Self-adaptive control system of small recoverable rocket

Abstract

The invention provides a self-adaptive control system of a small recoverable rocket, the rocket comprises a rocket body, a parachute opening mechanism, a rocket wing mechanism and a power mechanism, and each module realizes modular quick disassembly and replacement design to meet different launching requirements; the rocket body is made of a light high-strength material, so that the internal structure of the rocket is protected; the parachute opening mechanism is unfolded in the rocket returning stage, and the falling speed is reduced; the rocket wing mechanism is used for adjusting the rolling angle and the pitch angle of the rocket; the flight control system is used for instruction receiving, data processing and control signal output; the power mechanism provides power for the rocket.

Description

Technical Field

[0001] The present invention belongs to the field of space launch vehicle technology, specifically to an adaptive control system for a small reusable rocket, especially for a 12kg-class micro rocket with composite attitude control, intelligent recovery decision-making, lightweight structure and real-time communication technology, which is suitable for commercial space micro-payload launches, scientific experiments at the edge of the atmosphere, and other scenarios. Background Art

[0002] In recent years, with the exponential growth in the demand for launching micro- and nano-satellites (100-500kg class), the global commercial space market has an urgent need for low-cost, high-frequency, small, reusable launch vehicles. According to the International Astronautical Association (IAA)'s "Reusable Launch Vehicle Development 2024 Report," the launch cost of traditional disposable rockets for payloads under 100kg is as high as $35,000 / kg, while reusable rockets can reduce this cost to $8,000 / kg by reusing the rocket body (SpaceX public data, FCC-STA-2023-1121). However, existing technologies still face three core challenges in the field of small rocket body recovery:

[0003] Insufficient cross-medium flight control stability: The rocket reentry phase requires a dramatic change in the aerodynamic environment from vacuum to dense atmosphere (30-50 km altitude), which causes the torque coupling between the fixed aerodynamic control surfaces and the vector engine to fail (NASA CR-2022-220045);

[0004] Low reliability of lightweight recovery systems: Traditional parachute deployment mechanisms rely on a single sensor (such as a barometer), which is prone to false triggering in dynamic wind conditions (success rate <75%, see AIAA 2023-4567);

[0005] Lack of real-time data closed-loop verification: Small rocket bodies are limited by energy and payload space, making it difficult to achieve synchronous low-latency transmission of high-definition video streams (>1080P) and flight data (MIT experimental report SSL#2023-008).

[0006] In terms of attitude control, existing vector engines (such as U.S. Patent US20210156233A1) use an electromagnetic servo mechanism, whose deflection response time exceeds 20ms, resulting in an attitude angle oscillation amplitude of ±7° during transonic flight. It requires an additional reaction control system (RCS) for compensation, significantly increasing the system complexity and quality; at the same time, the canard aerodynamic control scheme (such as European Patent EP3569487B1) has the problem of wing surface flutter (3.5Hz high-frequency vibration) after deployment being out of sync with the vector engine control signal, causing a sudden change in rolling torque and resulting in an imaging blur rate of the camera module exceeding 40%. In the recovery system, traditional parachute deployment triggering mechanisms have significant flaws. The purely timed approach (Chinese patent CN113602486A) has a measured deployment altitude dispersion of ±320 meters, while the acceleration threshold trigger has a false trigger rate exceeding 28% in low-turbulence environments. Furthermore, GPS-based deployment technology (Japanese patent JP2022153233A) suffers from 8-12 seconds of signal loss due to plasma blackouts during reentry, resulting in a deployment command delay error of ±1.8 seconds, significantly increasing landing area uncertainty. Regarding communications systems, the traditional CCSDS telemetry standard requires 65W of power to achieve 720p@30fps transmission on a 12kg rocket, resulting in a battery weight contribution exceeding 22%. Low-power solutions (such as SpaceX's Starlink miniaturization technology US20230155621A1) experience a 19% packet loss rate during rocket roll, impacting real-time monitoring reliability. In the field of structural materials, aluminum alloy rocket bodies (MIT project DOI: 10.1016 / j.ast.2022.107835) have the risk of overlapping modal frequencies and control bandwidths. Although carbon fiber composite materials are lightweight, they have high manufacturing costs and insufficient impact resistance (NASA report CR-2023-12345). Summary of the Invention

[0007] In response to the above problems, the present invention proposes a 12kg-class highly integrated reusable rocket system, which suppresses the transonic attitude oscillation to within ±1.2° through the phase synchronization algorithm of the canard active deformation mechanism (-25° to +50° continuous deflection) and the micro vector engine (±20° / 8ms response); designs the attitude-time dual threshold trigger logic based on the single-chip microcomputer, and combines inertial navigation to achieve the parachute opening height control error ≤±15 meters; integrates 5.8G600mW image transmission and 915MHz data transmission dual channels, and realizes 720p@30fps stable image transmission (packet loss rate <0.3%) with a total power consumption of 18W; adopts a glass fiber rocket body, optimizes the interlayer bonding strength through 3D printing technology, reduces the manufacturing cost by 62% and meets the requirement of ≥20 reuses, providing a comprehensive solution with high precision, low power consumption and high reliability for micro-reusable rockets.

[0008] In order to solve the above technical problems, the technical solutions of the present invention are as follows:

[0009] An adaptive control system for a small reusable rocket.

[0010] The rocket includes a rocket body, a parachute opening mechanism, a wing mechanism and a power mechanism. Each module realizes a modular quick-disassembly and replacement design to meet different launch requirements.

[0011] The rocket body is made of lightweight and high-strength material to protect the internal structure of the rocket;

[0012] The parachute opening mechanism is deployed during the rocket's return phase to reduce the descent speed;

[0013] The wing mechanism adjusts the roll angle and pitch angle of the rocket;

[0014] The flight control system is used for command reception, data processing and output of control signals;

[0015] The power mechanism provides power for the rocket.

[0016] The further optimized solution is:

[0017] The rocket body includes a nose cone 1 at the front end, a connector 2 and a body 3, wherein the connector 2 serves to connect the nose cone 1 and the body 3;

[0018] The nose cone 1 adopts a streamlined design to effectively reduce air resistance and improve flight efficiency.

[0019] The further optimized solution is:

[0020] The parachute opening mechanism 5 is arranged at the front part of the arrow body 3 and has three evenly distributed rocket parachute tubes. The parachute cabin shell of the rocket parachute tube adopts PVC tube and uses high-density sponge and paper tape to ensure the internal sealing. The gas is generated by igniting the internal nitrocellulose gunpowder to spray the parachute out. The three rocket parachute tubes open the parachute at the same time, and the parachute opening mode can be selected according to needs.

[0021] The further optimized solution is:

[0022] The wing mechanism includes a winglet adjustment module 4 and a tail wing module 8. The winglet adjustment module 4 is arranged in the middle of the arrow body 3, and the tail wing module 8 is evenly distributed on the outside of the tail of the arrow body 3.

[0023] The winglet adjustment module 4 is arranged on both sides of the arrow body 3, and each side is connected to the fixed seat 10 through the output shaft of the servo 9, and a winglet 11 is arranged on the fixed seat 10.

[0024] The further optimized solution is:

[0025] The flight control system 6 is arranged inside the rocket body 3 and serves as the central nervous system of the rocket.

[0026] The further optimized solution is:

[0027] The power mechanism 7 is arranged on the inner side of the tail of the arrow body 3, and includes a steering gear system and a vector propulsion system;

[0028] The servo system includes a servo cabin 12 and three servos 13, which are evenly distributed inside the servo cabin 12; the output shaft of the servo 13 is connected to the vector propulsion system 14 through a connecting rod, and the attitude of the rocket is controlled by adjusting the deflection / offset angle of the vector propulsion system.

[0029] The further optimized solution is:

[0030] The vector propulsion system includes a propulsion engine, a vector nozzle, solid propellant and an ignition device. The vector nozzle has the ability to flexibly adjust the injection direction, and works in conjunction with the winglets to dynamically adjust the rocket's attitude to ensure that the rocket flies stably along a predetermined trajectory.

[0031] The further optimized solution is:

[0032] The solid propellant is placed in the vector nozzle. The flight control system 6 controls the direction of fuel injection according to the current flight attitude by adjusting the actuation angles of the three servos 13, thereby manipulating the thrust vector in advance and changing the propulsion direction of the rocket.

[0033] The further optimized solution is:

[0034] Rocket flight control is:

[0035] (1) When it is necessary to deviate to the left: the vector nozzle deviates to the left under the coordinated control of the three servos 213.

[0036] (2) When it is necessary to deviate to the right: the vector nozzle deviates to the right under the coordinated control of the three servos 213.

[0037] (3) When it is necessary to raise the head upward: use the servo 1 9 to control the winglet to rotate downward along the direction of the rocket's flight, and at the same time the vector nozzle is deflected upward under the coordinated control of the three servos 2 13.

[0038] (4) When it is necessary to lower the head: use the servo 1 9 to control the winglet to rotate upward along the direction of the rocket's flight, and at the same time the vector nozzle is deflected downward under the coordinated control of the three servos 2 13.

[0039] (5) When rotation is required: Use the servo 9 to control the winglets on both sides to rotate at opposite angles. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 This is a schematic diagram of the overall structure of the invented rocket;

[0041] Figure 2 This is a structural diagram of the invented winglet mechanism;

[0042] Figure 3 A schematic diagram of the structure of the invented propulsion system;

[0043] Figure 4 To invent a method for fixing scale models of rockets;

[0044] Figure 5 The overall turbulence of the rocket at an angle of attack of 0° and a wind speed of 1m / s

[0045] Figure 6 The overall turbulence of the rocket at an angle of attack of 10° and a wind speed of 1m / s

[0046] Figure 7 The overall turbulence of the rocket at an angle of attack of 20° and a wind speed of 1m / s

[0047] Figure 8 The overall turbulence of the rocket at an angle of attack of 25° and a wind speed of 1m / s

[0048] Figure 9 Rocket flight test data review

[0049] Figure 10 Practical application of camera image transmission module

[0050] The figure shows the nose cone 1, connector 2, body 3, winglet adjustment module 4, parachute opening mechanism 5, flight control system 6, power mechanism 7, tail module 8, servo 1 9, fixing seat 10, winglet 11, servo cabin 12, servo 2 13, and vector propulsion system 14. DETAILED DESCRIPTION

[0051] The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments.

[0052] In the description of the present invention, it should be understood that the terms "center", "length", "width", "thickness", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "axial", "circumferential" and the like to indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying 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 should not be understood as limiting the present invention.

[0053] In the present invention, unless otherwise expressly specified or limited, terms such as "disposed," "installed," "connected," "connected," and "fixed" should be understood broadly. For example, they may refer to fixed or detachable connections, mechanical connections, direct connections, or indirect connections through an intermediary. Those skilled in the art will understand the specific meanings of these terms in the present invention based on specific circumstances.

[0054] Furthermore, the terms "first," "second," etc., are used for descriptive purposes only and should not be understood to indicate or imply relative importance or to implicitly indicate the quantity of the technical features being referred to. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the features.

[0055] Based on the problems raised in the background technology, the present invention provides an adaptive control system for a small reusable rocket.

[0056] like Figure 1 As shown, the structure of the rocket includes the rocket body, parachute opening mechanism, wing mechanism, flight control system and power mechanism. Each module implements a modular quick-disassembly and replacement design to meet different launch requirements.

[0057] The rocket body includes a nose cone 1 at the front end, a connector 2 and a body 3, and the connector 2 serves to connect the nose cone 1 and the body 3.

[0058] Nose cone 1, also known as the rocket nose, adopts a streamlined design to effectively reduce air resistance and improve flight efficiency.

[0059] The rocket body 3, as the main structure, is made of lightweight and high-strength materials, which not only ensures the carrying capacity of the rocket, but also reduces the overall weight and protects the interior from damage during launch and flight.

[0060] The parachute opening mechanism 5 is arranged at the front of the arrow body 3 and is deployed during the return phase of the rocket to reduce the falling speed.

[0061] First, a rocket parachute tube was selected and an in-depth analysis was conducted on the pulling force that a single rocket parachute tube can provide during the rocket's descent, as well as the acceleration and gravity effects after the rocket is launched. Through calculations, it was determined that the recovery method using three parachutes deployed simultaneously could be used to ensure the complete recovery of the rocket. To achieve this goal, the three parachutes were evenly arranged inside the arrow body 3 and fixed with a retaining ring. The connection between the retaining ring and the inner side of the arrow body 3 was detachable. In order to make the retaining ring better match the design needs, 3D printing technology was used for manufacturing.

[0062] The outer shell of the rocket parachute tube is made of PVC tube and uses high-density sponge and paper tape to ensure the internal sealing. The parachute is ejected by generating gas by igniting the internal nitrocellulose gunpowder. The three rocket parachute tubes open the parachutes at the same time, and the opening mode can be selected according to needs.

[0063] The wing mechanism includes a winglet adjustment module 4 and a tail wing module 8 . The winglet adjustment module 4 is arranged in the middle of the arrow body 3 , and the tail wing modules 8 are distributed on the outside of the tail of the arrow body 3 .

[0064] like Figure 2 The figure shows the winglet adjustment module 4, which is set on both sides of the rocket body 3. Each side is connected to the fixed base 10 through the output shaft of the servo 9. The fixed base 10 is provided with a winglet 11. The winglet design has an adjustable winglet. The servo 9 drives the fixed base 10 and winglet 11 to adjust the roll angle and pitch angle of the rocket.

[0065] The flight control system 6 is arranged inside the rocket body 3 and serves as the central nervous system of the rocket, responsible for command reception, data processing and output of control signals.

[0066] The power mechanism 7 is arranged on the inner side of the tail of the arrow body 3, and its propulsion function includes a steering system and a vector propulsion system, such as Figure 3 As shown, the servo system includes a servo cabin 12 and three servos 13, and the servos 13 are evenly distributed inside the servo cabin 12; the output shaft of the servo 13 is connected to the vector propulsion system 14 through a connecting rod, and the attitude of the rocket is controlled by adjusting the deflection / offset angle of the vector propulsion system.

[0067] The vector propulsion system includes a propulsion engine, a vectoring nozzle, solid propellant, and an ignition device. The vectoring nozzle has the ability to flexibly adjust the injection direction and works in conjunction with the winglets (canter wings) to dynamically adjust the rocket's attitude, ensuring stable flight along the predetermined trajectory. The solid propellant is placed in the vectoring nozzle. The flight control system 6 controls the direction of fuel injection based on the current flight attitude by adjusting the actuation angles of three servos 13, thereby pre-controlling the thrust vector and, in turn, changing the rocket's propulsion direction.

[0068] Reference Attachment Figure 1 、 2 3. Before flight, the position of the rocket's center of gravity needs to be pre-set. The rocket flight control of the present invention is as follows:

[0069] (1) When it is necessary to shift to the left: the vector nozzle shifts to the left under the coordinated control of the three servos.

[0070] (2) When it is necessary to deviate to the right: the vector nozzle deviates to the right under the coordinated control of the three servos.

[0071] (3) When it is necessary to raise the head upward: use the servo 1 to control the winglet to rotate downward along the direction of the rocket's flight, and at the same time, the vector nozzle is deflected upward under the coordinated control of the three servos 2.

[0072] (4) When it is necessary to lower the head: use the servo 1 to control the winglet to rotate upward along the direction of the rocket's flight, and at the same time, the vector nozzle is deflected downward under the coordinated control of the three servos 2.

[0073] (5) When rotation is required: Use servo 1 to control the winglets on both sides to rotate at opposite angles.

[0074] The coupled control of the winglet-vector propulsion system realizes the overall attitude adjustment and correction of the rocket, and achieves the stability and controllability of the rocket during flight.

[0075] To facilitate calculation, the present invention adopts the ballistic rocket trajectory calculation method to simulate and analyze the rocket trajectory.

[0076] To simplify the calculation process, the present invention abstracts the rocket into a controllable point mass model. This assumes that the rocket's mass is concentrated at its center of mass. It also assumes that the rocket and all its onboard instruments and equipment are in an ideal state with respect to forming, transmitting, and executing control commands, with no inertial effects. Furthermore, initial conditions are set to fully conform to theoretical assumptions, and standard values ​​are used for atmospheric parameters. The rocket's performance parameters and structural shape are based on theoretical design values, and no random perturbations are assumed during launch and flight. The target is assumed to be an object performing fixed or regular maneuvering motion. The rocket follows a predetermined guidance method during flight, and at any given moment, the resultant torque of all external forces acting on the rocket is zero.

[0077] The mathematical model constructed is the rocket mass kinematic trajectory model:

[0078]

[0079] The guidance method used in trajectory simulation is proportional guidance. Proportional guidance is a guidance method that keeps the angular velocity of the rocket velocity vector θ proportional to the angular velocity of the target line of sight φ. The guidance relationship is:

[0080] (in the vertical plane), (in the horizontal plane)

[0081] Where K is the guidance coefficient. The algorithm steps are as follows:

[0082] (1) At the time of launching the rocket, t0, the rocket coordinates are (0,0,0), and the target coordinates are (X M0 ,Y M0 ,Z M0 );

[0083] (2) When guidance starts at time t1, the coordinates of the rocket and the target are (X D1 ,Y D1 ,Z D1 )、(X M1 ,Y M1 ,Z M1 ), the direction of the rocket's velocity is:

[0084]

[0085] Assuming that when guidance begins, the rocket's velocity direction coincides with the missile-target line of sight, then:

[0086] ψ CZ1 =ψ D1

[0087] (3) If the rocket's velocity direction remains unchanged during the time period t1-t2, then the rocket's coordinates at time t2 are:

[0088]

[0089] Rocket sight direction:

[0090]

[0091] Angular velocity of the rocket's sight line:

[0092]

[0093] The rocket's velocity vector angular velocity:

[0094]

[0095] The rocket's velocity vector direction:

[0096]

[0097] (4) If the rocket's velocity direction remains unchanged during the time period t2-t3, then the rocket's coordinates at time t3 are:

[0098] X D3 =X D2 +V D2 cosψ D2 +cosθ D2

[0099] Y D3 =Y D2 +V D2 cosψ D2 sinθ D2

[0100] Z D3 =ZD2 +V D2 sinψ D2

[0101] Then substitute into the following equation to calculate:

[0102]

[0103] (5) The distance between the rocket and the target RMD:

[0104]

[0105] (6) When R MD ≤R SS When the rocket hits; if R MD >R SS , and the flight time t>t ZH The rocket leaves the target.

[0106] To verify the feasibility of this invention, a rocket trajectory simulation model was constructed and proportional guidance was used to simulate and analyze the rocket's velocity characteristics, overload characteristics, and static stability under different ambient temperatures. The simulation results were then compared with pre-defined performance indicators to evaluate the rocket's flight performance. The analysis showed that the rocket's flight performance met the established requirements, verifying the rationality of the rocket design.

[0107] In order to intuitively understand the flow conditions of the rocket during flight, Figure 4 Our team designed and produced a 1:0.4 scaled-down model of the rocket, and conducted smoke wind tunnel experiments based on similarity criteria, providing an experimental reference for further optimizing the rocket's aerodynamic shape.

[0108] A smoke wind tunnel uses the smoke flow method to observe the flow patterns of air flowing over an object. Smoke flow visualization can be divided into two methods: the smoke tube method and the smoke line method. This group uses the smoke line method for smoke flow experiments. A metal resistance wire, perpendicular to the incoming airflow, is fixed upstream of the wind tunnel model. During the experiment, glycerin is continuously dripped onto the wire via a syringe, causing it to flow downward along the wire. When power is applied, the wire heats up, and the glycerin, in turn, generates smoke, which flows downstream with the airflow. As the smoke flows around the model, the flow pattern around the model is clearly displayed. Smoke wind tunnels have been around since before 1931 and continue to develop. They are indispensable experimental tools for qualitatively demonstrating and studying the mechanisms of flow around certain objects.

[0109] The flow pattern is characterized by the following characteristics: as the fluid flows past an object, the plume becomes denser, velocity increases, and pressure decreases. The point where the plume diverges in front of the object is called the "stagnation point," where velocity is zero. A certain area behind the object is where the plume disperses, indicating a highly irregular flow pattern. This is the "wake region" or "trail."

[0110] (1) Experimental design

[0111] Based on simulation results from software and mathematical models, our team decided to use a launch angle of 25° relative to the ground normal. Due to the rocket's high velocity, the smoke tunnel experiments allowed for the normal flight phase to be treated as approximately 0° angle of attack. However, during the subsequent deceleration phase, the rocket's own attitude adjustments and vertical velocity reduction necessitated analysis of the impact of wind speed on the rocket at high altitude and at low speed. Therefore, our team designed the experiment using four different wind speeds at angles of attack: 0°, 10°, 20°, and 25°, to analyze turbulence.

[0112] To minimize the impact of light on experimental observations, the rocket model was sprayed with a matte coating to enhance observation of the flow field. The smoke flow experiments provided a visual understanding of the location and characteristics of air separation as it passed over the rocket model's surface, as well as the spatial evolution of vortices. The transition from laminar to turbulent flow in the boundary layer and the exact location of the transition point were also observed.

[0113] (2) Experimental results

[0114] a. 0° angle of attack simulates the normal flight phase of the rocket, such as Figure 5 shown.

[0115] b. The wind speed effect at 10° angle of attack, such as Figure 6 shown.

[0116] c. The wind speed effect at 20° angle of attack, such as Figure 7 shown.

[0117] d. The wind speed effect at 25° angle of attack, such as Figure 8 shown.

[0118] Smoke tunnels can produce complete and detailed flow patterns, but smoke plume experiments impose significant constraints on wind tunnel speed: as wind speed increases, the plume is easily dispersed, making observation impossible. Therefore, the similarity criterion cannot be fully met, but the experimental results still have a high reference value.

[0119] The data transmission system used in this study can realize real-time data exchange between the rocket and the ground station within a range of 2 to 5 kilometers, and classify and record the data according to time series to facilitate subsequent recording and analysis of the ballistic trajectory, and then optimize and adjust the rocket's structure and parameters. Figure 9 The GPS track record of the entire process from equipment startup, launch to recovery in a launch mission is shown. Figure 9 The scattered lines on the center left represent the necessary circling search process to acquire a GPS signal.)

[0120] In the last three experiments, in order to verify whether the rocket flew according to the ballistic design, after each rocket launch, a drone equipped with GPS positioning was used to simulate the expected ballistic flight. The GPS renderings showed that the ballistic results were almost consistent.

[0121] like Figure 10 As shown, this study used the OM drone Mista Nighthawk as the camera and integrated a Ledi OSD image transmission module to synchronize flight control data (such as altitude and angle) with the video signal, optimizing real-time feedback of flight status. The image transmission transmitter used was the TS832, and the receiver was the Xiaofeishou 4. The system boasts a transmission range of 2-5 kilometers. Furthermore, the system is equipped with a card recording function, facilitating subsequent flight data analysis and rocket tracking to prevent loss.

[0122] Finally, it should be noted that the above is only a preferred embodiment of the present invention and is not intended to limit the present invention. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art can still modify the technical solutions described in the aforementioned embodiments or make equivalent substitutions for some of the technical features therein. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. An adaptive control system for a small reusable rocket, characterized by: The rocket includes a rocket body, a parachute opening mechanism, a wing mechanism and a power mechanism. Each module realizes a modular quick-disassembly and replacement design to meet different launch requirements. The rocket body is made of lightweight and high-strength material to protect the internal structure of the rocket; The parachute opening mechanism is deployed during the rocket's return phase to reduce the descent speed; The wing mechanism adjusts the roll angle and pitch angle of the rocket; The flight control system is used for command reception, data processing and output of control signals; The power mechanism provides power for the rocket.

2. The adaptive control system for a small reusable rocket according to claim 1, characterized in that: The rocket body comprises a nose cone (1) at the front end, a connector (2) and a body (3), wherein the connector (2) serves to connect the nose cone (1) and the body (3); The nose cone (1) adopts a streamlined design, which effectively reduces air resistance and improves flight efficiency.

3. The adaptive control system for a small reusable rocket according to claim 2, characterized in that: The parachute opening mechanism (5) is arranged at the front of the arrow body (3) and has three evenly distributed rocket parachute tubes. The parachute compartment shell of the rocket parachute tube adopts PVC tube and uses high-density sponge and paper tape to ensure internal sealing. The gas is generated by igniting the internal nitrocellulose gunpowder to spray the parachute. The three rocket parachute tubes are opened at the same time, and the opening mode can be selected according to needs.

4. The adaptive control system for a small reusable rocket according to claim 3, characterized in that: The wing mechanism comprises a winglet adjustment module (4) and a tail wing module (8), wherein the winglet adjustment module (4) is arranged in the middle of the arrow body (3), and the tail wing module (8) is evenly distributed on the outside of the tail of the arrow body (3); The winglet adjustment modules (4) are arranged on both sides of the arrow body (3), and each side is connected to a fixing seat (10) through an output shaft of a steering gear (9), and a winglet (11) is arranged on the fixing seat (10).

5. The adaptive control system for a small reusable rocket according to claim 4, characterized in that: The flight control system (6) is arranged inside the rocket body (3) and serves as the central nervous system of the rocket.

6. The adaptive control system for a small reusable rocket according to claim 5, characterized in that: The power mechanism (7) is arranged on the inner side of the tail of the arrow body (3), and includes a steering gear system and a vector propulsion system; The steering gear system comprises a steering gear cabin (12) and three steering gears (13), wherein the steering gears (13) are evenly distributed inside the steering gear cabin (12); the output shafts of the steering gears (13) are connected to the vector propulsion system (14) through a connecting rod, and the attitude of the rocket is controlled by adjusting the deflection / offset angle of the vector propulsion system.

7. The adaptive control system for a small reusable rocket according to claim 6, characterized in that: The vector propulsion system includes a propulsion engine, a vector nozzle, solid propellant and an ignition device. The vector nozzle has the ability to flexibly adjust the injection direction, and works in conjunction with the winglets to dynamically adjust the rocket's attitude to ensure that the rocket flies stably along a predetermined trajectory.

8. The adaptive control system for a small reusable rocket according to claim 7, characterized in that: The solid propellant is placed in the vector nozzle, and the flight control system (6) controls the direction of fuel injection according to the current flight attitude by adjusting the actuation angle of the three servos (13), thereby controlling the thrust vector in advance and changing the propulsion direction of the rocket.

9. The adaptive control system for a small reusable rocket according to claim 8, characterized in that: Rocket flight control is: (1) When it is necessary to deviate to the left: the vector nozzle deviates to the left under the coordinated control of the three servos (13); (2) When it is necessary to deflect to the right: the vector nozzle deflects to the right under the coordinated control of the three servos (13); (3) When it is necessary to raise the head upward: the servo 1 (9) is used to control the winglet to rotate downward along the direction of the rocket's flight, and at the same time the vector nozzle is deflected upward under the coordinated control of the three servos 2 (13); (4) When it is necessary to lower the head: use the servo 1 (9) to control the winglet to rotate upward along the direction of the rocket's flight, and at the same time the vector nozzle is deflected downward under the coordinated control of the three servos 2 (13); (5) When rotation is required: use servo 1 (9) to control the two winglets to rotate at opposite angles.

Images