Three-dimensional guidance law and system for intercepting terminal angle and view field constraint of maneuvering target

Abstract

The invention relates to the field of advanced guidance control of aircrafts, and discloses a terminal angle and view field constraint three-dimensional guidance law and system for intercepting a maneuvering target, and the specific steps are as follows: firstly, designing a sight angle rate profile expected to be tracked, and then designing a terminal angle and view field constraint guidance law of preset time. The advanced preset time control technology and the novel sight angle rate forming technology are adopted for achieving aircraft guidance control, the guidance law with the terminal angle and view field constraint is designed, and the maneuvering target can be intercepted at the expected terminal angle.

Description

Technical Field

[0001] The present invention relates to the field of advanced guidance and control of aircraft, and specifically to a three-dimensional guidance law and system with terminal angle and field of view constraints for intercepting maneuvering targets. Background Art

[0002] With the rapid development of anti-missile defense systems, the proportional navigation guidance law that can only meet the terminal miss distance constraint can no longer meet the requirements of modern missile combat missions. Therefore, advanced guidance laws with multiple constraints have been proposed. Among them, the attack-time guidance law can conduct saturation strikes on the target to be struck, the attack-angle guidance law can improve the strike effect, and the guidance law with field-of-view constraints can ensure that the seeker always locks on to the target. A guidance law that has both field-of-view constraints and strike-angle constraints can significantly enhance combat capabilities, yet this is also a more challenging problem.

[0003] Currently, a large amount of research has been devoted to the development of attack-angle guidance laws. According to the implemented algorithm principles, they can be divided into offset proportional navigation guidance laws, nonlinear control guidance laws, and optimal guidance laws. However, most of the existing research only focuses on stationary targets, and there is less research on the attack-angle guidance law for maneuvering targets.

[0004] Patent Publication No. CN119396167A discloses a variable-gain backstepping maneuvering target interception guidance method considering attack-angle constraints. This guidance law designs a polynomial function with the flown distance as the independent variable to represent the line-of-sight angle reference profile and designs a guidance law to achieve the tracking of the line-of-sight angle profile. However, this guidance law is only designed in a two-dimensional engagement plane and is difficult to apply in actual combat scenarios. Patent Publication No. CN116227343A discloses a design method for a terminal angle attack guidance law for intercepting maneuvering targets that satisfies the field-of-view angle constraint. This invention uses a neural network estimation model to estimate the disturbance caused by target maneuvering and uses a barrier Lyapunov function to ensure the field-of-view constraint. However, the introduction of the barrier Lyapunov function will significantly reduce the control effect of the guidance law and sacrifice the control speed while ensuring the field-of-view constraint. Patent Publication No. CN117452962A discloses an aircraft angle constraint guidance and control method. This invention uses a timed disturbance observer to observe the disturbance caused by target maneuvering and compensates for it. However, this guidance law does not consider the field-of-view constraint of the aircraft and cannot ensure that the seeker always locks on to the target. Patent Publication No. CN119148509A discloses an angle constraint geometric guidance method for maneuvering targets. This invention patent conducts the design in a relative guidance coordinate system, uses a second-order Bezier curve to generate the relative motion trajectory between the aircraft and the target, and uses a trajectory tracking algorithm to solve the guidance command. However, this invention patent does not consider the field-of-view constraint and only conducts the design in a two-dimensional engagement plane, which is relatively limited in practical applications.

[0005] The reasons for the current research status of strike angle rate guidance can be summarized as follows: ① When designing the strike angle guidance law in a two-dimensional engagement space, there is no need to consider the coupling in the pitch and yaw directions, and the design is more convenient and fast; ② In the case of target maneuvering, it is a very difficult problem to calculate the analytical solution of the lead angle of the aircraft, and it is difficult to ensure the realization of the field of view constraint. ③ Most of the current guidance laws are designed in the velocity system, and there are inevitable problems in the velocity coordinate system, such as the coupling in the pitch and yaw directions, which also leads to the slow development of the current three-dimensional angle constraint guidance law.

[0006] Based on the above-mentioned research and development problems of the strike angle guidance law, it is necessary to design an advanced guidance law with terminal angle and field of view angle constraints for intercepting maneuvering targets in a three-dimensional engagement space. Summary of the Invention

[0007] The purpose of the present invention is to provide a three-dimensional guidance law and system with terminal angle and field of view constraints for intercepting maneuvering targets, so as to solve the problems raised in the above background technology.

[0008] To achieve the above purpose, the present invention provides the following technical solution: A three-dimensional guidance law with terminal angle and field of view constraints for intercepting maneuvering targets, including the following steps:

[0009] Step 1: Based on the vector relative guidance coordinate system, design a line-of-sight angular rate profile that realizes terminal angle and field of view constraints;

[0010] Step 2: Design a preset time guidance law so that the aircraft can track the expected line-of-sight angular rate profile designed in Step 1 within the preset time;

[0011] Step 3: Based on the preset time control technology, design a preset time disturbance observer to observe the disturbance caused by target maneuvering within the preset time;

[0012] Step 4: Based on the expected line-of-sight angular rate profile in Step 1, the preset time guidance law in Step 2, and the preset time disturbance observer in Step 3, it is possible to intercept the maneuvering target at the expected terminal angle.

[0013] Step 5: By analyzing the relationship between the expected line-of-sight angular rate profile designed in Step 1 and the maximum relative lead angle, ensure that the field of view constraint is always satisfied.

[0014] Preferably: The three-dimensional strike guidance law in Step 1 is designed in the relative vector guidance coordinate system, and the specific engagement model can be described as follows: First, the construction of the guidance law described in Step 1 is completed in the relative vector guidance coordinate system. In this combat scenario, the motion state of the missile can be represented by the velocity vector V M and the control acceleration vector AM is represented, while the motion state of the target is represented by the velocity vector V T and the control acceleration vector A T described. Define the line-of-sight vector between the missile and the target as R, where R = ||R|| is the missile-target distance, and V M = ||V M || and V T = ||V T || are the motion speeds of the missile and the target respectively. Let and be unit basis vectors. In the system description, the relative velocity vector V R = V M - V T and its unit vector as well as the relative acceleration vector A R = A M - A T are introduced, where A R and V R satisfy the orthogonal relationship. Based on the above, the following basic relationship for differentiation with respect to time is established:

[0015]

[0016]

[0017] In the formula, Ω L and Ω VR represent the rotational angular velocities of the line-of-sight vector and the relative velocity vector respectively. Define the lead angle as σ M = acos(r·v M ), where acos() is the arccosine function. At the same time, define the relative lead angle σ R = acos(r·v R ). Differentiating it gives:

[0018]

[0019] where, In addition, performing a differentiation operation on the rotational angular velocity of the line-of-sight vector gives:

[0020]

[0021] Here, D represents the interference vector generated by the target maneuvering characteristics. In addition, the derivative of the missile-target distance R with respect to time is:

[0022]

[0023] where, V R = ||V R || is the magnitude of the vector V R .

[0024] Preferably, based on the vector relative guidance coordinate system in step 1, a line-of-sight angular rate profile that can achieve terminal angle and field-of-view constraints is designed. The specific desired line-of-sight angular rate profile is:

[0025]

[0026] where k > 0, t go = t f - t, t f is the terminal strike time, t is the actual flight time of the missile, e is the terminal angle error, and r d is the unit vector in the desired terminal angle direction.

[0027] Preferably, in step 2, a preset-time guidance law is designed so that the aircraft can track the desired line-of-sight angular rate profile designed in step 1 within a preset time. The specific guidance law design is:

[0028]

[0029] where S = Ω L - Ω Ld , is the disturbance observation value caused by target maneuver, S0 is the initial value of S, t s1 is the convergence time of the tracking error vector S that needs to be designed, and 0 < p < 1 is a constant.

[0030] Preferably, based on the preset-time control technology in step 3, a preset-time disturbance observer is designed to be able to observe the disturbance caused by target maneuver within a preset time. The specific preset-time disturbance observer is:

[0031]

[0032] where

[0033]

[0034] where it is assumed that the disturbance caused by target maneuver satisfies ||D|| ≤ D max , and the derivative of the disturbance is bounded and satisfies λ min (·) is the minimum eigenvalue of the matrix, p is the estimated initial value of the disturbance, L is the matrix that needs to be designed, c is the parameter that needs to be designed, and it needs to satisfy c ≥ (δ + ||L||D max ) / (λ min (L)η0), where η0 is the estimated initial error of the disturbance, t s2 is the convergence time that needs to be designed for the disturbance observer, αmax >0 is a parameter to be designed.

[0035] Preferably: In step 5, by analyzing the relationship between the designed desired line-of-sight angular rate profile and the maximum relative lead angle, specifically, the relationship between the maximum relative lead angle and the designed line-of-sight angular rate profile is:

[0036] σ R,max = a tan(ke0) (11)

[0037] where a tan() is the arctangent function and e0 is the initial value of the impact angle error. By analyzing the relationship between σ R,max and the maximum feasible maximum relative lead angle, and then designing the guidance parameters k and the feasible impact error angle region, the constraint of the relative lead angle can be achieved.

[0038] Preferably: In step 5, by constraining the relative lead angle, the constraint of the missile's lead angle is ensured, and thus the target is always locked. The relationship between the missile's relative lead angle and the lead angle is as follows:

[0039]

[0040] where is the maximum lead angle of the missile, and asin() is the arcsine function.

[0041] According to the above three-dimensional guidance system for intercepting a maneuvering target with terminal angle and field of view constraints, it includes: an aircraft and target real-time position calculation unit, an aircraft real-time guidance law calculation unit, a target disturbance real-time disturbance estimation unit, and an aircraft guidance control unit, where:

[0042] The aircraft and target real-time position calculation unit is used to perform real-time iterative calculations of the aircraft and target positions according to the real-time aircraft controller and the target maneuver amount;

[0043] The aircraft real-time guidance law calculation unit calculates the real-time control amount of the aircraft according to the designed terminal angle and field of view constraint guidance law;

[0044] The target disturbance real-time estimation unit performs real-time target disturbance estimation according to the designed preset time disturbance observer to compensate for the error caused by target maneuvering.

[0045] The aircraft guidance control unit iteratively calculates the aircraft control amount in real time according to the terminal angle and field of view constraint guidance law and the preset time disturbance observer, drives the servo system of the aircraft to respond, realizes flight control, and guides the aircraft to intercept the maneuvering target at the desired terminal angle.

[0046] The beneficial effects of the present invention compared with the prior art are:

[0047] The present invention proposes a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints, designs a method for tracking the vector line-of-sight angular rate, and can intercept maneuvering targets with the desired terminal angle. According to the designed preset-time disturbance observer, the disturbance caused by target maneuvering can be observed within the preset time, and then the disturbance caused by target maneuvering can be compensated. The design process of this guidance law is clear and concise, and it can be adapted to intercept targets with the terminal desired angle for stationary, moving, and maneuvering targets in a three-dimensional engagement space. The parameter design is clear, and users can adjust the parameters according to the actual scenario requirements, and it has a wide range of applications. BRIEF DESCRIPTION OF THE DRAWINGS

[0048] Figure 1 is the three-dimensional relative guidance coordinate system of the missile and the target of the present invention;

[0049] Figure 2 is the three-dimensional relative vector guidance coordinate system of the missile and the target of the present invention;

[0050] Figure 3 is the flow chart of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints;

[0051] Figure 4 is the three-dimensional trajectory of the aircraft and the three-dimensional trajectory diagram of the target of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints;

[0052] Figure 5 is the missile overload curve in the embodiment;

[0053] Figure 6 is the missile-target distance diagram of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints;

[0054] Figure 7 is the line-of-sight inclination angle, line-of-sight declination angle, and strike angle error diagram of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints;

[0055] Figure 8 is the curve diagram of the relative lead angle and the change of the lead angle of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints; Figure 9 is the line-of-sight angular rate tracking error diagram of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints; Figure 10 is the disturbance observation diagram of the preset-time disturbance observer of a three-dimensional guidance law and system for intercepting maneuvering targets with terminal angle and field of view constraints. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0056] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0057] Embodiment

[0058] Please refer to Figures 1-3 , a three-dimensional guidance law for terminal angle and field of view constraints for intercepting a maneuvering target in the illustration, including the following steps:

[0059] Step 1: Based on the vector relative guidance coordinate system, design a line-of-sight angular rate profile that realizes terminal angle and field of view constraints;

[0060] Step 2: Design a preset-time guidance law so that the aircraft can track the desired line-of-sight angular rate profile designed in Step 1 within a preset time;

[0061] Step 3: Based on the preset-time control technology, design a preset-time disturbance observer to observe the disturbance caused by the target maneuver within a preset time;

[0062] Step 4: Based on the desired line-of-sight angular rate profile in Step 1, the preset-time guidance law in Step 2, and the preset-time disturbance observer in Step 3, it is possible to intercept the maneuvering target at the desired terminal angle;

[0063] Step 5: By analyzing the relationship between the desired line-of-sight angular rate profile designed in Step 1 and the maximum relative lead angle, ensure that the field of view constraint is always satisfied.

[0064] Figure 1 This is the three-dimensional relative guidance coordinate system of the missile and the target of the present invention. In combination with Figure 1 , the engagement model between the missile and the moving target is specifically described as follows:

[0065] The motion state of the missile can be represented by the velocity vector V M and the control acceleration vector A M , while the motion state of the target is described by the velocity vector V T and the control acceleration vector A T . Define the line-of-sight vector between the missile and the target as R, R = ||R|| is the missile-target distance, V M = ||V M || and V T = ||V T || are the motion speeds of the missile and the target respectively, and let and Taking the unit basis vector, the relative velocity vector V is introduced in the system description. R = V M - V T and its unit vector as well as the relative acceleration vector A R = A M - A T , where A R and V R satisfy the orthogonal relationship. Based on these definitions, the following basic relationship of derivative with respect to time can be established:

[0066]

[0067]

[0068] Ω in the formula L and represent the rotational angular velocities of the line-of-sight vector and the relative velocity vector respectively. The lead angle is defined as σ M = acos(r·v M ), where acos() is the arccosine function. At the same time, the relative lead angle σ R = acos(r·v R ). Taking the derivative of it, we can get:

[0069]

[0070] where In addition, by taking the derivative operation on the rotational angular velocity of the line-of-sight vector, we can get:

[0071]

[0072] where D represents the interference vector generated by the target maneuvering characteristics. In addition, the derivative of the missile-to-target distance R with respect to time is:

[0073]

[0074] where V R = ||V R || is the magnitude of the vector V R .

[0075] Furthermore, based on the vector relative guidance coordinate system described in step 1, a line-of-sight angular rate profile that can achieve terminal angle and field-of-view constraints is designed. The specific steps are as follows:

[0076] Step 1-1: The desired line-of-sight angular rate profile is designed as:

[0077]

[0078] where k > 0, t go = t f -t, t f is the terminal strike time, t is the actual flight time of the missile, e is the terminal angle error, r d is the unit vector in the desired terminal angle direction;

[0079] Step 1-2: According to the formula cos e = r·r d , take the derivative and substitute to obtain:

[0080]

[0081] Step 1-3: Assume that the designed control law can stably track the designed line-of-sight angular rate before time t s1 . Substitute the designed desired line-of-sight angular rate designed in Step 1-1 into formula (20) to obtain that the terminal angle error satisfies the following dynamic equation:

[0082]

[0083] According to formula (21), the analytical solution of the terminal angle error can be further obtained as:

[0084]

[0085] where e0 is the initial terminal angle error, t go,0 is the initial value of t go ;

[0086] Step 1-4: The terminal angle error obtained in Step 1-3 indicates that the designed desired line-of-sight angular rate profile in Step 1-1 can achieve the terminal angle and field-of-view constraints.

[0087] Furthermore, based on the designed preset time guidance law described in Step 2, the aircraft can track the desired line-of-sight angular rate profile designed in Step 1 within the preset time. The specific steps are as follows:

[0088] Step 2-1: Construct the line-of-sight angular rate tracking error as

[0089] S = Ω L - Ω Ld (56)

[0090] Step 2-2: Take the derivative of formula (23) to obtain:

[0091]

[0092] Step 2-3: Using the preset time control technique, make formula (24) converge to zero within the preset time, and design the control quantity as:

[0093]

[0094] where, is the disturbance observation value brought by the target maneuver, S0 is the initial value of S, t s1 is the convergence time of the tracking error vector S that needs to be designed, and 0 < p < 1 is a constant;

[0095] Step 2-4: Construct the Lyapunov function as

[0096] V1 = 2S T S (59)

[0097] Next, take the derivative of formula (26) and substitute the guidance law (25) into it, and we can get:

[0098]

[0099] where, is the disturbance observation error. According to formula (27), it can be obtained that the terminal angle error can converge to zero within the preset time t s1 .

[0100] Furthermore, based on the preset time control technique described in Step 3, a preset time disturbance observer is designed, which can observe the disturbance brought by the target maneuver within the preset time. The specific steps are as follows:

[0101] Step 3-1: The specific preset time disturbance observer is designed as:

[0102]

[0103] where

[0104]

[0105] where, it is assumed that the disturbance brought by the target maneuver satisfies ||D|| ≤ D max , the derivative of the disturbance is bounded and satisfies λ min (·) is the minimum eigenvalue of the matrix, p is the estimated initial value of the disturbance, L is the matrix that needs to be designed, c is the parameter that needs to be designed, and it needs to satisfy c ≥ (δ + ||L||D max ) / (λ min (L)η0), where η0 is the estimated initial error of the disturbance, t s2 is the convergence time that the disturbance observer needs to be designed, and α max > 0 is the parameter that needs to be designed;

[0106] Step 3-2. The disturbance estimation error is:

[0107]

[0108] Step 3-3. Take the derivative of formula (30) and substitute the disturbance observer designed in Step 3-1, and we can get:

[0109]

[0110] Step 3-4. Construct the Lyapunov function as:

[0111]

[0112] Take the derivative of formula (32) and substitute formula (31) into it, and we can get:

[0113]

[0114] Step 3-5. According to the Cauchy-Schwarz inequality, we can get:

[0115]

[0116] Since L is positive definite and there exists a minimum eigenvalue λ min (L)>0, we can further get:

[0117]

[0118] Further, according to ||D||≤D max and the properties of matrix norms, we can get:

[0119]

[0120] Step 3-6. Substitute formulas (34), (35) and (36) obtained in Step 3-5 into it. Therefore, formula (33) can be further transformed into:

[0121]

[0122] Step 3-7. Substitute α(t) designed in Step 3-1 into (37), and we can get:

[0123]

[0124] Therefore, when there is the error gradually decays.

[0125] If Using the comparison theorem, we can get:

[0126]

[0127] Furthermore, it can be obtained that:

[0128]

[0129] According to formula (40), it can be obtained that when time t → t s2 the disturbance estimation error can strictly converge to zero.

[0130] Furthermore, based on the desired line-of-sight angular rate profile in step 1, the preset time guidance law in step 2, and the preset time disturbance observer in step 3 described in step 4, it is possible to intercept a maneuvering target at a desired terminal angle. The specific steps are as follows:

[0131] Step 4-1: Load the k parameter of the desired line-of-sight angular rate profile described in step 1;

[0132] Step 4-2: Load the t s1 parameter and p parameter of the preset time guidance law described in step 2;

[0133] Step 4-3: Load the p parameter, L parameter, c parameter, α max parameter, and t s2 parameter of the preset time disturbance observer described in step 3;

[0134] Step 4-4: Generate a guidance command according to the parameters loaded in steps 4-1, 4-2, and 4-3 and the real-time aircraft state, and guide the aircraft to intercept the maneuvering target at a desired terminal angle.

[0135] Furthermore, based on the relationship between the desired line-of-sight angular rate profile designed in step 1 and the maximum relative lead angle analyzed in step 5, it can be ensured that the field-of-view constraint is always satisfied. The specific steps are as follows:

[0136] Step 5-1: According to the kinematic equation of the aircraft, it can be obtained that:

[0137]

[0138] where

[0139] Also, since the design objective of the present invention is to make the real-time rotation angular rate of the aircraft track the designed line-of-sight angular rate profile, that is

[0140]

[0141] By combining formulas (41) and (42), σ

[0142] σR = atan(ke) (76)

[0143] where atan() is the arctangent function;

[0144] Step 5-2: According to Step 5-1, the maximum relative lead angle of the missile can be obtained as:

[0145] σ R.max = atan(ke0) (77)

[0146] where e0 is the initial strike angle error of the missile;

[0147] Step 5-3: Based on the relationship between the maximum relative lead angle derived in Step 5-2, the initial strike angle error, and the parameter k, adjusting the parameters e0 and k can achieve the constraint of the maximum relative lead angle.

[0148] Furthermore, based on the relationship between the expected line-of-sight angular rate profile designed in Step 1 and the maximum relative lead angle analyzed in Step 5, it can be ensured that the field-of-view constraint is always satisfied. Specifically, the lead angle constraint of the aircraft is converted into a relative lead angle constraint, and the specific equation relationship is:

[0149]

[0150] where is the maximum lead angle of the missile, and asin() is the arcsine function. Therefore, the problem of constraining the lead angle can be converted into the problem of constraining the relative lead angle.

[0151] A three-dimensional guidance system for terminal angle and field-of-view constraints for intercepting a maneuvering target, comprising: an aircraft and target real-time position calculation unit, an aircraft real-time guidance law calculation unit, a target disturbance real-time disturbance estimation unit, and an aircraft guidance and control unit, where:

[0152] The aircraft and target real-time position calculation unit is used to perform real-time iterative calculations of the aircraft and target positions according to the real-time aircraft controller and the target maneuver amount;

[0153] The aircraft real-time guidance law calculation unit is used to calculate the real-time control amount of the aircraft according to the designed terminal angle and field-of-view constraint guidance law;

[0154] The target disturbance real-time estimation unit is used to perform real-time target disturbance estimation according to the designed preset time disturbance observer to compensate for the error caused by target maneuvering.

[0155] The aircraft guidance and control unit is used to calculate the aircraft control quantity in real-time iteration according to the terminal angle and the field of view constraint guidance law and the preset time disturbance observer, drive the servo system of the aircraft to respond, realize flight control, and guide the aircraft to intercept a maneuvering target at the desired terminal angle.

[0156] Then, taking an anti-ship missile intercepting a ship as an example, the initial position of the missile is (-13.38 km, 7.726 km, 4.142 km), the initial velocity is 300 m / s, the initial ballistic inclination angle is 5.4768°, the initial ballistic deflection angle is -15.858°, the initial position of the warship is (0 km, 0 km, 0 km), the initial velocity is 15 m / s, the initial ballistic inclination angle is 0°, the initial ballistic deflection angle is 0°, the designed desired terminal line-of-sight inclination angle is -70°, the line-of-sight deflection angle is 0°, the warship performs an S-shaped maneuver, the acceleration is [0, sin(t / 11), 0], the maximum lead angle is set to 60°, and according to the equation relationship between the maximum lead angle and the maximum relative lead angle proposed by the present invention, the maximum relative lead angle can be calculated to be 51°. To ensure the field-of-view constraint, the parameter k = 1.2, and the parameters of the guidance law and the preset time disturbance observer are respectively: p = 4, t s1 = 11, L = [1, 1, 1] T , t s2 = 10, c = 4, α max = 10, p0 = [-0.0021, -0.0012, -0.004] T , this embodiment only provides a set of simulation results. For simulations in other scenarios, only the relevant parameters need to be changed accordingly. Figures 5-9 It is the simulation result of single-missile guidance.

[0157] As Figure 4 shows the motion trajectories of the anti-ship missile and the warship (the three-dimensional guidance law for the terminal angle and field-of-view constraint of intercepting a maneuvering target and the three-dimensional trajectories of the aircraft and the target of the system). Figure 5 shows the overload curve of the missile. Figure 6 shows the trajectory of the missile-target distance. Figure 7 shows the variation curves of the line-of-sight inclination angle, the line-of-sight deflection angle, and the terminal strike angle error. Figure 8 shows the variation curves of the relative lead angle and the lead angle. Figure 9 shows the line-of-sight angular rate tracking error. Figure 10 shows the observed value of the preset time disturbance observer. It can be seen from Figure 7 , Figure 8 , Figure 9 , Figure 10 that the designed guidance law can be at the desired time t s1The designed line of sight angular rate can be accurately tracked within s, and the designed preset time disturbance observer can accurately track the designed line of sight angular rate within the preset time t s2 The disturbance caused by the target maneuver is observed internally, and the maneuvering target can be intercepted at the desired terminal angle, and the constraint that the field of view angle is less than 60° is always met. The above simulation results demonstrate the advancement and effectiveness of the guidance law proposed in the present invention.

[0158] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device.

[0159] Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and variations may be made to the embodiments without departing from the principles and spirit of the present invention, and that the scope of the present invention is defined by the appended claims and their equivalents.

Claims

1. A three-dimensional guidance law with terminal angle and field-of-view constraints for intercepting maneuvering targets, characterized in that, It includes the following steps: Step 1: Based on the vector relative guidance coordinate system, design and implement the line-of-sight angular rate profile that realizes the terminal angle and field-of-view constraints; Step 2: Design a preset-time guidance law so that the aircraft can track the desired line-of-sight angular rate profile designed in Step 1 within the preset time; Step 3: Based on the preset-time control technology, design a preset-time disturbance observer to observe the disturbance caused by the target maneuver within the preset time; Step 4: Based on the desired line-of-sight angular rate profile in Step 1, the preset-time guidance law in Step 2, and the preset-time disturbance observer in Step 3, it is possible to intercept the maneuvering target at the desired terminal angle; Step 5: By analyzing the relationship between the desired line-of-sight angular rate profile designed in Step 1 and the maximum relative lead angle, ensure that the field-of-view constraint is always satisfied.

2. The three-dimensional guidance law for intercepting a maneuvering target with terminal angle and field-of-view constraints according to claim 1, characterized in that: The three-dimensional strike guidance law in step 1 is designed in the relative vector guidance coordinate system. The specific engagement model can be described as follows: First, the construction of the guidance law described in step 1 is completed in the relative vector guidance coordinate system. In this combat scenario, the motion state of the missile can be represented by the velocity vector V M and the control acceleration vector A M , while the motion state of the target is described by the velocity vector V T and the control acceleration vector A T . Define the line-of-sight vector between the missile and the target as R, where R = ||R|| is the missile-to-target distance. Let V M = ||V M || and V T = ||V T || be the motion speeds of the missile and the target respectively, and let and be unit basis vectors. In the system description, the relative velocity vector V R = V M - V T and its unit vector as well as the relative acceleration vector A R = A M - A T are introduced, where A R is orthogonal to V R . Based on the above, the following basic relationship of differentiation with respect to time is established: Ω in the formula L and Ω VR respectively represent the angular velocities of rotation of the line-of-sight vector and the relative velocity vector. Define the lead angle as σ M =acos(r·v M ), where acos() is the inverse cosine function. At the same time, define the relative lead angle σ R =acos(r·v R ). Taking the derivative of it, we get: Among them, In addition, by taking the derivative of the angular velocity of the line-of-sight vector rotation, we obtain: Where D represents the disturbance vector generated by the target maneuver characteristics. In addition, the derivative of the missile-to-target distance R with respect to time is: where, V R = ||V R || is the magnitude of the vector V R .

3. The three-dimensional guidance law for terminal angle and field-of-view constraints of intercepting a maneuvering target according to claim 2, characterized in that: The step 1 of designing the line-of-sight angular rate profile that can realize the terminal angle and field-of-view constraints based on the vector relative guidance coordinate system, and the specific desired line-of-sight angular rate profile is: where k > 0, t go = t f - t, t f is the terminal strike time, t is the actual flight time of the missile, e is the terminal angle error, r d is the unit vector in the desired terminal angle direction.

4. A three-dimensional guidance law for terminal angle and field-of-view constraints for intercepting maneuvering targets according to claim 3, characterized in that: The step 2 of designing the preset-time guidance law so that the aircraft can track the desired line-of-sight angular rate profile designed in Step 1 within the preset time, and the specific guidance law design is: where \(S = \Omega\) L -\(\Omega\) Ld , is the disturbance observation value brought by the target maneuver, \(S_0\) is the initial value of \(S\), \(t\) s1 is the convergence time of the tracking error vector \(S\) required to be designed, and \(0 \lt p \lt 1\) is a constant.

5. A three-dimensional guidance law for terminal angle and field of view constraints of intercepting a maneuvering target according to claim 4, characterized in that: The step 3 of designing the preset-time disturbance observer based on the preset-time control technology to observe the disturbance caused by the target maneuver within the preset time, and the specific preset-time disturbance observer is: Where wherein, it is assumed that the disturbance caused by the target maneuver satisfies ||D||≤D max , the derivative of the disturbance is bounded and satisfies λ min (·) is the minimum eigenvalue of the matrix, p is the initial value of the estimated disturbance, L is the matrix to be designed, c is the parameter to be designed, and it is required to satisfy c≥(δ + ||L||D max ) / (λ min (L)η0), where η0 is the initial error of the estimated disturbance, t s2 is the convergence time to be designed for the disturbance observer to be designed, α max >0 is the parameter to be designed.

6. The three-dimensional guidance law for intercepting a maneuvering target with terminal angle and field-of-view constraints according to claim 5, wherein: The step 5 of analyzing the relationship between the desired line-of-sight angular rate profile designed in Step 1 and the maximum relative lead angle, and the specific relationship between the maximum relative lead angle and the designed line-of-sight angular rate profile is: σ R,max = atan(ke0) (11) where atan() is the arctangent function, e0 is the initial value of the impact angle error, and analyze σ R,max the relationship with the maximum feasible maximum relative lead angle, and then design the guidance parameter k and the feasible impact error angle region, so as to achieve the constraint of the relative lead angle.

7. A three-dimensional guidance law for terminal angle and field-of-view constraints of intercepting a maneuvering target according to claim 6, characterized in that: The step 5 of ensuring the constraint of the missile lead angle by constraining the relative lead angle, and then always locking the target. The relationship between the missile relative lead angle and the lead angle is as follows: Among them, is the maximum lead angle of the missile, and asin() is the arcsine function.

8. A three-dimensional guidance system for intercepting maneuvering targets with terminal angle and field-of-view constraints, according to any one of claims 1-7, characterized in that, It includes: An aircraft and target real-time position calculation unit, an aircraft real-time guidance law calculation unit, a target disturbance real-time disturbance estimation unit, and an aircraft guidance and control unit. Among them: The aircraft and target real-time position calculation unit is used to perform real-time iterative calculations of the aircraft and target positions according to the real-time aircraft controller and the target maneuver amount; The aircraft real-time guidance law calculation unit calculates the real-time control amount of the aircraft according to the designed terminal angle and field-of-view constraint guidance law; The target disturbance real-time estimation unit performs real-time target disturbance estimation according to the designed preset-time disturbance observer to compensate for the error caused by the target maneuver; The aircraft guidance and control unit iteratively calculates the aircraft control amount in real time according to the terminal angle and field-of-view constraint guidance law and the preset-time disturbance observer, drives the servo system of the aircraft to respond, realizes flight control, and guides the aircraft to intercept the maneuvering target at the desired terminal angle.

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