Terminal guidance trajectory planning method based on missile-target relative distance
By calculating the relative position and speed of the UAV and the target, the angular velocity under the line of sight is obtained, and the proportional guidance method is used for guidance trajectory planning, which solves the problem of insufficient response speed and anti-interference ability of the video seeker control strategy, and achieves It provides direct adaptation and efficient guidance to traditional flight control systems.
Patent Information
- Application Number
- CN202411488185.5
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2024-10-24
- Publication Date
- 2025-01-24
AI Technical Summary
The existing video seeker control strategy relies on image information, resulting in the system's early response too fast and the later response speed is too slow, the anti-interference ability is weak, and the recursive calculations are not possible independently, and the adaptability is poor, making it difficult to directly connect with traditional flight control systems.
By obtaining the relative position and speed of the drone and the target, the relative distance is calculated, and the angular speed of the line of sight is obtained, and the proportion guide method is used for the guide trajectory planning.
It achieves direct adaptation to conventional flight control systems, improves the response speed and stability of the guidance system, enhances anti-interference capabilities, and has wide-scale promotion value.
Smart Images
Figure CN119354201A_ABST
Abstract
Description
Technical Field
[0001] The invention relates to the technical field of unmanned aerial vehicle guidance trajectory planning, and in particular to a terminal guidance trajectory planning method based on the relative distance between a missile and a target. Background Art
[0002] The core idea of the traditional infrared / radar seeker is to detect the target, collect echoes, and obtain the changes in the target's angular velocity in the pitch and yaw directions relative to itself in the line of sight coordinate system, and then use the corresponding guidance rate to design the guidance trajectory and plan the trajectory.
[0003] However, the new visible light video seeker mainly tracks and processes the target through image recognition methods, which results in the monocular vision seeker being unable to directly obtain the line of sight angular velocity information, and can only obtain the relative distance between the target and the aircraft through the miss distance information in the pixel coordinate system. On this basis, the method of designing the real-time guidance trajectory through the miss distance information has problems such as poor adaptability to traditional flight control manufacturers and weak reliability of the new method compared with the traditional guidance rate method. Specifically: At present, the control strategy of the video seeker is relatively simple. The core idea is to plan and design the motion trajectory of the aircraft through the distance information between the target and the center point in the pixel coordinate system. On the technical level, compared with the traditional seeker, it is overly dependent on image information, and the system responds too fast in the early stage and too slow in the later stage. At the same time, it cannot perform recursive calculations independently, and has weak anti-interference ability; in terms of the market, the video image seeker has weak dominance over the traditional flight control manufacturers. There is a problem that the manufacturer's flight control cannot directly connect with the seeker, and the control algorithm needs to be improved to adapt to the seeker, which increases the development cycle and development risk points.
[0004] Therefore, how to transform the missile-target relative position information evolved from the miss distance in the pixel coordinate system into the line-of-sight angular velocity information in the line-of-sight coordinate system, and then directly adapt the conventional flight control is the research focus of technical personnel in this field. Summary of the invention
[0005] Therefore, the technical problem to be solved by the present invention is to overcome the defects existing in the above-mentioned prior art, thereby providing an unmanned aerial vehicle guidance trajectory planning method based on the relative position distance between the aircraft and the target quality inspection, and the unmanned aerial vehicle is hereinafter referred to as UAV.
[0006] A terminal guidance trajectory planning method based on the relative distance between a missile and a target, comprising: S1. According to the position and speed of the drone pre-acquired by GPS, the relative position of the missile and the target and the three-axis speed of the drone body coordinate system are obtained; S2. multiply the relative position of the missile and the three-axis velocity of the drone body coordinate system to obtain the relative angular velocity in the inertial system; S3. Based on the spatial geometric relationship, the relative position of the missile and the target is processed to obtain the pitch or yaw angle based on the inertial system line of sight; and the pitch or yaw angle of the inertial system is processed by the coordinate system conversion operation to obtain the direction cosine matrix from the inertial coordinate system i system to the line of sight coordinate system; S4. Obtain the inertial system angular velocity by using the line of sight coordinate system direction cosine matrix and the inertial system angular velocity relative angular velocity, and convert it to the line of sight system angular velocity; obtain the line of sight system pitch or yaw angular velocity based on the line of sight system angular velocity; S5. Use the pitch or yaw angular velocity of the line of sight system to perform proportional guidance.
[0007] Preferably, in step S1, obtaining the relative position of the missile and the target based on the position speed of the drone specifically includes the following steps: Converting the drone position velocity into the drone inertial system position; The inertial system position of the UAV and the inertial system position of the target are jointly solved to obtain the relative position R of the missile and the target.
[0008] Preferably, the inertial system position of the UAV and the inertial system position of the target are jointly solved to obtain the relative position R of the missile and the target. Specifically: the target position and target speed are pre-acquired according to the target kinematics; The target position and target speed in the inertial system and the missile speed and missile position used to describe the position of the UAV in the inertial system are jointly solved based on the relative kinematics of the missile and the target to obtain the relative position and relative speed of the missile and the target. .
[0009] Preferably, in step S3, the Euler angle and the line of sight angular velocity expression relative to the inertial space are obtained based on the relative motion parameters of the missile and the target, which specifically includes: Establish the sight line coordinate system; Based on the line of sight coordinate system, the inertial coordinate system rotates around the y axis Angle, rotation around the z-axis of the first coordinate system Angle, rotation around the x-axis of the second coordinate system The angle is obtained, and the direction cosine matrix in the line of sight coordinate system is obtained; among them, the inertial coordinate system rotates around the y axis The angle forms the first coordinate system, and the inertial coordinate system rotates around the z-axis of the first coordinate system Angle, forming a second coordinate system, inertia rotates around the x-axis of the second coordinate system Angle to get the sight coordinate system; Using the pre-acquired components of the relative position of the guidance target in the inertial system, the Euler angles described by the relative distances x, y, and z in the horizontal, pitch, and yaw directions between the missile and the target are extracted from the direction cosine matrix.
[0010] Preferably, and The expression is: .
[0011] Preferably, the relative angular velocity expression in the inertial system obtained based on the relative position of the missile and the target and the three-axis velocity of the drone body coordinate system is: ; represents the velocity scalar value of this system, Indicates the relative distance of the three axes in the inertial system.
[0012] Preferably, based on the relative kinematics of the projectile and the target , , And R, further deduced: The projection in the line of sight coordinate system is: ; In the formula, represents the rolling angular velocity in the line of sight, represents the pitch angular velocity in the line of sight system, represents the yaw angular velocity in the line of sight system, i represents the inertial coordinate system; Approach speed ; The direction cosine matrix of the sight coordinate system is: .
[0013] Preferably, the method further includes constructing a guidance loop: the relative position and relative speed between the projectile and the target are obtained by solving the guidance head to form a guidance loop.
[0014] The technical solution of the present invention has the following advantages: The modeling of the present invention refers to the model in the book "Flight Dynamics of Winged Missiles", and uses a conventional six-degree-of-freedom dynamics kinematics model to perform dynamic simulation modeling of the entire process of climbing-level flight-terminal guidance to verify the effect of the algorithm. The climbing and level flight stages are not described in detail, and the focus is on solving the line of sight angular velocity through the relative information of the missile and the target in the terminal guidance stage, and completing the closed-loop strike full process design through the proportional guidance method.
[0015] During the cruise operation, the seeker is turned on to search for the enemy target, and then enters the terminal guidance stage. The seeker calculates the relative position information and speed information between the missile and the target, and the flight control subsystem uses the proportional guidance method to carry out terminal guidance strikes to complete the precise strikes on the designated moving targets; the missile-target line-of-sight angular rate obtained by the method provided by the present invention is accurate and reliable, and has excellent adaptability to the proportional guidance method, and has a wide range of promotion value for converting the miss amount in the pixel coordinate system into the line-of-sight angular rate. BRIEF DESCRIPTION OF THE DRAWINGS
[0016] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.
[0017] Figure 1 It is a block diagram of the guidance loop of the present invention; Figure 2 It is a schematic diagram of the simulation results of the three-axis position of the aircraft in the inertial system at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Figure 3 It is a schematic diagram of the simulation results of the flight trajectory of the unmanned aerial vehicle and the target at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Figure 4 It is a schematic diagram of the simulation results of the target three-axis velocity information at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Figure 5 This is a schematic diagram of the simulation results of the rudder deflection angle of the servo at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Figure 6 It is a schematic diagram of the three-axis aerodynamic simulation results of the aircraft at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Figure 7 It is a schematic diagram of the simulation results of the three-axis aerodynamic torque of the aircraft at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Figure 8 It is a schematic diagram of the simulation results of the sideslip angle at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Fig. 9 It is a schematic diagram of the simulation results of the trajectory inclination / deflection angle at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000,0,0]; Fig.10 It is a schematic diagram of simulation results of expected and actual pitch angles at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Fig.11 It is a schematic diagram of the simulation results of the expected and actual yaw angles at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Fig.12It is a schematic diagram of simulation results of expected and actual rolling angles at the moment of missile launch of the present invention, with no lateral displacement and a distance from the target of [10000, 0, 0]; Fig.13 It is a schematic diagram of simulation results of expected and actual angular velocity of the missile of the present invention at the moment of launch, with no lateral displacement and at a distance of [10000, 0, 0] from the target; Fig.14 It is a schematic diagram of the simulation results of the expected and actual angular velocities of the missile of the present invention at the moment of launch, with no lateral displacement and a distance from the target of [10000, 0, 0]; Fig.15 It is a schematic diagram of simulation results of expected and actual angular velocities of the missile of the present invention at the moment of launch, with no lateral displacement and at a distance of [10000, 0, 0] from the target; Fig.16 It is a schematic diagram of the simulation results of the three-axis position of the aircraft in the inertial system at the moment of missile launch of the present invention, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig.17 The instant when the missile of the present invention is launched, the displacement in the +z direction is 300 meters, and the distance from the target is [10000, 0, 300]. The flight trajectory of the unmanned aerial vehicle and the target. Fig.18 It is a schematic diagram of the simulation results of the target three-axis velocity information at the moment of missile launch of the present invention, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig.19 This is a schematic diagram of the simulation results of the rudder deflection angle of the servo at the moment of missile launch of the present invention, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig. 20 It is a schematic diagram of the three-axis aerodynamic simulation results of the aircraft at the moment of missile launch of the present invention, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig.21 It is a schematic diagram of the simulation results of the three-axis aerodynamic torque of the aircraft at the moment of missile launch of the present invention, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig. 22 It is a schematic diagram of the simulation results of the sideslip angle of attack at the moment of missile launch of the present invention, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig.23 It is a schematic diagram of the simulation results of the trajectory inclination / deflection angle of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance from the target of [10000, 0, 300]; Fig.24It is a schematic diagram of simulation results of expected and actual pitch angles of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance of [10000, 0, 300] from the target; Fig.25 It is a schematic diagram of the simulation results of the expected and actual yaw angles of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance of [10000, 0, 300] from the target; Fig.26 It is a schematic diagram of simulation results of expected and actual rolling angles of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance of [10000, 0, 300] from the target; Fig. 27 It is a schematic diagram of the simulation results of the expected and actual angular velocity of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance of [10000, 0, 300] from the target; Fig.28 It is a schematic diagram of the simulation results of the expected and actual angular velocities of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance of [10000, 0, 300] from the target; Fig.29 It is a schematic diagram of simulation results of expected and actual angular velocity of the missile of the present invention at the moment of launch, with a displacement of 300 meters in the +z direction and a distance of [10000, 0, 300] from the target; Fig.30 A schematic diagram of the simulation results of the three-axis position of the aircraft in the inertial system at the moment of missile launch of the present invention, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.31 Schematic diagram of simulation results of the flight trajectory of the unmanned aerial vehicle and the target at the moment of missile launch of the present invention, moving at a speed of +X10° in the direction of [10000,0,0] from the target; Fig.32 Schematic diagram of simulation results of target three-axis velocity information at the moment of missile launch of the present invention, moving at a speed of +X10° direction and at a distance of [10000,0,0] from the target; Fig.33 Schematic diagram of the simulation results of the rudder deflection angle of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.34 Schematic diagram of the three-axis aerodynamic simulation results of the aircraft at the moment of missile launch of the present invention, moving at a speed of +X10° in the direction of [10000,0,0] from the target; Fig.35 Schematic diagram of the simulation results of the three-axis aerodynamic torque of the aircraft at the moment of missile launch of the present invention, moving at a speed of +X10° in the direction of [10000,0,0] from the target; Fig.36Schematic diagram of simulation results of the sideslip angle at the moment of launch of the missile of the present invention, moving at a speed of +X10° in the direction, and at a distance from the target of [10000,0,0]; Fig.37 Schematic diagram of simulation results of trajectory inclination / deflection angle of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of [10000,0,0] from the target; Fig.38 Schematic diagram of simulation results of expected and actual pitch angles of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.39 Schematic diagram of simulation results of expected and actual yaw angles of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.40 Schematic diagram of simulation results of expected and actual rolling angles of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.41 Schematic diagram of simulation results of expected and actual angular velocity of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.42 Schematic diagram of simulation results of expected and actual angular velocity of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of [10000,0,0] from the target; Fig.43 A schematic diagram of simulation results of expected and actual angular velocities of the missile of the present invention at the moment of launch, moving at a speed of +X10° in the direction of the target at a distance of [10000,0,0]; Fig.44 The figure is a flow chart of the method of the present invention. DETAILED DESCRIPTION
[0018] The technical solution of the present invention will be described clearly and completely below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.
[0019] In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicating the orientation or positional relationship, are based on the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the present invention. In addition, the terms "first", "second", and "third" are used for descriptive purposes only, and cannot be understood as indicating or implying relative importance.
[0020] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.
[0021] In addition, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.
[0022] Embodiment 1: The main function of the seeker is to obtain the relative position and speed information between the missile and the target. This information is crucial for calculating the flight path required for the missile to reach the target. In this embodiment, a method for calculating the line of sight angular velocity through the relative information of the missile and the target in the terminal stage and completing the closed-loop strike full-process design through the proportional guidance method will be disclosed.
[0023] Specifically: like Fig.44 As shown, a terminal guidance trajectory planning method based on the relative distance between the missile and the target includes: S1. According to the position and speed of the drone pre-acquired by GPS, the relative position of the missile and the target and the three-axis speed of the drone body coordinate system are obtained; S2. multiply the relative position of the missile and the three-axis velocity of the drone body coordinate system to obtain the relative angular velocity in the inertial system; S3. Based on the spatial geometric relationship, the relative position of the missile and the target is processed to obtain the pitch or yaw angle based on the inertial system line of sight; and the pitch or yaw angle of the inertial system is processed by the coordinate system conversion operation to obtain the direction cosine matrix from the inertial coordinate system i system to the line of sight coordinate system; S4. Obtain the inertial system angular velocity by using the line of sight coordinate system direction cosine matrix and the inertial system angular velocity relative angular velocity, and convert it to the line of sight system angular velocity; obtain the line of sight system pitch or yaw angular velocity based on the line of sight system angular velocity; S5. Use the pitch or yaw angular velocity of the line of sight system to perform proportional guidance.
[0024] In step S1, the relative position of the missile and the target is obtained based on the position speed of the drone, which specifically includes the following steps: Converting the drone position velocity into the drone inertial system position; The inertial system position of the UAV and the inertial system position of the target are jointly solved to obtain the relative position R of the missile and the target.
[0025] The inertial system position of the UAV and the inertial system position of the target are jointly solved to obtain the relative position R of the missile and the target. Specifically: the target position and target speed are pre-acquired according to the target kinematics; The target position and target speed in the inertial system and the missile speed and missile position used to describe the position of the UAV in the inertial system are jointly solved based on the relative kinematics of the missile and the target to obtain the relative position and relative speed of the missile and the target. .
[0026] In step S3, the Euler angle and the angular velocity expression of the line of sight relative to the inertial space are obtained based on the relative motion parameters of the missile and the target, which specifically include: Establish the sight line coordinate system; Based on the line of sight coordinate system, the inertial coordinate system rotates around the y axis Angle, rotation around the z-axis of the first coordinate system Angle, rotation around the x-axis of the second coordinate system The angle is obtained, and the direction cosine matrix in the line of sight coordinate system is obtained; among them, the inertial coordinate system rotates around the y axis The angle forms the first coordinate system, and the inertial coordinate system rotates around the z-axis of the first coordinate system Angle, forming a second coordinate system, inertia rotates around the x-axis of the second coordinate system Angle to get the sight coordinate system; Using the pre-acquired components of the relative position of the guidance target in the inertial system, the Euler angles described by the relative distances x, y, and z in the horizontal, pitch, and yaw directions between the missile and the target are extracted from the direction cosine matrix.
[0027] and The expression is: .
[0028] The relative angular velocity expression in the inertial system obtained based on the relative position of the missile and the target and the three-axis velocity of the drone body coordinate system is: ; represents the velocity scalar value of this system, Indicates the relative distance of the three axes in the inertial system.
[0029] Based on the relative kinematics of the projectile and the target , , And R, further deduced: The projection in the line of sight coordinate system is: ; In the formula, represents the rolling angular velocity in the line of sight, represents the pitch angular velocity in the line of sight system, represents the yaw angular velocity in the line of sight system, i represents the inertial coordinate system; Approach speed ; The direction cosine matrix of the sight coordinate system is: .
[0030] In addition, the present embodiment also includes constructing a guidance loop: the relative position and relative speed between the projectile and the target are obtained by calculating with the guidance head to form a guidance loop.
[0031] Example 2 Based on Example 1, this example further discloses the following contents: During the cruise operation, the seeker starts searching for the target and enters the terminal guidance phase. The seeker calculates the relative position and speed information between the missile and the target, and the flight control subsystem uses the proportional guidance method to perform terminal guidance strikes to achieve precise strikes on designated moving targets. The specific terminal guidance detailed design is as follows: (1) Relevant assumptions a) Assuming the missile has a cross-shaped aerodynamic rudder layout; b) Use roll-stabilized STT skid turn control; c) Pitch and yaw channels are fully decoupled; d) Adopting the instantaneous equilibrium assumption to achieve the moment balance of the missile at any moment; e) The aerodynamic parameters are reference values and are different from the actual XFD aerodynamic parameters; (2) Missile three-degree-of-freedom dynamic equation ; is the missile speed, Ballistic inclination angle, Ballistic angle, Speed roll angle, is the missile displacement, m is the missile mass, P is the thrust value, and g is 9.8; is the aerodynamic drag, lift and lateral force, and the specific expression is: ; represents the partial derivative of lift due to the angle of attack, represents the lift coefficient caused by the angle of attack, Y represents the lift value, q represents the lift coefficient caused by the speed, represents the windward area, represents the partial derivative of the lateral force caused by the sideslip angle, represents the lateral force coefficient caused by the sideslip angle; Assuming angle of attack and sideslip angle Less than or equal to 15°, we get , , , ,but: ; The angle of attack can be obtained by inverting and sideslip angle for: ; Indicates the velocity modulus; (3) Autopilot The autopilot is simplified by considering it as a first-order link, and its transfer function for ; here represents the autopilot transfer function term, Input items, Represents the output item, is the autopilot time constant, J and They represent the product of inertia after the conventional and first-order links, s represents the complex variable of the constant Laplace transform, is a constant coefficient, here it is taken as 1, the longitudinal and lateral autopilots are the same.
[0032] (4) Seeker measurement Since the terminal guidance is carried out by the seeker to measure the line of sight angular velocity, the Euler angle and line of sight angular velocity expressions relative to the inertial space are given for the relative motion parameters of the missile and the target. First, establish the line of sight coordinate system : Located in the center of the seeker; The axis is located on the line from the center of the seeker to the target and points to the target; Axis and The axes are orthogonal and lie in the projectile pitch plane pointing upwards; axis axis The axis forms a right-handed coordinate system, and the sight coordinate system is formed by rotating the inertial coordinate system in sequence. Axis rotation Corner, around the new Axis rotation Angle, around the final Axis rotation The angle is obtained, and the projection of the inertial coordinate system to the line of sight coordinate system is: is the direction cosine matrix, and the component of the missile target relative position R in the inertial system is: ; Subscript Represent the missile and the target respectively. So we can get: ; x, y, z represent the relative distances between the missile and the target in horizontal, pitch and yaw directions respectively.
[0033] From this we can get: ; Relative speed , based on the method of Example 1 and the above formula, the line of sight angular velocity under inertia is expressed as: ; represents the velocity scalar value of this system, Indicates the relative distance of the three axes in the inertial system.
[0034] The projection in the line of sight coordinate system is: ; Approach speed ; The dynamic tracking characteristics of the seeker are considered as the first-order inertial link ,here Take it as 0.2s. It should be noted that The approach speed is commonly seen in textbooks in this field, and c has no special definition and is simply used for distinction.
[0035] (5) Target motion equation ; ; is the position of the target inertial system, Target three-axis speed, is the target three-axis acceleration, Indicates the horizontal x-position of the target, is the target height, is the target lateral position, Relative speed in horizontal direction, Indicates the target speed in the pitch direction, Indicates the target speed information in the yaw direction.
[0036] Pitch acceleration in the autopilot output velocity coordinate system and yaw acceleration ; ; In this embodiment and Substitute with to express, the pitch acceleration and yaw acceleration Substituting in: ; ; Where g is the gravitational acceleration and is taken as 9.8; if the guidance law takes gravity compensation into account, then Remove.
[0037] (6) Based on the above content, the guidance loop block diagram is formed as follows: Figure 1 shown.
[0038] In the guidance loop, the missile velocity and missile position are output based on the missile's three-degree-of-freedom dynamic equations to represent the UAV's inertial system position; Based on the target kinematics, the target velocity and target position used to represent the target inertial system position are obtained, and the relative position of the missile and the missile is calculated jointly with the missile velocity and missile position; Output of relative position of projectile and target based on relative kinematics , , And R, further combined with the relative angular velocity in the inertial system, the inertial system is converted to the line of sight system and enters the seeker link; The longitudinal terminal guidance law and the lateral terminal guidance law in the terminal guidance law are designed in coordination with the saturation link, and the longitudinal and lateral directions of the autopilot are considered as first-order links respectively, and the output result of the autopilot is converted from the velocity system overload to the ballistic system, and the output and Perform Laplace transform inverse to obtain the angle of attack and sideslip angle; The angle of attack and sideslip angle are substituted into the missile's three-degree-of-freedom dynamic equation, and the missile velocity and missile position used to represent the position of the UAV's inertial system are output.
[0039] Summarize the technical effects of this embodiment: 1) Fixed target in trajectory 1 pitch plane At the moment of launch, the UAV and the target are in the same horizontal plane. The target has no displacement in the lateral plane of the launch coordinate system. At the moment of launch, the three-axis distance from the target is [10000, 0, 0]. The simulation results are as follows: Figure 2-15As shown, specifically: Figure 2 is the three-axis position of the aircraft in the inertial system, Figure 3 is the flight trajectory of the UAV and the target, Figure 4 is the target three-axis speed information, Figure 5 is the rudder deflection angle, Figure 6 is the three-axis aerodynamic force of the aircraft, Figure 7 is the aerodynamic moment of the aircraft along three axes, Figure 8 is the angle of attack and the sideslip angle, Fig. 9 is the ballistic inclination / deflection angle, Fig.10 is the expected and actual pitch angle, Fig.11 is the expected and actual yaw angle, Fig.12 is the expected and actual roll angle, Fig.13 are the desired and actual angular velocity of pitch, Fig.14 is the desired and actual yaw angular velocity, Fig.15 is the expected and actual angular velocity of roll; Figure 2-15 The simulation results show that for a fixed target 10,000 meters away, the UAV hits the target in 316.6 seconds, with a hit accuracy of 0.499 meters under simulation, a stability of 0.26° / s in the terminal guidance stage, and a full-process stability of 0.95° / s, which meets the use requirements.
[0040] 2) Trajectory 2: Fixed target on the lateral plane At the moment of launch, the UAV and the target are on the same horizontal plane. The target is displaced 300 meters in the +Z direction in the lateral plane of the launch coordinate system. The three-axis distance from the target at the moment of launch is [10000, 0, 300]. The simulation results are as follows Figure 16-29 As shown, specifically: Fig.16 The three-axis position of the aircraft in the inertial system, Fig.17 Unmanned aerial vehicle and target flight trajectory, Fig.18 Target three-axis speed information, Fig.19 Rudder deflection angle, Fig. 20 Aircraft three-axis aerodynamic force, Fig.21 Aerodynamic moment of the aircraft in three axes, Fig. 22 Angle of attack Side slip angle, Fig.23 Ballistic inclination / angle, Fig.24 Expected and actual pitch angles, Fig.25 Yaw desired and actual angles, Fig.26 Roll desired and actual angles, Fig. 27 Pitch desired and actual angular velocity, Fig.28 Yaw desired and actual angular velocity, Fig.29 The desired and actual angular velocity of roll. Figure 16-29The simulation results show that for a fixed target with a displacement of 300 meters outside 10,000 meters, the UAV hits the target in 303.83 seconds, with a simulation hit accuracy of 1.545 meters and a full-range stability of 0.97° / s, which meets the use requirements.
[0041] 3) Moving target in the trajectory 3 plane At the moment of launch, the UAV and the target are in the same horizontal plane. The target moves at a speed of 5 km / h in the lateral plane of the launch coordinate system along the direction with an angle of 10° to +X. At the moment of launch, the three-axis distance from the target is [10000, 0, 0]. The simulation results are as follows Figure 30-43 As shown, specifically: Fig.30 The three-axis position of the aircraft in the inertial system, Fig.31 Unmanned aerial vehicle and target flight trajectory, Fig.32 Target three-axis speed information, Fig.33 Rudder deflection angle, Fig.34 Aircraft three-axis aerodynamic force, Fig.35 Aerodynamic moment of the aircraft in three axes, Fig.36 Angle of attack Side slip angle, Fig.37 Ballistic inclination / angle, Fig.38 Expected and actual pitch angles, Fig.39 Yaw desired and actual angles, Fig.40 Roll desired and actual angles, Fig.41 Pitch desired and actual angular velocity, Fig.42 Yaw desired and actual angular velocity, Fig.43 The desired and actual angular velocity of roll is given by Figure 30-43 The simulation results show that for the low-speed moving target 10,000 meters away, the UAV hits the target in 316.2 seconds, the hit accuracy is 0.8114 meters under simulation, and the whole-process stability is 0.946° / s, which meets the use requirements.
[0042] Table 1 Comparison results of three working conditions
[0043] Through simulation demonstration, it can be seen that the missile-target line-of-sight angular rate obtained by the algorithm is accurate and reliable, has excellent adaptability to the proportional guidance method, and has large-scale promotion value. It is a feasible method to convert the miss distance in the pixel coordinate system into the line-of-sight angular rate.
[0044] Obviously, the above embodiments are merely examples for the purpose of clear explanation, and are not intended to limit the implementation methods. For those skilled in the art, other different forms of changes or modifications can be made based on the above description. It is not necessary and impossible to list all the implementation methods here. The obvious changes or modifications derived therefrom are still within the scope of protection of the invention.
Claims
1. A terminal guidance trajectory planning method based on the relative distance between the missile and the target, characterized in that: include: S1. According to the position and speed of the drone pre-acquired by GPS, the relative position of the missile and the target and the three-axis speed of the drone body coordinate system are obtained; S2. multiplying the relative position of the missile and the three-axis velocity of the drone body coordinate system to obtain the relative angular velocity in the inertial system; S3. Based on the spatial geometric relationship, the relative position of the missile and the target is processed to obtain the pitch or yaw angle from the inertial system to the line of sight system; and the pitch or yaw angle of the inertial system is processed by the coordinate system conversion operation to obtain the direction cosine matrix from the inertial coordinate system to the line of sight coordinate system; S4. using the line of sight coordinate system direction cosine matrix and the inertial system angular velocity relative angular velocity, obtain the inertial system angular velocity, and convert it to the line of sight system angular velocity; Obtaining the line-of-sight pitch or yaw angular velocity based on the line-of-sight angular velocity; S5. Use the pitch or yaw angular velocity of the line of sight system to perform proportional guidance.
2. The terminal guidance trajectory planning method according to claim 1, characterized in that: In step S1, the relative position of the missile and the target is obtained based on the position speed of the drone, which specifically includes the following steps: Converting the drone position velocity into the drone inertial system position; The inertial system position of the UAV and the inertial system position of the target are jointly solved to obtain the relative position R of the missile and the target.
3. The terminal guidance trajectory planning method according to claim 2, characterized in that: The inertial system position of the UAV and the inertial system position of the target are jointly solved to obtain the relative position R of the missile and the target. Specifically: the target position and target speed are pre-acquired according to the target kinematics; The target position and target speed in the inertial system and the missile speed and missile position used to describe the position of the UAV in the inertial system are jointly solved based on the relative kinematics of the missile and the target to obtain the relative position and relative speed of the missile and the target. .
4. The terminal guidance trajectory planning method according to claim 3, characterized in that: In step S3, the Euler angle and the angular velocity expression of the line of sight relative to the inertial space are obtained based on the relative motion parameters of the missile and the target, which specifically include: Establish the sight line coordinate system; Based on the line of sight coordinate system, the inertial coordinate system rotates around the y axis Angle, rotation around the z-axis of the first coordinate system Angle, rotation around the x-axis of the second coordinate system The angle is obtained, and the direction cosine matrix in the line of sight coordinate system is obtained; among them, the inertial coordinate system rotates around the y axis The angle forms the first coordinate system, and the inertial coordinate system rotates around the z-axis of the first coordinate system Angle, forming a second coordinate system, inertia rotates around the x-axis of the second coordinate system Angle to get the sight coordinate system; Using the pre-acquired components of the relative position of the guidance target in the inertial system, the Euler angles described by the relative distances x, y, and z in the horizontal, pitch, and yaw directions between the missile and the target are extracted from the direction cosine matrix.
5. The terminal guidance trajectory planning method according to claim 4, characterized in that: and The expression is: .
6. The terminal guidance trajectory planning method according to claim 5, characterized in that: The relative angular velocity expression in the inertial system obtained based on the relative position of the missile and the target and the three-axis velocity of the drone body coordinate system is: ； represents the velocity scalar value of this system, Indicates the relative distance of the three axes in the inertial system.
7. The terminal guidance trajectory planning method according to claim 6, characterized in that: Based on the relative kinematics of the projectile and the target , , And R, further deduced: The projection in the line of sight coordinate system is: ； In the formula, represents the rolling angular velocity in the line of sight, represents the pitch angular velocity in the line of sight system, represents the yaw angular velocity in the line of sight system, i represents the inertial coordinate system; Approach speed ; The direction cosine matrix of the sight coordinate system is: .
8. The terminal guidance trajectory planning method according to claim 7, characterized in that: It also includes constructing a guidance loop: the relative position and relative speed of the projectile and target are obtained through the seeker to form a guidance loop.