Folding-wing unmanned aerial vehicle and miniature carrier rocket for launching folding-wing unmanned aerial vehicle

Abstract

The utility model relates to the technical field of unmanned aerial vehicles and launching of the unmanned aerial vehicles, in particular to a folding wing unmanned aerial vehicle and a miniature carrier rocket for launching of the folding wing unmanned aerial vehicle. The rocket comprises a power assembly, a parachute cabin is fixedly installed above the power assembly, an avionics cabin is fixedly installed above the parachute cabin, a variable-diameter connecting piece is fixedly installed above the avionics cabin, a load cabin is fixedly installed above the variable-diameter connecting piece, and a nose cone is fixedly installed above the load cabin. The unmanned aerial vehicle comprises a fuselage, an aircraft head used for fixing a battery and a receiver is fixedly installed at the front end of the fuselage, a wing fixing mechanism is fixedly installed in the middle of the fuselage, rotatable wings are installed above the wing fixing mechanism, and an empennage is fixedly installed at the rear end of the fuselage. The unmanned aerial vehicle launching process is simplified, the problem that control is complex due to the fact that a traditional rocket boosted unmanned aerial vehicle deviates in the thrust line and the gravity center of the unmanned aerial vehicle changes in the launching process is solved, the unmanned aerial vehicle launching efficiency is improved, the structure is simple, and operation is convenient.

Description

Technical Field

[0001] This utility model relates to the field of unmanned aerial vehicles (UAVs) and their deployment technology, and in particular to a folding-wing UAV and a miniature launch vehicle for its deployment. Background Technology

[0002] Folding-wing UAVs can fold their components through a folding mechanism, thereby effectively reducing their size and enabling them to be launched or deployed from various weapon platforms. They can perform one or more tasks such as reconnaissance and damage assessment, communication relay, target designation, and precision strikes. They are characterized by low cost, high cost-effectiveness, small size, and strong stealth capabilities.

[0003] Currently, the main launch methods for drones include rail-assisted launch, rocket-assisted launch, and launch tube launch. Traditional rocket-assisted launch uses a booster rocket mounted on the aircraft to provide thrust, which is automatically jettisoned after takeoff. However, due to assembly errors and other issues, rocket-assisted launch suffers from thrust line misalignment, significantly impacting takeoff safety with launch parameters. The launch process also exhibits complex dynamic responses, placing high demands on the control system. Furthermore, installing a booster rocket shifts the drone's center of gravity significantly rearward, reducing its static stability. While launch tube launch facilitates transportation and storage, its valve control methods and piping systems are complex, the power source cannot be stored long-term, and reliable power output cannot be guaranteed.

[0004] This invention proposes a folding-wing UAV and a launch vehicle system for its deployment. By installing the UAV inside a miniature launch vehicle, the UAV is transported to a designated airspace for deployment by the rocket. This avoids problems such as thrust line deviation and changes in the UAV's center of gravity during deployment that exist in traditional rocket-assisted launches, reduces the requirements for the control system, and improves the system's reliability. At the same time, the proposed system has the advantages of easy storage and transportation, avoiding the complex valve control and pipeline equipment of launch tube ejection, and providing a reliable power system for the UAV. Summary of the Invention

[0005] The purpose of this invention is to overcome the shortcomings of existing technologies by providing a folding-wing drone and a micro launch vehicle for its deployment, which solves the problems of complex structure, cumbersome operation, and insufficient power source in existing drone deployment technologies.

[0006] To achieve the above objectives, the technical solution adopted by this utility model is as follows: a folding wing drone, including a fuselage, with the nose of the aircraft fixedly installed at the front end of the fuselage, a battery installed above the nose of the aircraft, and a receiver installed below the nose of the aircraft.

[0007] A wing mount is fixedly installed in the middle of the fuselage, a tail fin is fixedly installed at the rear of the fuselage, a wing connector is installed above the wing mount, wings are symmetrically installed on both sides of the wing connector, and servos are fixedly installed on the wings.

[0008] Furthermore, the wing mount and the wing connector are connected by washers and screws; the screws pass through the through holes on the wing mount and the wing connector to achieve connection, and the wing connector rotates around the screw while driving the entire wing to rotate. The washers are placed on the contact surface between the wing connector and the wing mount to reduce friction during the rotation process.

[0009] Furthermore, the wing connector is fixed to the fuselage by a spring. When the wing rotates and folds, the spring is in a stretched state. When the wing is fully extended, the spring returns to its initial state. The wing connector is connected to a limiting slot on the wing mount to achieve limiting and fixing.

[0010] Furthermore, the outer layer of the wing is a composite material made of carbon fiber and epoxy resin, and the interior of the wing is filled with foam. The wing adopts the NACA3412 airfoil, with a wing chord length of 5cm and a wingspan of 44-45cm. It adopts an upward dihedral mounting method, which has good gliding performance. The battery and receiver adopt a detachable structure, and the tail fin adopts an integrated structure of vertical tail and horizontal tail.

[0011] The present invention also provides a miniature launch vehicle for launching the folding-wing UAV as described above, comprising a power assembly, a solid rocket engine fixedly installed inside the power assembly, a parachute compartment fixedly installed above the power assembly, a deceleration parachute disposed inside the parachute compartment, an avionics compartment fixedly installed above the parachute compartment, a variable diameter connector fixedly installed above the avionics compartment, the variable diameter connector being fixedly connected to a payload compartment above the payload compartment, and a fairing fixedly installed above the payload compartment.

[0012] Furthermore, the payload compartment includes an upper payload compartment and a lower payload compartment. The lower payload compartment is fixedly installed above a lower payload compartment connector. The lower payload compartment connector is fixedly connected to the upper payload compartment connector on both sides via hinges. The upper payload compartment is fixedly connected to the upper payload compartment connector. An avionics bracket for fixing and installing the avionics system board is fixedly installed inside the avionics compartment. An avionics switch is installed on the variable diameter connector.

[0013] Furthermore, the solid rocket motors are configured to be four in number, and the deceleration parachutes are configured to be two 50cm diameter deceleration parachutes, which are symmetrically fixed inside the parachute compartment.

[0014] Furthermore, two second explosive bolts are symmetrically fixedly installed on the upper part of the parachute compartment. The second explosive bolts are connected to the parachute compartment cover by screws. Under the control of the avionics system board, the two explosive bolts are ignited at the same time. The upper end of the parachute compartment cover flips outward under the action of the explosive bolts. The guide rope between the parachute compartment cover and the deceleration parachute pulls the deceleration parachute out of the parachute compartment. The deceleration parachute is then inflated to achieve the deceleration effect, thus realizing rocket recovery.

[0015] The fairing includes a first half-fairing and a second half-fairing. The surface area of ​​the first half-fairing is larger than that of the second half-fairing. The first half-fairing and the second half-fairing are fixedly connected by screws and a first explosive bolt. When the folding-wing UAV is launched, the first explosive bolt is ignited, and the first half-fairing and the second half-fairing separate under the action of the first explosive bolt, causing the payload compartment to rotate and open around the hinge, and the folding-wing UAV separates from the rocket.

[0016] Furthermore, the upper part of the payload compartment is composed of two semi-cylinders, and the first semi-fairing and the second semi-fairing are respectively fixedly installed on the two semi-cylinders constituting the upper part of the payload compartment. The payload compartment is made of a composite material of carbon fiber and epoxy resin.

[0017] Furthermore, the avionics bracket is also equipped with a lithium battery that supplies power to the avionics system board; the avionics switch includes a main avionics switch and an avionics safety switch.

[0018] The beneficial effects of this utility model are as follows: The folding-wing UAV and the micro launch vehicle used for its deployment allow the UAV's wings to be rotated and then inserted into the rocket's payload bay, enabling convenient and quick remote deployment of the UAV and rapid deployment of a large-scale UAV swarm. This simplifies the UAV deployment process and avoids problems such as thrust line deviation and control complexity caused by changes in the UAV's center of gravity during deployment, as well as the complex launch tube ejection system and insufficient power source of traditional rocket-assisted UAVs. It also improves the efficiency of UAV deployment, has a simple structure, and is easy to operate. Attached Figure Description

[0019] Figure 1 This is an isometric structural diagram of the folding-wing drone of this utility model;

[0020] Figure 2 This is a schematic diagram of the wing mounting structure of the folding-wing UAV of this utility model;

[0021] Figure 3 This is a front view schematic diagram of the micro launch vehicle of this utility model used for launching folding-wing drones;

[0022] Figure 4 This is a partial structural schematic diagram of the micro launch vehicle of this utility model used for launching folding-wing drones;

[0023] Figure 5 This is a schematic diagram of the connection structure between the fairing and the first explosive bolt of the micro launch vehicle used for launching folding-wing UAVs according to this utility model.

[0024] Figure 6 This is a partial structural diagram of the payload compartment of the micro launch vehicle used for launching folding-wing drones according to this utility model;

[0025] Figure 7 This is a schematic diagram of the folding-wing drone when it is deployed according to this utility model, with the drone's wings not unfolded.

[0026] Figure 8 This is a schematic diagram of the folding-wing drone when it is deployed according to this utility model.

[0027] 1. Powertrain; 2. Parachute compartment; 3. Avionics compartment; 4. Variable diameter connector; 5. Lower part of payload compartment; 6. Lower part of payload compartment connector; 7. Hinge; 8. Upper part of payload compartment connector; 9. Upper part of payload compartment; 10. First half fairing; 11. Second fairing; 12. Solid rocket motor; 13. Parachute compartment cover; 14. Avionics support frame; 15. Avionics switch; 15-1. Main avionics switch; 15-2. Avionics safety switch; 16. First explosive bolt; 17. Second explosive bolt; 18-1. Fuselage; 18-2. Nose section; 18-3. Wing connector; 18-4. Wing mount; 18-5. Wing; 18-6. Servo; 18-7. Tail section; 18-8. Battery; 18-9. Receiver; 18-10. Screw; 18-11. Gasket. Detailed Implementation

[0028] The principles and features of this utility model are described below with reference to the accompanying drawings. The examples given are only for explaining this utility model and are not intended to limit the scope of this utility model.

[0029] To achieve the above objectives, the present invention provides the following specific embodiments:

[0030] Example 1: As Figure 1 and Figure 2 As shown, a folding-wing drone includes a fuselage 18-1, a nose cone 18-2 fixedly mounted at the front end of the fuselage 18-1, a battery 18-8 mounted above the nose cone 18-2, a receiver 18-9 mounted below the nose cone 18-2, a wing mount 18-4 fixedly mounted in the middle of the fuselage 18-1, a tail fin 18-7 fixedly mounted at the rear end of the fuselage 18-1, a wing connector 18-3 mounted above the wing mount 18-4, wings 18-5 symmetrically mounted on both sides of the wing connector 18-3, and servos 18-6 fixedly mounted on the wings 18-5.

[0031] The wing mount 18-4 and the wing connector 18-3 are connected by a washer 18-11 and a screw 18-10. The screw 18-10 passes through the through hole on the wing mount 18-4 and the wing connector 18-3 to achieve connection. The wing connector 18-3 rotates around the screw 18-10, which drives the wing 18-5 to rotate as a whole. The washer 18-11 is placed on the contact surface of the wing connector 18-3 and the wing mount 18-4 to reduce friction during rotation. When the wing 18-5 rotates to be collinear with the fuselage 18-1, the UAV is installed into the rocket payload bay.

[0032] The wing connector 18-3 is fixed to the fuselage 18-1 by a spring. When the wing 18-5 is rotated and folded, the spring is in a stretched state. When the wing 18-5 is fully extended, the spring returns to its initial state. When the wing 18-5 is fully extended, the wing connector 18-3 is fixed by being connected to the wing mount 18-4 by a limiting slot.

[0033] The outer layer of the wing 18-5 is made of a composite material of carbon fiber and epoxy resin, and the interior of the wing 18-5 is filled with foam. The wing adopts the NACA3412 airfoil, with a wing chord length of 5cm and a wingspan of 44-45cm. It adopts an upward dihedral mounting method and has good gliding performance.

[0034] The battery 18-8 and receiver 18-9 are detachable, and the tail fin 18-7 is an integrated structure of vertical and horizontal tail fins; it is made using 3D printing technology.

[0035] Example 2: As Figures 3 to 6 As shown, a micro launch vehicle for launching the folding-wing UAV provided in Embodiment 1 includes a power assembly 1, in which a solid rocket engine 12 is fixedly installed. A parachute compartment 2 is fixedly installed above the power assembly 1, and a deceleration parachute is provided inside the parachute compartment 2. An avionics compartment 3 is fixedly installed above the parachute compartment 2. A variable diameter connector 4 is fixedly installed above the avionics compartment 3. The variable diameter connector 4 is fixedly connected to the payload compartment. A fairing is fixedly installed above the payload compartment.

[0036] The payload compartment includes an upper payload compartment 9 and a lower payload compartment 5. The lower payload compartment 5 is fixedly installed above the lower payload compartment connector 6. The lower payload compartment connector 6 is fixedly connected to the upper payload compartment connector 8 on both sides via hinges 7. The upper payload compartment 9 is fixedly connected to the upper payload compartment connector 8. The avionics compartment 3 is fixedly equipped with an avionics bracket 14 for fixing and installing the avionics system board. The variable diameter connector 4 is equipped with an avionics switch 15.

[0037] The solid rocket motor 12 is configured as four units, and the deceleration parachute is configured as two deceleration parachutes with a diameter of 50cm. The two deceleration parachutes are symmetrically fixed inside the parachute compartment 2. The avionics switch 15 includes an avionics master switch 15-1 and an avionics safety switch 15-2.

[0038] Two second explosive bolts 17 are symmetrically fixedly installed on the upper part of the parachute compartment 2. The second explosive bolts 17 are connected to the parachute compartment cover 13 by screws. Under the control of the avionics system board, the two explosive bolts are ignited simultaneously. The upper end of the parachute compartment cover 13 flips outward under the action of the explosive bolts. The guide rope between the parachute compartment cover 13 and the deceleration parachute pulls the deceleration parachute out of the parachute compartment 2. The deceleration parachute is then inflated to achieve a deceleration effect, realizing rocket recovery. The fairing includes a first half fairing 10 and a second half fairing 11. The surface area of ​​the first half fairing 10 is larger than that of the second half fairing 11. The first half fairing 10 and the second half fairing 11 are fixedly connected by screws and first explosive bolts 16. When the folding-wing UAV is deployed, the first explosive bolts 16 are ignited. The first half fairing 10 and the second half fairing 11 separate under the action of the first explosive bolts 16, causing the payload compartment to rotate and open around the hinge 7, and the folding-wing UAV separates from the rocket.

[0039] The upper part 9 of the payload compartment is composed of two semi-cylinders. The first semi-fairing 10 and the second semi-fairing 11 are respectively fixedly installed on the two semi-cylinders constituting the upper part 9 of the payload compartment. The payload compartment is made of a composite material made of carbon fiber and epoxy resin.

[0040] The avionics support 14 is also equipped with a lithium battery that supplies power to the avionics system board. It controls the rocket to reach the designated airspace through an external line, and completes the launch of the aircraft and the recovery of the rocket. The payload compartment has a diameter of 8cm and a total length of 43cm. The avionics compartment and the powertrain have a diameter of 6cm.

[0041] like Figure 7 and Figure 8 As shown, in implementing this utility model, after installing the folding-wing UAV onto the rocket, the rocket is placed on the launch pad, and the launch pad jumper wire is connected to the rocket signal line. The avionics system master switch and safety switch are turned on, and after verifying that the system is working correctly, the rocket is ignited and launched via an external ignition device or the avionics system. After the rocket reaches the designated airspace, the UAV is released, and the UAV's wings unfold to cruise and perform the designated mission. The specific process is as follows:

[0042] In this implementation, the wing 18-5 is first rotated to be parallel to the fuselage 18-1. Then, the UAV is installed into the rocket payload bay 5, and the first half-fairing 10, the second half-fairing 11, and the first explosive bolt 16 are secured using M2.5 screws. The rocket is placed on the launch pad, and the rocket signal line is connected to the jumper wire on the launch pad. The jumper wire provides the avionics system with the rocket launch signal and the initial altitude detection signal. With the jumper wire pre-connected, the avionics main switch 15-1 is turned on. The avionics system will record the rocket reference pressure after seven flashes and beeps of the indicator lights and buzzer on the panel. This reference pressure is used for subsequent altitude calculations by the pressure sensor. Afterward, the avionics system is in standby mode. The avionics safety switch 15-2 is turned on to check if the avionics system is working properly. After confirming that the avionics system is working properly, the solid rocket motor is ignited via an external igniter. At launch, the jumper wire disconnects, and the avionics system begins full operation, constantly monitoring the rocket's flight time, altitude, and other information. These measurements are stored in a FLASH memory chip, and various operations are executed according to preset parameters. When the avionics system detects that the rocket's flight parameters meet the preset requirements, it executes the FUSE1 ignition command. Under this command, the second explosive bolt 17 ignites, causing the upper part of the parachute compartment door 13 to flip outwards. The guide rope between the parachute compartment door and the deceleration chute pulls the chute out of the compartment, inflating it and controlling the rocket's deceleration. By controlling the parachute deployment time, the rocket can be controlled at a designated altitude. The avionics system continues to monitor the rocket's flight parameters. When separation requirements are met, it executes the FUSE2 ignition command, igniting the first explosive bolt 16. The first half-fairing 10 and the second half-fairing 11 separate under the action of the explosive bolt, causing the semi-cylindrical payload bay to open around the hinge, separating the UAV from the rocket. The rocket then descends and is recovered under the influence of the deceleration chute, while the UAV performs its designated mission. The avionics system stores rocket flight data while simultaneously shutting down the ignition port to ensure rocket safety and data integrity. This system can control the drone's deployment altitude error to within +0.1m.

[0043] Furthermore, the avionics system is developed based on the STM32F103C8T6 microcontroller. The system board is equipped with an external crystal oscillator, passive buzzer, indicator lights, a barometric pressure sensor, a FLASH memory chip, a USART communication output interface, a PWM output interface, and ignition output interfaces (FUSE1, FUSE2, FUSE3). USART communication allows for communication with external devices, enabling real-time modification of preset rocket flight parameters and real-time reading of rocket flight data, thus achieving remote control of the launch system.

[0044] The above description is only a preferred embodiment of the present utility model and is not intended to limit the present utility model. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present utility model should be included within the protection scope of the present utility model.

Claims

1. A folding wing drone comprising a fuselage (18-1), characterized in that, The front end of the fuselage (18-1) is fixedly installed with a plane head (18-2), the upper part of the plane head (18-2) is installed with a battery (18-8), and the lower part of the plane head (18-2) is installed with a receiver (18-9); The middle part of the fuselage (18-1) is fixedly installed with a wing seat (18-4), the rear end of the fuselage (18-1) is fixedly installed with a tail wing (18-7), the upper part of the wing seat (18-4) is installed with a wing connecting piece (18-3), the two sides of the wing connecting piece (18-3) are symmetrically installed with wings (18-5), and the wings (18-5) are fixedly installed with rudders (18-6).

2. The folding-wing drone of claim 1, wherein, The wing seat (18-4) and the wing connecting piece (18-3) are connected through a gasket and a screw; the screw passes through through holes on the wing seat (18-4) and the wing connecting piece (18-3) to realize connection, the wing connecting piece (18-3) rotates around the screw while driving the wings (18-5) to rotate as a whole, and the gasket is arranged on the contact surface of the wing connecting piece (18-3) and the wing seat (18-4) to reduce friction in the rotating process.

3. The folding-wing drone of claim 1 or 2, wherein, The wing connecting piece (18-3) and the fuselage (18-1) are fixedly connected through a spring, the spring is in a stretched state when the wings (18-5) are rotated and folded, the spring returns to the initial state when the wings (18-5) are completely unfolded, and the wing connecting piece (18-3) is connected to the wing seat (18-4) through a limiting clamping groove arranged on the wing seat (18-4) to realize limiting and fixing.

4. A folding-wing drone as claimed in claim 3, wherein, The outer layer of the wings (18-5) is a composite material made of carbon fiber and epoxy resin, and the inside of the wings (18-5) is filled with foam; the wings (18-5) adopt a NACA3412 airfoil, the wing chord length is 5 cm, the wingspan is 44-45 cm, an installation mode with an upper negative angle is adopted, and good gliding performance is achieved; the battery (18-8) and the receiver (18-9) adopt a detachable structure, and the tail wing (18-7) adopts an integrated structure of a vertical tail wing and a horizontal tail wing.

5. A micro launch vehicle for launching the folding wing drone of claims 1-4, characterized in that, The power assembly (1) is internally fixedly installed with a solid rocket engine (12), the upper part of the power assembly (1) is fixedly installed with a parachute cabin (2), the inside of the parachute cabin (2) is provided with a deceleration parachute, the upper part of the parachute cabin (2) is fixedly installed with an avionics cabin (3), the upper part of the avionics cabin (3) is fixedly installed with a variable-diameter connecting piece (4), the upper part of the variable-diameter connecting piece (4) is fixedly connected to a load cabin, and the upper part of the load cabin is fixedly installed with a fairing.

6. The microorbiter of claim 5, wherein, The load cabin comprises a load cabin upper part (9) and a load cabin lower part (5), the upper part of the load cabin lower part (5) is fixedly installed with a load cabin lower part connecting piece (6), the two sides of the load cabin lower part connecting piece (6) are fixedly connected to a load cabin upper part connecting piece (8) through hinges (7), and the load cabin upper part (9) is fixedly connected to the load cabin upper part connecting piece (8); the avionics cabin (3) is internally fixedly provided with an avionics support (14) for fixedly installing an avionics system board, and the variable-diameter connecting piece (4) is provided with an avionics switch (15).

7. The microorbiter of claim 5, wherein, The solid rocket engine (12) is arranged as four, the deceleration parachute is arranged as two deceleration parachutes with a diameter of 50 cm, and the two deceleration parachutes are symmetrically fixed and installed in the interior of the parachute cabin (2).

8. The microorbiter of claim 5, wherein, The upper portion of the parachute cabin (2) is symmetrically fixed and installed with two second explosion bolts (17), the second explosion bolts (17) are connected with the parachute cover (13) through screws, the two explosion bolts are simultaneously ignited under the control of the avionics system board, the upper end of the parachute cover (13) is turned outwards under the action of the explosion bolts, the guide rope between the parachute cover (13) and the deceleration parachute pulls the deceleration parachute out of the parachute cabin (2), the deceleration parachute is inflated to achieve the deceleration effect, and the rocket is recycled. The fairing comprises a first half fairing (10) and a second half fairing (11), the surface area of the first half fairing (10) is greater than that of the second half fairing (11), and the first half fairing (10) and the second half fairing (11) are fixedly connected through screws and a first explosion bolt (16); when the folded wing unmanned plane is launched, the first explosion bolt (16) is ignited, the first half fairing (10) and the second half fairing (11) are separated under the action of the first explosion bolt (16), the load cabin is rotated and opened around the hinge (7), and the folded wing unmanned plane is separated from the rocket.

9. The microorbiter of claim 8, wherein, The upper portion (9) of the load cabin is composed of two half cylinders, the first half fairing (10) and the second half fairing (11) are respectively fixedly installed on the two half cylinders constituting the upper portion (9) of the load cabin, and the load cabin is made of a composite material processed from carbon fibers and epoxy resin.

10. The microorbiter of claim 6, wherein, The avionics support (14) is further provided with lithium batteries for supplying power to the avionics system board; and the avionics switch (15) comprises an avionics master switch and an avionics safety switch.

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