CN114489101A - Terminal guidance control method and system for unmanned aerial vehicle - Google Patents
Terminal guidance control method and system for unmanned aerial vehicle Download PDFInfo
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Abstract
The invention provides a terminal guidance control method and a terminal guidance control system of an unmanned aerial vehicle, which construct a terminal guidance control system based on a homing guidance mode according to the guidance characteristics of an electronic warfare unmanned aerial vehicle, provide a corresponding terminal guidance control method, realize the requirements of diagonal constraint and guidance precision through a passive radar guidance method, simultaneously realize terminal maneuvering and simulated attack control, accurately intercept a target, defend missile attack and improve the penetration resistance.
Description
Technical Field
The invention belongs to the technical field of aerospace unmanned aerial vehicles, and particularly relates to a terminal guidance control method and a terminal guidance control system for an unmanned aerial vehicle.
Background
Electronic warfare drones are aircraft used to attack a given target, requiring that the aircraft have the ability to accurately fly toward the target and meet or collide with it within the killing radius. The electronic warfare unmanned aerial vehicle guidance control system is used for guiding the aircraft to overcome various interference factors and autonomously and accurately fly to a target according to a determined rule and requirements. The guidance control system is composed of a guidance system and a control system, wherein the guidance system measures the deviation between the actual motion condition of the unmanned aerial vehicle and the required motion condition, or measures the relative position and the deviation between the unmanned aerial vehicle and a target to form a guidance instruction part for controlling the unmanned aerial vehicle to fly, and the guidance control system comprises a guidance head, a sensor subsystem and a guidance instruction forming module. The seeker intercepts a signal of an enemy target radar, measures angle information of the unmanned aerial vehicle and the target radar in real time, completes tasks of finding and tracking a target and measuring the position of the target, the sensor subsystem completes measurement of motion parameters of the unmanned aerial vehicle, the guiding instruction forming module carries out transformation and operation on various parameters according to a guiding rule and forms a guiding instruction according to relative motion states and relative position deviation, and the guiding instruction is sent to the control system to control the unmanned aerial vehicle to track in real time so that the unmanned aerial vehicle can hit the target finally. The electronic warfare unmanned aerial vehicle guidance control technology is different from the guidance control technology of a conventional unmanned aerial vehicle, the searching, capturing, tracking and attacking of a target are mainly completed through the seeking guidance, the seeking guidance rule is reasonably designed so as to meet the requirements of the falling angle constraint and the guidance precision, the control of terminal maneuvering and simulated attack is realized, the target is accurately intercepted, the attack of a missile is resisted, and the penetration resistance is improved.
Conventional self-seeking guidance includes tracking, parallel approach, and proportional guidance. In order to continuously improve the hit precision, new guidance technologies such as a generalized proportional guidance method, a modified proportional guidance method, optimal guidance, predicted trajectory guidance and the like are continuously derived on the basis of the traditional target guidance technology.
The target-seeking guidance is a guidance technology for automatically guiding the unmanned aerial vehicle to a target by utilizing a seeker arranged on the unmanned aerial vehicle to receive certain characteristic energy of target radiation or anti-radiation, determining the relative position of the target and the unmanned aerial vehicle, and further forming a control instruction according to a preset guidance instruction, and the target-seeking guidance is an important technical basis for realizing accurate automatic tracking and accurate striking of the unmanned aerial vehicle to a moving target in real time. According to different modes of acquiring target characteristic energy, the target guidance is divided into an active mode, a semi-active mode and a passive mode, and according to different physical characteristics of energy, the target guidance can be divided into radar guidance, infrared guidance, television guidance, laser guidance and the like.
The proportional guidance law used by Songlong of China air-to-air missile research institute in a diving attack section requires that the missile keeps a given proportional relation between the rotation angular velocity of a velocity vector and the rotation angular velocity of a target sight line in the flight process, and the proportional guidance method is a widely applied guidance method and has obvious advantages in terms of guidance precision and feasibility of engineering application.
The optimal terminal guidance law with the constraint of the drop point and the drop angle is deduced by utilizing the Lagrange method, which is often superordinate to Beijing university of Physician, the guidance law is complex in form and low in realizability.
A thesis of designing a guidance law of a banked turning aircraft with angle constraint is published by chaga of Harbin industrial university and the like, an optimal guidance law based on a sliding mode surface is designed by considering the dynamic characteristic of a control loop, buffeting is eliminated by adjusting parameters of a sliding mode approach law, the guidance law can prove convergence, but the form of the guidance law is still complex, and 5 parameters in total need to be determined.
Most of the prior art at present is not easy to be applied in engineering, and most of the prior proportion-guided reference objects are guided missiles, so that the unmanned aerial vehicle and the missiles have great difference in aerodynamic appearance and control mode. Nowadays, for an electronic war unmanned aerial vehicle guidance control system with a terminal attack function, a great deal of problems and requirements still exist for the problems of the electronic war unmanned aerial vehicle guidance control system architecture and terminal guidance.
Disclosure of Invention
The invention provides a terminal guidance control method and a terminal guidance control system of an Unmanned Aerial Vehicle (UAV) aiming at the defects and requirements of the prior art, the invention constructs a terminal guidance control system based on a homing guidance mode according to the guidance characteristics of the UAV for electronic warfare, provides a corresponding terminal guidance control method, realizes the requirements of diagonal constraint and guidance precision by a passive radar guidance method, simultaneously realizes terminal maneuvering and simulated attack control, accurately intercepts a target, defends missile attack and improves the penetration resistance.
The specific implementation content of the invention is as follows:
the invention provides a terminal guidance control method of an unmanned aerial vehicle, which is based on a terminal guidance control system of the unmanned aerial vehicle and adopts a passive radar guidance mode to carry out terminal guidance control; the terminal guidance control system comprises a guidance subsystem, a control subsystem and a position sensor;
the guidance subsystem comprises a passive radar seeker unit and a guidance instruction generating device; the control subsystem comprises a flight control unit, an actuating mechanism and an attitude sensor; the passive radar seeker unit, the guidance instruction generating device, the flight control unit and the executing mechanism are sequentially connected and are in control connection with the unmanned aerial vehicle body through the executing mechanism; the attitude sensor and the position sensor are respectively arranged on the machine body, the attitude sensor is connected with the flight control unit, and the position sensor is connected with the guide instruction generating device;
the control method specifically comprises the following steps:
step 1: after the unmanned aerial vehicle transmits, continuously measuring target motion parameters through the passive radar seeker to obtain deviation relative to a required track, and sending the deviation relative to the required track to the guidance instruction generating device;
step 2: converting and calculating the received deviation signal through a guidance instruction generating device to form a guidance instruction; the guidance instruction requires the unmanned aerial vehicle to change course or speed;
and step 3: and the guidance instruction is sent to a control subsystem, and after the guidance instruction is converted and amplified to obtain instruction content, the control plane is driven to deflect through an executing mechanism to change the flight direction of the unmanned aerial vehicle or/and change the flight speed of the unmanned aerial vehicle, so that the unmanned aerial vehicle returns to the required track.
In order to better implement the present invention, further, in step 2, the specific operations of forming the guidance instruction are:
step 2.1: performing information modeling and filtering processing on the passive radar seeker;
step 2.2: resolving a guidance instruction;
step 2.3: and carrying out attack target position estimation.
In order to better implement the present invention, further, the specific operations of step 2.1 include:
step 2.1.1: acquiring aircraft state quantity information through a passive radar seeker and a position sensor, calculating to obtain a relative motion relation between the unmanned aerial vehicle and a target, and constructing a relative motion model;
step 2.1.2: calculating by using a relative motion model to obtain line of sight inclination angle data without noise, line of sight declination angle data without noise, line of sight inclination angle speed data without noise, line of sight declination angle speed data without noise, line of sight distance and line of sight speed; the calculated noise-free sight line inclination angle data, noise-free sight line deflection angle data, noise-free sight line inclination angle rate data and noise-free sight line deflection angle rate data are used as ideal real value data of the passive radar seeker;
step 2.1.3: constructing a seeker characteristic model according to the output data characteristic of the passive radar seeker; inputting the calculated noise-free line of sight inclination angle data and the noise-free line of sight declination angle data into a seeker characteristic model, and adding a measurement error and a noise characteristic to the noise-free line of sight inclination angle data and the noise-free line of sight declination angle data in the seeker characteristic model to obtain a noise-containing line of sight declination angle and a noise-containing line of sight inclination angle; using the obtained noise-containing line-of-sight declination angle and the obtained noise-containing line-of-sight inclination angle as line-of-sight angle characteristic data for simulating the output of a real passive radar seeker;
step 2.1.4: constructing a filtering module, and sending the obtained noise-containing sight line deflection angle and the noise-containing sight line inclination angle into the filtering module for processing to obtain a sight line deflection angle and a sight line inclination angle; and estimating to obtain the line of sight declination angle rate and the line of sight inclination angle rate.
In order to better implement the present invention, further, in step 2.1.4, the filtering module processes the input noisy view line declination angle, noisy view line inclination angle, and estimated view line declination angle rate and view line inclination angle rate by using a first-order inertia element low-pass filtering method; and simultaneously, a Kalman filter is arranged to carry out joint estimation on the view declination angle rate and the view dip angle rate.
In order to better implement the present invention, further, the operation flow of the kalman filter is as follows:
scheme 1: initializing a parameter matrix and an intermediate variable matrix;
and (2) a flow scheme: inputting a set parameter matrix;
and (3) a flow path: and circularly updating the estimated value data.
In order to better implement the present invention, further, the specific operations of the process 3 are: the following processes are circularly carried out:
scheme 3.1: updating the process estimation value;
scheme 3.2: updating the covariance of the process estimate;
scheme 3.3: calculating a Kalman gain;
scheme 3.4: updating the optimal estimation value;
scheme 3.5: and updating the covariance of the optimal estimation value.
In order to better implement the present invention, further, in step 2.1.2, a ground reference inertial coordinate system is established, and the calculation of the non-noise line of sight inclination angle data, the non-noise line of sight declination angle data, the non-noise line of sight inclination angle rate data, the non-noise line of sight declination angle rate data, the line of sight distance and the line of sight speed is performed according to the positions of the target and the unmanned aerial vehicle in the ground reference inertial coordinate system;
in the step 2.1.3, a machine body coordinate system and a sight line coordinate system are established, and sight line inclination angle data without noise and sight line deflection angle data without noise are converted into a ground reference inertial coordinate system from an implementation coordinate system according to the output data characteristics of the passive radar seeker; then, converting the ground inertial coordinate system into a body coordinate system; and finally, adding a measurement error and noise characteristics to obtain a visual line deflection angle containing noise and a visual line inclination angle containing noise.
In order to better implement the present invention, in step 2.1.4, in the filtering module, the obtained noise-containing view declination angle and the obtained noise-containing view inclination angle are converted from the body coordinate system to the ground reference inertial coordinate system, and then processed.
In order to better implement the present invention, further, the specific operations of step 2.2 are:
converging the information acquired by the passive radar seeker in the step 2.1.1 to obtain guidance parameters, attitude angles and execution terminal guidance sign information; and integrating the guidance parameters, the attitude angle and the information of the execution terminal guidance sign, the line of sight inclination angle rate data, the line of sight deflection angle rate data, the line of sight distance and the line of sight speed which are obtained in the step 2.1.2, and the line of sight deflection angle, the line of sight inclination angle, the line of sight deflection angle rate and the line of sight inclination angle rate which are obtained in the step 2.2.4, resolving the guidance parameters, generating a roll angle instruction, an overload overrun sign and an overload instruction, and sending the roll angle instruction, the overload overrun sign and the overload instruction to the flight control unit.
The invention also provides a terminal guidance control system of the unmanned aerial vehicle, which is used for carrying out the terminal guidance control method of the unmanned aerial vehicle; the terminal guidance control system comprises a guidance subsystem, a control subsystem and a position sensor;
the guidance subsystem comprises a passive radar seeker unit and a guidance instruction generating device; the control subsystem comprises a flight control unit, an actuating mechanism and an attitude sensor; the passive radar seeker unit, the guidance instruction generating device, the flight control unit and the executing mechanism are sequentially connected and are in control connection with the unmanned aerial vehicle body through the executing mechanism; the attitude sensor and the position sensor are respectively installed on the machine body, the attitude sensor is connected with the flight control unit, and the position sensor is connected with the guide instruction generating device.
Compared with the prior art, the invention has the following advantages and beneficial effects:
the invention provides a terminal guidance control method and a terminal guidance control system of an unmanned aerial vehicle, which construct a terminal guidance control system based on a homing guidance mode according to the guidance characteristics of an electronic warfare unmanned aerial vehicle, provide a corresponding terminal guidance control method, realize the requirements of diagonal constraint and guidance precision through a passive radar guidance method, simultaneously realize terminal maneuvering and simulated attack control, accurately intercept a target, defend missile attack and improve the penetration resistance.
Drawings
FIG. 1 is a schematic view of a guidance control system;
FIG. 2 is a schematic diagram of the estimation of the angle of view in the ground reference inertial frame;
FIG. 3 is a schematic view of a line-of-sight coordinate system versus a body coordinate system;
FIG. 4 is a schematic diagram of a Kalman filtering algorithm;
FIG. 5 is a schematic diagram of the relative motion of the drone and the target;
FIG. 6 is a schematic view of the configuration of the end guidance control system;
fig. 7 is a schematic view of an attack simulation.
Detailed Description
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it should be understood that the described embodiments are only a part of the embodiments of the present invention, and not all embodiments, and therefore should not be considered as a limitation to the scope of protection. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
In the description of the present invention, it is to be noted that, unless otherwise explicitly specified or limited, the terms "disposed," "connected," and "connected" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
Example 1:
the embodiment provides a terminal guidance control method of an unmanned aerial vehicle, which is based on a terminal guidance control system of the unmanned aerial vehicle, and adopts a passive radar guidance mode to perform terminal guidance control as shown in fig. 1; the terminal guidance control system comprises a guidance subsystem, a control subsystem and a position sensor;
the guidance subsystem comprises a passive radar seeker unit and a guidance instruction generating device; the control subsystem comprises a flight control unit, an actuating mechanism and an attitude sensor; the passive radar seeker unit, the guidance instruction generating device, the flight control unit and the executing mechanism are sequentially connected and are in control connection with the unmanned aerial vehicle body through the executing mechanism; the attitude sensor and the position sensor are respectively arranged on the machine body, the attitude sensor is connected with the flight control unit, and the position sensor is connected with the guide instruction generating device;
the control method specifically comprises the following steps:
step 1: after the unmanned aerial vehicle transmits, continuously measuring target motion parameters through the passive radar seeker to obtain deviation relative to a required track, and sending the deviation relative to the required track to the guidance instruction generating device;
step 2: converting and calculating the received deviation signal through a guidance instruction generating device to form a guidance instruction; the guidance instruction requires the unmanned aerial vehicle to change course or speed;
and step 3: and the guidance instruction is sent to a control subsystem, and after the guidance instruction is converted and amplified to obtain instruction content, the control plane is driven to deflect through an executing mechanism to change the flight direction of the unmanned aerial vehicle or/and change the flight speed of the unmanned aerial vehicle, so that the unmanned aerial vehicle returns to the required track.
The working principle is as follows: FIG. 1 is a schematic view of a guidance control principle, according to which an electronic warfare unmanned aerial vehicle workflow is decomposed as follows:
1) after the unmanned aerial vehicle launches, the seeker continuously measures the deviation relative to the required track, and sends the deviation to the guidance instruction forming module;
2) the guidance instruction forming module transforms and calculates the deviation signal to form a guidance instruction, and the instruction requires the unmanned aerial vehicle to change the course or the speed;
3) the guidance instruction is sent to a control system, and after transformation and amplification, a servo actuation system drives a control plane to deflect, so that the flight direction of the unmanned aerial vehicle is changed, and the unmanned aerial vehicle returns to the required track;
4) if so, the interference. Attitude angle changes, and attitude sensor detects out the attitude deviation, sends into the computer with the signal of telecommunication form, and control unmanned aerial vehicle resumes original state, guarantees that unmanned aerial vehicle is stable flies according to original orbit.
Example 2:
in this embodiment, on the basis of the above embodiment 1, to better implement the present invention, further, as shown in fig. 6, fig. 6 is a schematic diagram of a terminal guidance control system, which combines guidance head input and guidance instruction output, and implements full-digital simulation closed-loop control through a simulink building model to perform guidance precision influence analysis. In the step 2, the specific operation of forming the guidance instruction is as follows:
step 2.1: performing information modeling and filtering processing on the passive radar seeker;
step 2.2: resolving a guidance instruction;
step 2.3: and carrying out attack target position estimation.
Further, the specific operation of step 2.1 includes:
step 2.1.1: acquiring aircraft state quantity information through a passive radar seeker and a position sensor, calculating to obtain a relative motion relation between the unmanned aerial vehicle and a target, and constructing a relative motion model;
step 2.1.2: calculating by using a relative motion model to obtain line of sight inclination angle data without noise, line of sight declination angle data without noise, line of sight inclination angle speed data without noise, line of sight declination angle speed data without noise, line of sight distance and line of sight speed; the calculated noise-free sight line inclination angle data, noise-free sight line deflection angle data, noise-free sight line inclination angle rate data and noise-free sight line deflection angle rate data are used as ideal real value data of the passive radar seeker;
step 2.1.3: constructing a seeker characteristic model according to the output data characteristic of the passive radar seeker; inputting the calculated noise-free line of sight inclination angle data and the noise-free line of sight declination angle data into a seeker characteristic model, and adding a measurement error and a noise characteristic to the noise-free line of sight inclination angle data and the noise-free line of sight declination angle data in the seeker characteristic model to obtain a noise-containing line of sight declination angle and a noise-containing line of sight inclination angle; using the obtained noise-containing line-of-sight declination angle and the obtained noise-containing line-of-sight inclination angle as line-of-sight angle characteristic data for simulating the output of a real passive radar seeker;
step 2.1.4: constructing a filtering module, and sending the obtained noise-containing sight line deflection angle and the noise-containing sight line inclination angle into the filtering module for processing to obtain a sight line deflection angle and a sight line inclination angle; and estimating to obtain the line of sight declination angle rate and the line of sight inclination angle rate.
The working principle is as follows: according to the terminal guidance control principle, the design of the terminal guidance control system mainly comprises the following steps:
a) seeker information modeling and filtering
b) Guidance instruction solution
c) Attack target location estimation
1. Seeker information modeling and filtering
The part mainly aims at the radar seeker to carry out line-of-sight angle modeling and angular rate estimation filtering.
Firstly, the target line-of-sight angle and the angular rate are calculated according to the positions of the target and the aircraft in an inertial system, and the calculated value is used as an ideal real value of the seeker to be processed.
And secondly, converting the ideal line angle into a machine body coordinate system according to the output data characteristic of the seeker, namely the real line angle information output by the seeker is relative to the machine body coordinate system.
From the ground reference inertial system C0To the organism system C1Is C0→1From the ground reference inertial system C0To the line of sight coordinate system C4Is C0→4Setting the visual angle of the seeker output target-aircraft relative to the coordinate system of the aircraft body as delta qεAnd Δ qβEyes of peopleThe unit vector of the target-aircraft sight line has a projection of [ 100 ] in the sight line coordinate system]TThe projection of the sight line vector in the inertial coordinate system is
According to the transformation matrix from the inertia system to the body system, the projection of the unit sight line vector in the body coordinate system is obtained as C1xyzFrom the body system C1To the line of sight coordinate system C4Is C1→4The projection of the unit sight line vector in the body coordinate system is
The angle of the line of sight relative to the airframe can thus be expressed as
Δqε＝-arcsin c1z (1)
And thirdly, the measurement error and the noise characteristic of the seeker are added into the line-of-sight angle calculation value, so that the line-of-sight angle characteristic output by a real seeker can be simulated.
Under the condition of considering the slope error of the radome of the seeker, the sight line measured by the seeker relative to the body
The angle at which the system is formed can be described as
Δqε＝-(1+bε)arcsinc1Z (3)
Wherein, bεAnd bβRepresenting the radome aiming error slopes of the pitch channel and the yaw channel, respectively.
In actual flight, due to the influence of environmental factors, the slope of the aiming error of the antenna housing is randomly changed
So that its variation process is regarded as random white gaussian noise, i.e.
Wherein wεAnd wβRepresenting seeker measurement mechanism pitch channel and yaw channel gaussian white noise, respectively.
And fourthly, information output by the seeker generally needs to be filtered to be used for calculating a guidance command and providing a filter for jointly estimating the line-of-sight angular rate.
Other parts of this embodiment are the same as those of embodiment 1, and thus are not described again.
Example 3:
in this embodiment, on the basis of any one of the foregoing embodiments 1-2, in order to better implement the present invention, further, in step 2.1.4, the filtering module uses a first-order inertia element low-pass filtering method to process the input noisy view declination, noisy view dip, and the estimated view declination rate and view dip rate; and simultaneously, a Kalman filter is arranged to carry out joint estimation on the view declination angle rate and the view dip angle rate.
The working principle is as follows: firstly, the line-of-sight angle output by the seeker relative to a machine body coordinate system is converted into an inertial coordinate system. The line-of-sight angular rate is then calculated in such a way that the time intervals are taken according to the operating cycle of the control system.
And (3) performing low-pass filtering on the line-of-sight angle output by the seeker and the estimated line-of-sight angular rate by using a first-order inertia element:
wherein T is2Is a time constant. In order to enhance timeliness, a set of Kalman filter is designed for the line-of-sight angular rate joint estimation.
Other parts of this embodiment are the same as any of embodiments 1-2 described above, and thus are not described again.
Example 4:
in this embodiment, on the basis of any one of the above embodiments 1 to 3, in order to better implement the present invention, as shown in fig. 4, further, the operation flow of the kalman filter is as follows:
scheme 1: initializing a parameter matrix and an intermediate variable matrix;
and (2) a flow scheme: inputting a set parameter matrix;
and (3) a flow scheme: and circularly updating the estimated value data.
Further, the specific operations in the process 3 are as follows: the following processes are circularly carried out:
scheme 3.1: updating the process estimation value;
scheme 3.2: updating the covariance of the process estimate;
scheme 3.3: calculating a Kalman gain;
scheme 3.4: updating the optimal estimation value;
scheme 3.5: and updating the covariance of the optimal estimation value.
Other parts of this embodiment are the same as any of embodiments 1 to 3, and thus are not described again.
Example 5:
in this embodiment, on the basis of any one of the foregoing embodiments 1 to 4, in order to better implement the present invention, further, in step 2.1.2, a ground reference inertial coordinate system is established, and the calculation of the line-of-sight inclination angle data without noise, the line-of-sight declination angle data without noise, the line-of-sight inclination angle rate data without noise, the line-of-sight declination angle rate data without noise, the line-of-sight distance, and the line-of-sight speed is performed according to the position of the target and the unmanned aerial vehicle in the ground reference inertial coordinate system;
in the step 2.1.3, a machine body coordinate system and a sight line coordinate system are established, and sight line inclination angle data without noise and sight line deflection angle data without noise are converted into a ground reference inertial coordinate system from an implementation coordinate system according to the output data characteristics of the passive radar seeker; then, converting the ground inertial coordinate system into a body coordinate system; and finally, adding a measurement error and noise characteristics to obtain a visual line deflection angle containing noise and a visual line inclination angle containing noise.
Further, in the step 2.1.4, in the filtering module, the obtained noise-containing line-of-sight declination angle and the obtained noise-containing line-of-sight inclination angle are converted from the body coordinate system to the ground reference inertial coordinate system, and then are processed.
Other parts of this embodiment are the same as any of embodiments 1 to 4, and thus are not described again.
Example 6:
in this embodiment, on the basis of any one of the above embodiments 1 to 5, in order to better implement the present invention, further, the specific operation of step 2.2 is:
converging the information acquired by the passive radar seeker in the step 2.1.1 to obtain guidance parameters, attitude angles and execution terminal guidance sign information; and integrating the guidance parameters, the attitude angle and the information of the execution terminal guidance sign, the line of sight inclination angle rate data, the line of sight deflection angle rate data, the line of sight distance and the line of sight speed which are obtained in the step 2.1.2, and the line of sight deflection angle, the line of sight inclination angle, the line of sight deflection angle rate and the line of sight inclination angle rate which are obtained in the step 2.2.4, resolving the guidance parameters, generating a roll angle instruction, an overload overrun sign and an overload instruction, and sending the roll angle instruction, the overload overrun sign and the overload instruction to the flight control unit.
The working principle is as follows: the resolving specifically comprises the following steps:
firstly, establishing a relative motion process of the unmanned aerial vehicle and the target
The mathematical model of the equation of relative motion is
Wherein R represents the miss distance, and epsilon-0 represents the guidance relation established by different guidance methods
Second, select proper guidance law for searching according to task requirement
When in use
That is, the rate of change in the velocity direction is proportional to the rate of change in the target line-of-sight angle during flight to the target.
The proportional guidance method is a guidance method between tracking method and parallel approach method, and satisfies the requirements
Under the condition of convergence, the front section of the trajectory is bent, the maneuvering capability can be fully utilized, and the rear section of the trajectory is straight, so that the device has rich maneuvering capability. The guiding relation is rewritten into the overload form, which is the generalized proportional guidance method.
And (3) respectively calculating overload instructions in the pitching direction and the azimuth direction in the sight system, giving the overload instructions in the longitudinal direction and the lateral direction by a guidance law according to the angular velocity of the sight, and rotating the main lifting surface of the aircraft to the direction of the overload instruction vector sum to obtain normal overload perpendicular to the symmetrical surface in the sight system.
And projecting the overload instruction in the sight line coordinate system into the quasi-machine system, adopting the tilt turning control, and outputting the normal overload and roll angle instruction which is vertical to the symmetrical plane in the normal overload computer system in the quasi-machine system as a guidance instruction.
Other parts of this embodiment are the same as any of embodiments 1 to 5, and thus are not described again.
Example 7:
in this embodiment, on the basis of any one of embodiments 1 to 6, for the attack target position estimation:
the position information of the target should be obtained as soon as possible after entering the terminal guidance segment, which helps to achieve the expected hit effect on the one hand, and on the other hand, if the target is shut down, the aircraft continues to fly to the predetermined area according to this information, waits for the target to be searched and identified in the predetermined area, and implements the attack.
Determining the specific direction and position of the target
Before the unmanned plane for electronic warfare takes off, the attack targets are initially bound by parameters, and the initial parameters comprise target positions (Xt, Yt and Zt) and three movement speeds (Xt _ dot, Yt _ dot and Zt _ dot) of the targets.
Determining the approximate target area range if the specific direction and position of the target are not determined
Before or during take-off, the region range of the target is bound or the flight line is searched, and the region range is not limited to the shapes of irregular polygons, circles, ellipses and the like (binding of an attack region can be realized through flight line binding). If a search route is set, searching along the search route; if the search route is not set, the center of the target area range is used as the circle center, the detection distance of the seeker is used as the radius to circle and search for the target, the current height of the airplane is obtained during the circle, and the optimal cruise speed is obtained at the speed. When receiving an 'attack mode' instruction of the ground station, guiding the unmanned aerial vehicle to an attack target airspace, and accessing terminal guidance to execute an attack task.
If the specific direction and position of the target are not determined, and the approximate target area range is not determined
The method comprises the steps of conducting guidance by using guidance information of a seeker before the radar is shut down, conducting positioning calculation by using angle measurement information of the seeker and position information of a navigation system, estimating the relative position of a target-electronic warfare unmanned aerial vehicle, and conducting guidance information calculation by using the positioning information and the position information of the navigation system after the radar is shut down.
The attack target position estimation function is as follows: the position information of the target should be obtained as soon as possible after entering the terminal guidance segment, which on the one hand helps to achieve the desired hit effect, and on the other hand if the target is shut down, the aircraft continues to fly to the predetermined area according to this information, waiting for the target to be identified and performing the attack in the predetermined area. And designing estimation methods for determining the specific position and position of the target, not determining the specific position and position of the target, determining the range of the approximate target area, not determining the specific position and position of the target and not determining the target position in three scenes of the range of the approximate target area. FIG. 7 is a model diagram of the present invention for simulating attack control. The three-dimensional coordinates in the figure are all in meters.
Other parts of this embodiment are the same as any of embodiments 1 to 6, and thus are not described again.
Example 8:
in this embodiment, on the basis of any one of the above embodiments 1 to 7, in order to better implement the present invention, further:
with respect to step 2.1.2, as shown in fig. 2: FIG. 2 is a schematic view of an estimate of the angle of view in the inertial system from which a target angle of view in the inertial system can be calculated as
Deriving line-of-sight angular rate information from line-of-sight angle
Wherein:
Δx＝xt-x，Δy＝yt-y，ΔZ＝zt-Z；
xt，yt，Zt-coordinates of the target in the inertial frame of reference;
x, y, z-the coordinates of the aircraft in the inertial frame of reference;
xr，yr，Zr——C[Δx，Δy，ΔZ]wherein C is derived from rotation of the line of sight declination about the z-axis by the reference inertial frame;
The actual line-of-sight angle information output by the seeker is relative to the body coordinate system, and the ideal line-of-sight angle is converted into the body coordinate system.
From the ground reference inertial system C0To the organism system C1The transformation matrix of (2):
from the ground reference inertial system C0To the line of sight coordinate system C4The transformation matrix of (a) is:
setting the visual angle of the seeker output target-aircraft relative to the coordinate system of the aircraft body as delta qεAnd Δ qβThe unit vector of the target-aircraft line of sight has a projection in the line of sight coordinate system of [ 100 ]]TProjection of a line-of-sight vector in an inertial coordinate system
Is composed of
According to the transformation matrix from the inertia system to the body system, the projection of the unit sight line vector in the body coordinate system can be obtained as
From the body system C1To the line of sight coordinate system C4Is converted into
Projection of unit sight line vector in body coordinate system
Is composed of
As can be seen from the above equation, the angle of the line of sight relative to the machine system can be expressed as
Δqε＝-arc Sinclz (11)
Wherein c is1x，c1y，c1zSee in C1xyz。
Other parts of this embodiment are the same as any of embodiments 1 to 7, and thus are not described again.
Example 9:
this embodiment is based on any of the above embodiments 1 to 8, and further, as shown in fig. 3, in order to better implement the present invention, with respect to step 2.1.3:
the estimation of the visual angle rate, the filtering and the design of a guidance law are carried out in an inertial system, the information output by the seeker is relative to a body coordinate system, and the visual angle output by the seeker relative to the body coordinate system is converted into the inertial coordinate system as follows.
The projection of the unit sight line vector in the body coordinate system is
From the ground reference inertial system C0To the organism system C1Is C0→1The projection of the unit line-of-sight vector in the inertial coordinate system is
The line-of-sight angle of the line-of-sight vector with respect to the inertial system is calculated as follows:
other parts of this embodiment are the same as any of embodiments 1 to 8, and thus are not described again.
Example 10:
in this embodiment, based on any one of the above embodiments 1 to 9, in order to better implement the present invention, as shown in fig. 4, a specific example of the operation of the kalman filter is as follows:
taking a longitudinal channel as an example, a Kalman filter for jointly estimating the line-of-sight angular rate is given.
Firstly, a system state equation and a measurement state equation are established.
Assuming the target is fixed, the longitudinal channel state variable is defined as x1＝qε，
A controlled variable isεThe longitudinal channel joint estimation line-of-sight angular rate state equation is
As a result of the fixation of the target,
can be calculated by using speed information, aεThe acceleration information of the machine system is converted into a visual system to obtain the acceleration information, and the calculation formula is as follows:
note the book
Formula (13) can be rewritten as
Discretizing the linear time-varying system to obtain the system state equation
Xε(k+1)＝Φε(k)Xε(k)+Bε(k)uε(k)
Wherein Xε(k) Is a state quantity, uε(k) For controlling the quantity, there are
uε(k)＝-aε(k)
Adding a zero-mean Gaussian random process noise vector W due to model errors and the likeε(k) Obtaining XεForecast estimation of
The measurement equation takes the angle of the sight line obtained by the seeker through measurement relative to the aircraft system.
Zε(k)＝HεXε(k)+Vε(k) (16)
Wherein Vε(k) Is a zero-mean Gaussian random process noise vector, observation matrix
Wε(k)、Vε(k) The two process noises are in accordance with Gaussian distribution and are independent of each other, and the mathematical expressions of the process noise covariance matrix and the observation noise covariance matrix are
In the formula, muwIs WεMean matrix of (u)vIs VεThe mean matrix of (a); qεIs the system process noise WεThe covariance matrix of (a); rεFor measuring process noise VεIs the covariance matrix of (S ε is W)εAnd Vεδ is a kronecker symbol.
The first step is as follows: the empirical estimation value at this moment is updated based on the state variable and the controlled variable at the previous moment.
The second step is that: according to the system process parameter matrix phiε(k) Sum system process noise covariance matrix Qε(k) Estimating a state variable
Corresponding covariance matrix Pε(k+1/k)。
Wherein P isε(k) Representing the covariance matrix of the state variables, characterizing the confidence in the initial state, Qε(k) And obtaining a noise covariance matrix for the system process according to the variance of the noise.
The third step: updating Kalman gain matrix K of systemk
Wherein HεTo observe the matrix, Rε(k +1) is the noise V of the measurement processkThe covariance matrix of (2).
The fourth step: and updating the process estimator to be a comprehensive optimal estimator according to the measured value and the measured parameter matrix.
Wherein
yε(k +1) is an actual measurement value.
The fifth step: updating the covariance matrix P of the optimal estimatork。
Pε(k+1)＝(I-KkHε)Pε(k+1/k) (22)
Therefore, the design of the pitch channel promotes the Kalman filter as
Wherein
And
each represents XεFiltered and forecast estimates of, state quantity XεIs initially provided with
State variable corresponding covariance matrix PεAt an initial state of
Observation matrix HεIs composed of
System process noise WεOf the covariance matrix QεIs composed of
Measurement Process noise VεOf the covariance matrix RεIs composed of
Other parts of this embodiment are the same as any of embodiments 1 to 9, and thus are not described again.
Example 11:
in this embodiment, on the basis of any one of the above embodiments 1 to 10, in order to better implement the present invention, further, as shown in fig. 5, fig. 5 is a schematic diagram of a relative motion relationship between the unmanned aerial vehicle and the target, and a generalized guidance law is established according to the relative motion relationship. The guiding rule is designed mainly by the generalized proportional guiding rule in the terminal guiding section, and the guiding relation equation is
Wherein
For line-of-sight angular rate, K is the steering coefficient and n is the overload command (in the line-of-sight coordinate system).
Respectively calculating overload instructions in the pitching and azimuth directions in a sight system, wherein the expression is
Wherein, KyAnd KzIs a coefficient of proportionality that is,
for the angular rate of the pitch line of sight,
is the azimuthal line-of-sight angular rate.
The guidance law usually gives overload instructions in the longitudinal direction and the lateral direction according to the line-of-sight angular rate, and the normal overload perpendicular to the symmetrical plane in the aircraft system is obtained by rotating the main lifting surface of the aircraft to the direction of the overload instruction vector.
Firstly, projecting an overload instruction in a sight line coordinate system into a quasi-machine system
By means of banked turn control, according to normal overload n in quasi-aircraft systemy5And nz5Normal overload and roll angle commands perpendicular to the plane of symmetry in a computer system, namely
γc＝atan2(nzc，nyc)
Wherein:
nyc、nzc-a lateral, longitudinal overload command;
nc-overload instruction and vector size;
γc-a roll angle command.
If the falling angle constraint is required to be ensured, a falling angle constraint item can be introduced
Instead of the former
K1As a proportional guidance coefficient, K2For the coefficients of the corner constraint term, in general K2Larger effects are better, but at the expense of an increased amount of off-target, a compromise design is required.
Other parts of this embodiment are the same as any of embodiments 1 to 10 described above, and thus are not described again.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention in any way, and all simple modifications and equivalent variations of the above embodiments according to the technical spirit of the present invention are included in the scope of the present invention.
Claims (10)
1. A terminal guidance control method of an unmanned aerial vehicle is based on a terminal guidance control system of the unmanned aerial vehicle, and terminal guidance control is carried out in a passive radar guidance mode; the terminal guidance control system is characterized by comprising a guidance subsystem, a control subsystem and a position sensor;
the guidance subsystem comprises a passive radar seeker unit and a guidance instruction generating device; the control subsystem comprises a flight control unit, an actuating mechanism and an attitude sensor; the passive radar seeker unit, the guidance instruction generating device, the flight control unit and the executing mechanism are sequentially connected and are in control connection with the unmanned aerial vehicle body through the executing mechanism; the attitude sensor and the position sensor are respectively arranged on the machine body, the attitude sensor is connected with the flight control unit, and the position sensor is connected with the guide instruction generating device;
the control method specifically comprises the following steps:
step 1: after the unmanned aerial vehicle transmits, continuously measuring target motion parameters through the passive radar seeker to obtain deviation relative to a required track, and sending the deviation relative to the required track to the guidance instruction generating device;
step 2: converting and calculating the received deviation signal through a guidance instruction generating device to form a guidance instruction; the guidance instruction requires the unmanned aerial vehicle to change course or speed;
and step 3: and the guidance instruction is sent to a control subsystem, and after the guidance instruction is converted through conversion and amplification processing to obtain instruction content, the control surface is driven to deflect through an execution mechanism to change the flight direction of the unmanned aerial vehicle or/and change the flight speed of the unmanned aerial vehicle, so that the unmanned aerial vehicle returns to the required track.
2. The terminal guidance control method of the unmanned aerial vehicle as claimed in claim 1, wherein in the step 2, the specific operation of forming the guidance instruction is:
step 2.1: performing information modeling and filtering processing on the passive radar seeker;
step 2.2: resolving a guidance instruction;
step 2.3: and carrying out attack target position estimation.
3. The terminal guidance control method of the unmanned aerial vehicle as claimed in claim 2, wherein the specific operations of step 2.1 include:
step 2.1.1: acquiring aircraft state quantity information through a passive radar seeker and a position sensor, calculating to obtain a relative motion relation between the unmanned aerial vehicle and a target, and constructing a relative motion model;
step 2.1.2: calculating by using a relative motion model to obtain line of sight inclination angle data without noise, line of sight declination angle data without noise, line of sight inclination angle speed data without noise, line of sight declination angle speed data without noise, line of sight distance and line of sight speed; the calculated noise-free sight line inclination angle data, noise-free sight line deflection angle data, noise-free sight line inclination angle rate data and noise-free sight line deflection angle rate data are used as ideal real value data of the passive radar seeker;
step 2.1.3: constructing a seeker characteristic model according to the output data characteristic of the passive radar seeker; inputting the calculated noise-free line of sight inclination angle data and the noise-free line of sight declination angle data into a seeker characteristic model, and adding a measurement error and a noise characteristic to the noise-free line of sight inclination angle data and the noise-free line of sight declination angle data in the seeker characteristic model to obtain a noise-containing line of sight declination angle and a noise-containing line of sight inclination angle; using the obtained noise-containing line-of-sight declination angle and the obtained noise-containing line-of-sight inclination angle as line-of-sight angle characteristic data for simulating the output of a real passive radar seeker;
step 2.1.4: constructing a filtering module, and sending the obtained noise-containing sight line deflection angle and the noise-containing sight line inclination angle into the filtering module for processing to obtain a sight line deflection angle and a sight line inclination angle; and estimating to obtain the line of sight declination angle rate and the line of sight inclination angle rate.
4. The terminal guidance control method of the unmanned aerial vehicle as claimed in claim 3, wherein in the step 2.1.4, the filtering module processes the input noisy view declination, noisy view dip and the estimated view declination rate and view dip rate by using a first-order inertia element low-pass filtering method; and simultaneously, a Kalman filter is arranged to carry out joint estimation on the view declination angle rate and the view dip angle rate.
5. The terminal guidance control method of the unmanned aerial vehicle as claimed in claim 4, wherein the operation flow of the kalman filter is as follows:
scheme 1: initializing a parameter matrix and an intermediate variable matrix;
and (2) a flow scheme: inputting a set parameter matrix;
and (3) a flow path: and circularly updating the estimated value data.
6. The terminal guidance control method of the unmanned aerial vehicle as claimed in claim 5, wherein the specific operations of the process 3 are as follows: the following processes are circularly carried out:
scheme 3.1: updating the process estimation value;
scheme 3.2: updating the covariance of the process estimate;
scheme 3.3: calculating a Kalman gain;
scheme 3.4: updating the optimal estimation value;
scheme 3.5: and updating the covariance of the optimal estimation value.
7. The terminal guidance control method for the unmanned aerial vehicle as claimed in claim 3, 4, 5 or 6, wherein in step 2.1.2, a ground reference inertial coordinate system is established, and the calculation of the noise-free line of sight inclination angle data, the noise-free line of sight declination angle data, the noise-free line of sight inclination angle rate data, the noise-free line of sight declination angle rate data, the line of sight distance and the line of sight speed is performed according to the position of the target and the unmanned aerial vehicle in the ground reference inertial coordinate system;
in the step 2.1.3, a machine body coordinate system and a sight line coordinate system are established, and sight line inclination angle data without noise and sight line deflection angle data without noise are converted into a ground reference inertial coordinate system from an implementation coordinate system according to the output data characteristics of the passive radar seeker; then, converting the ground inertial coordinate system into a body coordinate system; and finally, adding a measurement error and noise characteristics to obtain a visual line deflection angle containing noise and a visual line inclination angle containing noise.
8. The terminal guidance control method for the unmanned aerial vehicle as claimed in claim 7, wherein in step 2.1.4, the obtained noise-containing view declination angle and the noise-containing view inclination angle are converted from the body coordinate system to the ground reference inertial coordinate system in the filtering module and then processed.
9. An end guidance control method for a drone according to claim 3 or 4 or 5 or 6, characterized in that the specific operations of step 2.2 are:
converging the information acquired by the passive radar seeker in the step 2.1.1 to obtain guidance parameters, attitude angles and execution terminal guidance sign information; and integrating the guidance parameters, the attitude angle and the information of the execution terminal guidance sign, the line of sight inclination angle rate data, the line of sight deflection angle rate data, the line of sight distance and the line of sight speed which are obtained in the step 2.1.2, and the line of sight deflection angle, the line of sight inclination angle, the line of sight deflection angle rate and the line of sight inclination angle rate which are obtained in the step 2.2.4, resolving the guidance parameters, generating a roll angle instruction, an overload overrun sign and an overload instruction, and sending the roll angle instruction, the overload overrun sign and the overload instruction to the flight control unit.
10. An end guidance control system of an unmanned aerial vehicle for carrying out an end guidance control method of an unmanned aerial vehicle of the above claim 1 or 2 or 3 or 4 or 5 or 6 or 7 or 8 or 9; the terminal guidance control system is characterized by comprising a guidance subsystem, a control subsystem and a position sensor;
the guidance subsystem comprises a passive radar seeker unit and a guidance instruction generating device; the control subsystem comprises a flight control unit, an actuating mechanism and an attitude sensor; the passive radar seeker unit, the guidance instruction generating device, the flight control unit and the executing mechanism are sequentially connected and are in control connection with the unmanned aerial vehicle body through the executing mechanism; the attitude sensor and the position sensor are respectively installed on the machine body, the attitude sensor is connected with the flight control unit, and the position sensor is connected with the guide instruction generating device.
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