Guidance control method and system for a remote multi-mode composite guided vehicle

Abstract

The application discloses a guidance control method and system for a remote multi-mode composite guided aircraft, in the system, an active radar / infrared multi-mode composite guidance strategy is used in the terminal guidance section, the advantages of the two are complementary, the information acquisition capability of the detection system is enhanced, the detection aspect guarantees the capability of the aircraft to accurately hit the target under the influence of various interference factors, in the middle guidance section, the inertial guidance is adopted for the aircraft, and the over-weight compensation proportional guidance law control is used to make the aircraft to carry out the range-increasing gliding method, the accurate hitting range of the aircraft is improved from 20km to 40km, and in the terminal guidance section, the new terminal sliding mode guidance law is adopted to control the terminal guidance section of the aircraft, on the basis of guaranteeing the accurate hitting, the landing angle error is controlled to be within 2° from 10°.

Description

TECHNICAL FIELD

[0001] The present application relates to a guidance control method of an aircraft, in particular to a guidance control method and system for a long-range multi-mode compound guided aircraft. BACKGROUND

[0002] With the continuous development of science and technology, accurate attack on targets has become an important technology in modern warfare. Having the ability to accurately attack the opponent's key targets not only can quickly establish an advantage in confrontation, but also has the advantages of not being easy to cause misjudgment and low cost. Therefore, accurate attack technology has gradually become the key to confrontation. As an important equipment for accurate attack, the requirements for precise guided aircraft are also increasing, but problems also arise.

[0003] Firstly, in terms of detection accuracy, with the increasing richness of interference concealment means, the existing technology of duplex state compound guidance is difficult to determine the state transition opportunity, which leads to errors in obtaining information and affects the accuracy of the aircraft hitting;

[0004] In order to ensure the safety of our side, the aircraft needs to have a longer range, but a longer range often means a decrease in accuracy and a change in landing angle; a technical solution that increases the range without reducing the accuracy is an important goal pursued by researchers;

[0005] In addition, the weak link of whether the aircraft can hit the target is a key factor in causing greater damage. In the existing technology, the landing angle error can be constrained 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 present inventors have made in-depth research on the aircraft guidance control system, adopted a range control scheme in the mid-guidance section, designed a new detection method in the terminal guidance section, and adopted a new control method to improve the hitting accuracy and landing angle accuracy, thereby proposing a guidance control method and system for a long-range multi-mode compound guided aircraft. SUMMARY

[0007] In order to overcome the above problems, the inventors have made intensive research and designed a guidance control method and system for a remote multi-mode composite guided aircraft, in which an active radar / infrared multi-mode composite guidance strategy is used in the terminal guidance section, the advantages of both are complementary, the information acquisition capability of the detection system is enhanced, the detection capability of the aircraft under the influence of various interference factors is still ensured, the method of using inertial guidance for the aircraft in the middle guidance section and controlling it to increase the range of gliding through the over-weight compensation proportional guidance law, realizes the improvement of the accurate hitting range of the aircraft from 20km to 40km, and the new terminal sliding mode guidance law is used in the terminal guidance section to control the terminal guidance section of the aircraft, on the basis of ensuring accurate hitting, the landing angle error is controlled to be within 2° from 10°; thereby completing the present application.

[0008] Specifically, the present application aims to provide a guidance control method for a remote multi-mode composite guided aircraft, which comprises

[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 to the central processing module as input quantities,

[0010] The position, relative speed and line-of-sight angle of the target are measured by the active radar guidance module, and the measured information is transmitted to the central processing module as input quantities,

[0011] The line-of-sight angle of the missile is measured by the infrared guidance module, and is transmitted to the central processing module as an input quantity,

[0012] The central processing module obtains the required overload of the aircraft in real time based on the received information, and decomposes the required overload according to the roll angle information to generate rudder deflection instructions, and transmits the finally obtained rudder deflection instructions to the actuator to control the aircraft to fly towards the target.

[0013] Wherein, when the aircraft is in the middle guidance section, the central processing module obtains the required overload a m ,

[0014] When the aircraft is in the terminal guidance section, the central processing module obtains the required overload a m .

[0015] The over-weight compensation proportional guidance law obtains the required overload of the aircraft by the following formula (I):

[0016]

[0017] Wherein, N represents the proportional guidance coefficient,

[0018] represents the relative velocity of the projectile and the target,

[0019] represents the line-of-sight angular rate of the projectile and the target,

[0020] c represents a gravity compensation term,

[0021] g represents the gravity acceleration.

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

[0023]

[0024] wherein f1(x) represents a basic overload function,

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

[0026] μ represents an overload coefficient,

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

[0028] s represents a sliding mode variable,

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

[0030] dt represents a time integration step.

[0031] The application also provides a guidance control system for a remote multi-mode composite guided aircraft, which comprises 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 the position information are transmitted to the central processing module as input quantities,

[0033] The position of the target, the speed of the target and the line-of-sight angle of the projectile are measured by the active radar guidance module, and the measured information is transmitted to the central processing module as input quantities,

[0034] The line-of-sight angle of the projectile is measured by the infrared guidance module, and is transmitted to the central processing module as an input quantity,

[0035] The required overload of the aircraft is obtained by the central processing module, and the required overload is decomposed according to the roll angle information to generate a rudder deflection command, and the final rudder deflection command is transmitted to an executing mechanism to control the aircraft to fly towards the target.

[0036] The application has the beneficial effects including:

[0037] (1), the guidance control method for the remote multi-mode composite guided aircraft according to the present application, the central processing module adopts inertial guidance for the aircraft in the middle guidance stage and controls it to carry out range-increasing glide through the over-weight compensation proportional guidance law, which realizes the promotion of the precise hitting range of the aircraft from 20 km to 40 km;

[0038] (2), the guidance control method for the remote multi-mode composite guided aircraft according to the present application, the central processing module adopts a new type of terminal sliding mode guidance law in the terminal guidance stage, which enhances the anti-interference ability of the aircraft, has stronger robustness, makes the process of the aircraft reaching the expected impact angle more rapid and stable, and realizes the control of the impact angle error from 10° to within 2° on the basis of ensuring the precise hitting.

[0039] (3), the guidance control method for the remote multi-mode composite guided aircraft according to the present application, the central processing module adopts a new type of terminal sliding mode surface, which makes the convergence speed of the sliding mode control faster, but at the same time has the same robustness as the terminal sliding mode surface.

[0040] (4), the guidance control method for the remote multi-mode composite guided aircraft according to the present application, the convex combination fusion algorithm is used to call the angle information obtained by the active radar guidance module and the infrared guidance module, the relative distance between the missile and the target provided by the active radar guidance module is combined, the line-of-sight angle of the missile and the target is obtained and transmitted to the central processing module, and then the required overload of the aircraft is obtained, which realizes the purpose of obtaining target information with higher precision and better stability; if one of the active radar guidance module and the infrared guidance module fails, the other module alone provides information. BRIEF DESCRIPTION OF DRAWINGS

[0041] Figure 1 The overall logical structure schematic diagram of the guidance control method for the remote multi-mode composite guided aircraft according to the present application is shown.

[0042] Figure 2 The trajectory schematic diagram of the aircraft in the embodiment is shown.

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

[0044] Figure 4 The trajectory schematic diagram of the aircraft in the embodiment and the comparative example 2 is shown. DETAILED DESCRIPTION

[0045] The present application will be further described in detail by the drawings and embodiments. Through these descriptions, the characteristics and advantages of the present application will become more clear and explicit.

[0046] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. While the implementations are illustrated with reference to various figures, the figures are not necessarily drawn to scale.

[0047] According to the present application, a guidance control method for a remote multi-mode composite guided vehicle is provided, in which,

[0048] The roll angle and position of the vehicle are measured by an inertial navigation module, and the roll angle information and position information are transmitted to the central processing module as input quantities. The inertial navigation module includes a mechanical gyroscope which can sensitively measure the three-axis angular rate of the projectile body immediately after being enabled without consuming time for determination of zero reference; the inertial navigation module also includes three-axis MEMS gyroscope, accelerometer and other INS inertial elements, which need to be aligned in zero position after being enabled, and the attitude information measurement needs to be converged to the true value after a certain time, i.e. the inertial navigation module is started to align in advance before the vehicle enters the mid-guidance stage.

[0049] The position information (x t ,y t ) of the target is preloaded for the vehicle in advance before the vehicle is launched; the vehicle is in an uncontrolled climbing state for a period of time after being launched, and climbs by relying on the lift of the vehicle itself. The spatial orientation gyroscope is unlocked when the programmed time is reached; the rudder is opened when the control starting time is reached, the guidance module on the vehicle is started, and the vehicle enters the mid-guidance stage; the mid-guidance stage of the vehicle is also called the range-increasing gliding stage.

[0050] The inertial navigation module is aligned in zero position after being enabled for a period of time, and the attitude information measurement needs to be converged to the true value after a certain time, so as to obtain the position information (x m ,y m ) of the vehicle; the central processing module calculates the relative distance R between the projectile and the target and the line-of-sight angle q between the projectile and the target according to the position information of the vehicle and the target;

[0051]

[0052]

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

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

[0055] The active radar guidance module is composed of an antenna system, a receiver, a signal processing system, a servo mechanism, a transmitter and the like. The microwave / millimeter wave generated by the missile-borne transmitter is radiated out through the seeker antenna, and the echo reflected by the target is received by the same antenna, thus completing the search phase of the target. The process from searching for the target to switching into 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 highly directional seeker antenna is driven by the data processor and the servo mechanism, and the angle tracking of the target is realized.

[0056] In the method, the line-of-sight angle of the missile-target is measured by the infrared guidance module and is sent to the central processing module as an input quantity.

[0057] The infrared guidance module is composed of an optical system, a modulator, an infrared detector, a refrigerator, a servo mechanism and electronic circuits. The optical system receives the infrared radiation of the target, processes it into an optical signal containing target information through the modulator, converts the optical signal into an electrical signal easy to process through the infrared detector, and then performs filtering, amplification and processing on the signal through the electronic circuits to detect the target angular position information and send the information to the servo mechanism to make the optical axis move towards the target direction, thus realizing the continuous tracking of the target by the guidance system.

[0058] In the method, the required overload of the aircraft is obtained in real time by the central processing module according to the received information, the required overload is decomposed according to the roll angle information to generate a rudder deflection command, and the final rudder deflection command is transmitted to the execution mechanism to control the aircraft to fly towards the target. The execution mechanism is a rudder servo mechanism.

[0059] In a preferred embodiment, when the aircraft is in the mid-guidance phase, the central processing module obtains the required overload a m ,

[0060] When the aircraft is in the terminal guidance phase, the central processing module obtains the required overload a m .

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

[0062] In a preferred embodiment, the required overload of the aircraft is obtained by the over-weight compensation proportional guidance law through the following formula (I):

[0063]

[0064] wherein N represents a proportional guidance coefficient, generally taken as 2-6 in engineering, preferably taken as 4;

[0065] represents a relative velocity of the missile and the target, the relative distance between the missile and the target can be obtained by processing the active radar guidance module by using the convex combination fusion algorithm, and the relative distance between the missile and the target can also be obtained by the infrared guidance module, and the relative velocity of the missile and the target can be obtained by differentiating the relative distance between the missile and the target;

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

[0067] c represents a gravity compensation coefficient, taken as 1-1.5;

[0068] g represents a gravity acceleration, taken as 9.81.

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

[0070]

[0071] wherein f1(x) represents a basic overload function,

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

[0073] μ represents an overload coefficient,

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

[0075] s represents a sliding mode variable,

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

[0077] dt represents a time integration step.

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

[0079]

[0080] wherein, represents a relative velocity of the missile and the target,

[0081] R represents a relative distance between the missile and the target,

[0082] represents an expected line-of-sight angular rate, which is generally a fixed value, preferably taken as 0,

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

[0084] represents the line-of-sight angular rate of the missile.

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

[0086]

[0087] wherein R represents the relative distance of the missile,

[0088] represents the expected line-of-sight angular rate,

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

[0090] q d represents the expected impact angle of the aircraft,

[0091] x1 represents a state variable, the state variable x1 is selected as x1=q-q d ,

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

[0093] The line-of-sight angle q of the missile in the terminal guidance stage is obtained by processing the line-of-sight angle of the missile obtained by the active radar guidance module and the line-of-sight angle of the missile obtained by the infrared guidance module by using a convex combination fusion algorithm; the convex combination fusion algorithm is described in Liu Weidong, Liu Yang, Gao Lie'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 (five):

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

[0096] wherein g(x1) is obtained by the following formula (six):

[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, the design parameters are specifically taken as k0=1, k1=2, k2=20, and α / β=3 / 5.

[0100] In the mid-guidance stage and the terminal guidance stage, the central processing module decomposes the calculated required overload according to the roll angle given by the mechanical gyro of the inertial navigation module to obtain the rudder deflection command, and transmits the finally obtained rudder deflection command to the actuator to control the attitude of the aircraft, and finally controls the aircraft to the desired impact angle q d Precise hit target.

[0101] The application also provides a guidance control system for a remote multi-mode composite guided aircraft, which comprises a guidance control system, a central processing module and a rudder servo mechanism; the guidance 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 to the central processing module as input quantities,

[0103] The position of the target, the speed of the target and the line-of-sight angle of the missile-target are measured by the active radar guidance module, and the measured information is transmitted to the central processing module as input quantities,

[0104] The line-of-sight angle of the missile-target is measured by the infrared guidance module, and is transmitted to the central processing module as an input quantity,

[0105] The required overload of the aircraft is obtained by the central processing module, and the required overload is decomposed according to the roll angle information to generate a rudder deflection command, and the finally obtained rudder deflection command is transmitted to the actuator to control the aircraft to fly towards the target;

[0106] When the aircraft is in the mid-guidance stage, the central processing module obtains the required overload a m ,

[0107] When the aircraft is in the terminal guidance stage, the central processing module obtains the required overload a m ;

[0108] The required overload of the aircraft is obtained by the over-weight compensation proportional guidance law through the following formula (I):

[0109]

[0110] Wherein, N represents the proportional guidance coefficient,

[0111] represents the relative motion speed of the missile-target,

[0112] represents the line-of-sight angle rate of the missile-target,

[0113] c represents the gravity compensation term,

[0114] g represents the acceleration of gravity;

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

[0116]

[0117] wherein f1(x) represents a basic overload function,

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

[0119] μ represents an overload coefficient,

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

[0121] s represents a sliding mode variable,

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

[0123] dt represents a time integral step.

[0124] The system further comprises a power master control device, which is connected with a thermal power source on the aircraft and is responsible for power supply of the entire system, so that each module can normally work under rated power according to the required voltage.

[0125] Embodiment 1

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

[0127] The speed of the aircraft is V m0 =600m / s;

[0128] The initial trajectory inclination angle of the aircraft is ;

[0129] The initial line-of-sight angle between the aircraft and the target is q0=0°.

[0130] The target is set as a fixed building on the ground.

[0131] After the aircraft enters the midcourse guidance phase, the roll angle of the aircraft, the position of the aircraft and the line-of-sight angle between the aircraft and the target are simulated in real time by the simulation system, and the relative motion speed between the missile and the target, i.e., the speed of the aircraft, is obtained based on the speed of the position change of the aircraft; the central processing module obtains the required overload of the aircraft through the following formula (one):

[0132]

[0133] According to the roll angle information, the required overload is decomposed to generate rudder deflection instructions, and the final rudder deflection instructions are transmitted to the actuator to control the aircraft to fly to the target.

[0134] The specific values in the formula (I) are N=4, c=1.5, and g=9.8.

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

[0136]

[0137] 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 (V):

[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 stage are k0=1, k1=2, k2=20, and α / β=3 / 5.

[0147] The expected landing angle q d is 90 degrees.

[0148] According to the roll angle information, the required overload is decomposed to generate rudder deflection instructions, and the final rudder deflection instructions are transmitted to the actuator to control the aircraft to fly to the target, and the final flight trajectory is shown by the solid trajectory in Figure 2 、 Figure 3 and Figure 4 .

[0149] Example 2

[0150] An example is set up basically the same as in Example 1, with the only difference being that the expected landing angle qd is 80 degrees; the final flight trajectory is shown in FIG. 8. Figure 2

[0151] Example 3

[0152] An example similar to Example 1 is set up, with the only difference being that the desired impact angle q is 70 degrees. d is 70 degrees; the final flight trajectory is shown in FIG. 9. Figure 2

[0153] Example 4

[0154] An example similar to Example 1 is set up, with the only difference being that the desired impact angle q is 60 degrees. d is 60 degrees; the final flight trajectory is shown in FIG. 10. Figure 2

[0155] Comparative Example 1

[0156] An experimental process similar to Example 1 is set up, with the only difference being that, in the midcourse guidance stage, proportional guidance is used for guidance, i.e.

[0157] The final flight trajectory is shown in the dashed curve of FIG. 11. Figure 3

[0158] Comparative Example 2

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

[0160] The final flight trajectory is shown in FIG. 12. Figure 4

[0161] From the above Examples 1 to 4 and Comparative Example 1, it can be seen that the guidance control system and method for a remote multi-mode composite guided aircraft provided by the present application can control the aircraft to increase the precise hit range of the aircraft from 20 km to 35 km;

[0162] ​​​​​From the above embodiments 1 to 4 and the comparative example 2, it can be seen that the guidance control system and method for remote multi-mode composite guided aircraft provided by the application improves the robustness of the control process, and at the same time enhances the anti-interference ability of the aircraft, so that the process of the aircraft reaching the expected landing angle is faster and more stable. The method adopted by the patent can make the aircraft realize the progress of the landing angle error from 10° to 2° on the basis of ensuring accurate hitting.

[0163] The above describes the application in combination with preferred embodiments, but these embodiments are only exemplary and serve only to illustrate. On this basis, various substitutions and improvements can be made to the application, and these all fall within the protection scope of the application.

Claims

1. A guidance control method applied to a remote multi-mode composite guided vehicle, characterized in that, In the method, the roll angle and the position of the aircraft are measured by the inertial navigation module, and the roll angle information and the position information are transmitted to the central processing module as input quantities, the position of the target, the relative velocity of the missile-target, and the line-of-sight angle of the missile-target are measured by the active radar guidance module, and the measured information is transmitted to the central processing module as input quantities, the line-of-sight angle of the missile-target is measured by the infrared guidance module, and is transmitted to the central processing module as an input quantity, the required overload of the aircraft is obtained in real time by the central processing module based on the received information, and the required overload is decomposed according to the roll angle information to generate a rudder deflection command, and the final rudder deflection command is transmitted to the actuator to control the aircraft to fly towards the target; The central processing module obtains the required overloads a of the aircraft by means of an over-weighted proportional guidance law when the aircraft is in the mid-guidance phase m , The central processing module obtains the required overload a of the aircraft by a new terminal sliding mode guidance law when the aircraft is in the terminal guidance phase m ; the required overload of the aircraft is obtained by the over-recovery proportional guidance guidance law through the following formula (I): wherein N represents a proportional guidance coefficient, Indicates the relative velocity between the projectile and the target. surface the line-of-sight angle rate of the missile-target, c represents a gravity compensation term, g represents the acceleration of gravity; the required overload of the aircraft is obtained by the new terminal sliding mode guidance law through the following formula (II): wherein f1(x) represents a basic overload function, f2(x) represents an overload gain function, μ represents an overload coefficient, g(x1) represents a basic sliding mode function, s represents a sliding mode variable, d(g(x1)) represents the time rate of change of g(x1), dt represents the time integration step.

2. The guidance control method for a remote multi-mode composite guided aircraft according to claim 1, wherein the expression of f1(x) is shown in the following formula (III): wherein, represents the velocity of the relative motion between the projectile and the target, R represents the relative distance of the missile-target, Indicates the desired line-of-sight angular rate. x2represents a state variable, the state variable x2being selected as represents the ballistic line-of-sight angular rate.

3. The guidance control method for a remote multi-mode composite guided aircraft according to claim 1, wherein the expression of f2(x) is shown in the following formula (IV): wherein R represents the relative distance of the missile-target, Indicates the desired line-of-sight angular rate. θ represents the trajectory inclination angle of the aircraft, q d denotes the desired angle of landing of the aircraft, x1 represents a state variable, the state variable x1 being chosen as x1 = q - q d , q represents the line-of-sight angle of the missile-target.

4. The guidance control method for a remote multi-mode composite guided aircraft according to claim 1, wherein the sliding mode variable s is obtained through the following formula (V): s=g(x1)+x2(V) wherein g(x1) is obtained through the following formula (VI): k0, k1, k2, α, β each independently represent a design parameter.

5. The guidance control method for a remote multi-mode composite guided aircraft according to claim 1, wherein when the relative distance of the missile-target R is less than a set distance, the aircraft enters the terminal guidance phase; the set distance is 3-8 km.

6. The guidance control method for a remote multi-mode composite guided aircraft according to claim 1, wherein in the central processing module, the line-of-sight angle of the missile-target obtained by the active radar guidance module and the line-of-sight angle of the missile-target obtained by the infrared guidance module are processed by a convex combination fusion algorithm to obtain a final line-of-sight angle of the missile-target, and the required overload is obtained based on the final line-of-sight angle of the missile-target.

7. A guidance control system for a remote multi-mode composite guided vehicle, characterized by, The system comprises an inertial navigation module, an active radar guidance module, an infrared guidance module, and a central processing module, the roll angle and the position of the aircraft are measured by the inertial navigation module, and the roll angle information and the position information are transmitted to the central processing module as input quantities, The position of the target, the speed of the target and the line-of-sight angle of the missile and the target are measured by the active radar guidance module, and the measured information is transmitted to the central processing module as input, The line-of-sight angle of the missile and the target is measured by the infrared guidance module, and is transmitted to the central processing module as input, The required overload of the aircraft is obtained by the central processing module, and the required overload is decomposed according to the roll angle information to generate a rudder deflection command, and the finally obtained rudder deflection command is transmitted to an executing mechanism to control the aircraft to fly towards the target; wherein, when the aircraft is in the mid-guidance phase, the central processing module obtains the required overloading a of the aircraft through the overloading compensation proportional guidance law m , The central processing module obtains the required overload a of the aircraft by a new terminal sliding mode guidance law when the aircraft is in the terminal guidance phase m ; The required overload of the aircraft is obtained by the overloading compensation proportional guidance law through the following formula (I): Wherein, N represents a proportional guidance coefficient, Indicates the relative velocity between the projectile and the target. surface The line-of-sight angle rate of the missile and the target is shown, c represents a gravity compensation term, g represents a gravity acceleration; The required overload of the aircraft is obtained by the new terminal sliding mode guidance law through the following formula (II): Wherein, f1(x) represents a basic overload function, f2(x) represents an overload gain function, μ represents an overload coefficient, g(x1) represents a basic sliding mode function, s represents a sliding mode variable, d(g(x1)) represents the time rate of change of g(x1), dt represents a time integral step length.

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