Middle guidance method of patrol missile with space-time constraint

Abstract

The invention discloses a guidance method in a patrol missile with space-time constraint, and belongs to the field of guidance control. The implementation method comprises the following steps: constructing a mid-guidance problem, and planning a mid-guidance expected track point according to the mid-guidance problem by using computing resources of a launching platform before launching; a Douglas-Peucker algorithm is adopted to dilute the track points planned by the launching platform, and the extracted track points including the position, the expected incident angle and the arrival time information are bound to a guidance control module of the patrol ammunition; the incident angle constraint guidance law is used as a reference guidance law, and the remaining flight time of the patrol missile adopting the reference guidance law is estimated; designing an offset item for controlling the arrival time by using the residual flight time estimation value, and constructing an incident angle and arrival time constraint guidance law by using the reference guidance law and the offset item; according to the incident angle and arrival time constraint guidance law, the patrolling ammunition sequentially passes through the bound mid-guidance expected track points, and cruising section mid-guidance of the patrolling ammunition is achieved.

Description

Technical Field

[0001] The invention relates to a mid-range guidance method with time and space constraints, in particular to a guidance method with incident angle constraints and arrival time constraints applicable to cruise missiles, belonging to the field of guidance control. Background Art

[0002] The modern battlefield environment is becoming increasingly complex. High-value enemy targets usually have multi-layered defense systems, which poses a great challenge to the penetration of guided weapons. Relying on a single high-performance weapon can effectively penetrate, but the cost of research and development and use is high. How to improve the penetration capability of guided weapons at a lower cost has become a research hotspot in the field of guided weapons.

[0003] Cruise missiles are a new type of intelligent ammunition that can cruise in a designated mission area and perform combat missions such as hovering reconnaissance, target strikes, and effect evaluation. Compared with hypersonic missiles, cruise missiles are low-cost and efficient and flexible to use. The coordinated combat of cruise missile clusters can improve the penetration capability without significantly increasing the cost of guided weapons, which is of great significance.

[0004] The existing theoretical research and methods of collaborative guidance are mainly concentrated in the terminal guidance stage. Cruise missiles have a long range. In order to achieve the coordinated strike mission of multiple cruise missiles against the same target, coordinated terminal guidance is required after coordinated mid-range guidance. In the mid-range stage, the aircraft reaches the specified airspace within the expected time and meets certain incident angle constraints, which can provide favorable initial conditions for terminal guidance. The existing methods for implementing mid-range guidance mainly include numerical optimization methods such as pseudo-spectral method. In practical applications, there are problems such as large search space and slow solution speed. Due to the limitation of the computing power of the onboard computer, it is difficult to ensure the real-time performance of mid-range guidance.

[0005] Considering the above factors, it is very necessary to propose a mid-range guidance method for cruise missiles that can meet the constraints of incident angle and arrival time. Summary of the invention

[0006] In view of the shortcomings of the existing cruise missile mid-range guidance method, the main purpose of the present invention is to provide a mid-range guidance method with time and space constraints, taking the optimal control energy of the cruise missile as the optimization goal, and taking the no-fly zone detour requirement constraint, the minimum turning radius constraint, the incident angle and the arrival time constraint based on the cruise missile in the constant altitude and constant speed cruise segment as the constraint conditions, to construct a mid-range guidance problem, and before launching, the computing resources of the launch platform are used to plan the expected mid-range guidance track points of the cruise missile cruise segment according to the mid-range guidance problem; the Douglas-Peucker algorithm is used to thin the track points planned by the launch platform, and the extracted track points containing the position, the expected entry angle and the target position are extracted. The track points of the incident angle and arrival time information are bound to the guidance control module of the cruise missile; the incident angle constraint guidance law is used as the reference guidance law to estimate the remaining flight time of the cruise missile using the reference guidance law; the estimated value of the remaining flight time is used to design the bias term for controlling the arrival time, and the incident angle and arrival time constraint guidance law is constructed using the reference guidance law and the bias term; according to the incident angle and arrival time constraint guidance law, the cruise missile passes through the bound mid-guidance expected track points in sequence to avoid the no-fly zone and arrive at the mid-terminal guidance handover position at the expected time and incident angle, thereby realizing the mid-guidance of the cruise missile in the cruise segment. Since the mid-guidance expected track points tracked by the cruise missile according to the incident angle and arrival time constraint guidance law are obtained by utilizing the computing resources of the launch platform, the computing power requirements of the cruise missile onboard computer are reduced. The present invention has the advantages of taking into account the arrival time, incident angle constraints and the need to fly around the no-fly zone at the same time, and has low requirements for the computing power of the cruise missile onboard computer and high real-time performance of mid-guidance. The expected track point includes position, expected incident angle and arrival time information.

[0007] The present invention discloses a mid-range guidance method for a cruise missile with time and space constraints, comprising the following steps:

[0008] Step 1: Taking the optimal control energy of the cruise missile as the optimization goal, and taking the no-fly zone detour requirement, minimum turning radius constraint, incident angle and arrival time constraint of the cruise missile in the constant altitude and constant speed cruise phase as the constraint conditions, the mid-range guidance problem is constructed. The computational resources of the launch platform are used to solve the mid-range guidance problem with numerical methods to obtain the expected mid-range guidance track points that meet the constraints.

[0009] Taking the optimal control energy of the cruise missile as the optimization goal, and taking the no-fly zone flight requirement constraint, minimum turning radius constraint, incident angle and arrival time constraint based on the cruise missile in the constant altitude and constant speed cruise phase as the constraint conditions, the mid-range guidance problem is constructed as shown in formula (1):

[0010]

[0011] Where J represents the objective function, u represents the virtual control quantity and satisfies |u|≤1, t f represents the expected time to reach the target point, x M represents the x-axis coordinate of the cruise missile in the inertial system, yM represents the coordinate of the cruise missile on the y-axis of the inertial system, V M represents the flight speed of the patrol vehicle, γ M Represents the ballistic deviation angle of the cruise missile, R min Indicates the minimum turning radius of the patrol vehicle, x o,i represents the coordinate of the center of the i-th no-fly zone on the x-axis of the inertial system, y o,i represents the coordinate of the center of the i-th no-fly zone on the y-axis of the inertial system, R i represents the distance from the center of the i-th no-fly zone, x d and d represents the target point position, γ d represents the expected angle of incidence to the target point, x0 and y0 represent the initial position of the cruise missile, and γ0 represents the initial ballistic deviation angle of the cruise missile.

[0012] The mid-range guidance problem shown in equation (1) is solved by numerical method to obtain the mid-range guidance expected track point that meets the no-fly zone detour requirement, minimum turning radius constraint, incident angle and arrival time constraint. The expected track point includes position, expected incident angle and arrival time information.

[0013] Step 2: Use the Douglas-Peucker algorithm to thin out the expected mid-range guidance track points planned in step 1, and inject the extracted track points including the expected position, incident angle and arrival time information into the guidance and control module of the cruise missile. By thinning out the expected mid-range guidance track points, the number of expected track points that need to be tracked is reduced. The incident angle constraint guidance law is used as the reference guidance law, and the analytical remaining flight time prediction formula is derived based on the reference guidance law. Using the position of the cruise missile and the current track point, the speed of the cruise missile and the ballistic deviation angle as input, the remaining flight time t is determined according to the remaining flight time prediction formula. go,j .

[0014] Step 2.1: Use the Douglas-Peucker algorithm to thin out the expected track planned by the launch platform, and obtain a total of M track points. The information of each track point includes the expected position, incident angle and arrival time.

[0015] Step 2.2: Establish the relative motion equation between the cruise missile and the jth (j=2,3,...M) track point, design a reference guidance law with the ability to control the angle of incidence, and predict the remaining flight time for the cruise missile to reach the jth track point under the action of the reference guidance law.

[0016] The relative motion equation between the cruise missile and the jth track point is expressed as

[0017]

[0018]

[0019]

[0020]

[0021] Among them, r j represents the relative distance between the cruise missile and the jth track point, σ j represents the lead angle of the patrol vehicle, λ j represents the sight angle, a M Indicates the normal acceleration command of the aircraft.

[0022] The normal acceleration command of the reference guidance law is

[0023]

[0024] Where N represents a positive navigation ratio, γ d,j represents the expected angle of incidence of the cruise missile when it reaches the jth track point, represents the predicted value of the incident angle of the cruise missile arriving at the jth track point and satisfies

[0025]

[0026] Under the reference guidance law, the remaining flight time for the cruise missile to reach the jth track point is

[0027]

[0028] According to formula (8), the remaining flight time for the cruise missile to reach the jth track point under the reference guidance law is predicted.

[0029] Step 3: Expected time t to reach the jth track point j and the remaining flight time t obtained in step 2 go,j The difference between the current time and j Defined as the arrival time control error. Design the expected change rate based on the initial value of the arrival time control error and the expected convergence time Construct a bias term to control the arrival time error δ of the cruise missile j Follow the designed expected rate of change The reference guidance law and the bias term are used to construct the guidance law constrained by the incident angle and arrival time. The guidance law constrained by the incident angle and arrival time ensures that the arrival time control error δ j Convergence is achieved before the cruise missile reaches the jth track point, so that the cruise missile arrives at the jth track point at the expected incident angle and expected time, that is, the guidance of the cruise missile with incident angle and arrival time constraints is realized.

[0030] Step 3.1: Determine the arrival time control error δ jThe expected rate of change

[0031] Arrival time control error δ j represents the expected arrival time t j and the remaining time prediction value t go,j and the difference between the current time. When controlling δ j When it converges to zero, the cruise missile can reach the jth track point at the expected time. The arrival time control error δ j The definition of

[0032]

[0033] Since the flight time from the j1th track point to the jth track point is limited, it is necessary to ensure that the arrival time control error δ j It is very important that the cruise missile converges to zero before it reaches the jth track point. At the same time, it is necessary to ensure that the distance between the cruise missile and the jth track point decreases monotonically during the guidance process to ensure that the lead angle σ j Less than 90°. In order to ensure the arrival time control error δ j Converges to zero before the cruise missile reaches the jth track point, constructing the arrival time control error δ j The expected rate of change is as follows:

[0034]

[0035] Among them, sig p (·)=sign(·)|·| p , sign() represents the sign function, 0<p<1, δ j,0 Denotes δ j The initial value of t s is a constant value.

[0036] Step 3.2: Determine the guidance method bias term so that the arrival time control error δ j Following the expected rate of change in equation (10), the guidance of the rover with incident angle and arrival time constraints is achieved.

[0037] Design the following guidance law with incident angle and arrival time constraints

[0038] a M =a B +a T (11)

[0039] Among them, a B represents the reference guidance law, which is used to ensure zero miss and desired angle of incidence; a T Represents the bias term, which is used to ensure the time control error δ jThe expected rate of change in equation (10) is satisfied, thereby satisfying the arrival time constraint.

[0040] From equations (4), (8) and (9), we get

[0041]

[0042] in,

[0043]

[0044] According to equations (10) and (12), the bias term is

[0045]

[0046] Combining equations (6), (11) and (14), we can obtain the normal acceleration command with incident angle and arrival time constraints, so that the arrival time control error of the cruise missile satisfies the expected change rate in equation (10), ensuring that the arrival time control error δ j Convergence is achieved before the cruise missile reaches the jth track point, so that the cruise missile reaches the jth track point at the expected incident angle and expected time.

[0047] Step 4: In step 1, the computational resources of the launch platform are used to solve the mid-range guidance problem using numerical methods to obtain the expected mid-range guidance track points that meet the constraints. Step 2: Reduce the number of expected track points that need to be tracked by thinning the expected mid-range guidance track points obtained in step 1, and inject the track points obtained after thinning, including the expected position, incident angle, and arrival time information, into the guidance control module of the cruise missile. The cruise missile derives the normal acceleration command according to the bound track point information and the current flight status information, and the guidance law with incident angle and arrival time constraints constructed in step 3, until the cruise missile arrives at the target area at the expected incident angle and arrival time, completing the mid-range guidance that meets the time and space constraints.

[0048] Step 4.1: Step 1 uses the computing resources of the launch platform to solve the mid-range guidance problem using numerical methods to obtain the mid-range guidance desired track points that meet the constraints. Step 2 reduces the number of desired track points that need to be tracked by thinning the mid-range guidance desired track points obtained in step 1. Bind the M track points obtained after thinning to the guidance and control module of the cruiser. Take a certain value R m , preset the initial value j=2, and take the jth track point as the current target track point.

[0049] Step 4.2: The cruise missile calculates the normal acceleration command based on the current target track point information and flight status information according to the guidance law constructed in step 3, and performs real-time feedback tracking on the desired track.

[0050] Step 4.3: Calculate the distance R between the cruise missile's current position and the current target track pointj If R j >R m , repeat step 4.2; if R j ≤R m And j<M, let j=j+1, and take the jth track point as the current target track point; if R j ≤R m And j = M, the cruise missile reaches the designated position and completes the mid-range guidance that meets the time and space constraints.

[0051] Beneficial effects:

[0052] 1. The present invention discloses a mid-course guidance method with time and space constraints. Before launching, the computing resources of the launch platform are used to plan the expected mid-course guidance track points of the cruise missile in the cruise phase according to the mid-course guidance problem. The number of expected track points that need to be tracked is reduced by thinning the expected mid-course guidance track points. On this basis, the incident angle constraint guidance law is used as the reference guidance law, the remaining flight time of the cruise missile using the reference guidance law is estimated, and the estimated value of the remaining flight time is used to design the bias term for controlling the arrival time. The reference guidance law and the bias term are used to construct the incident angle and arrival time constraint guidance law, the expected mid-course guidance track points are tracked, and the complex nonlinear cruise missile mid-course guidance problem is converted into a guidance problem with incident angle and arrival time constraints between a finite number of track points, thereby reducing the complexity of the cruise missile mid-course guidance problem and the computing power requirement for the cruise missile onboard computer, and improving the real-time performance of the mid-course guidance.

[0053] 2. The present invention discloses a mid-range guidance method with time and space constraints. Aiming at the problem that a single angle and time constraint guidance method cannot take into account the demand for circumventing a no-fly zone, the computing resources of the launch platform are used to plan the expected mid-range guidance track points of the cruise missile taking into account the demand for no-fly zones according to the mid-range guidance problem. A finite number of expected track points are obtained by thinning out the expected mid-range guidance track points, and the expected track points are tracked using the guidance law with incident angle and arrival time constraints. The expected track that meets the demand for circumventing a no-fly zone is introduced into the angle and time constraint guidance, thereby meeting the demand for circumventing a no-fly zone while meeting the arrival time and incident angle constraints.

[0054] 3. The present invention discloses a mid-range guidance method with time and space constraints. To solve the problem that the arrival time control error needs to converge within a finite time, the specified time control theory is used to design an error feedback term to correct the arrival time. This can ensure that the arrival time control error converges to zero within the specified time, thereby improving the arrival time control performance. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] Figure 1 It is a flow chart of a loitering missile mid-range guidance method with time and space constraints of the present invention;

[0056] Figure 2is the expected track and the extracted track point diagram of the cruise missile according to the embodiment of the present invention;

[0057] Figure 3 It is a schematic diagram of the relative motion relationship between the cruise missile and the current track point based on the present invention;

[0058] Figure 4 is a flight track diagram of a cruise missile according to an embodiment of the present invention;

[0059] Figure 5 is a graph showing a change in the incident angle error of a cruise missile according to an embodiment of the present invention;

[0060] Figure 6 is a graph showing a change in the arrival time control error of a cruise missile according to an embodiment of the present invention;

[0061] Figure 7 4 is a diagram showing changes in the normal acceleration instruction of the cruise missile according to an embodiment of the present invention. DETAILED DESCRIPTION

[0062] In order to better illustrate the purpose and advantages of the present invention, the invention is further described below with reference to the accompanying drawings and examples.

[0063] In the following embodiments, numerical simulation is used to verify the effectiveness of the proposed guidance method for cruise missiles with time and space constraints. The parameters are set as: N = 3, p = 0.8, t s =0.5|δ j,0 |.

[0064] like Figure 1 As shown, the present embodiment discloses a mid-range guidance method for a cruise missile with time and space constraints, and the specific implementation steps are as follows:

[0065] Step 1: The initial conditions of the cruise missile are x0 = 0m, y0 = 0m, γ0 = 180°. The terminal conditions of the cruise missile are x d =6000m,y d =6000m,γ d =0°. The flight speed of the cruise missile is V M =50m / s, the minimum turning radius is R min = 125m. The expected arrival time at the target point is t f = 200s. The no-fly zone conditions are shown in Table 1.

[0066] Table 1 No-fly zone conditions

[0067] serial number <![CDATA[x o,i (m)]]> <![CDATA[y o,i (m)]]> <![CDATA[R i (m)]]> 1 30 2000 250 2 2600 3200 500 3 4700 4400 400

[0068] Under the above conditions, the mid-course guidance problem considering the requirement of flying around the no-fly zone with the constraints of the incident angle and arrival time is numerically solved, and the expected trajectory is obtained as follows: Figure 2 shown.

[0069] Step 2: Use the Douglas-Peucker algorithm to thin out the expected track planned by the launch platform, and obtain a total of 9 track points including the starting point and the target point. The track points obtained by thinning are as follows: Figure 2 The specific information of the track points is shown in Table 2.

[0070] Table 2 Track point information

[0071] Track point number <![CDATA[x d,i (m)]]> <![CDATA[y d,i (m)]]> <![CDATA[γ d,i (you)]]> <![CDATA[t d,i (s)]]> 1 0 0 180 0 2 -419.77 167.43 136.92 9.26 3 -625.79 557.72 99.87 18.24 4 -283.48 1456.71 44.85 38.24 5 436.35 1912.14 22.79 55.38 6 2910.62 2808.19 35.67 108.24 7 4867.02 5417.97 46.44 173.82 8 5435.38 5855.83 26.76 188.24 9 6000 6000 0 200

[0072] The relative motion between the cruise missile and the current track point is as follows: Figure 3 As shown, where M and T j In the guidance process, according to the state of the cruise missile and the information of the current track point, such as the position of the cruise missile and the current track point, the speed of the cruise missile and the ballistic deviation angle, the remaining time prediction value of the cruise missile to reach the current track point under the action of the reference guidance law is calculated according to the following formula:

[0073]

[0074] Step 3: Calculate the arrival time control error at the current moment by the following formula:

[0075] δ j =t go,j +tt j

[0076] According to the status of the aircraft and the current track point, the reference guidance law, the bias term, and the guidance law constrained by the incident angle and arrival time are determined by the following formulas:

[0077]

[0078]

[0079] a M =a B +a T

[0080] Step 4: The cruise missile calculates the normal acceleration command based on the bound track point information and current flight status information and the guidance law with incident angle and arrival time constraints constructed in step 3, tracks the expected track point, and completes the mid-course guidance that meets the time and space constraints.

[0081] Step 4.1: Bind the 9 track points obtained after thinning to the guidance and control module. m =2m, j=2 is preset, and the jth track point is taken as the current target track point.

[0082] Step 4.2: The cruise missile calculates the normal acceleration command based on the current target track point information and flight status information according to the guidance law constructed in step 3, and performs real-time feedback tracking on the desired track.

[0083] Step 4.3: Calculate the distance R between the cruise missile's current position and the current target track point j If R j >R m , repeat step 4.2; if R j ≤R m And j<M, let j=j+1, and take the jth track point as the current target track point; if R j ≤R m And j=M, the cruise missile reaches the designated position and the mid-course guidance ends.

[0084] The guidance results are as follows Figures 4 to 7 The flight path of the cruise missile is as follows. Figure 4 As shown. Figure 4 It can be seen that the patrol vehicle successfully feedback tracked the desired trajectory and reached the target airspace while satisfying the time and space constraints. Figure 5 and Figure 6 The incident angle error and arrival time control error curves of the patrol vehicle reaching each desired track point during flight are given. Figure 5 and Figure 6 It can be seen that when reaching each desired track point, the incident angle error and the arrival time control error converge to zero. The above results verify the effectiveness of the loitering missile guidance method with time and space constraints proposed in the present invention.

[0085] The above implementation process and exemplary examples have been described in detail the present invention, but these descriptions cannot be understood as limiting the present invention. It is understood by those skilled in the art that, without departing from the spirit and scope of the present invention, various equivalent substitutions, modifications or improvements can be made to the technical solution of the present invention and its implementation methods, which all fall within the scope of the present invention.

Claims

1. A mid-range guidance method for a loitering missile with time and space constraints, characterized in that: The steps include: Step 1: Taking the optimal control energy of the cruise missile as the optimization goal, and taking the no-fly zone detour requirement constraint, minimum turning radius constraint, incident angle and arrival time constraint of the cruise missile in the constant altitude and constant speed cruise phase as the constraint conditions, construct the mid-range guidance problem; using the computing resources of the launch platform to adopt numerical methods to solve the mid-range guidance problem, and obtain the mid-range guidance expected track points that meet the constraints; Step 2: Use the Douglas-Peucker algorithm to thin out the expected mid-range guidance track points planned in step 1, and inject the extracted track points including the expected position, incident angle and arrival time information into the guidance control module of the cruise missile. By thinning out the expected mid-range guidance track points, the number of expected track points that need to be tracked is reduced; the incident angle constraint guidance law is used as the reference guidance law, and the analytical remaining flight time prediction formula is derived according to the reference guidance law; the position of the cruise missile and the current track point, the speed of the cruise missile and the ballistic deviation angle are used as input, and the remaining flight time t is determined according to the remaining flight time prediction formula. go,j ; Step 3: Expected time t to reach the jth track point j and the remaining flight time t obtained in step 2 go,j The difference between the current time and j Defined as the arrival time control error; design the expected rate of change based on the initial value of the arrival time control error and the expected convergence time Construct a bias term to control the arrival time error δ of the cruise missile j Follow the expected rate of change as designed The reference guidance law and the bias term are used to construct the guidance law constrained by the incident angle and arrival time. The guidance law constrained by the incident angle and arrival time is used to ensure that the arrival time control error δ j Convergence is achieved before the cruise missile reaches the jth track point, so that the cruise missile reaches the jth track point at the expected incident angle and expected time, that is, the guidance of the cruise missile with incident angle and arrival time constraints is realized; Step 4: In step 1, the computing resources of the launch platform are used to solve the mid-range guidance problem using numerical methods to obtain the expected mid-range guidance track points that meet the constraints. Step 2: Thinning the expected mid-range guidance track points obtained in step 1 reduces the number of expected track points that need to be tracked, and injects the track points obtained after thinning including the expected position, incident angle and arrival time information into the guidance control module of the cruise missile; the cruise missile derives the normal acceleration command according to the bound track point information and the current flight status information and the guidance law with incident angle and arrival time constraints constructed in step 3, until the cruise missile arrives at the target area with the expected incident angle and arrival time, completing the mid-range guidance that meets the time and space constraints.

2. A mid-range guidance method for a loitering missile with time and space constraints as claimed in claim 1, characterized in that: The implementation method of step one is: Taking the optimal control energy of the cruise missile as the optimization goal, and taking the no-fly zone flight requirement constraint, minimum turning radius constraint, incident angle and arrival time constraint based on the cruise missile in the constant altitude and constant speed cruise phase as the constraint conditions, the mid-range guidance problem is constructed as shown in formula (1): Where J represents the objective function, u represents the virtual control quantity and satisfies |u|≤1, t f represents the expected time to reach the target point, x M represents the x-axis coordinate of the cruise missile in the inertial system, y M represents the coordinate of the cruise missile on the y-axis of the inertial system, V M represents the flight speed of the patrol vehicle, γ M Represents the ballistic deviation angle of the cruise missile, R min Indicates the minimum turning radius of the patrol vehicle, x o,i represents the coordinate of the center of the i-th no-fly zone on the x-axis of the inertial system, y o,i represents the coordinate of the center of the i-th no-fly zone on the y-axis of the inertial system, R i represents the distance from the center of the i-th no-fly zone, x d and d represents the target point position, γ d represents the expected angle of incidence to the target point, x0 and y0 represent the initial position of the cruise missile, and γ0 represents the initial ballistic deviation angle of the cruise missile; The mid-range guidance problem shown in formula (1) is solved by a numerical method to obtain the mid-range guidance expected track point that meets the no-fly zone detour requirement, the minimum turning radius constraint, the incident angle and the arrival time constraint; the expected track point includes the position, the expected incident angle and the arrival time information.

3. A mid-range guidance method for a loitering missile with time and space constraints as claimed in claim 2, characterized in that: The implementation method of step 2 is: Step 2.1: Use the Douglas-Peucker algorithm to thin out the expected track planned by the launch platform, and obtain a total of M track points; the information of each track point includes the expected position, incident angle and arrival time; Step 2.2: Establish the relative motion equation between the cruise missile and the jth (j=2,3,...M) track point, design a reference guidance law with the ability to control the incident angle, and predict the remaining flight time for the cruise missile to reach the jth track point under the action of the reference guidance law; The relative motion equation between the cruise missile and the jth track point is expressed as Among them, r j represents the relative distance between the cruise missile and the jth track point, σ j represents the lead angle of the patrol vehicle, λ j represents the sight angle, a M Indicates the normal acceleration command of the aircraft; The normal acceleration command of the reference guidance law is Where N represents a positive navigation ratio, γ d,j represents the expected angle of incidence of the cruise missile when it reaches the jth track point, represents the predicted value of the incident angle of the cruise missile arriving at the jth track point and satisfies Under the reference guidance law, the remaining flight time for the cruise missile to reach the jth track point is According to formula (8), the remaining flight time for the cruise missile to reach the jth track point under the reference guidance law is predicted.

4. A mid-range guidance method for a loitering missile with time and space constraints as claimed in claim 3, characterized in that: The implementation method of step three is: Step 3.1: Determine the arrival time control error δ j The expected rate of change Arrival time control error δ j represents the expected arrival time t j and the remaining time prediction value t go,j and the difference between the current time; when controlling δ j When it converges to zero, the cruise missile can reach the jth track point at the expected time; the arrival time control error δ j The definition of δ j =t go,j +t-t j (9) Since the flight time from the j1th track point to the jth track point is limited, it is necessary to ensure that the arrival time control error δ j It is very important that the cruise missile converges to zero before it reaches the jth track point; at the same time, it is necessary to ensure that the distance between the cruise missile and the jth track point decreases monotonically during the guidance process to ensure that the lead angle σ j Less than 90°; To ensure the arrival time control error δ j Converges to zero before the cruise missile reaches the jth track point, constructing the arrival time control error δ j The expected rate of change is as follows: Among them, sig p (·)=sign()|·| p , sign() represents the sign function, 0<p<1, δ j,0 Denotes δ j The initial value of t s is a constant value; Step 3.2: Determine the guidance method bias term so that the arrival time control error δ j Following the expected rate of change in equation (10), the guidance of the rover with incident angle and arrival time constraints is achieved; Design the following guidance law with incident angle and arrival time constraints a M =a B +a T (11) Among them, a B represents the reference guidance law, which is used to ensure zero miss and desired angle of incidence; a T Represents the bias term, which is used to ensure the time control error δ j Satisfy the expected rate of change in equation (10), and thus meet the arrival time constraint; From equations (4), (8) and (9), we get in, According to equations (10) and (12), the bias term is Combining equations (6), (11) and (14), we can obtain the normal acceleration command with incident angle and arrival time constraints, so that the arrival time control error of the cruise missile satisfies the expected change rate in equation (10), ensuring that the arrival time control error δ j Convergence is achieved before the cruise missile reaches the jth track point, so that the cruise missile reaches the jth track point at the expected incident angle and expected time.

5. A mid-range guidance method for a loitering missile with time and space constraints as claimed in claim 4, characterized in that: The implementation method of step 4 is: Step 4.1: Step 1: Use the computing resources of the launch platform to solve the mid-range guidance problem with a numerical method to obtain the mid-range guidance desired track points that meet the constraints. Step 2: Reduce the number of desired track points that need to be tracked by thinning the mid-range guidance desired track points obtained in step 1; bind the M track points obtained after thinning to the guidance control module of the cruiser; take a certain value R m , preset the initial value j=2, and take the jth track point as the current target track point; Step 4.2: The cruise missile calculates the normal acceleration command based on the current target track point information and flight status information according to the guidance law constructed in step 3, and performs real-time feedback tracking on the desired track; Step 4.3: Calculate the distance R between the cruise missile's current position and the current target track point j If R j >R m , repeat step 4.2; if R j ≤R m And j<M, let j=j+1, and take the jth track point as the current target track point; if R j ≤R m And j = M, the cruise missile reaches the designated position and completes the mid-range guidance that meets the time and space constraints.

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