CN110645843B - High-dynamic compensation guidance control system and method for high-speed maneuvering target - Google Patents

High-dynamic compensation guidance control system and method for high-speed maneuvering target Download PDF

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CN110645843B
CN110645843B CN201910759932.7A CN201910759932A CN110645843B CN 110645843 B CN110645843 B CN 110645843B CN 201910759932 A CN201910759932 A CN 201910759932A CN 110645843 B CN110645843 B CN 110645843B
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aircraft
maneuvering target
dynamic aircraft
relative
speed maneuvering
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CN110645843A (en
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王伟
赵健廷
南宇翔
曹先彬
杜文博
肖振宇
纪毅
王雨辰
师兴伟
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Beihang University
Beijing Institute of Technology BIT
China North Industries Corp
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Beihang University
Beijing Institute of Technology BIT
China North Industries Corp
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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
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Abstract

The invention discloses a high dynamic compensation guidance control system and method for a high speed maneuvering target, the system comprises a measuring module (1), a processing module (2) and an executing module (3), the measuring module (1) is used for measuring the relative position information of a high dynamic aircraft and the high speed maneuvering target and the attitude information of the high dynamic aircraft in real time, the processing module (2) is used for obtaining a rudder deflection instruction signal according to the information measured by the measuring module in real time, and the executing module (3) is used for receiving the rudder deflection instruction signal obtained by the processing module and converting the signal into a signal form required by a steering engine. According to the system provided by the invention, the real-time relative position information of the aircraft and the target is obtained through the active radar, the attitude information of the aircraft is obtained through the MEMS gyroscope and the geomagnetic sensor, and the relative acceleration of the aircraft and the target is obtained through calculation of the airborne microprocessor, so that the guidance control instruction of the high-dynamic aircraft is compensated, and the tracking performance of the aircraft is improved.

Description

High-dynamic compensation guidance control system and method for high-speed maneuvering target
Technical Field
The invention belongs to the field of automatic control, relates to a compensation guidance control system and method for a high-dynamic aircraft, and particularly relates to a high-dynamic compensation guidance control system and method for a high-speed maneuvering target.
Background
The high dynamic aircraft is the backbone strength of military forces of various countries, and has the capabilities of accurate striking, efficient damage, battlefield survival and the like. The tracking targets of the high-dynamic aircraft are mostly high-speed maneuvering targets, and due to the fact that the maneuvering targets and the aircraft have high relative movement speed and maneuverability, a serious challenge is brought to a guidance system of the aircraft.
In the prior art, a guidance system of a high-dynamic aircraft mostly adopts a proportional guidance law or an overweight compensation proportional guidance law, guidance control instructions of the high-dynamic aircraft are not compensated or are fixedly compensated, real-time prediction of motion information of a high-speed maneuvering target is difficult to achieve, the tracking performance, the target maneuvering adapting capacity and the tracking precision of the aircraft are poor, and the development of the high-dynamic aircraft is severely restricted.
Therefore, designing a high dynamic compensation guidance control system and method for a high speed maneuvering target to make a high dynamic aircraft have the capability of adapting to target maneuvering and high tracking accuracy is a technical problem to be solved at present.
Disclosure of Invention
In order to overcome the problems, the inventor of the present invention has conducted intensive research to design a high dynamic compensation guidance control system and method for a high speed maneuvering target, where the system includes a measurement module for measuring the relative position information of a high dynamic aircraft and the high speed maneuvering target and the attitude information of the high dynamic aircraft in real time, a processing module for obtaining a rudder deflection command signal according to the information measured by the measurement module in real time, and an execution module for receiving the rudder deflection command signal obtained by the processing module and converting the signal into a signal form required by a steering engine. The real-time relative position information of the aircraft and the target is obtained through the active radar, the attitude information of the aircraft is obtained through the MEMS gyroscope and the geomagnetic sensor, the relative acceleration of the aircraft and the target is obtained through calculation of the airborne microprocessor, the relative acceleration is used for compensating a guidance control instruction of the high-dynamic aircraft, and the tracking performance of the aircraft is improved, so that the invention is completed.
In particular, the invention aims to provide a high dynamic compensation guidance control system for high speed maneuvering targets, wherein the system comprises a measuring module 1, a processing module 2 and an execution module 3, wherein,
the measuring module 1 is used for measuring the relative position information of the high-dynamic aircraft and the high-speed maneuvering target and the attitude information of the high-dynamic aircraft in real time,
the processing module 2 is used for obtaining a rudder deflection instruction signal according to the information measured by the measuring module in real time,
and the execution module 3 is used for receiving the rudder deflection command signal obtained by the processing module and converting the signal into a signal form required by the steering engine.
Wherein the measuring module 1 comprises a relative position information measuring submodule 11 and an aircraft attitude information measuring submodule 12, wherein,
the relative position information measuring submodule 11 is used for measuring the position information of the high-dynamic aircraft relative high-speed maneuvering target in real time,
the aircraft attitude information measurement submodule 12 is used for obtaining attitude information of a high-dynamic aircraft in real time.
The relative position information measuring submodule 11 is a radar measuring element, and includes a radar transmitter 111, a radar receiver 112 and an information processor 113;
wherein, the radar transmitter 111 is used for transmitting microwave or millimeter wave to the high-speed maneuvering target,
the radar receiver 112 is used for receiving echo information of a high-speed maneuvering target,
the information processor 113 is configured to process the received echo information to obtain position information of the high-dynamic aircraft relative to the high-speed maneuvering target.
Wherein the aircraft attitude information measuring submodule 12 comprises a MEMS gyroscope 121 and a geomagnetic sensor 122, wherein,
the MEMS gyroscope 121 is used for measuring the pitch angle of the high-dynamic aircraft
Figure BDA0002169949270000031
And the yaw angle psi, and the yaw angle,
the geomagnetic sensor 122 is used for measuring a roll angle γ of the high-dynamic aircraft.
Wherein the processing module 2 comprises a relative position calculating submodule 21, a relative speed calculating submodule 22, a relative acceleration calculating submodule 23, a compensation guidance instruction calculating submodule 24 and a rudder deflection instruction calculating submodule 25, wherein,
the relative position resolving submodule 21 is used for obtaining the relative position of the high-dynamic aircraft and the high-speed maneuvering target under a ground coordinate system according to the measured distance, the sight declination angle and the sight inclination angle of the high-dynamic aircraft relative to the high-speed maneuvering target in real time;
the relative speed calculating submodule 22 is used for obtaining the relative speed of the high-dynamic aircraft and the high-speed maneuvering target according to the relative position of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system in real time;
the relative acceleration resolving submodule 23 is used for obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target according to the relative speed of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system in real time;
the compensation guidance instruction resolving submodule 24 is used for obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system in real time according to the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system, and the relative acceleration is information of a compensation aircraft guidance control instruction;
the rudder deflection instruction resolving submodule 25 is used for obtaining a rudder deflection instruction signal in real time according to the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system.
Wherein the relative position calculating submodule 21 obtains the relative position of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the following formula (I),
Figure BDA0002169949270000041
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA00021699492700000415
denotes the inclination of the line of sight, and θ denotes the declination of the line of sight.
Wherein, the relative speed calculating submodule 22 obtains the relative speed of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the following formula (II),
Figure BDA0002169949270000042
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA0002169949270000043
the first differential of the expression r is shown,
Figure BDA0002169949270000044
the inclination of the line of sight is indicated,
Figure BDA0002169949270000045
to represent
Figure BDA00021699492700000416
The first differential of the time domain is,
theta represents the declination of the line of sight,
Figure BDA0002169949270000046
the first differential of the theta is represented,
Figure BDA0002169949270000047
the first differential of x is represented by,
Figure BDA0002169949270000048
the first differential of y is represented by,
Figure BDA0002169949270000049
represents the first differential of z.
Wherein, the relative acceleration resolving submodule 23 obtains the relative acceleration of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the following formula (III),
Figure BDA00021699492700000410
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA00021699492700000411
the first differential of the expression r is shown,
Figure BDA00021699492700000412
the second derivative of r is represented as,
Figure BDA00021699492700000417
the inclination of the line of sight is indicated,
Figure BDA00021699492700000413
to represent
Figure BDA00021699492700000418
The first differential of the time domain is,
Figure BDA00021699492700000414
to represent
Figure BDA00021699492700000419
The second order differential of (a) is,
theta represents the declination of the line of sight,
Figure BDA0002169949270000051
the first differential of the theta is represented,
Figure BDA0002169949270000052
the second derivative of theta is represented by,
Figure BDA0002169949270000053
the second derivative of x is represented as,
Figure BDA0002169949270000054
the second derivative of y is represented by,
Figure BDA0002169949270000055
represents the second derivative of z.
The invention also provides a high dynamic compensation guidance control method for a high speed maneuvering target, which is preferably realized by the system of the first aspect, wherein the method comprises the following steps:
step 1, obtaining the relative position information of a high-dynamic aircraft and a high-speed maneuvering target and the attitude information of the high-dynamic aircraft in real time through a measurement module 1;
step 2, obtaining a rudder deflection instruction signal of the high-dynamic aircraft in real time through the processing module 2;
and 3, converting the rudder deflection command signal in real time through the execution module 3 so as to control the steering engine to move.
Wherein, step 2 comprises the following substeps:
step 2-1, obtaining the relative position of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the relative position resolving submodule 21;
2-2, obtaining the relative speed of the high-dynamic aircraft and the high-speed maneuvering target in the ground coordinate system in real time through the relative speed resolving submodule 22;
2-3, obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the ground coordinate system in real time through the relative acceleration resolving submodule 23;
2-4, obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in a projectile coordinate system in real time through a compensation guidance instruction resolving submodule 24;
and 2-5, obtaining a rudder deflection instruction signal of the high-dynamic aircraft in real time through the rudder deflection instruction resolving submodule 25.
The invention has the advantages that:
(1) according to the high-dynamic compensation guidance control system for the high-speed maneuvering target, the real-time relative position information of the aircraft and the target is obtained through the active radar, the attitude information of the aircraft is obtained through the MEMS gyroscope and the geomagnetic sensor, the relative acceleration of the aircraft and the target is obtained through calculation of the airborne microprocessor, the guidance control instruction of the high-dynamic aircraft is compensated, and the tracking performance of the aircraft is improved;
(2) according to the high-dynamic compensation guidance control method for the high-speed maneuvering target, provided by the invention, the relative motion information between the high-dynamic aircraft and the high-speed maneuvering target is used as feedback and is used as a dynamic compensation item of a proportional guidance law, so that the high-dynamic aircraft has the capability of adapting to target maneuvering and is high in tracking precision.
Drawings
FIG. 1 shows an overall logic diagram of a high dynamic compensation guidance control system for a high speed maneuvering target according to a preferred embodiment of the invention;
FIG. 2 is a diagram illustrating the relative position information of a high dynamic vehicle and a high speed maneuvering target according to a preferred embodiment of the invention;
FIG. 3 illustrates a trajectory curve for tracking a stationary target in a simulation experiment;
FIG. 4 shows a trajectory curve for tracking a uniform motion target in a simulation experiment;
FIG. 5 shows a trajectory curve for tracking a high-speed maneuver target in a simulation experiment;
FIG. 6 shows a flow chart of the steps of a high dynamic compensation guidance control method for high speed maneuvering targets according to a preferred embodiment of the invention.
The reference numbers illustrate:
1-a measurement module;
11-relative position information measurement submodule;
111-a radar transmitter;
112-a radar receiver;
113-an information processor;
12-an aircraft attitude information measurement submodule;
121-MEMS gyroscope;
122-a geomagnetic sensor;
2-a processing module;
21-relative position resolving submodule;
22-relative speed resolving submodule;
23-relative acceleration resolving submodule;
24-compensation guidance instruction resolving submodule;
25-a rudder deflection instruction resolving submodule;
3-an execution module;
31-an amplifying transformer;
32-a steering engine;
4-a power supply module;
m-a high dynamic aircraft;
t-high speed maneuver targets.
Detailed Description
The present invention will be described in further detail below with reference to the accompanying drawings and embodiments. The features and advantages of the present invention will become more apparent from the description. In which, although various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
The invention provides a high-dynamic compensation guidance control system for a high-speed maneuvering target, which is arranged on a high-dynamic aircraft to predict and compensate a guidance control command signal in real time as shown in figure 1,
the system comprises a measuring module 1, a processing module 2 and an execution module 3, wherein,
the measuring module 1 is used for measuring the relative position information of the high-dynamic aircraft and the high-speed maneuvering target and the attitude information of the high-dynamic aircraft in real time,
the processing module 2 is used for obtaining a rudder deflection instruction signal according to the information measured by the measuring module in real time,
and the execution module 3 is used for receiving the rudder deflection command signal obtained by the processing module and converting the signal into a signal form required by the steering engine.
The high dynamic state means that the aircraft can fly with large maneuvering and has large normal acceleration (generally, the flying condition with the normal acceleration of more than 10g is called large maneuvering and g represents gravity acceleration).
In the invention, the high dynamic aircraft refers to a rotating aircraft with the speed of more than 800m/s and the rotating speed of more than 15 r/s.
According to a preferred embodiment of the invention, the measurement module 1 comprises a relative position information measurement submodule 11 and an aircraft attitude information measurement submodule 12, wherein,
the relative position information measuring submodule 11 is used for measuring the position information of the high-dynamic aircraft relative high-speed maneuvering target in real time,
the aircraft attitude information measurement submodule 12 is used for obtaining attitude information of a high-dynamic aircraft in real time.
In a further preferred embodiment, the relative position information measured by the relative position information measurement sub-module 11 includes the distance, the view declination angle and the view inclination angle of the high-dynamic aircraft relative to the high-speed maneuvering target;
the attitude information of the high-dynamic aircraft comprises a pitch angle, a yaw angle and a roll angle.
In the present invention, the position information of the high dynamic aircraft relative to the high speed maneuvering target is shown in fig. 2, in which the moving coordinate system takes the centroid of the aircraft M as the origin, and the x, y, and z axes are respectively parallel to the x, y, and z axes of the ground coordinate system. The distance between the mass center of the aircraft M and the mass center of the high-speed maneuvering target T is the distance r of the high-dynamic aircraft relative to the high-speed maneuvering target; the included angle between the distance r and the projection of the distance r on the horizontal plane is the inclination angle of the sight line
Figure BDA0002169949270000081
Projection of distance r on horizontal plane and x-axisThe included angle between the two is the declination angle theta of the sight line.
According to a preferred embodiment of the present invention, the relative position information measuring submodule 11 is a radar measuring element, and includes a radar transmitter 111, a radar receiver 112 and an information processor 113;
wherein, the radar transmitter 111 is used for transmitting microwave or millimeter wave to the high-speed maneuvering target,
the radar receiver 112 is used for receiving echo information of a high-speed maneuvering target,
the information processor 113 is configured to process the received echo information to obtain position information of the high-dynamic aircraft relative to the high-speed maneuvering target.
In the invention, the radar measuring element is an active radar measuring element, and the relative position information (distance and angle) of the high-dynamic aircraft and the high-speed maneuvering target can be effectively measured by adopting an active radar measuring method. The active radar measurement method adopts a radar self-seeking guidance system, and comprises microwave radar self-seeking guidance and millimeter wave radar self-seeking guidance.
The microwave active seeking guidance can independently capture and track a target, the closer to the target, the stronger the resolution capability of the target on the angular position, and the higher the guidance accuracy.
Millimeter wave guidance is mainly used for accurate guidance, millimeter waves generally refer to electromagnetic waves with the wavelength of 1-10 mm, the corresponding frequency is 30-300 GHz, the wavelength and the frequency of the millimeter waves are between the microwave band and the infrared band, the inherent characteristics of the two bands are combined, and the millimeter wave guidance is an ideal selection band for a high-performance guidance system. The millimeter wave guidance system has less loss through the atmosphere; the guidance equipment has small volume and light weight; the measurement precision is high, and the resolution capability is strong; the anti-interference capability is strong; strong capability of identifying metal targets, and the like.
In the invention, the signal processor processes the received echo information to obtain the information about the target and the electromagnetic environment, and further analyzes and logically operates the information to form a control command.
In a further preferred embodiment, the aircraft attitude information measurement submodule 12 includes a MEMS gyroscope 121 and a geomagnetic sensor 122, wherein,
the MEMS gyroscope 121 is used for measuring the pitch angle of the high-dynamic aircraft
Figure BDA0002169949270000091
And the yaw angle psi, and the yaw angle,
the geomagnetic sensor 122 is used for measuring a roll angle γ of the high-dynamic aircraft.
According to a preferred embodiment of the invention, the processing module 2 comprises a relative position calculating submodule 21, a relative speed calculating submodule 22, a relative acceleration calculating submodule 23, a compensation guidance instruction calculating submodule 24 and a rudder deflection instruction calculating submodule 25, wherein,
the relative position resolving submodule 21 is used for obtaining the relative position of the high-dynamic aircraft and the high-speed maneuvering target under a ground coordinate system according to the measured distance, the sight declination angle and the sight inclination angle of the high-dynamic aircraft relative to the high-speed maneuvering target in real time;
the relative speed calculating submodule 22 is used for obtaining the relative speed of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system according to the relative position of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system in real time;
the relative acceleration resolving submodule 23 is used for obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system in real time according to the relative speed of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system;
the compensation guidance instruction resolving submodule 24 is used for obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system in real time according to the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system, and the relative acceleration is information of a compensation aircraft guidance control instruction;
the rudder deflection instruction resolving submodule 25 is used for obtaining a rudder deflection instruction signal in real time according to the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system.
In a further preferred embodiment, the relative position calculating submodule 21 obtains the relative position of the high-dynamic aircraft and the high-speed maneuvering target in the ground coordinate system in real time by the following formula (one),
Figure BDA0002169949270000101
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA0002169949270000102
denotes the inclination of the line of sight, and θ denotes the declination of the line of sight.
Preferably, the relative speed calculating submodule 22 obtains the relative speed of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the following formula (II),
Figure BDA0002169949270000111
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA0002169949270000113
the first differential of the expression r is shown,
Figure BDA00021699492700001120
the inclination of the line of sight is indicated,
Figure BDA0002169949270000114
to represent
Figure BDA00021699492700001121
The first differential of the time domain is,
theta represents the declination of the line of sight,
Figure BDA0002169949270000115
the first differential of the theta is represented,
Figure BDA0002169949270000116
the first differential of x is represented by,
Figure BDA0002169949270000117
the first differential of y is represented by,
Figure BDA0002169949270000118
represents the first differential of z;
wherein the content of the first and second substances,
Figure BDA0002169949270000119
and
Figure BDA00021699492700001110
namely the relative speed of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system.
More preferably, the relative acceleration calculating submodule 23 obtains the relative acceleration of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the following formula (three),
Figure BDA0002169949270000112
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA00021699492700001111
the first differential of the expression r is shown,
Figure BDA00021699492700001112
the second derivative of r is represented as,
Figure BDA00021699492700001122
the inclination of the line of sight is indicated,
Figure BDA00021699492700001113
to represent
Figure BDA00021699492700001123
The first differential of the time domain is,
Figure BDA00021699492700001114
to represent
Figure BDA00021699492700001124
The second order differential of (a) is,
theta represents the declination of the line of sight,
Figure BDA00021699492700001115
the first differential of the theta is represented,
Figure BDA00021699492700001116
the second derivative of theta is represented by,
Figure BDA00021699492700001117
the second derivative of x is represented as,
Figure BDA00021699492700001118
the second derivative of y is represented by,
Figure BDA00021699492700001119
the second derivative of z is represented by,
wherein the content of the first and second substances,
Figure BDA0002169949270000123
and
Figure BDA0002169949270000124
namely the relative acceleration of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system.
According to a preferred embodiment of the invention, the compensation guidance instruction resolving submodule 24 obtains the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the missile coordinate system in real time through the following formulas (four) and (five),
Figure BDA0002169949270000121
Figure BDA0002169949270000122
wherein gamma represents the roll angle of the high dynamic aircraft,
Figure BDA0002169949270000125
representing the pitch angle of the high dynamic aircraft, and psi representing the yaw angle of the high dynamic aircraft;
Figure BDA0002169949270000126
the second derivative of x is represented as,
Figure BDA0002169949270000127
the second derivative of y is represented by,
Figure BDA0002169949270000128
the second derivative of z is represented by,
Figure BDA0002169949270000129
and
Figure BDA00021699492700001210
namely the relative acceleration of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system;
ax1representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the direction of the body axis of the aircraft in a missile coordinate system, ay1Representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the longitudinal symmetry plane of the aircraft and the direction vertical to the body axis under the elastic body coordinate system, az1The relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the direction perpendicular to the two directions under the elastic coordinate system is shown.
In a further preferred embodiment, the rudder deflection command resolving submodule 25 obtains a rudder deflection command signal by the following equation (six),
Figure BDA0002169949270000131
wherein, aMθIs highAcceleration of the dynamic aircraft in the direction of the declination of the line of sight, aMφThe acceleration of the high dynamic aircraft in the direction of the line of sight inclination is represented, N represents a guidance coefficient, N is generally equal to 4,
Figure BDA0002169949270000132
the first differential of the expression r is shown,
Figure BDA0002169949270000133
the first differential of the theta is represented,
Figure BDA0002169949270000134
to represent
Figure BDA0002169949270000135
First order differential of ay1Representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the longitudinal symmetry plane of the aircraft and the direction vertical to the body axis under the elastic body coordinate system, az1The relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the direction perpendicular to the two directions under the elastic coordinate system is shown.
Because the acceleration along the speed direction of the aircraft is not controllable, and the proportion guidance law design a is usually adopted in engineeringMθAnd aMφSo that
Figure BDA0002169949270000136
And
Figure BDA0002169949270000137
and the steering deviation signal is obtained through the formula (six) in the invention.
In the invention, the processing module realizes the function of compensating the high dynamic aircraft guidance control instruction in real time through an airborne microprocessor, and the microprocessor is a common device in the prior art and is not particularly limited. In particular, the main chip on the microprocessor may resolve relative position information, relative velocity information, and relative acceleration information between the aircraft and the target in the ground coordinate system. After the coordinate system is converted, the relative acceleration information in the missile coordinate system is utilized to compensate the guidance control instruction of the high-dynamic aircraft, so that the aircraft has the capability of adapting to target maneuver, and the tracking precision is obviously improved.
According to a preferred embodiment of the invention, the actuator module 3 comprises an amplifying transducer 31 and a steering engine 32, wherein,
the amplification converter 31 is used for amplifying the rudder deflection instruction signal output by the processing module and converting the rudder deflection instruction signal into a signal form required by the steering engine;
the steering engine 32 moves according to the signal output by the amplification converter.
In the invention, the execution module receives the rudder deflection command signal output by the processing module, firstly amplifies the signal through the amplification converter, and converts the signal into a signal form required by the steering engine according to the type of the steering engine. The steering engine can generate enough rotation torque under the action of the output signal of the amplification changer, overcomes the reaction torque of the control surface, and enables the control surface to deflect rapidly or fix the control surface at a required angle.
In a further preferred embodiment, the steering engine 32 is an air-cooled steering engine.
The cold air type steering engine controls the high-pressure gas valve according to the rudder deflection command signal, so that high-pressure gas (high-pressure air or helium) pushes the actuating device, and the movement of the control surface is controlled.
According to a preferred embodiment of the invention, the high dynamic compensation guidance control system for the high-speed maneuvering target further comprises a power supply module 4, so as to provide the required rated voltage for the measurement module 1 and the processing module 2, and ensure the continuous and stable operation of the whole system.
In the invention, the power supply module can also provide a stable working environment according to the power consumption requirements of the radar part device, the MEMS gyroscope and the geomagnetic sensor in the measurement module and the rated voltage of the microprocessor in the processing module.
Meanwhile, the power supply module is connected to a thermal power supply on the aircraft, and the whole circuit is checked to prevent the components from being burnt out due to the occurrence of problems such as local short circuit.
The invention also provides a high dynamic compensation guidance control method for a high-speed maneuvering target, which is preferably realized by the system, as shown in FIG. 6, and comprises the following steps:
step 1, obtaining the relative position information of the high-dynamic aircraft and the high-speed maneuvering target and the attitude information of the high-dynamic aircraft in real time through a measuring module 1.
Wherein, step 1 comprises the following substeps:
step 1-1, measuring the position information of the high-dynamic aircraft relative to the high-speed maneuvering target in real time through a relative position information measuring submodule 11.
The relative position information measuring submodule 11 is a radar measuring element, and includes a radar transmitter 111, a radar receiver 112 and an information processor 113;
the radar transmitter 111 is used for transmitting microwaves or millimeter waves to a high-speed maneuvering target,
the radar receiver 112 is used for receiving echo information of a high-speed maneuvering target,
the information processor 113 is configured to process the received echo information to obtain position information of the high-dynamic aircraft relative to the high-speed maneuvering target.
Step 1-2, measuring the attitude information of the high-dynamic aircraft in real time through the aircraft attitude information measuring submodule 12.
The aircraft attitude information measurement sub-module 12 includes a MEMS gyro 121 and a geomagnetic sensor 122,
the MEMS gyroscope 121 is used for measuring the pitch angle of the high-dynamic aircraft
Figure BDA0002169949270000151
And the yaw angle psi, and the yaw angle,
the geomagnetic sensor 122 is used for measuring a roll angle γ of the high-dynamic aircraft.
And 2, acquiring a rudder deflection command signal of the high-dynamic aircraft in real time through the processing module 2.
Wherein the step 2 comprises the following substeps:
step 2-1, obtaining the relative position of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through the relative position resolving submodule 21;
2-2, obtaining the relative speed of the high-dynamic aircraft and the high-speed maneuvering target in the ground coordinate system in real time through the relative speed resolving submodule 22;
2-3, obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the ground coordinate system in real time through the relative acceleration resolving submodule 23;
2-4, obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in a projectile coordinate system in real time through a compensation guidance instruction resolving submodule 24;
and 2-5, obtaining a rudder deflection instruction signal of the high-dynamic aircraft in real time through the rudder deflection instruction resolving submodule 25.
And 3, converting the rudder deflection command signal in real time through the execution module 3 so as to control the steering engine to move.
In the invention, the relative position information of the high-dynamic aircraft and the high-speed maneuvering target can be effectively measured by adopting an active radar measuring method, the relative acceleration information between the aircraft and the target is further obtained, the motion information is used as feedback, and the guidance control instruction of the aircraft is compensated in real time, so that the aircraft has the capability of adapting to target maneuvering, and the tracking precision is obviously improved.
Experimental example:
experimental example 1
A simulation experiment of tracking a static target by a high-dynamic aircraft is carried out through a computer, wherein the simulation conditions are as follows: the flying speed of the high dynamic aircraft is 800m/s, and the rotating speed is 15 r/s.
The conditions of the high dynamic compensation guidance control instruction method, the proportional guidance method without the compensation term and the proportional guidance method with the fixed compensation when the static target is tracked are simulated respectively, and the result is shown in figure 3.
Wherein the content of the first and second substances,
(I) the invention discloses a high dynamic compensation guidance control instruction method, which is a high dynamic compensation guidance control method aiming at a high speed maneuvering target, and comprises the following specific steps:
(1) the high dynamic aircraft transmits microwaves to the static target through the radar transmitter, then the radar receiver receives echo information and transmits the echo information to the information processor for processing, and the distance r, the sight declination angle theta and the sight inclination angle of the high dynamic aircraft relative to the static target are obtained
Figure BDA0002169949270000162
And transmitting to a processing module;
(2) measuring the pitch angle of the high-dynamic aircraft through the MEMS gyroscope
Figure BDA0002169949270000161
The yaw angle psi is measured through the geomagnetic sensor, and the roll angle gamma of the high-dynamic aircraft is transmitted to the processing module;
(3) the processing module sequentially obtains the relative position of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system, the relative speed of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system, the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system and a rudder deflection instruction signal according to the information measured in the steps (1) and (2);
wherein, the relative position of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system is obtained in real time through the following formula (I),
Figure BDA0002169949270000171
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA00021699492700001711
the inclination angle of the sight line is shown, and theta represents the declination angle of the sight line;
the relative speed of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system is obtained in real time through the following formula (II),
Figure BDA0002169949270000172
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA0002169949270000173
the first differential of the expression r is shown,
Figure BDA00021699492700001712
the inclination of the line of sight is indicated,
Figure BDA0002169949270000174
to represent
Figure BDA00021699492700001713
The first differential of the time domain is,
theta represents the declination of the line of sight,
Figure BDA0002169949270000175
the first differential of the theta is represented,
Figure BDA0002169949270000176
the first differential of x is represented by,
Figure BDA0002169949270000177
the first differential of y is represented by,
Figure BDA0002169949270000178
represents the first differential of z;
wherein the content of the first and second substances,
Figure BDA0002169949270000179
and
Figure BDA00021699492700001710
namely the relative speed of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system;
the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system is obtained in real time through the following formula (III),
Figure BDA0002169949270000181
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure BDA0002169949270000184
the first differential of the expression r is shown,
Figure BDA0002169949270000185
the second derivative of r is represented as,
Figure BDA0002169949270000188
the inclination of the line of sight is indicated,
Figure BDA0002169949270000186
to represent
Figure BDA00021699492700001820
The first differential of the time domain is,
Figure BDA0002169949270000187
to represent
Figure BDA00021699492700001821
The second order differential of (a) is,
theta represents the declination of the line of sight,
Figure BDA0002169949270000189
the first differential of the theta is represented,
Figure BDA00021699492700001810
the second derivative of theta is represented by,
Figure BDA00021699492700001811
the second derivative of x is represented as,
Figure BDA00021699492700001812
the second derivative of y is represented by,
Figure BDA00021699492700001813
the second derivative of z is represented by,
wherein the content of the first and second substances,
Figure BDA00021699492700001814
and
Figure BDA00021699492700001815
namely the relative acceleration of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system;
the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the elastic body coordinate system is obtained in real time through the following formulas (four) and (five),
Figure BDA0002169949270000182
Figure BDA0002169949270000183
wherein gamma represents the roll angle of the high dynamic aircraft,
Figure BDA00021699492700001816
representing the pitch angle of the high dynamic aircraft, and psi representing the yaw angle of the high dynamic aircraft;
Figure BDA00021699492700001817
the second derivative of x is represented as,
Figure BDA00021699492700001818
the second derivative of y is represented by,
Figure BDA00021699492700001819
the second derivative of z is represented by,
Figure BDA0002169949270000192
and
Figure BDA0002169949270000193
namely the relative acceleration of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system;
ax1representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the direction of the body axis of the aircraft in a missile coordinate system, ay1Representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the longitudinal symmetry plane of the aircraft and the direction vertical to the body axis under the elastic body coordinate system, az1Representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the direction perpendicular to the two directions under the elastic body coordinate system;
the rudder deflection command signal is obtained by the following formula (six),
Figure BDA0002169949270000191
wherein, aMθRepresenting the acceleration of the highly dynamic aircraft in the direction of the declination of the line of sight, aMφThe acceleration of the high dynamic aircraft in the direction of the line of sight inclination is represented, N represents a guidance coefficient, N is generally equal to 4,
Figure BDA0002169949270000194
the first differential of the expression r is shown,
Figure BDA0002169949270000195
the first differential of the theta is represented,
Figure BDA0002169949270000196
to represent
Figure BDA0002169949270000197
First order differential of ay1Representing the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the longitudinal symmetry plane of the aircraft and the direction vertical to the body axis under the elastic body coordinate system, az1The relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in the direction perpendicular to the two directions under the elastic coordinate system is shown.
(4) The rudder deflection instruction signal is amplified by the amplifying converter and converted into a signal form required by the steering engine, the steering engine adopts a cold air type steering engine, and the high-pressure gas valve is controlled in sequence according to the rudder deflection instruction, so that the high-pressure lift pushes the actuating device, and the control surface is operated to move.
The track is shown as the high dynamic compensation proportion guidance law in fig. 3;
(II) the proportion guidance method without a compensation term refers to that in the guidance process, the traditional proportion guidance law in the prior art is adopted for guidance control, and the method comprises the following specific steps:
(1) the high dynamic aircraft transmits microwaves to the static target through the radar transmitter, then the radar receiver receives echo information and transmits the echo information to the information processor for processing, and the distance r, the sight declination angle theta and the sight inclination angle of the high dynamic aircraft relative to the static target are obtained
Figure BDA0002169949270000209
And transmitting to a processing module;
(2) the processing module obtains a rudder deflection command signal through the following formula (seven) according to the information measured in the step (1),
Figure BDA0002169949270000201
wherein, aMθRepresenting the acceleration of the highly dynamic aircraft in the direction of the declination of the line of sight, aMφThe acceleration of the high dynamic aircraft in the direction of the line of sight inclination is represented, N represents a guidance coefficient, N is generally equal to 4,
Figure BDA0002169949270000203
the first differential of the expression r is shown,
Figure BDA0002169949270000204
the first differential of the theta is represented,
Figure BDA0002169949270000205
to represent
Figure BDA0002169949270000206
First differential of (a).
(3) The rudder deflection instruction signal is amplified by the amplifying converter and converted into a signal form required by the steering engine, the steering engine adopts a cold air type steering engine, and the high-pressure gas valve is controlled in sequence according to the rudder deflection instruction, so that the high-pressure lift pushes the actuating device, and the control surface is operated to move.
The track is shown as the 'proportional guidance law' in fig. 3;
(III) the proportional guidance method for strengthening the fixed term compensation refers to that in the guidance process, the existing proportional guidance law (overweight compensation proportional guidance law) with the fixed term compensation is adopted for guidance control, and the specific steps are as follows:
(1) the high dynamic aircraft transmits microwaves to the static target through the radar transmitter, then the radar receiver receives echo information and transmits the echo information to the information processor for processing, and the distance r, the sight declination angle theta and the sight inclination angle of the high dynamic aircraft relative to the static target are obtained
Figure BDA00021699492700002010
And transmitting to a processing module;
(2) the processing module obtains a rudder deflection command signal through the following formula (eight) according to the information measured in the step (1),
Figure BDA0002169949270000202
wherein, aMθRepresenting the acceleration of the highly dynamic aircraft in the direction of the declination of the line of sight, aMφThe acceleration of the high dynamic aircraft in the direction of the line of sight inclination is represented, N represents a guidance coefficient, N is generally equal to 4,
Figure BDA0002169949270000207
the first differential of the expression r is shown,
Figure BDA0002169949270000208
the first differential of the theta is represented,
Figure BDA0002169949270000211
to represent
Figure BDA0002169949270000213
First differential of (a).
Wherein the content of the first and second substances,
Figure BDA0002169949270000212
for the over-weight compensation term of the proportional guidance law, N is generally taken1＝2。
(3) The rudder deflection instruction signal is amplified by the amplifying converter and converted into a signal form required by the steering engine, the steering engine adopts a cold air type steering engine, and the high-pressure gas valve is controlled in sequence according to the rudder deflection instruction, so that the high-pressure lift pushes the actuating device, and the control surface is operated to move.
The trajectory is shown as the "over-compensated proportional guidance law" in fig. 3.
Also shown in fig. 3 is a partial enlarged view around point a.
As can be seen from fig. 3, the curves shown in the "high dynamic compensation proportion guidance law", the "overweight compensation proportion guidance law" and the "target motion trajectory" are all overlapped, which indicates that guidance methods using different guidance laws can track the high dynamic aircraft to the position of the stationary target.
Experimental example 2
A simulation experiment of tracking a uniform motion target by a high dynamic aircraft is carried out through a computer, wherein the simulation conditions are as follows: the flying speed of the high dynamic aircraft is 800m/s, and the rotating speed is 15 r/s.
The results of the conditions of the method for simulating the high dynamic compensation guidance control instruction, the proportional guidance method without the compensation term and the proportional guidance method with the fixed compensation when tracking the uniform motion target are shown in fig. 4.
Wherein the content of the first and second substances,
(I) the high dynamic compensation guidance control instruction method is a high dynamic compensation guidance control method aiming at a high speed maneuvering target, the specific steps are similar to those in the experimental example 1, and the difference is that: a radar transmitter transmits microwaves to a constant-speed moving target;
the track is shown as the high dynamic compensation proportion guidance law in fig. 4;
(II) the proportional guidance method without a compensation term refers to the guidance control by adopting the traditional proportional guidance law in the prior art in the guidance process, and the specific steps are similar to those in the experimental example 1, and are different from the following steps: a radar transmitter transmits microwaves to a constant-speed moving target;
the track is shown as the 'proportional guidance law' in fig. 4;
(III) the fixed compensation proportion guidance method is to adopt the existing fixed term compensation proportion guidance law (overweight compensation proportion guidance law) to perform guidance control in the guidance process, the specific steps are similar to those in the experimental example 1, and the difference is that: a radar transmitter transmits microwaves to a constant-speed moving target;
the trajectory is shown as the "over-compensated proportional guidance law" in fig. 4.
Also shown in fig. 4 is a partial enlarged view in the vicinity of point B.
As can be seen from fig. 4, the three guidance methods with different guidance laws can enable the high dynamic aircraft to track the position of the target moving at a constant speed.
Experimental example 3
A simulation experiment of tracking a high-speed maneuvering target by a high-dynamic aircraft is carried out through a computer, and the simulation conditions are as follows: the flying speed of the high dynamic aircraft is 800m/s, and the rotating speed is 15 r/s.
The conditions of a high dynamic compensation guidance control instruction method, a proportional guidance method without a compensation term and a proportional guidance method with fixed compensation when tracking a high-speed maneuvering target are simulated respectively, and the result is shown in FIG. 5.
Wherein the content of the first and second substances,
(I) the high dynamic compensation guidance control instruction method is a high dynamic compensation guidance control method aiming at a high speed maneuvering target, the specific steps are similar to those in the experimental example 1, and the difference is that: the radar transmitter transmits microwaves to the high-speed maneuvering target;
the track is shown as the 'high dynamic compensation proportion guidance law' in fig. 5;
(II) the proportional guidance method without a compensation term refers to the guidance control by adopting the traditional proportional guidance law in the prior art in the guidance process, and the specific steps are similar to those in the experimental example 1, and are different from the following steps: the radar transmitter transmits microwaves to the high-speed maneuvering target;
the track is shown as the 'proportional guidance law' in fig. 5;
(III) the fixed compensation proportion guidance method is to adopt the existing fixed term compensation proportion guidance law (overweight compensation proportion guidance law) to perform guidance control in the guidance process, the specific steps are similar to those in the experimental example 1, and the difference is that: the radar transmitter transmits microwaves to the high-speed maneuvering target;
the trajectory is shown as the "over-compensated proportional guidance law" in fig. 5.
Also shown in fig. 5 is a close-up view around point C.
As can be seen from FIG. 5, in the process of tracking the movement of the high-speed maneuvering target, the high-dynamic compensation guidance control instruction method can accurately track the high-speed maneuvering target, and the proportional guidance method without the compensation term and the proportional guidance method with the reinforcement compensation cannot accurately track the high-speed maneuvering target.
In addition, the high-dynamic aircraft for the high-speed maneuvering target can compensate a guidance control instruction in real time and change the flight track continuously, so that the aim of tracking the target accurately is fulfilled.
The present invention has been described above in connection with preferred embodiments, but these embodiments are merely exemplary and merely illustrative. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.

Claims (9)

1. A high dynamic compensation guidance control system for high speed maneuvering targets, characterized in that the system comprises a measuring module (1), a processing module (2) and an execution module (3), wherein,
the measuring module (1) is used for measuring the relative position information of the high-dynamic aircraft and the high-speed maneuvering target and the attitude information of the high-dynamic aircraft in real time,
the processing module (2) is used for obtaining a rudder deflection command signal in real time according to the information measured by the measuring module,
the execution module (3) is used for receiving the rudder deflection instruction signal obtained by the processing module and converting the signal into a signal form required by the steering engine;
the processing module (2) comprises a relative position calculating submodule (21), a relative speed calculating submodule (22), a relative acceleration calculating submodule (23), a compensation guidance instruction calculating submodule (24) and a rudder deflection instruction calculating submodule (25), wherein,
the relative position resolving submodule (21) is used for obtaining the relative position of the high-dynamic aircraft and the high-speed maneuvering target under a ground coordinate system in real time according to the measured distance, the sight declination angle and the sight inclination angle of the high-dynamic aircraft relative to the high-speed maneuvering target;
the relative speed resolving submodule (22) is used for obtaining the relative speed of the high-dynamic aircraft and the high-speed maneuvering target according to the relative position of the high-dynamic aircraft and the high-speed maneuvering target under a ground coordinate system in real time;
the relative acceleration resolving submodule (23) is used for obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in real time according to the relative speed of the high-dynamic aircraft and the high-speed maneuvering target under a ground coordinate system;
the compensation guidance instruction resolving submodule (24) is used for obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system in real time according to the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system, and the relative acceleration is information of a compensation aircraft guidance control instruction;
and the rudder deflection instruction resolving submodule (25) is used for acquiring a rudder deflection instruction signal in real time according to the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the missile coordinate system.
2. The system according to claim 1, characterized in that the measurement module (1) comprises a relative position information measurement submodule (11) and an aircraft attitude information measurement submodule (12), wherein,
the relative position information measuring submodule (11) is used for measuring the position information of the high-dynamic aircraft relative high-speed maneuvering target in real time,
the aircraft attitude information measurement submodule (12) is used for obtaining attitude information of the high-dynamic aircraft in real time.
3. The system according to claim 2, wherein the relative position information measurement submodule (11) is a radar measurement element including a radar transmitter (111), a radar receiver (112) and an information processor (113);
wherein the radar transmitter (111) is used for transmitting microwave or millimeter wave to a high-speed maneuvering target,
the radar receiver (112) is used for receiving echo information of a high-speed maneuvering target,
the information processor (113) is used for processing the received echo information to obtain the position information of the high-dynamic aircraft relative high-speed maneuvering target.
4. The system according to claim 2, wherein the aircraft attitude information measurement submodule (12) includes a MEMS gyroscope (121) and a geomagnetic sensor (122), wherein,
the MEMS gyroscope (121) is used for measuring the pitch angle of the high-dynamic aircraft
Figure FDA0002479325330000021
And the yaw angle psi, and the yaw angle,
the geomagnetic sensor (122) is used for measuring a rolling angle gamma of the high-dynamic aircraft.
5. The system according to claim 1, characterized in that the relative position resolving submodule (21) obtains in real time the relative position of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system by the following equation (I),
Figure FDA0002479325330000031
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure FDA00024793253300000312
denotes the inclination of the line of sight, and θ denotes the declination of the line of sight.
6. The system according to claim 1, characterized in that the relative velocity resolving submodule (22) obtains in real time the relative velocity of the high dynamic aircraft and the high speed maneuvering target under the ground coordinate system by the following equation (two),
Figure FDA0002479325330000032
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure FDA0002479325330000033
the first differential of the expression r is shown,
phi denotes the inclination of the line of sight,
Figure FDA0002479325330000034
which represents the first differential of phi,
theta represents the declination of the line of sight,
Figure FDA0002479325330000035
the first differential of the theta is represented,
Figure FDA0002479325330000036
the first differential of x is represented by,
Figure FDA0002479325330000037
the first differential of y is represented by,
Figure FDA0002479325330000038
represents the first differential of z.
7. The system according to claim 1, wherein the relative acceleration resolving submodule (23) obtains in real time the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system by the following formula (III),
Figure FDA0002479325330000039
wherein r represents the distance between the high dynamic aircraft and the high speed maneuvering target,
Figure FDA00024793253300000310
the first differential of the expression r is shown,
Figure FDA00024793253300000311
the second derivative of r is represented as,
phi denotes the inclination of the line of sight,
Figure FDA0002479325330000041
which represents the first differential of phi,
Figure FDA0002479325330000042
which represents the second derivative of phi,
theta represents the declination of the line of sight,
Figure FDA0002479325330000043
the first differential of the theta is represented,
Figure FDA0002479325330000044
the second derivative of theta is represented by,
Figure FDA0002479325330000045
the second derivative of x is represented as,
Figure FDA0002479325330000046
to representThe second derivative of y is the sum of,
Figure FDA0002479325330000047
represents the second derivative of z.
8. A high dynamic compensation guidance control method for a high speed maneuvering target, the method being implemented by the system of one of claims 1 to 7, characterized in that the method comprises the steps of:
step 1, obtaining the relative position information of a high-dynamic aircraft and a high-speed maneuvering target and the attitude information of the high-dynamic aircraft in real time through a measurement module (1);
step 2, obtaining a rudder deflection instruction signal of the high-dynamic aircraft in real time through the processing module (2);
and 3, converting the rudder deflection command signal in real time through the execution module (3) to control the steering engine to move.
9. The method according to claim 8, characterized in that step 2 comprises the following sub-steps:
step 2-1, obtaining the relative position of the high dynamic aircraft and the high speed maneuvering target in the ground coordinate system in real time through a relative position resolving submodule (21);
2-2, obtaining the relative speed of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system in real time through a relative speed resolving submodule (22);
2-3, obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target under the ground coordinate system in real time through a relative acceleration resolving submodule (23);
2-4, obtaining the relative acceleration of the high-dynamic aircraft and the high-speed maneuvering target in a missile coordinate system in real time through a compensation guidance instruction resolving submodule (24);
and 2-5, obtaining a rudder deflection instruction signal of the high-dynamic aircraft in real time through a rudder deflection instruction resolving submodule (25).
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