Foldable semi-tilting composite wing aircraft and control method thereof

Abstract

According to the foldable semi-tilting composite wing aircraft and the control method thereof, streamline design is adopted for the whole aircraft body, integrated aerodynamic layout of four-axis eight-propeller tilting rotors and vertical fins is matched, and through redundancy design and airflow optimization, the stability and the wind resistance are remarkably improved. The half tilting function of the rotor wing can realize the conversion from a helicopter mode to a fixed wing mode, and the fuselage in the whole working envelope is always kept horizontal. The overall size of the rotorcraft arms and the wings can be reduced through a folding mechanism and a folding mode, and the transportation requirement is met. The aircraft is provided with a battery management system (BMS), can monitor the state of a battery in real time in the whole journey, has the functions of low electric quantity reminding and emergency landing, and adopts a power distribution optimization scheme for different flight modes, so that the flight efficiency is improved, the energy loss is reduced, and the task endurance is increased. In the control aspect, the flight mode, the power requirement and the BMS response form a closed loop, driving force is obtained by adjusting the tilting angle and the rotating speed of a propeller during forward flight, and mode conversion is achieved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aircraft, in particular to a foldable semi-inclined compound wing aircraft and a control method thereof. BACKGROUND

[0002] With the rapid development of electric aviation technology, electric vertical take-off and landing aircraft (eVTOL) has made a series of key technological breakthroughs, giving birth to a new concept of urban air mobility (UAM). At present, eVTOL aircraft is mainly divided into multi-rotor configuration, compound wing configuration and tilt wing configuration. However, although the technology is constantly improving, eVTOL aircraft still faces many challenges in practical application.

[0003] For example, the Chinese patent with publication number CN114671006B introduces a kind of aircraft rotor branch arm folding mechanism, which adopts a kind of folding mechanism outside the way and has a complex shape design, which will produce additional aerodynamic interference in flight. The rotor folding mechanism of the aircraft design adopts a kind of built-in folding mechanism, the external fixation adopts a kind of thread connection form, and the shape is still cylindrical, and the aerodynamic shape is simple.

[0004] The Chinese patent with publication number CN118770544A introduces an electric vertical take-off and landing tandem wing aircraft with full tilt wing. The aircraft adopts four sets of tilt mechanism to realize the tilt of the wing and the rotor, reduces the production cost and improves the safety redundancy. It has rotor flight mode and fixed wing flight mode, can vertically take off and fly at high speed. In addition, the aircraft has various control methods in different flight modes, and the wing does not need aerodynamic rudder, reduces the mechanical mechanism and the weight of the additional structure. The first and second wings and the rotor components are arranged staggered, which improves the aerodynamic efficiency, and the landing gear has the functions of support and drooping tail, which has the advantages of cost and redundancy. However, this design still has problems such as large change of aircraft attitude angle, low efficiency of multi-rotor aircraft, large difficulty of full tilt transition control, difficulty of aircraft transportation, and insufficient power storage for commercial operation.

[0005] Secondly, the Chinese patent with publication number CN118877199A introduces an electric vertical take-off and landing aircraft using vector thrust technology. The aircraft realizes accurate control of compound wing tilt through vector thrust technology, and can flexibly convert flight attitude. Its power redundancy design sets up multiple independent power propulsion systems according to seat version, which can redundantly set up multiple power units to ensure flight safety. In addition, the aircraft has full-automatic and manual driving modes, the front seat door adopts special design, and a large number of carbon fiber materials are used in the aircraft, which optimizes the flight performance. However, this aircraft also has problems such as short endurance time, short mileage, slow flight speed and limited flight distance.

[0006] Furthermore, Chinese patent publication number CN116853486A introduces an electric vertical-wing tilt-quadrotor aircraft that combines tilt-rotor technology with conventional wing design. The aircraft changes the direction of the rotors by tilting the vertical wings to achieve different flight modes. The tilt disc drives the tilt vertical wings to rotate, allowing the rotors to turn in any direction. The tail wing and main wing are connected to the horizontal rudder, and the tilt vertical wing is connected to the vertical rudder. Together with the tilt disc and the rotor, a variety of flight attitude controls can be achieved, such as pitching up, pitching down, flying left and right, and turning when hovering. When the four rotors stop working, the aircraft can glide on the main wings and tail wing, effectively reducing the chance of crash and the risk of casualties in the event of a crash. Although the aircraft has a unique power mode and good control performance, and a high safety factor, it still faces problems such as short flight time and difficulty in transportation.

[0007] Therefore, the transportation cost of traditional aircraft is extremely high and relatively difficult due to the size issues of fixed-wing and multi-rotor aircraft. On the other hand, when flying forward, the aircraft needs to tilt the fuselage at a certain angle to obtain the forward driving force, which requires the pilot to tilt synchronously with the fuselage, affecting the driving experience. In addition, although various wings have been designed or composite wings have been used to increase the flight time, the operating time still needs to be improved compared to commercial operation missions. After completing a flight, a long charging time is required, which further limits its feasibility of commercial operation.

[0008] In summary, the key challenges that eVTOL aircraft urgently need to address include large changes in attitude angle during flight, difficulty in controlling full-tilt transitions, transportation difficulties caused by the large size of the aircraft, and insufficient power storage for commercial operations. Summary of the Invention

[0009] In order to improve the problems of large attitude angle changes during aircraft flight, difficulty in full-tilt transition control, and transportation difficulties caused by the large size of the aircraft, the present application provides a foldable semi-tilt composite wing aircraft and a control method thereof.

[0010] In a first aspect, the present application provides a foldable semi-tilting composite wing aircraft, which adopts the following technical solution: A foldable semi-tilting composite wing aircraft comprises a fuselage, wings, a vertical tail, a tilt mechanism, an arm folding mechanism, a wing folding mechanism, a rotor power system, and a flight control system. The rotor power system comprises: There are two front rotor mounting frames and two rear rotor mounting frames, each of which is provided on both sides of the fuselage, and the horizontal height of the rear rotor mounting frame is greater than that of the front rotor mounting frame; The upper rotor and the lower rotor are mounted on the upper and lower sides of the free ends of the front rotor mounting frame and the rear rotor mounting frame; The tilting mechanism comprises: Four arms are provided, one end of which is rotatably mounted on the side of the fuselage and the other end of which is mounted on a front rotor mounting frame or a rear rotor mounting frame; A linear actuator is fixed to the fuselage, and an output end thereof is connected to a tilt link, the other end of which is hinged to the fuselage arm; The flight control system is configured to control the linear actuator to increase the forward tilt angle of the front rotor mounting frame and simultaneously increase the rotation speed of all the upper rotors and the lower rotors on the front rotor mounting frame when in a tilted forward flight state, and to balance the fuselage by adjusting the tilt angle of the rear rotor mounting frame and increasing the rotation speed of all the upper rotors and the lower rotors on the rear rotor mounting frame.

[0011] Furthermore, the machine arm is configured as a two-section machine arm, and the machine arm folding mechanism includes a first folding member and a second folding member provided at adjacent ends of the two sections of the machine arm, and the first folding member is hinged to the second folding member; The end of the machine arm provided with the first folding component is sleeved with a sleeve lock buckle, and the end of the machine arm provided with the second folding component is fixedly connected with an external screw barrel, and the sleeve lock buckle is threadedly connected to the external screw barrel.

[0012] Furthermore, the wing includes a central wing and two side wings hinged to both ends of the central wing, and the central wing is fixed to the top of the fuselage; The wing folding mechanism comprises: a central connecting rod connected to the end of the central wing; Side connecting rods, fixedly connected to the ends of the side wings, and hingedly connected to the central connecting rod; The side fixing pins pass through the assembled center wing and side wings.

[0013] Furthermore, the wing folding mechanism further comprises: A center positioning cylinder is fixedly connected to the middle of the center wing. The end of the center connecting rod is plugged and adapted to the end of the center positioning cylinder, and a first pin hole is formed through the corresponding ends of the two. A central fixing pin is plugged and connected to the top of the central wing and correspondingly plugged and connected to the first pin hole. Two central fixing pins are provided.

[0014] Furthermore, the wing is a high-speed airfoil and is set at a mounting angle of 6°, and the wing is located between the front rotor mounting frame and the rear rotor mounting frame.

[0015] Furthermore, a power box body with a tail opening is built into the rear of the fuselage, a battery pack is slidably arranged in the power box body, and an automatic lock for locking the battery pack is provided at the open end of the power box body; A hatch is provided at the rear of the fuselage corresponding to the position of the power box body, and a hatch cover and a linear actuator for driving the hatch cover to flip to close or open the hatch are hingedly connected to the hatch.

[0016] Furthermore, a canopy is arranged on the upper side of the front section of the fuselage, and the propeller plane of the upper rotor is located below the canopy.

[0017] Furthermore, a battery management system is also included, which adopts the BMS architecture of a multi-pack system and includes: The first layer of master control is responsible for communicating with the flight control system, managing the multi-packet charging logic, coordinating the key information of the single packet and reporting it to the flight control system, and performing insulation testing after parallel high voltage connection; The second-layer master is responsible for single-packet energy management and fault diagnosis, reporting single-packet information to the first-layer master; and is responsible for single-packet thermal management and controlling the high-voltage power on and off during single-pack charging; The third-layer slave controller is responsible for collecting the cell voltage in a single pack, collecting the module temperature, and reporting the voltage / temperature data to the second-layer master controller.

[0018] In a second aspect, the present application provides a method for controlling a foldable semi-tilt composite wing aircraft, based on the above-mentioned foldable semi-tilt composite wing aircraft, comprising the following steps: Determine the flight mode based on the aircraft's flight speed; Based on the flight mode, power demand and battery management system response, a closed loop is formed to dynamically adjust the tilt angle of the front and rear rotors and the propeller speed in different flight modes; When the aircraft is in forward flight, the tilt mechanism is used to adjust the front and rear rotors to tilt forward. The flight control system controls the front rotor speed to increase to increase forward thrust, and the rear rotor speed is synchronously adjusted to balance the fuselage pitch moment. The flight control system simultaneously sends instructions to the battery management system. The first-layer master control of the battery management system calculates the discharge rate corresponding to the target power of a single packet, and activates the high-power mode through the second-layer master control of the battery management system. The sampling period of the third-layer slave control of the battery management system is shortened to 50ms. When the aircraft transitions from a hovering state to a forward flight state, the tilt angle is first increased, and then the rotation speed is gradually increased.

[0019] Further, when the aircraft is in a hovering state and a vertical take-off and landing stage, the flight control system controls all rotor tilt angles to be 0°, and the upper rotor and the lower rotor on the same arm rotate in opposite directions; at the same time, the flight control system sends instructions to the battery management system, the master control layer adjusts the power distribution of the multiple battery packs and divides them into four packs by receiving the "hovering" instruction from the flight control system, and the slave control layer shortens the sampling period to 100 ms; When the aircraft is in a high-speed cruising stage, the wings generate starting lift due to the forward flight speed, the flight control system controls all rotor tilt angles to further increase, while reducing the rotation speed of all propellers, and the flight control system calculates the wing lift ratio in real time and dynamically allocates rotor power; wherein, if one battery pack fails, the remaining three battery packs output the same power, and the master control layer forces the remaining battery pack SOC to maintain ≥23.4%; when any battery pack SOC <20%, the master control layer sends a "return request" to the flight control system, while limiting the maximum power of a single pack ≤30kW.

[0020] In summary, the present application includes at least one of the following beneficial technical effects: 1. The foldable semi-tilt composite wing aircraft of the present application adopts a composite configuration design combining tilting rotors and fixed wings, which belongs to an innovative flight principle and aerodynamic layout design. Compared with multi-rotor aircraft, this foldable composite wing aircraft has more advantages in terms of endurance and range, and can fly for a longer time and a longer distance. Compared with a composite fixed-wing aircraft, it does not have additional dead weight and control surfaces, which effectively reduces the potential risks during flight, making safety more prominent; 2. Compared with the full-tilt configuration, the foldable semi-tilt composite wing aircraft of the present application does not have complex hub structures and control surfaces, which not only reduces the difficulty of control and makes flight control more convenient, but also further improves safety. One of the significant features of this aircraft is that the fuselage remains horizontal throughout the entire operating envelope. This feature is of great significance for the transportation of valuable instruments, valuable drugs, and other items that require high flight stability; 3. The foldable semi-tilt composite wing aircraft of the present application has the characteristic of easy folding, which greatly alleviates the difficulties faced by the aircraft during transportation or storage, making it easier to manage and store in a non-flying state. The automatic quick battery replacement module and the BMS battery management system provided by the aircraft can effectively extend the endurance time, improve the operating efficiency, and thus increase the overall benefits of low-altitude economy. BRIEF DESCRIPTION OF DRAWINGS

[0021] In order to more clearly illustrate the technical solutions in the present invention or the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0022] Figure 1 It is a schematic diagram of the overall structure of an embodiment of the present application; Figure 2 This is a side view of the overall structure of an embodiment of the present application; Figure 3 is a schematic structural diagram of the tilting mechanism of an embodiment of the present application; Figure 4 is a side view of the embodiment of the present application in a tilted forward flight state; Figure 5 This is a schematic structural diagram of the arm folding mechanism of an embodiment of the present application; Figure 6 is a schematic structural diagram of the wing folding mechanism of an embodiment of the present application; Figure 7 This is a structural diagram of the automatic rapid battery replacement module according to an embodiment of the present application; Figure 8 This is a diagram of the battery management system architecture of an embodiment of the present application; Figure 9 It is a control logic diagram of the flight control system of an embodiment of the present application.

[0023] Reference numerals: 1. Fuselage; 11. Wings; 111. Center wing; 112. Side wings; 12. Vertical tail; 13. Canopy; 14. Landing gear; 21. Front rotor mounting bracket; 22. Rear rotor mounting bracket; 23. Upper rotor; 24. Lower rotor; 3. Arm; 311. Linear actuator; 312. Tilt link; 321. First folding member; 322. Second folding member; 323. Sleeve lock; 324. External screw; 41. Center connecting rod; 42. Side connecting rod; 43. Side fixing pin; 44. Center positioning cylinder; 45. Center fixing pin; 46. Side positioning cylinder; 51. Power supply box; 52. Battery pack; 53. Automatic lock; 54. Hatch cover; 55. Linear actuator. DETAILED DESCRIPTION

[0024] To make the objectives, technical solutions, and advantages of the present invention more clear, the technical solutions of the present invention will be clearly and completely described below in conjunction with the accompanying drawings. Obviously, the embodiments described are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts shall fall within the scope of protection of the present invention.

[0025] Reference Figure 1 and Figure 2 The present application discloses a foldable, semi-tilting composite wing aircraft comprising a fuselage 1, wings 11, a vertical tail 12, a tilt mechanism, an arm-folding mechanism, a wing-folding mechanism, a rotor propulsion system, and a flight control system. The fuselage 1 is streamlined, but not limited to a streamlined fuselage 1. An openable canopy 13 is located on the upper front section of the fuselage 1. The canopy 13 opens upward, and a cabin within the canopy 13 houses seats and the aircraft's control module. The rear section of the fuselage 1 houses system modules such as the power supply, flight control system, electromechanical system, and avionics system. A landing gear 14 is located on the underside of the fuselage 1. The landing gear 14 has four landing points, with the legs connected by aluminum alloy connectors. The bottom of the landing gear 14 consists of four aluminum alloy plates welded to the four legs. The upper portion of the landing gear 14 is bolted to the underside of the fuselage 1. There are two vertical tails 12 symmetrically arranged on both sides, and are connected to the rear of the fixing frame of the fuselage 1 by bolts.

[0026] The rotor propulsion system includes: The front rotor mounting frame 21 and the rear rotor mounting frame 22 are distributed on the front and rear sides of the fuselage 1, and two are provided and arranged on both sides of the fuselage 1. The horizontal height of the rear rotor mounting frame 22 is greater than that of the front rotor mounting frame 21. The upper rotor 23 and the lower rotor 24 are mounted on the upper and lower sides of the free ends of the front rotor mounting frame 21 and the rear rotor mounting frame 22. The upper rotor 23 and the lower rotor 24 both include motors and propellers. In order to provide the pilot with a better forward view, the propeller plane of the upper rotor 23 is located below the canopy 13.

[0027] Reference Figure 1 and Figure 3 , the tilting mechanism includes: Four arms 3 are provided and are symmetrically arranged on both sides of the fuselage 1. One end of the arm 3 is rotatably mounted on the side of the fuselage 1, and the other end is mounted on the front rotor mounting frame 21 or the rear rotor mounting frame 22; The linear actuator 311 is fixed to the fuselage 1, and its output end is connected to the tilt link 312, the other end of which is hinged to the arm 3. The rotor power system can tilt the front and rear rotors up to ±30° through the tilt mechanism.

[0028] Reference Figure 1 and Figure 4 The flight control system is configured to control the linear actuator 311 to increase the forward tilt angle of the front rotor mounting frame 21 and simultaneously increase the rotation speed of all upper rotors 23 and lower rotors 24 on the front rotor mounting frame 21 when in the tilted forward flight state, and to balance the fuselage 1 by adjusting the tilt angle of the rear rotor mounting frame 22 and increasing the rotation speed of all upper rotors 23 and lower rotors 24 on the rear rotor mounting frame 22.

[0029] Reference Figure 1 and Figure 5 The arm 3 is configured in two sections. The arm folding mechanism includes a first folding member 321 and a second folding member 322, located adjacent to each other. The first folding member 321 and the second folding member 322 are hingedly connected. The end of the arm 3 provided with the first folding member 321 is fitted with a sleeve lock 323, while the end of the arm 3 provided with the second folding member 322 is fixedly connected to an external screw 324, with the sleeve lock 323 being threadedly connected to the external screw 324. To fold the arm 3, the sleeve lock 323 is first rotated to remove it from the external screw 324, and then the end of the arm 3 away from the fuselage 1 is rotated to complete the folding.

[0030] Reference Figure 1 and Figure 6 The wing 11 includes a central wing 111 and two side wings 112 hinged at both ends of the central wing 111. The central wing 111 is fixed to the top of the fuselage 1; the wing 11 is a high-speed airfoil and is set at an installation angle of 6°. The wing 11 is located between the front rotor mounting frame 21 and the rear rotor mounting frame 22.

[0031] The wing folding mechanism includes: a central connecting rod 41 connected to the end of the central wing 111, one end of which has a hinge joint; The side connecting rod 42 is fixed to the end of the side wing 112, and the side connecting rod 42 is hinged to the hinge joint on the central connecting rod 41; The side fixing pin 43 passes through the assembled center wing 111 and the side wing 112; The center positioning cylinder 44 is fixed to the middle of the center wing 111. The end of the center connecting rod 41 away from the side connecting rod 42 is plugged and adapted to the end of the center positioning cylinder 44. A first pin hole is formed through the corresponding ends of the center connecting rod 41 and the center positioning cylinder 44. The end of the center connecting rod 41 also has a third pin hole formed through it, the axis of which is perpendicular to the axis of the first pin hole. A central fixing pin 45 is plugged into the top of the central wing 111 and correspondingly plugged into the first pin hole. Two central fixing pins 45 are provided. The side positioning cylinder 46 is fixed to the end of the side wing 112 and is inserted into the end of the center wing 111. A second pin hole is formed through the side positioning cylinder 46. When the side wing 112 is completely combined with the center wing 111, the side fixing pin 43 is inserted into the center wing 111 and is connected with the second pin hole.

[0032] When the wing 11 is installed, the center connecting rod 41 is turned over so that the center connecting rod 41 hinged to the side wing 112 is in line with the side wing 112. Then, the side wing 112 is combined with one end of the center wing 111. At this time, the center connecting rod 41 is inserted into the center wing 111 and is connected with one end of the center positioning cylinder 44. The side positioning cylinder 46 is also inserted into the end of the center wing 111. Then, the center fixing pin 45 is inserted into the center wing 111 and is connected with the first pin hole on the center positioning cylinder 44 and the center connecting rod 41, so as to fix the center connecting rod 41 and prevent it from being turned over. Then, the side fixing pin 43 is inserted into the center wing 111 and is connected with the second pin hole on the side positioning cylinder 46, so as to fix the side wing 112 and ensure that the side wing 112 is stably connected with the center wing 111.

[0033] When the wing 11 is folded, the two side fixing pins 43 and the two center fixing pins 45 are pulled out first. Then, the side wing 112 is moved away from the center wing 111 until the side positioning cylinder 46 is pulled out of the end of the center wing 111. Then, the center connecting rod 41 is turned over by 90°, so that the side wing 112 is perpendicular to the center wing 111. Then, the side wing 112 is turned over by 90° by means of the hinge on the center connecting rod 41, so that the side wing 112 is folded. In order to prevent the folded side wing 112 from slipping off, the center fixing pin 45 can be further inserted into the first pin hole of the center positioning cylinder 44 and passes through the third pin hole on the center connecting rod 41. At this time, the center connecting rod 41 is locked, so that the folded state of the side wing 112 is locked.

[0034] Further, referring to Figure 1 and Figure 7 , the fuselage 1 is further provided with an automatic quick battery replacement module. Specifically, a tail opening power box 51 is arranged in the tail of the fuselage 1. A battery pack 52 is slidably arranged in the power box 51. An automatic lock 53 for locking the battery pack 52 is arranged at the opening end of the power box 51. The automatic lock 53 is a conventional technical means in the art, which will not be described here. A hatch is formed at the part of the tail of the fuselage 1 corresponding to the power box 51. A hatch cover 54 is hinged to the hatch. A direct-acting actuator 55 is arranged to turn over the hatch cover 54 to close or open the hatch.

[0035] In addition, the aircraft of the present application further comprises a battery management system. Referring to Figure 7 and Figure 8The battery management system (BMS) monitors the key parameters of the battery pack 52 in real time to estimate its status and diagnose faults. It then selects the appropriate power supply solution based on different modes, improving flight efficiency and reducing energy loss. The power box 51 features a multi-porous design for excellent heat dissipation, and a metal dust screen is installed on the inner surface of the power box 51 to protect the battery management system. Specifically, the battery management system utilizes a multi-pack BMS architecture, including: The first layer of master control is responsible for communicating with the flight control system, managing the multi-packet charging logic, coordinating the key information of a single packet and reporting it to the flight control system, and performing insulation testing after parallel high voltage connection; The second-layer master control is responsible for single-packet energy management and fault diagnosis, reporting single-packet information to the first-layer master control; it is also responsible for single-packet thermal management and controlling the high-voltage power on and off during single-pack charging; The third-layer slave controller is responsible for collecting the cell voltage in a single pack, collecting the module temperature, and reporting the voltage / temperature data to the second-layer master controller.

[0036] When performing automatic fast battery replacement operation, perform the following steps: Hatch opening and battery unlocking (0-10s): The linear actuator 55 contracts, pulling the hatch cover 54 to rotate 90° around the hinge toward the inside of the fuselage 1, completely exposing the battery replacement interface; the automatic lock 53 unlocks the battery pack 52, and the battery pack 52 is unlocked and slides along the guide rail in the power box body 51 to a pluggable position; the battery management system disconnects the power supply circuit of the battery pack 52 and sends a "power off completion" signal to the battery replacement workbench.

[0037] Battery swap workstation interaction (10-30s): The aircraft sends battery swap task information, including battery model, remaining power, and fault code, to the battery swap workstation via wireless communication; the battery swap workstation robotic arm grabs the old battery pack 52, inserts the new battery pack 52, and confirms the physical connection is in place; the battery management system verifies the parameters of the new battery pack 52 (voltage, capacity, safety certification), and establishes a power supply connection after a successful match.

[0038] Hatch closing and battery locking (30-40s): The automatic lock 53 locks the battery pack 52 with a locking force ≥ 50N; the linear actuator 55 extends, pushing the hatch cover 54 to reset and seal, with a closing accuracy of ±1mm; the cockpit interface displays "Battery replacement completed", the battery management system switches to the new battery pack 52 for power supply, and updates the real-time power data.

[0039] And the automatic quick battery replacement module has a safety protection mechanism during the entire battery replacement process, as follows: when the hatch cover 54 is not completely closed, the automatic lock catch 53 cannot lock the battery pack 52, and the aircraft is prohibited from starting the power system; when the battery pack 52 is not locked, the battery replacement hatch cover 54 remains open and an alarm is sounded. During the battery replacement process, the flight control system suspends the tilt mechanism and rotor power output, leaving only basic avionics equipment power supply. After the new battery pack 52 is connected, it needs to pass through 3 cycles of voltage detection (interval 2s) to ensure that there is no abnormality before restoring full system power supply. In the event of a sudden power failure during battery replacement, the standby lithium battery pack maintains the open state of the hatch cover 54 for 10 minutes, facilitating manual intervention; when communication interruption of the battery replacement workstation is detected, the last valid battery parameter stored locally is automatically switched to.

[0040] The embodiment of the application also discloses a folding semi-tilt composite wing aircraft control method, based on the folding semi-tilt composite wing aircraft, referring to Figure 1 、 Figure 4 and Figure 9 , the following technical scheme is adopted: A folding semi-tilt composite wing aircraft control method comprises the following steps: determining a flight mode according to a flight speed of the aircraft; forming a closed loop based on the flight mode, power demand and battery management system response, and dynamically adjusting the tilt angles of the front and rear rotors and the propeller speeds in different flight modes.

[0041] Specifically, the aircraft of the application is a four-axle eight-propeller left-right symmetrical aerodynamic layout with the rear rotor higher than the front rotor. Eight motors drive eight propellers, and the eight motors independently control the corresponding rotors. The increase or decrease of the motor speed will directly cause the corresponding change of the rotor speed, and there are two groups of rotor tilt mechanisms corresponding to the tilt of the front and rear rotors, and the tilt control of the front and rear rotors is also independent. The two groups of tilt mechanisms are tilted forward, and under the cooperation of the eight motors, the rotor tilt mechanism and the battery management system (BMS), the aircraft can exhibit good performance and stability in different flight stages and flight modes.

[0042] When the aircraft is hovering or performing vertical takeoff and landing (VTOL), the flight control system controls the tilt angle of all rotors to 0°, and the upper rotor 23 and lower rotor 24 on the same arm 3 rotate in opposite directions to offset torque and improve hovering stability. The avionics system monitors attitude sensor data in real time. Simultaneously, the flight control system sends instructions to the battery management system, which synchronously monitors the voltage and temperature of the battery pack 52 and outputs a "battery health" signal to the flight control system. The first-level master control receives the "hover" command from the flight control system and adjusts the power distribution among the multiple packs, evenly distributing it among the four packs to prevent overloading of a single pack. The third-level slave control system collects the battery cell voltage at high frequency. The sampling period is shortened to 100ms to prevent voltage drops caused by high current surges. The propeller speed of each motor is fine-tuned to compensate for minor offsets and ensure hovering stability.

[0043] When the aircraft is in forward flight, specifically transitioning from hovering to tilted forward flight, the tilt mechanism adjusts the forward tilt of the front and rear rotors. The flight control system increases the front rotor's speed to increase forward thrust, while the rear rotor's speed is simultaneously adjusted to balance the fuselage's pitch moment. Because the rear rotor is vertically higher than the front rotor, the rear rotor's tilt angle and motor speed are different from those of the front rotor. These adjustments are made based on the aircraft's attitude, ensuring that the component of the rotor's thrust in the direction of gravity equals the aircraft's own gravity and that the cockpit remains horizontal. Furthermore, the tilt of the rotors also provides the aircraft with the force component required for forward flight. The flight control system also sends instructions to the battery management system. The first-level master control of the battery management system calculates the discharge rate corresponding to the target power of each battery pack and activates high-power mode through the second-level master control of the battery management system. The third-level slave control of the battery management system shortens the sampling period to 50ms, monitors the voltage drop of the battery cells in real time, calculates the energy consumption of each battery pack every 10 seconds, and updates the remaining SOC. The first-level master control compares the SOC difference between each battery pack. If it exceeds 3%, it triggers active balancing by the second-level master control. If one battery pack fails, the first-level master control immediately redistributes power to the remaining three packs and notifies the control system to "adjust rotor power distribution" to compensate for the torque imbalance. When the aircraft transitions from hovering to forward flight, it first increases the tilt angle and then gradually increases the rotation speed.

[0044] When the aircraft is in the high-speed cruising stage, wing 11 generates starting lift due to the forward flight speed. The flight control system controls all rotor tilt angles to further increase, while reducing the speed of all propellers. The flight control system also calculates the lift ratio of wing 11 and dynamically distributes rotor power. The first-layer master control evenly distributes power to four packs and enters "economic mode". The sampling period of the third-layer slave control is extended to 500ms to reduce the power consumption of the controller; the second-layer master control fails to receive data from the third-layer slave control for three consecutive times, and determines that "single pack failure"; the second-layer master control disconnects the main positive relay of the faulty pack and reports to the first-layer master control. The first-layer master control will adjust the power of the remaining three packs and send a "power redundancy enable" signal to the flight control system. After receiving the signal, the flight control system adjusts the rotor speed to compensate for the pitch moment deviation caused by uneven power; specifically, if one battery pack 52 fails, the remaining three battery packs 52 each output the same power, and the first-layer master control forces the SOC of the remaining battery packs 52 to maintain ≥23.4%; when the SOC of any battery pack 52 is less than 20%, the first-layer master control sends a "return request" to the flight control system and limits the maximum power of a single pack to ≤30kW to avoid deep discharge.

[0045] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A foldable semi-tilting composite wing aircraft, comprising a fuselage, wings, a vertical tail, a tilt mechanism, an arm folding mechanism, a wing folding mechanism, a rotor power system, and a flight control system, characterized in that: The rotor power system includes: There are two front rotor mounting frames and two rear rotor mounting frames, each of which is provided on both sides of the fuselage, and the horizontal height of the rear rotor mounting frame is greater than that of the front rotor mounting frame; The upper rotor and the lower rotor are mounted on the upper and lower sides of the free ends of the front rotor mounting frame and the rear rotor mounting frame; The tilting mechanism comprises: Four arms are provided, one end of which is rotatably mounted on the side of the fuselage and the other end of which is mounted on a front rotor mounting frame or a rear rotor mounting frame; A linear actuator is fixed to the fuselage, and an output end thereof is connected to a tilt link, the other end of which is hinged to the fuselage arm; The flight control system is configured to control the linear actuator to increase the forward tilt angle of the front rotor mounting frame and simultaneously increase the rotation speed of all the upper rotors and the lower rotors on the front rotor mounting frame when in a tilted forward flight state, and to balance the fuselage by adjusting the tilt angle of the rear rotor mounting frame and increasing the rotation speed of all the upper rotors and the lower rotors on the rear rotor mounting frame.

2. The foldable semi-tilting composite wing aircraft according to claim 1, characterized in that: The machine arm is configured as a two-section machine arm, and the machine arm folding mechanism includes a first folding member and a second folding member provided at adjacent ends of the two sections of the machine arm, wherein the first folding member is hinged to the second folding member; The end of the machine arm provided with the first folding component is sleeved with a sleeve lock buckle, and the end of the machine arm provided with the second folding component is fixedly connected with an external screw barrel, and the sleeve lock buckle is threadedly connected to the external screw barrel.

3. The foldable semi-tilting composite wing aircraft according to claim 1, characterized in that: The wing comprises a central wing and two side wings hinged at both ends of the central wing, and the central wing is fixed to the top of the fuselage; The wing folding mechanism comprises: a central connecting rod connected to the end of the central wing; Side connecting rods, fixedly connected to the ends of the side wings, and hingedly connected to the central connecting rod; The side fixing pins pass through the assembled center wing and side wings.

4. The foldable semi-tilting composite wing aircraft according to claim 3, characterized in that: The wing folding mechanism further comprises: A center positioning cylinder is fixedly connected to the middle of the center wing. The end of the center connecting rod is plugged and adapted to the end of the center positioning cylinder, and a first pin hole is formed through the corresponding ends of the two. A central fixing pin is plugged and connected to the top of the central wing and correspondingly plugged and connected to the first pin hole. Two central fixing pins are provided.

5. The foldable semi-tilting composite wing aircraft according to claim 1, characterized in that: The wing is a high-speed airfoil and is set at an installation angle of 6°. The wing is located between the front rotor mounting frame and the rear rotor mounting frame.

6. The foldable semi-tilting composite wing aircraft according to claim 1, characterized in that: A power box body with a tail opening is built into the rear of the fuselage, a battery pack is slidably arranged in the power box body, and an automatic lock for locking the battery pack is provided at the open end of the power box body; A hatch is provided at the rear of the fuselage corresponding to the position of the power box body, and a hatch cover and a linear actuator for driving the hatch cover to flip to close or open the hatch are hingedly connected to the hatch.

7. The foldable semi-tilting composite wing aircraft according to claim 1, characterized in that: A canopy is arranged on the upper side of the front section of the fuselage, and the propeller plane of the upper rotor is located below the canopy.

8. The foldable semi-tilting composite wing aircraft according to claim 1, characterized in that: The battery management system also includes a battery management system that adopts a multi-pack system BMS architecture, including: The first layer of master control is responsible for communicating with the flight control system, managing the multi-packet charging logic, coordinating the key information of the single packet and reporting it to the flight control system, and performing insulation testing after parallel high voltage connection; The second-layer master is responsible for single-packet energy management and fault diagnosis, reporting single-packet information to the first-layer master; and is responsible for single-packet thermal management and controlling the high-voltage power on and off during single-pack charging; The third-layer slave controller is responsible for collecting the cell voltage in a single pack, collecting the module temperature, and reporting the voltage / temperature data to the second-layer master controller.

9. A method for controlling a foldable semi-tilting composite wing aircraft, based on the foldable semi-tilting composite wing aircraft according to any one of claims 1 to 8, characterized in that: The following steps are involved: Determine the flight mode based on the aircraft's flight speed; Based on the flight mode, power demand and battery management system response, a closed loop is formed to dynamically adjust the tilt angle of the front and rear rotors and the propeller speed in different flight modes; When the aircraft is in forward flight, the tilt mechanism is used to adjust the front and rear rotors to tilt forward. The flight control system controls the front rotor speed to increase to increase forward thrust, and the rear rotor speed is synchronously adjusted to balance the fuselage pitch moment. The flight control system simultaneously sends instructions to the battery management system. The first-layer master control of the battery management system calculates the discharge rate corresponding to the target power of a single packet, and activates the high-power mode through the second-layer master control of the battery management system. The sampling period of the third-layer slave control of the battery management system is shortened to 50ms. When the aircraft transitions from a hovering state to a forward flight state, the tilt angle is first increased, and then the rotation speed is gradually increased.

10. The method for controlling a foldable semi-tilting composite wing aircraft according to claim 9, characterized in that: When the aircraft is in a hovering state or in the vertical takeoff and landing phase, the flight control system controls the tilt angle of all rotors to 0°, and the upper rotor and the lower rotor on the same arm rotate in opposite directions. At the same time, the flight control system sends instructions to the battery management system. The first-layer master control adjusts the multi-packet power distribution and evenly distributes it to four packets by receiving the "hover" instruction from the flight control system. The sampling period of the third-layer slave control is shortened to 100ms. When the aircraft is in the high-speed cruise stage, the wing generates starting lift due to the forward flight speed. The flight control system controls the tilt angle of all rotors to further increase, while reducing the speed of all propellers. The flight control system calculates the wing lift ratio in real time and dynamically distributes rotor power. Among them, if one battery pack fails, the remaining three battery packs each output the same power, and the first-layer master control forces the SOC of the remaining battery packs to maintain ≥23.4%; when the SOC of any battery pack is less than 20%, the first-layer master control sends a "return request" to the flight control system and limits the maximum power of a single pack to ≤30kW.

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