Two-dimensional trajectory correction fuze device for empennage stabilizing projectile

Abstract

The two-dimensional trajectory correction fuze device comprises a wave-transparent cover, a base, a control cabin, a rotary cross-shaped rudder and a battery cabin which are sequentially arranged in the direction from a bullet head to a bullet tail, an electric steering engine is arranged in the rotary cross-shaped rudder, and a radar module is arranged between the base and the wave-transparent cover; a flight control board, a trajectory resolving board and an attitude measurement module are arranged in the control cabin, a navigation module is embedded in the outer wall of the control cabin, a control module is loaded in the flight control board, the navigation module is used for detecting coordinate parameters and speed parameters, the radar module is used for measuring the distance between a warhead and the ground surface, and the attitude measurement module is used for detecting attitude parameters. The trajectory resolving board is used for recognizing aerodynamic parameters, and the control module is used for generating a steering engine control instruction and controlling the electric steering engine to adjust the aerodynamic force vector of the rotary cross rudder according to the steering engine control instruction to conduct trajectory correction. The two-dimensional trajectory correction fuze device has the beneficial effects that the two-dimensional trajectory correction fuze device can be quickly launched, adapts to complex meteorological conditions, is compact in structure and can be interchanged with a standard fuze.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft trajectory correction, and in particular to a two-dimensional trajectory correction fuze device for a tail-stabilized projectile. Background Art

[0002] Traditional fin-stabilized missiles lack guidance capabilities, rely on initial launch parameters and fixed ballistic designs, and are difficult to adapt to complex and changing weather conditions and initial errors, resulting in low hit accuracy. Therefore, in the use environment, a large amount of ammunition is usually consumed, which puts a lot of pressure on logistics and is difficult to adapt to complex and changing weather environments. In order to improve the hit rate, some technical routes use multi-channel deflectable rudders or pulse engines, which require a redesign of the entire system, and cannot be upgraded to low-cost standard ammunition. In addition, it cannot be interchangeable with standard fuses.

[0003] In addition, current flight control technology often requires meteorological parameters, aerodynamic parameters, preset trajectory, and magnetic field parameters to be set before launch. Each parameter requires theoretical calculation and pre-launch measurement, and there will be a large deviation between the theoretical calculation and the actual flight parameters. Pre-launch measurement requires the arrangement of measurement equipment in advance, and the preparation time is too long. In summary, the existing trajectory flight control technology has poor adaptability to complex and changeable meteorological environments, and the launch preparation time is too long, making it difficult to achieve efficient and real-time trajectory adjustment. Summary of the invention

[0004] The technical problem to be solved by the present invention is to provide a two-dimensional trajectory correction fuze device for tail-stabilized projectiles, which can be launched quickly, adapt to complex meteorological conditions, has a compact structure, and is interchangeable with standard fuzes. In order to overcome the defects of the above-mentioned prior art (or related technology), the present invention provides a two-dimensional trajectory correction fuze device for tail-stabilized projectiles.

[0005] The present invention provides a two-dimensional trajectory correction fuze device for a tail-stabilized projectile, comprising a wave-transmitting cover, a base, a control cabin, a rotating cross rudder and a battery cabin, which are arranged in sequence from the warhead to the tail of the warhead, the two ends of the base are respectively fixedly connected to the wave-transmitting cover and the control cabin, the two ends of the rotating cross rudder are respectively fixedly connected to the control cabin and the battery cabin, an electric steering gear is arranged inside the rotating cross rudder, and the electric steering gear is meshed with an internal gear of an inner ring of the rotating cross rudder through a planetary gear, a radar module is arranged between the base and the wave-transmitting cover, a flight control board, a trajectory solver board and an attitude measurement module are arranged in the control cabin, a navigation module is embedded on the outer wall of the control cabin, a control module is loaded in the flight control board, and the control module is respectively connected to the electric steering gear. The aircraft, the radar module, the trajectory calculation board, the attitude measurement module and the navigation module are electrically connected, the navigation module is used for real-time detection to obtain coordinate parameters and speed parameters, the radar module is used for real-time measurement of the distance between the warhead and the ground surface, and outputs a valid signal after the distance from the ground is less than a preset threshold, the attitude measurement module is used for real-time detection to obtain attitude parameters, the trajectory calculation board is used for real-time identification of aerodynamic parameters during flight, the control module is used for obtaining a steering gear control instruction according to the coordinate parameters, the speed parameters, the attitude parameters and the aerodynamic parameters after receiving the valid signal, and controls the electric steering gear according to the steering gear control instruction to adjust the aerodynamic force vector of the rotating cross rudder to perform trajectory correction.

[0006] Compared with the prior art, the two-dimensional trajectory correction fuze device for a fin-stabilized projectile of the present invention has the following advantages: In the present invention, a radar module, a flight control board, a trajectory calculation board, a posture measurement module and a navigation module are arranged in the missile body. The radar module is used to realize the real-time detection function of the distance between the warhead and the ground, the flight control board is used to realize the overall control function, the trajectory calculation board is used to realize the real-time detection function of the aerodynamic parameters, the posture measurement module is used to realize the real-time detection function of the posture parameters, the navigation module is used to realize the real-time monitoring function of the coordinate parameters and the speed parameters, and the control module is used to calculate the steering gear control instruction to control the electric steering gear to adjust the aerodynamic force vector of the rotating cross rudder to perform trajectory correction. Compared with the existing guidance scheme, the launch preparation time is shorter and the links are fewer, and rapid launch can be realized. In the present invention, the steering gear control instruction can be obtained by calculating the real-time moving trend of the target point error according to the coordinate parameters, the speed parameters, the posture parameters and the aerodynamic parameters, so that the meteorological disturbance is compensated, and the present invention has higher adaptability to complex meteorological conditions. At the same time, the overall structural design is more compact, and it has interchangeability with the traditional standard fuze, and it can be quickly used for the guidance upgrade of conventional ammunition.

[0007] In a possible implementation, the radar module includes an induction coil and a millimeter-wave radar, the millimeter-wave radar is fixed on the base, and the induction coil is sleeved on the outside of the millimeter-wave radar.

[0008] Compared with the existing technology, the above technical solution can realize real-time detection of the distance of the warhead from the ground through the millimeter wave radar, and output the effective signal to the control module in time through the signal line to ensure the timely transmission of the signal.

[0009] In a possible implementation, the navigation module includes a satellite receiver and two satellite antennas, the two satellite antennas are embedded on the outer wall of the control cabin, and the satellite receiver is disposed in the control cabin and electrically connected to the two satellite antennas.

[0010] In a possible implementation manner, the attitude measurement module, the trajectory solver board, the flight control board, and the satellite receiver are sequentially stacked and spaced from the warhead to the tail.

[0011] Compared with the prior art, the above technical solution can reduce the installation space inside the projectile, making the overall structure more compact.

[0012] In a possible implementation, two antenna slots are provided on the outer wall of the control cabin for fixedly installing the two satellite antennas, and antenna covers are provided at the slot openings of the two antenna slots.

[0013] In a possible implementation, the electric servo is externally sleeved with two deep groove ball bearings, the inner walls of the two deep groove ball bearings are fixedly connected to the outer wall of the electric servo, and the outer walls of the two deep groove ball bearings are interference fit with the inner wall of the rotating cross rudder.

[0014] In a possible implementation manner, an aluminum ring is provided between the two deep groove ball bearings, and the aluminum ring is sleeved on the outside of the electric steering gear.

[0015] In a possible implementation manner, a compression screw is provided between the deep groove ball bearing close to the wave-transmitting cover and the control cabin, and the compression screw sleeve is provided on the outside of the electric steering gear.

[0016] Compared with the prior art, the above technical solution can fix the control cabin and the deep groove ball bearing by pressing the screws, thereby axially fixing the deep groove ball bearing and avoiding axial displacement.

[0017] In a possible implementation, a disc spring is provided between the aluminum ring and the two deep groove ball bearings, and the two disc springs are sleeved on the outside of the electric servo.

[0018] Compared with the prior art, the above technical solution can reduce the vibration of the two deep groove ball bearings through disc springs, improve the control effect and control accuracy of the electric servo on the rotating cross rudder, and reduce the error impact caused by vibration.

[0019] In a possible implementation, a lithium battery and an overload switch are provided in the battery compartment. The overload switch is electrically connected to the lithium battery and the control module, respectively, and is used to close to power on the control module after receiving a launch impact.

[0020] Compared with the existing technology, the above technical solution can control the power-on through the overload switch, reasonably distribute the lithium battery power, and improve the overall operating endurance. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 It is a schematic diagram of the overall structure of the present invention; Figure 2 It is a schematic diagram of the explosion structure of the present invention; Figure 3 It is a cross-sectional structural schematic diagram of the present invention; Figure 4 It is a cross-sectional structural schematic diagram of the steering gear assembly of the present invention; Figure 5 It is a schematic diagram of the system framework of the present invention; Figure 6 The Huygens bridge schematic diagram of the attitude measurement module of the present invention; Figure 7 A schematic diagram of the magnetic field parameter calibration process during flight of the present invention; Figure 8 It is a schematic diagram of the pneumatic parameter identification process of the present invention; Explanation of the reference numerals: 1. battery compartment; 2. millimeter-wave radar; 3. base; 4. control compartment; 5. satellite antenna; 6. antenna cover; 7. wave-transparent cover; 8. electric servo; 9. aluminum ring; 10. pressure screw; 11. deep groove ball bearing; 12. disc spring; 13. rotating cross rudder; 14. fuze security mechanism; 15. lithium battery; 16. battery compartment cover; 17. overload switch; 18. induction coil; 19. attitude measurement module; 20. satellite receiver; 21. trajectory solver board; 22. flight control board. DETAILED DESCRIPTION

[0022] First, those skilled in the art should understand that these implementations are only used to explain the technical principles of the embodiments of the present invention, and are not intended to limit the protection scope of the embodiments of the present invention. Those skilled in the art can make adjustments to them as needed to adapt to specific application scenarios.

[0023] The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0024] See also Figure 1-Figure 8 The embodiment of the present invention discloses a two-dimensional trajectory correction fuze device for a tail-stabilized projectile. The device can realize longitudinal and lateral correction of the trajectory by only using a single-channel servo. The device has a compact structure and can be interchanged with standard fuzes. The device can upgrade traditional ammunition to guided ammunition without redesigning the projectile body. The flight control system adopted has an online aerodynamic parameter identification function, which can collect various parameters during flight and obtain the optimal aerodynamic parameters through iterative successive approximation, thereby improving the accuracy and adaptability of trajectory calculation. During flight, magnetic field data is collected in real time, and parameter identification is performed through MP generalized inverse calculation to obtain calibration parameters of the magnetic field. No calibration is required before launch, thereby improving launch efficiency. The moving trend away from the target is indirectly measured through efficient trajectory calculation, and the angle of the aerodynamic wing surface is changed to align the trajectory with the target. The device has stronger adaptability to complex meteorological conditions than traditional guidance laws.

[0025] Continue to see Figure 2 and Figure 3 In the two-dimensional trajectory correction fuze device of the tail-stabilized projectile of the embodiment of the present invention, the induction coil 18 is sleeved on the outer circle of the millimeter-wave radar 2, and the bottom is adhered to the upper surface of the control cabin 4 with epoxy resin. The millimeter-wave radar 2 is fixed to the end surface of the control cabin 4 with screws in the form of a flange. The flight control board 22 provides power and communication interface for the millimeter-wave radar 2. The induction coil 18 serves as the antenna of the near-field induction setting module to power the near-field induction setting module and transmit wireless data. The near-field induction setting module is integrated on the flight control board 22. The millimeter-wave radar 2 and the induction coil 18 are covered with wave-transparent material on the outside. The wave-transparent cover 7 is connected to the control cabin 4 by threads, and the intermediate gap is filled with foam material.

[0026] Continue to see Figure 2 and Figure 3 The control cabin 4 is made of aluminum alloy, and two symmetrical antenna slots are opened on the side wall along the circumferential direction for installing the satellite antenna 5. The satellite antenna 5 is fixed in the antenna slot by screws. The outside of the satellite antenna 5 is covered by an antenna cover 6. The antenna cover 6 is a wave-transmitting material, and the gap is filled with foam glue; the attitude measurement module 19, the trajectory solver board 21, the flight control board 22 and the satellite receiver 20 are stacked and placed, and the circuit boards communicate through wires. The attitude measurement module 19, the trajectory solver board 21, and the satellite receiver 20 are all powered by the flight control board 22. The gap in the control cabin 4 is filled with epoxy resin. The attitude measurement module 19 sends the measured and calculated attitude parameters to the control module in the flight control board 22 at a regular interval. The trajectory solver board 21 receives the coordinate parameters and speed parameters sent by the satellite receiver 20, and sends the trajectory end point to the control module in the flight control board 22 after running the trajectory equation solver.

[0027] Continue to see Figure 4The steering gear assembly consists of a rotating cross rudder 13 with a fixed deflection angle, a deep groove ball bearing 11, a disc spring 12, an aluminum ring 9, a pressure screw 10 and an electric steering gear 8. The inner ring of the rotating cross rudder 13 with a fixed deflection angle is provided with an internal gear, which meshes with the planetary gear of the electric steering gear 8. The inner ring of the deep groove ball bearing 11 is connected to the outer circle of the electric steering gear 8, and the outer ring is connected to the rotating cross rudder 13 with a fixed deflection angle through an interference fit. The disc spring 12 is used to reduce vibration between the two deep groove ball bearings 11, and the pressure screw 10 axially fixes the deep groove ball bearings 11. The control module in the flight control board 22 sends a steering gear control instruction to the electric steering gear 8. The electric steering gear 8 drives the rotation of the aerodynamic wing surface on the outer surface of the rotating cross rudder 13 through gear transmission, thereby adjusting the aerodynamic force vector for trajectory correction.

[0028] Continue to see Figure 2 and Figure 3 The battery compartment 1 contains a lithium battery 15, a fuze security mechanism 14, and an overload switch 17. The bottom of the battery compartment 1 is a battery compartment cover 16. The gap in the compartment is filled with epoxy resin, and then the battery compartment cover 16 is fixed to the battery compartment 1 by threads. The overload switch 17 is closed after the launch impact, and the fuze security mechanism 14 is powered on. The lithium battery 15 supplies power to the electric servo 8 and the flight control board 22 respectively. The flight control board 22 is stepped down by a low-voltage difference linear regulator to supply power to the satellite receiver 20, the trajectory solver board 21, and the attitude measurement system 19. The fuze security mechanism 14 receives the signal of the flight control board 22 and then acts.

[0029] Before launching, the external wireless induction locator activates the near-field induction setting module in the missile body through the induction coil 18, transmits data to the near-field induction setting module and stores it in the internal memory; after launching, the flight control board 22 is powered on, and the flight control board 22 sends a read instruction to the near-field induction setting module. The near-field induction setting module sends various parameters for setting, including target coordinates, meteorological parameters, ephemeris data, etc. After the flight control board 22 successfully reads the ephemeris data, it forwards the ephemeris data to the satellite receiver 20. In the period of time after launching when the satellite is not positioned, the attitude measurement module 19 reads the original value of the magnetic field during the flight.

[0030] Continue to see Figure 6The core part of the attitude measurement module 19 is two groups of orthogonally arranged Wheatstone bridges. Each group of Wheatstone bridges consists of four anisotropic magnetoresistive elements (AMR), namely R1, R2, R3, R4, R5, R6, R7 and R8. The magnetization direction is shown in the figure. After applying an external magnetic field to the bridge, the resistance value of each magnetoresistive element changes, so that the output positive and output negative terminal voltages change. Within a certain range, the external magnetic field strength is linearly related to the terminal voltage. The attitude measurement module 19 amplifies and converts this voltage into analog-to-digital form, calculates the corresponding XY axis magnetic field strength, and then calculates the Z axis magnetic field strength based on the strength of the local geomagnetic field, and then obtains the three-dimensional geomagnetic field vector. In the absence of other magnetic field interference, theoretically, this geomagnetic field vector should be distributed on a uniform spherical surface, but because there is always hard magnetic interference and soft magnetic interference outside, the execution Figure 7 The calibration algorithm shown can obtain the calibration parameters of the geomagnetic field. After calibration, the measured raw data can obtain accurate magnetic field vectors for attitude solution. After launch, the system is powered on to collect more than 100 groups of effective magnetic field data to form a sample space matrix. The MP generalized inverse of the sample space matrix is ​​calculated to obtain the least squares solution of the calibration parameters. After obtaining the calibration parameters, the attitude measurement module 19 translates, rotates and scales the read sensor raw values ​​to finally obtain the calibrated magnetic field measurement values. The final attitude parameters can be calculated based on the flight trajectory inclination, azimuth and local geomagnetic field invariant vector.

[0031] Continue to see Figure 8 Before the satellite positioning in the ascending phase of the trajectory, the flight control board 22 drives the electric servo 8 to rotate at a constant speed to eliminate the trajectory correction force. When the satellite is positioned for the first time, the flight control board 22 records the current position and current speed at the current moment as the starting point of the trajectory calculation. When the trajectory reaches the top, the trajectory solution board 21 sets an initial aerodynamic drag coefficient according to the above starting point, and calculates the predicted coordinates of reaching the top of the trajectory according to the meteorological parameters stored in the near-field sensing setting module, and compares the calculated predicted coordinates with the actual coordinates measured by the satellite receiver 20, and uses the deviation between the two as the basis for correcting the drag coefficient. If the calculated value is far from the measured value, the drag coefficient is increased, and if it is close, the drag coefficient is reduced, so as to obtain the drag coefficient correction amount, and then the corrected drag coefficient is input into the trajectory solution board 21 for continued calculation, and the above process is iterated repeatedly until the above deviation converges to the threshold range, and the final effective aerodynamic drag coefficient is obtained.

[0032] During the ballistic descent stage, the trajectory solver 21 calculates the moving speed of the theoretical landing point in real time based on the above-mentioned effective aerodynamic drag coefficient and the meteorological parameters stored in the near-field sensing setting module, taking the coordinates and speed measured by the navigation module as initial conditions, and calculates the vector of the moving speed, the line connecting the theoretical landing point and the target point, and the angle between the two vectors. This angle is used as the input of the control module (i.e., the angular increment of the servo), driving the rotating cross rudder 13 with a fixed deflection angle to generate aerodynamic lift to change the trajectory. The rotation of the electric servo 8 controls the direction of the lift generated by the rotating cross rudder 13. After the servo driver of the electric servo 8 receives the above-mentioned angular increment instruction, i.e., the servo control instruction, it controls the electric servo 8 to rotate, changes the lift direction of the aerodynamic wing surface, and thus achieves the purpose of correcting the trajectory.

[0033] In the description of the present invention, the description with reference to the terms "one embodiment", "some embodiments", "in the present embodiment", "specific example", or "some examples" etc. means that the specific features, mechanisms, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, mechanisms, materials or characteristics described may be combined in a suitable manner in any one or more embodiments or examples. In addition, those skilled in the art may combine and combine the different embodiments or examples described in this specification and the features of the different embodiments or examples, without contradiction.

[0034] The above is only a specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed by the present invention should be included in the protection scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.

Claims

1. A two-dimensional trajectory correction fuze device for a fin-stabilized projectile, characterized in that: The invention comprises a wave-transmitting cover, a base, a control cabin, a rotating cross rudder and a battery cabin which are arranged in sequence from the warhead to the tail of the warhead, the two ends of the base are respectively fixedly connected to the wave-transmitting cover and the control cabin, the two ends of the rotating cross rudder are respectively fixedly connected to the control cabin and the battery cabin, an electric steering gear is arranged in the rotating cross rudder and the electric steering gear is meshed with the internal gear of the inner ring of the rotating cross rudder through a planetary gear, a radar module is arranged between the base and the wave-transmitting cover, a flight control board, a ballistic solution board and an attitude measurement module are arranged in the control cabin, a navigation module is embedded on the outer wall of the control cabin, a control module is loaded in the flight control board, and the control module is respectively connected to the electric steering gear, the radar module, the trajectory calculation board and the attitude measurement module. The trajectory solver board, the attitude measurement module and the navigation module are electrically connected, the navigation module is used for real-time detection to obtain coordinate parameters and speed parameters, the radar module is used for real-time measurement of the distance between the warhead and the ground surface, and outputs a valid signal after the distance from the ground is less than a preset threshold, the attitude measurement module is used for real-time detection to obtain attitude parameters, the trajectory solver board is used for real-time identification of aerodynamic parameters during flight, the control module is used for obtaining a servo control instruction according to the coordinate parameters, the speed parameters, the attitude parameters and the aerodynamic parameters after receiving the valid signal, and controlling the electric servo to adjust the aerodynamic force vector of the rotating cross rudder to perform trajectory correction according to the servo control instruction.

2. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 1, characterized in that: The radar module includes an induction coil and a millimeter wave radar. The millimeter wave radar is fixed on the base, and the induction coil is sleeved on the outside of the millimeter wave radar.

3. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 1, characterized in that: The navigation module includes a satellite receiver and two satellite antennas. The two satellite antennas are embedded on the outer wall of the control cabin. The satellite receiver is arranged in the control cabin and is electrically connected to the two satellite antennas.

4. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 3, characterized in that: The attitude measurement module, the trajectory calculation board, the flight control board and the satellite receiver are sequentially stacked and spaced from the warhead to the tail.

5. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 3, characterized in that: Two antenna slots are provided on the outer wall of the control cabin for fixedly installing the two satellite antennas, and antenna covers are provided at the slot openings of the two antenna slots.

6. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 1, characterized in that: The outer sleeve of the electric steering gear is provided with two deep groove ball bearings, the inner walls of the two deep groove ball bearings are fixedly connected to the outer wall of the electric steering gear, and the outer walls of the two deep groove ball bearings are interference fit with the inner wall of the rotating cross rudder.

7. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 6, characterized in that: An aluminum ring is arranged between the two deep groove ball bearings and the aluminum ring is sleeved on the outside of the electric steering gear.

8. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 6, characterized in that: A compression screw is arranged between the deep groove ball bearing close to the wave-transmitting cover and the control cabin, and the compression screw sleeve is arranged outside the electric steering gear.

9. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 7, characterized in that: Disc springs are arranged between the aluminum ring and the two deep groove ball bearings, and the two disc springs are sleeved on the outside of the electric servo.

10. The two-dimensional trajectory correction fuze device for a fin-stabilized projectile according to claim 1, characterized in that: A lithium battery and an overload switch are provided in the battery compartment. The overload switch is electrically connected to the lithium battery and the control module respectively, and is used for closing to power on the control module after receiving a launch impact.

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