High Dynamic Aircraft Guidance and Control Method Based on Fusion Differential Method

Abstract

This invention discloses a guidance and control method for high-dynamic aircraft based on a fusion differential method. To improve the system's response speed and continuous accuracy, this method designs a fusion differential algorithm that includes nonlinear and linear terms to ensure rapid response in the initial response stage and smooth tracking in the subsequent tracking stage during line-of-sight angular velocity estimation. This allows for timely and accurate acquisition of the line-of-sight angular velocity required for guidance and control. Furthermore, based on a gravity-compensated proportional guidance method, the method designs the required overload for the high-dynamic aircraft. The required overload derives control commands according to the transfer ratio of different aircraft and transmits them to the servo mechanism. The servo mechanism then controls the servo motor to control the aircraft's flight, achieving higher-precision guidance and control.

Description

Technical Field

[0001] This invention relates to guidance and control of highly dynamic aircraft, particularly guidance and control of highly dynamic aircraft carrying strapdown seekers, and specifically to a guidance and control method for highly dynamic aircraft based on a fusion differential method. Background Technology

[0002] High-dynamic guided aircraft often employ strapdown seekers to acquire targets due to their high overload characteristics, cost requirements, and limitations in guidance systems. However, because strapdown seekers are fixed to the aircraft, they cannot directly obtain the line-of-sight angular velocity of the aircraft relative to the target and apply it to the guidance and control system, unlike platform seekers. Instead, they need to process the angle signal using certain algorithms to estimate the line-of-sight angular velocity.

[0003] There will inevitably be a discrepancy in timing and frequency between the angular velocity estimated by the algorithm system and the actual angular velocity. This difference determines the effectiveness of guidance; better tracking results in more precise guidance and control. However, the fast response characteristics and tracking accuracy of existing angular velocity acquisition methods still need further improvement. A more superior differential method needs to be designed to obtain the line-of-sight angular velocity for better guidance and control of highly dynamic guided aircraft.

[0004] For the reasons mentioned above, the inventors have conducted in-depth research on the method for obtaining the line-of-sight angular velocity of a high-dynamic strapdown seeker aircraft, and based on this, have carried out guidance and control of the aircraft, in order to design a high-dynamic aircraft guidance and control method based on the fusion differential method that can solve the above problems. Summary of the Invention

[0005] To overcome the aforementioned problems, the inventors conducted intensive research and designed a high-dynamic aircraft guidance and control method based on a fusion differential method. This method, to improve system response speed and continuous accuracy, incorporates a fusion differential algorithm including nonlinear and linear terms to ensure rapid response in the initial response phase and smooth tracking in the subsequent tracking phase during line-of-sight angular velocity estimation. This allows for timely and accurate acquisition of the line-of-sight angular velocity required for guidance and control. Furthermore, based on a gravity-compensated proportional guidance system, the required overload is designed for the high-dynamic aircraft. The required overload derives control commands according to the transfer ratio of different aircraft and transmits them to the servo mechanism. The servo mechanism then controls the servo motor to control the aircraft's flight, achieving higher-precision guidance and control; thus, this invention is complete.

[0006] Specifically, the purpose of this invention is to provide a high-dynamic aircraft guidance and control method based on the fusion differential method. In this guidance and control method, the required overload of the aircraft is obtained in real time. The required overload is used to derive control commands according to the transfer ratio of the aircraft and transmit them to the servo mechanism. The servo mechanism controls the servo motor to control the aircraft, thereby enabling the aircraft to hit the target.

[0007] Among them, a strapdown seeker is provided on the aircraft, and the target is captured through the strapdown seeker, and the line-of-sight angle between the missile and the target is obtained in real time.

[0008] Among them, the required overload is obtained through the following formula (1):

[0009] a = NV(t)z2(t) + K4g (1)

[0010] Among them, a represents the required overload, [[ID=,12]]

[0011] N represents the proportional guidance coefficient,

[0012] V(t) represents the speed of the aircraft,

[0013] z2(t) represents the line-of-sight angular velocity between the missile and the target,

[0014] K4 represents the gravity compensation coefficient,

[0015] g represents the gravity coefficient.

[0016] Among them, at the initial moment, the value of the line-of-sight angular velocity z2(t) between the missile and the target is 0, and at subsequent moments, the line-of-sight angular velocity z2(t) between the missile and the target is obtained by integrating the line-of-sight angular acceleration between the missile and the target ;

[0017] Preferably, the line-of-sight angular acceleration between the missile and the target is obtained through the following formula (2):

[0018]

[0019] Among them, K1, K2, and K3 each represent design parameters;

[0020] z′2(t) represents the estimated value of the line-of-sight angular velocity between the missile and the target;

[0021] e(t) represents the difference between the integral value of the estimated value of the line-of-sight angular velocity between the missile and the target and the line-of-sight angle between the missile and the target obtained by the strapdown seeker.

[0022] Among them, the e(t) is obtained through the following formula (3):

[0023] <00,00054>

[0024] Among them, q(t) represents the line-of-sight angle between the missile and the target obtained by the strapdown seeker;

[0025] is equal to z′2(t), and both represent the estimated value of the line-of-sight angular velocity between the missile and the target;

[0026] z1(t) represents the integral value of the estimated value of the line-of-sight angular velocity between the missile and the target.

[0027] ]]

[0027] The estimated value of the projectile-eye line-of-sight angular velocity z′2(t) is the same as the projectile-eye line-of-sight angular velocity z2(t) at the previous moment.

[0028] The beneficial effects of this invention include:

[0029] The high-dynamic aircraft guidance and control method based on the fusion differential method provided by the present invention utilizes the missile-target line-of-sight angle obtained by the strapdown seeker and obtains a more accurate line-of-sight angular velocity in real time through fusion differential method, providing a data basis for the precise control of high-dynamic aircraft and enabling the aircraft to have higher guidance accuracy. Attached Figure Description

[0030] Figure 1 This illustration shows a comparison between the real-time line-of-sight angular velocity of the missile and the actual line-of-sight angular velocity obtained by the aircraft in an embodiment of this application.

[0031] Figure 2 This diagram illustrates a comparison between the real-time line-of-sight angular velocity of the missile and the actual line-of-sight angular velocity obtained by the second aircraft in this embodiment of the application.

[0032] Figure 3 The diagram shows the flight trajectories of aircraft one and aircraft two in the embodiments of this application. Detailed Implementation

[0033] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present invention will become clearer and more apparent.

[0034] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0035] According to the high-dynamic aircraft guidance and control method based on the fusion differential method provided by the present invention, the required overload of the aircraft is obtained in real time. The required overload is used to derive control commands according to the transfer ratio of the aircraft, which are then transmitted to the servo mechanism. The servo mechanism controls the servo motors to control the aircraft, thereby enabling the aircraft to hit the target. Preferably, different aircraft have different corresponding transfer ratios, and matching is selected according to the specific characteristics of the aircraft model. Preferably, the servo mechanism includes servo motors on the aircraft.

[0036] Among them, a strapdown seeker is provided on the aircraft. The target is captured through the strapdown seeker, and the line-of-sight angle between the missile and the target is obtained in real time. After the strapdown seeker captures the target, the aircraft enters the terminal guidance section. The strapdown seeker transmits the obtained line-of-sight angle information between the missile and the target to the microprocessor module on the aircraft in real time, and the microprocessor module is used to perform the conversion processing from the line-of-sight angle between the missile and the target to the line-of-sight angular velocity between the missile and the target, and further obtain the required overload for controlling the aircraft.

[0037] In a preferred embodiment, the required overload is obtained by the following formula (1):

[0038] a = NV(t)z2(t) + K4g (1)

[0039] Among them, a represents the required overload,

[0040] N represents the proportional guidance coefficient, and its value ranges from 4 to 6, preferably 4;

[0041] V(t) represents the aircraft speed, which is obtained by real-time measurement of the inertial system on the aircraft;

[0042] z2(t) represents the line-of-sight angular velocity between the missile and the target,

[0043] K4 represents the gravity compensation coefficient, preferably 1.1;

[0044] g represents the gravity coefficient, and its value is 9.8.

[0045] In this application, the gravity compensation proportional guidance method is used to design the overload for the high-dynamic aircraft, which can compensate for the gravity effect while optimizing the guidance, and further improve the stability and hitting accuracy of the aircraft.

[0046] In a preferred embodiment, at the initial moment, the value of the line-of-sight angular velocity z2(t) between the missile and the target is 0, and the initial moment refers to the moment when the strapdown seeker starts to work;

[0047] At subsequent moments, the line-of-sight angular velocity z2(t) between the missile and the target is obtained by integrating the line-of-sight angular acceleration between the missile and the target The strapdown seeker in this application obtains the line-of-sight angle information between the missile and the target at a predetermined frequency, and the time interval between the acquisition times of two adjacent line-of-sight angle information between the missile and the target is consistent with the time interval between the adjacent moments;

[0048] Preferably, each obtained line-of-sight angle information between the missile and the target corresponds to a line-of-sight angular velocity between the missile and the target.

[0049] Preferably, the line-of-sight angular acceleration between the missile and the target is obtained by the following formula (2):

[0050] <

[0051] Among them, K1, K2, and K3 each represent design parameters; preferably, their values are: K1 = -150, K2 = -5 × 10 , , , , , , ,

[0064] ,

[0063] ,

[0062] ,

[0061] ,

[0060] ,

[0066] ,

[0065] , K3 = -250;

[0052] z′2(t) represents the estimated value of the missile - target line - of - sight angular velocity;

[0053] e(t) represents the difference between the integral value of the estimated missile - target line - of - sight angular velocity and the missile - target line - of - sight angle obtained by the strapdown seeker. <000​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​​The difference is that in the first aircraft, the high-dynamic aircraft guidance and control method based on the fusion differential method provided in this application is adopted, in which the initial value of the missile-target line-of-sight angular velocity z2(t) is 0.

[0067] At subsequent moments, the aforementioned projectile-eye line-of-sight angular velocity z2(t) is obtained by adjusting the projectile-eye line-of-sight angular acceleration. Calculate the points to obtain;

[0068] The acceleration of the line of sight of the bullet. Obtained through the following formula (ii);

[0069]

[0070] The design parameters are: K1 = -150, K2 = -5 × 10⁻⁶. 4 K3 = -250;

[0071] The e(t) is obtained by the following equation (iii):

[0072]

[0073] The estimated value of the line-of-sight angular velocity z′2(t) is related to the line-of-sight angular acceleration at the previous moment. The values ​​of the angular velocity z2(t) obtained by integration are the same.

[0074] In the second aircraft, the angular velocity of the missile-eye line of sight is obtained in real time by the sliding mode differentiation method; the sliding mode differentiation method is described in Wang Jianhao, Hu Jianbo, Zhang Liang, Zhang Pengtao, Song Min. Inversion terminal sliding mode flight control based on sliding mode disturbance observer [J]. Systems Engineering and Electronics, 2018, 40(06):1345-1350.

[0075] During the aircraft's flight, the simulation system provides the actual and accurate angular velocity of the missile's line of sight in real time, such as... Figure 1 and Figure 2 As shown by the solid lines in the text;

[0076] Correspondingly, the angular velocity of the missile-target line of sight obtained by the aircraft in real time is as follows: Figure 1 As shown by the dashed line in the image;

[0077] The real-time angular velocity of the missile's line of sight obtained by the aircraft is as follows: Figure 2 The dotted lines in the text are shown.

[0078] pass Figure 1 and Figure 2The comparison shows that the line-of-sight angular velocity obtained by the high-dynamic aircraft guidance and control method based on the fusion differential method provided in this application is closer to the actual line-of-sight angular velocity, and the initial response speed and tracking stability are better than the line-of-sight angular velocity obtained by the sliding mode differential method.

[0079] Furthermore, the overall flight trajectories of spacecraft one and spacecraft two are as follows: Figure 3 As shown; Figure 3 The solid line in the diagram represents the flight path of spacecraft one, and the dashed line represents the flight path of spacecraft two. Figure 3 As can be seen from this, the high-dynamic aircraft guidance and control method based on the fusion differential method provided in this application can enable the aircraft to have a more ideal ballistic trajectory when the target distance is 5km, and the final impact point of the aircraft is more effective in resolving the target.

[0080] The present invention has been described above with reference to preferred embodiments; however, these embodiments are merely exemplary and illustrative. Various substitutions and modifications can be made to the present invention based on these embodiments, all of which fall within the scope of protection of the present invention.

Claims

1. A high-dynamic aircraft guidance and control method based on a fusion differential method, characterized in that in this guidance and control method, the required overload of the aircraft is obtained in real time. The required overload derives a control command according to the transfer ratio of the aircraft and transmits it to the servo mechanism. The servo mechanism controls the actuator to control the aircraft, so that the aircraft hits the target; wherein, a strapdown seeker is provided on the aircraft to capture the target through the strapdown seeker and obtain the missile-to-target line-of-sight angle in real time; the required overload is obtained by the following formula (1): a = NV(t)z2(t) + K4g (1) where, a represents the required overload, N represents the proportional guidance coefficient, V(t) represents the aircraft speed, z2(t) represents the missile-to-target line-of-sight angular velocity, K4 represents the gravity compensation coefficient, g represents the gravity coefficient; at the initial moment, the value of the missile-to-target line-of-sight angular velocity z2(t) is 0, At subsequent moments, the aforementioned projectile-eye line-of-sight angular velocity z2(t) is obtained by adjusting the projectile-eye line-of-sight angular acceleration. Calculate the points to obtain; The acceleration of the line of sight of the bullet. We obtain it through the following formula (ii): where, K1, K2, and K3 each represent design parameters; z′2(t) represents the estimated value of the missile-to-target line-of-sight angular velocity; e(t) represents the difference between the integral value of the estimated value of the missile-to-target line-of-sight angular velocity and the missile-to-target line-of-sight angle obtained by the strapdown seeker.

2. The high-dynamic aircraft guidance and control method based on the fusion differential method according to claim 1, characterized in that the e(t) is obtained by the following formula (3): where, q(t) represents the missile-to-target line-of-sight angle obtained by the strapdown seeker; Equal to z′2(t), both represent estimated values ​​of the angular velocity of the projectile's line of sight; z1(t) represents the integral value of the estimated value of the missile-to-target line-of-sight angular velocity.

3. The high-dynamic aircraft guidance and control method based on the fusion differential method according to claim 1, characterized in that the value of the estimated value of the missile-to-target line-of-sight angular velocity z′2(t) is the same as the value of the missile-to-target line-of-sight angular velocity z2(t) at the previous moment.

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