Guidance control method and system for remote multi-mode composite guidance aircraft

Abstract

The invention discloses a guidance control method and system for a remote multi-mode composite guidance aircraft, in the system, an active radar / infrared multi-mode composite guidance strategy is used in a terminal guidance section, and the information acquisition capability of a detection system is enhanced by utilizing complementary advantages of the active radar and the infrared multi-mode composite guidance strategy; the ability of the aircraft to accurately hit a target under the influence of various interference factors is guaranteed from the aspect of detection, the aircraft is subjected to inertial guidance in the middle guidance section, and the aircraft is controlled through an overweight complementary proportion guidance law to perform range-extending gliding, so that the accurate hit range of the aircraft is increased from 20 km to 40 km; the terminal guidance section of the aircraft is controlled by adopting a novel terminal sliding mode guidance law at the terminal guidance section, and the falling angle error is controlled to be within 2 degrees from 10 degrees on the basis of ensuring accurate hit.

Description

Technical Field

[0001] The invention relates to a guidance control method for an aircraft, and in particular to a guidance control method and system for a long-range multi-mode composite guidance aircraft. Background Art

[0002] With the continuous development of science and technology, precision strikes on targets have become an important technology in modern warfare. The ability to accurately strike the opponent's key targets can not only enable one's side to quickly establish an advantage in the confrontation, but also has the advantages of being less likely to cause accidental injuries and low cost. Therefore, precision strike technology has gradually become the key to confrontation. As an important equipment for precision strikes, the requirements for precision-guided aircraft have also increased, but problems have also emerged.

[0003] First, in terms of detection accuracy, as the means of interference hiding become more and more abundant, the dual-state composite guidance in the existing technology is difficult to determine the timing of state transition, which leads to errors in information acquisition and thus affects the accuracy of aircraft hits;

[0004] In order to ensure our safety, the aircraft needs to have a longer range, but this often means reducing accuracy and changing the landing angle. The technical solution of increasing the range without reducing accuracy is an important goal pursued by R&D personnel.

[0005] In addition, whether the aircraft can hit the weak link of the target is a key factor in whether it can cause greater damage. The existing technology can constrain the landing angle error to within 10°. If it can be further refined to within 2°, the actual damage effect will be greatly improved.

[0006] Based on the above problems, the inventors have conducted in-depth research on the aircraft guidance and control system, adopted a high-precision control scheme in the mid-guidance stage, proposed a new detection method in the terminal guidance stage, and adopted a new control method to improve the hit accuracy and landing angle accuracy, thus proposing a guidance and control method and system for a long-range multi-mode composite guided aircraft. Summary of the invention

[0007] In order to overcome the above problems, the inventors have conducted intensive research and designed a guidance control method and system for long-range multi-mode composite guided aircraft. In this system, an active radar / infrared multi-mode composite guidance strategy is used in the terminal guidance stage. The advantages of the two are complementary to each other, the information acquisition capability of the detection system is enhanced, and the ability of the aircraft to accurately hit the target under the influence of various interference factors is guaranteed from the detection aspect. In the middle guidance stage, the aircraft is inertially guided and controlled by the overweight proportional guidance law to perform extended-range gliding, which realizes an increase in the accurate hit range of the aircraft from 20km to 40km. In the terminal guidance stage, a new terminal sliding mode guidance law is used to control the terminal guidance stage of the aircraft. On the basis of ensuring accurate hits, the angle of impact error is controlled from 10° to within 2°; thereby completing the present invention.

[0008] Specifically, the object of the present invention is to provide a guidance control method for a long-range multi-mode composite guided aircraft, the method comprising:

[0009] The roll angle and position of the aircraft are measured by the inertial navigation module, and the roll angle information and position information are transmitted as input to the central processing module.

[0010] The active radar guidance module measures the target's position, the relative speed between the missile and the target, and the missile-target sight angle, and transmits the measured information as input to the central processing module.

[0011] The infrared guidance module measures the sight angle between the missile and the target, and sends it as input to the central processing module.

[0012] Through the central processing module, the required overload of the aircraft is obtained in real time based on the received information, and then the required overload is decomposed according to the roll angle information to generate a rudder deviation command. The final rudder deviation command is transmitted to the actuator to control the aircraft to fly towards the target.

[0013] When the aircraft is in the middle guidance stage, the central processing module obtains the required overload a of the aircraft through the overload compensation proportional guidance law. m ,

[0014] When the aircraft is in the terminal guidance phase, the central processing module obtains the required overload a of the aircraft through the new terminal sliding mode guidance law. m .

[0015] The overload compensation proportional guidance law obtains the required overload of the aircraft through the following formula (1):

[0016]

[0017] Where N is the proportional guidance coefficient,

[0018] Indicates the relative speed of the projectile and the target.

[0019] represents the missile-to-target line-of-sight angular velocity,

[0020] c represents the gravity compensation term,

[0021] g represents the acceleration due to gravity.

[0022] The novel terminal sliding mode guidance law obtains the required overload of the aircraft through the following formula (II):

[0023]

[0024] Among them, f1(x) represents the basic overload function,

[0025] f2(x) represents the overload gain function,

[0026] μ represents the overload factor,

[0027] g(x1) represents the basic sliding mode function,

[0028] s represents the sliding mode variable,

[0029] d(g(x1)) represents the rate of change of g(x1) with respect to time,

[0030] dt is the time integration step.

[0031] The present application also provides a guidance control system for a long-range multi-mode composite guided aircraft, the system comprising an inertial navigation module, an active radar guidance module, an infrared guidance module and a central processing module.

[0032] The roll angle and position of the aircraft are measured by the inertial navigation module, and the roll angle information and position information are transmitted as input to the central processing module.

[0033] The active radar guidance module measures the position of the target, the speed of the target and the sight angle between the missile and the target, and transmits the measured information as input to the central processing module.

[0034] The infrared guidance module measures the sight angle between the missile and the target, and sends it as input to the central processing module.

[0035] The required overload of the aircraft is obtained through the central processing module, and then the required overload is decomposed according to the roll angle information to generate a rudder deviation instruction, and the final rudder deviation instruction is transmitted to the actuator to control the aircraft to fly to the target.

[0036] The beneficial effects of the present invention include:

[0037] (1) According to the guidance and control method for a long-range multi-mode composite guided aircraft provided by the present invention, the central processing module adopts inertial guidance for the aircraft in the mid-guidance stage and controls it through the overweight compensation proportional guidance law to make it glide with extended range, thereby achieving an increase in the aircraft's precise hit range from 20km to 40km;

[0038] (2) According to the guidance and control method for a long-range multi-mode composite guided aircraft provided by the present invention, the central processing module adopts a new terminal sliding mode guidance law in the terminal guidance stage, which enhances the anti-interference capability of the aircraft and has stronger robustness, so that the process of the aircraft reaching the desired landing angle is faster and smoother, and on the basis of ensuring accurate hits, the landing angle error is controlled from 10° to within 2°.

[0039] (3) According to the guidance and control method for a long-range multi-mode composite guided aircraft provided by the present invention, a new terminal sliding surface is used in the central processing module to accelerate the convergence speed of the sliding mode control, while having the same robustness as the terminal sliding surface.

[0040] (4) According to the guidance and control method for a long-range multi-mode composite guided aircraft provided by the present invention, the angle information obtained by the convex combination fusion algorithm for the active radar guidance module and the infrared guidance module is retrieved, and the missile-target line-of-sight angle is obtained and transmitted to the central processing module in combination with the relative distance between the missile and the target provided by the active radar guidance module, thereby obtaining the required overload of the aircraft, thereby achieving the purpose of obtaining target information with higher accuracy and better stability; if one of the active radar guidance module and the infrared guidance module fails, the other module will provide information separately. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] Figure 1 A schematic diagram showing the overall logical structure of the guidance and control method for a long-range multi-mode composite guided aircraft according to the present application;

[0042] Figure 2 A schematic diagram of the trajectory of an aircraft in an embodiment is shown;

[0043] Figure 3 A schematic diagram of the trajectory of the aircraft in the embodiment and comparative example 1 is shown;

[0044] Figure 4 Schematic diagrams of aircraft trajectories in the embodiment and comparative example 2 are shown. DETAILED DESCRIPTION

[0045] The present invention will be further described in detail below through the accompanying drawings and embodiments. Through these descriptions, the characteristics and advantages of the present invention will become more clear and distinct.

[0046] 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. Although various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless otherwise noted.

[0047] According to a guidance control method for a long-range multi-mode composite guided aircraft provided by the present invention, in this method,

[0048] The roll angle and position of the aircraft are measured by the inertial navigation module, and the roll angle information and position information are transmitted to the central processing module as input. The inertial navigation module includes a mechanical gyroscope, which does not need to spend time to determine the zero reference, and can perform sensitive measurements of the three-axis angular rate of the projectile immediately after activation; the inertial navigation module also includes INS inertial elements such as a three-axis MEMS gyroscope and an accelerometer. The INS inertial elements need to be zero-aligned after activation, and its attitude information measurement needs a certain amount of time to converge to the true value, that is, the inertial navigation module starts alignment in advance before the aircraft enters the mid-guidance stage.

[0049] The position information of the target is pre-installed in the aircraft before the aircraft is launched (x t ,y t ); the aircraft is in an uncontrolled climbing state for a period of time after launch, relying on the aircraft's own lift to climb. When the loading time is reached, the space directional gyro is unlocked; when the control start time is reached, the rudder is opened, the guidance module on the aircraft is started, and it enters the mid-range guidance stage; the mid-range guidance stage of the aircraft is also called the extended-range gliding stage.

[0050] The inertial navigation module performs zero alignment after being activated for a period of time. Its attitude information measurement needs a certain period of time to converge to the true value and obtain the position information of the aircraft (x m ,y m ) The central processing module calculates the relative distance R between the missile and the target and the sight angle q between the missile and the target according to the received information of the position of the aircraft and the target;

[0051]

[0052]

[0053] In the mid-guidance stage, the required overload is obtained through the relative distance between the missile and the target and the sight angle between the missile and the target obtained above.

[0054] In this method, the position of the target, the speed of the target and the sight angle between the missile and the target are measured by the active radar guidance module, and the measured information is transmitted as input to the central processing module;

[0055] The active radar guidance module is composed of the antenna system, receiver, signal processing system, servo mechanism, transmitter and other main parts. The microwave / millimeter wave generated by the missile-borne transmitter is radiated through the seeker antenna, and the echo reflected by the target is received by the same antenna, completing the search phase for the target. The process from searching for the target to entering the tracking phase is the interception phase, which is mainly completed by the signal processor and the data processor. In the tracking phase, the speed tracking and distance tracking of the target are realized through the processing of the receiver and the signal processor. The seeker antenna with strong directivity is driven by the data processor and the servo mechanism to realize the angle tracking of the target.

[0056] In this method, the sight angle between the missile and the target is measured by the infrared guidance module and sent to the central processing module as input;

[0057] The infrared guidance module is composed of an optical system, a modulator, an infrared detector, a refrigerator, a servo mechanism and an electronic circuit. The optical system receives the infrared radiation of the target, which is processed by the modulator into an optical signal including the target information. The infrared detector converts the optical signal into an easy-to-process electrical signal, which is then filtered, amplified and processed by the electronic circuit to detect the target angular position information, and the information is sent to the servo mechanism to make the optical axis move in the target direction, so as to realize the continuous tracking of the target by the guidance system.

[0058] In the method, the central processing module obtains the required overload of the aircraft in real time according to the received information, and then decomposes the required overload according to the roll angle information to generate a rudder deviation instruction, and transmits the final rudder deviation instruction to the actuator to control the aircraft to fly to the target. The actuator is the steering gear servo mechanism.

[0059] In a preferred embodiment, when the aircraft is in the middle guidance stage, the central processing module obtains the required overload a of the aircraft through the overload compensation proportional guidance law. m ,

[0060] When the aircraft is in the terminal guidance phase, the central processing module obtains the required overload a of the aircraft through the new terminal sliding mode guidance law. m .

[0061] When the relative distance R between the missile and the target is less than the set distance, the aircraft enters the terminal guidance phase; the set distance is 3km-8km. When entering the terminal guidance phase, the active radar / infrared seeker is activated to capture the target. Under the control of the new sliding mode guidance law, the aircraft finally hits the target at the expected angle of impact.

[0062] In a preferred embodiment, the overload compensation proportional guidance law obtains the required overload of the aircraft through the following formula (1):

[0063]

[0064] Wherein, N represents the proportional guidance coefficient, which is generally 2-6 in engineering, and preferably 4;

[0065] It represents the relative motion speed of the projectile and the target. The relative distance between the projectile and the target can be obtained by processing the active radar guidance module using the convex combination fusion algorithm. The relative distance between the projectile and the target can also be obtained by the infrared guidance module. The relative motion speed of the projectile and the target can be obtained by differentiating the relative distance between the projectile and the target.

[0066] represents the missile-target line-of-sight angular rate, obtained by differentiating the missile-target line-of-sight angle;

[0067] c represents the gravity compensation coefficient, which ranges from 1 to 1.5;

[0068] g represents the acceleration due to gravity, and its value is 9.81.

[0069] In a preferred embodiment, the novel terminal sliding mode guidance law obtains the required overload of the aircraft through the following formula (II):

[0070]

[0071] Among them, f1(x) represents the basic overload function,

[0072] f2(x) represents the overload gain function,

[0073] μ represents the overload factor,

[0074] g(x1) represents the basic sliding mode function,

[0075] s represents the sliding mode variable,

[0076] d(g(x1)) represents the rate of change of g(x1) with respect to time,

[0077] dt is the time integration step.

[0078] Preferably, the expression of f1(x) is as shown in the following formula (III):

[0079]

[0080] in, Indicates the relative speed of the projectile and the target.

[0081] R represents the relative distance between the projectile and the target.

[0082] Indicates the expected line of sight angular velocity, which is generally a fixed value, preferably 0.

[0083] x2 represents the state variable, and the state variable x2 is selected as

[0084] Indicates the missile-target line-of-sight angular velocity.

[0085] Preferably, the expression of f2(x) is as shown in the following formula (iv):

[0086]

[0087] Among them, R represents the relative distance between the projectile and the target,

[0088] represents the expected line of sight angular velocity,

[0089] θ represents the ballistic inclination angle of the aircraft, which is calculated by the mechanical gyroscope of the inertial navigation module.

[0090] q d represents the desired landing angle of the aircraft,

[0091] x1 represents the state variable, and the state variable x1 is selected as x1=qq d ,

[0092] q represents the sight angle of the missile.

[0093] The missile-target sight angle q in the terminal guidance stage is obtained by processing the missile-target sight angle obtained by the active radar guidance module and the missile-target sight angle obtained by the infrared guidance module using a convex combination fusion algorithm; for the convex combination fusion algorithm, see Liu Weidong, Liu Yang, Gao Li'e, Research on track fusion algorithm based on convex combination and Bar-Shalom-Campo, Computer Engineering and Applications.

[0094] Preferably, the sliding mode variable s is obtained by the following formula (V):

[0095] s=g(x1)+x2 (V)

[0096] Among them, g(x1) is obtained by the following formula (VI):

[0097]

[0098] k0, k1, k2, α, β each independently represent a design parameter, wherein k0>0, k1, k2>1, α, β are both odd numbers and 1 / 2<α / β<1;

[0099] Preferably, specific values ​​of the design parameters are: k0=1, k1=2, k2=20, α / β=3 / 5.

[0100] In the mid-guidance and terminal guidance stages, the central processing module will decompose the calculated required overload according to the roll angle given by the mechanical gyroscope of the inertial navigation module, obtain the rudder deflection command, and transmit the final rudder deflection command to the actuator to control the aircraft attitude and finally control the aircraft to the desired landing angle q d Hit the target precisely.

[0101] The present application also provides a guidance and control system for a long-range multi-mode composite guided aircraft, the system comprising a guidance and control system, a central processing module and a steering gear servo mechanism; the guidance and control system specifically comprises an inertial navigation module, an active radar guidance module and an infrared guidance module.

[0102] The roll angle and position of the aircraft are measured by the inertial navigation module, and the roll angle information and position information are transmitted as input to the central processing module.

[0103] The active radar guidance module measures the position of the target, the speed of the target and the sight angle between the missile and the target, and transmits the measured information as input to the central processing module.

[0104] The infrared guidance module measures the sight angle between the missile and the target, and sends it as input to the central processing module.

[0105] The required overload of the aircraft is obtained through the central processing module, and then the required overload is decomposed according to the roll angle information to generate a rudder deviation instruction, and the final rudder deviation instruction is transmitted to the actuator to control the aircraft to fly to the target;

[0106] When the aircraft is in the middle guidance stage, the central processing module obtains the required overload a of the aircraft through the overload compensation proportional guidance law. m ,

[0107] When the aircraft is in the terminal guidance phase, the central processing module obtains the required overload a of the aircraft through the new terminal sliding mode guidance law. m ;

[0108] The overload compensation proportional guidance law obtains the required overload of the aircraft through the following formula (1):

[0109]

[0110] Where N is the proportional guidance coefficient,

[0111] Indicates the relative speed of the projectile and the target.

[0112] represents the missile-to-target line-of-sight angular velocity,

[0113] c represents the gravity compensation term,

[0114] g represents the acceleration due to gravity;

[0115] The novel terminal sliding mode guidance law obtains the required overload of the aircraft through the following formula (II):

[0116]

[0117] Among them, f1(x) represents the basic overload function,

[0118] f2(x) represents the overload gain function,

[0119] μ represents the overload factor,

[0120] g(x1) represents the basic sliding mode function,

[0121] s represents the sliding mode variable,

[0122] d(g(x1)) represents the rate of change of g(x1) with respect to time,

[0123] dt is the time integration step.

[0124] The system also includes a power supply master control device, which is connected to the thermal power supply on the aircraft and is responsible for supplying power to the entire system. It can enable each module to work normally at rated power according to the required voltage requirements.

[0125] Example 1

[0126] The initial position of the aircraft in the inertial coordinate system is the origin of the coordinate system. The initial relative distance between the missile and the target in the terminal guidance stage is: R0=8000m.

[0127] Aircraft speed: V m0 =600m / s;

[0128] The initial ballistic inclination of the aircraft: ;

[0129] The initial sight angle between the aircraft and the target: q0 = 0°;

[0130] Set the target to a fixed building on the ground.

[0131] After the aircraft enters the mid-guidance stage, the simulation system simulates and outputs the aircraft's roll angle, aircraft position, and missile-target sight angle in real time, and then obtains the missile-target relative motion speed, i.e., the aircraft speed, based on the aircraft's position change speed. The central processing module obtains the required overload of the aircraft through the following formula (I):

[0132]

[0133] Then, based on the roll angle information, the required overload is decomposed to generate a rudder deviation command, which is then transmitted to the actuator to control the aircraft to fly towards the target.

[0134] The specific values ​​in formula (1) are N=4, c=1.5, g=9.8;

[0135] After the missile-target distance is reduced to 8 km, the aircraft enters the terminal guidance phase, and the target position, missile-target relative speed, and missile-target sight angle are simulated and output in real time by the simulation system, and then the preset expected landing angle q is called. d In the final guidance stage, the central processing module obtains the required overload of the aircraft through the following formula (II):

[0136]

[0137] Among them, the expression of f1(x) is shown in the following formula (III):

[0138]

[0139] The expression of f2(x) is shown in the following formula (IV):

[0140]

[0141] The value of μ is 0.5,

[0142] The sliding mode variable s is obtained by the following formula (5):

[0143] s=g(x1)+x2 (V)

[0144] g(x1) is obtained by the following formula (VI):

[0145]

[0146] The specific values ​​in the terminal guidance phase are k0=1, k1=2, k2=20, α / β=3 / 5;

[0147] Expected fall angle q d The value is 90 degrees.

[0148] Then, according to the roll angle information, the required overload is decomposed to generate the rudder deflection command, and the final rudder deflection command is transmitted to the actuator to control the aircraft to fly to the target. The final flight trajectory is as follows: Figure 2 , Figure 3 and Figure 4 The solid line trajectory is shown in .

[0149] Example 2

[0150] The configuration is basically the same as that of embodiment 1, except that the desired landing angle qd The value is 80 degrees; the final flight trajectory is as follows Figure 2 as shown in .

[0151] Example 3

[0152] The configuration is basically the same as that of embodiment 1, except that the desired landing angle q d The value is 70 degrees; the final flight trajectory is as follows Figure 2 as shown in .

[0153] Example 4

[0154] The configuration is basically the same as that of embodiment 1, except that the desired landing angle q d The value is 60 degrees; the final flight trajectory is as follows Figure 2 as shown in .

[0155] Comparative Example 1

[0156] The experimental process is similar to that of Example 1, except that in the middle guidance section, the proportional guidance rate is used for guidance, that is,

[0157] The final flight trajectory is as follows Figure 3 This is shown by the middle dotted line before the extended range curve.

[0158] Comparative Example 2

[0159] An experimental process similar to that of Example 1 is set up, except that, in the terminal guidance stage, the aircraft is guided and controlled by a conventional sliding film control method, as described in Zhao Yao, Sheng Yongzhi Liu Xiangdong, Sliding mode-based continuous guidance law with terminal angle constraint, Chinese Journal of Aeronautics and Astronautics (English Edition).

[0160] The final flight trajectory is as follows Figure 4 Shown by the dashed line.

[0161] It can be seen from the above-mentioned Examples 1 to 4 and Comparative Example 1 that the guidance control system and method for a long-range multi-mode composite guided aircraft provided by the present application can control the aircraft so that the precise hit range of the aircraft is increased from 20km to 35km;

[0162] It can be seen from the above-mentioned Examples 1 to 4 and Comparative Example 2 that the guidance control system and method for a long-range multi-mode composite guided aircraft provided by the present application improve the robustness of the control process, and at the same time enhance the anti-interference ability of the aircraft, so that the process of the aircraft reaching the desired landing angle is faster and smoother. The method adopted in this patent can enable the aircraft to achieve an improvement in the landing angle error from 10° to 2° on the basis of ensuring accurate hits.

[0163] The present invention has been described above in conjunction with preferred embodiments, but these embodiments are only exemplary and serve only as an illustration. On this basis, the present invention may be subjected to a variety of substitutions and improvements, all of which fall within the scope of protection of the present invention.

Claims

1. A guidance control method applied to a long-range multi-mode composite guided aircraft, characterized in that: In this method, The roll angle and position of the aircraft are measured by the inertial navigation module, and the roll angle information and position information are transmitted as input to the central processing module. The active radar guidance module measures the target's position, the relative speed between the missile and the target, and the sight angle between the missile and the target, and transmits the measured information as input to the central processing module. The infrared guidance module measures the sight angle between the missile and the target and sends it as input to the central processing module. Through the central processing module, the required overload of the aircraft is obtained in real time based on the received information, and then the required overload is decomposed according to the roll angle information to generate a rudder deviation command. The final rudder deviation command is transmitted to the actuator to control the aircraft to fly towards the target.

2. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 1, characterized in that: When the aircraft is in the middle guidance stage, the central processing module obtains the required overload a of the aircraft through the overload compensation proportional guidance law. m , When the aircraft is in the terminal guidance phase, the central processing module obtains the required overload a of the aircraft through the new terminal sliding mode guidance law. m .

3. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 2, characterized in that: The overload compensation proportional guidance law obtains the required overload of the aircraft through the following formula (1): Where N is the proportional guidance coefficient, Indicates the relative speed of the projectile and the target. represents the missile-to-target line-of-sight angular velocity, c represents the gravity compensation term, g represents the acceleration due to gravity.

4. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 2, characterized in that: The novel terminal sliding mode guidance law obtains the required overload of the aircraft through the following formula (II): Among them, f1(x) represents the basic overload function, f2(x) represents the overload gain function, μ represents the overload factor, g(x1) represents the basic sliding mode function, s represents the sliding mode variable, d(g(x1)) represents the rate of change of g(x1) with respect to time, dt is the time integration step.

5. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 4, characterized in that: The expression of f1(x) is shown in the following formula (III): in, Indicates the relative speed of the projectile and the target. R represents the relative distance between the projectile and the target. represents the expected line of sight angular velocity, x2 represents the state variable, and the state variable x2 is selected as Indicates the missile-target line-of-sight angular velocity.

6. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 4, characterized in that: The expression of f2(x) is shown in the following formula (IV): Among them, R represents the relative distance between the projectile and the target, represents the expected line of sight angular velocity, θ represents the ballistic inclination angle of the aircraft, q d represents the desired landing angle of the aircraft, x1 represents the state variable, and the state variable x1 is selected as x1=qq d , q represents the sight angle of the missile.

7. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 4, characterized in that: The sliding mode variable s is obtained by the following formula (V): s = g(x1) + x2(V) Among them, g(x1) is obtained by the following formula (VI): k0, k1, k2, α, and β each independently represent a design parameter.

8. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 2, characterized in that: When the relative distance R between the missile and the target is less than the set distance, the aircraft enters the terminal guidance stage; the set distance is 3km-8km.

9. The guidance control method for a long-range multi-mode composite guided aircraft according to claim 1, characterized in that: In the central processing module, a convex combination fusion algorithm is used to process the missile-target line-of-sight angle obtained by the active radar guidance module and the missile-target line-of-sight angle obtained by the infrared guidance module to obtain a final missile-target line-of-sight angle, and the required overload is obtained based on the final missile-target line-of-sight angle.

10. A guidance and control system for a long-range multi-mode composite guided aircraft, characterized in that: The system includes an inertial navigation module, an active radar guidance module, an infrared guidance module and a central processing module. The inertial navigation module measures the roll angle and position of the aircraft, and transmits the roll angle information and position information as input to the central processing module. The active radar guidance module measures the position of the target, the speed of the target and the sight angle between the missile and the target, and transmits the measured information as input to the central processing module. The infrared guidance module measures the sight angle between the missile and the target, and sends it as input to the central processing module. The required overload of the aircraft is obtained through the central processing module, and then the required overload is decomposed according to the roll angle information to generate a rudder deviation instruction, and the final rudder deviation instruction is transmitted to the actuator to control the aircraft to fly to the target; When the aircraft is in the middle guidance stage, the central processing module obtains the required overload a of the aircraft through the overload compensation proportional guidance law. m , When the aircraft is in the terminal guidance phase, the central processing module obtains the required overload a of the aircraft through the new terminal sliding mode guidance law. m ; The overload compensation proportional guidance law obtains the required overload of the aircraft through the following formula (1): Where N is the proportional guidance coefficient, Indicates the relative speed of the projectile and the target. represents the missile-to-target line-of-sight angular velocity, c represents the gravity compensation term, g represents the acceleration due to gravity; The novel terminal sliding mode guidance law obtains the required overload of the aircraft through the following formula (II): Among them, f1(x) represents the basic overload function, f2(x) represents the overload gain function, μ represents the overload factor, g(x1) represents the basic sliding mode function, s represents the sliding mode variable, d(g(x1)) represents the rate of change of g(x1) with respect to time, and dt represents the time integration step.

Images