A cruise missile guidance method with space-time constraints

Abstract

The application discloses a midcourse guidance method of a loitering munition with space-time constraints and belongs to the field of guidance control. The application realizes the method as follows: constructing a midcourse guidance problem, planning expected midcourse guidance track points according to the midcourse guidance problem by using the computing resources of a launch platform before launching; adopting a Douglas-Puck algorithm to thin out the track points planned by the launch platform, and binding the extracted track points containing position, expected incident angle and arrival time information to a guidance control module of the loitering munition; taking an incident angle constraint guidance law as a benchmark guidance law, estimating the residual flight time of the loitering munition adopting the benchmark guidance law; designing a bias term to the control arrival time by using the residual flight time estimation value, constructing an incident angle and arrival time constraint guidance law by using the benchmark guidance law and the bias term; and according to the incident angle and arrival time constraint guidance law, the loitering munition sequentially passes through the bound midcourse guidance expected track points, so that the midcourse guidance of the loitering munition in a loitering segment is realized.

Description

TECHNICAL FIELD

[0001] The present application relates to a midcourse guidance method with space-time constraints, in particular to a guidance method with incidence angle constraints and time of arrival constraints for a loitering munition, and belongs to the field of guidance control. BACKGROUND

[0002] The modern battlefield environment is increasingly complex, and the enemy's high-value targets usually have multi-level defense systems, which brings great challenges to the penetration of guided weapons. Relying on a single high-performance weapon can effectively penetrate, but the research and use cost 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] Loitering munitions are a new type of intelligent ammunition that can loiter in a designated mission area to perform combat tasks such as loitering reconnaissance, target attack, and effect evaluation. Compared with hypersonic missiles, loitering munitions have lower costs and are more efficient and flexible. Through the cooperative combat of loitering munition clusters, the penetration capability can be improved without significantly increasing the cost of guided weapons, which is of great significance.

[0004] The existing theoretical research and methods of cooperative guidance mainly focus on the terminal guidance phase. Loitering munitions have a long range, and to achieve the cooperative attack task of multiple loitering munitions on the same target, cooperative midcourse guidance is needed before cooperative terminal guidance. In the midcourse phase, the aircraft needs to arrive at the specified airspace within the expected time and meet certain incidence angle constraints to provide favorable initial conditions for terminal guidance. The existing methods for implementing midcourse guidance mainly include numerical optimization methods such as pseudo-spectral method, which have problems such as large search space and slow solving speed in actual application. Due to the limitations of the computing power of the missile-borne computer, it is difficult to ensure the real-time performance of midcourse guidance.

[0005] Considering the above factors, it is necessary to propose a midcourse guidance method for loitering munitions that can meet the incidence angle and time of arrival constraints. SUMMARY

[0006] Aiming at the deficiency of the existing cruise missile midcourse guidance method, the main purpose of the present application is to provide a midcourse guidance method with space-time constraints, taking the optimal control energy of the cruise missile as the optimization target, taking the flight-prohibited area circumnavigation demand constraint, the minimum turning radius constraint, the incident angle and the arrival time constraint of the cruise missile in the constant altitude and constant speed cruise stage as the constraint condition, constructing the midcourse guidance problem, and using the computing resources of the launch platform to plan the midcourse guidance expected track points of the cruise missile cruise stage before launching according to the midcourse guidance problem; the Douglas-Puck algorithm is used to thin out the track points planned by the launch platform, and the track points containing position, expected incident angle and arrival time information are bound to the guidance control module of the cruise missile; the incident angle constraint guidance law is taken as the benchmark guidance law, and the remaining flight time of the cruise missile using the benchmark guidance law is estimated; the bias term to the control arrival time is designed by using the remaining flight time estimation value, and the incident angle and arrival time constraint guidance law is constructed by using the benchmark guidance law and the bias term; according to the incident angle and arrival time constraint guidance law, the cruise missile passes through the bound midcourse guidance expected track points in turn, realizes the flight-prohibited area avoidance and reaches the midcourse and terminal guidance handover position at the expected time and incident angle, and realizes the midcourse guidance of the cruise missile in the cruise stage. Since the midcourse guidance expected track points tracked by the cruise missile according to the incident angle and arrival time constraint guidance law are calculated by using the computing resources of the launch platform, the cruise missile computer power requirement is reduced. The present application has the advantages of considering the arrival time, incident angle constraint and flight-prohibited area circumnavigation demand at the same time, low cruise missile computer power requirement, high midcourse guidance real-time performance. The expected track points include position, expected incident angle and arrival time information.

[0007] The present application discloses a cruise missile midcourse guidance method with space-time constraints, comprising the following steps:

[0008] Step one, taking the optimal control energy of the cruise missile as the optimization target, taking the flight-prohibited area circumnavigation demand constraint, the minimum turning radius constraint, the incident angle and the arrival time constraint of the cruise missile in the constant altitude and constant speed cruise stage as the constraint condition, constructing the midcourse guidance problem. The midcourse guidance problem is solved by using the computing resources of the launch platform and the numerical method, and the midcourse guidance expected track points satisfying the constraints are obtained.

[0009] Taking the optimal control energy of the cruise missile as the optimization target, taking the flight-prohibited area circumnavigation demand constraint, the minimum turning radius constraint, the incident angle and the arrival time constraint of the cruise missile in the constant altitude and constant speed cruise stage as the constraint condition, constructing the midcourse guidance problem as shown in formula (1)

[0010]

[0011] Wherein, 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 coordinate of the cruise missile in the x-axis of the inertial system, yM represents the coordinate of the cruise missile in the y-axis of the inertial system, V M represents the flight speed of the cruise missile, γ M represents the trajectory angle of the cruise missile, R min represents the minimum turning radius of the cruise missile, x o,i represents the coordinate of the center of the i-th no-fly zone in the x-axis of the inertial system, y o,i represents the coordinate of the center of the i-th no-fly zone in 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 y d represents the target point position, γ d represents the desired impact angle to the target point, x0and y0represent the initial position of the cruise missile, γ0represents the initial trajectory angle of the cruise missile.

[0012] The guidance problem shown in equation (1) is solved according to a numerical method to obtain the desired trajectory points of the midcourse guidance that satisfy the no-fly zone fly-around requirement, the minimum turning radius constraint, the impact angle constraint, and the arrival time constraint. The desired trajectory points include position, desired impact angle, and arrival time information.

[0013] Step two, the Douglas-Peucker algorithm is used to thin out the desired trajectory points of the midcourse guidance planned in step one, and the extracted trajectory points including desired position, impact angle, and arrival time information are injected into the guidance control module of the cruise missile to reduce the number of desired trajectory points that need to be tracked by thinning out the desired trajectory points of the midcourse guidance. The impact angle constraint guidance law is used as the reference guidance law, and an analytical residual flight time prediction formula is derived based on the reference guidance law. The position of the cruise missile and the current trajectory point, the speed and trajectory angle of the cruise missile are used as inputs to determine the residual flight time t go,j .

[0014] Step 2.1: The Douglas-Peucker algorithm is used to thin out the desired trajectory planned by the launch platform, and a total of M trajectory points are obtained. Each trajectory point information includes desired position, impact angle, and arrival time.

[0015] Step 2.2: The relative motion equation between the cruise missile and the j-th (j=2, 3,...M) trajectory point is established, and a reference guidance law with impact angle control capability is designed to predict the residual flight time of the cruise missile to the j-th trajectory point under the action of the reference guidance law.

[0016] The relative motion equation between the cruise missile and the j-th trajectory point is represented as

[0017]

[0018]

[0019]

[0020]

[0021] where r j denotes the relative distance between the loitering munition and the jthwaypoint, σ j denotes the lead angle of the loitering munition, λ j denotes the line-of-sight angle, a M denotes the normal acceleration command of the loitering munition.

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

[0023]

[0024] where N denotes the positive navigation ratio, γ d,j denotes the desired impact angle of the loitering munition at the jthwaypoint, denotes the predicted value of the impact angle of the loitering munition at the jthwaypoint and satisfies

[0025]

[0026] The remaining flight time of the loitering munition to the jthwaypoint under the baseline guidance law is

[0027]

[0028] The remaining flight time of the loitering munition to the jthwaypoint under the baseline guidance law is predicted according to equation (8).

[0029] Step three, the difference δ j between the desired time t go,j to the jthwaypoint obtained in step two and the current time is defined as the arrival time control error. The desired change rate is designed based on the initial value of the arrival time control error and the desired convergence time. The bias term is constructed to make the arrival time control error δ j of the loitering munition follow the designed desired change rate The impact angle and arrival time constrained guidance law is constructed using the baseline guidance law and the bias term. According to the impact angle and arrival time constrained guidance law, the arrival time control error δ j is ensured to converge before the loitering munition reaches the jthwaypoint, so that the loitering munition arrives at the jthwaypoint with the desired impact angle and at the desired time, i.e., the guidance of the loitering munition with impact angle and arrival time constraints is realized.

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

[0031] time-to-go control error δ j denotes the desired time-to-go t j and the predicted value of the remaining time t go,j to the jthwaypoint. When the control δ j converges to zero, the cruise missile can reach the jthwaypoint at the desired time t j . The definition of the time-to-go control error δ

[0032]

[0033] Since the flight time from the j1thwaypoint to the jthwaypoint is finite, it is important to ensure that the time-to-go control error δ j converges to zero before the cruise missile reaches the jthwaypoint. At the same time, it is necessary to ensure that the cruise missile maintains a monotonically decreasing distance to the jthwaypoint during the guidance process, and that the lead angle σ j is less than 90°. In order to ensure that the time-to-go control error δ j converges to zero before the cruise missile reaches the jthwaypoint, the desired rate of change of the time-to-go control error δ j is defined as follows:

[0034]

[0035] where sig p (·) = sign(·) |·| p , sign() denotes the sign function, 0 < p < 1, δ j,0 denotes the initial value of δ j , and t s is a constant.

[0036] Step 3.2: Determine the bias term of the guidance method to make the time-to-go control error δ j follow the desired rate of change in equation (10), and achieve guidance of the cruise vehicle with both the angle-of-arrival and time-to-go constraints.

[0037] Design a guidance law with both the angle-of-arrival and time-to-go constraints as follows

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

[0039] where a B denotes the reference guidance law, which is used to ensure zero miss distance and desired angle-of-arrival; a T denotes the bias term, which is used to ensure that the time-to-go control error δ jThe desired rate of change in equation (10) is satisfied, and the arrival time constraint is satisfied.

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

[0041]

[0042] wherein,

[0043]

[0044] The bias term is obtained from equations (10) and (12)

[0045]

[0046] The normal acceleration command with the incidence angle and arrival time constraints is obtained by combining equations (6), (11) and (14), so that the arrival time control error of the cruise missile satisfies the desired rate of change in equation (10), and the arrival time control error δt is ensured to satisfy the arrival time constraint. j The cruise missile converges before reaching the jthwaypoint, and reaches the jthwaypoint at the desired incidence angle and the desired time.

[0047] Step four, step one uses the computing resources of the launch platform to solve the midcourse guidance problem by using a numerical method, and obtains the desired waypoint of the midcourse guidance that satisfies the constraints. Step two reduces the number of desired waypoints to be tracked by thinning out the desired waypoint obtained in step one, and injects the waypoint obtained after thinning out, which includes the desired position, incidence angle and arrival time information, into the guidance control module of the cruise missile. The cruise missile calculates the normal acceleration command according to the guidance law with the incidence angle and arrival time constraints constructed in step three according to the waypoint information and the current flight state information, until the cruise missile flies to the target area at the desired incidence angle and arrival time, and completes the midcourse guidance that satisfies the space-time constraints.

[0048] Step 4.1: Step one uses the computing resources of the launch platform to solve the midcourse guidance problem by using a numerical method, and obtains the desired waypoint of the midcourse guidance that satisfies the constraints. Step two reduces the number of desired waypoints to be tracked by thinning out the desired waypoint obtained in step one. The M waypoints obtained after thinning out are bound to the guidance control module of the cruise vehicle. Take a certain value R m , preset the initial value j=2, and take the jthwaypoint as the current target waypoint.

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

[0050] Step 4.3: Calculate the distance R between the current position of the cruise missile and the current target waypointj 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 waypoint as the current target waypoint; if R j ≤R m and j = M, the cruise missile reaches the specified position, and the midcourse guidance satisfying the space-time constraints is completed.

[0051] Advantages:

[0052] 1. The midcourse guidance method with space-time constraints disclosed in the application uses the computing resources of a launch platform to plan the midcourse expected waypoint of the cruise missile in the cruise phase according to the midcourse guidance problem before launch, reduces the number of expected waypoints to be tracked by thinning the midcourse expected waypoints, and on this basis, takes the angle-of-arrival constraint guidance law as the reference guidance law, estimates the remaining flight time of the cruise missile using the reference guidance law and designs a bias term to the control arrival time using the remaining flight time estimate, constructs the angle-of-arrival and arrival time constraint guidance law using the reference guidance law and the bias term, tracks the midcourse expected waypoint, and converts the complex nonlinear midcourse guidance problem of the cruise missile into a guidance problem with angle-of-arrival and arrival time constraints between a limited number of waypoints, reduces the complexity of the midcourse guidance problem of the cruise missile and the computing power requirement of the missile-borne computer, and improves the real-time performance of the midcourse guidance.

[0053] 2. The midcourse guidance method with space-time constraints disclosed in the application solves the problem that a single angle and time constraint guidance method cannot meet the demand for flying around the no-fly zone by using the computing resources of the launch platform to plan the midcourse expected waypoint of the cruise missile considering the demand for the no-fly zone according to the midcourse guidance problem, reducing the expected waypoints by thinning the midcourse expected waypoints, tracking the expected waypoints using the angle-of-arrival and arrival time constraint guidance law, and introducing the expected waypoint meeting the demand for flying around the no-fly zone into the angle and time constraint guidance, so as to meet the demand for flying around the no-fly zone while meeting the arrival time and angle-of-arrival constraints.

[0054] 3. The midcourse guidance method with space-time constraints disclosed in the application solves the problem that the arrival time control error needs to converge within a limited time by using the specified time control theory to design an error feedback term to correct the arrival time, which can ensure that the arrival time control error converges to zero within the specified time, and improves the arrival time control performance. DETAILED DESCRIPTION

[0055] Figure 1 is a flowchart of the midcourse guidance method with space-time constraints of the cruise missile of the application;

[0056] Figure 2This is a map showing the desired trajectory and extracted track points of a loitering munition according to an embodiment of the present invention;

[0057] Figure 3 This is a schematic diagram illustrating the relative motion relationship between the loitering munition and the current waypoint on which this invention is based;

[0058] Figure 4 This is a flight path diagram of a loitering munition according to an embodiment of the present invention;

[0059] Figure 5 This is a graph showing the variation of the incident angle error of the loitering munition according to an embodiment of the present invention;

[0060] Figure 6 This is a graph showing the variation of the loitering munition arrival time control error according to an embodiment of the present invention;

[0061] Figure 7 This is a diagram showing the variation of the normal acceleration command of a loitering munition according to an embodiment of the present invention. Detailed Implementation

[0062] To better illustrate the purpose and advantages of the present invention, the invention will be further described below in conjunction with the accompanying drawings and examples.

[0063] The following embodiments will verify the effectiveness of the proposed guidance method for loitering munitions with spatiotemporal constraints through numerical simulation. The parameters are set as follows: N = 3, p = 0.8, t s =0.5|δ j,0 |

[0064] like Figure 1 As shown in the figure, this embodiment discloses a guidance method for a loitering munition with spatiotemporal constraints, and the specific implementation steps are as follows:

[0065] Step 1: The initial conditions for the loitering munition are x0 = 0m, y0 = 0m, γ0 = 180°. The terminal conditions for the loitering munition are x d =6000m, y d =6000m, γ d =0°. The loitering munition's flight speed is V. M =50m / s, minimum turning radius is R min =125m. The expected arrival time to the target point is t. f =200s. No-fly zone conditions are shown in Table 1.

[0066] Table 1 No-Fly Zone Conditions

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

[0068] Under the above conditions, a numerical solution is performed on the mid-course guidance problem that considers the need to bypass no-fly zones and has constraints on the angle of incidence and time of arrival. The desired trajectory obtained is as follows:Figure 2

[0069] Step two, the Douglas-Pok algorithm is used to thin the desired track of the launch platform planning, and 9 track points including the starting point and target point are obtained. The track points obtained by thinning are shown in Figure 2

[0070] Table 2 Track point information

[0071] Waypoint Numbering x d,i (m)]]> y d,i (m)]]> d,i (deg)]]> ​ 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 of the cruise missile and the current track point is shown in Figure 3 j M and T represent the cruise missile and the current track point respectively. 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 angle, etc., the predicted value of the remaining time 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 three, the arrival time control error at the current time is calculated by the following formula

[0075] δ j = t go,j +t-t j

[0076] According to the state of the aircraft and the current track point, the reference guidance law, the bias term and the incidence angle and arrival time constraint guidance law are determined by the following formulas respectively

[0077]

[0078]

[0079] a M =a B +a T

[0080] Step four, the cruise missile tracks the expected track point according to the normal acceleration command calculated by the incidence angle and arrival time constraint guidance law constructed in step three based on the track point information and the current flight state information, and completes the midcourse guidance that meets the space-time constraints.

[0081] Step 4.1: The 9 track points obtained by thinning are installed in the guidance control module. Take R m = 2m, preset j = 2, and take the jth track point as the current target track point.

[0082] ​​​Step 4.2: According to the current target waypoint information and flight state information, the cruise missile calculates the normal acceleration command according to the guidance law constructed in step 3, and performs real-time feedback tracking on the expected trajectory.

[0083] Step 4.3: Calculate the distance R between the current position of the cruise missile and the current target waypoint 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 waypoint as the current target waypoint; if R j ≤R m and j = M, the cruise missile reaches the specified position, and the midcourse guidance ends.

[0084] The guidance results are shown in Figures 4 to 7 . The flight trajectory of the cruise missile is shown in Figure 4 . As can be seen from Figure 4 , the cruise missile successfully performs feedback tracking on the expected trajectory and reaches the target airspace under the time and space constraints. Figure 5 and Figure 6 show the angle of arrival error and the arrival time control error curves of the cruise missile when it reaches each expected waypoint during flight. From Figure 5 and Figure 6 , it can be seen that the angle of arrival error and the arrival time control error converge to zero when the cruise missile reaches each expected waypoint. The above results verify the effectiveness of the midcourse guidance method for cruise missiles with time and space constraints proposed in the present application.

[0085] The above implementation process and exemplary examples have been described in detail, but these descriptions should not be understood as limiting the present application. Those skilled in the art understand that various equivalent substitutions, modifications or improvements can be made to the technical solutions and implementation modes of the present application without departing from the spirit and scope of the present application, and these all fall within the scope of the present application.

Claims

1. A guidance method for a cruise missile with space-time constraints, characterized by: The method comprises the following steps: Step one, taking optimal control energy of the cruise missile as the optimization target, taking the flight-prohibited area circumnavigation demand constraint, minimum turning radius constraint, incident angle and arrival time constraint of the cruise missile in the constant height and constant speed cruise stage as the constraint condition, a midcourse guidance problem is constructed; a numerical method is used to solve the midcourse guidance problem by using the computing resources of the launch platform to obtain the midcourse expected track points meeting the constraints; Step two, the Douglas-Pork algorithm is used to thin out the expected track points of the midcourse guidance planned in step one, and the track points including the expected position, incident angle and arrival time information are injected into the guidance control module of the cruise missile, so as to reduce the number of expected track points to be tracked by thinning out the expected track points of the midcourse guidance; the incident angle constraint guidance law is taken as the reference guidance law, and the analytical residual 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 and the trajectory angle of the cruise missile are taken as the input, and the residual flight time t go,j is determined according to the residual flight time prediction formula. The implementation method of step two is, Step 2.1: the Douglas-Puck algorithm is used to thin the expected track planned by the launch platform, and M track points are obtained; each track point information comprises an expected position, an incident angle and an arrival time; Step 2.2: a relative motion equation between the cruise missile and the jth track point is established, j = 2, 3,..., M, a reference guidance law with incident angle control capability is designed, and the remaining flight time of the cruise missile to the jth track point under the action of the reference guidance law is predicted; The relative motion equation between the cruise missile and the jth track point is represented as where r j represents the relative distance between the loitering missile and the jthwaypoint, σ j represents the lead angle of the loitering vehicle, λ j represents the line-of-sight angle, a M represents the normal acceleration command of the vehicle; The normal acceleration instruction of the reference guidance law is where N denotes a positive navigation ratio, γ d,j denotes the expected impact angle of the cruise missile to the jthwaypoint, denotes the predicted value of the impact angle of the cruise missile to the jthwaypoint and satisfies The remaining flight time of the cruise missile to the jth track point under the action of the reference guidance law is The remaining flight time of the cruise missile to the jth track point under the action of the reference guidance law is predicted according to formula (8); Step three, the expected time t to reach the jth waypoint j and the remaining flight time t obtained in step two go,j and the current time t j defined as the time-to-go control error; the expected rate of change is designed based on the initial value of the time-to-go control error and the expected convergence time constructing the bias term to make the time-to-go control error δ of the cruise missile j following the designed expected rate of change constructing the angle-of-arrival and time-to-go constraint guidance law using the reference guidance law and the bias term; ensuring the time-to-go control error δ according to the angle-of-arrival and time-to-go constraint guidance law j achieving convergence before the cruise missile reaches the jth waypoint, so that the cruise missile reaches the jth waypoint at the expected angle-of-arrival and the expected time, i.e. achieving the guidance with angle-of-arrival and time-to-go constraints for the cruise missile Step four, step one uses the computing resources of the launch platform to solve the midcourse guidance problem by using the numerical method to obtain the midcourse expected track points meeting the constraints, step two thins the midcourse expected track points obtained in step one to reduce the number of expected track points to be tracked, and the track points obtained after thinning, comprising expected position, incident angle and arrival time information, are injected into the guidance control module of the cruise missile; the cruise missile obtains the normal acceleration instruction according to the track point information and the current flight state information and according to the guidance law with incident angle and arrival time constraint constructed in step three, until the cruise missile flies to the target area at the expected incident angle and arrival time, and the midcourse guidance meeting the space-time constraints is completed.

2. The guidance method for a loitering munition with spatiotemporal constraints as described in claim 1, characterized in that: The implementation method of step one is, Taking optimal control energy of the cruise missile as the optimization target, taking the flight-prohibited area circumnavigation demand constraint, minimum turning radius constraint, incident angle and arrival time constraint of the cruise missile in the constant height and constant speed cruise stage as the constraint condition, a midcourse guidance problem is constructed as shown in formula (1) where J denotes the objective function, u denotes the virtual control variable and satisfies |u|≤1, t f denotes the desired time to reach the target point, x M denotes the coordinate of the cruise missile in the x-axis of the inertial system, y M denotes the coordinate of the cruise missile in the y-axis of the inertial system, V M denotes the flight speed of the cruise missile, γ M denotes the trajectory angle of the cruise missile, R min denotes the minimum turning radius of the cruise missile, x o,i denotes the coordinate of the center of the ith forbidden flight zone in the x-axis of the inertial system, y o,i denotes the coordinate of the center of the ith forbidden flight zone in the y-axis of the inertial system, R i denotes the distance from the center of the ith forbidden flight zone, x d and y d denote the target point position, γ d denotes the desired angle of incidence to reach the target point, x0and y0denote the initial position of the cruise missile, γ0denotes the initial trajectory angle of the cruise missile; The midcourse guidance problem shown in formula (1) is solved by using the numerical method to obtain the midcourse expected track points meeting the flight-prohibited area circumnavigation demand, minimum turning radius constraint, incident angle and arrival time constraint; the expected track points comprise position, expected incident angle and arrival time information.

3. The guidance method for a cruise missile with space-time constraints as claimed in claim 2, characterized in that: The implementation method of step three is, Step 3.1 : Determining the time of arrival control error delta j Desired rate of change Arrival time control error δ j Indicates the expected arrival time t j and remaining time prediction t go,j And the difference between the current time and the current time; when controlling δ j When convergence reaches zero, the loitering munition can then reach the j-th waypoint at the desired time; the error δ is controlled by the arrival time. j The definition is δ j = t go,j + t - t j (9) Since the flight time from the (j-1)th waypoint to the jth waypoint is finite, it is necessary to ensure that the arrival time control error δ j It is important to converge to zero before the cruise missile reaches the jth waypoint; at the same time, it is necessary to ensure that the distance between the cruise missile and the jth waypoint monotonically decreases during the guidance process, and the lead angle σ j is less than 90°; in order to ensure that the arrival time control error δ j converges to zero before the cruise missile reaches the jth waypoint, the expected rate of change of the arrival time control error δ j is shown in the following formula: wherein sig p ()sign()|| p , sign(·) denotes a sign function, 0 < p < 1, δ j,0 denotes an initial value of δ j , t s is a constant value; Step 3.2: Determine the guidance method bias term to control the time of arrival error δ j Following the desired rate of change in equation (10), the guidance to the loiterer with both angle of arrival and time of arrival constraints is achieved; The guidance law with incident angle and arrival time constraint is designed as follows a M = a B + a T (11) where a B represents the reference guidance law, which guarantees zero miss distance and desired impact angle; a T represents the bias term, which guarantees the time-to-go control error δ j satisfies the desired rate of change in equation (10), and thus satisfies the time-to-go constraint; From formula (4), (8) and (9), it can be obtained that wherein, According to formula (10) and (12), the bias term is Combining equations (6) (11) and (14), the normal acceleration command with the constraints of the impact angle and the time of arrival is obtained, which makes the time of arrival control error satisfy the expected change rate in equation (10), and ensures that the time of arrival control error δ j The convergence is achieved before the cruise missile reaches the jthwaypoint, and the cruise missile reaches the jthwaypoint with the expected impact angle and the expected time.

4. The guidance method for a cruise missile with space-time constraints as claimed in claim 3, characterized in that: The implementation method of step four is, Step 4.1: using the computing resources of the launch platform to solve the midcourse guidance problem by numerical method to obtain the desired midcourse waypoints satisfying the constraints, step 2: reducing the number of desired waypoints to be tracked by thinning out the desired waypoints obtained in step 1; binding the M waypoints obtained after thinning to the guidance control module of the loitering munition; taking a certain value R m , presetting the initial value j = 2, and taking the jth waypoint as the current target waypoint; Step 4.2: the cruise missile obtains the normal acceleration instruction according to the current target track point information and flight state information and according to the guidance law constructed in step three, and performs real-time feedback tracking on the expected track; Step 4.3: Calculate the distance R between the current position of the cruise missile and the current target waypoint j ; if R j >R m , repeat step 4.2; if R j ≤R m and j j ≤R m and j = M, the cruise missile reaches the specified position, completing the midcourse guidance that meets the space-time constraints.

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