Terminal guidance control method based on all-wheel-drive system

Abstract

The invention discloses a terminal guidance control method based on an all-wheel-drive system, and the method comprises the steps: obtaining an expected acceleration of an aircraft based on a satellite signal and an attitude sensitive system at a mid-guidance section, controlling the aircraft to fly to a target, and in the process, if the aircraft is rejected by a satellite, controlling the aircraft to fly to the target; if yes, the expected acceleration of the aircraft is obtained based on the satellite signal applied at the last moment until the real-time satellite signal is obtained again, and the laser seeker is started after the aircraft is launched for a preset time; after a laser seeker captures a target, the expected acceleration of the aircraft is obtained in real time through a novel sight angle constraint guidance law, a rudder instruction is generated based on the expected acceleration of the aircraft, a steering engine is controlled to conduct rudder work, the aircraft is controlled to fly to the target, and the target is collided at an expected sight angle. The deviation caused by satellite denial is compensated and corrected through the novel sight angle constraint guidance law, and finally the aircraft hits the target.

Description

Technical Field

[0001] The invention relates to an aircraft guidance control method, and in particular to a terminal guidance control method based on an all-wheel drive system. Background Art

[0002] High-speed flying targets have high flying speeds, strong maneuverability, wide flight envelopes, complex flight trajectories, and large instantaneous overloads. In addition, they usually require kinetic damage by direct collision with the target. In order to effectively intercept such flying targets, higher requirements are placed on the guided flight control capabilities of the intercepting aircraft. In order to maximize the damage effect of the warhead, the guided aircraft needs to be able to hit the target at a certain interception angle. Therefore, in the interception mission of highly maneuverable targets, while ensuring the guidance accuracy, considering the interception angle constraint has also become an important requirement in the actual application process.

[0003] Based on the above problems, the inventors have conducted in-depth research on the terminal guidance control of the aircraft, and specifically designed a terminal guidance control method based on an all-wheel drive system in order to solve the above problems accordingly. Summary of the invention

[0004] In order to overcome the above-mentioned problems, the inventors have conducted intensive research and designed a terminal guidance control method based on an all-wheel drive system. In this method, the large maneuverability characteristics of the target are taken into account, a two-dimensional projectile-target relative motion model is established on the pitch channel, and combined with a high-order all-wheel drive system method, a second-order strict feedback guidance and control integrated nonlinear model is established, so as to obtain the system control input pitch rudder deflection angle in real time, and on this basis, the aircraft is controlled to intercept and hit the target at the desired interception angle, thereby completing the present invention.

[0005] Specifically, the purpose of the present invention is to provide a terminal guidance control method based on an all-wheel drive system, the method comprising the following steps:

[0006] Step 1: The aircraft captures the target through the seeker in the terminal guidance phase;

[0007] Step 2, obtaining the position and speed information of the aircraft itself through satellite signals, and obtaining the pitch rudder deflection angle of the system control input in real time according to the preset expected interception angle;

[0008] Step 3, generating a rudder instruction based on the system control input pitch rudder deflection angle and the system control input yaw rudder deflection angle, controlling the servo to steer, controlling the aircraft to fly toward the target, and colliding with the target at a desired interception angle.

[0009] Among them, the system control input pitch rudder deflection angle is obtained in real time through the following formula (1):

[0010]

[0011] Where u represents the system control input pitch rudder deflection angle,

[0012] b2 represents the model coefficient,

[0013] represents the control coefficient matrix,

[0014] represents the control coefficient matrix,

[0015] s2 represents the dynamic surface,

[0016] represents the dynamic surface integral,

[0017] f2 represents the model coefficient,

[0018] represents the observed estimate of the disturbance d2.

[0019] The beneficial effects of the present invention include:

[0020] (1) According to the terminal guidance control method based on the all-wheel drive system provided by the present invention, the actual operation processing steps in the aircraft are simpler, the model recursion process is reduced, the adjustment parameters are fewer, and the calculation amount in the aircraft is reduced, which helps to reduce the cost of the aircraft and also helps to ensure the calculation efficiency on the aircraft;

[0021] (2) According to the terminal guidance control method based on the all-wheel drive system provided by the present invention, the method designs a control scheme that meets the interception angle constraint under the framework of the backstepping method, realizes the linearization of the nonlinear system, and meets the accuracy and real-time requirements required for intercepting high-dynamic aircraft;

[0022] (3) According to the terminal guidance control method based on the all-wheel drive system provided by the present invention, the expected interception angle is defined by the expected line of sight angle, and the convergence speed of constraining the aircraft to the expected interception angle is faster and the error is smaller, which has an anti-interference effect on external disturbances such as aerodynamic parameters. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 The overall logic diagram of the terminal guidance control method based on the all-wheel drive system of the present application is shown;

[0024] Figure 2 A schematic diagram of the relative motion trajectory curve of the projectile and the target in the embodiment is shown;

[0025] Figure 3 A schematic diagram of a curve showing a change in the sight angle of the projectile in the embodiment is shown;

[0026] Figure 4 A schematic diagram of a curve showing a change in the angle of attack of an aircraft in an embodiment is shown;

[0027] Figure 5 A schematic diagram of an aircraft overload variation curve in an embodiment is shown;

[0028] Figure 6 A schematic diagram of a curve showing a change in the rudder angle of an aircraft in an embodiment is shown. DETAILED DESCRIPTION

[0029] 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.

[0030] 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.

[0031] According to a terminal guidance control method based on an all-wheel drive system provided by the present invention, Figure 1 As shown in , the method comprises the following steps:

[0032] Step 1: The aircraft captures the target through the seeker in the terminal guidance phase.

[0033] The seeker may be a laser seeker, through which the distance between the aircraft and the target is obtained in real time, and the missile-target sight angle is also obtained in real time.

[0034] Step 2, obtain the position and speed information of the aircraft itself through satellite signals, and obtain the system control input pitch rudder deflection angle in real time according to the preset expected interception angle.

[0035] In this application, the expected interception angle is defined using the expected line of sight angle, and the constraint problem of the interception angle at the guidance terminal is converted into the constraint problem of the missile-target line of sight angle at the guidance terminal. In order to ensure that the aircraft can accurately intercept the target and achieve effective damage, the derivative of the missile-target line of sight angle needs to be constrained to 0.

[0036] In this application, the system control input pitch rudder deflection angle is obtained in real time through the following formula (1):

[0037]

[0038] Where u represents the system control input pitch rudder deflection angle,

[0039] b2 represents the model coefficient,

[0040] represents the control coefficient matrix,

[0041] represents the control coefficient matrix,

[0042] s2 represents dynamic surface 2,

[0043] represents the differential of the dynamic surface 2,

[0044] f2 represents the model coefficient,

[0045] represents the observed estimate of the disturbance d2.

[0046] In a preferred embodiment, the Obtained by the following formula (II):

[0047]

[0048] Among them, p2 represents the auxiliary variable of the observer of disturbance d2, and its value range is p2>0.

[0049] β2 represents a parameter of the reduced-order state observer, and its value range is β2>0. Preferably, the specific value of β2 is β2=20.

[0050] α is the angle of attack,

[0051] The angular rate of the angle of attack is obtained as follows: in, is the pitch angle, which is obtained by integrating the pitch angle rate measured by the inertial measurement unit on the aircraft; θ M is the ballistic inclination angle, calculated by the formula: V y represents the velocity component of the image-guided vehicle on the y-axis of the inertial coordinate system, V x It represents the velocity component of the image-guided aircraft on the x-axis of the inertial coordinate system. The two velocity components can be obtained by integrating the acceleration measured by the accelerometer on the aircraft. get.

[0052] In a preferred embodiment, the p2 is obtained by reacting the following formula (III): Find the integral to obtain:

[0053]

[0054] in, represents the derivative of the auxiliary variable of the observer with respect to the disturbance d2,

[0055] p′2 represents the p2 value obtained at the last moment, and its initial value is 0.

[0056] u′ represents the system control input pitch rudder angle obtained at the last moment, and its initial value is 0.

[0057] Preferably, b2 is obtained by the following formula (four):

[0058]

[0059] in,

[0060] ρ represents the atmospheric density,

[0061] V M Indicates the flight speed of the aircraft, obtained through satellite signals.

[0062] S represents the reference area of ​​the aircraft,

[0063] l represents the characteristic length of the aircraft,

[0064] represents the partial derivative of the pitch moment coefficient with respect to the pitch rudder deflection angle,

[0065] J z It represents the moment of inertia of the aircraft about the longitudinal axis of the projectile.

[0066] In a preferred embodiment, the f2 is obtained by the following formula (V):

[0067]

[0068] Among them, the V M Indicates the flight speed of the aircraft, obtained through satellite signals.

[0069] g represents the acceleration due to gravity, and its value is 9.8.

[0070] θ M It represents the ballistic inclination of the aircraft, which is obtained by the ballistic inclination calculation formula: V y represents the velocity component of the image-guided vehicle on the y-axis of the inertial coordinate system, V x It represents the velocity component of the image-guided aircraft on the x-axis of the inertial coordinate system. The two velocity components can be obtained by integrating the acceleration measured by the accelerometer on the aircraft.

[0071] α is the angle of attack,

[0072] represents the angular rate of the angle of attack and is obtained by:

[0073]

[0074] w z It represents the pitch angle rate, which is obtained by measuring the inertial measurement unit on the aircraft.

[0075] a1, a2, and a3 each independently represent a system parameter.

[0076] Preferably, the system parameters are obtained by the following formula (thirteen):

[0077]

[0078] in,

[0079] ρ represents the atmospheric density,

[0080] V M Indicates the flight speed of the aircraft.

[0081] S represents the reference area of ​​the aircraft,

[0082] l represents the characteristic length of the aircraft,

[0083] represents the partial derivative of the lift coefficient with respect to the angle of attack,

[0084] represents the pitch moment coefficient with respect to the angle of attack,

[0085] represents the pitch moment coefficient to the pitch angular rate,

[0086] J z It represents the moment of inertia of the aircraft about the longitudinal axis of the projectile.

[0087] In a preferred embodiment, the dynamic surface s2 is obtained by the following formula (VI):

[0088] s2=α-α d (six)

[0089] Where α represents the angle of attack,

[0090] α d Indicates virtual control instructions;

[0091] Said Obtained by differentiating s2.

[0092] Preferably, the virtual control instruction α d Obtained by the following formula (VII):

[0093]

[0094] Wherein, b1 represents the model coefficient, which is obtained by the following formula (VIII):

[0095]

[0096] r represents the relative distance between the projectile and the target, which can be obtained by where x T and T Indicates the target position coordinates, x M and M Indicates the position coordinates of the aircraft,

[0097] q represents the sight angle between the missile and the target, which is obtained through the seeker.

[0098] represents the control coefficient matrix,

[0099] represents the control coefficient matrix,

[0100] f1 model coefficient,

[0101] The observed estimate of the disturbance d1;

[0102] s1 represents dynamic surface 1,

[0103] represents the integral of dynamic surface one;

[0104] Preferably, the dynamic surface s1 is obtained by the following formula (IX):

[0105] s1=qq d (Nine).

[0106] Among them, qd represents the desired interception angle, which is filled in the aircraft before the aircraft is launched;

[0107] Said Obtained by differentiating s1.

[0108] In this application, the and Obtained by the following formula (fourteen):

[0109]

[0110] Said and The following formula (XV) is used to obtain

[0111]

[0112] That is, the system state feedback matrices A1 and A2 in this application are designed as:

[0113]

[0114]

[0115] Where j represents the imaginary part of the complex number,

[0116] ω represents the expected performance parameter of the system, and its value is preferably 10.

[0117] In a preferred embodiment, the Obtained by the following formula (10):

[0118]

[0119] Among them, p1 represents the auxiliary variable of the observer of disturbance d1, and its value range is p1>0.

[0120] β1 represents a parameter of the reduced-order state observer, and its value range is β1>0. Preferably, the specific value of β1 is β1=1.

[0121] q represents the sight angle between the missile and the target, which is obtained through the seeker on the aircraft.

[0122] Indicates the angular rate of the barrage's line of sight, obtained through the seeker on the aircraft.

[0123] In a preferred embodiment, the p1 is obtained by Find the integral to obtain:

[0124]

[0125] in, represents the derivative of the auxiliary variable of the observer with respect to the disturbance d1,

[0126] p′1 represents the p1 value obtained at the last moment, and its initial value is 0. In a preferred embodiment, the f1 is obtained by the following formula (XII):

[0127]

[0128] It indicates the rate of change of the relative distance between the missile and the target, which is measured by the seeker on the aircraft.

[0129] r is the distance between the aircraft and the target,

[0130] It represents the angular rate of the missile-target line of sight, which is measured by the seeker on the aircraft.

[0131] Step 3, generating a rudder instruction based on the system control input pitch rudder deflection angle and the system control input yaw rudder deflection angle, controlling the servo to steer, controlling the aircraft to fly toward the target, and colliding with the target at a desired interception angle.

[0132] The system control input yaw rudder deflection angle described in this application can be obtained based on a method similar to that in step 2, or it can be obtained based on a traditional method, for example, using the method involved in Dong Bin and Li Yuhua's Design of Variable Structure Control System for BTT Missile Pitch / Yaw Channel [J]. Ordnance Technology, 3197-3200+3213.

[0133] Example

[0134] The simulation scenario in a two-dimensional plane simulates the task of an aircraft intercepting a high-maneuverability target in the air. The initial parameters of the aircraft and the target are shown in the following table:

[0135]

[0136]

[0137] The system parameters of the aircraft are obtained by the following formula:

[0138]

[0139]

[0140] in,

[0141] m represents the mass of the aircraft, and its value is m=12kg

[0142] ρ represents the atmospheric density, and its value is ρ=1.39kg / m 3 ,

[0143] V M Indicates the flight speed of the aircraft, the value is V M =600m / s 2 ,

[0144] S represents the reference area of ​​the aircraft, and its value is S = 0.013m 2 ,

[0145] l represents the characteristic length of the aircraft, which is l = 1m.

[0146] It represents the partial derivative of lift coefficient with respect to angle of attack and is taken as

[0147] It represents the pitch moment coefficient with respect to the angle of attack, and its value is

[0148] It represents the pitch moment coefficient to the pitch angular rate, and its value is

[0149] It represents the partial derivative of the pitch moment coefficient with respect to the pitch rudder deflection angle, and its value is

[0150] J z The moment of inertia of the aircraft about the longitudinal axis of the missile is J z =0.85kg·m 2 , and then we get:

[0151] a1=404.5868,a2=633671.532

[0152] a3=-18.2719,b2=-10963.1753

[0153] The system state feedback matrices A1 and A2 are designed as:

[0154]

[0155] The parameters of the reduced-order state observer are β1=1, β2=20,

[0156] Considering the saturation characteristics of the aircraft actuator, the normal overload of the aircraft is limited:

[0157]

[0158] Among them, a M,max is the maximum overload that the aircraft can provide. Due to the limited maneuverability of the aircraft, the acceleration command is limited to a M,max =80g.

[0159] The target mobility is: a T =50sin(0.3t)

[0160] Considering the saturation characteristics of the pitch rudder actuator, the maximum rudder deflection angle of the pitch rudder is also limited:

[0161]

[0162] Among them, δ z,max is the maximum rudder deflection angle that the aircraft actuator can provide, and its maximum rudder deflection angle is δ z,max =15°.

[0163] The simulation was repeated 5 times, and the expected interception angles preset in each simulation were different, namely q d =10°, 30°, 50°, 70°, 90°.

[0164] During each simulation, the control scheme of the aircraft is consistent, and the following control scheme is used:

[0165] Step 1: In the terminal guidance phase, the aircraft simulates the seeker capturing the target, and can provide the aircraft with the missile-target sight angle, missile-target distance, angle of attack, angle of attack angular rate, missile-target sight angular rate, aircraft ballistic inclination angle, and aircraft ballistic inclination angular rate in real time;

[0166] Step 2, obtaining the position and speed information of the aircraft itself through satellite signals, and obtaining the pitch rudder deflection angle of the system control input in real time according to the preset expected interception angle;

[0167] Among them, the system control input pitch rudder deflection angle is obtained in real time through the following formula (1):

[0168]

[0169] Said Obtained by the following formula (II):

[0170]

[0171] The p2 is obtained by Find the integral to obtain:

[0172]

[0173] p′2 represents the p2 value obtained at the last moment,

[0174] The f2 is obtained by the following formula (V):

[0175]

[0176] The dynamic surface s2 is obtained by the following formula (VI):

[0177] s2=α-α d (six)

[0178] Virtual control instructions α d Obtained by the following formula (VII):

[0179]

[0180] Wherein, b1 is obtained by the following formula (VIII):

[0181]

[0182] The dynamic surface s1 is obtained by the following formula (IX):

[0183] s1=qq d (Nine)

[0184] Obtained by the following formula (10):

[0185]

[0186] p1 is obtained by Get points

[0187]

[0188] p′1 represents the p1 value obtained at the last moment, and its initial value is 0.

[0189] The f1 is obtained by the following formula (XII):

[0190]

[0191] Step 3: Generate a rudder command based on the pitch rudder angle input by the system control, control the servo to steer, and control the aircraft to fly to the target.

[0192] The relative motion trajectory curve of the projectile and the target obtained by 5 simulations is as follows: Figure 2 As shown in Figure 2 It can be seen that under the constraints of different expected interception angles, the aircraft can successfully intercept the target; the obtained missile-target line-of-sight angle change curve is as follows: Figure 3 As shown in , the obtained aircraft angle of attack change curve is as follows Figure 4 As shown in , the obtained aircraft overload change curve is as follows Figure 5 As shown in , the obtained aircraft rudder angle change curve is as follows Figure 6 as shown in .

[0193] It can be seen from the above simulation results that under the control of the terminal guidance control method based on the all-wheel drive system provided in the present application, under the action of different expected interception angles, the actual missile-target line-of-sight angle can quickly converge to the expected value and continue to maintain it, which shows that the terminal guidance control method based on the all-wheel drive system provided in the present application is sufficiently effective.

[0194] 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 terminal guidance control method based on an all-wheel drive system, characterized in that: The method comprises the following steps: Step 1: The aircraft captures the target through the seeker in the terminal guidance phase. Step 2: Obtain the position and speed information of the aircraft itself through satellite signals, and obtain the pitch rudder angle of the system control input in real time according to the preset expected interception angle. Step 3, generating a rudder instruction based on the system control input pitch rudder deflection angle and the system control input yaw rudder deflection angle, controlling the servo to steer, controlling the aircraft to fly toward the target, and colliding with the target at a desired interception angle.

2. The terminal guidance control method based on the all-wheel drive system according to claim 1 is characterized in that: The system control input pitch rudder deflection angle is obtained in real time through the following formula (1): Where u represents the system control input pitch rudder deflection angle, b2 represents the model coefficient, represents the control coefficient matrix, represents the control coefficient matrix, s2 represents dynamic surface 2, represents the differential of the dynamic surface 2, f2 represents the model coefficient, represents the observed estimate of the disturbance d2.

3. The terminal guidance control method based on the all-wheel drive system according to claim 2 is characterized in that: Said Obtained by the following formula (II): Where p2 represents the auxiliary variable of the observer of disturbance d2, β2 represents the design parameter of the observer for disturbance d2, α is the angle of attack, The angular rate representing the angle of attack.

4. The terminal guidance control method based on the all-wheel drive system according to claim 3 is characterized in that: The p2 is obtained by Find the integral to obtain: in, represents the derivative of the auxiliary variable of the observer with respect to the disturbance d2, p′2 represents the p2 value obtained at the last moment, u′ represents the pitch rudder deflection angle of the system control input obtained at the last moment, Preferably, b2 is obtained by the following formula (four): in, ρ represents the atmospheric density, V M Indicates the flight speed of the aircraft. S represents the reference area of ​​the aircraft, l represents the characteristic length of the aircraft, represents the partial derivative of the pitch moment coefficient with respect to the pitch rudder deflection angle, J z It represents the moment of inertia of the aircraft about the longitudinal axis of the projectile.

5. The terminal guidance control method based on the all-wheel drive system according to claim 4 is characterized in that: The f2 is obtained by the following formula (V): Among them, the V M Indicates the flight speed of the aircraft. g is the acceleration due to gravity, θ M represents the ballistic inclination angle of the aircraft, α is the angle of attack, represents the angular rate of the angle of attack, a1, a2, and a3 each independently represent a system parameter.

6. The terminal guidance control method based on the all-wheel drive system according to claim 2 is characterized in that: The dynamic surface s2 is obtained by the following formula (VI): s2 = α - α d (VI) Where α represents the angle of attack, α d Indicates virtual control instructions; Said Obtained by differentiating s2.

7. The terminal guidance control method based on the all-wheel drive system according to claim 6 is characterized in that: The virtual control instruction α d Obtained by the following formula (VII): Wherein, b1 represents the model coefficient, which is obtained by the following formula (VIII): r is the distance between the aircraft and the target, q represents the sight angle of the projectile. represents the control coefficient matrix, represents the control coefficient matrix, f1 is the model coefficient, represents the observed estimate of the disturbance d1, s1 represents dynamic surface 1, represents the differential of the dynamic surface one; Preferably, the dynamic surface s1 is obtained by the following formula (IX): s1=qq d (Nine) Among them, q d represents the expected interception angle; Said Obtained by differentiating s1.

8. The terminal guidance control method based on the all-wheel drive system according to claim 7, characterized in that: Said Obtained by the following formula (10): Among them, p1 represents the auxiliary variable of the observer of disturbance d1, β1 represents the design parameter of the observer for disturbance d1, q represents the sight angle of the projectile. Indicates the missile-target line-of-sight angular velocity.

9. The terminal guidance control method based on the all-wheel drive system according to claim 8, characterized in that: The p1 is obtained by Find the integral to obtain: in, represents the derivative of the auxiliary variable of the observer with respect to the disturbance d1, p′1 represents the p1 value obtained at the previous moment.

10. The terminal guidance control method based on the all-wheel drive system according to claim 9, characterized in that: The f1 is obtained by the following formula (XII): Indicates the rate of change of the distance between the aircraft and the target, r is the distance between the aircraft and the target, Indicates the missile-target line-of-sight angular velocity.

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