An angle-constrained guidance method for aircraft that considers the dynamic characteristics of autopilot

Abstract

This invention discloses an angle-constrained guidance method for aircraft considering the dynamic characteristics of an autopilot, comprising the following steps: establishing the relative motion equations between the aircraft and the target; setting autopilot equations to characterize the autopilot with dynamic delays; setting aircraft angle constraints; establishing a guidance system model based on the relative motion equations and the autopilot equations; and obtaining the guidance law based on the guidance system model. The angle-constrained guidance method for aircraft considering the dynamic characteristics of an autopilot disclosed in this invention can control an aircraft to intercept maneuvering aerial targets in a shorter time, improving the guidance accuracy of the aircraft when attacking maneuvering aerial targets.

Description

Technical Field

[0001] This invention relates to an angle-constrained guidance method for aircraft that takes into account the dynamic characteristics of an autopilot, and belongs to the field of guidance and control. Background Technology

[0002] For guided aircraft, an autopilot is a device that automatically controls and stabilizes the aircraft's attitude and motion during flight according to certain technical requirements. It interacts with onboard or ground-based guidance systems to form the aircraft's guidance and control system, which mainly consists of sensitive elements, a computer, and servo mechanisms. When a disturbance threatens to destabilize the aircraft's attitude, the sensitive elements detect the change in attitude, the computer calculates the necessary correction, and the servo mechanisms, based on the proposed correction information, maneuver the aircraft to the desired attitude position to ensure stable flight.

[0003] However, guided aircraft flying within the atmosphere are subject to aerodynamic forces and the delay characteristics of the aircraft's hardware, resulting in dynamic delays in the autopilot and a decrease in guidance accuracy.

[0004] Therefore, it is necessary to conduct more in-depth research on guidance methods in order to reduce the impact of the dynamic delay of autopilot on guidance accuracy. Summary of the Invention

[0005] To overcome the above problems, the inventors conducted in-depth research and proposed an angle-constrained guidance method for aircraft that considers the dynamic characteristics of the autopilot, comprising the following steps:

[0006] S1. Establish the equations of relative motion between the aircraft and the target;

[0007] S2. Set up autopilot equations to characterize autopilots with dynamic delays;

[0008] S3. Set aircraft angle constraints and establish a guidance system model based on the relative motion equations and autopilot equations:

[0009] S4. Obtain the guidance law based on the guidance system model to control the flight of the aircraft.

[0010] In a preferred embodiment, in S1, the equation of relative motion is expressed as:

[0011]

[0012] Where r is the relative distance between the aircraft and the target, q is the tilt angle of the missile's line of sight, θ is the deflection angle of the missile's line of sight, and a mr Let r be the component of the aircraft's acceleration in the direction of r;

[0013] a mrLet r be the component of the aircraft's acceleration in the r direction;

[0014] a mq Let be the component of the aircraft's acceleration in the q direction;

[0015] a mθ Let θ be the component of the aircraft's acceleration in the θ direction;

[0016] a tr Let r be the component of the target's acceleration in the r direction;

[0017] a tq Let q be the component of the target's acceleration in the q direction;

[0018] a tθ Let θ be the component of the target's acceleration in the θ direction.

[0019] In a preferred embodiment, the autopilot equation is expressed as:

[0020]

[0021] Where ξ is the damping ratio, ω n u is the natural frequency of the autopilot. q For guidance commands in the q direction, u θ This is the guidance command in the θ direction.

[0022] In a preferred embodiment, the aircraft angle constraint is expressed as:

[0023] q(t f )=q d

[0024] θ(t f )=θ d

[0025] Among them, t f For the guidance terminal moment, q d θ is the desired terminal line-of-sight tilt angle. d To determine the desired terminal line-of-sight angle, q(t) f θ(t) represents the aircraft's line-of-sight tilt angle at the guidance terminal moment. f The angle of deflection of the aircraft's line of sight at the guidance terminal moment is 0.

[0026] In a preferred embodiment, the guidance system model is represented as follows:

[0027]

[0028] a=-2ξω n b2=ω n2 ;

[0029] Where x1, x2, x3, and x4 represent system states, set as x1 = [qq] d ,θ-θ d ] T , x3=[a mq ,a mθ ] T , u = [u q ,u θ ] T .

[0030] In a preferred embodiment, step S4 includes the following sub-steps:

[0031] S41. Set a fast terminal sliding surface so that system states x1 and x2 converge to the desired value in a fixed time.

[0032] S42. Based on the fast terminal sliding surface, obtain the virtual control law x of system state x3. 3d ;

[0033] S43. Based on the fast terminal sliding surface, obtain the virtual control law x of system state x4. 4d ;

[0034] S44. Obtain the actual guidance law based on the acquired virtual control law.

[0035] In a preferred embodiment, in S41, the rapid terminal sliding surface s1 is configured as follows:

[0036]

[0037] Among them, α1>0, β1>0, m1, n1, p1 and q1 are all positive odd numbers, and m1>n1 and p1<q1.

[0038] In a preferred embodiment, in S42, a virtual control quantity x is set. 3c :

[0039]

[0040] Where α² and β² are constant parameters, h is the maximum value of the saturation function constraint, and m², n², p², and q² are all positive odd numbers. Let be a saturation function, expressed as:

[0041]

[0042] By setting a first-order low-pass filter, the virtual control quantity x is... 3cProcess and obtain the virtual control law x of system state x3. 3d .

[0043] In a preferred embodiment, in S43, a virtual control quantity x is set. 4c :

[0044]

[0045] Where α3 and β3 are constant parameters, and m3, n3, p3 and q3 are all positive odd numbers.

[0046] By setting a nonlinear filter, the virtual control quantity x is... 4c Process and obtain the virtual control law x of system state x4. 4d .

[0047] In a preferred embodiment, in S44, the actual guidance law is:

[0048]

[0049] The beneficial effects of this invention include:

[0050] (1) Improved the guidance accuracy of aircraft when attacking aerial maneuvering targets;

[0051] (2) It can control the aircraft to intercept aerial maneuvering targets in a relatively short time. Attached Figure Description

[0052] Figure 1 This diagram illustrates a preferred embodiment of a multi-constraint terminal velocity optimal mid-course guidance method for long-range gliders according to the present invention.

[0053] Figure 2 The flight trajectories of the aircraft under different initial ballistic inclination angles and ballistic deflection angles are shown in Example 1.

[0054] Figure 3 The image shows the projectile-target line-of-sight tilt angle of the aircraft under different initial trajectory tilt angles and trajectory deflection angles in Example 1;

[0055] Figure 4 The image shows the projectile-to-target line-of-sight tilt rate of the aircraft under different initial trajectory tilt angles and trajectory deflection angles in Example 1;

[0056] Figure 5 The longitudinal acceleration of the aircraft under different initial trajectory inclination angles and trajectory deflection angles is shown in Example 1;

[0057] Figure 6 The image shows the missile-target line-of-sight angle of the aircraft under different initial trajectory inclination angles and trajectory deflection angles in Example 1;

[0058] Figure 7 The image shows the missile-to-target line-of-sight deflection rate of the aircraft under different initial trajectory inclination angles and trajectory deflection angles in Example 1;

[0059] Figure 8 The lateral acceleration of the aircraft under different initial trajectory inclination angles and trajectory deflection angles is shown in Example 1;

[0060] Figure 9 The diagram shows the flight trajectory of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles in Example 1;

[0061] Figure 10 The image shows the missile-eye line-of-sight tilt angle of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles in Example 1;

[0062] Figure 11 The image shows the missile-to-eye line-of-sight tilt rate of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles in Example 1;

[0063] Figure 12 The longitudinal acceleration of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles is shown in Example 1;

[0064] Figure 13 The image shows the missile-eye line-of-sight deflection angle of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles in Example 1;

[0065] Figure 14 This illustrates the missile-eye line-of-sight deflection rate of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles in Example 1;

[0066] Figure 15 The diagram shows the lateral acceleration of the aircraft under different desired terminal line-of-sight tilt angles and deflection angles in Example 1. Detailed Implementation

[0067] 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.

[0068] 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.

[0069] According to the present invention, an angle-constrained guidance method for aircraft considering the dynamic characteristics of an autopilot is provided, such as... Figure 1 As shown, it includes the following steps:

[0070] S1. Establish the equations of relative motion between the aircraft and the target;

[0071] S2. Set up autopilot equations to characterize autopilots with dynamic delays;

[0072] S3. Set aircraft angle constraints and establish a guidance system model based on the relative motion equations and autopilot equations:

[0073] S4. Obtain the guidance law based on the guidance system model to control the flight of the aircraft.

[0074] In S1, the equation of relative motion is expressed as:

[0075]

[0076] Where r is the relative distance between the aircraft and the target, q is the tilt angle of the missile's line of sight, θ is the deflection angle of the missile's line of sight, and a mr Let r be the component of the aircraft's acceleration in the direction of r;

[0077] a mr Let r be the component of the aircraft's acceleration in the r direction;

[0078] a mq Let be the component of the aircraft's acceleration in the q direction;

[0079] a mθ Let θ be the component of the aircraft's acceleration in the θ direction;

[0080] a tr Let r be the component of the target's acceleration in the r direction;

[0081] a tq Let q be the component of the target's acceleration in the q direction;

[0082] a tθ Let θ be the component of the target's acceleration in the θ direction.

[0083] In S2, the autopilot equation is expressed as:

[0084]

[0085] Where ξ is the damping ratio, ω n u is the natural frequency of the autopilot. q For guidance commands in the q direction, u θ This is the guidance command in the θ direction.

[0086] Compared to traditional autopilot equations, the above autopilot equations fully consider the dynamic delay problem, resulting in higher guidance accuracy of the subsequent guidance law.

[0087] According to the present invention, in S3, the aircraft angle constraint is expressed as:

[0088] q(t f )=q d

[0089] θ(t f )=θ d

[0090] Among them, t f For the guidance terminal moment, q d θ is the desired terminal line-of-sight tilt angle. d To determine the desired terminal line-of-sight angle, q(t) f θ(t) represents the aircraft's line-of-sight tilt angle at the guidance terminal moment. f The angle of deflection of the aircraft's line of sight at the guidance terminal moment is 0.

[0091] The aforementioned aircraft angle constraint transforms the problem of aircraft attack angle constraint into a terminal line-of-sight angle constraint problem, reducing the subsequent computational load.

[0092] In S3, the guidance system model is represented as follows:

[0093]

[0094] a=-2ξω n b2=ω n 2 ;

[0095] Where x1, x2, x3, and x4 represent system states, set as x1 = [qq] d ,θ-θ d ] T , x3=[a mq ,a mθ ] T , u = [u q ,u θ ] T .

[0096] This guidance model fully considers the dynamic characteristics of the autopilot and the attack angle constraints of the aircraft, providing a foundation for the accurate acquisition of guidance commands.

[0097] In a preferred embodiment, S4 includes the following sub-steps:

[0098] S41. Set a fast terminal sliding surface so that system states x1 and x2 converge to the desired value in a fixed time.

[0099] S42. Based on the fast terminal sliding surface, obtain the virtual control law x of system state x3. 3d ;

[0100] S43. Based on the fast terminal sliding surface, obtain the virtual control law x of system state x4. 4d ;

[0101] S44. Obtain the actual guidance law based on the acquired virtual control law.

[0102] In S41, by making the system states x1 and x2 converge to the desired values ​​over a fixed time, the aircraft can accurately hit the target at the desired attack angle.

[0103] Furthermore, the rapid terminal sliding surface s1 is configured as follows:

[0104]

[0105] Among them, α1>0, β1>0, m1, n1, p1 and q1 are all positive odd numbers, and m1>n1 and p1<q1.

[0106] The first derivative of the fast terminal sliding surface s1 can be expressed as:

[0107]

[0108]

[0109] In S42, the sliding mode error surface is defined as...

[0110]

[0111] Differentiating it and combining it with the guidance system model, we can obtain:

[0112]

[0113]

[0114] The aforementioned terminal sliding mode exhibits a singularity problem, and traditional methods of setting virtual control rates cannot effectively yield results.

[0115] In this invention, the singularity problem is overcome by the following method: specifically, a virtual control quantity x is set. 3c :

[0116]

[0117] Where α² and β² are constant parameters, h is the maximum value of the saturation function constraint, and m², n², p², and q² are all positive odd numbers. It is a saturation function used to limit singular terms in the control input. The amplitude of the saturation function is expressed as:

[0118]

[0119] By setting a first-order low-pass filter, the virtual control quantity x is... 3c Process and obtain the virtual control law x of system state x3. 3d .

[0120] The first-order low-pass filter is configured as follows:

[0121]

[0122] Where τ is the filter coefficient, and λ1 and λ2 are constant parameters. Preferably, λ1 > 1 and 0 < λ2 < 1.

[0123] Traditional sliding mode control often requires multiple derivatives of the virtual control law, leading to the "differential expansion" problem. In this invention, a new virtual control variable x is introduced, drawing on the dynamic surface design method. 3d It is composed of x 3c It is obtained by passing through a first-order low-pass filter.

[0124] Furthermore, traditional dynamic surface design methods often employ first-order linear filters, which cannot guarantee finite / fixed-time convergence of the system. In this invention, by setting the aforementioned first-order low-pass filter, the fixed-time convergence characteristics of the system are guaranteed, thereby ensuring the guidance accuracy of the aircraft.

[0125] In S43,

[0126] Define the sliding mode error surface as

[0127] s3 = x3 - x 3d

[0128] Differentiating it and combining it with the guidance system model, we can obtain:

[0129]

[0130] Similar to S42, in S43, the virtual control quantity x is set. 4c :

[0131]

[0132] Where α3 and β3 are constant parameters, and m3, n3, p3 and q3 are all positive odd numbers.

[0133] By setting a nonlinear filter, the virtual control quantity x is... 4c Process and obtain the virtual control law x of system state x4. 4d .

[0134] The nonlinear filter is configured as follows:

[0135]

[0136] In S44, the sliding mode error surface is defined as...

[0137] s4 = x4 - x 4d

[0138] Differentiating it and combining it with the guidance system model, we can obtain:

[0139]

[0140] This leads to the actual guidance law, expressed as:

[0141]

[0142] Example

[0143] Example 1

[0144] To obtain a guidance law with aircraft angle constraints through simulation experiments, the following steps are included:

[0145] S1. Establish the equations of relative motion between the aircraft and the target;

[0146] S2. Set up autopilot equations to characterize autopilots with dynamic delays;

[0147] S3. Set aircraft angle constraints and establish a guidance system model based on the relative motion equations and autopilot equations:

[0148] S4. Obtain the guidance law based on the guidance system model.

[0149] In S1, the equation of relative motion is expressed as:

[0150]

[0151] In S2, the autopilot equation is expressed as:

[0152]

[0153] Where ξ is set to 0.8, ω n = 8 rad / s.

[0154] In S3, the aircraft angle constraint is expressed as:

[0155] q(t f )=q d

[0156] θ(t f )=θ d

[0157] The guidance system model is represented as follows:

[0158]

[0159] a=-2ξω n b2=ω n 2

[0160] S4 includes the following sub-steps:

[0161] S41. Set a fast terminal sliding surface so that system states x1 and x2 converge to the desired value in a fixed time.

[0162] S42. Based on the fast terminal sliding surface, obtain the virtual control law x of system state x3. 3d ;

[0163] S43. Based on the fast terminal sliding surface, obtain the virtual control law x of system state x4. 4d ;

[0164] S44. Obtain the actual guidance law based on the acquired virtual control law.

[0165] In S41, the fast terminal sliding surface s1 is configured as follows:

[0166]

[0167] In S42, the sliding mode error surface is defined as...

[0168]

[0169] Set virtual control quantity x 3c :

[0170]

[0171]

[0172] By setting a first-order low-pass filter, the virtual control quantity x is... 3c Process and obtain the virtual control law x of system state x3. 3d .

[0173] The first-order low-pass filter is configured as follows:

[0174]

[0175] In S43,

[0176] Define the sliding mode error surface as

[0177] s3 = x3 - x 3d

[0178] Set virtual control quantity x 4c :

[0179]

[0180] By setting a nonlinear filter, the virtual control quantity x is... 4c Process and obtain the virtual control law x of system state x4. 4d .

[0181] The nonlinear filter is configured as follows:

[0182]

[0183] In S44, the sliding mode error surface is defined as...

[0184] s4 = x4 - x 4d

[0185] The actual guidance law is obtained, expressed as:

[0186]

[0187] Among them, set m1=m2=m3=m4=5, n1=n2=n3=n4=3, p1=p2=p3=p4=3, q1=q2=q3=q4=5, τ=0. 1, α1=0.4, β1=0.1, α2=10, β2=0.01, α3=α4=1, β3=β4=1, H=100, λ1=1.3, λ2=0.5.

[0188] In the simulation experiment, the initial position of the aircraft was set to (0m, 0m, 0m), the initial position of the target was set to (8000m, 6000m, 2000m), and the velocity of the aircraft was v. m =500m / s, the target's velocity is v m =300m / s², the maximum acceleration of the aircraft is set to a. mmax =200m / s 2 .

[0189] Multiple sets of tests were conducted, in which the desired terminal line-of-sight tilt angle of the aircraft was set to q. d =50°, the desired terminal line-of-sight deviation angle of the aircraft is set to θ. d = -20°, target acceleration set to a ty2 =a tz2 =30m / s 2 The initial trajectory inclination of the target is γ. m0 =150°, the initial ballistic deviation angle of the target is ψ vm0 = -10°, set the flight with different initial trajectory inclination angles γ m0 and deflection angle ψ vm0 The targets of the attack are as follows:

[0190] Set the initial trajectory inclination angle γ of the aircraft. m0 30°, initial ballistic deflection angle ψ vm0 -30°;

[0191] Set the initial trajectory inclination angle γ of the aircraft. m0 60°, initial ballistic deflection angle ψ vm0 It is 0°;

[0192] Set the initial trajectory inclination angle γ of the aircraft. m0 90°, initial ballistic deflection angle ψ vm0 The angle was set to 30° to observe the simulation results of the aircraft under different initial trajectory tilt angles and deflection angles.

[0193] Simulation results are as follows Figure 2-8 As shown, where, Figure 2 The flight trajectories of the aircraft are shown under different initial ballistic inclination angles and ballistic deflection angles; Figure 3 The missile-target line-of-sight tilt angle of the aircraft is shown under different initial ballistic tilt angles and ballistic deflection angles; Figure 4 The projectile-target line-of-sight tilt rate of the aircraft is shown under different initial ballistic tilt angles and ballistic deflection angles; Figure 5 The longitudinal acceleration of the aircraft is shown under different initial trajectory inclination angles and trajectory deflection angles; Figure 6 The missile-target line-of-sight angle of the aircraft is shown under different initial trajectory inclination angles and trajectory deflection angles; Figure 7 The missile-target line-of-sight deflection rate of the aircraft is shown under different initial trajectory inclination angles and trajectory deflection angles. Figure 8 The lateral acceleration of the aircraft is shown under different initial trajectory inclination angles and trajectory deflection angles.

[0194] from Figure 2 It can be seen that the aircraft can successfully intercept targets under different initial trajectory inclination angles and trajectory deflection angles. Figure 3 , 4 As can be seen from points 6 and 7, this guidance law can adjust the projectile-target line-of-sight tilt angle q and the projectile-target line-of-sight tilt angle rate. Bullet line of sight deflection angle θ, bullet line of sight deflection rate It gradually converges towards the expected value, and can converge smoothly to the expected value. From Figure 5 , 8 It can be seen that the aircraft can converge within a finite time, and the convergence speed is fast and relatively stable. From Figure 2-8 It can be seen that the aircraft can effectively intercept aerial maneuvering targets in a relatively short period of time.

[0195] Example 2

[0196] The same simulation experiment as in Example 1 was conducted, except that multiple sets of tests were performed, in which the initial trajectory inclination angle γ of the flight was set. m060°, initial ballistic deflection angle ψ vm0 The angle is 0°, and the target acceleration is set to a. ty2 =a tz2 =30m / s 2 The initial trajectory inclination of the target is γ. m0 =150°, the initial ballistic deviation angle of the target is ψ vm0 = -10°, set the flight to attack targets with different desired terminal line-of-sight tilt and deflection angles, respectively:

[0197] Set the desired terminal line-of-sight tilt angle of the aircraft to q. d =20°, the expected terminal line-of-sight angle of the aircraft is θ d = -20°;

[0198] Set the desired terminal line-of-sight tilt angle of the aircraft to q. d =30°, the desired terminal line-of-sight angle of the aircraft is θ d = -30°;

[0199] Set the desired terminal line-of-sight tilt angle of the aircraft to q. d =60°, the expected terminal line-of-sight angle of the aircraft is θ d = -10°;

[0200] The simulation results were observed under different desired terminal line-of-sight tilt angles and deflection angles.

[0201] Simulation results are as follows Figure 9-15 As shown, where, Figure 9 The flight trajectory diagrams of the aircraft are shown under different expected terminal line-of-sight tilt angles and deflection angles; Figure 10 The missile-target line-of-sight tilt angle of the aircraft is shown under different expected terminal line-of-sight tilt angles and deflection angles; Figure 11 The missile-to-eye line-of-sight tilt rate of the aircraft under different expected terminal line-of-sight tilt angles and deflection angles is shown. Figure 12 The longitudinal acceleration of the aircraft is shown under different desired terminal line-of-sight tilt angles and deflection angles. Figure 13 The missile-target line-of-sight deflection angle of the aircraft is shown under different expected terminal line-of-sight tilt angles and deflection angles. Figure 14 The missile-to-eye line-of-sight deflection rate of the aircraft under different expected terminal line-of-sight tilt angles and deflection angles is shown. Figure 15 The lateral acceleration of the aircraft is shown under different desired terminal line-of-sight tilt and deflection angles.

[0202] from Figure 9 It can be seen that the aircraft can successfully intercept targets under different expected terminal line-of-sight tilt angles and deflection angles. Figure 10 , 11 As can be seen from 13 and 14, this guidance law can make the missile-target line-of-sight tilt angle q and the missile-target line-of-sight tilt angle rate... Bullet line of sight deflection angle θ, bullet line of sight deflection rate It gradually converges towards the expected value, and can converge smoothly to the expected value. From Figure 12 , 15 It can be seen that the aircraft can converge within a finite time, and the convergence speed is fast and relatively stable. From Figure 9-15 It can be seen that the aircraft can effectively intercept aerial maneuvering targets in a relatively short period of time.

[0203] In the description of this invention, it should be noted that the terms "upper," "lower," "inner," "outer," "front," and "rear," etc., indicate the orientation or positional relationship based on the orientation or positional relationship in the working state of this invention, and are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this invention. Furthermore, the terms "first," "second," "third," and "fourth" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0204] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal communication between two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0205] 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 method for aircraft angular constraint guidance considering autopilot dynamics, characterized in that, The method comprises the following steps: S1, establishing a relative motion equation of the aircraft and the target; S2, setting an autopilot equation for representing the autopilot with dynamic delay; S3, setting an aircraft angle constraint, and establishing a guidance system model based on the relative motion equation and the autopilot equation; S4, obtaining a guidance law according to the guidance system model to control the flight of the aircraft; In S1, the relative motion equation is expressed as: where r is the relative distance between the vehicle and the target, q is the line-of-sight angle between the vehicle and the target, θ is the line-of-sight angle between the vehicle and the target, a mr is the component of the acceleration of the vehicle in the direction of r a mq is the component of the acceleration of the aircraft in the q direction; a mθ is the component of the acceleration of the aircraft in the θ direction; a tr The component of the targeted acceleration in the r direction; a tq Component of the acceleration targeted in the q direction; a tθ a component of the targeted acceleration in the θ direction; The autopilot equation is expressed as: wherein ξ is a damping ratio, ω n is a natural frequency of the autopilot, u q is a guidance command in the q direction, and uθ is a guidance command in the θ direction. The aircraft angle constraint is expressed as: q(t f ) = q d θ(t f ) = θ d where t f is the guidance terminal time, q d is the desired terminal line-of-sight elevation angle, θ d is the desired terminal line-of-sight azimuth angle, q(t f ) is the vehicle line-of-sight elevation angle at the guidance terminal time, and θ(t f ) is the vehicle line-of-sight azimuth angle at the guidance terminal time. The guidance system model is expressed as: a = -2ξω n , b2= ω n 2 ; wherein x1, x2, x3, x4 are system states, set as x1 = [q - q d , θ - θ d ] T , x3 = [a mq , a mθ ] T , u = [u q , u θ ] T In S4, the following sub-steps are included: S41, setting a fast terminal sliding surface, so that the system states x1 and x2 are fixed-time convergent to the expected values; S42, based on the fast terminal sliding mode surface, a virtual control law x of the system state x3 is obtained 3d ; S43, based on the fast terminal sliding mode surface, a virtual control law x4 of the system state x is obtained 4d ; S44, obtaining an actual guidance law according to the obtained virtual control rate; In S41, the fast terminal sliding surface s1 is set as: Wherein, α1>0, β1>0, m1, n1, p1 and q1 are all positive odd numbers, and m1>n1, p1 In S42, the virtual control amount x is set 3c : where s2 is a sliding mode error surface, a2, b2 are constant parameters, h is the maximum value of the saturation function limit, m2, n2, p2 and q2 are all positive odd numbers, is a saturation function, and is expressed as: By setting a first-order low-pass filter to the virtual control variable x 3c The processing obtains the virtual control law x of the system state x3 3d In S43, the virtual control variable x is set 4c : s3 = x3 - x 3d Wherein, α3, β3 are constant parameters, and m3, n3, p3 and q3 are all positive odd numbers; By setting a non-linear filter on the virtual control variable x 4c processing, a virtual control law x 4d for the system state x4 is obtained In S44, the actual guidance law is:

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