A hyper-sonic interceptor missile multi-constraint midcourse guidance method system considering information loss

Abstract

The application discloses a kind of high supersonic interceptor multi-constraint midcourse guidance method systems considering information loss, belong to interceptor technical field.The method is first established under the ground trajectory system Intercept missile mathematical model, then the end attack angle constraint is converted into terminal line-of-sight constraint tracking problem, simplify the problem difficulty;Information is estimated using a filter, then a trajectory classifier is designed based on long short memory neural network and a trajectory prediction database is trained, the database is used for trajectory prediction as the support of guidance rate design, finally the expected shooting angle is designed by the estimated target aircraft information, and the guidance rate is derived, which improves the accuracy of guidance, and the effectiveness of the method is verified by MATLAB simulation software.

Description

Technical Field

[0001] This invention belongs to the field of interceptor missile technology and relates to a multi-constraint guidance method system for hypersonic interceptor missiles that takes into account information loss. Background Technology

[0002] Many aircraft possess supersonic and stealth capabilities, while near-space hypersonic aircraft have extremely high flight altitudes and speeds.

[0003] Near-space hypersonic interceptors often face the problem of missing enemy information during interception. Accurate acquisition and processing of target vehicle information is crucial for effective guidance. However, the characteristics of near-space hypersonic vehicles result in small radar cross-sections, high maneuverability, and significantly increased radar detection and tracking difficulties at hypersonic speeds, making them highly susceptible to information loss or delays. This information loss may include not only lags in updating basic parameters such as target vehicle position and velocity, but also the lack of advanced information such as target vehicle trajectory prediction and maneuver intent recognition, posing a severe challenge to the real-time performance and accuracy of the interception guidance system. Therefore, addressing the issue of hypersonic interception guidance that considers information loss has become an urgent problem to be solved. Summary of the Invention

[0004] The purpose of this invention is to overcome the shortcomings of the prior art and provide a multi-constraint guidance method system for hypersonic interceptor missiles that takes into account the lack of information, so as to solve the problem that near-space hypersonic interceptor missiles often face the problem of missing target aircraft information during the interception process.

[0005] To achieve the above objectives, the present invention employs the following technical solution:

[0006] A multi-constraint guidance method for hypersonic interceptor missiles that considers information loss includes the following steps:

[0007] S1. Establish a motion model for the interceptor missile and set multiple constraints for the motion model, including maximum stagnation point heat flux density constraint, maximum dynamic pressure constraint, maximum overload constraint, control quantity constraint, and attack angle constraint; the attack angle constraint is converted into simple terminal line-of-sight tracking.

[0008] S2, the position of the target aircraft in the observation coordinate system is obtained through the tracking filter;

[0009] S3, based on the position of the target aircraft in the observation coordinate system, the position, velocity and trajectory category of the target aircraft in the geocentric rectangular coordinate system are obtained through a ballistic identification classifier; the ballistic identification classifier is obtained by training an LSTM network;

[0010] S4. When target aircraft information is missing, based on the ballistic category of the target aircraft in the geocentric rectangular coordinate system obtained in S3, the interceptor missile and the target aircraft's firing angle are obtained, and the desired velocity of the interceptor missile is further obtained. The desired ballistic deflection angle is obtained through the desired velocity. Based on simple line-of-sight tracking at the terminal, a lateral tracking guidance law is designed. Through the lateral tracking guidance law, the mid-course guidance correction of the interceptor missile is performed in the lateral plane. The mid-course guidance correction is to eliminate the ballistic deflection angle error through the sideslip angle command, and to guide the interceptor missile to the target aircraft through mid-course guidance.

[0011] A further improvement of the present invention is that:

[0012] Preferably, in S1, the maximum stagnation point heat flux density constraint is:

[0013]

[0014] Among them, Q max ρ is the maximum allowable heat flux density of the interceptor missile, k is the proportionality coefficient for a specific interceptor missile shape, and ρ is the atmospheric density.

[0015] Preferably, in S1, the maximum dynamic pressure constraint is:

[0016] q=0.5ρv 2 ≤q max (6)

[0017] In the formula, q max This indicates the maximum permissible dynamic pressure value of the interceptor missile, measured in P. a ρ is the atmospheric density, and v is the aircraft speed.

[0018] Preferably, the maximum overload constraint is:

[0019]

[0020] Where L and D represent the lift and drag of the interceptor missile, respectively, α is the angle of attack, m is the mass of the interceptor missile, g0 is the gravitational acceleration at sea level, and n ymax Indicates the maximum allowable overload value for the interceptor missile;

[0021] The control quantity constraint is:

[0022] α min ≤α≤α max (8)

[0023] Where α is the angle of attack, α min α is the minimum angle of attack allowed for the interceptor missile. max This is the maximum angle of attack allowed for the interceptor missile.

[0024] Preferably, in S1, the formula for converting the attack angle constraint into simple line-of-sight tracking is:

[0025]

[0026] Where, q f The line-of-sight angle at the end of guidance. The heading angle of the target aircraft at the end of the final guidance phase. V is the angle of attack of the missile. t V m These are the velocity vectors of the target aircraft and the missile, respectively.

[0027] Preferably, in S2, the tracking filter is:

[0028]

[0029] In the formula, X∈R 9 , Y∈R 3 The measured output of the filter is denoted by W, and the measured noise is denoted by V, which represent the system noise and the measurement noise, respectively.

[0030] Preferably, in S2, the position of the target in the observation coordinate system is:

[0031]

[0032] In the formula, R is the distance measurement value of the observation device; ε and κ are the angle measurement values ​​of the observation device.

[0033] Preferably, in S4, the desired ballistic deflection angle is:

[0034]

[0035] Among them, V G The velocity of the interceptor missile is given in the launch coordinate system.

[0036] Preferably, in S4, the sideslip angle command is:

[0037]

[0038] Among them, F aero,z Aerodynamic force in velocity coordinate system The lateral force coefficient affected by the sideslip angle, q dynamic pressure, and S is the reference area of ​​the interceptor missile.

[0039] A multi-constraint guidance system for hypersonic interceptors that takes into account information loss includes:

[0040] The motion model unit is used to establish the motion model of the interceptor missile and set multiple constraints for the motion model. The constraints include maximum stagnation point heat flux density constraint, maximum dynamic pressure constraint, maximum overload constraint, control quantity constraint, and attack angle constraint. The attack angle constraint is converted into simple terminal line-of-sight tracking.

[0041] The tracking unit is used to obtain the position of the target aircraft in the observation coordinate system through a tracking filter;

[0042] The classification unit is used to obtain the position, velocity, and trajectory category of the target aircraft in the geocentric rectangular coordinate system based on the position of the target aircraft in the observation coordinate system through a trajectory identification classifier; the trajectory identification classifier is obtained by training an LSTM network.

[0043] The guidance unit, when enemy information is lacking, obtains the trajectory category of the target aircraft in the geocentric rectangular coordinate system obtained by the classification unit, acquires the interceptor missile and the target aircraft's firing angle, further obtains the interceptor missile's desired velocity, and obtains the desired trajectory deflection angle through the desired velocity; based on simple line-of-sight tracking at the terminal, a lateral tracking guidance law is designed, and through the lateral tracking guidance law, the interceptor missile's mid-course guidance correction is performed in the lateral plane. The mid-course guidance correction is to eliminate the trajectory deflection angle error through sideslip angle commands, and guide the interceptor missile to the target aircraft through mid-course guidance.

[0044] Compared with the prior art, the present invention has the following beneficial effects:

[0045] This invention discloses a multi-constraint guidance method for hypersonic interceptor missiles considering information gaps. The method first establishes a mathematical model of the interceptor missile within a ground-based ballistic system. Second, it considers process constraints, control quantity constraints, initial state constraints, and terminal state constraints, providing specific mathematical descriptions. The terminal attack angle constraint is transformed into a terminal line-of-sight constraint tracking problem, simplifying the problem's difficulty. A filter is designed to estimate the acquired information. Then, a ballistic classifier is designed based on a long short-term memory neural network and trained to obtain a ballistic prediction database. This database is used for ballistic prediction to support guidance rate design. Finally, the guidance rate is derived by designing the expected firing angle based on the estimated target vehicle information, improving guidance accuracy. The method is validated using MATLAB simulation software to determine its effectiveness.

[0046] Furthermore, the ballistic recognition classifier is obtained through training an LSTM network. A comprehensive database of different ballistic forms is established through extensive training. When information is missing, guidance commands are designed by calling upon the data for ballistic prediction. Simulation verification shows that this scheme can effectively solve the interception problem when information is missing and has good guidance effect. Attached Figure Description

[0047] Figure 1 This is a training diagram of the network model of the present invention;

[0048] Figure 2 This is a diagram of the multi-constraint guidance method for hypersonic interceptor missiles of the present invention;

[0049] Figure 3 This is a schematic diagram of the interceptor missile's guided strike according to the present invention. Detailed Implementation

[0050] Hereinafter, the terms "first," "second," "third," and "fourth" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined as "first," "second," "third," or "fourth" may explicitly or implicitly include one or more of that feature.

[0051] The synchronization method provided in this application can be applied to terminal devices such as mobile phones, tablets, wearable devices, in-vehicle devices, augmented reality (AR) / virtual reality (VR) devices, laptops, ultra-mobile personal computers (UMPCs), netbooks, and personal digital assistants (PDAs). This application does not impose any restrictions on the specific type of terminal device.

[0052] It should be noted that the terms "first," "second," etc., used in the specification and drawings of this invention are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the invention described herein can be implemented in orders other than those illustrated or described herein. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover non-exclusive inclusion; for example, a process, method, system, product, or apparatus that comprises a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0053] The first aspect of this invention discloses a multi-constraint guidance method for hypersonic interceptor missiles that considers information loss, the method specifically including the following steps:

[0054] Step 1: Establish a mathematical model of the interceptor missile's motion.

[0055] The kinematic and dynamic model of the interceptor's center of mass is established in the ballistic coordinate system as follows:

[0056]

[0057] Where: P—angle of attack; α B —Angle of attack; β B —Side slip angle; γ v — Velocity tilt angle; X— Air resistance; Y— Lift; Z— Lateral force; m— Missile mass; g— Gravitational acceleration.

[0058] Where X and Y can typically be described as

[0059]

[0060] in, Represents dynamic pressure; ρ—atmospheric density at flight altitude; v—air-to-air velocity of the interceptor missile; S—reference area of ​​the interceptor missile; C x C y C z These are aerodynamic coefficients, representing drag coefficient, lift coefficient, and lateral force coefficient, respectively.

[0061] The dynamic and kinematic model of the interceptor missile rotating around its center of mass is established as follows:

[0062]

[0063] The torque model is as follows:

[0064]

[0065] Among them, M x M y M z It is a triaxial torque;

[0066] ω x ω y ω z This provides the three-axis attitude and angular velocity information for the interceptor missile;

[0067] J x J y J z The moment of inertia is the rotational inertia of the three axes;

[0068] The rolling moment coefficient caused by the rudder deflection angle;

[0069] The yaw moment coefficient caused by the rudder deflection angle;

[0070] This is the pitching moment coefficient caused by the rudder deflection angle.

[0071] Step 2: Establish a multi-constraint mathematical model

[0072] Hypersonic interceptor missiles experience strong drag and deceleration upon re-entry into the atmosphere, resulting in a dramatic increase in surface temperature. Therefore, designing the re-entry trajectory requires consideration of a series of constraints, including process constraints, control quantity constraints, initial state constraints, and terminal state constraints. The mathematical descriptions of these constraints are given below.

[0073] a) Maximum stagnation point heat flux density constraint

[0074] During reentry, the interceptor missile's mechanical energy is converted into heat energy through braking. This enormous energy is transferred from the atmosphere to the aircraft surface, causing high-temperature ablation of the aircraft structure, particularly at the stagnation point of the nose where the heat flux density is higher, resulting in more severe ablation. Therefore, the interceptor missile's structure can be protected by limiting the stagnation point heat flux density. The following empirical formula is used to calculate the value of the stagnation point heat flux density:

[0075]

[0076] Among them, Q max This represents the maximum permissible heat flux density of the interceptor missile, where k is a proportionality coefficient for a specific interceptor missile shape, which is taken as 5.5 × 10⁻⁶ in this invention. -8 ρ is the atmospheric density.

[0077] b) Maximum dynamic pressure constraint

[0078] To protect the structural safety of the interceptor missile, the maximum dynamic pressure during flight needs to be limited to restrict the aerodynamic forces and torques during reentry. Let q represent the dynamic pressure during flight, which can be obtained from the following equation:

[0079] q=0.5ρv 2 ≤q max (6)

[0080] In the formula, q max This indicates the maximum permissible dynamic pressure value of the interceptor missile, measured in P. a

[0081] c) Maximum overload constraint

[0082] During reentry, the amplitude of the maximum overload needs to be limited to protect the interceptor missile's structure from damage caused by aerodynamic forces and moments. In engineering design, constraints are typically placed on the interceptor missile's normal overload or total overload. This invention constrains the total overload, and its calculation formula is as follows:

[0083]

[0084] In the formula, L and D represent the lift and drag of the interceptor missile, respectively, α is the angle of attack, m is the mass of the interceptor missile, g0 is the gravitational acceleration at sea level, and n ymax This indicates the maximum allowable overload value for the interceptor missile.

[0085] d) Control constraints

[0086] During reentry, the interceptor missile primarily controls its aerodynamic forces and torque through angle-of-attack and billet commands, thereby controlling its flight trajectory. While protecting the interceptor missile's structural safety, the amplitude of the angle of attack must be constrained to ensure reliable aerodynamic forces generated by the actuators. Let α... max Let represent the maximum angle of attack allowed for the interceptor missile. Then, the control constraints can be expressed as:

[0087] α min ≤α≤α max (8)

[0088] e) Attack Angle Constraint

[0089] Constraining the attack angle meets the need for interceptor missiles to adjust their line-of-sight angle to ensure good side window detection conditions in the operation of intercepting high-speed maneuvering target aircraft. To facilitate the problem-solving, this section mainly focuses on modeling the longitudinal plane.

[0090] Since there are three definitions of the angle of attack, the definition used here is: the angle between the missile's velocity vector and the target aircraft's velocity vector when the interceptor missile hits the target aircraft.

[0091] The guidance law is designed based on the idea of ​​zero line-of-sight angular rate, that is, the missile-target line-of-sight angular rate is 0 at the moment the missile hits the target vehicle.

[0092]

[0093] in, and q represents the heading angle of the target aircraft and the trajectory inclination angle of the interceptor missile at the end of terminal guidance, respectively. f The line-of-sight angle at the end of guidance.

[0094] Define attack angle The angle between the velocity vectors of the missile and the target vehicle at the moment of impact is called the angle between their respective velocity vectors.

[0095]

[0096] achievable

[0097]

[0098] Assume V t <V m The final line-of-sight angle obtained by performing trigonometric operations on the above equation is:

[0099]

[0100] if Given that, for a given angle of attack It can be obtained. Therefore, it can be seen from the above formula that there always exists a q. f Correspondingly, the terminal attack angle constraint problem can be transformed into a simple terminal line-of-sight tracking problem.

[0101] Step 3: Design of the tracking filter on the interceptor missile

[0102] Define the observation coordinate system as (ox) g ,y g ,z g ), where the origin is the location of the observation radar, oy g In a vertical plane, pointing upwards is considered positive. ox g Within the attack plane formed by the origin and the target aircraft detected by the radar, and with oy g Vertical, pointing towards the target aircraft is positive. oz g The right-hand rule is used to determine this.

[0103] Within the observation coordinate system, define the following three state variables:

[0104]

[0105] Among them, C x C y C z The aerodynamic coefficient; m t Let S be the mass of the target aircraft; and S be the equivalent reference area of ​​the target aircraft. Selecting state variables, the following state equations are established:

[0106]

[0107] in:

[0108]

[0109] Where ρ is the air density; The speed of the target aircraft; μ = 3.98199 × 10 14 ;R e The radius of the Earth; This is the transformation matrix from the velocity coordinate system to the observation coordinate system, and its specific representation is as follows:

[0110]

[0111] The following EKF model is established:

[0112]

[0113] In the formula, X∈R 9 , Y∈R 3 Let W be the measured output of the filter, and let W and V be the system noise and measurement noise, respectively, both being zero-mean Gaussian white noise. This model is a hybrid system with continuous state and discrete output.

[0114]

[0115] In practice, radar and other detection devices can only acquire the distance and angle information of the target aircraft within the observation coordinate system. Let ε and κ represent the elevation angle and azimuth angle, respectively. Then, within the observation coordinate system, the measured position coordinates of the target aircraft... It can be obtained from the following formula, namely:

[0116]

[0117] In the formula, The distance measurement value is obtained by the observation device; and Let be the measured angle value obtained by the observation device. The variance of the distance measurement by the measuring device, E(ΔR), is known. 2 The variance of the angle measurement, E(Δε). 2 E(Δκ) 2 Distance measurement noise, elevation noise, and azimuth noise are independent of each other, meaning the noise variance matrix R can be measured. n .

[0118] Step 4, Ballistic prediction based on LSTM

[0119] Hypersonic target vehicles can glide without power during reentry in several possible modes, primarily including balanced gliding and balanced skip gliding. These represent two typical flight trajectories. Balanced gliding refers to maintaining approximately constant altitude flight conditions according to equilibrium equations; a typical trajectory is the famous Qian Xuesen trajectory. Balanced skip gliding can be figuratively described as a skipping-on-water trajectory, also known as the Sanger trajectory.

[0120] Since the unpowered reentry glide trajectory of hypersonic vehicles is far more complex than that of inertial trajectories, trajectory classification must first be performed based on measurement data. A trajectory classification classifier is constructed, with the target vehicle information as input and the trajectory type as output; the main body of the trajectory classification classifier is an LSTM network.

[0121] The reentry phase ballistic data of hypersonic glide vehicles is divided into balanced glide and balanced jump glide, denoted as Glide and Jump, respectively. The total ballistic dataset can be represented as:

[0122] DATA={TD_glide,TD_jump} (23)

[0123] In the formula, TD_glide and TD_jump represent the datasets for various ballistic trajectories, respectively.

[0124] TD = {tra 1 ,tra 2 ,…,tra N} (twenty four)

[0125] In the formula, tra N The ballistic sample is a complete trajectory of the target aircraft, which is a ballistic time series.

[0126] tra={D x D y D z D Vx D Vy D Vz ,flag} (25)

[0127] In the formula, D x D y D z D Vx D Vy D Vz The flags represent the target aircraft's position, velocity, and trajectory category in a geocentric rectangular coordinate system. The geocentric rectangular coordinate system is defined as follows: the x-axis is aligned with the Earth's rotation axis, the y-axis points to the intersection of the launch point meridian and the equator in the equatorial plane, and the z-axis forms a right-handed rectangular coordinate system with the two axes.

[0128] Preprocessing normalizes the possible range of values ​​for the ballistic data roots to reduce the amplitude range, and then manually assigns labels to them to form the formal network training sample set.

[0129] After constructing the ballistic recognition classifier, the network is trained using the network training sample set to obtain the trained ballistic recognition classifier.

[0130] After the network construction of the ballistic identification classifier is completed, it is also necessary to classify the ballistics according to the current situation and perform ballistic prediction to support the design of the interceptor's maneuverability and guidance rate.

[0131] Step 5: Orientation-based mid-course guidance design

[0132] When target vehicle information is missing, the predicted position information of the target vehicle is obtained through the ballistic category in step 4. Guidance design is then based on this predicted position information as a hypothetical position of the target vehicle. Assuming the attack angle constraint is met, the aforementioned attack angle constraint has already addressed the vertical plane; subsequent design processes mainly focus on the lateral plane. Therefore, the mid-course guidance problem of the missile transforms into a guidance correction problem in the lateral plane. The main task of lateral guidance is to eliminate heading angle errors by designing a sideslip angle command β, guiding the interceptor to the target vehicle. The ballistic deflection angle is defined as the angle between the projection of the missile's velocity vector onto the horizontal plane (i.e., the Oxz plane of the ground coordinate system) and the ox axis of the ground coordinate system.

[0133] In the above process, the predicted position information of the target aircraft can be obtained by existing methods based on the ballistic category.

[0134] See Figure 3 The guidance law design process includes the following steps:

[0135] Define the coordinates of the current position A of the interceptor missile as follows: The latitude and longitude coordinates of the target aircraft T are: The firing angle Ad from the current point A to the target aircraft T can then be calculated. It is easy to see that when the interceptor's velocity direction coincides with the current firing direction, the interceptor can accurately strike the target aircraft. Therefore, this velocity direction is defined as the desired velocity direction, and a north-sky-east coordinate system is established at point A to represent this velocity, as follows:

[0136]

[0137] The conversion of this velocity back to the original launch system is as follows:

[0138]

[0139] In the formula, R ET-NRE The transformation matrix from the North-Eastern coordinate system to the Geocentric Cartesian coordinate system is shown below:

[0140]

[0141] R G-ET The transformation matrix from the geocentric rectangular coordinate system to the launch coordinate system is represented as follows:

[0142]

[0143] Therefore, the expected trajectory deflection angle σ is... d It can be represented as follows:

[0144]

[0145] The lateral tracking guidance law is designed using the concept of proportional feedback control, as shown in equation (37), where the tracking error is defined as e. σ =σ-σ d Therefore, the expected rate of change of tracking error is:

[0146]

[0147] The dynamic equation for the reentry trajectory deflection is given as follows:

[0148]

[0149] Where the subscript 3 represents the ballistic coordinate system, F aero This is due to aerodynamic forces. Therefore, the actual rate of change of the tracking error is:

[0150]

[0151] Considering This can be simplified to:

[0152]

[0153] The required aerodynamic force is:

[0154] F aero,z3 =k1e σ mVcosθ-mg z3 (35)

[0155] Under BTT control conditions, when the tilt angle is 0, the ballistic coordinate system coincides with the velocity coordinate system, therefore:

[0156]

[0157] The sideslip angle command for the next moment can be obtained as follows:

[0158]

[0159] A second aspect of the present invention discloses a multi-constraint guidance system for hypersonic interceptors that takes into account information loss, comprising:

[0160] The motion model unit is used to establish the motion model of the interceptor missile and set multiple constraints for the motion model. The constraints include maximum stagnation point heat flux density constraint, maximum dynamic pressure constraint, maximum overload constraint, control quantity constraint, and attack angle constraint. The attack angle constraint is converted into simple terminal line-of-sight tracking.

[0161] The tracking unit is used to obtain the position of the target aircraft in the observation coordinate system through a tracking filter;

[0162] The classification unit is used to obtain the position, velocity, and trajectory category of the target aircraft in the geocentric rectangular coordinate system based on the position of the target aircraft in the observation coordinate system through a trajectory identification classifier; the trajectory identification classifier is obtained by training an LSTM network.

[0163] The guidance unit, when enemy information is lacking, obtains the trajectory category of the target aircraft in the geocentric rectangular coordinate system obtained by the classification unit, acquires the interceptor missile and the target aircraft's firing angle, further obtains the interceptor missile's desired velocity, and obtains the desired trajectory deflection angle through the desired velocity; based on simple line-of-sight tracking at the terminal, a lateral tracking guidance law is designed, and through the lateral tracking guidance law, the interceptor missile's mid-course guidance correction is performed in the lateral plane. The mid-course guidance correction is to eliminate the trajectory deflection angle error through sideslip angle commands, and guide the interceptor missile to the target aircraft through mid-course guidance.

[0164] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method for multi-constraint guidance of a hypersonic interceptor missile considering information loss, characterized in that, The method comprises the following steps: S1, a motion model of the interceptor is established, and a plurality of constraint conditions are set for the motion model, the constraint conditions comprising a maximum stagnation point heat flux density constraint, a maximum dynamic pressure constraint, a maximum overload constraint, a control quantity constraint, and an attack angle constraint; the attack angle constraint is converted into terminal line-of-sight simple tracking; S2, a position of the target aircraft in an observation coordinate system is obtained through a tracking filter; S3, based on the position of the target aircraft in the observation coordinate system, a position, a speed, and a trajectory category of the target aircraft in a geocentric rectangular coordinate system are obtained through a trajectory identification classifier; the trajectory identification classifier is obtained through LSTM network training; S4, when target aircraft information is missing, based on the trajectory category of the target aircraft in the geocentric rectangular coordinate system obtained in S3, a shooting angle of the interceptor and the target aircraft is obtained, and further, an expected speed of the interceptor is obtained, and an expected trajectory deflection angle is obtained through the expected speed; based on the terminal line-of-sight simple tracking, a lateral tracking guidance law is designed, and through the lateral tracking guidance law, a midcourse guidance correction of the interceptor in a lateral plane is performed; the midcourse guidance correction is to eliminate the trajectory deflection angle error through a sideslip angle command, and the interceptor is guided into the target aircraft through the midcourse guidance.

2. The method of claim 1, wherein, In S1, the maximum stagnation point heat flux density constraint is: where Q max is the maximum heat flux allowed for the interceptor, k is a proportionality factor for a certain interceptor shape, and p is the atmospheric density.

3. The method of claim 1, wherein, In S1, the maximum dynamic pressure constraint is: q = 0.5 p v 2 ≤ q max (6) where q max represents the maximum dynamic pressure allowed for the interceptor missile, in P a , p is the atmospheric density, and v is the vehicle speed.

4. The method of claim 1, wherein, The maximum overload constraint is: where L, D represent the lift and drag of the interceptor, respectively, a is the angle of attack, m is the mass of the interceptor, g0is the acceleration due to gravity at sea level, n ymax represents the maximum allowable value of the overload of the interceptor; The control quantity constraint is: α min ≤α≤α max (8) where a is the angle of attack, a min is the minimum angle of attack allowed for the interceptor, a max is the maximum angle of attack allowed for the interceptor.

5. The method of claim 1, wherein, In S1, a formula for converting the attack angle constraint into the terminal line-of-sight simple tracking is: where q f is the line-of-sight angle at the end of guidance, is the heading angle of the target vehicle at the end of terminal guidance, θ is the attack angle of the missile, V t , V m are the velocity vectors of the target vehicle and the missile, respectively.

6. The method of claim 1, wherein, In S2, the tracking filter is: where X ∈ R 9 , Y ∈ R 3 is the measurement output of the filter, W, V are the system and measurement noise, respectively.

7. The method of claim 1, wherein, In S2, the position of the target in the observation coordinate system is: wherein is a distance measurement by the observation device; and is an angle measurement by the observation device.

8. The method of claim 1, wherein, In S4, the expected trajectory deflection angle is: where V G is the velocity of the interceptor in the launch frame.

9. The method of claim 1, wherein, In S4, the sideslip angle command is: where F aero,z is the aerodynamic force in the velocity coordinate system, the side force coefficient influenced by the side slip angle, q is the dynamic pressure, and S is the reference area of the interceptor.

10. A multi-constrained guidance system for a hypersonic interceptor missile considering information loss, characterized in that, It comprises: A motion model unit, configured to establish a motion model of the interceptor, and set a plurality of constraint conditions for the motion model, the constraint conditions comprising a maximum stagnation point heat flux density constraint, a maximum dynamic pressure constraint, a maximum overload constraint, a control quantity constraint, and an attack angle constraint; the attack angle constraint is converted into terminal line-of-sight simple tracking; A tracking unit, configured to obtain a position of the target aircraft in an observation coordinate system through a tracking filter; A classification unit, configured to obtain a position, a speed, and a trajectory category of the target aircraft in a geocentric rectangular coordinate system through a trajectory identification classifier based on the position of the target aircraft in the observation coordinate system; the trajectory identification classifier is obtained through LSTM network training; A guidance unit, configured to, when enemy information is missing, obtain a shooting angle of the interceptor and the target aircraft based on the trajectory category of the target aircraft in the geocentric rectangular coordinate system obtained by the classification unit, and further, obtain an expected speed of the interceptor, and obtain an expected trajectory deflection angle through the expected speed; based on the terminal line-of-sight simple tracking, a lateral tracking guidance law is designed, and through the lateral tracking guidance law, a midcourse guidance correction of the interceptor in a lateral plane is performed; the midcourse guidance correction is to eliminate the trajectory deflection angle error through a sideslip angle command, and the interceptor is guided into the target aircraft through the midcourse guidance.

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