Otherwise, judging that each missile is in the primary guidance section at the current moment.
In some possible implementations, the calculating the acceleration command of each missile in the preliminary guidance segment in S4 includes:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure BDA0002635954700000041
the acceleration instruction of the missile on the vertical plane is as follows:
Figure BDA00026359547000000410
wherein k is b K is a predetermined constant > 0 e And g is a gravity constant, and 0 is a preset constant.
In some possible implementations, the calculating the acceleration command of each missile in the middle guidance stage in S4 includes:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure BDA0002635954700000042
the acceleration instruction of the missile on the vertical plane is as follows:
Figure BDA0002635954700000043
wherein k is C > 0 is the time coordination term proportionality coefficient, Δ R i Is the distance of the eyes, V i Is the magnitude of the flying speed of the missile,
Figure BDA0002635954700000044
the remaining time of flight of each missile is averaged.
In some of the possible implementations of the present invention,
desired vertical face drop angle
Figure BDA0002635954700000045
The value range is as follows:
Figure BDA0002635954700000046
wherein, θ min ,θ max Respectively the minimum probing angle and the maximum probing angle of the seeker.
In some possible implementations, the time t remaining for the ith missile go,i Obtained by the following formula:
Figure BDA0002635954700000047
mean value of time of flight remaining for each missile
Figure BDA0002635954700000048
Obtained by the following formula:
Figure BDA0002635954700000049
wherein eta is e,i ＝q e,i -θ i Representing the lead angle and N is the total number of missiles.
The invention has the following beneficial effects:
the invention provides a design method of a multi-missile cooperative guidance law, which provides a possible solution for the problem of multi-missile composite cooperative guidance considering attack time and falling angle constraints in a three-dimensional space, utilizes a mode of decoupling control of a vertical surface and a transverse plane of a missile to respectively give an acceleration instruction of a vertical surface and a transverse plane of a primary, a middle and a final guidance section, can guide a plurality of missiles to simultaneously hit a target at an expected trajectory inclination angle in the three-dimensional space, has certain engineering realizability, can hit the target at the same time by a plurality of missiles at a specific falling angle, can improve the penetration capability of a warhead on a firm shelter and an armor, adopts the composite guidance strategy of primary guidance, middle guidance and final guidance, can reduce the requirement on the hardware capability of the aircraft as far as possible while ensuring the striking precision, and further improve the efficiency-cost ratio of combat
Drawings
In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings used in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained according to the drawings without creative efforts.
Fig. 1 shows a flow chart of a multi-missile cooperative guidance law design method provided by the embodiment of the invention.
Fig. 2 shows a flight trajectory diagram of each missile in a rectangular coordinate system, which is proposed by the embodiment of the invention.
Fig. 3 shows the distance between each missile and the target according to the embodiment of the present invention.
Figure 4 shows a graph of the drop angles of each of the missiles as proposed by the embodiment of the present invention.
FIG. 5 shows vertical plane acceleration commands for each missile according to an embodiment of the present invention.
FIG. 6 shows the lateral planar acceleration commands for each missile as proposed by an embodiment of the present invention.
Detailed Description
In order to make the technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
As shown in fig. 1, an embodiment of the present invention provides a method for designing a multi-missile cooperative guidance law, including:
s1, establishing a space rectangular coordinate system according to the initial positions of the target and each missile, and obtaining the initial value (x) of the target in the space rectangular coordinate system T ,y T ,z T ) Τ And each leadInitial value of bullet (x) i ,y i ,z i ) Τ Wherein the subscript i represents each missile number;
s2, calculating expected trajectory deflection angle of each guided missile primary guiding section based on the initial value of the target and the initial value of each guided missile
Figure BDA0002635954700000051
And desired ballistic inclination angle theta i * ；
In some embodiments, step S2 includes the following sub-steps
S21, expected trajectory deflection angle of primary guiding section of each missile
Figure BDA0002635954700000052
Obtained by the following formula:
Figure BDA0002635954700000053
expected trajectory inclination angle theta of primary guiding section of each missile i * Obtained by the following formula:
Figure BDA0002635954700000061
wherein, c i In order to obtain the initial guidance expected falling angle scaling coefficient, the selection can be carried out according to the initial combat battle array positions of a plurality of aircrafts in practical application. c. C i One possible range of values for (a) is as follows:
Figure BDA0002635954700000062
s3, judging whether each missile enters a final guidance segment at the current moment or not based on the working state of each missile guidance head, and when each missile is in the final guidance segment, determining the trajectory deflection angle psi of each missile at the current moment based on i And ballistic inclination angle theta i Calculating the acceleration instruction of each guided missile under the final guidance section, and when each guided missile is not in the final guidance section, each guided missile is in the middleA guidance-making section or a preliminary guidance-making section;
in some embodiments, step S3 includes the following sub-steps
S31, judging whether each missile enters a final guide section at the current moment based on the working state of each missile guide head comprises the following steps: each missile seeker feeds back a preset signal, and each missile enters a final guidance section; for example, if the seeker sends out the steady tracking state word for a plurality of consecutive times (which may be 5 times, 6 times, etc., and this is not limited in this application), the aircraft proceeds to the final leader. The method for determining whether the guided missile enters the terminal guidance section is not limited to be realized by the feedback of the guided missile seeker information, and the judgment can be carried out by judging the distance between the guided missile and the target, for example, when the distance between the guided missile and the target is smaller than a preset threshold value, the guided missile is judged to enter the terminal guidance section.
S32, ballistic deflection angle psi of each missile at current moment i And ballistic inclination angle theta i Can be obtained by the following steps:
ballistic deflection angle psi of each missile at current moment i ：
ψ i ＝atan2(-V z,i /V x,i )；
Trajectory inclination angle theta of each missile at current moment i ：
Figure BDA0002635954700000063
Wherein, V x,i ,V y,i ,V z,i Respectively are the velocity components of the missile in the navigation reference coordinate system.
S33, based on the trajectory deflection angle psi of each missile at the current moment i And ballistic inclination angle theta i Calculating the acceleration instruction of each missile under the final guidance section comprises the following steps:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure BDA0002635954700000064
the acceleration instruction of the missile on the vertical plane is as follows:
Figure BDA0002635954700000065
wherein k is P ∈[3,6]Is the ratio of the navigation to the navigation,
Figure BDA0002635954700000066
is the rate of approach of the projectile to the eye,
Figure BDA0002635954700000067
is the line-of-sight azimuth rate,
Figure BDA0002635954700000068
is the angular velocity of the line of sight, q e,i Is the elevation angle of the line of sight, k A Greater than 0 is the proportional coefficient of the falling angle constraint term of the vertical plane, t go,i The remaining flight time of the ith missile is shown,
Figure BDA0002635954700000071
is the desired fall angle and g is the gravitational constant.
Wherein the desired vertical drop angle
Figure BDA0002635954700000072
The value range is as follows:
Figure BDA0002635954700000073
wherein, theta min ,θ max Respectively the minimum probing angle and the maximum probing angle of the seeker.
In some embodiments, the time t remaining for the ith missile of the plurality of missiles go,i Can be obtained by the following formula:
Figure BDA0002635954700000074
wherein eta is e,i ＝q e,i -θ i Representing the lead angle;
s4, based on the trajectory deflection angle psi of each missile at the current moment i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure BDA0002635954700000075
And desired ballistic inclination
Figure BDA0002635954700000076
Judging whether each missile is in a primary guidance section or a middle guidance section at the current moment, calculating the acceleration instruction of each missile under the primary guidance section when each missile is in the primary guidance section, and calculating the acceleration instruction of each missile under the middle guidance section when each missile is in the middle guidance section;
in some embodiments, step S4 includes the following sub-steps
S41, ballistic deflection angle psi of each missile at current moment i Obtained by the following formula:
ψ i ＝atan2(-V z,i /V x,i )；
trajectory inclination angle theta of each missile at current moment i Obtained by the following formula:
Figure BDA0002635954700000077
wherein, V x,i ,V y,i ,V z,i Respectively is the velocity component of the missile under the navigation reference coordinate system;
s42, based on the trajectory deflection angle psi of each missile at the current moment i And ballistic inclination angle theta i Expected deviation angle of the primary guide section from the target trajectory
Figure BDA0002635954700000078
And desired ballistic inclination
Figure BDA0002635954700000079
Judging whether each missile at the current moment is in the primary guidance section or the intermediate guidance section comprises the following steps:
trajectory deviation of each missile at current momentAngle psi i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure BDA00026359547000000710
And desired ballistic inclination angle theta i * When the requirement of the following formula is met, each guided missile at the current moment is judged to be in the middle brake guide section,
Figure BDA00026359547000000711
|θ i -θ i * |＜δ e
wherein, delta b Error range allowed for ballistic declination, δ e Error range allowed for ballistic inclination angle, δ b And delta e In practice a small positive number, e.g. delta, may be used b ＝0.2°，δ e ＝0.2°；
Otherwise, judging that each missile is in the primary guidance section at the current moment.
It should be understood by those skilled in the art that after the missile enters the intermediate guidance section, the condition is no longer effective, that is, after the missile enters the intermediate guidance section, the missile only judges whether the missile is in the intermediate guidance section or the final guidance section, and the missile does not enter the initial guidance section.
S43, calculating the acceleration instruction of each missile under the primary guidance section comprises the following steps:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure BDA0002635954700000081
the acceleration instruction of the missile on the vertical plane is as follows:
a eI,i ＝-k e (θ i -θ i * )+g cos θ i
wherein k is b K is a predetermined constant > 0 e > 0 being a predetermined constant, e.g. k b ＝1，k e G is the gravity constant (g ≈ 9.807) 1.
S44, calculating the acceleration instruction of each missile under the middle guidance section comprises the following steps:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure BDA0002635954700000082
the acceleration instruction of the missile on the vertical plane is as follows:
Figure BDA0002635954700000083
wherein k is C > 0 is the time coordination term proportionality coefficient, Δ R i Is the distance of the eyes, V i Is the magnitude of the flying speed of the missile,
Figure BDA0002635954700000084
the remaining time of flight of each missile is averaged.
Mean value of time of flight remaining for each missile
Figure BDA0002635954700000085
Obtained by the following formula:
Figure BDA0002635954700000086
wherein N is the total number of missiles.
And S5, controlling each missile to hit the target at the same time according to the acceleration instruction of each missile at the current moment.
Examples
This embodiment gives a specific example to verify the effect of the present application:
setting 3 missiles to attack the target, setting the position of the target in a space coordinate system as (0, 0, 10KM),
the initial conditions for 3 missiles are shown in table 1 below:
TABLE 1 missile initial conditions
Figure BDA0002635954700000087
The guidance law parameters take the following values:
a primary guide section: k is a radical of e ＝1，k b ＝1，δ e ＝δ b ＝0.01rad；
A middle guidance section: n is 3, k P ＝4，k C ＝0.025，k A ＝2，
Figure BDA0002635954700000091
And (3) final guide section: k is a radical of P ＝4，k A ＝2。
The numerical simulation calculation results are shown in fig. 2-6, and it can be seen that the present invention can guide a plurality of aircrafts to hit a target at a specific landing angle at the same time, the variation of the generated acceleration instruction is relatively stable,
it should be understood that the above-described embodiments of the present invention are examples for clearly illustrating the invention, and are not to be construed as limiting the embodiments of the present invention, and it will be obvious to those skilled in the art that various changes and modifications can be made on the basis of the above description, and it is not intended to exhaust all embodiments, and obvious changes and modifications can be made on the basis of the technical solutions of the present invention.

Claims (9)

1. A multi-missile cooperative guidance law design method divides the flight process of each missile into a primary guidance section, a middle guidance section and a final guidance section, and is characterized by comprising the following steps:
s1, establishing a space rectangular coordinate system according to the initial positions of the target and each missile, and obtaining the initial value (x) of the target in the space rectangular coordinate system T ,y T ,z T ) T And initial value (x) of each missile i ,y i ,z i ) T Wherein, subscript i represents each missile serial number;
s2 radicalCalculating expected trajectory deflection angles of the primary guidance sections of all the missiles according to the initial values of the targets and the initial values of all the missiles
Figure FDA0003600364910000011
And desired ballistic inclination
Figure FDA0003600364910000019
Wherein the expected trajectory deviation angle of the primary guiding section of each missile
Figure FDA0003600364910000012
Obtained by the following formula:
Figure FDA0003600364910000013
expected trajectory inclination angle of primary guiding section of each missile
Figure FDA00036003649100000110
Obtained by the following formula:
Figure FDA0003600364910000014
wherein, c i Desired landing angle scaling factor for initial guidance, c i The value ranges are as follows:
Figure FDA0003600364910000015
s3, judging whether each missile enters a final guidance segment at the current moment or not based on the working state of each missile guidance head, and when each missile is in the final guidance segment, determining the trajectory deflection angle psi of each missile at the current moment based on i And ballistic inclination angle theta i Calculating the acceleration instruction of each guided missile under the final guidance section, and when each guided missile is not in the final guidance section, each guided missile is in the intermediate guidance section or the primary guidance section;
s4, based on the trajectory deflection angle psi of each missile at the current moment i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure FDA0003600364910000016
And desired ballistic inclination
Figure FDA0003600364910000017
Judging whether each missile is in a primary guidance section or a middle guidance section at the current moment, calculating an acceleration instruction of each missile under the primary guidance section when each missile is in the primary guidance section, and calculating an acceleration instruction of each missile under the middle guidance section when each missile is in the middle guidance section;
and S5, controlling each missile to hit the target at the same time according to the acceleration instruction of each missile at the current moment.
2. The method of claim 1, wherein the step S3 of determining whether each missile enters the terminal guidance segment at the current moment based on the operating state of each missile guidance head comprises:
each missile seeker feeds back a preset signal, and each missile enters a final guidance section.
3. The method of claim 1, wherein the ballistic declination ψ of each missile at the current time in S3 and S4 i Obtained by the following formula:
ψ i ＝atan2(-V z,i /V x,i )；
the trajectory inclination angle theta of each missile at the current moment in the S3 and S4 i Obtained by the following formula:
Figure FDA0003600364910000018
wherein, V x,i ,V y,i ,V z,i And the velocity components of the missile in the space rectangular coordinate system are respectively.
4. The method of claim 3, wherein the step of calculating the acceleration command of each missile under the final guidance segment in S3 comprises the steps of:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure FDA0003600364910000021
the acceleration instruction of the missile on the vertical plane is as follows:
Figure FDA0003600364910000022
wherein k is P ∈[3,6]Is the ratio of the navigation to the navigation,
Figure FDA0003600364910000023
is the rate of approach of the projectile to the eye,
Figure FDA0003600364910000024
is the line-of-sight azimuth rate,
Figure FDA0003600364910000025
is the angular velocity of the line of sight, q e,i Is the elevation angle of the line of sight, k A Greater than 0 is the proportional coefficient of the falling angle constraint term of the vertical plane, t go,i The remaining flight time of the ith missile is shown,
Figure FDA0003600364910000026
is the desired fall angle and g is the gravitational constant.
5. The method of claim 4, wherein the S4 is based on the ballistic deflection angle ψ of each missile at the current time i And ballistic inclination angle theta i Desired ballistic declination from the primary guidance segment
Figure FDA0003600364910000027
And desired ballistic inclination
Figure FDA00036003649100000213
Judging whether each missile at the current moment is in the primary guidance section or the intermediate guidance section comprises the following steps:
ballistic deflection angle psi of each missile at current moment i And ballistic inclination angle theta i Desired ballistic declination from the primary guiding section
Figure FDA0003600364910000028
And desired ballistic inclination
Figure FDA00036003649100000214
When the requirement of the following formula is met, each guided missile at the current moment is judged to be in the middle brake guide section,
Figure FDA0003600364910000029
Figure FDA00036003649100000210
wherein, delta b Error range allowed for ballistic declination, δ e The allowable error range of the ballistic inclination angle;
otherwise, judging that each missile is in the primary guidance section at the current moment.
6. The method of claim 5, wherein the step of calculating the acceleration command of each missile in the initial guidance segment in S4 comprises the following steps:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure FDA00036003649100000211
the acceleration instruction of the missile on the vertical plane is as follows:
Figure FDA00036003649100000215
wherein k is b K is a predetermined constant > 0 e And g is a gravity constant, and 0 is a preset constant.
7. The method of claim 6, wherein the step of calculating the acceleration commands of the missiles in the middle guidance distance in S4 comprises the steps of:
the acceleration instruction of the missile on the transverse lateral plane is as follows:
Figure FDA00036003649100000212
the acceleration instruction of the missile on the vertical plane is as follows:
Figure FDA0003600364910000031
wherein k is C > 0 is the time coordination term proportionality coefficient, Δ R i Is the distance of the eyes, V i Is the magnitude of the flying speed of the missile,
Figure FDA0003600364910000032
the remaining time of flight of each missile is averaged.
8. The method of claim 7,
desired vertical face drop angle
Figure FDA0003600364910000033
The value range is as follows:
Figure FDA0003600364910000034
wherein, theta min ,θ max Respectively the minimum probing angle and the maximum probing angle of the seeker.
9. The method of claim 7, wherein the time t remaining for the ith missile go,i Obtained by the following formula:
Figure FDA0003600364910000035
mean value of time of flight remaining for each missile
Figure FDA0003600364910000036
Obtained by the following formula:
Figure FDA0003600364910000037
wherein eta is e,i ＝q e,i -θ i Representing the lead angle and N is the total number of missiles.
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CN112731965B (en) * 2020-12-17 2022-04-08 哈尔滨工业大学 Guidance method based on target maneuver identification
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Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5435503A (en) * 1993-08-27 1995-07-25 Loral Vought Systems Corp. Real time missile guidance system
GB201300848D0 (en) * 2013-01-15 2013-07-24 Mbda Uk Ltd A missile seeker and guidance method
CN103245256A (en) * 2013-04-25 2013-08-14 北京理工大学 Multi-missile cooperative attack guidance law designing method
CN106843265A (en) * 2016-12-30 2017-06-13 哈尔滨工业大学 Three-dimensional many guided missile cooperative guidance method and systems of finite time convergence control
CN108168381A (en) * 2018-01-04 2018-06-15 北京理工大学 A kind of control method of more pieces of guided missile cooperations
CN109597423A (en) * 2019-01-08 2019-04-09 北京航空航天大学 A kind of design method of the multiple constraint Terminal Guidance Laws based on reference sight angle signal
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
CN111306998A (en) * 2020-03-24 2020-06-19 北京理工大学 Parameter perturbation self-adaptive guided rocket projectile vertical attack guidance method

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5435503A (en) * 1993-08-27 1995-07-25 Loral Vought Systems Corp. Real time missile guidance system
GB201300848D0 (en) * 2013-01-15 2013-07-24 Mbda Uk Ltd A missile seeker and guidance method
CN103245256A (en) * 2013-04-25 2013-08-14 北京理工大学 Multi-missile cooperative attack guidance law designing method
CN106843265A (en) * 2016-12-30 2017-06-13 哈尔滨工业大学 Three-dimensional many guided missile cooperative guidance method and systems of finite time convergence control
CN108168381A (en) * 2018-01-04 2018-06-15 北京理工大学 A kind of control method of more pieces of guided missile cooperations
CN109597423A (en) * 2019-01-08 2019-04-09 北京航空航天大学 A kind of design method of the multiple constraint Terminal Guidance Laws based on reference sight angle signal
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
CN111306998A (en) * 2020-03-24 2020-06-19 北京理工大学 Parameter perturbation self-adaptive guided rocket projectile vertical attack guidance method

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