Bounded acceleration meeting time control guidance method based on sight angle forming

Abstract

The invention relates to the technical field of aircraft guidance, in particular to a bounded acceleration encounter time control guidance method based on line-of-sight angle shaping, which comprises the following steps of: firstly, establishing a relative motion model of an aircraft and a target, and constructing an expected preposed angle track based on a real-time relative distance between the aircraft and the target; a sliding mode controller based on a superhelix algorithm is adopted to track the trajectory, a bounded integral control mechanism with a saturation function and state switching logic is integrated, and it is ensured that a control instruction meets physical constraints and is resistant to integral saturation; estimating residual flight time through a discretization numerical algorithm, comparing the residual flight time with expected arrival time, and dynamically adjusting parameters of the front angle trajectory; for the maneuvering target, updating the guidance distance by predicting the position of the maneuvering target at the expected arrival moment; adopting a continuous smooth function to suppress buffeting; the method does not depend on time prediction, is high in precision and robustness, and enables the aircraft to still meet the double constraints of angle and time at the same time in a complex environment.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft guidance technology, in particular to a bounded acceleration encounter time control guidance method based on line-of-sight angle shaping. BACKGROUND

[0002] In the field of aircraft guidance, the proportional guidance law is the most widely used aircraft guidance method, which requires the rotation rate of the aircraft velocity vector to be proportional to the rotation rate of the aircraft-target line of sight. Its main features are simple implementation, low terminal overload and smooth trajectory, so it has been widely studied and applied in engineering practice. However, with the development of technology, new demands for control of the terminal impact angle and flight time of the aircraft have been put forward. Therefore, guidance laws with terminal angle constraints and terminal time constraints have received extensive attention and in-depth study.

[0003] Existing guidance laws that meet time and angle constraints usually rely on time prediction. For example, a high-speed aircraft time coordination guidance method and system based on sliding mode control is disclosed in Chinese Patent Publication No. CN119847176A, which quickly predicts the remaining flight time of multiple aircrafts through a centralized strategy and corrects the time deviation according to the "who follows who" principle. The bottom layer uses a sliding mode time controllable guidance law, which introduces a saturation function to suppress chattering and avoid singularities. Combined with three-dimensional biased proportional guidance, the lateral trajectory is curved to adjust the flight time, and variable step integration is used to improve the accuracy and efficiency of the remaining time calculation. Chinese Patent Publication No. CN119644850A discloses a distributed time coordination guidance method for cruise missiles with limited control inputs. The remaining flight time estimate is used as the coordination variable, and the remaining flight time estimates of each missile are exchanged through inter-missile communication. The expected coordination variable value of each missile is calculated by the coordination strategy, which is used as the input of the lower-level time constraint guidance law. The lower-level time constraint guidance law is decomposed into the horizontal and lateral planes, and the attitude angle and throttle are calculated by the segmented guidance law as the guidance control input to the autopilot. According to the input attitude angle and throttle command, the rudder and throttle quantities are obtained to control the cruise missile attitude and speed to achieve time coordination. However, in actual dynamic environments, time prediction is easily affected by target motion state changes, environmental disturbances and system uncertainties, making it difficult to guarantee prediction accuracy and further improving guidance performance. SUMMARY

[0004] To overcome the shortcomings of the prior art, the present application proposes a bounded acceleration encounter time control guidance method based on line-of-sight angle shaping. This method does not rely on time prediction, has high accuracy and strong robustness, and enables the aircraft to meet both angle and time constraints in complex adversarial environments.

[0005] To achieve the above-mentioned purposes, the present application adopts the following technical solutions:

[0006] This invention proposes a bounded acceleration encounter time control guidance method based on line-of-sight angle shaping, comprising the following steps:

[0007] S1. Establish a two-dimensional relative motion model between the aircraft and the target;

[0008] S2. Based on the two-dimensional relative motion model, the real-time relative distance between the aircraft and the target is used as the independent variable, and the desired leading angle trajectory is constructed using a hyperbolic tangent cubic polynomial function.

[0009] S3. Based on the expected lead angle trajectory and the actual lead angle of the aircraft, a tracking error is constructed. The tracking error is processed by a sliding mode controller based on the super spiral algorithm to generate the expected normal jerk of the aircraft.

[0010] S4. A bounded integral control mechanism is integrated into the sliding mode controller to perform anti-integral saturation and amplitude limiting processing on the desired normal acceleration, thereby generating normal acceleration.

[0011] S5. Estimate the remaining flight time of the aircraft in real time based on a discretized numerical algorithm, and dynamically adjust the parameters of the desired leading angle trajectory according to the deviation between the remaining flight time and the desired arrival time.

[0012] Furthermore, the method also includes:

[0013] S6. Predict the target's position at the expected arrival time based on the target's current motion state, and update the real-time relative distance with the predicted position;

[0014] S7. Replace the sign function in the sliding mode controller with a continuously differentiable smooth function.

[0015] Furthermore, S1 specifically includes:

[0016] Establish the geometric relationship between the aircraft and the target in a two-dimensional plane, and define... Forward angle, As the line of view, Let be the track angle. The expression for the two-dimensional relative motion model is:

[0017]

[0018] In the formula, for Time derivative, The normal acceleration of the aircraft. For the speed of the aircraft, This represents the relative distance between the aircraft and the target.

[0019] Furthermore, the desired leading angle trajectory is expressed as follows:

[0020]

[0021]

[0022] In the formula, For the desired leading angle, This refers to the real-time relative distance between the aircraft and the target. The initial relative distance between the spacecraft and the target is a parameter. Here, k represents the intermediate calculation coefficients, and k is the dynamically adjusted parameter. This is the actual initial leading angle;

[0023] The desired leading angle trajectory satisfies the following boundary conditions:

[0024] hour, ;

[0025] hour, ; .

[0026] Furthermore, in S3, the formula for calculating the tracking error e is:

[0027]

[0028] The sliding mode controller processes tracking errors as follows:

[0029] S301, Based on the tracking error e and its derivative Construct the sliding surface function s:

[0030]

[0031] In the formula, For design parameters, ;

[0032] S302. Differentiate the sliding surface function s to obtain the sliding surface rate of change.

[0033] S303, Set the rate of change of the sliding surface. To obtain the equivalent control term Based on equivalent control items and switching control items Generate the desired normal accelerometer. :

[0034]

[0035]

[0036] In the formula, To switch gain, For integral gain, >0, where τ is the integral variable over time.

[0037] Furthermore, in S4, the bounded integral control mechanism is specifically as follows:

[0038] S401, based on the maximum normal acceleration of the aircraft Construct saturation boundary curves, the set of which is:

[0039]

[0040] Where w1 is the x-coordinate of a point on the saturation boundary curve, and w2 is the y-coordinate of a point on the saturation boundary curve;

[0041] The saturation boundary curve The expression is:

[0042]

[0043] In the formula, m≥1;

[0044] S402. Based on the set of saturation boundary curves, a composite control law is formed by introducing a saturation function and state switching logic into the control law of the sliding mode controller. The composite control law is expressed as follows:

[0045] A M = ω 1 [ ω ˙ 1 ω ˙ 2 ] = [ − k u | ε | 1 2 s i g n ( ε ) α A ˙ d ω 2 2 m − 1 − α A ˙ d ω 2 ω 1 2 m − 2 A max 2 m − k u | ε | 1 2 s i g n ( ε ) ] [ ω 1 ω 2 ]

[0046] A M = ω 1 [ ω ˙ 1 ω ˙ 2 ] = [ − k u | ε | 1 2 s i g n ( ε ) α A ˙ d ω 2 2 m − 1 − α A ˙ d ω 2 ω 1 2 m − 2 A max 2 m − k u | ε | 1 2 s i g n ( ε ) ] [ ω 1 ω 2 ]

[0047]

[0048] In the formula, For proportional gain, >0, where α is the integral state weighting coefficient, α>0. This is the maximum normal acceleration of the aircraft;

[0049] S403. Solve the composite control law to obtain the normal acceleration.

[0050] Furthermore, in S5, the method for dynamically adjusting the parameters of the preceding angle trajectory is as follows:

[0051] S501. Based on the current aircraft state and the current value of k, a discretized numerical algorithm is used to simulate the aircraft's motion along the desired forward angular trajectory and estimate the total flight time. ;

[0052] S502, Calculated and estimated total flight time Compared with the expected total flight time deviation :

[0053]

[0054] like Output the value of parameter k; The first preset threshold;

[0055] like According to the first preset adjustment amount After updating parameter k, enter S503;

[0056] S503, Repeat S501 to S502 until... Output the value of parameter k;

[0057] S504. Using the value of parameter k output from S502 or S503 as the initial value, a discretized numerical algorithm is used to simulate the aircraft's forward motion along the desired leading angle trajectory, and the remaining flight time is estimated. ;

[0058] S505, Calculate and estimate the remaining flight time Deviation from expected arrival time :

[0059]

[0060] In the formula, Current flight time;

[0061] like Output the value of parameter k; The second preset threshold;

[0062] like According to the first preset adjustment amount After updating parameter k, proceed to S506;

[0063] S506, Repeat S504 to S505 until... Output the value of parameter k.

[0064] Furthermore, the continuously differentiable smooth function is an approximate function or a Sigmoid function.

[0065] The present invention also proposes an aircraft cooperative guidance system, including a memory and a processor. The memory stores a computer program, and when the computer program is executed by the processor, the above-mentioned guidance method is implemented.

[0066] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0067] (1) This invention constructs a predetermined flight path in the range domain by designing a hyperbolic tangent cubic polynomial desired lead angle trajectory based on relative distance, thus avoiding direct dependence on flight time and improving the initial planning robustness of the guidance system. Furthermore, a super-spiral sliding mode controller is used to track the desired trajectory, and combined with a bounded integral control mechanism, the aircraft can accurately converge to the predetermined trajectory within a finite time, while ensuring that all control commands satisfy the physical acceleration constraints of the aircraft, thereby ensuring the physical realizability and stability of the system while achieving high-precision tracking.

[0068] (2) The present invention can evaluate and correct flight time deviation in real time through an online time estimation and trajectory parameter dynamic adjustment mechanism based on discretized numerical algorithm, thereby accurately controlling multiple aircraft to reach the target in coordination. At the same time, for maneuvering targets, the present invention automatically compensates for the maneuverability of the target by extrapolating the current position of the target to the expected arrival time and updating the relative distance in the guidance, thereby significantly improving the accuracy of reaching maneuvering targets. Attached Figure Description

[0069] Figure 1 This is a schematic diagram illustrating the framework of the method of the present invention;

[0070] Figure 2 This is a schematic diagram of the relative motion between the aircraft and the target in an example of the present invention;

[0071] Figure 3 Different in the examples of the present invention The expected leading angle trajectory of the submersible;

[0072] Figure 4 This is a schematic diagram of the bounded integral control principle in an example of the present invention;

[0073] Figure 5 The simulation results curves for the three aircraft working together to reach a stationary target in the example of this invention are: (a) the trajectory of the aircraft and the target, (b) the curve of the aircraft's leading angle, (c) the curve of the aircraft's normal acceleration, and (d) the curve of the relative distance between the aircraft and the target.

[0074] Figure 6The simulation results curves for the three aircraft working together to reach the maneuvering target in the example of this invention are: (a) the trajectory of the aircraft and the target, (b) the curve of the aircraft's leading angle, (c) the curve of the aircraft's normal acceleration, and (d) the curve of the relative distance between the aircraft and the target.

[0075] Figure 7 The following are terminal miss distance distribution diagrams of three aircraft in Monte Carlo simulations in this invention example: (a) aircraft 1, (b) aircraft 2, and (c) aircraft 3. Detailed Implementation

[0076] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0077] Example

[0078] refer to Figure 1 This embodiment proposes a bounded acceleration encounter time control guidance method based on line-of-sight shaping, including the following steps:

[0079] S1. Establish a two-dimensional relative motion model between the aircraft and the target;

[0080] like Figure 2 As shown, the geometric relationship between the aircraft M and the target T is established in a two-dimensional plane, and the leading angle of the aircraft is defined as... That is, the line of sight between the velocity vector and the angle, and the line of view is... track angle is All angles are positive in the counter-clockwise direction, and the normal acceleration of the aircraft is... The speed is , Indicates the location of the aircraft. Indicates the target's location and the relative distance between the aircraft and the target. ,

[0081]

[0082] The pre-angle can be obtained from the established geometric relationships. The dynamic equations, i.e., the two-dimensional relative motion model between the aircraft and the target, are expressed as follows:

[0083]

[0084] In the formula, for The time derivative.

[0085] In this step, the two-dimensional relative motion model describes the relationship between aircraft maneuvers, target motion, and changes in lead angle.

[0086] S2. To achieve terminal angle constraints and shape the flight trajectory, this step uses the real-time relative distance between the aircraft and the target. As the independent variable, a hyperbolic tangent cubic polynomial function is used to construct the expected leading angle trajectory, which is expressed in the following form:

[0087]

[0088]

[0089] In the formula, For the desired leading angle, This refers to the real-time relative distance between the aircraft and the target. The initial relative distance between the spacecraft and the target is a parameter. These are intermediate calculation coefficients determined by k, where k is a dynamically adjusted parameter. This is the actual initial leading angle;

[0090] refer to Figure 3 It can be seen that by adjusting the parameter k, the flight trajectory can be changed, thereby adjusting the flight time. To ensure the physical feasibility of the trajectory, the terminal arrival conditions, and the stability of the aircraft before reaching the target, the expression of the lead angle trajectory should satisfy the following boundary conditions:

[0091] hour, ;

[0092] hour, ; ;

[0093]

[0094] Solving for: , .

[0095] S3. Based on the desired lead angle trajectory and the real-time lead angle of the aircraft... Construct tracking error : ;

[0096] The tracking error is processed using a sliding mode controller based on a superhelical algorithm to generate the desired normal jerk of the aircraft. The specific process is as follows:

[0097] S301, Based on the tracking error e and its derivative Construct the sliding surface function s:

[0098]

[0099] In the formula, For design parameters, ;

[0100] S302. Differentiate the sliding surface function s to obtain the sliding surface rate of change. :

[0101]

[0102] S303, Set the rate of change of the sliding surface. To obtain the equivalent control term :

[0103]

[0104] Based on equivalent control items and switching control items Generate the desired normal accelerometer. :

[0105]

[0106]

[0107] In the formula, To switch gain, For integral gain, >0, where τ is the integral variable over time.

[0108] S4, Reference Figure 4 The sliding mode controller integrates a bounded integral control mechanism to perform anti-integral saturation and amplitude limiting processing on the desired normal acceleration, generating the final normal acceleration. The bounded integral control mechanism includes:

[0109] S401, based on the maximum normal acceleration of the aircraft Construct saturation boundary curves, the set of which is:

[0110]

[0111] Where w1 is the x-coordinate of a point on the saturation boundary curve, and w2 is the y-coordinate of a point on the saturation boundary curve;

[0112] The saturation boundary curve The expression is:

[0113]

[0114] In the formula, , m≥1.

[0115] S402. Based on the set of saturation boundary curves, a composite control law is formed by introducing a saturation function and state switching logic into the control law of the sliding mode controller. The composite control law is expressed as follows:

[0116] A M = ω 1 [ ω ˙ 1 ω ˙ 2 ] = [ − k u | ε | 1 2 s i g n ( ε ) α A ˙ d ω 2 2 m − 1 − α A ˙ d ω 2 ω 1 2 m − 2 A max 2 m − k u | ε | 1 2 s i g n ( ε ) ] [ ω 1 ω 2 ]

[0117] A M = ω 1 [ ω ˙ 1 ω ˙ 2 ] = [ − k u | ε | 1 2 s i g n ( ε ) α A ˙ d ω 2 2 m − 1 − α A ˙ d ω 2 ω 1 2 m − 2 A max 2 m − k u | ε | 1 2 s i g n ( ε ) ] [ ω 1 ω 2 ]

[0118]

[0119] In the formula, For proportional gain, >0, where α is the integral state weighting coefficient, α>0. This is the maximum normal acceleration of the aircraft;

[0120] S403. Solve the composite control law to obtain the normal acceleration:

[0121] .

[0122] S5. Based on a discretized numerical algorithm, the remaining flight time of the aircraft is estimated in real time, and the parameters of the leading angular trajectory are dynamically adjusted according to the deviation between the remaining flight time and the expected arrival time. The specific process is as follows:

[0123] S501. Input the value of parameter k, and use a discretized numerical algorithm to simulate the aircraft's motion along the desired forward angular trajectory to estimate the total flight time. ;

[0124] S502, Calculated and estimated total flight time Compared with the expected total flight time deviation :

[0125]

[0126] like Output the value of parameter k; The first preset threshold;

[0127] like According to the first preset adjustment amount After updating parameter k, enter S503;

[0128] S503, Repeat S501 to S502 until... Output the value of parameter k;

[0129] S504. Using the value of parameter k output from S502 or S503 as input, a discretized numerical algorithm is employed to simulate the aircraft's forward motion along the desired leading angle trajectory and estimate the remaining flight time. ;

[0130] S505, Calculate and estimate the remaining flight time Deviation from expected arrival time :

[0131]

[0132] In the formula, Current flight time;

[0133] like Output the value of parameter k; The second preset threshold;

[0134] like According to the first preset adjustment amount After updating parameter k, proceed to S506;

[0135] S506, Repeat S504 to S505 until... Output the value of parameter k.

[0136] For example, in this embodiment, , , As the simulation step size per unit time, this embodiment estimates the total flight time according to the following algorithm 1. and remaining flight time :

[0137] Algorithm 1:

[0138] Input: The value of parameter k, the initial lead angle aircraft speed Simulation step size per unit time The initial relative distance between the aircraft and the target Maximum normal acceleration ;

[0139] Output: Flight time t;

[0140] Step 1: Initialization hour, , ;

[0141] Step 2, when At that time, the following loop is executed:

[0142] Calculate the desired aircraft lead angle ;

[0143] Calculate the tracking error e;

[0144] Calculate the smooth film function s;

[0145] Calculate normal acceleration ;

[0146] Update the aircraft's trajectory:

[0147]

[0148]

[0149]

[0150] End the loop

[0151] Step 3: Output t.

[0152] When Algorithm 2 calls Algorithm 1, the output t is: When Algorithm 3 calls Algorithm 1, the output t is ;

[0153] The parameter k is selected according to Algorithm 2, that is, Algorithm 2 is used to find the corresponding parameter k. of .

[0154] Algorithm 2:

[0155] Input: Initial leading angle aircraft speed Simulation step size per unit time =0.001, the initial relative distance between the aircraft and the target. Current flight time Expected total flight time ;

[0156] Output: parameter k

[0157] Step 1: Initialization ; Let k be the initial value;

[0158] Step 2, when Execute the following loop:

[0159] Calculate the flight time corresponding to the current k using Algorithm 1. ,

[0160] , ;

[0161] ,

[0162] End of judgment;

[0163] End the loop;

[0164] Step 3: Output k.

[0165] The output value of k in this algorithm is... corresponding ;

[0166] The parameter k is adjusted online using the following algorithm 3. That is, during the flight, the remaining distance is constantly checked according to the current trajectory and compared with the expected arrival time (planned remaining time), and fine adjustments are made at any time.

[0167] Algorithm 3:

[0168] Input: Initial leading angle aircraft speed The value of parameter k Simulation step size per unit time =0.001, the real-time relative distance between the aircraft and the target. Expected total flight time Current flight time ;

[0169] Output: parameter k;

[0170] Step 1, when Execute the following loop:

[0171] Step 2: Calculate the remaining flight time corresponding to the current k using Algorithm 1. ,

[0172] , ;

[0173] ,

[0174] End of judgment;

[0175] End the loop;

[0176] Step 2: Output k.

[0177] Update using the k value output by this algorithm .

[0178] S6. If the target is a maneuvering target, in order to compensate for the maneuvering target's motion, in each guidance cycle, its position at the expected arrival time is predicted based on the target's current motion state, and the real-time relative distance is updated with the predicted position; assuming the target maintains uniform linear motion after the current moment, the target's position P at the expected arrival time is:

[0179]

[0180] The target's current location.

[0181] The relative distance between the aircraft and position P is used as the updated real-time relative distance and input into the desired leading angle trajectory in S2.

[0182] S7. To suppress the chattering effect of the sign function in the sliding mode controller, a continuously differentiable smooth function is used to replace the sign function in the sliding mode controller for the excited target. The continuously differentiable smooth function is either an approximation function or a Sigmoid function. In this embodiment, the Sigmoid function is selected.

[0183]

[0184] In the formula, Replace the symbolic functions in S3 and S4 with this function.

[0185] To eliminate tracking error when it approaches 0, The resulting singularity problem, for Make corrections:

[0186]

[0187] In the formula: ƍ is a very small constant value, ƍ=0.0001.

[0188] To verify the effectiveness of the present invention, a systematic numerical simulation was performed according to the implementation process of the above embodiments. The parameters of the simulation platform were set as follows: , =3.5Mach, =5Mach, =0.001s, =6g, α=0.9, , , stationary targets =32.366s, maneuvering target =30.199s. For stationary targets, the initial parameters of the aircraft are shown in Table 1. For maneuvering targets, the initial parameters of the aircraft are shown in Table 2.

[0189] surface Initial parameters for multiple aircraft to reach a stationary target

[0190]

[0191] Table 2 Initial parameters of multiple aircraft arriving at the maneuvering target

[0192]

[0193] During the simulation, the given expected total flight time is first determined. For each spacecraft, Algorithm 2 is executed to obtain its respective... During flight, each spacecraft independently runs Algorithm 3, adjusting parameter k in real time online.

[0194] Simulation results reference Figure 5 Figure 6 Tables 3 and 4 Figure 5 (a)-(d) show the curves of the trajectory, lead angle, normal acceleration and relative distance between the three aircraft and the stationary target when they arrive at the same expected time. Figure 6 (a)-(d) show the corresponding results of reaching the maneuvering target. From Figure 5 , Figure 6 Tables 3 and 4 show that the positions of the aircraft converge to the target positions, and the aircraft's leading angle, normal acceleration, and relative distance between the aircraft and the target all converge to near 0. This indicates that the guidance method proposed in this invention effectively controls multiple aircraft to accurately reach stationary or maneuvering targets.

[0195] Table 3. Error values ​​of distance and impact time between multiple aircraft and stationary targets

[0196]

[0197] Table 4. Distance and angular error values ​​of multiple aircraft towards maneuvering targets

[0198]

[0199] To test robustness, 500 Monte Carlo shooting simulations were conducted, with random perturbations added to the initial conditions based on those in Tables 1 and 2. The statistical results of the terminal miss distance are as follows: Figure 7 As shown in Table 5, Figure 7 (a)-(c) show the terminal miss distance distributions of the three aircraft in Monte Carlo simulations. Figure 7 As can be seen from Table 5, the results of 500 Monte Carlo simulations show that the terminal miss distances of the three aircraft are less than 0.88 meters, 0.58 meters, and 0.98 meters, respectively, which are less than the maximum relative displacement of 1.7 meters per simulation step, proving the robustness and effectiveness of the guidance method of the present invention.

[0200] Table 5 Monte Carlo simulation results

[0201]

[0202] In summary, the guidance method of this invention, through desired trajectory design, inner-loop tracking control, effective anti-saturation mechanism, and phased parameter adjustment, can accurately reach the target.

[0203] The specific embodiments of the present invention are provided to enable those skilled in the art to understand or implement the invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the invention.

[0204] It should be understood that the present invention is not limited to the content already described above, and various modifications and changes can be made without departing from its scope. The scope of the present invention is limited only by the appended claims.

Claims

1. A bounded acceleration encounter time control guidance method based on line-of-sight angle shaping, characterized in that, Includes the following steps: S1. Establish a two-dimensional relative motion model between the aircraft and the target; S2. Based on the two-dimensional relative motion model, the real-time relative distance between the aircraft and the target is used as the independent variable, and the desired leading angle trajectory is constructed using a hyperbolic tangent cubic polynomial function. S3. Based on the expected lead angle trajectory and the actual lead angle of the aircraft, a tracking error is constructed. The tracking error is processed by a sliding mode controller based on the super spiral algorithm to generate the expected normal jerk of the aircraft. S4. A bounded integral control mechanism is integrated into the sliding mode controller to perform anti-integral saturation and amplitude limiting processing on the desired normal acceleration, thereby generating normal acceleration. S5. Estimate the remaining flight time of the aircraft in real time based on a discretized numerical algorithm, and dynamically adjust the parameters of the desired leading angle trajectory according to the deviation between the remaining flight time and the desired arrival time.

2. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 1, characterized in that, The method also includes: S6. Predict the target's position at the expected arrival time based on the target's current motion state, and update the real-time relative distance with the predicted position; S7. Replace the sign function in the sliding mode controller with a continuously differentiable smooth function.

3. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 1, characterized in that, S1 specifically includes: Establish the geometric relationship between the aircraft and the target in a two-dimensional plane, and define... Forward angle, As the line of view, Let be the track angle. The expression for the two-dimensional relative motion model is: In the formula, The normal acceleration of the aircraft. For the speed of the aircraft, This represents the relative distance between the aircraft and the target.

4. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 1, characterized in that, The desired leading angle trajectory is expressed as follows: In the formula, For the desired leading angle, This refers to the real-time relative distance between the aircraft and the target. The initial relative distance between the spacecraft and the target is a parameter. Here, k represents the intermediate calculation coefficients, and k is the dynamically adjusted parameter. This is the actual initial leading angle; The desired leading angle trajectory satisfies the following boundary conditions: hour, ; hour, ; .

5. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 3, characterized in that, In S3, the formula for calculating the tracking error e is: The sliding mode controller processes tracking errors as follows: S301, Based on the tracking error e and its derivative Construct the sliding surface function s: In the formula, For design parameters, ; S302. Differentiate the sliding surface function s to obtain the sliding surface rate of change. S303, Set the rate of change of the sliding surface. To obtain the equivalent control term Based on equivalent control items and switching control items Generate the desired normal accelerometer. : In the formula, To switch gain, For integral gain, >0, where τ is the integral variable over time.

6. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 1, characterized in that, In S4, the bounded integral control mechanism is specifically as follows: S401, based on the maximum normal acceleration of the aircraft Construct saturation boundary curves, the set of which is: Where w1 is the x-coordinate of a point on the saturation boundary curve, and w2 is the y-coordinate of a point on the saturation boundary curve; The saturation boundary curve The expression is: In the formula, m≥1; S402. Based on the set of saturation boundary curves, a composite control law is formed by introducing a saturation function and state switching logic into the control law of the sliding mode controller. The composite control law is expressed as follows: In the formula, For proportional gain, >0, where α is the integral state weighting coefficient, α>0. This is the maximum normal acceleration of the aircraft; S403. Solve the composite control law to obtain the normal acceleration.

7. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 4, characterized in that, In S5, the method for dynamically adjusting the parameters of the desired leading angle trajectory is as follows: S501. Based on the current aircraft state and the current value of k, a discretized numerical algorithm is used to simulate the aircraft's motion along the desired forward angular trajectory and estimate the total flight time. ; S502, Calculated and estimated total flight time Compared with the expected total flight time deviation : like Output the value of parameter k; The first preset threshold; like According to the first preset adjustment amount After updating parameter k, enter S503; S503, Repeat S501 to S502 until... Output the value of parameter k; S504. Using the value of parameter k output from S502 or S503 as the initial value, a discretized numerical algorithm is used to simulate the aircraft's forward motion along the desired leading angle trajectory, and the remaining flight time is estimated. ; S505, Calculate and estimate the remaining flight time Deviation from expected arrival time : In the formula, Current flight time; like Output the value of parameter k; The second preset threshold; like According to the first preset adjustment amount After updating parameter k, proceed to S506; S506, Repeat S504 to S505 until... Output the value of parameter k.

8. The bounded acceleration encounter time control guidance method based on line-of-sight shaping according to claim 1, characterized in that, The continuously differentiable smooth function is an approximate function or a Sigmoid function.

9. A cooperative guidance system for aircraft, characterized in that, It includes a memory and a processor, the memory storing a computer program that, when executed by the processor, implements the guidance method as described in any one of claims 1 to 8.

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