An aircraft

Abstract

The application discloses a kind of aircraft, the aircraft includes fairing, aircraft body and vector thrust mechanism;The inside of the aircraft body is provided with oil tank, the oil tank is used to provide fuel for the vector thrust mechanism;One end of the aircraft body is connected with the fairing, the other end of the aircraft body is connected with the vector thrust mechanism;The vector thrust mechanism includes turbojet and vector nozzle, the vector nozzle includes the drive mechanism erected in the head of the turbojet, and vector jet is erected in the tail of the turbojet;The vector thrust mechanism is used to adjust the current attitude position of the aircraft;The outer side wall of the vector thrust mechanism is provided with landing mechanism, for stabilizing the overall attitude when the aircraft lands.The aircraft provided by the application can greatly reduce the research and development risk of the aircraft, improve the iteration speed and reduce the iteration cost.

Description

Technical Field

[0001] This invention belongs to the field of aerospace technology, and specifically relates to an aircraft. Background Technology

[0002] With the continuous development of the aerospace field both domestically and internationally, reducing space launch costs is one of the major challenges facing the entire aerospace industry. This is especially true for low-altitude, high-speed flight test launches, where the recovery and reuse of launch vehicles and their payloads is a crucial cost-reduction measure. From a configuration and technical perspective, reusable launch vehicles can be broadly categorized into horizontally recoverable reusable launch vehicles and vertically recoverable reusable launch vehicles. Achieving reuse through vehicle recovery and refueling can save over 90% of costs.

[0003] However, for aircraft, due to their need to pass through the atmosphere and be recovered and landed, high requirements are placed on the structural design of aircraft. Furthermore, due to the extremely high cost of aircraft, the cost of iterative trial and error is huge. These factors together result in extremely high cost, extremely long research cycle and great research risk for aircraft. In this paper, an aircraft is proposed to solve the above problems. Summary of the Invention

[0004] This invention provides an aircraft to address the problems of high cost, long research cycle, and high research risk in existing aircraft technologies.

[0005] To solve the above-mentioned technical problems, the present invention provides an aircraft, including a fairing, an aircraft body, and a vector thrust mechanism;

[0006] The aircraft body is equipped with a fuel tank, which is used to provide fuel for the vector thrust mechanism.

[0007] One end of the aircraft body is connected to the fairing, and the other end of the aircraft body is connected to the vector thrust mechanism; the vector thrust mechanism includes a turbojet engine and a vector nozzle, the vector nozzle includes a drive mechanism mounted on the head of the turbojet engine and a vector nozzle mounted on the tail of the turbojet engine; the vector thrust mechanism is used to adjust the current attitude position of the aircraft.

[0008] The outer wall of the vector thrust mechanism is provided with a landing mechanism, which includes several support structures arranged around the outer wall of the vector thrust mechanism to stabilize the overall attitude of the aircraft when it lands.

[0009] As a further improvement of the present invention, the turbojet engine is connected to the oil outlet of the oil tank;

[0010] The output shaft of the drive mechanism is rotatably connected to a rocker arm. A connecting rod is provided between the drive mechanism and the vector nozzle. One end of the connecting rod is hinged to the outer wall of the vector nozzle, and the other end of the connecting rod is connected to the rocker arm bearing.

[0011] As a further improvement of the present invention, the oil tank includes an oil tank wall, an oil tank top cover and an oil tank bottom cover, the oil tank top cover is provided with an oil inlet and the oil outlet is provided at the bottom of the oil tank bottom cover;

[0012] The fuel tank is equipped with a connecting pipe and a partition. The two ends of the connecting pipe are connected to the top cover and the bottom cover of the fuel tank, respectively. One end of the partition is connected to the inner wall of the bottom cover of the fuel tank, and the other end of the partition forms an angle with the plane of the bottom cover of the fuel tank, and there is a gap between the partition and the outer wall of the connecting pipe.

[0013] As a further improvement of the present invention, the fuel tank is located at the center of gravity of the aircraft, and the fuel outlet of the fuel tank is connected to the turbojet engine to provide fuel for the turbojet engine.

[0014] As a further improvement of the present invention, each of the support structures includes a support rod and a tie rod, one end of the support rod and the tie rod being hinged to the outer wall of the vector thrust mechanism through an ear plate, and the other end of the support rod and the tie rod being connected through a support foot;

[0015] The support foot includes an inclined end, a grounding end, and a hinged end. The inclined end is inclinedly disposed on the top of the grounding end and fixedly connected to the other end of the support rod. The hinged end is disposed on the side wall of the grounding end and hinged to the other end of the pull rod.

[0016] As a further improvement of the present invention, the radius of the inscribed circle of the polygon formed by the contact points between the several supporting structures and the ground is greater than or equal to 1 / 3 of the height of the center of mass of the aircraft.

[0017] As a further improvement of the present invention, a roll angle control mechanism is provided on the outer side wall of the aircraft body, and the roll angle control mechanism is used to adjust the current roll angle attitude of the aircraft.

[0018] The roll angle control mechanism includes a ring-shaped mount and two adjustment components, which are symmetrically arranged around the side wall of the mount with the mount as the center point.

[0019] Each of the adjustment components includes a drive motor and a support arm, the support arm being disposed on the outer side wall of the mount and extending outward, and an electric speed controller connected to the drive motor being disposed on the support arm;

[0020] The drive motor is located at the end of the support arm, and its output shaft is connected to a propeller.

[0021] As a further improvement of the present invention, the roll angle control mechanism is disposed on the outer side wall of the aircraft body and near the center of gravity of the aircraft.

[0022] As a further improvement of the present invention, a control mechanism is also provided within the main body of the aircraft. The control mechanism includes attitude and position information, a main control unit, and a numerical control transmission unit.

[0023] The attitude position sensor is used to acquire the current attitude position information of the aircraft; the numerical control transmission unit is used to receive the command information sent by the ground station and transmit the command information to the main control unit; the main control unit is used to generate control information based on the current attitude position information of the aircraft and the command information, and send the control information to the actuator.

[0024] The actuator includes the vector thrust mechanism and the roll angle control mechanism, and the actuator is used to receive and execute the control information.

[0025] As a further improvement of the present invention, the attitude position sensor, the vector thrust mechanism, the roll angle control mechanism, and the numerical control transmission unit are connected to the main control unit via wireless and / or wired means, and the numerical control transmission unit and the ground station are wirelessly connected.

[0026] The present invention provides an aircraft, including a fairing and an aircraft body. The aircraft body is equipped with a control mechanism and a fuel tank. The side wall of the aircraft body is equipped with a roll angle control mechanism. The bottom of the aircraft body is equipped with a vector thrust mechanism and a landing mechanism. By setting a fuel tank inside the aircraft body, it is ensured that the air bubbles in the fuel tank can be quickly discharged during the fuel export process, thus ensuring the continuity of fuel supply and stably supplying fuel to the vector thrust mechanism.

[0027] By installing a vector thrust mechanism at the bottom of the aircraft body, the linkage drives the vector nozzle to produce a pitch swing of ±23° relative to the turbojet engine, thereby adjusting the thrust direction generated by the vector thrust mechanism. Two servos are set up to precisely control the pitch angle, yaw angle, and thrust direction of the vector nozzle. A roll angle control mechanism is set up to achieve safe and stable landing of the aircraft. The roll angle attitude of the aircraft can be controlled even at low speeds, which solves the problem of lack of roll angle control freedom at low speeds in existing technologies. At the same time, the side baffles further reduce wind resistance and enhance safety. By setting a landing mechanism on the side wall of the vector thrust mechanism, the overall attitude of the aircraft is guaranteed during landing.

[0028] Meanwhile, the main control unit, numerical control transmission unit, and attitude position sensor are integrated within the main body of the aircraft. The numerical control transmission unit transmits the current attitude position information of the aircraft monitored by the attitude position sensor to the main control unit, and also transmits the command information received from the ground station to the main control unit. The main control unit combines the current attitude position information of the aircraft and the command information sent by the ground station to generate control information, and then transmits it to other actuators set up inside the aircraft. This can effectively simulate the working conditions of the aircraft under medium and low speed conditions. After the algorithm is tested and matured, it is deployed on the aircraft. It is low-cost and suitable for mass production, which greatly reduces the research and development risk of the aircraft, improves the iteration speed, and reduces the iteration cost. Attached Figure Description

[0029] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0030] Figure 1 This is a schematic diagram of the structure of an aircraft provided in an embodiment of the present invention;

[0031] Figure 2 This is a functional block diagram of an aircraft provided in an embodiment of the present invention;

[0032] Figure 3 This is a schematic diagram of the structure of a fuel tank for an aircraft provided in an embodiment of the present invention;

[0033] Figure 4 This is a structural schematic diagram of an aircraft that directly uses the side wall of the aircraft body as the fuel tank wall, according to an embodiment of the present invention.

[0034] Figure 5 This is a perspective view of an aircraft that uses the side wall of the aircraft body as the fuel tank wall, as provided in an embodiment of the present invention.

[0035] Figure 6 This is a schematic diagram of the structure of a fuel tank top cover for an aircraft provided in an embodiment of the present invention;

[0036] Figure 7 This is a schematic diagram of the structure of a fuel tank bottom cover for an aircraft provided in an embodiment of the present invention;

[0037] Figure 8 This is a schematic diagram of the internal structure of a vector thrust mechanism for an aircraft provided in an embodiment of the present invention;

[0038] Figure 9This is a schematic diagram of the structure of an aircraft after the vector nozzle is deflected, according to an embodiment of the present invention;

[0039] Figure 10 This is a schematic diagram of the structure of a vector thrust mechanism for an aircraft provided in an embodiment of the present invention;

[0040] Figure 11 This is a schematic diagram of the structure of a servo motor, servo motor frame, and servo motor board of an aircraft after installation, according to an embodiment of the present invention.

[0041] Figure 12 This is a schematic diagram of the structure of a vector nozzle of an aircraft provided in an embodiment of the present invention;

[0042] Figure 13 This is a schematic diagram of the structure of a servo motor for an aircraft provided in an embodiment of the present invention;

[0043] Figure 14 This is a schematic diagram of the structure of a servo circuit board for an aircraft provided in an embodiment of the present invention;

[0044] Figure 15 This is a schematic diagram of the structure of a servo mount for an aircraft provided in an embodiment of the present invention;

[0045] Figure 16 This is a schematic diagram of the horizontal angle formed by the two connecting arms of an aircraft to the center of the vector nozzle, according to an embodiment of the present invention.

[0046] Figure 17 This is a schematic diagram of the structure of a landing mechanism for an aircraft provided in an embodiment of the present invention;

[0047] Figure 18 This is a schematic diagram of the structure of a support foot for an aircraft provided in an embodiment of the present invention;

[0048] Figure 19 This is a schematic diagram of the roll angle control mechanism for an aircraft provided in an embodiment of the present invention.

[0049] Explanation of reference numerals in the attached figures

[0050] 100 - Fairing;

[0051] 200 - Main body of the aircraft;

[0052] 300-Fuel tank; 3010-Fuel tank wall; 3101-Connecting pipe; 3020-Fuel tank top cover; 3201-Oil inlet; 3202-First connecting part; 3203-Oil guide pipe; 3203a-First end; 3203b-Second end; 3204-Quick connector; 3205-Support plate; 3205a-First side wall; 3205b-Second side wall; 3206-Second fastening bolt; 3207-Second fixing groove; 3030-Fuel tank bottom cover; 3301-Oil outlet; 3302-Second connecting part; 3303-Partition plate; 3304-Baffle plate; 3304a-First side wall; 3304b-Second side wall; 3305-First fastening bolt; 3306-First fixing groove; 3040-Sealing groove; 3401-Sealing ring;

[0053] 400-Vector thrust mechanism; 4010-Drive mechanism; 4101-Servo; 4101a-Extension plate; 4102-Output shaft; 4103-Rocker arm; 4103a-Adjustment hole; 4104-Servo frame; 4104a-Extension; 4105-Servo board; 4105a-First locking post; 4105b-Second locking post; 4105c-Fixing post; 4106-Reinforcing plate; 4107-Opening; 4108-Guide plate; 4108a-First groove; 4108b-Second groove; 4020-Turbojet engine; 4201-Connecting rod; 4202-Second coupling; 4203-Universal Joint; 4204-Fisheye joint bearing; 4205-Turbojet mounting bracket; 4206-Turbojet mount; 4207-Fireproof plate; 4207a-Third groove; 4030-Vector nozzle; 4030a-Air inlet; 4030b-Air outlet; 4301-Connecting arm; 4301a-Connecting end; 4301b-Extension end; 4302-First coupling; 4303-Universal ring; 4304-First shaft pin; 4305-External connector; 4305a-Fixed end; 4305b-Free end; 4306-Second shaft pin; 4306-Reinforcing rib; 4040-Support body; 4401-Air inlet;

[0054] 500 - Landing mechanism; 5010 - Support structure; 5011 - Support rod; 5012 - Tie rod; 5013 - Ear plate; 5020 - Support foot; 5021 - Inclined end; 5022 - Grounding end; 5023 - Hinge end;

[0055] 600 - Roll angle control mechanism; 6010 - Side baffle; 6012 - Support frame; 6013 - Propeller; 6014 - Drive motor; 6015 - Support arm; 6016 - Electronic speed controller; 6017 - Mounter. Detailed Implementation

[0056] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0057] To address the problems of high cost, long research cycle, and high research risk in existing aircraft technologies, this invention provides an aircraft. Please refer to... Figures 1-19 The present invention provides an aircraft, please refer to the following for details. Figure 1 This is a schematic diagram of the structure of an aircraft provided in an embodiment of the present invention, including a fairing 100, an aircraft body 200, a vector thrust mechanism 400 and a landing mechanism 500 arranged from top to bottom; wherein, the fairing 100 is generally conical in shape and is used to guide the airflow at the head of the aircraft and reduce the air resistance of the aircraft during flight.

[0058] Furthermore, the material used to manufacture the main body 200 of the aircraft can be determined based on the speed of the aircraft. Commonly available materials include, but are not limited to, composite materials such as plastic, carbon fiber, and glass fiber, stainless steel, or aluminum. The temperature resistance of the selected material should be higher than the aerodynamic heating rate of the outer surface of the aircraft. For example, when the speed of the aircraft is 200 m / s, composite materials such as plastic and carbon fiber can be used. The material of the main body 200 of the aircraft can be selected according to the specific circumstances, which will not be elaborated on here.

[0059] Furthermore, the main body 200 of the aircraft is equipped with an attitude position sensor, a main control unit, and a numerical control transmission unit. Specifically, the attitude position sensor is used to acquire the current attitude position information of the aircraft, the numerical control transmission unit is used to receive the command information sent by the ground station and transmit the received command information to the main control unit, and the main control unit is used to generate control information based on the current attitude position information of the aircraft and the command information, and send the above control information to the actuators such as the vector thrust mechanism.

[0060] Specifically, the aforementioned attitude and position sensors can be strapdown inertial navigation systems, including but not limited to AHRS (Attitude and Heading Reference System) + GNSS (Global Navigation Satellite System) integrated navigation, etc., and can also use MEMS (Micro-electromechanical Systems) and fiber optic gyroscopes; among which, the models of MEMS can include but are not limited to MPU9250, IST8310, RM3100, and the models of BMI088 fiber optic gyroscopes can include but are not limited to FS-70FS-210; the current attitude and position information of the aircraft can also be obtained through platform-type inertial navigation systems such as mechanical gyroscopes and electrostatic gyroscopes.

[0061] The aforementioned numerical control transmission unit is used to realize long-distance transmission of commands between the ground station and the main control unit, and at the same time to transmit the current status information of the aircraft back to the ground station. The numerical control transmission unit can use embedded data transmission solutions, such as P900, P800, and NRF24L01; or it can use Wi-Fi data transmission solutions, such as ESP32 and ESP8266. Since the data transmission between devices via wireless communication is a very mature application in the existing technology, the specific data transmission solutions available in wireless communication will not be elaborated on here.

[0062] The main control unit is used to generate control information based on the current attitude and position information of the aircraft and the command information received from the ground station, and to send the above control information to the actuators inside the aircraft. Specifically, it can be implemented by a microcontroller, a microcontroller and a flight controller, and needs to have interfaces such as CAN bus, UART interface, and PWM. Among them, the microcontroller models that can be selected include, but are not limited to, Jetson Nano, Jetson TX2, Jetson Xavier, Raspberry Pi 4B, etc. The microcontroller models that can be selected include, but are not limited to, the entire Arduino series, the entire STM32 series, etc. The flight controller can be PX4, APM, etc. There are no restrictions on the specific configuration of the main control unit.

[0063] Please refer to Figure 2This is a functional block diagram of an aircraft provided in an embodiment of the present invention. Since the main control unit, the numerical control transmission unit, and the attitude position sensor are all located inside the aircraft body 200, the attitude position sensor, the main control unit, and the numerical control transmission unit can all be connected through signal lines. Thus, the numerical control transmission unit transmits the current attitude position information of the aircraft monitored by the attitude position sensor to the main control unit, and also transmits the command information received from the ground station to the main control unit. The main control unit combines the current attitude position information of the aircraft and the command information sent by the ground station to generate control information, and then transmits it to other actuators set inside the aircraft.

[0064] It should be noted that, except for the wireless communication method between the numerical control transmission unit and the ground station, the main control unit, numerical control transmission unit, attitude and position sensors installed in the main body 200 of the aircraft, as well as other actuators installed on the aircraft, can be connected by cables or wirelessly. Those skilled in the art can adjust according to the actual situation, and no specific restrictions are made here.

[0065] Furthermore, the ground station is used to transmit control commands of the aircraft to the main control unit via the numerical control transmission unit, such as changing the flight position or flight attitude. After receiving this command information from the ground station, the main control unit performs specific analysis based on the current attitude and position information of the aircraft, and controls other actuators to achieve the control commands from the ground station. It can also be understood that the ground station sends a command information, which does not include how to adjust the various components inside the aircraft. Instead, the main control unit analyzes the received command information in conjunction with the current attitude information of the aircraft and issues control information to the actuators set on the aircraft to realize the command information sent by the ground station.

[0066] Furthermore, the ground station can also display the aircraft's status information, such as battery level, fuel level, flight altitude, and flight speed. Operators can set the display information of the ground station according to actual needs. In this embodiment of the invention, the ground station can be a model aircraft remote controller, such as ET16 or TX18S; it can be a Windows or Linux system laptop or desktop computer, or it can be existing autopilot software. Any system that can achieve the above-mentioned technical effects of the ground station is feasible.

[0067] Furthermore, a power module is also provided inside the main body 200 of the aircraft to supply power to all electrical components on the aircraft. This power module should be isolated from the aforementioned numerical control transmission unit, main control unit, attitude position sensor and other control mechanisms to ensure the integrity of signal transmission between the aforementioned control mechanisms.

[0068] Furthermore, a fuel tank 300 is also installed inside the main body 200 of the aircraft for supplying fuel to the vector thrust mechanism 400. Please refer to [reference needed]. Figure 3 This is a schematic diagram of the structure of a fuel tank 300 for an aircraft according to an embodiment of the present invention. The fuel tank 300 is preferably located at the center of gravity of the aircraft and is connected to the turbojet engine 4020 of the vector thrust mechanism 400 to provide fuel for the turbojet engine 4020. The location at the center of gravity can reduce the change in the center of gravity caused by the fuel inside the fuel tank 300. The aircraft fuel tank 300 includes a fuel tank wall 3010, a fuel tank top cover 3020, and a fuel tank bottom cover 3030. The fuel tank top cover 3020 is provided with a fuel inlet 3201, and the fuel tank bottom cover 3030 is provided with a fuel outlet 3301. It should be noted that the internal cavity formed by the fuel tank wall 3010, the fuel tank top cover 3020, and the fuel tank bottom cover 3030 is used to place fuel. Its shape is adapted to the shape of the aircraft body 200 and can be placed inside the aircraft body 200 for use. Of course, the inner side wall of the aircraft body 200 can also be used directly as the fuel tank wall 3010.

[0069] Please refer to Figure 4 This is a structural diagram of an aircraft that uses the side wall of the aircraft body 200 as the fuel tank wall 3010, as provided in an embodiment of the present invention. This can minimize the weight of the fuel tank 300 and improve the space utilization rate inside the aircraft body 200. In this case, it is necessary to set a fuel tank top cover 3020 and a fuel tank bottom cover 3030 that are adapted to the inner side wall of the aircraft body 200. In practical applications, it is feasible to choose either a fuel tank wall 3010 that is adapted to the inner wall of the aircraft body 200 or to use the inner side wall of the aircraft body 200 as the fuel tank wall 3010 directly. No further restrictions are imposed here.

[0070] Further, please refer to Figure 5This is a perspective view of an aircraft that uses the side wall of the aircraft body 200 as the fuel tank wall 3010, as provided in an embodiment of the present invention. Both the fuel tank top cover 3020 and the fuel tank bottom cover 3030 are frustum structures, with the maximum diameter of the frustum structure matching the inner diameter of the fuel tank wall 3010. The bottom surfaces of the two frustum structures, being the maximum diameter, are fixed to both ends of the fuel tank wall 3010, while the top surfaces of the two frustum structures face away from the fuel tank wall 3010. A connecting pipe 3101 is also provided inside the fuel tank wall 3010 for passing wires through the fuel tank 300. Specifically, both ends of the connecting pipe 3101 are connected to the fuel tank top cover 3020 and the fuel tank bottom cover 3030, respectively, and sequentially pass through the center of the fuel tank top cover 20 and the fuel tank bottom cover 30. The fuel tank top cover 3020 and fuel tank bottom cover 3030 are fixed at both ends of the fuel tank wall 3010. It should be noted that the connecting pipe 3101 is a hollow tubular structure. The connecting pipe 3101 is set to provide a passage for the wires installed inside the aircraft, thereby facilitating the installation of control cables. The connection part of the connecting pipe 3101 with the fuel tank top cover 3020 is referred to as the first connection part 3202, and the connection part with the fuel tank bottom cover 3030 is referred to as the second connection part 3302. The first connection part 3202 and the second connection part 3302 are sealed together. Specifically, they can be bonded with industrial adhesive to prevent fuel in the fuel tank 300 from leaking out from the first connection part 3202 and / or the second connection part 3302, which would affect the fuel storage effect.

[0071] Preferably, the oil inlet 3201 is located near the first connecting part 3202, and the oil outlet 3301 is located near the second connecting part 3302.

[0072] Further, please refer to Figure 7 This is a schematic diagram of the structure of a fuel tank bottom cover 3030 of an aircraft according to an embodiment of the present invention. A partition 3303 is also provided inside the fuel tank 300. The partition 3303 is used to divide the internal cavity of the fuel tank 300 into a fuel storage area and a bubble-free area. The partition 3303 is preferably an annular partition. One end of the annular partition is fixedly connected to the inner wall of the fuel tank bottom cover 3030, and the other end of the annular partition is arranged along the direction of the fuel inlet 3201 and tapers inward, i.e., the other end of the partition 3303 forms a plane with the plane of the fuel tank bottom cover 3030. After tilting, a gap is also left between it and the outer wall of the connecting pipe 3101; when the direction of the top cover 3020 of the oil tank is upward and the direction of the bottom cover 3030 of the oil tank is downward, the above description can be understood as follows: one end of the partition 3303 is fixedly connected to the inner wall of the bottom cover 3030 of the oil tank, and the other end of the partition 3303 is tilted upward. While its end gradually shrinks, a gap is left between it and the outer wall of the connecting pipe 3101. The upward tilted partition 3303 here can ensure that the bubbles at the corners of the bubble-free area can also be discharged smoothly.

[0073] It should be noted that when the above-mentioned partition 3303 is tilted upward, the angle between its tilt angle and the horizontal direction is preferably set to 3-15°; the above-mentioned partition 3303 is a double-layer structure, which can be formed by directly bonding two partitions 3303 with an adhesive, or by printing a double-layer partition structure using 3D printing technology. These methods are all feasible; the double-layer partition structure here can continuously and stably supply fuel to the vector thrust mechanism 400 without bubbles even when the aircraft is swaying at a large angle.

[0074] Specifically, the aforementioned interval 3303 can be an irregular interval or a relatively regular annular interval. However, in principle, the width of the interval 3303 should be kept within the range of 2-8mm. This setting ensures that the fuel in the oil storage area can smoothly enter the bubble-free zone through the interval 3303. At the same time, since there is also an oil outlet 3301 on the bottom 3030 of the oil tank, when the vector thrust mechanism 400 needs to use fuel, the fuel in the oil tank 300 will be discharged through the oil outlet 3301 of the bubble-free zone for use by the vector thrust mechanism 400.

[0075] Further, please continue to refer to Figure 7 This is a schematic diagram of the structure of a fuel tank bottom cover 3030 for an aircraft fuel tank according to an embodiment of the present invention. To enable rapid removal of air bubbles after fuel enters the bubble-free zone, this embodiment of the present invention also provides several baffles 3304 along the circumferential direction on the inner sidewall of the fuel tank bottom cover 3030. To adapt to the frustum structure of the fuel tank bottom cover 3030, this embodiment of the present invention sets the baffles 3304 into a triangular structure. Since the fuel tank bottom cover 3030 has a frustum structure, this triangular structure is designed to fit the inner sidewall of the fuel tank bottom cover 3030. The structure can actually be set as an obtuse triangle, with the two adjacent sides of the triangle denoted as the first sidewall 3304a and the second sidewall 3304b, respectively. The first sidewall 3304a is fixedly connected to the inner sidewall of the vertical end of the fuel tank bottom cover 3030, and the second sidewall 3304b is fixedly connected to the inner sidewall of the inclined end of the fuel tank bottom cover 3030. It should be added that another reason why the baffle 3304 is preferably a triangular structure is that the triangular structure can strengthen the support, thereby improving the overall strength of the fuel tank 300.

[0076] In this embodiment of the invention, the first sidewall 3304a of the baffle 3304 is fixedly connected to the inner sidewall of the vertical end of the fuel tank bottom cover 3030 by a first fastening bolt 3305. Specifically, the first sidewall 3304a of the baffle 3304 is provided with a first fixing groove 3306 through which the first fastening bolt 3305 passes. The first fixing groove 3306 is provided with an internal thread (not shown in the figure) that is compatible with the first fastening bolt 3305. When fixing, after the first fastening bolt 3305 is passed through the sidewall of the fuel tank bottom cover 3030 and the first fixing groove 3306 in sequence and the thread is tightened, the baffle 3304 can be fixed to the fuel tank bottom cover 3030. Alternatively, the first fastening bolt 3305 passes through the tank wall 3010, the side wall of the tank bottom cover 3030, and the first fixing groove 3306 in sequence and is tightened to fix the baffle 3304 on the tank bottom cover 3030. At the same time, the tank bottom cover 3030 and the tank wall 3010 are fixedly connected together. This connection method is also feasible. No further restrictions are placed on the specific connection method here. Meanwhile, the second side wall 3304b of the baffle 3304 passes through the partition 3303 and is fixedly connected to the inner side wall of the inclined end of the tank bottom cover 3030. The end of the second side wall 3304b is fixedly connected to the second connecting part 3302.

[0077] When the aircraft is operating, the fuel tank 300 will shake or vibrate, causing the fuel level in the fuel storage area to sway significantly. However, due to the partition 3303, the less than 1cm gap between the partition 3303 and the connecting pipe 101 prevents the swaying of the liquid in the fuel storage area from being transmitted to the bubble-free area. At the same time, several baffles 3304 are installed in the bubble-free area to create obstruction and further reduce the swaying of the fuel in the bubble-free area. Moreover, the fuel outlet 3301 is located at the bottom of the fuel tank bottom cover 3030. Even if bubbles are generated in the fuel in the bubble-free area due to swaying, they can be discharged during the process of being discharged from the fuel outlet 301, further ensuring the continuity of fuel supply.

[0078] Further, please refer to Figure 6 This is a schematic diagram of the structure of a fuel tank top cover 3020 of an aircraft provided in an embodiment of the present invention. The fuel tank top cover 3020 is provided with an oil inlet 3201, and the number of oil inlets 3201 can be several. The fuel tank top cover 3020 is also provided with an L-shaped oil guide pipe 3203 equal to the number of oil inlets. In a specific embodiment provided by the present invention, the fuel tank top cover 3020 is provided with two oil inlets 3201, and correspondingly, two oil guide pipes 3203 are provided.

[0079] Specifically, the oil guide pipe 3203 includes a first end 3203a and a second end 3203b. The first end 3203a of the oil guide pipe 3203 is connected to the oil inlet 3201, and the oil inlet 3201 is provided with a quick connector 3204 for quickly connecting to an external fuel conduit to introduce fuel. The second end 3203b of the oil guide pipe 3203 passes through the top cover 3020 of the fuel tank and is connected to the interior of the fuel tank 300. A filter element (not shown in the figure) is also provided at the bottom of the second end of the oil guide pipe 3203. The filter element can filter out fuel impurities entering the fuel tank 300, thereby preventing impurities in the fuel and air from entering the fuel tank 300 and blocking the turbojet engine 4020 installed in the vector thrust mechanism 400 for long-term use. It should be noted that the L-shaped fuel guide pipe 2203 provided by the present invention can also balance the air pressure inside and outside the fuel tank 300 after the fuel is refueled. Due to the surface tension of the fuel liquid, the occasional splashing caused by shaking during flight will not cause the fuel in the fuel tank 300 to leak after passing through the filter element and the fuel guide pipe 3203.

[0080] Further, please continue to refer to Figure 6 The inner wall of the fuel tank top cover 3020 is provided with a plurality of support plates 3205 along the circumferential direction. The support plates 3205 can be of any shape, but in a specific embodiment of the present invention, in order to adapt to the frustum structure of the fuel tank top cover 3020, the support plates 3205 are provided with a triangular structure, specifically an obtuse triangle. The two adjacent sides of the triangle are respectively designated as the first side wall 3205a and the second side wall 3205b. The first side wall 3205a is fixedly connected to the inner wall of the vertical end of the fuel tank top cover 3020, and the second side wall 3205b is fixedly connected to the inner wall of the inclined end of the fuel tank top cover 3020. It should be noted that the reason why the triangular structure is preferred is that the triangular structure can strengthen the support, thereby improving the overall strength of the fuel tank 300.

[0081] Specifically, the first sidewall 3205a of the support plate 3205 is fixedly connected to the inner sidewall of the vertical end of the fuel tank top cover 3020 by a second fastening bolt 3206. The first sidewall 3205a of the support plate 3205 is provided with a second fixing groove 3207 for the second fastening bolt 3206 to pass through. The second fixing groove 3207 is provided with an internal thread (not shown in the figure) that matches the second fastening bolt 3206. When fixing, after passing the second fastening bolt 3206 through the sidewall of the fuel tank top cover 3020 and the second fixing groove 3207 and tightening the thread, the support plate 3205 can be fixed to the fuel tank top cover 3020; or, the second fastening bolt... After the bolt 3206 passes through the tank wall 3010, the side wall of the tank top cover 3020, and the second fixing groove 3207 in sequence and is tightened, the support plate 3205 is fixed on the tank top cover 3020. At the same time, the tank top cover 3020 and the tank wall 3010 are fixedly connected together. This connection method is also feasible. No further restrictions are placed on the specific connection method here. Meanwhile, the second side wall 3205b of the support plate 3205 is set along the direction of the oil inlet 3201 and is fixedly connected to the inner side wall of the inclined end of the tank top cover 3020. The end of the second side wall 3205b of the support plate 3205 is fixedly connected to the first connecting part 3202.

[0082] Furthermore, the outer walls of the top cover 3020 and the bottom cover 3030 of the fuel tank are provided with annular and inwardly recessed sealing grooves 3040, and sealing rings 3401 are provided in the sealing grooves 3040. The top cover 3020 and the bottom cover 3030 of the fuel tank are connected to the upper and lower ends of the fuel tank wall 3010 through the sealing rings 3401 to prevent fuel leakage in the fuel tank 300.

[0083] Please continue to refer to Figure 8 The present invention provides a schematic diagram of the internal structure of a vector thrust mechanism 400 for an aircraft; it includes a turbojet engine 4020, a drive mechanism 4010 mounted on the head of the turbojet engine 4020, and a vector nozzle 4030 mounted on the tail of the turbojet engine 4020; specifically, the turbojet engine 20 can be selected from models such as SW220B, SW300B, and Jetcat400Pro.

[0084] Please refer to Figure 12This is a schematic diagram of the structure of a vector nozzle 4030 for an aircraft according to an embodiment of the present invention. The vector nozzle 4030 is a cylindrical structure that tapers inward at the bottom. It can be made of high-temperature resistant alloys such as 316 stainless steel, nickel-based alloys, and nickel-copper alloys, and its strength must be guaranteed even at temperatures above 750°C. Since the vector nozzle 4030 is a cylindrical structure that tapers inward at the bottom and is located below the nozzle of the turbojet engine 4020, in its specific installation, it is necessary to ensure that the diameter of the air inlet 4030a of the vector nozzle 4030, i.e., the maximum diameter, is 1.25-1.4 times that of the turbojet engine 4020. The nozzle diameter of the vector nozzle 4030 should be 0. The minimum diameter of the outlet 4030b of the vector nozzle 4030 should also be 1.07-1.2 times the nozzle diameter of the turbojet engine 4020. The overall length of the vector nozzle 4030 should be 1.3-1.6 times the diameter of the inlet 4030a of the vector nozzle 4030. At the same time, the setting of the inlet 4030a and the outlet 4030b should use a smooth transition curve that conforms to hydrodynamics, so as to achieve a better adjustment effect on the thrust direction. When the vector nozzle 4030 returns to the positive state, the central axis of the vector nozzle 4030 should coincide with the central axis of the aircraft in the Z-axis direction.

[0085] Furthermore, the sidewall of the vector nozzle 4030 is provided with several L-shaped connecting arms 4301. Both the sidewall of the vector nozzle 4030 and the connecting arms 4301 are provided with reinforcing ribs 4306. The reinforcing ribs 4306 are provided here to improve the overall strength of the vector nozzle 4030. The connecting arm 4301 includes a connecting end 4301a and an extension end 4301b. The connecting end 4301a of the connecting arm 4301 is fixedly connected to the outer sidewall of the vector nozzle 4030, and the extension end 4301b of the connecting arm 4301 extends along the head direction of the turbojet engine 4020.

[0086] Furthermore, the drive mechanism 4010 is mounted on the head of the turbojet engine 4020, and several connecting rods 4201 are provided between the drive mechanism 4010 and the vector nozzle 4030. It should be noted that the drive mechanism 4010 can be a drive motor, a geared motor, or a servo motor, which are all common mechanisms on the market capable of driving the output shaft to rotate. In this embodiment of the invention, the drive mechanism 4010 is specifically set as a servo motor 4101, and the servo motor 4101 can be of models such as MG995 and MG996R, but is not limited to the above models. In a specific embodiment of the invention, the number of servo motors 4101 is specifically set to two, and there are two connecting rods 4201 between the drive mechanism 4010 and the vector nozzle 4030. The outer wall of the vector nozzle 4030 is also provided with two connecting arms 4301.

[0087] Specifically, each connecting arm 4301 has a first coupling 4302 at its extended end 4301b, and each connecting rod 4201 has a second coupling 4202 adapted to the first coupling 4302 at its end near the vector nozzle 4030. The first coupling 4302 and the second coupling 4202 are hinged together by a universal joint 4203, that is, the connecting rod 4201 and the corresponding connecting arm 4301 are connected by the universal joint 4203. Since the vector nozzle 4030 needs to swing ±23°, a universal joint 4203 needs to be set between the connecting rod 4201 and the connecting arm 4301 to achieve the above effect. The universal joint 4203 can be made of 304 stainless steel. It should be noted that other hinge methods can also be used to connect the connecting rod 4201 and the connecting arm 4301. Any structural improvement that enables the vector nozzle 4030 to swing is feasible.

[0088] Furthermore, each servo motor 4101 has a rocker arm 4103 coaxially rotatably connected to its output shaft 4102. The rocker arm 4103 is used to connect the servo motor 4101 and the connecting rod 4201. The rocker arm 4103 can be of various models, including but not limited to 40T servo arms and 25T servo arms. The rocker arm 4103 is provided with several adjustment holes 4103a. The end of the connecting rod 4201 near the head of the turbojet engine 4020 is provided with a fisheye spherical bearing 4204. The end of the connecting rod 4201 is connected to one of the adjustment holes 4103a of the rocker arm 4103 through the fisheye spherical bearing 4204. As for how the rocker arm 4103 and the connecting rod 4201 are connected through the fisheye spherical bearing 4201, this is a very common application in the prior art and will not be described in detail here.

[0089] Specifically, the connecting rod 4201 is made of stainless steel or carbon fiber. In a specific embodiment of the present invention, the end of the connecting rod 4201 is connected to the adjustment hole 4103a at the far end of the rocker arm 4103 through a fisheye joint bearing 4201. Of course, those skilled in the art can select the position of the adjustment hole 4103a according to the actual situation, and can also improve the bearing connection method of the fisheye joint bearing 4201. As long as the connection between the servo motor 4101 and the connecting rod 4201 can be achieved, and the connecting rod 4201 is allowed to swing slightly, any other connection method is feasible and no further restrictions are imposed here.

[0090] In this embodiment of the invention, a cylindrical support body 4040 is also provided, which is connected to the end of the aircraft body 200. Please refer to... Figure 10This is a schematic diagram of the structure of a vector thrust mechanism 400 for an aircraft provided in an embodiment of the present invention. The support body 4040 is made of aluminum alloy, stainless steel, or composite materials such as carbon fiber. It should be made of low-density and high-strength materials and also needs to have certain high-temperature resistance so that it can still maintain its strength at temperatures above 200°C. The specific material of the support body 4040 is not further restricted here. Then, the drive mechanism 4010, the turbojet engine 4020 and the vector nozzle 4030 are sequentially arranged in the support body 4040.

[0091] Furthermore, the drive mechanism 4010 is mounted on the head of the turbojet engine 4020 via a servo mount 4104 and a servo board 4105. Please refer to [reference needed]. Figure 11 This is a schematic diagram of the structure of an aircraft after the servo motor 4101, servo motor frame 4104, and servo motor board 4105 are installed together, according to an embodiment of the present invention. Specifically, the servo motor frame 4104 and the servo motor board 4105 are both set as plate structures, and the servo motor frame 4104 can be made of aluminum alloy.

[0092] Specifically, since both the servo bracket 4104 and the servo board 4105 are located inside the support body 4040, the maximum diameter of both the servo bracket 4104 and the servo board 4105 does not exceed the inner diameter of the support body 4040. Preferably, the maximum diameter of the servo bracket 4104 is exactly equal to the inner diameter of the support body 4040, that is, the servo bracket 4104 is fixedly connected to the inner wall of the support body 4040. When installing the servo 4101, the two servos 4101 are snapped onto the edge of the servo bracket 4104.

[0093] In the embodiments of the present invention, please refer to Figure 15 This is a schematic diagram of the structure of a servo frame 4104 for an aircraft provided in an embodiment of the present invention. The servo frame 4104 is configured as a similar triangular structure, and an extension 4104a is provided on the edge of the servo frame 4104. The servo frame 4104 is fixedly connected to the inner wall of the support body 4040 through the extension 4104a. Specifically, two servos 4101 are snapped into the adjacent two edges of the servo frame 4104, and the output shaft 102 of the servo 4101 extends away from the center of the servo frame 104. Only when the servos 4101 are set at the edge of the servo frame 4104 can the output shaft 4102 drive the rocker arm 4103 to rotate, thereby causing the connecting rod 4201 to move in space.

[0094] Further, please refer to Figure 13This is a schematic diagram of the structure of a servo motor 4101 of an aircraft provided in an embodiment of the present invention. An extension plate 4101a is also provided at one end of the servo motor 4101 near the output shaft 4102. When fixing the servo motor 4101, the two servo motors 4101 are first fixed at the edge positions of the adjacent two sides of the servo motor frame 4104, and then the servo motor plate 4105 is fixed on the servo motor 4101.

[0095] For details, please refer to Figure 14 This is a schematic diagram of the structure of a servo board 4105 for an aircraft according to an embodiment of the present invention. The servo board 4105 is provided with a plurality of first latching posts 4105a, which are arranged along the edge of the servo 4101 and the extension length of the first latching posts 4105a is equal to the height of the servo 4101. The servo board 4105 is also provided with a plurality of second latching posts 4105b, which are arranged along the edge of the extension plate 4101a and the extension length of the second latching posts 4105b is less than the height of the servo 4101. The servo board 4105 is also provided with a fixing post 4105c, the extension length of the fixing post 4105c is equal to the height of the servo 4101, and the diameter is slightly larger than the diameter of the first latching posts 4105a.

[0096] It should be noted that the servo board 4105 is used to fix the servo 4101 onto the servo frame 4104, and the servo board 4105 is provided with several first locking pins 4105a, second locking pins 4105b and fixing pins 4105c. Therefore, when manufacturing the servo board 4105, it can be made by 3D printing technology or by milling. It can be made of high-strength engineering plastics such as PC, POM, ABS, or lightweight materials such as aluminum alloy. No further restrictions are placed on the specific manufacturing method and materials of the servo board 4105.

[0097] In this embodiment of the invention, when fixing the servo motor 4101, servo motor frame 4104 and servo motor plate 4105, they can be fixed by setting a number of screws (not shown in the figure), such as using screws to pass through the first snap-fit ​​post 4105a and the servo motor frame 4104 in sequence, and / or using screws to pass through the fixing post 4105c and the servo motor frame 4104 in sequence, thereby snapping the two servo motors 4101 between the servo motor plate 4105 and the servo motor frame 4104, and further mounting the two servo motors 4101 at the head position of the turbojet engine 4020.

[0098] Furthermore, a reinforcing plate 4106 is provided above the servo board 4105. The reinforcing plate 4106 is made of carbon fiber or aluminum plate and is used to reinforce the servo 4101, servo frame 4104 and servo board 4105. Optionally, openings 4107 can also be provided on the servo frame 4104, servo board 4105 and reinforcing plate 4106. The openings 4107 can reduce the weight of each component to a certain extent, thereby reducing the overall weight of the aircraft. The shape of the openings 4107 can be circular, triangular, hollow, etc. Those skilled in the art can adjust the shape of the openings 4107 according to actual needs, and no specific limitations are made here.

[0099] Furthermore, a guide plate 4108 is provided below the servo frame 4104. The guide plate 4108 is located above the turbojet engine 4020 and is fixedly connected to the inner wall of the support body 4040. The aforementioned servo frame 4104, servo 4101, and servo plate 4105 are sequentially installed on the top of the guide plate 4108. At the same time, the end of the servo frame 4104 extends outward and is fixedly connected to the inner wall of the support body 4040. Specifically, the side wall of the guide plate 4108 is recessed inward and has a first groove 4108a for the two connecting rods 4201 to pass through, and a second groove 4108b for wires or oil pipes to pass through. The first groove 108a and the second groove 108b are provided here because the guide plate 4108 cannot affect the direction of the internal wiring or internal oil pipes of the aircraft, nor can it affect the spatial position movement required by the two connecting rods 4201.

[0100] Please continue to refer to Figure 8 Next, we will describe how to fix the turbojet engine 4020 inside the support body 4040. A ring-shaped turbojet fixing bracket 4205 is set on the outer wall of the turbojet engine 4020. The inner diameter of the turbojet fixing bracket 4205 is equal to the outer diameter of the turbojet engine 4020, thereby fixing the turbojet fixing bracket 4205 to the outer wall of the turbojet engine 4020. The diameter of the turbojet fixing bracket 4205 also needs to be smaller than the inner diameter of the support body 4040 in order to fix the turbojet engine 4020 with the turbojet fixing bracket 4205 installed inside the support body 4040.

[0101] Furthermore, an L-shaped turbojet mount 4206 is fixed to the outer wall of the turbojet mounting bracket 4205. The turbojet mount 4206 is made of aluminum or stainless steel. In a specific embodiment of the present invention, two turbojet mounts 4206 are arranged along one diameter of the turbojet mounting bracket 205. In principle, the arrangement of these two turbojet mounts 4206 should not affect the spatial movement of the two connecting rods 4201. When connecting, one end of the turbojet mount 4206 is fixedly connected to the outer wall of the turbojet mounting bracket 4205, and the other end of the turbojet mount 4206 is fixedly connected to the inner wall of the support body 4040, thereby fixing the turbojet engine 4020 inside the support body 4040.

[0102] Furthermore, the support body 4040 is provided with several air inlets 4401 near the head of the turbojet engine 4020 for air intake; a fireproof plate 4207 made of high-temperature resistant material is also provided above the turbojet mounting bracket 4205. The thickness of the fireproof plate 4207 is preferably 0.5-1.5mm, and it can be made of high-temperature resistant metals such as 304 or 316L stainless steel, which can still guarantee its strength at temperatures above 400℃. The fireproof plate 4207 is provided here to prevent the flame generated when the turbojet engine 4020 starts from being sucked back and damaged. Specifically, the side wall of the fireproof plate 4207 is also provided with an inwardly recessed third groove 4207a at the position where the two connecting rods 4201 pass through, so as not to affect the spatial movement of the two connecting rods 201.

[0103] Next, we will describe how the vector nozzle 4030 is set inside the support body 4040. An elliptical universal joint 4303 is movably connected to the outer wall of the vector nozzle 4030. The universal joint 4303 is made of a high-temperature resistant metal such as 316L stainless steel, which can still guarantee its strength at temperatures above 600°C. It should be noted that the movable connection here can be understood as a hinge.

[0104] Specifically, the universal joint 4303 is positioned above the connecting arm 4301, and the universal joint 4303 and the vector nozzle 4030 are not in direct contact, but are connected by a pivot pin connection, which is the previously mentioned movable connection; in a specific embodiment provided by the present invention, two first pivot pins 4304 are provided along the diameter direction of the vector nozzle 4030, and the two first pivot pins 4304 pass through the outer sidewalls of the universal joint 4303 and the vector nozzle 4030 in sequence, please refer to Figure 9 This is a schematic diagram of the structure of a vector nozzle 4030 of an aircraft after deflection, provided by an embodiment of the present invention. By movably connecting the vector nozzle 4030 to the universal joint 4303, the vector nozzle 4030 can pitch and rotate around the first shaft pins 4304 at both ends.

[0105] Furthermore, the outer wall of the universal joint 4303 is provided with two “Ω”-shaped external connectors 4305 along its long axis. These external connectors 4305 are made of high-temperature resistant metals such as 316L stainless steel, which can still guarantee their strength at temperatures above 600°C. Since the universal joint 4303 is elliptical, and an ellipse includes a major axis and a minor axis, in this embodiment of the invention, the two external connectors 4305 are located along the major axis of the ellipse, at a position that does not affect the spatial movement of the two connecting rods 4201.

[0106] Specifically, the external connector 4305 includes a fixed end 4305a and two outwardly bent free ends 4305b. The fixed end 4305a is movably connected to the outer wall of the universal joint 4303 through a second shaft pin 4306. That is, the universal joint 4303 and the external connector 4305 are movably connected through the second shaft pin 4306. It should be noted that the movable connection here can be understood as a hinge. The two free ends 4305b are fixedly connected to the inner wall of the support body 4040. Therefore, the two free ends 4305b need to be provided with a certain curvature, which matches the curvature of the inner wall of the support body 4040. The vector nozzle 4030 is mounted below the turbojet engine 4020 by means of screw connection.

[0107] Please refer to Figure 16 This is a schematic diagram of the horizontal angle α formed by the two connecting arms 4301 of an aircraft to the center of the vector nozzle 4030, according to an embodiment of the present invention. In a specific setting, the vector nozzle 4030 and the two connecting arms 4301 are located on the same plane. The center of the vector nozzle 4030 is taken as the endpoint, and the lines connecting the two connecting arms 4301 to the endpoint are taken as the two sides. The angle α formed should be set in the range of 60°-120°.

[0108] It should be noted that the control mechanism is located in the empty space inside the main body 200 of the aircraft after the fuel tank 300 and the vector thrust mechanism 400 are installed, and the center of gravity of the control mechanism should coincide with the central axis of the aircraft in the Z-axis direction. The straight-line distance from the center of gravity of the aircraft to the air outlet 4030b of the vector nozzle 4030 should not be less than 1 / 2.8 of the length of the aircraft's Z-axis, i.e., the height of the aircraft. In other words, the aircraft cannot be set too short and thick.

[0109] Further, please refer to Figure 17 This is a schematic diagram of the structure of a landing mechanism 500 for an aircraft provided in an embodiment of the present invention. The landing mechanism 500 includes a plurality of support structures 5010 disposed on the outer wall of a vector thrust mechanism 400. In a specific embodiment provided by the present invention, the outer wall of the vector thrust mechanism 400 is provided with three support structures 5010.

[0110] Specifically, each support structure 5010 includes a support rod 5011 and a tie rod 5012, wherein one end of the support rod 5011 and the tie rod 5012 are hinged to the outer wall of the vector thrust mechanism 400 through an ear plate 5013, and the other end of the support rod 5011 and the tie rod 5012 are connected through a support foot 5020.

[0111] Furthermore, the support foot 5020 includes an inclined end 5021, a grounding end 5022, and a hinged end 5023. Please refer to further details. Figure 18 This is a schematic diagram of the structure of a support foot 5020 for an aircraft according to an embodiment of the present invention. The inclined end 5021 is inclinedly disposed on the top of the grounding end 5022, and is fixedly connected to the other end of the support rod 5011. The hinged end is disposed on the side wall of the grounding end 5022 and is hinged to the other end of the pull rod 5012. In this embodiment of the invention, the radius of the inscribed circle of the polygon formed by the contact points of several support feet 5020 with the ground is greater than or equal to 1 / 3 of the height of the aircraft's center of gravity; otherwise, the aircraft will be unstable during landing and prone to tipping over. The landing mechanism 500 disclosed in this embodiment of the invention is a rigid landing gear, which has the characteristics of simple structure, light weight, and high strength, and is suitable for use when the aircraft is at low speed or when the aerodynamic shape requirements are low.

[0112] It should be noted that, in addition to the landing mechanism 500 provided in the embodiments of the present invention, existing automatic folding landing gear, such as the automatic folding landing gear used in SpaceX Falcons-9, can also be applied. This type of automatic folding landing gear can open the folded landing gear during the landing phase through a drive structure, such as a motor or push-pull electromagnet, by gravity, hydraulic pressure, or spring action. This type of automatic folding landing gear is manually folded when the aircraft takes off, which can significantly reduce air resistance and optimize flight performance during the aircraft's operation. When the aircraft lands, the previously folded landing gear is opened through a drive structure to achieve a smooth landing. Since the above-mentioned automatic folding landing gear is already existing technology disclosed in the art, its specific structure will not be described in detail here. The above only provides an alternative solution for the landing mechanism 500. Furthermore, when using automatic folding landing gear, since it has an internal drive structure, the landing mechanism 500 can be connected to the main control unit via wired / wireless means as an actuator. The main control unit can then control the operation of the drive structure, thereby achieving the goal of automatically opening the landing gear during landing and ensuring a smooth landing of the aircraft.

[0113] Furthermore, a roll angle control mechanism 600 is provided on the outer wall of the aircraft body 200, near the center of the aircraft. The rotation axis of this roll angle control mechanism 600 coincides with the central axis of the aircraft in the Z-axis direction, and is used to adjust the roll angle attitude of the aircraft. Please refer to further details. Figure 19 This is a schematic diagram of the roll angle control mechanism 600 of an aircraft provided in an embodiment of the present invention; wherein, two adjustment components are arranged symmetrically around the side wall of the mount 6017 with the mount 6017 as the center point. It should be noted that the mount 6017 is not limited to a ring structure. As long as it has a through hole in a certain structure that can be installed on the outer side wall of the aircraft body 200, it is feasible.

[0114] Furthermore, each adjustment component includes a drive motor 6014 and a support arm 6015. The support arm 6015 is disposed on the outer side wall of the mount 6017 and extends outward. The drive motor 6014 is disposed on the support arm 6015. The output shaft of the drive motor 6014 is connected to a propeller 6013. The support arm 6015 is also provided with an electronic speed controller 6016 connected to the drive motor 6014. The electronic speed controller 6016 is used to control the output power and rotation direction of the drive motor 6014, thereby controlling the speed and rotation effect of the propeller 6013 disposed on the drive motor 6014. The connection between the drive motor 6014 and the electronic speed controller 6016 can be an electrical connection or a wireless connection. Those skilled in the art can choose according to actual needs, and no specific limitation is made here.

[0115] Specifically, the aforementioned drive motor 6014 can be configured as a brushless motor, and the electronic speed controller 6016 can be configured as a brushless bidirectional electronic speed controller. The brushless motor is driven by the electronic speed controller 6016, which in turn drives the propeller 6013 on the drive motor 6014 to rotate. Preferably, the drive motor 6014 is located at the end of the support arm 6015. Of course, the drive motor 6014 can also be located at any position on the support arm 6015. In principle, the positions of the drive motor 6014 and propeller 6013 on each adjustment component should meet the requirement of symmetry with the installer 6017 as the central point.

[0116] As an optional embodiment of the present invention, each adjustment component further includes a side baffle 6010, which is a hollow cylindrical structure with a thickness just enough to enclose the propeller 6013 within the side baffle 6010. The side baffle 6010 does not directly contact the propeller 6013 and should not affect the normal rotation of the propeller 6013. A support frame 6012 adapted to the side baffle 6010 is provided between the drive motor 6014 and the support arm 6015. The support frame 6012 can be a plate-like structure or a more aesthetically pleasing structure similar to the propeller 6013. In principle, as long as the side baffle can be enclosed within the propeller 6013, the support frame 6012 can be used. Any support plate 6012 structure selected for mounting the side baffle 6010 around the propeller 6013 is feasible. After mounting, a certain gap is provided between the side baffle 6010 and the propeller 6013 to ensure that the normal rotation of the propeller 6013 is not affected. The support frame 6012 is fixedly connected to the side baffle 6010, and the diameter of the support frame 6012 is not less than the inner diameter of the side baffle 6010. This arrangement can better mount the side baffle 6010 around the propeller 6013, and the position of the side baffle 6010 and the support frame 6012 during the mounting process cannot affect the normal rotation of the propeller 6013.

[0117] Furthermore, to reduce wind resistance and the weight of the roll angle controller, the support frame 6012 can also be made of lightweight, high-modulus sheet metal such as carbon fiber or aerospace aluminum. Both the support frame 6012 and the support arm 6015 employ a hollow structure. As another optional embodiment of the invention, the support frame 6012 is made of lightweight, high-modulus sheet metal such as carbon fiber or aerospace aluminum, and the support arm 6015 is made of lightweight, high-modulus tubing such as carbon fiber or aerospace aluminum. The support frame 6012 and support arm 6015 are made as hollow as possible to minimize wind resistance and the weight of the roll angle controller.

[0118] Furthermore, a support frame 6012 is also provided on the side of the side baffle 6010 away from the support arm 6015. At this time, support frames 6012 are provided on both sides of the side baffle 6010. Both support frames 6012 are provided with hollow structures, and the position of the two support frames 6012 does not affect the normal operation of the propeller 6013.

[0119] It should be further explained that the control mechanism provided in this embodiment of the invention includes a numerical control transmission unit, a main control unit, an attitude position sensor, etc., and the execution mechanism provided in this embodiment of the invention includes a vector thrust mechanism 400 and a roll angle control mechanism 600. Specifically, the numerical control transmission unit controls the servo motor 4101 of the vector thrust mechanism 400 and the drive motor 6014 and ESC 6016 of the roll angle control mechanism 600. When the landing mechanism 500 is set as an automatic folding landing gear such as that of SpaceX Falcons-9, the landing mechanism 500 can also act as an execution mechanism to execute the control information initiated by the numerical control transmission unit. The above-mentioned aircraft provided in this embodiment of the invention can be specifically applied to reusable launch vehicles.

[0120] It is understood that the above embodiments are merely exemplary embodiments used to illustrate the principles of the present invention, and the present invention is not limited thereto. For those skilled in the art, various modifications and improvements can be made without departing from the spirit and essence of the present invention, and these modifications and improvements are also considered to be within the scope of protection of the present invention.

Claims

1. An aircraft, characterized in that, Including the fairing, the main body of the aircraft, and the vector thrust mechanism; The aircraft body is equipped with a fuel tank, which is used to provide fuel for the vector thrust mechanism. One end of the aircraft body is connected to the fairing, and the other end of the aircraft body is connected to the vector thrust mechanism; the vector thrust mechanism includes a turbojet engine and a vector nozzle, the vector nozzle includes a drive mechanism mounted on the head of the turbojet engine and a vector nozzle mounted on the tail of the turbojet engine; the vector thrust mechanism is used to adjust the current attitude position of the aircraft. The vector nozzle has two L-shaped connecting arms on its sidewall. Each connecting arm has a connecting end and an extension end. The connecting end of the connecting arm is fixedly connected to the outer sidewall of the vector nozzle, and the extension end of the connecting arm extends along the head direction of the turbojet engine. Two connecting rods are provided between the drive mechanism and the vector nozzle. The connecting rods are hinged to the extension ends of the connecting arms. The included angle formed by the two connecting arms to the center of the vector nozzle is in the range of 60°-120°. An elliptical universal joint is movably connected to the outer wall of the vector nozzle. Two first shaft pins are arranged along the diameter direction of the vector nozzle. The two first shaft pins pass through the universal joint and the outer wall of the vector nozzle in sequence to movably connect the vector nozzle inside the universal joint. Two external connectors are arranged along the long axis direction of the outer wall of the universal joint. Each external connector includes a fixed end and two outwardly bent free ends, and the two free ends are fixedly connected to the inner wall of the support body. The outer side wall of the main body of the aircraft is provided with a roll angle control mechanism, which is used to adjust the current roll angle attitude of the aircraft. The outer wall of the vector thrust mechanism is provided with a landing mechanism, which includes several support structures arranged around the outer wall of the vector thrust mechanism to stabilize the overall attitude of the aircraft when it lands.

2. The aircraft as described in claim 1, characterized in that, The turbojet engine is connected to the oil outlet of the fuel tank; The output shaft of the drive mechanism is coaxially rotatably connected to a rocker arm, and the other end of the connecting rod is connected to the rocker arm bearing.

3. An aircraft as described in claim 2, characterized in that, The fuel tank includes a fuel tank wall, a fuel tank top cover, and a fuel tank bottom cover. The fuel tank top cover is provided with an oil inlet, and the oil outlet is located at the bottom of the fuel tank bottom cover. The fuel tank is equipped with a connecting pipe and a partition. The two ends of the connecting pipe are connected to the top cover and the bottom cover of the fuel tank, respectively. One end of the partition is connected to the inner wall of the bottom cover of the fuel tank, and the other end of the partition forms an angle with the plane of the bottom cover of the fuel tank, and there is a gap between the partition and the outer wall of the connecting pipe.

4. An aircraft as described in claim 3, characterized in that, The fuel tank is located at the center of gravity of the aircraft and provides fuel for the turbojet engine.

5. An aircraft as described in claim 1, characterized in that, Each of the support structures includes a support rod and a tie rod, one end of which is hinged to the outer wall of the vector thrust mechanism via an ear plate, and the other end of which is connected via a support foot. The support foot includes an inclined end, a grounding end, and a hinged end. The inclined end is inclinedly disposed on the top of the grounding end and fixedly connected to the other end of the support rod. The hinged end is disposed on the side wall of the grounding end and hinged to the other end of the pull rod.

6. An aircraft as described in claim 5, characterized in that, The radius of the inscribed circle of the polygon formed by several of the contact points between the support structures and the ground is greater than or equal to 1 / 3 of the height of the aircraft's center of mass.

7. An aircraft as described in claim 1, characterized in that, The roll angle control mechanism includes a ring-shaped mount and two adjustment components, which are symmetrically arranged around the side wall of the mount with the mount as the center point. Each of the adjustment components includes a drive motor and a support arm, the support arm being disposed on the outer side wall of the mount and extending outward, and an electric speed controller connected to the drive motor being disposed on the support arm; The drive motor is located at the end of the support arm, and its output shaft is connected to a propeller.

8. An aircraft as described in claim 7, characterized in that, The roll angle control mechanism is located near the center of gravity of the aircraft.

9. An aircraft as described in claim 1, characterized in that, The aircraft body also includes a control mechanism, which comprises an attitude and position sensor, a main control unit, and a numerical control transmission unit. The attitude position sensor is used to acquire the current attitude position information of the aircraft; the numerical control transmission unit is used to receive the command information sent by the ground station and transmit the command information to the main control unit; the main control unit is used to generate control information based on the current attitude position information of the aircraft and the command information, and send the control information to the actuator. The actuator includes the vector thrust mechanism and the roll angle control mechanism, and the actuator is used to receive and execute the control information.

10. An aircraft as described in claim 9, characterized in that, The attitude position sensor, the vector thrust mechanism, the roll angle control mechanism, and the numerical control transmission unit are connected to the main control unit via wireless and / or wired means, and the numerical control transmission unit and the ground station are wirelessly connected.

Images