CN112445230B - Multi-mode guidance system and guidance method for high dynamic aircraft in large cross-domain complex environment - Google Patents

Abstract

The invention discloses a multi-mode guidance system and a guidance method for a high-dynamic aircraft in a large-span complex environment. The guidance system is applied to a high-dynamic aircraft in a large-span complex environment. The aircraft is provided with a satellite guidance module and a laser. Seeker, in order to ensure that the laser seeker can capture the laser signal of the target when entering the final guidance section, the aircraft's traveling direction is controlled in the middle guidance section by setting the virtual target position, so that the aircraft can pass the virtual target position, thereby Ensure that the target can enter the field of view of the laser seeker. In addition, due to the installation of various guidance modules, more accurate parameter information can be screened out by setting specific screening and judgment conditions, thereby providing more accurate information for subsequent solutions. based on the data, and then obtain the guidance command through the solution process that can eliminate the interference factors.

Description

Multi-mode guidance system and guidance method for high dynamic aircraft in large cross-domain complex environment

technical field

The invention relates to a guidance method, in particular to a multi-mode guidance method for a high-dynamic aircraft in a large cross-domain complex environment.

Background technique

High dynamic guidance aircraft can complete the guidance task stably and efficiently under conventional environmental conditions. However, the complex environment puts forward stricter requirements for guided aircraft. In the face of harsh environments, aircraft with a single guidance mode may even be unable to obtain guidance information due to large errors. In the end, the flight mission cannot be accurately completed due to the guidance information, and in the aircraft with the fusion of multiple guidance modes, there are deficiencies in the cooperative operation and weight distribution among the multiple guidance modes; especially the guidance of the composite laser seeker The aircraft, under some special conditions, cannot perform laser guidance when entering the terminal guidance segment because the target does not appear in the field of view of the seeker; some aircraft have a combination of satellite guidance and laser guidance. In the process of guidance processing, the anti-interference ability is weak. When one guidance method is limited, it cannot be switched to another guidance method in time, resulting in a decrease in the final guidance control accuracy; in order to improve its ability to deal with complex environments, the laser can be The seeker is set as a strapdown seeker. For a guided aircraft equipped with a strapdown laser seeker, it is difficult to directly detect the line-of-sight angular velocity between the aircraft and the target when calculating the overload. The calculation accuracy of the line-of-sight angular velocity will also affect the final guidance accuracy.

Due to the above reasons, the inventors have conducted in-depth research on the existing guidance methods for high dynamic aircraft, and designed a multi-mode guidance system and guidance method for high dynamic aircraft in large-span complex environments that can solve the above problems.

SUMMARY OF THE INVENTION

In order to overcome the above-mentioned problems, the inventors of the present invention have carried out keen research and designed a multi-mode guidance system and a guidance method for a high-dynamic aircraft in a large-scale complex environment. The guidance system is applied in a high-dynamic aircraft in a large-scale complex environment , The aircraft is equipped with a satellite guidance module and a laser seeker. In order to ensure that the laser seeker can capture the laser signal of the target when entering the terminal guidance section, the travel direction of the aircraft is controlled in the middle guidance section by setting the virtual target position. , so that the aircraft can pass through the virtual target position, so as to ensure that the target can enter the field of view of the laser seeker. In addition, due to the installation of a variety of guidance modules, more accurate parameters can be screened by setting specific screening and judgment conditions. information, so as to provide a more accurate data basis for the subsequent calculation, and then obtain the line-of-sight angular velocity of the aircraft and the target through the calculation process that can eliminate the interference factors, and further obtain the guidance command, thereby completing the present invention.

Specifically, an object of the present invention is to provide a multi-mode guidance system for a highly dynamic aircraft in a large cross-domain complex environment. The guidance system includes a satellite module 1, a laser seeker 2, an attitude sensitive module 3, and a virtual target module 4. , a correction module (5), a line-of-sight angular velocity calculation module 6 and an overload calculation module 7;

The satellite module 1 is used for receiving satellite signals, and according to the satellite signals, the position information, speed information and the line-of-sight angle information between the aircraft and the target are calculated in real time;

The laser seeker 2 is used to detect the line-of-sight information between the aircraft and the target in the terminal guidance segment;

The attitude sensitive module 3 is used to obtain the flight parameter information of the aircraft sensitively in real time,

The virtual target module 4 is used to provide a virtual target position for the aircraft, so that the aircraft flies to the virtual target position in the middle guidance section,

The correction module 5 is respectively connected with the satellite module 1 and the laser seeker 2, and outputs the line-of-sight angle information of the aircraft and the target that is close to the true value;

The line-of-sight angular velocity calculation module 6 is used to obtain the line-of-sight angular velocity of the aircraft and the target in real time according to the line-of-sight angle information of the aircraft and the target that is close to the true value output by the correction module 5;

The overload calculation module 7 is used for obtaining the required overload according to the line-of-sight angular velocity of the aircraft and the target obtained in real time by the line-of-sight angular velocity calculation module 6 .

Another object of the present invention is to provide a multi-mode guidance method for a high-dynamic aircraft in a large cross-domain complex environment, the method comprising the following steps:

Step 1, receive satellite signals through satellite module 1, and calculate the position information, speed information and line-of-sight information of the aircraft and the target in real time according to the satellite signals;

The flight parameter information of the aircraft is obtained sensitively in real time through the attitude sensitive module 3;

Step 2, in the middle guidance section, the virtual target position is provided for the aircraft by the virtual target module 4, so that the aircraft flies to the virtual target position in the middle guidance section,

Step 3, in the terminal guidance section, the laser seeker 2 is used to detect the line-of-sight information between the aircraft and the target in the terminal guidance section;

Step 4, in the terminal guidance section, output the line-of-sight angle information of the aircraft and the target that is close to the true value through the correction module 5;

Step 5, in the terminal guidance section, obtain the line-of-sight angular velocity of the aircraft and the target in real time through the line-of-sight angular velocity calculation module 6 according to the line-of-sight angle information of the aircraft and the target that is close to the true value output by the correction module 5;

Step 6, in the final guidance section, the overload calculation module 7 obtains the required overload according to the line-of-sight angular velocity of the aircraft and the target obtained by the line-of-sight angular velocity calculation module 6 in real time.

The beneficial effects of the present invention include:

(1) The multi-mode guidance system and guidance method for a high-dynamic aircraft in a large-scale and complex environment provided by the present invention can adapt to the complex flight environment. Capable of capturing laser information;

(2) The multi-mode guidance system and guidance method for a high-dynamic aircraft in a large cross-domain complex environment provided according to the present invention can combine two guidance means of satellite guidance and laser guidance, and can automatically operate according to the advantages and disadvantages of the guidance means in each flight stage Switch to a more suitable guidance method for guidance;

(3) The multi-mode guidance system and guidance method for a high-dynamic aircraft in a large-span complex environment provided according to the present invention can be applied to a strapdown guided aircraft, thereby further improving its anti-interference ability to deal with complex environments, and can timely and accurately The line-of-sight angular velocity between the aircraft and the target is estimated, and the overload is required to ensure its guidance accuracy.

Description of drawings

FIG. 1 shows a schematic diagram of the overall structure of a multi-mode guidance system for a high-dynamic aircraft in a large-span complex environment according to a preferred embodiment of the present invention;

FIG. 2 shows a schematic structural diagram of a four-piece composite antenna of a satellite module in a multi-mode guidance system for a high-dynamic aircraft in a large-span complex environment according to a preferred embodiment of the present invention;

Figure 3 shows the change trajectory of the real value of the line-of-sight angular velocity between the aircraft and the target and the calculated value of the line-of-sight angular velocity between the aircraft and the target without considering the influence of internal and external disturbances on the system in the experimental example;

Fig. 4 shows the change trajectory of the real value of the line-of-sight angular velocity of the aircraft and the target and the calculated value of the line-of-sight angular velocity of the aircraft and the target under the situation of considering the influence of internal and external disturbances on the system in the experimental example;

Description of reference numbers:

1- Satellite module

11- Four composite antennas

12-Anti-interference module

13-Satellite solver sub-module

14-Accommodating slot

15-Protective baffle

2-Laser seeker

3- Attitude Sensitive Module

4- Virtual Object Module

5- Correction module

6-Line of sight angular velocity calculation module

7-Overload solving module

Detailed ways

The present invention will be further described in detail below through the accompanying drawings and embodiments. The features and advantages of the present invention will become more apparent from these descriptions.

The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. While various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless otherwise indicated.

On the one hand, according to the multi-mode guidance system for a high-dynamic aircraft provided in the present invention, the system includes a satellite module 1, a laser seeker 2, an attitude sensitive module 3, a virtual target module 4, a correction module 5, a line of sight Angular velocity calculation module 6 and overload calculation module 7;

Among them, the virtual target module 4 , the correction module 5 , the line-of-sight angular velocity calculation module 6 and the overload calculation module 7 may be integrated in one circuit board or installed in different circuit boards respectively.

In a preferred embodiment, the satellite module 1 is used to receive satellite signals, and calculate the position information and speed information of the aircraft in real time according to the satellite signals.

In order to make the position and velocity information obtained by the satellite module 1 more accurate and the source of the information more reliable, and to reduce the possibility of losing the position and velocity information periodically due to lost satellites, in this application, preferably, the satellite module includes four pieces of information. The synthetic antenna 11, the anti-jamming module 12 and the satellite solving sub-module 13, wherein:

Four synthetic antennas are used to receive satellite signals,

The anti-jamming module 12 is connected with the four composite antennas 11, and is used for filtering the satellite signal to eliminate noise interference in the satellite signal;

The satellite solving sub-module 13 receives the filtered satellite signal, converts the satellite signal into a navigation message, and then calculates the position and speed information of the aircraft at the current moment according to the navigation message; the navigation message described in the present invention is composed of The message that the navigation satellite broadcasts to the user describing the operating state parameters of the navigation satellite, including the system time, ephemeris, almanac, correction parameters of the satellite clock, the health status of the navigation satellite and the parameters of the ionospheric delay model; the parameters of the navigation message are given to the user. The time information is provided, and the user's position coordinates and speed can be calculated by using the parameters of the navigation text; and then the line-of-sight angle information between the aircraft and the target can be obtained in real time according to the position coordinates of the aircraft and the position coordinates of the target point.

As shown in FIG. 2 , the four composite antennas 11 include four sheet-like antenna plates for receiving satellite signals under high overload conditions. The antennas can be rectangular flat plates or arcs with radians. The shape of a plate can be set according to the outline of the aircraft. In this application, it is preferably an arc-shaped plate with an arc, which is matched with the outline of the aircraft, and in the process of rolling the aircraft, the arc-shaped plate with an arc is Antennas receive satellite signals longer and with better signal strength.

Preferably, the four composite antennas 11 are evenly distributed around the aircraft, and are arranged along the circumferential direction of the aircraft's rolling, so as to ensure that the satellite signal reception capability will not be weakened when the aircraft is rolling at high speed.

Compared with the traditional cone antenna or loop antenna, the four-piece composite antenna 11 in the present application is less susceptible to external noise or interference due to the small space occupied by the chip antenna, and the chip antenna has a higher degree of integration, and its satellite The signal reception ability is stronger.

Preferably, the four-piece composite antenna 11 can be made of the same material as the traditional loop antenna or cone antenna, and the thickness of the four-piece composite antenna can be reduced as much as possible on the basis of ensuring stability and physical strength to reduce costs.

Preferably, four composite antennas 11 are arranged on the outer wall of the aircraft,

More preferably, a concave accommodating groove 14 is provided on the outer wall of the aircraft, the four composite antennas 11 are installed in the accommodating groove 14, and the depth dimension of the accommodating groove 14 is larger than that of the four composite antennas. The thickness of the antenna 11 is determined, and a protective baffle 15 is provided outside the four composite antennas 11 .

The outer shape of the protective baffle is adapted to the outline of the aircraft, which can be arc-shaped or flat. The four composite antennas will not move and break during acceleration.

The protective baffle 15 is used to protect the four composite antennas on the inside of the aircraft during the acceleration phase, and prevent the four composite antennas from being damaged during the acceleration process. When the aircraft enters the guidance stage, the protective baffle 15 is removed from the aircraft. The upper part is disengaged, so that the four composite antennas are exposed, so that it is convenient to receive satellite signals with the four composite antennas, so as to avoid the shielding baffle 15 from shielding/interfering with satellite signals. Preferably, the four composite antennas are similar to the steering gear on the aircraft, and need to be started only in the guidance stage, so the protective baffle 15 and the baffle outside the steering gear of the aircraft can be controlled and disengaged synchronously.

The laser seeker 2 is a strapdown laser seeker, which can be used for real-time detection to obtain the line-of-sight angle between the aircraft and the target. The strapdown seeker is fixed on the missile body to save the mechanical frame structure, so the attitude of the missile body is reduced. The motion is all coupled into the information measured by the seeker, and it is difficult to directly obtain the line-of-sight angular velocity information of the aircraft and the target through the laser seeker.

In a preferred embodiment, as shown in FIG. 1 , the attitude sensing module 3 is used to obtain the flight parameter information of the aircraft sensitively in real time. Specifically, the attitude sensitive module 3 includes an inertial gyroscope and a geomagnetic sensor, the pitch angle information and yaw angle information of the aircraft are obtained in real time through the inertial gyro, and the roll angle information of the aircraft is obtained in real time through the geomagnetic sensor. The flight parameter information includes pitch angle information, yaw angle information and roll angle information of the aircraft. After the pitch angle information and yaw angle information are known, angle values in other coordinate systems can be obtained through coordinate transformation.

The aircraft described in this application is suitable for a high-dynamic aircraft with a large span and a complex environment, wherein the large span refers to that the aircraft has strong adaptability to changes in altitude. Flying at an altitude of 3000m and above; the high dynamic refers to the high speed of the aircraft. Generally speaking, it is called high dynamic when the speed of the aircraft is above 10 rpm; the aircraft is flying towards the target. During the process, the sensitive components and steering gear on the aircraft are called start-up control when they start to work, and the flight stage after start-up control is called the guidance section. The guidance section includes the middle guidance section and the terminal guidance section. , Generally speaking, when a guided aircraft with a laser seeker is 3km away from the target, and the laser seeker starts to capture the target, the aircraft enters the terminal guidance section. After taking control, the flight stage before entering the terminal guidance section is Middle guidance section.

In the process of the aircraft flying to the target direction, in the middle guidance process, the aircraft in the prior art controls the flight with the target position as the desired end point. Due to the interference of factors, the flight trajectory of the aircraft will change, and it will have different flight trajectories due to different external interference conditions, so that when entering the terminal guidance segment, that is, when the aircraft is 3km away from the target and the laser seeker starts to capture the target, the aircraft may It has deviated from the optimal flight trajectory, and the position of the target is not necessarily within the field of view of the laser seeker. The laser seeker cannot capture the laser signal reflected by the target, so laser guidance cannot be performed, resulting in a decrease in hit accuracy.

In view of these problems, in a preferred embodiment of the present invention, the flight trajectory of the aircraft is adjusted in advance by setting the virtual target position, so that the aircraft passes near the area where the virtual target position is located, and then the specific position of the virtual target point is set by setting the specific position of the virtual target point. To ensure that the aircraft can capture the laser signal when it is near the virtual target position, so that the aircraft can technically capture the laser signal reflected by the target when entering the terminal guidance segment.

Further preferably, the multi-mode guidance system is provided with a virtual target module 4, wherein a virtual target position is stored, and in the middle guidance section, the virtual target module 4 transfers the virtual target position to the overload calculation module 7 in real time, and the overload solution The arithmetic module 7 takes the virtual target position as the target position, and controls the aircraft to fly to the virtual target position.

The target position is a set of three-dimensional space coordinate data information, and specifically includes longitude coordinates, latitude coordinates and altitude values where a target point is located.

In particular, the virtual target module 4 obtains the virtual target position by the following formula:

Among them, x _F represents the x-axis coordinate in the plumb plane where the virtual target position is located, y _F represents the y-axis coordinate in the plumb plane where the virtual target position is located,

为弹道偏角，α为飞行器攻角，β为飞行器侧滑角，OT表示发射原点和目标间的横向距离，L_c表示飞行器滑翔的距离，T_s为预设的姿态调整时间。c _t and c _g are design gains, c _t =3, c _g =4, v _g is the speed of the aircraft, θ is the ballistic inclination,

is the ballistic declination angle, α is the attack angle of the aircraft, β is the sideslip angle of the aircraft, OT is the lateral distance between the launch origin and the target, L _c is the gliding distance of the aircraft, and T _s is the preset attitude adjustment time.

为预先设计好的期望值，α和β是根据期望的θ和

值通过坐标转换得到的，其坐标转换过程参见《导弹力学》钱杏芳等编著，北京理工大学出版社；where θ,

are pre-designed expected values, α and β are based on the expected θ and

The value is obtained by coordinate transformation. For the coordinate transformation process, please refer to "Missile Mechanics" edited by Qian Xingfang et al. Beijing Institute of Technology Press;

Before launching, the aircraft will simulate the flight trajectory according to its own known parameters and the information of the launch point and the target point. During the flight trajectory simulation process, the speed value of the aircraft when entering the final guidance section is the above v _g , so The gliding distance L _c of the aircraft is the gliding distance from the start control point to the end guidance section during flight trajectory simulation.

For the simulation process of the flight trajectory simulation, please refer to Han Zipeng et al., edited by Han Zipeng et al., Beijing Institute of Technology Press.

Setting the virtual target position can ensure that the target can be in the field of view of the laser seeker when the aircraft enters the terminal guidance segment, preventing the problem that the laser seeker cannot capture the laser signal.

Controlling the aircraft to fly towards the virtual target position in the middle guidance segment can ensure that the aircraft finally reaches the virtual target position or near the virtual target position, and the virtual target position is obtained through reasonable calculation. When the aircraft reaches the virtual target position, that is, It can enter the terminal guidance section, and can ensure that the laser seeker on the aircraft can capture the laser signal at this position, that is, it can ensure that the target is within the field of view of the laser seeker.

The multi-mode guidance system refers to a composite aircraft with multiple guidance modes, such as an aircraft with satellite guidance mode and laser guidance mode. In the multi-mode guidance system, multiple modes can be repeatedly detected to obtain some common parameters, such as The line-of-sight angle information between the aircraft and the target. However, whether the information is accurate or not can only be known through judgment and comparison. However, how to timely and accurately select parameters that are closer to the true values from the same parameters provided by different modes can directly affect the subsequent guidance control accuracy.

In view of the above problems, in a preferred embodiment, by setting a judgment module and setting corresponding reasonable judgment parameters to quickly and efficiently solve the selection problem of the same kind of parameters, specifically, to quickly and efficiently select a parameter that is closer to the true value. Line-of-sight information between the aircraft and the target.

Specifically, a correction module 5 is set in the multi-mode guidance system. The correction module 5 is connected to the satellite module 1 and the laser seeker 2 respectively, and outputs the line-of-sight angle information of the aircraft and the target that is close to the real value.

The correction module 5 receives the line-of-sight information of the aircraft and the target transmitted by the satellite module 1 and the laser seeker 2 in real time; the line-of-sight information of the aircraft and the target obtained according to the satellite module 1 is called the line-of-sight of the satellite aircraft and the target. The angle q _G , according to the line-of-sight angle information of the aircraft and the target obtained by the laser seeker 2 is called the line-of-sight angle q _laser of the laser aircraft and the target.

In the terminal guidance section, the correction module 5 receives the sight angle q _G between the satellite vehicle and the target and the sight angle q _laser between the laser vehicle and the target in real time, and compares the difference between the two. When the difference is greater than or equal to a predetermined value, It is considered that the line-of-sight angle q between the laser aircraft and the target _laser has not yet entered the working state, and the line-of-sight angle q _G between the satellite aircraft and the target is considered to be closer to the true value, and the line-of-sight angle q _G between the satellite aircraft and the target is passed to the line-of-sight angular velocity calculation module 6 Carry out the calculation of the line-of-sight angular velocity between the aircraft and the target; when the difference is less than the predetermined value, it is considered that the line-of-sight angle q _laser between the laser aircraft and the target has entered the working state, and accurate line-of-sight angle information can be obtained. The line-of-sight angle q between the _laser aircraft and the target is transmitted to the line-of-sight angular velocity calculation module 6 to calculate the line-of-sight angular velocity between the aircraft and the target.

Preferably, the predetermined value is 0.2° to 1°, more preferably 0.5°. The inventors have found through research that when the difference between the sight angle between the satellite vehicle and the target and the sight angle between the laser vehicle and the target is 0.5 When the angle is below °, the line-of-sight angular velocity between the aircraft and the target detected by the laser seeker is closer to the true value, so the line-of-sight angular velocity between the aircraft and the target obtained by calculating the line-of-sight angle between the laser aircraft and the target will also be closer to the true value. Thereby improving the guidance accuracy.

其中导航比N取值为2～4，优选为4，a_M为飞行器的需用过载，V为飞行器的速度，θ为滚转角，由惯性陀螺直接测量得到；

为飞行器与目标的视线角速度，此时的飞行器与目标的视线角速度由下式得到：Preferably, in the middle guidance process, in order to ensure that the aircraft can still pass through the virtual target position when it is flying towards the virtual target position under the condition of being interfered by the external large span and complex environment, the guidance rate used in the middle guidance section is Compensation Guidance Law, i.e.

The navigation ratio N is 2 to 4, preferably 4, a _M is the required overload of the aircraft, V is the speed of the aircraft, and θ is the roll angle, which is directly measured by the inertial gyroscope;

is the line-of-sight angular velocity between the aircraft and the target. The line-of-sight angular velocity between the aircraft and the target is obtained by the following formula:

为飞行器与目标的视线角速度；x_F表示虚拟目标位置所在的铅锤平面内x轴坐标，y_F表示虚拟目标位置所在的铅锤平面内y轴坐标，x_t表示卫星模块实时探测得到的飞行器自身当前位置的铅锤平面内x轴坐标，y_t表示卫星模块实时探测得到的飞行器自身当前位置的铅锤平面内y坐标。γ is the ballistic inclination angle, and its value is equivalent to the pitch angle of the aircraft, which can be directly measured by the attitude gyro; q is the line-of-sight angle between the aircraft and the target,

is the line-of-sight angular velocity between the aircraft and the target; x _F represents the x-axis coordinate in the plumb plane where the virtual target position is located, y _F represents the y-axis coordinate in the plumb plane where the virtual target position is located, and x _t represents the real-time detection of the aircraft by the satellite module The x-axis coordinate in the plumb plane of its current position, and y _t represents the y-coordinate in the plumb plane of the aircraft's current position detected in real time by the satellite module.

In order to improve the adaptability of the aircraft to large-scale cross-domain flight, the laser seeker is set as a strapdown seeker, but at the same time, it is difficult for the laser seeker to directly detect the line-of-sight angular velocity information of the aircraft and the target, and it is guided at the end. In the process, the accuracy of satellite guidance is not high enough, and it is easy to fail due to problems such as signal shielding. Therefore, it is necessary to analyze and give the line-of-sight angular velocity information of the aircraft and the target in real time according to the parameter information that can be obtained;

In view of the above-mentioned problems, in a preferred embodiment, as shown in FIG. 1 , the line-of-sight angular velocity calculation module 6 is used in the terminal guidance segment to calculate the aircraft and the target in real time according to the received line-of-sight angle information of the aircraft and the target. The line-of-sight angular velocity of the aircraft and the target is transmitted to the calculation module for the guidance and control of the final guidance section, that is, the calculation needs to use overload.

Among them, the line-of-sight angular velocity calculation module 6 is connected with the correction module 5. During the final guidance segment, the line-of-sight angular velocity calculation module 6 only receives the line-of-sight angle information of the aircraft and the target that is closer to the true value and transmitted by the correction module 5,

The line-of-sight angular velocity calculation module 6 calculates the line-of-sight angular velocity information of the aircraft and the target in real time through the following equations (1), (2) and (3),

表示x₂的导数，

表示x₁的导数，

表示x₀的导数，上一时刻得到的数值作为下一时刻迭代的初始值。Among them, q _g represents the line-of-sight angle between the aircraft and the target transmitted by the correction module 5, and q ₀ represents the estimated value of the line-of-sight angle between the aircraft and the target, that is, the above formulas (1), (2), ( 3) The estimated value of the line-of-sight angle between the aircraft and the target, q ₁ represents the estimated value of the line-of-sight angular velocity between the aircraft and the target, that is, in the calculation process, it is estimated by the above equations (1), (2), (3) The estimated value of the line-of-sight angular velocity between the aircraft and the target;

represents the derivative of _x2 ,

represents the derivative of x ₁ ,

Represents the derivative of x ₀ , and the value obtained at the previous moment is used as the initial value of the iteration at the next moment.

The initial moment x ₀ =0, x ₁ =0, x ₂ =0, every 0.001s is used as the integration step, iterates, and obtains the values of x ₀ , x ₁ and x ₂ at the next moment.

进而得到下一时刻的初始值x₀、x₁和x₂；再将得到的x₀、x₁、x₂和接收到的q_g值代入到式(一)、式(二)和式(三)中，从而得到再下一时刻对应的

如此持续循环迭代即可持续得到每次积分得到的对应x₀、x₁、x₂。Specifically, in the first iteration moment, the initial moment x ₀ =0, x ₁ =0, x ₂ =0 and the received q _g value are substituted into equations (1), (2) and (3) ), so that it can be solved

Then obtain the initial values x ₀ , x ₁ and x ₂ at the next moment; then substitute the obtained x ₀ , x ₁ , x ₂ and the received q _g values into formula (1), formula (2) and formula ( 3), so as to obtain the corresponding

In this way, the corresponding x ₀ , x ₁ , and x ₂ obtained by each integration can be continuously obtained by continuing the loop iteration.

表示飞行器与目标的视线角速度，实时输出该飞行器与目标的视线角速度给过载解算模块7即可用于解算需用过载。in,

Indicates the line-of-sight angular velocity between the aircraft and the target, and outputting the line-of-sight angular velocity between the aircraft and the target in real time to the overload calculation module 7 can be used to calculate the required overload.

Wherein, the a ₀ , a ₁ , a ₂ , δ, k ₁ and k ₂ are all design parameters. In this application, preferably, a ₀ =1-1.5, a ₁ =7-10, a ₂ =10- 15. δ=1～2, k ₁ =0.1～0.4 and k ₂ =0.2～0.4;

More preferably, a ₀ =1.1, a ₁ =8.5, a ₂ =11.5, δ=1.5, k ₁ =0.3, k ₂ =0.3.

导航比N取值为4，a_M为飞行器的需用过载，V为飞行器的速度，

为飞行器与目标的视线角速度，即为视线角速度解算模块6实时给出的飞行器与目标的视线角速度。In a preferred embodiment, as shown in FIG. 1 , the overload resolution module 7 uses a proportional guidance guidance law to perform guidance control, that is,

The value of the navigation ratio N is 4, a _M is the required overload of the aircraft, V is the speed of the aircraft,

is the line-of-sight angular velocity of the aircraft and the target, that is, the line-of-sight angular velocity of the aircraft and the target given by the line-of-sight angular velocity calculation module 6 in real time.

On the other hand, the present invention also provides a multi-mode guidance method for a high-dynamic aircraft in a large-span complex environment. The method is implemented based on the multi-mode guidance system for a high-dynamic aircraft in a large-span complex environment. It includes the following steps:

Step 1, receive satellite signals through satellite module 1, and calculate the position information, speed information and line-of-sight information of the aircraft and the target in real time according to the satellite signals;

The flight parameter information of the aircraft is obtained sensitively in real time through the attitude sensitive module 3;

Step 2, in the middle guidance section, the virtual target position is provided for the aircraft by the virtual target module 4, so that the aircraft flies to the virtual target position in the middle guidance section,

Step 3, in the terminal guidance section, the laser seeker 2 is used to detect the line-of-sight information between the aircraft and the target in the terminal guidance section;

Step 4, in the terminal guidance section, output the line-of-sight angle information of the aircraft and the target that is close to the true value through the correction module 5;

Step 5, in the terminal guidance section, obtain the line-of-sight angular velocity of the aircraft and the target in real time through the line-of-sight angular velocity calculation module 6 according to the line-of-sight angle information of the aircraft and the target that is close to the true value output by the correction module 5;

Step 6, in the final guidance section, the overload calculation module 7 obtains the required overload according to the line-of-sight angular velocity of the aircraft and the target obtained by the line-of-sight angular velocity calculation module 6 in real time.

In a preferred embodiment, the satellite module described in step 1 includes four composite antennas 11, an anti-jamming module 12 and a satellite solving sub-module 13;

As shown in FIG. 2 , four composite antennas 11 are used to receive satellite signals, and the four composite antennas 11 include four sheet-like antenna plates for receiving satellite signals under high overload conditions. The antenna may be rectangular. A flat plate shape, or an arc-shaped plate shape with a radian, can be set according to the outline of the aircraft. In this application, it is preferably an arc-shaped plate shape with a radian, which matches the outline of the aircraft and can be used when the aircraft rolls. During the process, the curved plate antenna with radian can receive satellite signals for a longer time, and the signal strength is better.

Preferably, the four composite antennas 11 are evenly distributed around the aircraft, and are arranged along the circumferential direction of the aircraft's rolling, so as to ensure that the satellite signal reception capability will not be weakened when the aircraft is rolling at high speed.

The anti-jamming module 12 is connected with the four composite antennas 11, and is used for filtering the satellite signal to eliminate noise interference in the satellite signal;

The satellite solving sub-module 13 receives the filtered satellite signal, converts the satellite signal into a navigation message, and then calculates the position and speed information of the aircraft at the current moment according to the navigation message; the parameters of the navigation message provide the user with time information , the user's position coordinates and speed can be calculated by using the parameters of the navigation text; and then the line-of-sight angle information between the aircraft and the target can be obtained in real time according to the position coordinates of the aircraft and the position coordinates of the target point.

Preferably, the flight parameter information in step 1 includes pitch angle information, yaw angle information and roll angle information of the aircraft.

In a preferred embodiment, in step 2, in the middle guidance section, the virtual target position information is calculated by the following formula in the virtual target module 4;

Among them, x _F represents the x-axis coordinate in the plumb plane where the virtual target position is located, y _F represents the y-axis coordinate in the plumb plane where the virtual target position is located,

为弹道偏角，α为飞行器攻角，β为飞行器侧滑角，OT表示发射原点和目标间的横向距离，L_c表示飞行器滑翔的距离，T_s为预设的姿态调整时间。c _t and c _g are design gains, c _t =3, c _g =4, v _g is the speed of the aircraft, θ is the ballistic inclination,

is the ballistic declination angle, α is the attack angle of the aircraft, β is the sideslip angle of the aircraft, OT is the lateral distance between the launch origin and the target, L _c is the gliding distance of the aircraft, and T _s is the preset attitude adjustment time.

为预先设计好的期望值，α和β是根据期望的θ和

值通过坐标转换得到，其坐标转换过程参见《导弹力学》钱杏芳等编著，北京理工大学出版社；where θ,

are pre-designed expected values, α and β are based on the expected θ and

The value is obtained by coordinate transformation. For the coordinate transformation process, please refer to "Missile Mechanics" edited by Qian Xingfang et al. Beijing Institute of Technology Press;

In the process of flight trajectory simulation, the speed value of the aircraft when entering the terminal guidance section is the v _g , and the gliding distance L _c of the aircraft is the gliding distance from the start control point to the terminal guidance section during the flight trajectory simulation .

The virtual target module 4 transmits the virtual target position to the overload calculation module 7 in real time, and the overload calculation module 7 takes the virtual target position as the target position, and controls the aircraft to fly to the virtual target position;

Existing guided aircraft all use the target location as the desired end point to control the flight. Before entering the final guidance section, due to the interference of external factors such as disturbance and wind resistance, the flight trajectory of the aircraft will change. It has different flight trajectories, so that when entering the terminal guidance stage, that is, when the aircraft is 3km away from the target and the laser seeker starts to capture the target, the aircraft may have deviate from the optimal flight trajectory, and the target position may not be in the laser guidance. In the field of view of the seeker, the laser seeker cannot capture the laser signal reflected at the target, so it cannot carry out laser guidance, resulting in a decrease in hit accuracy.

The virtual target position can ensure that the target can be in the field of view of the laser seeker when the aircraft enters the terminal guidance section, preventing the problem that the laser seeker cannot capture the laser signal.

控制飞行器飞向虚拟目标位置，In step 2, the following formula is passed in the load calculation module 7

Control the aircraft to fly to the virtual target position,

为飞行器与目标的视线角速度，此时的飞行器与目标的视线角速度由下式得到：Among them, the navigation ratio N is 2 to 4, preferably 4, a _M is the required overload of the aircraft, V is the speed of the aircraft, θ is the roll angle, which is directly measured by the inertial gyroscope,

is the line-of-sight angular velocity between the aircraft and the target. The line-of-sight angular velocity between the aircraft and the target is obtained by the following formula:

为飞行器与目标的视线角速度；x_F表示虚拟目标位置对应的经度坐标，y_F表示虚拟目标位置对应的纬度坐标，x_t表示卫星模块实时探测得到的飞行器自身当前位置的经度坐标，y_t表示卫星模块实时探测得到的飞行器自身当前位置的纬度坐标。γ is the ballistic inclination angle, and its value is equivalent to the pitch angle of the aircraft, which can be directly measured by the attitude gyro; q is the line-of-sight angle between the aircraft and the target,

is the line-of-sight angular velocity between the aircraft and the target; x _F represents the longitude coordinate corresponding to the virtual target position, y _F represents the latitude coordinate corresponding to the virtual target position, x _t represents the longitude coordinate of the current position of the aircraft itself detected in real time by the satellite module, and y _t represents The latitude coordinates of the current position of the aircraft itself detected by the satellite module in real time.

In a preferred embodiment, in step 3, in the terminal guidance section, the laser seeker is a strapdown seeker on the aircraft, which can be used to detect the line-of-sight angle between the aircraft and the target in real time, but it is difficult to The line-of-sight angular velocity information of the aircraft and the target is directly obtained through the laser seeker.

In a preferred embodiment, in step 4, in the terminal guidance section, the correction module 5 receives the line-of-sight angle q _G between the satellite aircraft and the target and the line-of-sight angle q _laser between the laser aircraft and the target in real time, and compares and receives the satellite aircraft. The difference between the sight angle q _G with the target and the sight angle q _laser of the laser aircraft and the target, when the difference is greater than or equal to a predetermined value, output the sight angle q _G of the satellite aircraft and the target; when the difference is less than the predetermined value When the value is , the line-of-sight angle q _laser between the laser aircraft and the target is output; the predetermined value is 0.2° to 1°, preferably 0.5°.

Through the above method, by setting a specific judgment threshold value in the correction module 5, it is possible to timely and accurately select parameters that are closer to the true value from the same parameters provided by different modules, and can quickly and efficiently select a more accurate value. The line-of-sight information between the aircraft and the target that is close to the true value.

In a preferred embodiment, in step 5, in the final guidance section, the line-of-sight angular velocity calculation module 6 calculates the line-of-sight angular velocity information of the aircraft and the target in real time through the following equations (1), (2) and (3) ,

表示x₂的导数，

表示x₁的导数，

表示x₀的导数，上一时刻得到的数值作为下一时刻迭代的初始值；Among them, q _g represents the line-of-sight angle between the aircraft and the target transmitted by the correction module 5, and q ₀ represents the estimated value of the line-of-sight angle between the aircraft and the target, that is, the above formulas (1), (2), ( 3) The estimated value of the line-of-sight angle between the aircraft and the target, q ₁ represents the estimated value of the line-of-sight angular velocity between the aircraft and the target, that is, in the calculation process, it is estimated by the above equations (1), (2), (3) The estimated value of the line-of-sight angular velocity between the aircraft and the target;

represents the derivative of _x2 ,

represents the derivative of x ₁ ,

Represents the derivative of x ₀ , and the value obtained at the previous moment is used as the initial value of the iteration at the next moment;

At the initial moment x ₀ =0, x ₁ =0, x ₂ =0, every 0.001s is used as the integration step, iterates, and obtains the values of x ₀ , x ₁ and x ₂ at the next moment;

进而得到下一时刻的初始值x₀、x₁和x₂；再将得到的x₀、x₁、x₂和接收到的q_g值代入到式(一)、式(二)和式(三)中，从而得到再下一时刻对应的

如此持续循环迭代即可持续得到每次积分得到的对应x₀、x₁、x₂。Specifically, in the first iteration moment, the initial moment x ₀ =0, x ₁ =0, x ₂ =0 and the received q _g value are substituted into equations (1), (2) and (3) ), so that it can be solved

Then obtain the initial values x ₀ , x ₁ and x ₂ at the next moment; then substitute the obtained x ₀ , x ₁ , x ₂ and the received q _g values into formula (1), formula (2) and formula ( 3), so as to obtain the corresponding

In this way, the corresponding x ₀ , x ₁ , and x ₂ obtained by each integration can be continuously obtained by continuing the loop iteration.

表示飞行器与目标的视线角速度，实时输出该飞行器与目标的视线角速度给过载解算模块7即可用于解算需用过载。in,

Indicates the line-of-sight angular velocity between the aircraft and the target, and outputting the line-of-sight angular velocity between the aircraft and the target in real time to the overload calculation module 7 can be used to calculate the required overload.

Wherein, the a ₀ , a ₁ , a ₂ , δ, k ₁ and k ₂ are all design parameters. In this application, preferably, a ₀ =1-1.5, a ₁ =7-10, a ₂ =10- 15. δ=1～2, k ₁ =0.1～0.4 and k ₂ =0.2～0.4;

More preferably, a ₀ =1.1, a ₁ =8.5, a ₂ =11.5, δ=1.5, k ₁ =0.3, k ₂ =0.3.

导航比N取值为4，a_M为飞行器的需用过载，V为飞行器的速度，

为飞行器与目标的视线角速度，即为视线角速度解算模块6实时给出的飞行器与目标的视线角速度。In a preferred embodiment, in step 6, in the final guidance section, the overload resolution module 7 uses a proportional guidance guidance law to perform guidance control, that is,

The value of the navigation ratio N is 4, a _M is the required overload of the aircraft, V is the speed of the aircraft,

is the line-of-sight angular velocity of the aircraft and the target, that is, the line-of-sight angular velocity of the aircraft and the target given by the line-of-sight angular velocity calculation module 6 in real time.

According to the multi-mode guidance method for a high-dynamic aircraft in a large-span complex environment provided by the present invention, a combination of strapdown seeker and satellite guidance is used to control the flight trajectory of the aircraft in the middle guidance section by setting a virtual target position. , so that the laser seeker can quickly capture the target and enter the working state when entering the terminal guidance segment; the correction module 5 is used to filter out the line-of-sight angle information between the aircraft and the target obtained by the satellite module and the laser seeker. Then, through the line-of-sight angular velocity calculation module 6, the line-of-sight angular velocity of the aircraft and the target with strong accuracy and robustness is calculated in real time, and then the required overload is calculated based on this to improve the guidance accuracy.

Experimental example

The simulation experiment of the aircraft is carried out through the computer. The simulation conditions of the rotating aircraft are: the flight speed of the rotating aircraft is 580m/s, and the rotation speed is 11.6r/s;

The rotating aircraft can be directly simulated by the computer, and the line-of-sight angle between the aircraft and the target corresponding to the rotating aircraft and the line-of-sight angular velocity between the aircraft and the target can be given in real time. The line-of-sight angle between the aircraft and the target is sent to the rotating aircraft in real time as an input value;

A line-of-sight angular velocity calculation module is arranged in the rotating aircraft, and the line-of-sight angular velocity calculation module can receive the line-of-sight angle information between the aircraft and the target in real time, and solve the real-time solution through the following formulas (1), (2) and (3) Calculate the line-of-sight angular velocity information of the aircraft and the target, and obtain the calculated value of the line-of-sight angular velocity of the aircraft and the target;

Wherein, a ₀ =1.1, a ₁ =8.5, a ₂ =11.5, δ=1.5, k ₁ =0.3, k ₂ =0.3.

表示x₂的导数，

表示x₁的导数，

表示x₀的导数，上一时刻得到的数值作为下一时刻迭代的初始值；q _g represents the line-of-sight angle information received in real time between the aircraft and the target, q ₀ represents the estimated value of the line-of-sight angle between the aircraft and the target, that is, in the calculation process, it is estimated by the above equations (1), (2), (3) The estimated value of the line-of-sight angle between the aircraft and the target, q ₁ represents the estimated value of the line-of-sight angular velocity between the aircraft and the target, that is, the aircraft estimated by the above formulas (1), (2), (3) during the calculation process. an estimate of the line-of-sight angular velocity with the target;

represents the derivative of _x2 ,

represents the derivative of x ₁ ,

Represents the derivative of x ₀ , and the value obtained at the previous moment is used as the initial value of the iteration at the next moment;

At the initial moment x ₀ =0, x ₁ =0, x ₂ =0, every 0.001s is used as the integration step, iterates, and obtains the values of x ₀ , x ₁ and x ₂ at the next moment;

进而得到下一时刻的初始值x₀、x₁和x₂；再将得到的x₀、x₁、x₂和接收到的q_g值代入到式(一)、式(二)和式(三)中，从而得到再下一时刻对应的

如此持续循环迭代即可持续得到每次积分得到的对应x₀、x₁、x₂。Specifically, in the first iteration moment, the initial moment x ₀ =0, x ₁ =0, x ₂ =0 and the received q _g value are substituted into equations (1), (2) and (3) ), so that it can be solved

Then obtain the initial values x ₀ , x ₁ and x ₂ at the next moment; then substitute the obtained x ₀ , x ₁ , x ₂ and the received q _g values into formula (1), formula (2) and formula ( 3), so as to obtain the corresponding

In this way, the corresponding x ₀ , x ₁ , and x ₂ obtained by each integration can be continuously obtained by continuing the loop iteration.

表示飞行器与目标的视线角速度，即为所述飞行器与目标的视线角速度的解算值。in,

Indicates the line-of-sight angular velocity between the aircraft and the target, which is the calculated value of the line-of-sight angular velocity between the aircraft and the target.

Experimental example 1 does not consider the influence of internal and external disturbances on the system

The line-of-sight angular velocity between the aircraft and the target given by the computer in real time is shown in the dotted line "the true value of the line-of-sight angular velocity between the aircraft and the target" in Figure 3. The line-of-sight angular velocity between the aircraft and the target remains unchanged within two seconds. The obtained line-of-sight angle between the aircraft and the target is passed to the aircraft filled with the line-of-sight angular velocity calculation module, and the final calculated value of the line-of-sight angular velocity between the aircraft and the target is shown in Fig. The solution value of ” is shown,

It can be seen from FIG. 3 that the calculated value of the line-of-sight angular velocity of the aircraft and the target calculated by the line-of-sight angular velocity calculation module provided by the application can quickly approach the real line-of-sight angular velocity, and the required time is less than 0.4 seconds, and the real line-of-sight angular velocity is tracked. Afterwards, it can change with the change of the real line of sight angular velocity, which can prove that the line-of-sight angular velocity calculation module has strong accuracy.

Experimental Example 2 Considering the influence of internal and external disturbances on the system

The line-of-sight angular velocity of the aircraft and the target given by the computer in real time is shown in the thin solid line "the true value of the line-of-sight angular velocity of the aircraft and the target" in Figure 4. The line-of-sight angle with the target is passed to the aircraft filled with the line-of-sight angular velocity calculation module, and the final calculated value of the line-of-sight angular velocity between the aircraft and the target is shown in Figure 4. The dotted line "The solution value of the line-of-sight angular velocity between the aircraft and the target" shown,

It can be seen from FIG. 4 that the calculated value of the line-of-sight angular velocity of the aircraft and the target calculated by the line-of-sight angular velocity calculation module provided by the present application can quickly approach the real line-of-sight angular velocity, and the required time is less than 0.4 seconds, and the real line-of-sight angular velocity is tracked. Afterwards, it can change with the change of the real line-of-sight angular velocity, which can prove that the line-of-sight angular velocity calculation module has strong robustness and accuracy.

The present invention has been described above with reference to the preferred embodiments, but these embodiments are merely exemplary and serve only for illustrative purposes. On this basis, various substitutions and improvements can be made to the present invention, which all fall within the protection scope of the present invention.

Claims (10)

1. a high dynamic aircraft multi-mode guidance system under a large cross-domain complex environment, is characterized in that, this guidance system comprises satellite module (1), laser seeker (2), attitude sensitive module (3), virtual target module (4), a correction module (5), a line-of-sight angular velocity calculation module (6) and an overload calculation module (7); The satellite module (1) is used for receiving satellite signals, and according to the satellite signals, the position information, speed information and line-of-sight information of the aircraft and the target are calculated in real time; The laser seeker (2) is used to detect the line-of-sight information between the aircraft and the target in the terminal guidance segment; The attitude sensitive module (3) is used to obtain the flight parameter information of the aircraft sensitively in real time, The virtual target module (4) is used to provide a virtual target position for the aircraft, so that the aircraft flies to the virtual target position in the middle guidance section, The correction module (5) is respectively connected with the satellite module (1) and the laser seeker (2), and outputs the line-of-sight angle information of the aircraft and the target that is close to the true value; The line-of-sight angular velocity calculation module (6) is used to obtain the line-of-sight angular velocity of the aircraft and the target in real time according to the line-of-sight angle information of the aircraft and the target that is close to the true value output by the correction module (5); The overload calculation module (7) is used for obtaining the required overload according to the line-of-sight angular velocity of the aircraft and the target obtained in real time by the line-of-sight angular velocity calculation module (6); The virtual target position is stored in the virtual target module (4), and the virtual target module (4) obtains the virtual target position by the following formula:

Among them, x _F represents the x-axis coordinate in the plumb plane where the virtual target position is located, y _F represents the y-axis coordinate in the plumb plane where the virtual target position is located, c _t and c _g are design gains, v _g is the speed of the aircraft, θ is the ballistic inclination,

is the ballistic declination angle, α is the attack angle of the aircraft, β is the sideslip angle of the aircraft, OT is the lateral distance between the launch origin and the target, L _c is the gliding distance of the aircraft, and T _s is the preset attitude adjustment time. 2. The high-dynamic aircraft multi-mode guidance system according to claim 1, characterized in that, The satellite module (1) includes four composite antennas (11), an anti-jamming module (12) and a satellite solving sub-module (13), Wherein, the four composite antennas (11) include four sheet-like antenna plates for receiving satellite signals; The anti-jamming module (12) is connected with the four composite antennas (11), and is used for filtering the satellite signal to eliminate noise interference in the satellite signal; The satellite solving sub-module 13 receives the filtered satellite signal, converts the satellite signal into a navigation message, and then calculates the position information, speed information and line-of-sight information of the aircraft and the target at the current moment according to the navigation message. 3. The high-dynamic aircraft multi-mode guidance system according to claim 2, characterized in that, The four composite antennas (11) are uniformly distributed around the aircraft and arranged along the circumferential direction of the aircraft's rolling. 4. The high-dynamic aircraft multi-mode guidance system according to claim 1, characterized in that, The attitude sensitive module (3) includes an inertial gyro and a geomagnetic sensor, The pitch angle information and yaw angle information of the aircraft are obtained sensitively in real time through the inertial gyroscope, The roll angle information of the aircraft is obtained sensitively in real time through the geomagnetic sensor. 5. The high-dynamic aircraft multi-mode guidance system according to claim 1, characterized in that, The correction module (5) receives the line-of-sight angle q _G between the satellite aircraft and the target and the line-of-sight angle q _laser between the laser aircraft and the target in real time in the terminal guidance section, When the difference between the two is greater than or equal to a predetermined value, output the line-of-sight angle q _G between the satellite vehicle and the target; When the difference between the two is less than a predetermined value, output the line-of-sight angle q _laser between the laser aircraft and the target. 6. The multi-mode guidance system for a high-dynamic aircraft in a large-span complex environment according to claim 5, characterized in that, The predetermined value is 0.2°˜1°. 7 . The multi-mode guidance system for a highly dynamic aircraft in a large-span complex environment according to claim 6 , wherein the predetermined value is 0.5°. 8 . 8. The high-dynamic aircraft multi-mode guidance system according to claim 1, characterized in that, In the line-of-sight angular velocity calculation module (6), the line-of-sight angular velocity information of the aircraft and the target is calculated in real time by the following equations (1), (2) and (3),

Among them, q _g represents the line-of-sight angle between the aircraft and the target transmitted by the correction module (5), q ₀ represents the estimated value of the line-of-sight angle between the aircraft and the target, q ₁ represents the estimated value of the line-of-sight angular velocity between the aircraft and the target,

represents the derivative of _x2 ,

represents the derivative of x ₁ ,

represents the derivative of x ₀ ; The initial moment x ₀ =0, x ₁ =0, x ₂ =0, for any moment, every 0.001s is used as the integration step, iterates, and obtains the values of x ₀ , x ₁ and x ₂ at the next moment; in,

Represents the line-of-sight angular velocity information of the aircraft and the target obtained by real-time calculation and transmits it to the overload calculation module (7); Wherein, the a ₀ , a ₁ , a ₂ , δ, k ₁ and k ₂ are all design parameters. 9. The multi-mode guidance system for a high-dynamic aircraft in a large-span complex environment according to claim 1, characterized in that, In the overload calculation module (7), the proportional guidance guidance law is used for guidance control, that is,

The value of the navigation ratio N is 4, a _M is the required overload of the aircraft, V is the speed of the aircraft,

is the line-of-sight angular velocity of the aircraft and the target, that is, the line-of-sight angular velocity of the aircraft and the target given in real time by the line-of-sight angular velocity calculation module (6). 10. A high-dynamic aircraft multi-mode guidance system based on the large-span complex environment described in one of claims 1 to 9 realizes a high-dynamic aircraft multi-mode guidance method in a large-span complex environment, the method comprising the steps: Step 1, receive the satellite signal through the satellite module (1), and calculate the position information, speed information and the line-of-sight information of the aircraft and the target in real time according to the satellite signal; The flight parameter information of the aircraft is obtained in real time by the attitude sensitive module (3); Step 2, in the middle guidance section, the virtual target position is provided for the aircraft by the virtual target module (4), so that the aircraft flies to this virtual target position in the middle guidance section, Step 3, in the terminal guidance section, the laser seeker (2) is used to detect the line-of-sight information between the aircraft and the target in the terminal guidance section; Step 4, in the terminal guidance section, output the line-of-sight information of the aircraft and the target that is close to the true value through the correction module (5); Step 5, in the terminal guidance section, obtain the angular velocity of sight of the aircraft and the target in real time according to the sight angle information of the aircraft and the target that is close to the true value output by the angular velocity of sight calculation module (6) according to the correction module (5); Step 6, in the final guidance section, the overload calculation module (7) obtains the required overload according to the line-of-sight angular velocity of the aircraft and the target obtained in real time by the line-of-sight angular velocity calculation module (6).

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