Pure electric aircraft and control method therefor

Abstract

A pure electric aircraft and a flight control method therefor. A pure electric aircraft (200) comprises an aircraft body (210), a frame (2210), a plurality of non-tilting rotors (220), and two ducted fans (230). The aircraft body (210) is provided with a manned flight cockpit (2110); the frame (2210) is arranged at the top of the aircraft body (210); the plurality of non-tilting rotors (220) are arranged on the frame (2210) at intervals around a vertical line where the center of gravity of the aircraft (200) is located, and when the aircraft body (210) is in a flight state, the rotation axis of each non-tilting rotor (220) is always in a fixed state relative to the position of the aircraft body (210); and the two ducted fans (230) are connected to the frame (2210) and are respectively located on the left side and the right side of the aircraft body (210), and when yawing moments provided by the non-tilting rotors (220) are in an unbalanced state, a compensation yawing moment is provided for the aircraft body (210). According to the pure electric aircraft (200), when the yawing moments formed by the plurality of non-tilting rotors (220) are in the unbalanced state, the compensation yawing moment can be provided for the aircraft body (210) by means of the two ducted fans (230), so that the flight safety is relatively high.

Description

Pure electric aircraft and control method thereof

[0001] This application claims priority to Chinese patent application No. 202311383016.0 filed on October 23, 2023, the entire contents of which are incorporated by reference into this application. Technical Field

[0002] The present application relates to the field of aircraft technology, and in particular to a pure electric aircraft and a control method thereof. Background Art

[0003] With the needs of people's lives and work, flying transportation equipment has been widely used, such as passenger planes, helicopters, and transport drones. Flying transportation equipment has very high requirements for its own flight power system. During the flight, all parts of the flight power system need to ensure normal operation to ensure the safety of passengers and flying transportation equipment.

[0004] However, when the flight power system of existing flying transportation equipment is underpowered or some equipment malfunctions, yaw moment imbalance may occur, resulting in the flying transportation equipment being unable to yaw. Existing flying transportation equipment cannot overcome the problem of being unable to yaw due to yaw moment imbalance. Technical issues

[0005] The main purpose of this application is to propose a pure electric aircraft and a control method thereof, aiming to overcome the problem of being unable to achieve yaw due to unbalanced yaw moment. Technical Solutions

[0006] To achieve the above-mentioned objectives, in a first aspect, the present application provides a pure electric aircraft, which includes an aircraft body, a frame, a plurality of non-tilt rotors and two ducted fans. The aircraft body is provided with a flight cabin for carrying drivers and passengers; the frame is arranged on the top of the aircraft body; the plurality of non-tilt rotors are connected to the frame, and the plurality of non-tilt rotors are arranged at intervals around a vertical line where the center of gravity of the aircraft is located. When the aircraft body is in flight, the position of the rotation axis of the non-tilt rotor relative to the aircraft body is always in a fixed state; the two ducted fans are connected to the frame and are respectively located on the left and right sides of the aircraft body, and provide a compensating yaw moment to the aircraft body when the yaw moment provided by the non-tilt rotor is in an unbalanced state.

[0007] In a second aspect, the present application also provides a flight control method for an aircraft, which is used to control the flight of an aircraft. The aircraft includes an aircraft body, multiple non-tilt rotors, and two ducted fans. The multiple non-tilt rotors are used to provide lift, thrust, and yaw moment to the aircraft body, and the two ducted fans are respectively located on the left and right sides of the aircraft body; the flight control method includes: obtaining a target yaw moment of the aircraft; during the flight of the aircraft, obtaining a first yaw moment generated by the multiple non-tilt rotors; when the difference between the first yaw moment and the target yaw moment is greater than a specified value, determining a compensatory yaw moment based on the difference; and controlling the operation of the ducted fan based on the compensatory yaw moment, wherein the difference between the sum of the second yaw moment generated by the operation of the ducted fan and the first yaw moment and the target yaw moment is less than or equal to the specified value. Beneficial effects

[0008] An embodiment of the present application provides a pure electric aircraft, which includes an aircraft body, a frame, multiple non-tilt rotors and two ducted fans. The aircraft body is provided with a flight cabin for carrying drivers and passengers; the frame is arranged on the top of the aircraft body; multiple non-tilt rotors are connected to the frame, and the multiple non-tilt rotors are arranged at intervals around a vertical line where the center of gravity of the aircraft is located. When the aircraft body is in flight, the position of the rotation axis of the non-tilt rotors relative to the aircraft body is always in a fixed state; the two ducted fans are connected to the frame and are respectively located on the left and right sides of the aircraft body. When the yaw moment provided by the non-tilt rotors is in an unbalanced state, the two ducted fans provide a compensating yaw moment to the aircraft body. In order to cope with the unbalanced yaw moment formed by multiple non-tilt rotors, such as yaw moment imbalance or failure of at least one non-tilt rotor among multiple non-tilt rotors, the aircraft in this embodiment also includes two ducted fans, which are connected to the aircraft body. When the yaw moment formed by multiple non-tilt rotors is in an unbalanced state, the ducted fans can operate and provide compensatory yaw torque to the aircraft body, so that the aircraft body can overcome the problem of being unable to yaw due to the imbalance of yaw moment. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] In order to more clearly illustrate the technical solution of the present application, the following is a brief introduction to the drawings required for use in the implementation. Obviously, the drawings described below are only some implementation methods of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0010] FIG1 is a schematic structural diagram of a flying car provided in an embodiment of the present application;

[0011] FIG2 is a schematic diagram of the structure of the flying car shown in FIG1 , wherein the vehicle and the aircraft are decoupled;

[0012] FIG3 is a schematic structural diagram of a vehicle provided in an embodiment of the present application;

[0013] FIG4 is a schematic structural diagram of an aircraft provided in an embodiment of the present application;

[0014] FIG5 is a schematic structural diagram of the automatic docking device of the aircraft shown in FIG4 ;

[0015] FIG6 is a schematic structural diagram of the aircraft shown in FIG4 with its arms folded and its landing gear mechanism deployed;

[0016] FIG7 is a schematic structural diagram of the flying car shown in FIG1 and the aircraft just after docking;

[0017] FIG8 is a schematic structural diagram of the aircraft shown in FIG7 moving toward the carrying platform;

[0018] FIG9 is a schematic structural diagram of the center of gravity of the aircraft shown in FIG7 being transferred to the carrying platform;

[0019] FIG10 is a schematic diagram of the structure of the flying car shown in FIG1 from above;

[0020] FIG11 is a schematic structural diagram of the rotor rotating surface formed by the foldable rotor of the aircraft shown in FIG10;

[0021] FIG12 is a schematic structural diagram of the positional relationship between the foldable rotor and the aircraft body of the aircraft shown in FIG10 ;

[0022] FIG13 is a schematic structural diagram of the aircraft shown in FIG10 when the arms are folded;

[0023] FIG14 is a schematic structural diagram of the aircraft shown in FIG10 when the arms are deployed;

[0024] FIG15 is a schematic structural diagram of the foldable rotor of the aircraft shown in FIG10 in an expanded state;

[0025] FIG16 is a schematic structural diagram of the foldable rotor of the aircraft shown in FIG10 in a folded state;

[0026] FIG17 is a schematic structural diagram of the foldable rotor of the aircraft shown in FIG10 in an expanded state;

[0027] FIG18 is a schematic flow chart of a flight control method for an aircraft provided in one embodiment of the present application;

[0028] FIG19 is a schematic diagram of the functional framework of a flight control device disclosed in an embodiment of the present application.

[0029] Implementation Methods of the Application

[0030] In order to enable those skilled in the art to better understand the present invention, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0031] Referring to Figures 1 and 2 , this embodiment provides a flying car 1000. Flying car 1000 comprises a land-based transportation device such as a car, a flying transportation device such as an airplane, or a transportation device such as a flying car that has both land and flying capabilities. This embodiment of the present application provides a flying car 1000 with both land and flying capabilities. Flying car 1000 comprises a vehicle 100 and an aircraft 200, which are detachably coupled.

[0032] Referring to Figure 3, the vehicle 100 in this embodiment of the present application is used for ground travel, enabling the flying car 1000 to operate on land. The vehicle 100 comprises a vehicle body 110, a plurality of wheels 120, and an automatic docking device 130. The vehicle body 110 is the main structure of the vehicle 100, carrying the vehicle 100 and other functional equipment. The vehicle body 110 comprises a land-based cabin 1110 and a supporting platform 1130, which are connected to each other. The land-based cabin 1110 is equipped with seats for passengers, and the supporting platform 1130 is used to dock the aircraft 200. The plurality of wheels 120 are mounted on the chassis of the vehicle body 110, enabling the vehicle body 110 to operate on land. The automatic docking device 130 is mounted on the supporting platform 1130 of the vehicle body 110 and is used to dock with and tow the aircraft 200.

[0033] Referring to Figure 4 , the aircraft 200 in the embodiment of the present application is used to travel in the air, enabling the flying car 1000 to achieve the function of traveling in the air. The aircraft 200 and the vehicle 100 together constitute the flying car 1000, which has both land-based and flight functions. The aircraft 200 includes an aircraft body 210 and a plurality of foldable rotors 220 disposed on the aircraft body 210. The aircraft body 210 is the main structure of the aircraft 200, and a flight cabin 2110 is provided on the aircraft body 210 for passengers to ride. The plurality of foldable rotors 220 are disposed on the aircraft body 210 to provide propulsion for the aircraft 200 to travel in the air. Driven by the plurality of foldable rotors 220, the aircraft 200 can rise into the air and travel. The aircraft 200 may further include two ducted fans 230 , which are connected to the aircraft body 210 . The two ducted fans 230 are used to provide compensating yaw moments to the aircraft body 210 when the yaw moments formed by the multiple foldable rotors 220 are in an unbalanced state.

[0034] Traditional flying cars 1000 require the aircraft 200 to land vertically directly on the vehicle 100's support platform 1130 during landing. Firstly, the large size and weight of a passenger-carrying aircraft 200 make it difficult to precisely control its posture and motion. Secondly, the overall size of the flying car 1000 is limited by existing road infrastructure, so the landing surface provided by the vehicle 100's support platform 1130 is relatively small, making it difficult for the aircraft 200 to land on the support platform 1130 quickly. Furthermore, the aircraft 200's rotors are in operation, making it susceptible to collision with the vehicle 100, potentially damaging both the aircraft 200 and the vehicle 100. Even if the aircraft 200 were to land on the support platform 1130, its position would not meet coupling requirements. After the aircraft 200 lands on the supporting platform 1130, adjusting the relative position of the aircraft 200 and the vehicle 100 not only requires overcoming the weight of the aircraft 200, but also requires considering whether the structure of the supporting platform 1130 can withstand the repeated deflection and movement of the aircraft 200 on the supporting platform 1130, and avoiding structural collision between the aircraft 200 and the vehicle 100. Therefore, it is extremely difficult to adjust the position of the aircraft 200 on the supporting platform 1130, and the user experience is relatively low.

[0035] The embodiment of the present application provides a new coupling method for the aircraft 200 and the vehicle 100 by providing an automatic docking device 130. Specifically, the aircraft 200 switches from a flying state to a ground-based state, and then the vehicle 100 drives near the aircraft 200 and controls the position of the vehicle 100 and the aircraft 200 to align so as to drive the automatic docking device 130 to automatically connect to the aircraft body 210, and drive the aircraft body 210 to move toward the vehicle body 110 until the aircraft body 210 is detachably coupled to the supporting platform 1130. When the aircraft 200 needs to take off, the aircraft 200 first releases the coupling relationship between the aircraft body 210 and the supporting platform 1130 before taking off. After the coupling relationship is released, the aircraft 200 first docks on the ground, and then starts to take off from the ground and rise into the air, while the vehicle 100 stays on the land. Therefore, the aircraft 200 in the embodiment of the present application does not need to be precisely positioned with the carrier platform 1130 and can quickly land on the ground. The vehicle 100 then moves to a position near the aircraft 200 and aligns with the aircraft, allowing the automatic docking device 130 to dock with the aircraft 200 and tow the aircraft 200 to the carrier platform 1130. The control difficulty of coupling the aircraft body 210 and the vehicle 100 is relatively low, which increases the speed of coupling the aircraft body 210 and the carrier platform 1130 and facilitates adjustment of the relative position of the aircraft 200 and the vehicle 100. During the process of coupling and decoupling the aircraft 200 with the carrier platform 1130, the foldable rotors 220 stop rotating, improving the safety of the flying car 1000 during use.

[0036] In this embodiment, the extent to which the carrying platform 1130 covers the aircraft 200 is not limited; the aircraft 200 may be partially or fully enclosed. In this embodiment, the vehicle 100 capable of partially enclosing the aircraft 200 has an appearance similar to that of an existing pickup truck. Specifically, the carrying platform 1130 further includes a carrying plate 1131 and a fence structure 1132. The carrying plate 1131 is connected to the land cabin 1110. The fence structure 1132 is disposed around the periphery of the carrying plate 1131. When the aircraft 200 is coupled to the carrying platform 1130, it can prevent the aircraft 200 from horizontally leaving the carrying plate 1131. The fence structure 1132, the land cabin 1110, and the carrying plate 1131 collectively define a first accommodating space 1133 that communicates with the exterior. In order to enable the aircraft 200 to enter and exit the first accommodating space 1133, the fence structure 1132 includes a movable fence (not marked in the figure), which is arranged on the side of the supporting plate 1131 away from the land cabin 1110. The movable fence can rotate relative to the supporting plate 1131 to provide a channel for the aircraft body 210 to enter the supporting platform 1130.

[0037] In this embodiment, the vehicle 100, which fully encloses the aircraft 200, has an appearance similar to that of an existing box truck. Specifically, the carrying platform includes a carrying plate 1131 and a box body (not shown). The carrying plate 1131 is connected to the land-based cabin 1110, and the box body cover is mounted on the carrying plate 1131. The box body and the carrying plate 1131 together define a second accommodation space (not shown) isolated from the outside. To allow the aircraft 200 to enter and exit the second accommodation space, the box body includes a door, which is movably disposed on the side of the box body facing away from the land-based cabin 1110. The door is used to open the box body to clear the aircraft body 210.

[0038] In this embodiment, vehicle 100 further includes a vehicle control system 140 for controlling wheels 120. Wheels 120 can change the motion state of vehicle body 110, such as acceleration, deceleration, forward movement, or reverse movement, in response to vehicle motion commands from vehicle control system 140. Similarly, aircraft 200 further includes a flight control system 240 for controlling foldable rotors 220. Foldable rotors 220 can change the motion state of aircraft body 210, such as takeoff, landing, yaw, or hovering, in response to aircraft motion commands from flight control system 240. Furthermore, aircraft 200 in this embodiment can travel on land, for example, to facilitate running, cushioning, and the like.

[0039] Referring to Figures 3 and 4 , in this embodiment, the vehicle control system 140 is disposed within the land cockpit 1110. The land cockpit 1110 can be understood as both the passenger and driving spaces. The driver can manipulate the vehicle control system 140 within the land cockpit 1110 to operate the vehicle 100. The vehicle control system 140 can be one or more of a driving instrument panel, a steering wheel, an accelerator, a brake, a gear lever, and the like. The flight control system 240 is disposed within the flight cockpit 2110 and includes one or more of a flight instrument panel, a center console, and an aircraft joystick. In some embodiments, the vehicle control system 140 and the flight control system 240 can be communicatively coupled to each other to transmit motion data or motion status information of the vehicle 100 and the aircraft 200. In other embodiments, vehicle 100 is an autonomous transportation device, and therefore the location of vehicle control system 140 is not specifically limited; vehicle control system 140 may be located in a location on vehicle body 110 other than cockpit 1110, or may be located within cockpit 1110 to assist passengers in driving vehicle 100. If aircraft 200 is also an autonomous transportation device, aircraft 200 is configured similarly to vehicle 100 and will not be further described here.

[0040] Next, the specific structure of the vehicle 100 and the aircraft 200 for docking coupling will be introduced, and how the vehicle 100 and the aircraft 200 achieve docking coupling through the specific structure will be introduced.

[0041] Referring to FIG. 5 , the automatic docking device 130 in this embodiment includes a traction mechanism 1310 and a linear motion mechanism 1320. The traction mechanism 1310 is movably mounted on the carrier platform 1130. When the vehicle 100 controls the carrier platform 1130 to align with the aircraft body 210, the traction mechanism 1310 tends to move toward or away from the aircraft body 210. The linear motion mechanism 1320 is the structural basis for the traction mechanism 1310 to move in a fixed direction. The linear motion mechanism 1320 is mounted on the carrier platform 1130, and the traction mechanism 1310 is connected to the linear motion mechanism 1320. The linear motion mechanism 1320 is used to drive the traction mechanism 1310 to move on the carrier platform. Correspondingly, the aircraft 200 also includes a docking mechanism 250 (see FIG. 2 ) connected to the aircraft body 210. The docking mechanism 250 is adapted to cooperate with the traction mechanism 1310. After the traction mechanism 1310 is connected with the docking mechanism 250 , the traction mechanism 1310 pulls the docking mechanism 250 to pull the aircraft 200 onto the carrying platform 1130 .

[0042] Please refer to Figure 5. First, the specific components of the linear motion mechanism 1320 will be introduced to reveal the mechanism by which the linear motion mechanism 1320 drives the traction mechanism 1310 to move. Specifically, the linear motion mechanism 1320 includes a screw 1321, a rotating motor 1322, and a slider 1323. The screw 1321 is arranged along the forward direction of the vehicle 100. The rotating motor 1322 is driven and connected to one end of the screw 1321. When the rotating motor 1322 is turned on, it can drive the screw 1321 to rotate. The slider 1323 is threadedly connected to the screw 1321 and is connected to the traction mechanism 1310. When the supporting platform 1130 moves to a position relative to the aircraft body 210, the screw 1321 rotates, and the slider 1323, together with the traction mechanism 1310, moves closer to or away from the aircraft body 210 relative to the supporting platform 1130. In this embodiment, there are two linear motion mechanisms 1320. The two linear motion mechanisms 1320 are arranged horizontally with the traction mechanism 1310 to avoid increasing the height of the supporting platform 1130 and preventing the overall structure of the flying car 1000 from being too high and unsuitable for existing transportation facilities and building structures (the dimensions of the overall structure of the flying car 1000 will be specifically defined later). The two linear motion mechanisms 1320 are arranged on opposite sides of the traction mechanism. On the one hand, this increases the driving force acting on the traction mechanism 1310. On the other hand, it can keep the traction mechanism 1310 in a roughly balanced state, making it less likely to tilt or deflect, thereby ensuring the traction effect.

[0043] Referring to Figure 5 , the following describes the specific structures of the traction mechanism 1310 and the docking mechanism 250 of the aircraft 200. Specifically, the traction mechanism 1310 includes a traction body 1311 and a traction ring 1312. Correspondingly, the docking mechanism 250 of the aircraft 200 includes a mating ring (not shown). The traction ring 1312 is disposed on the side of the traction body 1311 facing away from the landing cabin 1110. When the support platform 1130 moves to a position opposite to the aircraft body 210, the traction ring 1312 and the mating ring are positioned opposite and engage with each other. In other embodiments, the traction mechanism 1310 and the docking mechanism 250 may also be hook-shaped structures.

[0044] In order to enable the traction ring body 1312 to be interlocked with the matching ring body, in this embodiment, the traction ring body 1312 is partially exposed on the side of the traction body 1311 away from the land cabin 1110, and the traction ring body 1312 is provided with a notch. The traction ring body 1312 is rotatably set on the traction body 1311, and the traction ring body 1312 can rotate around its axis. When the traction ring body 1312 and the matching ring body are relatively set, the traction ring body 1312 can rotate and make the notch correspond to the matching ring body, so that the matching ring body and the traction ring body 1312 can be interlocked, and then the traction ring body 1312 continues to rotate to hide the notch inside the traction body 1311, thereby ensuring the stability of the interlocking relationship between the matching ring body and the traction ring body 1312.

[0045] Referring to Figure 5 , after the aircraft 200 is towed onto the supporting platform 1130 and before the aircraft body 210 is coupled to the supporting platform 1130, the direction of movement of the aircraft body 200 is maintained as straight as possible relative to the vehicle body 110. Therefore, the automatic docking device 130 in this embodiment also includes two guide slots 1330. Both guide slots 1330 are horizontally disposed on the supporting platform 1130, extending in the forward direction of the vehicle 100. Correspondingly, mating wheel assemblies (not shown) are provided on opposite sides of the bottom of the aircraft body 210. When the supporting platform 1130 is moved to a position opposite the aircraft body 210, the two guide slots 1330 and the two mating wheel assemblies are aligned one by one. Each mating wheel assembly can slide into and roll within a corresponding guide slot 1330, thereby limiting the trajectory of the aircraft body 210.

[0046] Please refer to Figure 5. In order to facilitate the mating wheel set to slide into the corresponding guide groove 1330, the guide groove 1330 in this embodiment includes a first groove body 1331 and a second groove body 1332 that are connected. The second groove body 1332 is connected to the end of the first groove body 1331 that is away from the land cabin 1110. The groove width of the second groove body 1332 gradually increases in the direction away from the land cabin 1110, that is, the groove width of the second groove body 1332 is wider at the end close to the mating wheel set, and the mating wheel set can slide smoothly into the second groove body 1332.

[0047] After the mating wheel set slides into the guide groove 1330, the traction mechanism 1310 pulls the aircraft body 210 to move in a straight line. The movement direction of the mating wheel set is not necessarily consistent with the extension direction of the guide groove 1330, which will cause the mating wheel set to be stuck in the guide groove 1330; therefore, the mating ring body in this embodiment is rotatably connected to the side of the chassis of the aircraft body 210 facing the ground, and the mating ring body can rotate linearly in the vertical direction relative to the chassis of the aircraft body 210. In the process of towing the aircraft body 210, the traction ring body 1312 adjusts the movement direction of the mating wheel set through the mating ring body, so that the mating wheel set and the guide groove 1330 cooperate smoothly.

[0048] Referring to FIG3 , the automatic docking device 130 in this embodiment further includes a limiting mechanism 1340. The application scenario of the limiting mechanism 1340 is as follows: when the aircraft body 210 is coupled to the carrying platform 1130, the limiting mechanism 1340 is connected between the aircraft body 210 and the carrying platform 1130 to limit the movement of the aircraft body 210 in any direction relative to the carrying platform 1130. Specifically, the limiting mechanism 1340 in this embodiment includes a horizontal limiting assembly (not shown in the figure), which can be connected between the carrying platform 1130 and the aircraft body 210 or between the vehicle body 110 and the aircraft body 210. The horizontal limiting assembly is used to limit the movement of the aircraft body 210 in the horizontal direction relative to the carrying platform 1130, for example, to limit the aircraft body 210 from swaying back and forth or swinging left and right in the forward direction of the vehicle body 110. The limiting mechanism 1340 in this embodiment also includes a vertical limiting assembly (not shown), which is connected between the support platform 1130 and the aircraft body 210 and is used to limit the vertical movement of the aircraft body 210 relative to the support platform 1130. The provision of the limiting mechanism 1340 stabilizes the position of the aircraft body 210 on the support platform 1130 and ensures the stability of the fixed connection between the aircraft body 210 and the support platform 1130. The limiting mechanism 1340 limits the movement of the aircraft body 210 on the support platform 1130, thereby reducing the impact damage caused by the movement of the aircraft body 210 to the aircraft 200 itself and the vehicle 100.

[0049] This embodiment does not limit the type and specific structure of the horizontal and vertical limiter assemblies; they may be connected by a hook and loop or a stopper that limits movement. In this embodiment, an electrical limiter mechanism 1340 may be used. When the aircraft body 210 and the carrying platform 1130 are coupled, the electrical limiter mechanism connects to the aircraft body 210 to restrict its movement. Before the aircraft body 210 is ready for takeoff, the electrical limiter mechanism disconnects the carrying platform 1130 or the vehicle body 110 from the aircraft body 210.

[0050] Please refer to Figure 3. In order to reduce the damage caused by the vibration of the aircraft body 210 to the aircraft 200 itself and the vehicle 100, the automatic docking device 130 in this embodiment also includes a buffer mechanism 1350. The buffer mechanism 1350 is arranged on the supporting platform 1130. When the aircraft body 210 and the supporting platform 1130 are coupled, the buffer mechanism 1350 is located between the aircraft body 210 and the supporting platform 1130, and between the aircraft body 210 and the vehicle body 110. The buffer mechanism 1350 can reduce the impact damage caused by the shaking of the aircraft body 210 to the vehicle body 110 and the aircraft body 210. This embodiment does not specifically limit the type and specific structure of the buffer mechanism 1350. The buffer mechanism 1350 can be a damper or a hydraulic buffer.

[0051] Referring to Figures 4 and 6 , the aircraft 200 in this embodiment also includes a landing gear mechanism 260, which is used to support the aircraft body 210 when the aircraft 200 is in a landing position. The landing gear mechanism 260 is rotatably connected to the aircraft body 210. When the aircraft 200 is in flight, the landing gear mechanism 260 rotates toward the aircraft body 210 to assume a folded position, thereby reducing the resistance experienced by the aircraft body 210 in mid-air. When the aircraft 200 is about to land, the landing gear mechanism 260 can rotate relative to the aircraft body 210 to assume a supported position and support the ground. In other embodiments, structures such as airbags or piers can also be used to achieve landing of the aircraft body 210.

[0052] Referring to FIG7 , in order to ensure that the left and right sides of the aircraft body 210 along the forward direction of the aircraft 200 are relatively balanced when the aircraft body 210 is in a parked state, the landing gear mechanism 260 in this embodiment includes two landing gears 2610. The two landing gears 2610 are rotatably connected to the left and right sides of the aircraft body 210, respectively. Each landing gear 2610 can be rotated relative to the aircraft body 210 to be deployed or folded. In other embodiments, the landing gear mechanism 260 can be a unitary structure. Such a unitary landing gear mechanism 260 can simultaneously support the left and right sides of the aircraft body 210 when the aircraft 200 is in a parked state.

[0053] Referring to Figure 4 , to ensure that the landing gear 2610 can automatically deploy or fold, the landing gear mechanism in this embodiment further includes at least two drive members 2630. The two drive members 2630 are provided in a one-to-one correspondence with the two landing gears 2610, and the drive members 2630 connect the aircraft body 210 and a corresponding landing gear 2610. The drive members 2630 can automatically drive the landing gear 2610 to rotate relative to the aircraft body 210 based on the driving state of the aircraft 200, thereby folding or deploying the landing gear 2610 relative to the aircraft body 210 without requiring manual intervention, thereby improving safety. The drive members 2630 in this embodiment can be hydraulic cylinders, linear motors, or other devices.

[0054] Referring to FIG4 , the landing gear 2610 in this embodiment includes a support rod 2611 and a support leg 2612. The support rod 2611 is connected between the support leg 2612 and the aircraft body 210. When the landing gear mechanism 260 is in the supported state, the support leg 2612 abuts the ground and supports the aircraft body 210 upward. The support leg 2612 is a rod structure and extends generally along the forward direction of the aircraft 200. At least one of the front and rear ends of the support leg 2612 in the forward direction of the aircraft 200 is tilted away from the ground. The tilted portion has an arcuate edge 2613 formed on the side of the support leg 2612 facing the ground. When the aircraft 200 is landing on the ground, the arcuate edge 2613 allows the support leg 2612 to adapt to the ground and prevent foreign objects on the ground from colliding violently with the support leg 2612, thereby preventing abnormal vibration of the aircraft body 210 and preventing damage to the support leg 2612 and the aircraft body 210.

[0055] After the traction mechanism 1310 on the carrying platform 1130 and the docking mechanism 250 on the aircraft body 210 dock, the traction mechanism 1310 pulls the aircraft body 210 toward the carrying platform 1130 until the aircraft body 210 is coupled to the carrying platform 1130. Therefore, the process of the aircraft body 210 moving from being independent of the carrying platform 1130 to being coupled to the carrying platform 1130 is a gradual process. Before the aircraft body 210 has mostly moved to the carrying platform 1130, or before the center of gravity of the aircraft body 210 has not moved to a safe position on the carrying platform 1130, the landing gear mechanism 260 still needs to be installed between the aircraft body 210 and the ground to maintain the balance of the aircraft body 210 and the vehicle 100. In this embodiment, the aircraft 200 includes at least two sets of landing gear mechanisms 260, which are arranged on the aircraft body 210 at intervals along the forward direction of the aircraft 200, with one set of landing gear mechanisms 260 located near the carrier platform 1130 and the other located away from the carrier platform 1130. In other words, one set of landing gear mechanisms 260 is located at the front end of the aircraft body 210, and the other set is located at the rear end of the aircraft body 210. Each set of landing gear mechanisms 260 may include two landing gears 2610, with the two landing gears 2610 in the set being located on the left and right sides of the aircraft body 210, respectively. It should be understood that in other embodiments, more landing gear mechanisms 260 may be provided based on the weight of the aircraft body 210 and the length of the aircraft.

[0056] In this embodiment, the two landing gear mechanisms 260 are folded one after another during the process of the aircraft body 210 being independent of the carrying platform 1130 and then being coupled to the carrying platform 1130. The folding time nodes of the two landing gear mechanisms 260 will be described in detail below.

[0057] Referring to Figures 7 and 8 , both landing gear mechanisms 260 are initially in a supported state. As the aircraft body 210 is dragged by the traction mechanism 1310, a smaller portion of the aircraft body 210 moves onto the support platform 1130 and is supported by the support platform 1130. When the support platform 1130 and the landing gear mechanism 260 farther from the support platform 1130 are able to overcome the weight of the aircraft body 210 and maintain equilibrium, the landing gear mechanism 260 closer to the support platform 1130 folds and retracts to switch to the folded state.

[0058] Please refer to Figure 9. The traction mechanism 1310 continues to pull the aircraft body 210 toward the carrying platform 1130. When most of the structure of the aircraft body 210 moves to the carrying platform 1130 and the center of gravity of the aircraft body 210 and the carrying platform 1130 and the vehicle body 110 reach a balanced state, that is, the carrying platform 1130 can independently support the aircraft body 210 without the assistance of a landing gear mechanism 260 away from the carrying platform 1130, a landing gear mechanism 260 away from the carrying platform 1130 folds and contracts to switch to a folded state.

[0059] Finally, the traction mechanism 1310 continues to traction the aircraft body 210 to move toward the carrying platform 1130 until the aircraft body 210 moves to a designated coupling position on the carrying platform 1130 , and then the aircraft body 210 is coupled to the carrying platform 1130 .

[0060] Referring again to FIG. 7 , the vehicle 100 in this embodiment is a pure electric vehicle. The vehicle 100 further includes a vehicle electrical system 170 disposed within the vehicle body 110 and electrically connected to the plurality of wheels 120 to provide electrical energy to the wheels 120, enabling the wheels 120 to propel the vehicle body 110 on land. In other embodiments, the vehicle 100 may also be an extended-range electric vehicle, meaning that in addition to the vehicle electrical system 170, the vehicle 100 also includes a fuel generator. Similarly, the aircraft 200 in this embodiment is a pure electric aircraft or a pure electric manned helicopter. The aircraft 200 further includes a flight electrical system 270 disposed within the aircraft body 210 and electrically connected to the plurality of foldable rotors 220 to provide electrical energy to the plurality of foldable rotors 220, enabling the foldable rotors 220 to propel the aircraft body 210 through the air.

[0061] Specifically, in this embodiment, the vehicle electrical system 170 includes a first power battery pack 1710, which is disposed in the vehicle body 110 and is used to store electrical energy to supply power to the land propulsion system and other electrical components of the vehicle 100. The flight electrical system 270 includes a second power battery pack 2710, which is disposed in the aircraft body 210 and is used to store electrical energy to supply power to the foldable rotors 220 and other electrical components of the aircraft 200. This embodiment does not limit the types of the first power battery pack 1710 and the second power battery pack 2710; they can be acid lithium batteries or alkaline zinc-manganese batteries.

[0062] In some embodiments, the vehicle electrical system 170 and the flight electrical system 270 are independent of each other, and no electrical energy is transferred between the first power battery pack 1710 and the second power battery pack 2710. In this embodiment, the first power battery pack 1710 and the second power battery pack 2710 are structurally independent of each other, but electrical energy can be transferred between the first power battery pack 1710 and the second power battery pack 2710 when the aircraft body 210 is coupled to the carrier platform 1130. Specifically, the vehicle electrical system 170 further includes a first connector 1720 electrically connected to the first power battery pack 1710, and the flight electrical system 270 further includes a second connector 2720 electrically connected to the second power battery pack 2710. When the aircraft body 210 and the carrying platform 1130 are in a coupled state, the first connector 1720 and the second connector 2720 are electrically connected to allow electrical energy to be transmitted between the first power battery pack 1710 and the second power battery pack 2710, so that the first power battery pack 1710 can be supplemented by the electrical energy of the second power battery pack 2710, or the second power battery pack 2710 can be supplemented by the electrical energy of the first power battery pack 1710. In addition, in this embodiment, the first power battery pack 1710 and the second power battery pack 2710 are both provided with power connection ports to receive electrical energy supplementation from a power source external to the transport vehicle.

[0063] First, the situation where the first power battery pack 1710 replenishes power from the second power battery pack 2710 will be described. After the aircraft 200 completes its flight and is coupled to the carrying platform 1130, the first connector 1720 connected to the first power battery pack 1710 can replenish power from the second power battery pack 2710 via the second connector 2720, ensuring that the flight electrical system 270 stores sufficient power for the next flight of the aircraft 200. Specifically, the vehicle electrical system 170 also includes an energy management module 1730, which is electrically connected to the first power battery pack 1710. Upon detecting that the aircraft body 210 and the carrying platform 1130 are coupled, the energy management module 1730 can control the first power battery pack 1710 to replenish power from the second power battery pack 2710.

[0064] Next, we will describe how the second power battery pack 2710 replenishes power from the first power battery pack 1710. The flying car 1000 in this embodiment also includes a power sensor (not shown) and a vehicle controller (not shown). The power sensor is electrically connected to the vehicle controller. The power sensor detects the remaining power in the first power battery pack 1710 and the second power battery pack 2710, and the vehicle controller controls whether the second power battery pack 2710 transfers power to the first power battery pack 1710. When the vehicle 100 and the aircraft 200 are coupled and the vehicle 100 is driving, if the power sensor detects a low remaining power in the first power battery pack 1710, the vehicle controller controls the second power battery pack 2710 to transfer power to the first power battery pack 1710. This embodiment specifically describes two specific manifestations of insufficient power in the first power battery pack 1710.

[0065] In the first specific manifestation, the power sensor detects the remaining power of the first power battery pack 1710, and the vehicle controller estimates the cruising range of the vehicle 100 based on the remaining power of the first power battery pack 1710. If the cruising range of the vehicle 100 is lower than the target cruising range, it means that the power in the first power battery pack 1710 is insufficient to support the flying car 1000 to travel to the target location. In this case, it can be determined that the remaining power in the first power battery pack 1710 is insufficient. The power sensor then detects the remaining power in the second power battery pack 2710. If the remaining power in the second power battery pack 2710 is greater than or equal to the second preset power, that is, if the power in the second power battery pack 2710 is sufficient, the vehicle controller controls the second power battery pack 2710 to replenish power to the first power battery pack 1710.

[0066] In a second specific embodiment, the flying car 1000 in this embodiment further includes an acceleration sensor (not shown). The acceleration sensor is used to detect the actual acceleration of the vehicle 100. When the actual acceleration of the vehicle 100 is less than or equal to a preset acceleration, it indicates that the remaining power in the first power battery pack 1710 is insufficient, and there is insufficient energy for the vehicle 100 to reach the preset speed within the preset time. The power sensor then detects the remaining power in the second power battery pack 2710. If the remaining power in the second power battery pack 2710 is greater than or equal to a second preset power level, i.e., if the second power battery pack 2710 has sufficient energy, the vehicle controller controls the second power battery pack 2710 to replenish energy from the first power battery pack 1710.

[0067] In both of the above scenarios where the second power battery pack 2710 replenishes power to the first power battery pack 1710, sufficient power is required in the second power battery pack 2710, i.e., the remaining power in the second power battery pack 2710 must be greater than or equal to the second preset power level. In special circumstances, even if the remaining power in the second power battery pack 2710 is less than the second preset power level, the second power battery pack 2710 can still replenish power to the first power battery pack 1710. However, when the power in the second power battery pack 2710 drops to the preset minimum power level, power transmission to the first power battery pack 1710 stops, and the first power battery pack 1710 needs to seek other power supply equipment for replenishment.

[0068] Referring to Figure 7 , this embodiment does not limit the positions of the first connector 1720 and the second connector 2720; their specific positions are determined based on the actual orientation of the vehicle body 110 and the aircraft body 210. As previously mentioned, in this embodiment, the vehicle body 110 and the carrying platform 1130 are arranged side by side in the forward direction of the vehicle 100. When the aircraft body 210 and the carrying platform 1130 are coupled, the aircraft body 210 and the vehicle body 110 are also arranged side by side in the forward direction of the vehicle 100. Furthermore, in this embodiment, the front of the aircraft body 210 faces opposite directions from the front of the vehicle body 110 and is arranged substantially parallel to the front of the vehicle body 110. Therefore, in this embodiment, the first connector 1720 is positioned on the side of the vehicle body 110 facing the aircraft body 210, and the second connector 2720 is correspondingly positioned on the side of the aircraft body 210 facing the vehicle body 110. This allows the first connector 1720 and the second connector 2720 to be positioned opposite each other and easily connected when the aircraft body 210 and the carrying platform 1130 are coupled.

[0069] Specifically, the vehicle body 110 in this embodiment has a first backplate 1120 facing the vehicle body 210. That is, the first backplate 1120 is disposed toward the carrying platform 1130. A first connector 1720 is disposed on the first backplate 1120 to accommodate the carrying platform 1130 and the vehicle body 210. Correspondingly, the vehicle body 210 has a second backplate 2120 facing the vehicle body 110. A second connector 2720 is disposed on the second backplate 2120 to accommodate the vehicle body 110. When the vehicle body 210 and the carrying platform 1130 are coupled, the first backplate 1120 and the second backplate 2120 are disposed opposite each other, and the first connector 1720 and the second connector 2720 are also disposed opposite each other and electrically connected, enabling electrical energy transmission between the first power battery pack 1710 and the second power battery pack 2710.

[0070] This embodiment does not specifically limit the location of the first power battery pack 1710 and the second power battery pack 2710. In this embodiment, the first power battery pack 1710 is fixedly mounted on the chassis of the vehicle body 110 and located within the vehicle body 110. This ensures that the heavier first power battery pack 1710 remains stable, ensuring power supply to the vehicle 100 and the structural stability of the vehicle body 110. A first high-voltage wire harness 1740 is disposed between the first power battery pack 1710 and the first connector 1720. The first high-voltage wire harness 1740 extends along the chassis of the vehicle body 110 and the structure of the first backplate 1120 and is electrically connected between the first power battery pack 1710 and the first connector 1720 to facilitate power transmission. Similarly, the second power battery pack 2710 is fixedly mounted on the chassis of the aircraft body 210 and is located inside the aircraft body 210. A second high-voltage wire harness 2730 is disposed between the second power battery pack 2710 and the second connector 2720. The second high-voltage wire harness 2730 extends along the structure of the chassis of the aircraft body 210 and the second backplane 2120 and is electrically connected between the second power battery pack 2710 and the second connector 2720. In other embodiments, the first power battery pack 1710 may be disposed in the ground cabin 1110, and the second power battery pack 2710 may be disposed in the flight cabin 2110.

[0071] Referring again to Figure 3 , the multiple wheel assemblies 120 in this embodiment include a first wheel assembly 1210, a second wheel assembly 1220, and a third wheel assembly 1230, as well as a vehicle drive motor. The first wheel assembly 1210, the second wheel assembly 1220, and the third wheel assembly 1230 are all disposed on the ground-facing side of the vehicle body 110, and at least one of the first wheel assembly 1210, the second wheel assembly 1220, and the third wheel assembly 1230 is driven by the vehicle drive motor. In some embodiments, the number of wheel assemblies 120 may be four, five, or six, etc., depending on the overall weight of the flying car 1000 and the load-bearing capacity of the wheel assemblies 120. In this embodiment of the present application, each wheel assembly 120 refers to a wheel assembly that shares a common wheel axle. Each wheel assembly 120 may include multiple wheels, such as two wheels, four wheels, six wheels, or eight wheels, etc. The multiple wheels in each wheel assembly 120 are symmetrically disposed at opposite ends of the common wheel axle.

[0072] In this embodiment, the positions of the various wheel assemblies 120 are determined based on the weight borne by various parts of the vehicle 100 when the aircraft 200 and the vehicle 100 are coupled. Specifically, the first wheel assembly 1210 is positioned corresponding to the land cabin 1110 of the vehicle body 110 to support the land cabin 1110. The second wheel assembly 1220 and the third wheel assembly 1230 are positioned corresponding to the support platform 1130 to support the support platform 1130. In other embodiments, the first wheel assembly 1210, the second wheel assembly 1220, and the third wheel assembly 1230 may be equidistantly spaced on the bottom of the vehicle body 110.

[0073] Referring again to FIG. 4 , the aircraft 200 in this embodiment further includes a frame 2210 . The frame 2210 is disposed on the aircraft body 210 and serves as a connecting medium between the aircraft body 210 and the plurality of foldable rotors 220 . The plurality of foldable rotors 220 are disposed on the frame 2210 and are used to provide lift, thrust, and yaw torque to the aircraft body 210 , enabling the aircraft body 210 to perform maneuvers such as takeoff, landing, forward movement, and yaw. In this embodiment, the plurality of foldable rotors 220 are selected to have the same output power. Multiple foldable rotors 220 with the same rated output power place relatively low demands on the overall structure of the aircraft 200 , thereby reducing the complexity of the overall structure of the aircraft 200 and lowering the development cost of the aircraft 200 . In some other embodiments, the multiple foldable rotors 220 can be a combination of foldable rotors 220 with multiple rated output powers. The advantage is that a more functional flight mode can be developed. For example, in this type of flight mode, the aircraft 200 can fly in harsh environments and climates, thereby improving the adaptability of the aircraft 200 to the flight environment and making the aircraft 200 more competitive in the market.

[0074] In this embodiment, foldable rotor 220 maintains a stable position relative to the structure of aircraft body 210, simplifying the installation structure of foldable rotor 220. The fixed position of foldable rotor 220 is specifically manifested as follows: foldable rotor 220 has a rotation axis, which serves as the rotation axis of the rotor of foldable rotor 220. During flight, the rotation axis of foldable rotor 220 remains fixed relative to aircraft body 210. Therefore, in this embodiment, foldable rotor 220 can also be referred to as a non-tilt rotor.

[0075] When the aircraft 200 is in level flight, the range of the angle between the rotation axis of the foldable rotor 220 and the vertical plane in this embodiment is approximately less than or equal to 10°. The settings of multiple foldable rotors 220 are not the same. For ease of understanding, the range setting of the angle of the rotation axis of the foldable rotor will be described in detail in some specific usage environments later, and no further explanation will be given here.

[0076] In order to improve the ability of the aircraft 200 to cope with emergencies while flying in the air and to improve the safety of the aircraft 200 during flight, this embodiment makes specific arrangements in terms of the number of foldable rotors 220, the distribution of the foldable rotors 220 on the aircraft body 210, the structure of the foldable rotors 220, and the compensation mechanism for abnormal situations of the foldable rotors 220, which will be described in detail below.

[0077] Referring to Figure 10 , in this embodiment, a vertical plane defined by the forward direction of the aircraft body 210 serves as the mid-axis plane of the aircraft body 210. The aircraft body 210 is symmetrical about the mid-axis plane, and the two sides of the aircraft body 210 in its forward direction are defined as the left and right sides of the aircraft body 210. The weight of the left and right sides of the aircraft body 210 is approximately equal or not much different, so that the left and right parts of the aircraft body 210 are essentially in a state of balance. The number of foldable rotors 220 can be even or odd. In this embodiment, the number of foldable rotors 220 is set to an even number, and the multiple foldable rotors 220 can be evenly distributed on the left and right sides of the aircraft body 210. That is, the number of foldable rotors 220 corresponding to the left and right sides of the aircraft body 210 is the same, so that the driving force on the left and right sides of the aircraft body 210 is also approximately the same. In other embodiments, corresponding settings can be made according to the structural characteristics of the aircraft body 210. For example, if the aircraft 200 is a fuel-type or extended-range aircraft 200, the number and layout of the foldable rotors 220 on the left and right sides of the aircraft body 210 can be determined based on the specific position of the oil storage device on the aircraft 200 and the impact of the weight change of the oil storage device in flight on the center of gravity of the aircraft body 210.

[0078] The specific number of foldable rotors 220 should be considered and set based on factors such as the weight of the aircraft body 210 and the output power of the foldable rotors 220. In traditional multi-rotor manned aircraft, six rotors are convenient for deployment, but if one of the six rotors malfunctions, yaw loss can occur. Therefore, eight or more rotors are chosen to provide redundant power and overcome the yaw loss issue. However, the overall size of the flying vehicle 1000 must also be considered to ensure it can adapt to public roads and existing parking systems. For example, in this embodiment, when the vehicle 100 and the aircraft 200 are coupled, the overall length of the flying vehicle 1000 does not exceed 5.9 meters, the height does not exceed 2 meters, and the width does not exceed 2.2 meters. However, providing eight or more rotors will make it difficult to meet the overall size requirements of the flying car 1000, and will significantly increase the overall weight of the aircraft 200, reduce the lightweight degree of the aircraft 200, and shorten the flight time and range of the aircraft 200.

[0079] Therefore, this application chooses to use six foldable rotors 220 as the power system of aircraft 200. While meeting the power requirements of aircraft 200, this also achieves lightweightness and improves the endurance of aircraft 200. When aircraft 200 is coupled with vehicle 100, the overall size of flying car 1000 can adapt to public road settings and existing parking systems. With six foldable rotors 220, a malfunction of one of these foldable rotors 220 could cause the entire aircraft 200 to lose yaw control. Therefore, this embodiment of the application provides two ducted fans 230 to overcome this safety issue, overcoming the difficulty of traditional six-rotor manned aircraft that lacks yaw control capabilities.

[0080] Referring to Figure 10 , the left and right sides of the aircraft body 210, in the direction of travel of the aircraft 200, each correspond to three foldable rotors 220, thereby maintaining a roughly balanced state in the left-right direction. Furthermore, the aircraft body 210 is also roughly balanced in the direction of travel of the aircraft. Specifically, the aircraft body 210 in this embodiment includes a first body portion 2170, a second body portion 2180, and a third body portion 2190. The first body portion 2170, the second body portion 2180, and the third body portion 2190 are sequentially arranged and connected in parallel along the direction of travel of the aircraft 200. The six foldable rotors 220 are divided into three groups, each group has two foldable rotors 220, and the three groups of foldable rotors 220 correspond to the first main body 2170, the second main body 2180 and the third main body 2190 respectively. The three groups of foldable rotors 220 bear the weight of the first main body 2170, the second main body 2180 and the third main body 2190 respectively, so that the aircraft body 210 is also in a roughly balanced state in the forward direction of the aircraft 200.

[0081] Referring to Figure 11 , in this embodiment, to provide sufficient operating space for each foldable rotor 220, multiple foldable rotors 220 are sequentially and spaced apart around a frame 2210. For ease of understanding, when the aircraft 200 is in level flight, projecting the aircraft 200 onto a horizontal plane, the projections of the multiple foldable rotors 220 are distributed around the periphery of the projection of the aircraft body 210. This allows the multiple foldable rotors 220 to diverge relative to the aircraft body 210, preventing structural collisions between adjacent or nearby foldable rotors 220 during transport and ensuring a high degree of safety for the aircraft 200 during flight. However, to prevent the multiple foldable rotors 220 from being too dispersed, the divergence of the multiple foldable rotors 220 is limited in this embodiment. In this embodiment, the foldable rotor 220 forms a rotor rotating surface 2224 when rotating. When the aircraft 200 is in a level flight state, the aircraft 200 is projected onto the horizontal plane, and the projected portions of the rotor rotating surfaces 2224 of at least two adjacent foldable rotors 220 overlap, so that the multiple foldable rotors 220 have a certain degree of integrity and correlation, and the aerodynamic efficiency of the foldable rotors 220 is improved.

[0082] Referring to Figure 12 , in this embodiment, multiple foldable rotors 220 are arranged in a divergent manner. The lines connecting the center points of the multiple foldable rotors 220 form a polygon 2225. The geometric center of polygon 2225 and the center of gravity of the aircraft body 210 lie on the same vertical line, or the vertical line containing the geometric center of polygon 2225 and the vertical line containing the center of gravity of the aircraft body 210 are parallel to each other at a small distance. This ensures that the aircraft 200 maintains a balanced and stable state during flight. It should be understood that the formation of polygon 2225 should be understood as forming polygon 2225 by connecting the center points of the multiple foldable rotors 220 when projected on a horizontal plane. Furthermore, the multiple foldable rotors 220 in this embodiment are arranged in a centrally symmetrical manner, which can be understood as the polygon 2225 formed by the multiple foldable rotors 220 being roughly a centrally symmetrical figure, so that the driving force formed by any three connected foldable rotors 220 in this embodiment is roughly the same as the driving force formed by the other three connected foldable rotors 220, further improving the balance and safety of the aircraft 200 in flight.

[0083] In this embodiment, two foldable rotors 220 located on the same diagonal of polygon 2225 form a pair of corresponding foldable rotors 220, with the two corresponding foldable rotors 220 arranged symmetrically about the center. When the two foldable rotors 220 have the same output power, the two foldable rotors 220 form a balanced external force group relative to the aircraft body 210. In the direction of the line connecting the two foldable rotors 220, the aircraft body 210 can maintain balance under the action of the two foldable rotors 220. Furthermore, in this embodiment, the rotation directions of the two foldable rotors 220 in the relative group are arranged in opposite directions to generate a yaw moment.

[0084] In order to keep the multiple rotors in a dispersed state, traditional multi-rotor manned aircraft have fixed wings extending on both sides of the multi-rotor manned aircraft along its flight direction, and the rotors are set on the fixed wings so that the multiple rotors are arranged around the multi-rotor manned aircraft. However, the fixed wings cannot be folded and occupy a very large space. The multi-rotor manned aircraft cannot be reduced in size and cannot be driven onto public roads. It can only take off or land on the apron, which seriously limits the scope of use of the multi-rotor manned aircraft.

[0085] Referring to Figures 13 and 14 , in this embodiment, to achieve divergent positions of the multiple foldable rotors 220 relative to the aircraft body 210, a frame 2210 is configured to include a bracket 2211 and multiple arms 2212. The bracket 2211 is disposed on the top of the aircraft body 210, and the multiple arms 2212 are mounted on the bracket 2211. The arms 2212 are used to mount the foldable rotors 220, and therefore the number of arms 2212 matches the number of foldable rotors 220. For example, in this embodiment, the number of arms 2212 matches the number of foldable rotors 220, both being six. The six foldable rotors 220 are disposed in a one-to-one correspondence with the six arms 2212, with each foldable rotor 220 mounted on a corresponding arm 2212. The multiple arms 2212 are sequentially spaced along the circumference of the frame. The arm 2212 extends relative to the bracket 2211 in a direction away from the center of the bracket 2211, so that the multiple foldable rotors 220 are connected to the aircraft body 210 in a divergent state. As can be seen from the above description, the bracket 2211 in this embodiment can be understood as a structurally independent frame structure, which is arranged on the top of the aircraft body 210 and the flight cabin 2110. The bracket 2211 can be connected to the top of the aircraft body 210 by welding to enhance the integrity between the aircraft body 210 and the bracket 2211. The bracket 2211 can also be fixedly connected to the aircraft body 210 by fixing components (such as threaded fasteners such as bolts, or connecting agents such as structural adhesives) to facilitate replacement or maintenance of the bracket 2211.

[0086] In other embodiments, the support 2211 can serve as the overall framework of the aircraft body 210, generally defining the exterior contour of the aircraft body 210. The support 2211 is formed by overlapping multiple rod-like and / or column-like structures and is generally rectangular. The rectangular structure of the support 2211 has an interior space, part or all of which serves as the flight cabin 2110. In this embodiment, the arm 2212 is disposed on top of the support 2211, that is, the arm 2212 is disposed on top of the flight cabin 2110. The arm 2212 can be integrally formed with the support 2211 to enhance the integrity between the arm 2212 and the support 2211. The arm 2212 can also be fixedly and detachably connected to the support 2211 via a connecting structure, which not only ensures the stability of the connection between the arm 2212 and the support 2211 but also facilitates replacement or maintenance of the arm 2212. If the arm 2212 and the bracket 2211 are connected by a connecting structure, the arm 2212 can be set to be rotatably connected to the frame 2211. When the aircraft 200 is in a landing state, the arm 2212 can be folded and moved close to the bracket 2211 to reduce the size of the overall structure of the aircraft body 210.

[0087] When the aircraft 200 is in a landing state, or when the aircraft 200 is coupled to the vehicle 100, in order to enable the multiple foldable rotors 220 to transition from a diverging state to a converged state, in this embodiment, multiple arms 2212 are rotatably connected to the bracket 2211. When the foldable rotors 220 are in the diverging state, the multiple arms 2212 are in an extended state. The arms 2212 can rotate relative to the bracket 2211 in a direction approaching the bracket 2211 to overlap the outer periphery of the bracket 2211, thereby converting the multiple arms 2212 into a folded state and the multiple foldable rotors 220 into a converged state. This reduces the outer profile of the foldable rotors 220, allowing the size of the aircraft 200 to adapt to the vehicle 100 and ensuring that the size of the flying car 1000 meets the specifications. Conversely, when the multiple arms 2212 are in the folded state and the multiple foldable rotors 220 are in the gathered state, the arms 2212 can rotate relative to the bracket 2211 to restore to the extended state, so that the multiple foldable rotors 220 restore to the divergent state to prepare for flight.

[0088] Referring to Figures 15 and 16 , in order to further reduce the external profile of foldable rotor 220 when aircraft 200 is in a stopped and landing state, in this embodiment, in addition to the foldable arms 2212, foldable rotor 220 is also foldable. Specifically, foldable rotor 220 in this embodiment includes at least two blades 2222 and a hub 2223. The two blades 2222 are connected to opposite sides of hub 2223. The two blades 2222 can be brought closer together to assume a folded state. In the folded state, the two blades 2222 are arranged along the length of arm 2212, i.e., the blades 2222 are stacked on arm 2212, thereby reducing the external profile of aircraft 200. In order to simplify the structure of the foldable rotor 220, the foldable rotor 220 in this embodiment uses a single-axis single-propeller rotor, and the foldable rotor 220 is a fixed-propeller rotor, that is, when the foldable rotor 220 is in a rotating state, the blades 2222 do not deflect or move relative to the hub 2223.

[0089] Please refer to Figure 10 again. For the convenience of explanation and distinction between the various foldable rotors 220, the six foldable rotors 220 in this embodiment are respectively defined as a first rotor 2230, a second rotor 2240, a third rotor 2250, a fourth rotor 2260, a fifth rotor 2270 and a sixth rotor 2280. The six foldable rotors 220 are arranged in a surrounding manner. In the counterclockwise direction of their surrounding arrangement, the six foldable rotors 220 are arranged in ascending order, that is, in the counterclockwise direction, the first rotor 2230, the second rotor 2240, the third rotor 2250, the fourth rotor 2260, the fifth rotor 2270 and the sixth rotor 2280 are arranged in sequence. When the aircraft 200 is projected onto a horizontal plane in flight, with the forward direction of the aircraft 200 as a reference direction, the second rotor 2240, the third rotor 2250, and the fourth rotor 2260 are located to the left of the aircraft body 210, while the first rotor 2230, the sixth rotor 2280, and the fifth rotor 2270 are located to the right of the aircraft body 210. Specifically, the third rotor 2250 is located on the left side of the aircraft body 210, the second rotor 2240 is located at the left front of the aircraft body 210, and the fourth rotor 2260 is located at the left rear of the aircraft body 210. The sixth rotor 2280 is located on the right side of the aircraft body 210, the first rotor 2230 is located at the right front of the aircraft body 210, and the fifth rotor 2270 is located at the right rear of the aircraft body 210.

[0090] Referring to Figure 10 , it can be deduced from the foregoing that first rotor 2230 and second rotor 2240 correspond to third body portion 2190, third rotor 2250 and sixth rotor 2280 correspond to second body portion 2180 of aircraft body 210, and fourth rotor 2260 and fifth rotor 2270 correspond to first body portion 2170. Furthermore, first rotor 2230 and fourth rotor 2260 are located on the same diagonal line of polygon 2225 and form a pair of twos, second rotor 2240 and fifth rotor 2270 are located on the same diagonal line and form a pair of twos, and third rotor 2250 and sixth rotor 2280 are located on the same diagonal line and form a pair of twos. Assuming that the first rotor 2230, the third rotor 2250 and the fifth rotor 2270 are configured to rotate in a first direction (for example, clockwise), and the second rotor 2240, the fourth rotor 2260 and the sixth rotor 2280 are configured to rotate in a second direction (for example, counterclockwise), the second direction is set opposite to the first direction.

[0091] Referring to Figure 17 , as previously mentioned, when the aircraft 200 is in level flight, the projected portions of the rotor rotation surfaces 2224 of at least two adjacent foldable rotors 220 are projected onto a horizontal plane, thereby limiting the divergence of the foldable rotors 220. However, in actual spatial structures, it is necessary to prevent the rotor rotation surfaces 2224 of two adjacent foldable rotors 220 from overlapping. The rotor rotation surfaces 2224 of multiple foldable rotors 220 are located on at least two planes to prevent structural collisions between adjacent foldable rotors 220 and ensure the safety and sustainability of the foldable rotors. In this embodiment, the two foldable rotors 220 corresponding to the first body portion 2170 and the two foldable rotors 220 corresponding to the third body portion 2190 are arranged on the same horizontal plane, defined as a first horizontal plane. The two foldable rotors 220 corresponding to the second body portion 2180 are arranged on another horizontal plane, defined as a second horizontal plane, with the first horizontal plane being higher than the second horizontal plane. That is, third rotor 2250 and sixth rotor 2280 are located on the ground-facing side of first rotor 2230, second rotor 2240, fourth rotor 2260, and fifth rotor 2270. The height of third rotor 2250 and sixth rotor 2280 is lower than that of first rotor 2230, second rotor 2240, fourth rotor 2260, and fifth rotor 2270 to prevent collision between third rotor 2250 and second rotor 2240 and fourth rotor 2260, and between sixth rotor 2280 and first rotor 2230 and fifth rotor 2270. To prevent collision between first rotor 2230 and second rotor 2240, and between fourth rotor 2260 and fifth rotor 2270, which are located in the same plane, reasonable avoidance measures can be implemented based on the lengths of arm 2212 and blades 2222. This embodiment will not be described in detail herein. In other embodiments, each foldable rotor 220 may be arranged on a different plane, which is not specifically limited in this embodiment.

[0092] In this embodiment, the first rotor 2230, the second rotor 2240, the fourth rotor 2260 and the fifth rotor 2270 are spaced apart at the four corners of the aircraft body 210 in the forward direction of the aircraft 200. The structures of the first rotor 2230, the second rotor 2240, the fourth rotor 2260 and the fifth rotor 2270 and the airflow generated do not interfere with each other. Therefore, in this embodiment, the rotation axes of the first rotor 2230, the second rotor 2240, the fourth rotor 2260 and the fifth rotor 2270 are arranged along the vertical direction, that is, the blades 2222 of the first rotor 2230, the second rotor 2240, the fourth rotor 2260 and the fifth rotor 2270 are arranged roughly horizontally, forming a roughly horizontal rotor rotating surface 2224, so as to provide vertical driving force to the aircraft body 210.

[0093] Referring to FIG. 11 , when aircraft 200 is projected onto a horizontal plane, there is a projected overlap between third rotor 2250 and second and fourth rotors 2240, 2260. To prevent the airflow directed by third rotor 2250 from overlapping significantly with the airflow directed by second and fourth rotors 2240, 2260 in the actual spatial structure, third rotor 2250 is tilted in this embodiment (as shown in FIG. 17 ), such that both the rotation axis and rotor rotating surface 2224 of third rotor 2250 are tilted. The tilted rotor rotating surface 2224 of third rotor 2250 reduces the mutual influence of airflow between third rotor 2250 and second and fourth rotors 2240, 2260, thereby improving the aerodynamic efficiency of foldable rotor 220.

[0094] Specifically, the two blades 2222 of the third rotor 2250 are substantially aligned in a straight line when deployed. When the two blades 2222 rotate to the same vertical plane, the blade 2222 closer to the aircraft body 210 is higher than the blade 2222 farther from the aircraft body 210. This causes the rotor tip 2229 of the third rotor 2250, which faces the sky, to be tilted outward relative to the aircraft body 210, and the rotor rotation surface 2224 of the third rotor 2250 to be tilted outward relative to the aircraft body 210. This outward tilt allows the radial direction of the third rotor 2250 to be offset from the flight cabin 2110, preventing the radial direction of the third rotor 2250 from being perpendicular to the flight cabin 2110. This reduces damage to the flight cabin 2110 and its occupants in the event of an accidental explosion.

[0095] To further protect the structure of each foldable rotor 220, in this embodiment, the rotor rotation surface 2224 of the third rotor 2250 is lower than the rotor rotation surfaces 2224 of the second rotor 2240 and the fourth rotor 2260. In this embodiment, the angle between the rotation axis of the third rotor 2250 and the vertical direction is greater than or equal to 5°. Combined with the above, it can be determined that the angle between the rotation axis of the third rotor 2250 and the vertical direction in this embodiment is in the range of [5°, 10°]. In other embodiments, the angle between the rotation axis of the third rotor 2250 and the vertical direction can be adaptively adjusted according to actual conditions.

[0096] Please refer to Figure 17. In this embodiment, the positional relationship between the sixth rotor 2280, the first rotor 2230 and the fifth rotor 2270 can be set with reference to the positional relationship between the third rotor 2250, the second rotor 2240 and the fourth rotor 2260 mentioned above. In this embodiment, the sixth rotor 2280, the first rotor 2230 and the fifth rotor 2270 are roughly mirror-symmetrical with the third rotor 2250, the second rotor 2240 and the fourth rotor 2260 about the central axis plane of the aircraft body 210.

[0097] When aircraft 200 is projected onto a horizontal plane, there is a projected overlap between sixth rotor 2280, first rotor 2230, and fifth rotor 2270. To prevent the airflow directed by sixth rotor 2280 from overlapping significantly with the airflow directed by first rotor 2230 and fifth rotor 2270 in the actual spatial structure, sixth rotor 2280 is tilted in this embodiment (as shown in FIG. 17 ), such that both the rotation axis and the rotor rotation surface of sixth rotor 2280 are tilted. The tilted configuration of rotor rotation surface 2224 of sixth rotor 2280 can reduce the mutual influence of airflow between sixth rotor 2280 and first rotor 2230 and fifth rotor 2270, thereby improving the aerodynamic efficiency of foldable rotor 220.

[0098] Specifically, the two blades 2222 of the sixth rotor 2280 are substantially aligned in a straight line when deployed. When the two blades 2222 rotate to the same vertical plane, the blade 2222 closer to the aircraft body 210 is higher than the blade 2222 farther from the aircraft body 210. This causes the skyward-facing rotor tip of the sixth rotor 2280 to be tilted outward relative to the aircraft body 210, and the rotor rotation surface 2224 of the sixth rotor 2280 is also tilted outward relative to the aircraft body 210. This outward tilt allows the radial direction of the motor of the sixth rotor 2280 to be offset from the flight cabin 2110. This prevents the radial direction of the motor of the sixth rotor 2280 from being perpendicular to the flight cabin 2110, thereby reducing damage to the flight cabin 2110 and its occupants in the event of an accidental explosion.

[0099] To further protect the structure of each foldable rotor 220, in this embodiment, the rotor rotation surface 2224 of the sixth rotor 2280 is lower than the rotor rotation surfaces 2224 of the first rotor 2230 and the fifth rotor 2270. In this embodiment, the angle between the rotation axis of the sixth rotor 2280 and the vertical direction is greater than or equal to 5°. Based on the above, the angle between the rotation axis of the sixth rotor 2280 and the vertical direction in this embodiment can be determined to be in the range of [5°, 10°]. In other embodiments, the angle between the rotation axis of the sixth rotor 2280 and the vertical direction can be adaptively adjusted according to actual conditions.

[0100] In this embodiment, four of the six arms 2212 are provided in one-to-one correspondence with the first rotor 2230, the second rotor 2240, the fourth rotor 2260, and the fifth rotor 2270, while the remaining two arms 2212 are provided in one-to-one correspondence with the third rotor 2250 and the sixth rotor 2280. For ease of distinction, the arm 2212 corresponding to the third rotor 2250 is defined as the first arm 2213, and the arm 2212 corresponding to the sixth rotor 2280 is defined as the second arm 2214. The first arm 2213 and the second arm 2214 are respectively provided on either side of the aircraft body 210 in the direction of travel of the aircraft 200.

[0101] Referring to Figure 6 , to achieve the staggered placement of the six foldable rotors 220, in this embodiment, when all six arms 2212 are in the folded state, the four arms 2212, excluding the first arm 2213 and the second arm 2214, are stacked side by side on the same plane, which is generally horizontal. The first arm 2213 and the second arm 2214 are stacked on the ground-facing side of the four arms 2212 to ensure the proper positioning of the third rotor 2250 and the sixth rotor 2280 relative to the first rotor 2230, the second rotor 2240, the fourth rotor 2260, and the fifth rotor 2270, thereby improving space utilization for the six arms 2212 and the six foldable rotors 220.

[0102] Referring to Figure 4 , the auxiliary yaw system of the aircraft 200 of this embodiment includes two ducted fans 230, mounted on the aircraft body 210. When the yaw torque generated by the six foldable rotors 220 is insufficient, or when one of the six foldable rotors 220 fails, the ducted fans 230 provide auxiliary yaw torque to compensate for the yaw torque deviation caused by the failed foldable rotor 220. In this embodiment, the ducted fans 230 remain stable relative to the structure of the aircraft body 210, simplifying their installation. The fixed position of the ducted fans 230 is specifically manifested in that the ducted fans 230 have a rotation axis, and during flight, the rotation axis of the ducted fans 230 remains fixed relative to the aircraft body 210. In other embodiments, the ducted fans 230 can be movably mounted on the aircraft body 210 to increase their flexibility and improve the performance of the auxiliary yaw system. In this embodiment, the ducted fan 230 is used as an auxiliary yaw system. The ducted fan 230 can generate a large yaw moment. When the aircraft 200 is in flight, the ducted fan 230 is subject to less resistance and has a higher overall working efficiency.

[0103] Additionally, when aircraft 200 is in level flight, the angle between the rotation axis of ducted fan 230 and the horizontal plane in this embodiment is approximately less than or equal to 5°. In other embodiments, the angle between the rotation axis of ducted fan 230 and the horizontal plane can be adaptively adjusted based on actual conditions.

[0104] Referring to Figure 4 , to facilitate illustration of the specific location of each ducted fan 230 on the aircraft body 210 and its relative positional relationship to other structures, in this embodiment, two ducted fans 230 are defined as a first ducted fan 2311 and a second ducted fan 2312. Specifically, the first ducted fan 2311 is positioned corresponding to the third rotor 2250 and is mounted on the first arm 2213. The second ducted fan 2312 is positioned corresponding to the sixth rotor 2280 and is mounted on the second arm 2214. The first arm 2213 and the second arm 2214 are located on the left and right sides of the aircraft body 210, respectively, in the direction of aircraft 200's travel. When the aircraft body 210 deflects, the first and second ducted fans 2311 and 2312, positioned on the first and second arms 2213 and 2214, exert a significant deflection moment. When the rated power of the first ducted fan 2311 and the second ducted fan 2312 remains unchanged, the compensation adjustment range formed by the first ducted fan 2311 and the second ducted fan 2312 being arranged on the first arm 2213 and the second arm 2214 is larger than the compensation adjustment range formed by the first ducted fan 2311 and the second ducted fan 2312 being arranged on the four arms 2212 except the first arm 2213 and the second arm 2214.

[0105] In this embodiment, the first ducted fan 2311 and the third rotor 2250 are both disposed on the first arm 2213. The first ducted fan 2311 and the third rotor 2250 can be located at the end of the first arm 2213 away from the aircraft body 210. This end is farther away from the center of gravity of the aircraft body 210, resulting in a relatively long lateral moment arm and a relatively large yaw moment, thereby increasing the power efficiency of the first ducted fan 2311.

[0106] Specifically, the first arm 2213 has a top side and a bottom side that face away from each other, with the top side of the first arm 2213 facing the sky and the bottom side of the first arm 2213 facing the ground. The third rotor 2250 is disposed on the top side of the first arm 2213, and the first ducted fan 2311 is disposed on the bottom side of the first arm 2213, so that the third rotor 2250 and the first ducted fan 2311 are spaced apart from each other to avoid structural collisions. Furthermore, in order to reduce the impact of the airflow at the end of the blades 2222 of the third rotor 2250 on the first ducted fan 2311, in this embodiment, the first ducted fan 2311 is disposed away from the blades 2222 of the third rotor 2250. Specifically, in this embodiment, the first ducted fan 2311 has a center of rotation (i.e., an axis of rotation). The hub 2223 of the third rotor 2250 forms a first maximum rotation surface when in a rotating state. When the aircraft 200 is projected toward the horizontal plane, the projection surface of the first maximum rotation surface covers the rotation center of the first ducted fan 2311, so that the rotation center of the first ducted fan 2311 and its nearby structure correspond to the hub 2223 of the third rotor 2250. From a macroscopic structural point of view, the rotation center (that is, the axis of rotation) of the first ducted fan 2311 is located directly below the hub 2223 of the third rotor 2250. Therefore, the airflow at the hub 2223 of the third rotor 2250 is relatively small and weak, and has little impact on the first ducted fan 2311.

[0107] Similarly, in this embodiment, the second ducted fan 2312 has a rotation center (i.e., a rotation axis). The hub 2223 of the sixth rotor 2280 forms a second maximum rotation surface when in a rotating state. When the aircraft 200 is projected onto the horizontal plane, the projection surface of the second maximum rotation surface covers the rotation center of the second ducted fan 2312, so that the rotation center of the second ducted fan 2312 and its surrounding structures correspond to the hub 2223 of the sixth rotor 2280. From a macroscopic structural point of view, the rotation center (i.e., the rotation axis) of the second ducted fan 2312 is located directly below the hub 2223 of the sixth rotor 2280. Therefore, the airflow below the hub 2223 of the sixth rotor 2280 is relatively small and weak, and has little impact on the second ducted fan 2312.

[0108] In order to monitor the operating status of each foldable rotor 220 and control the operating nodes of the ducted fans 230 according to the operating status of each foldable rotor 220, the aircraft 200 in this embodiment also includes a yaw control system (not shown in the figure). The yaw control system is electrically connected to the foldable rotor 220 and the ducted fan 230 respectively to enable information exchange and command execution. The control mechanism of the yaw control system for the failure of one of the foldable rotors 220 is as follows: when any one of the multiple foldable rotors 220 fails, the yaw control system controls the operation of the two ducted fans 230 to provide a compensating yaw moment to the aircraft 200. The magnitude of the compensating yaw moment is roughly equal to the difference between the actual yaw moment jointly formed by the remaining five foldable rotors 220 and the target yaw moment. The yaw control system enables the aircraft 200 to cope with the phenomenon of foldable rotor 220 failure and ensures the safety of the pilot and the aircraft 200.

[0109] In addition, the control mechanism of the yaw control system for the situation where the six foldable rotors 220 operate normally but the yaw moment is insufficient is as follows: the yaw control system controls the operation of the two ducted fans 230 to provide a compensatory yaw moment to the aircraft 200, and the magnitude of the compensatory yaw moment is roughly equal to the difference between the actual yaw moment jointly constituted by the six foldable rotors 220 and the target yaw moment.

[0110] As previously mentioned, the first rotor 2230 and the fourth rotor 2260 are paired in pairs, the second rotor 2240 and the fifth rotor 2270 are paired in pairs, and the third rotor 2250 and the sixth rotor 2280 are paired in pairs. Each set of foldable rotors 220 is capable of maintaining balance on the corresponding diagonal line of the aircraft body 210. To simplify the yaw control system's control of the six foldable rotors 220, the yaw control system in this embodiment is further configured to: if any foldable rotor 220 fails, the corresponding foldable rotor 220 is controlled to stop operating, and the two ducted fans 230 are controlled to operate to provide a compensating yaw moment to the aircraft body 210. This configuration has the advantage that, if the first rotor 2230 fails, the yaw control system controls the fourth rotor 2260 to stop operating, and controls the two ducted fans 230 to operate to compensate for the deviation in yaw moment caused by the failure of the first and fourth rotors 2230 and 2260. During this process, the yaw control system only needs to use the two ducted fans 230 to compensate for the yaw torque required to deflect the aircraft body 210. This eliminates or reduces the need to consider the fourth rotor 2260, for example, due to differences in position, output power, and drag between the fourth rotor 2260 and the two ducted fans 230. This simplifies the calculation steps of the yaw control system, improves operational efficiency, and allows the aircraft 200 to return to normal flight in a shorter period of time.

[0111] Furthermore, assuming that first rotor 2230 fails and the yaw control system stops fourth rotor 2260, second rotor 2240, fifth rotor 2270, third rotor 2250, and sixth rotor 2280 can still maintain the balance of aircraft body 210. The yaw control system only needs to control the four normally functioning foldable rotors 220 to increase their output power to compensate for the vertical traction and achieve the desired traction. There is no need to comprehensively consider the balance adjustment between second rotor 2240, third rotor 2250, fourth rotor 2260, fifth rotor 2270, and sixth rotor 2280.

[0112] In summary, the embodiment of the present application provides a purely electric aircraft 200, which includes an aircraft body 210, a frame 2210, and a plurality of non-tilting foldable rotors 220. The aircraft body 210 is provided with a flight cabin 2110 for carrying the driver and passengers. The aircraft 200 uses a plurality of foldable rotors as a power system, wherein the frame 2210 is located at the top of the aircraft body 210, the plurality of foldable rotors 220 are connected to the frame 2210, and the plurality of foldable rotors 220 are arranged at intervals around the vertical line where the center of gravity of the aircraft 200 is located. When the aircraft body 210 is in flight, the rotation axis of the foldable rotors 220 is always in a fixed state relative to the position of the aircraft body 210. To address unbalanced yaw moments generated by the multiple non-tilting foldable rotors 220, such as those resulting from yaw moment imbalance or a failure of at least one of the multiple foldable rotors 220, aircraft 200 is further provided with two ducted fans 230. These two ducted fans 230 are connected to a frame 2210 and located on the left and right sides of the aircraft body 210. These fans provide compensatory yaw moments to the aircraft body 210 in the event of an unbalanced yaw moment generated by the foldable rotors 220. Compared to conventional multi-rotor manned aircraft, aircraft 200 in this embodiment does not utilize fixed wings to disperse the multiple foldable rotors 220. Instead, the multiple foldable rotors are mounted on a partially foldable frame 2210, allowing the overall structure of aircraft 200 to be retracted or extended. When aircraft 200 is coupled to vehicle 100, the size of the flying car 1000 can adapt to the road. In addition, the aircraft 200 is provided with two ducted fans 230 on the left and right sides of its forward direction. When the yaw moment formed by multiple non-tilt foldable rotors 220 is in an unbalanced state, the two ducted fans 230 can operate and provide compensatory yaw moment to the aircraft body 210, so that when the aircraft body 210 is provided with six foldable rotors 220, the problem of being unable to achieve yaw due to the imbalance of yaw moment can be overcome.

[0113] Referring to FIG. 18 , an embodiment of the present application further provides a flight control method for an aircraft based on the aforementioned aircraft and the aircraft yaw control system. The flight control method includes the following steps:

[0114] S10: Obtain the target yaw moment of the aircraft.

[0115] The target yaw moment is the yaw moment required to yaw the aircraft by a certain angle (e.g., the target yaw angle, which can be set based on actual requirements). The yaw control system determines the target yaw moment required to achieve this yaw motion based on the target yaw moment and the aircraft's parameters. Specifically, these parameters include the aircraft's weight, the drag it encounters, the air pressure within the atmosphere, and the air velocity.

[0116] In some embodiments, the step of "S10: obtaining a target yaw moment of the aircraft" includes: obtaining a current yaw angle and a target yaw angle of the aircraft; and determining a target yaw moment of the aircraft based on the current yaw angle and the target yaw angle.

[0117] S20: During the flight of the aircraft, a first yaw moment generated by the foldable rotor is obtained.

[0118] The yaw control system obtains the actual output power of the multiple foldable rotors, calculates the first yaw moment generated by the multiple foldable rotors based on the actual output power of the multiple foldable rotors, and compares the first yaw moment with the target yaw moment to determine whether the first yaw moment can meet the requirements for the aircraft to deflect to the target yaw angle.

[0119] S30: When the difference between the first yaw moment and the target yaw moment is greater than a specified value, determine a compensating yaw moment according to the difference.

[0120] When the yaw control system determines that the difference between the first yaw moment and the target yaw moment is greater than a specified value, it indicates that the aircraft body cannot successfully achieve the target yaw angle under the first yaw moment. The specified value in this embodiment refers to the difference between the upper and lower limits of a yaw moment range that enables the aircraft body to achieve the target yaw angle, wherein the target yaw moment is within this range. In some examples, the specified value refers to the difference between the minimum yaw moment that enables the aircraft body to achieve the target yaw angle and the target yaw moment. In other embodiments, the specified value may also be a numerical range, where the minimum value within the numerical range indicates that when the difference between the target yaw moment and the first yaw moment is the minimum value within the numerical range, the first yaw moment can fully support the aircraft's yaw to the target yaw angle, and the maximum value within the numerical range indicates that when the difference between the target yaw moment and the first yaw moment is the maximum value within the numerical range, the first yaw moment can substantially support the aircraft's yaw to the target yaw angle.

[0121] The yaw control system then calculates the compensatory yaw moment that the auxiliary yaw system needs to provide according to the first yaw moment, the target yaw moment and the specified value.

[0122] S40: Controlling the auxiliary yaw system to operate according to the compensated yaw moment, wherein the difference between the sum of the second yaw moment and the first yaw moment generated by the operation of the auxiliary yaw system and the target yaw moment is less than or equal to a specified value.

[0123] The yaw control system controls the operation of the auxiliary yaw system, which activates the ducted fan and generates a second yaw moment. The difference between the sum of the second yaw moment and the first yaw moment and the target yaw moment must be less than or equal to a specified value, so that the aircraft can achieve the target yaw angle under the combined drive of the foldable rotor and the auxiliary yaw system.

[0124] In some embodiments, the operation of the auxiliary yaw system is controlled based on the compensated yaw moment, including: determining the target power of the ducted fan based on the compensated yaw moment; determining the target speed of the ducted fan based on the target power and the size and aerodynamic efficiency of the ducted fan; and controlling the operation of the ducted fan based on the target speed so that the ducted fan generates a second yaw moment.

[0125] Therefore, in this embodiment, an auxiliary yaw system is introduced. When the yaw moment generated by the foldable rotor is unbalanced, the auxiliary yaw system can operate and provide compensatory yaw torque to the aircraft body, thereby enabling the aircraft body to overcome the problem of being unable to yaw due to the unbalanced yaw moment. Specifically, the power of the auxiliary yaw system can change in real time with the flight attitude and flight power parameters. When a foldable rotor failure is detected, the auxiliary yaw system begins to operate. When the aircraft yaw deflects actively or passively, the aircraft controller or flight control system can control the aircraft yaw according to established control principles (such as heading hold or major nose turn). If the aircraft controller or flight control system finds that its control capability is insufficient, the auxiliary yaw system is used for supplementary control to meet the aircraft yaw control requirements.

[0126] In some embodiments, the step of “S20 obtaining a first yaw moment generated by the foldable rotor during the flight of the aircraft” in the above method includes the following steps S21 to S22.

[0127] S21: During the flight of the aircraft, the operating parameters of each foldable rotor and the fault index parameters of the foldable rotor are obtained, wherein the fault index parameters are monitored by the fault self-detection system of the aircraft, and the fault index parameters represent the degree of failure of the foldable rotor.

[0128] Specifically, the aircraft controller or yaw control system can obtain operating parameters of each foldable rotor. The operating parameters of the foldable rotors may include the output power of the foldable rotors, the drag experienced by the foldable rotors, etc. The aircraft also includes a fault self-diagnosis system electrically connected to the yaw control system. The fault self-diagnosis system is configured to detect the extent of a foldable rotor fault (e.g., the extent of damage or severity of the fault), record the extent of the foldable rotor fault as a fault indicator parameter, and transmit the foldable rotor fault indicator parameter to the yaw control system. Taking the rotor assembly as an example of a foldable rotor, as an example, the fault detection system may include a fault detection circuit connected to the motor drive circuit, which is used to determine whether the rotor motor is operating normally based on the electrical signal condition of the motor drive circuit (for example, at least one of the output power condition, current size, and voltage size), thereby deriving the degree of fault of the rotor motor; as another example, the fault detection system may also include a speed sensor, which may be connected to the rotor shaft, which is used to detect the rotation state of the rotor (for example, at least one of the rotation direction, rotation speed, and rotation acceleration), and determine whether the rotor is operating normally based on the rotation state, thereby deriving the degree of fault of the rotor.

[0129] In some examples, the information carried by the fault indicator parameters may include the specific faulty component and the extent of damage to the specific fault. For example, the fault indicator parameters may be represented by a code, such as the fault code EA2, where one bit of the code EA represents the specific fault location (e.g., a motor or rotor), and another bit of the code represents the severity of the fault (e.g., a primary or secondary fault). Therefore, the entire machine controller can determine the specific fault location and severity by reading the fault indicator parameters.

[0130] S22: When it is determined based on the operating parameters and the fault indicator parameters that at least one of the plurality of foldable rotors has a fault, a first yaw moment generated by the foldable rotor is obtained.

[0131] Specifically, because the fault indicator parameter carries information including the specific faulty component and the extent of damage caused by the specific fault, the entire machine controller or yaw control system can determine the number of faulty foldable rotors among the multiple foldable rotors based on the fault indicator parameter. If at least one of the multiple foldable rotors is faulty, the entire machine controller or yaw control system can obtain the first yaw moment generated by the foldable rotor based on the operating parameters of the foldable rotor. In some examples, if the foldable rotor is not faulty, it is not necessary to specifically obtain the first yaw moment generated by the foldable rotor. However, if the fault indicator parameter indicates that the foldable rotor is faulty, indicating that the generated yaw moment may deviate from the target yaw moment, it is necessary to obtain the first yaw moment generated by the foldable rotor. Therefore, in this embodiment, the system resources consumed by control and calculation are relatively small, which facilitates smooth operation of the entire machine controller.

[0132] In some embodiments, when a single foldable rotor among multiple foldable rotors is faulty, the step of "S22: determining, based on operating parameters and fault indicator parameters, that at least one of the multiple foldable rotors is faulty, and obtaining the first yaw moment generated by the foldable rotor" includes the following steps S221 to S222.

[0133] S221: When it is determined based on the operating parameters and the fault indicator parameters that one of the multiple foldable rotors is faulty, the foldable rotor corresponding to the faulty foldable rotor is controlled to stop operating.

[0134] The yaw control system, incorporating the aforementioned distribution of the power system, controls the corresponding foldable rotor to stop operating in the event of a failure of one of the multiple foldable rotors, thereby maintaining the horizontal balance of the aircraft. Because the multiple foldable rotors in this embodiment are arranged symmetrically about the central axis and in pairs, if one of the multiple foldable rotors fails, the corresponding foldable rotor's control process can essentially maintain the lift and thrust provided by the remaining foldable rotors. This also makes operating parameters easier to control and reduces the computational burden on the overall control system.

[0135] S222: Obtain a first yaw moment generated by all foldable rotors except the foldable rotor that has failed and its corresponding foldable rotor.

[0136] Taking six foldable rotors as an example, two of the multiple foldable rotors stop working, the yaw control system recalculates the first yaw moment generated by the remaining four foldable rotors, and controls the auxiliary yaw system to operate to provide a compensatory yaw moment.

[0137] Please refer to Figure 19, which is a schematic diagram of the functional framework of a flight control device disclosed in an embodiment of the present application. Based on the above-mentioned aircraft and flight control method thereof, an embodiment of the present application further provides a flight control device, which may include a target torque calculation module A10, an actual torque calculation module A20, a compensation torque calculation module A30, and a yaw compensation module A40.

[0138] The target torque calculation module A10 is used to obtain the target yaw moment of the aircraft, and the actual torque calculation module A20 is used to obtain the first yaw moment generated by the foldable rotor during the flight of the aircraft. The compensation torque calculation module A30 is used to determine the compensation yaw moment according to the difference between the first yaw moment and the target yaw moment when the difference between the first yaw moment and the target yaw moment is greater than a specified value. The compensation yaw module A40 is used to control the operation of the auxiliary yaw system according to the compensation yaw moment. Among them, the difference between the sum of the second yaw moment and the first yaw moment generated by the operation of the auxiliary yaw system and the target yaw moment is less than or equal to the specified value. The specific working process of each of the above-mentioned functional units / functional modules can refer to the corresponding process in the aforementioned method embodiment and will not be repeated here.

[0139] In the embodiments provided herein, the coupling between units / modules may be electrical, mechanical, or other forms of coupling. Furthermore, the functional units / functional modules in the various embodiments of the present invention may be integrated into a single processing unit, or each unit may exist physically separately, or two or more units may be integrated into a single unit. The aforementioned integrated units may be implemented in the form of hardware or software functional units.

[0140] Based on the aforementioned flying car and flight control method, embodiments of the present application may also provide an aircraft, which may include a processor and a memory. The memory may store one or more computer programs. The one or more computer programs are configured to execute the method described in the aforementioned method embodiment. The memory may be independent or integrated with the processor.

[0141] The processor may include one or more processing cores. The processor may utilize various interfaces and lines to connect various parts of the entire land body, and may execute various functions of the land body and process data by running or executing instructions, programs, code sets, or instruction sets stored in the memory, and calling data stored in the memory. In one embodiment, the processor may be implemented in at least one hardware form selected from digital signal processing (DSP), field programmable gate array (FPGA), and programmable logic array (PLA). The processor may integrate one or a combination of a central processing unit (CPU), a graphics processing unit (GPU), and a modem. The CPU primarily processes the operating system, user interface, and application programs; the GPU is responsible for rendering and drawing display content; and the modem is used to handle wireless communications. It is understandable that the aforementioned modem may not be integrated into the processor, but may be implemented separately through a communication chip.

[0142] The memory may include random access memory (RAM) or read-only memory (ROM). The memory may be used to store instructions, programs, codes, code sets, or instruction sets. The memory may include a program storage area and a data storage area, wherein the program storage area may store instructions for implementing an operating system, instructions for implementing at least one function (such as a touch function, a sound playback function, an image playback function, etc.), instructions for implementing the above-mentioned various method embodiments, etc. The data storage area may also store data created by the land body during use (such as a phone book, audio and video data, chat history data, etc.).

[0143] When the computer program instructions stored in the memory are executed, the processor can be used to perform various operations performed by the land body in the above method embodiment. The specific implementation of these operations can be found in the previous embodiments and will not be described in detail here.

[0144] The present application also provides a computer-readable storage medium. The computer-readable medium stores computer program code, which can be invoked by a processor to perform various operations in the above method embodiments. The specific implementation of each of the above operations can be found in the previous embodiments and will not be described in detail here.

[0145] The computer-readable storage medium may be an electronic memory such as a flash memory, an electrically erasable programmable read-only memory (EEPROM), an erasable programmable read-only memory (EPROM), a hard disk, or a ROM. In one embodiment, the computer-readable storage medium may include a non-transitory computer-readable storage medium. The computer-readable storage medium has storage space for program code for executing any of the method steps in the above method. These computer program codes can be read from or written to one or more computer program products. The computer program code can be compressed, for example, in an appropriate form.

[0146] In the specification of this application, certain words are used to refer to specific components in the specification and claims. Those skilled in the art should understand that hardware manufacturers may use different terms to refer to the same component. The specification and claims do not use the difference in name as a way to distinguish components, but use the difference in function of the components as the criterion for distinction. For example, "including" mentioned throughout the specification and claims is an open term and should be interpreted as "including but not limited to"; "substantially" means that those skilled in the art can solve the technical problem within a certain error range and basically achieve the technical effect.

[0147] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "inside", etc. indicating directions or positional relationships are based on the directions or positional relationships shown in the accompanying drawings. They are only used to simplify the description for the convenience of describing this application, and do not indicate or imply that the device or element referred to must have a specific direction, be constructed and operated in a specific direction. Therefore, they should not be understood as limitations on this application.

[0148] In this application, unless otherwise expressly specified or limited, terms such as "mounted," "connected," "connect," and "fixed" should be interpreted broadly. For example, these terms may refer to fixed, removable, or integral connections; mechanical or electrical connections; direct or indirect connections through an intermediary; internal communication between two components; or mere surface contact. Those skilled in the art will understand the specific meanings of these terms in this application based on the specific circumstances.

[0149] In the description of this specification, the description with reference to the terms "one embodiment", "some embodiments", "example", "specific example", or "some examples" means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art can combine and combine different embodiments or examples described in this specification and features of different embodiments or examples without contradiction.

[0150] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of the technical features being referred to. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of such features. Throughout the description of this application, "plurality" means at least two, for example, two, three, etc., unless otherwise specifically defined.

[0151] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit them. Although the present application 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 cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A pure electric aircraft, wherein: The pure electric aircraft comprises: An aircraft body, wherein the aircraft body is provided with a flight cabin for carrying a driver and passengers; A frame, located on the top of the aircraft body; a plurality of non-tilt rotors, the plurality of non-tilt rotors being connected to the frame, the plurality of non-tilt rotors being arranged at intervals around a vertical line where the center of gravity of the aircraft is located, and when the aircraft body is in flight, the position of the rotation axis of the non-tilt rotor relative to the aircraft body is always in a fixed state; and, Two ducted fans are connected to the frame and are respectively located on the left and right sides of the aircraft body. When the yaw moment provided by the non-tilt rotor is in an unbalanced state, the two ducted fans provide a compensating yaw moment to the aircraft body.

2. The pure electric aircraft according to claim 1, wherein: The aircraft body is symmetrical about a predetermined central axis plane, and the central axis plane is a vertical plane defined by the forward direction of the aircraft; the two ducted fans are respectively located on both sides of the central axis plane; The number of the non-tilt rotors is six, and the six non-tilt rotors are arranged at intervals around a vertical line where the center of gravity of the aircraft is located, and three non-tilt rotors are arranged on each side of the central axis plane.

3. The pure electric aircraft according to claim 2, wherein: When the aircraft is in flight, the position of the rotation axis of the ducted fan relative to the aircraft body is always fixed.

4. The pure electric aircraft according to claim 2 or 3, wherein: The non-tilt rotor has a rotor rotating surface; in a flight state, multiple rotor rotating surfaces of multiple non-tilt rotors are located in at least two planes; when the aircraft is projected onto a horizontal plane in a level flight state, the projections of at least two adjacent rotor rotating surfaces of the non-tilt rotors partially overlap.

5. The pure electric aircraft according to any one of claims 2 to 4, wherein: The aircraft also includes a yaw control system, which is electrically connected to the six non-tilt rotors and the two ducted fans, respectively. The yaw control system is configured to control the two ducted fans to operate so as to provide the compensating yaw moment to the aircraft body when any one of the non-tilt rotors fails.

6. The pure electric aircraft according to claim 5, wherein: In the direction in which the non-tilt rotors are distributed circumferentially, lines connecting the center points of the six non-tilt rotors in sequence form a polygon, and two of the non-tilt rotors located on the same diagonal of the polygon constitute a pair of corresponding non-tilt rotors; the yaw control system is specifically configured as follows: when any one of the non-tilt rotors fails, the non-tilt rotor corresponding to the failed non-tilt rotor is controlled to stop working, and the two ducted fans are controlled to operate to provide the compensating yaw moment to the aircraft body.

7. The pure electric aircraft according to claim 6, wherein: The six non-tilt rotors include a first rotor, a second rotor, a third rotor, a fourth rotor, a fifth rotor and a sixth rotor. In the direction in which the six non-tilt rotors are distributed around, the six non-tilt rotors are arranged in ascending order in a counterclockwise direction. Among them, the first rotor and the fourth rotor are paired in pairs, the second rotor and the fifth rotor are paired in pairs, and the third rotor and the sixth rotor are paired in pairs.

8. The pure electric aircraft according to claim 7, wherein: The first rotor, the third rotor, and the fifth rotor are configured to rotate in a first direction, and the second rotor, the fourth rotor, and the sixth rotor are configured to rotate in a second direction, which is opposite to the first direction.

9. The pure electric aircraft according to any one of claims 2 to 8, wherein: The six non-tilt rotors include a first rotor, a second rotor, a third rotor, a fourth rotor, a fifth rotor and a sixth rotor. In the direction in which the six rotors are distributed around, the six non-tilt rotors are arranged in ascending order in a counterclockwise direction. When the aircraft is projected onto a horizontal plane in a flight state, the projections of the third rotor and the sixth rotor are respectively located on both sides of the aircraft body in the forward direction, the first rotor is located at the right front of the aircraft body, the second rotor is located at the left front of the aircraft body, the fourth rotor is located at the left rear of the aircraft body, and the fifth rotor is located at the right rear of the aircraft body.

10. The pure electric aircraft according to claim 9, wherein: In the level flight state of the aircraft, the angle between the rotation axis of the third rotor and the vertical direction is greater than or equal to 5 degrees and less than or equal to 10 degrees, and the rotor tip of the third rotor is inclined outwardly relative to the aircraft body; or / and, The angle between the rotation axis of the sixth rotor and the vertical direction is greater than or equal to 5 degrees and less than or equal to 10 degrees, and the rotor tip of the sixth rotor is inclined outward relative to the aircraft body.

11. The pure electric aircraft according to claim 9 or 10, wherein: The rotation axes of the first rotor, the second rotor, the fourth rotor and the fifth rotor are arranged along the vertical direction.

12. The pure electric aircraft according to any one of claims 9 to 11, wherein: The rotor rotation surface of the third rotor is lower than the rotor rotation surface of any one of the second rotor and the fourth rotor; or / and The rotor rotation surface of the sixth rotor is lower than the rotor rotation surface of any one of the first rotor and the fifth rotor.

13. The pure electric aircraft according to any one of claims 2 to 12, wherein: The frame includes a bracket and six arms, the bracket is arranged on the top of the aircraft body, the six arms are arranged in sequence and spaced apart along the circumference of the bracket, and each arm extends relative to the bracket in a direction away from the center of the bracket; The six non-tilt rotors are arranged in one-to-one correspondence with the six machine arms, and each non-tilt rotor is arranged on a corresponding one of the machine arms.

14. The pure electric aircraft according to claim 13, wherein: Each of the arms is foldably connected to the bracket. The arm can be rotated relative to the bracket to be unfolded to be in an extended state. The arm can be rotated relative to the bracket and overlapped on the outer periphery of the bracket to be in a folded state.

15. The pure electric aircraft according to claim 14, wherein: The six non-tilt rotors include a first rotor, a second rotor, a third rotor, a fourth rotor, a fifth rotor and a sixth rotor, and in the direction in which the six rotors are distributed around, the six rotors are arranged in ascending order in a counterclockwise direction; In the folded state, the two arms corresponding to the third rotor and the sixth rotor are stacked on the side of the four arms corresponding to the first rotor, the second rotor, the fourth rotor and the fifth rotor facing the ground.

16. The pure electric aircraft according to claim 15, wherein: The aircraft body has two side portions on both sides of the forward direction respectively; the six arms include a first arm and a second arm, and the first arm and the second arm are arranged in one-to-one correspondence with the two side portions of the aircraft body; The third rotor is connected to the first arm, and the sixth rotor is connected to the second arm; The two ducted fans include a first ducted fan and a second ducted fan, the first ducted fan is connected to the first machine arm, and the second ducted fan is connected to the second machine arm.

17. The pure electric aircraft according to claim 16, wherein: In a level flight state, the first arm has a top side and a bottom side facing away from each other, the bottom side of the first arm faces the ground, the first ducted fan is installed on the bottom side of the first arm, and the third rotor is installed on the top side of the first arm; the first ducted fan has a rotation center, and the hub of the third rotor has a first maximum rotation surface in a rotating state, and when the aircraft is projected onto a horizontal plane, the projection surface of the first maximum rotation surface covers the rotation center of the first ducted fan; and / or, In the flight state, the second arm has a top side and a bottom side facing away from each other, the bottom side of the second arm faces the ground, the second ducted fan is installed on the bottom side of the second arm, and the sixth rotor is installed on the top side of the second arm; the second ducted fan has a rotation center, and the hub of the sixth rotor has a second maximum rotation surface in the rotating state. When the aircraft is projected onto the horizontal plane, the projection surface of the second maximum rotation surface covers the rotation center of the second ducted fan.

18. A pure electric aircraft according to any one of claims 2 to 17, wherein: In level flight, the angle between the rotation axis of the ducted fan and the horizontal plane is less than or equal to 5 degrees.

19. A pure electric aircraft according to any one of claims 2 to 18, wherein: The foldable rotor includes at least two blades and a hub, and the foldable rotor is the non-tilt rotor; the two blades are connected to the opposite sides of the hub, and the two blades are arranged to rotate relative to each other. The two blades can approach each other to be in a stacked state, or move away from each other to be in an unfolded state.

20. The pure electric aircraft according to any one of claims 1 to 19, wherein: The aircraft further comprises a docking mechanism, which is arranged on a side of the aircraft body away from the forward direction of the aircraft, and is suitable for cooperating with an external traction mechanism so that the aircraft body moves under the traction of the external traction mechanism.

21. The pure electric aircraft according to any one of claims 1 to 20, wherein: The docking mechanism comprises a matching ring body, which is rotatably connected to a side of the chassis of the aircraft body facing the ground, and the matching ring body can rotate linearly along a vertical direction relative to the chassis of the aircraft body.

22. The pure electric aircraft according to any one of claims 1 to 21, wherein: The aircraft further comprises at least two landing gear mechanisms, the two landing gear mechanisms being arranged in parallel and spaced relation on the aircraft body along the forward direction, the landing gear mechanisms comprising: two landing gears, the two landing gears are rotatably connected to the two sides of the aircraft body in the forward direction of the aircraft; and, Two driving members are arranged in one-to-one correspondence with the two landing gears, each of the driving members is connected between a corresponding one of the landing gears and the aircraft body, and the driving member drives the landing gear to rotate away from the aircraft body to unfold and be in a supporting state, or drives the landing gear to rotate in a direction close to the aircraft body to be in a folded state.

23. A flight control method for an aircraft, wherein: The flight control method is used to control the flight of an aircraft, wherein the aircraft comprises an aircraft body, a plurality of non-tilt rotors and two ducted fans, wherein the plurality of non-tilt rotors are used to provide lift, thrust and yaw moment to the aircraft body, and the two ducted fans are respectively located on the left and right sides of the aircraft body; the flight control method comprises: Obtaining a target yaw moment of the aircraft; During the flight of the aircraft, obtaining a first yaw moment generated by the plurality of non-tilt rotors; When a difference between the first yaw moment and the target yaw moment is greater than a specified value, determining a compensating yaw moment according to the difference; The operation of the ducted fan is controlled according to the compensated yaw moment, wherein a difference between a sum of a second yaw moment generated by the operation of the ducted fan and the first yaw moment and the target yaw moment is less than or equal to the specified value.

24. The flight control method according to claim 23, wherein: The step of obtaining a first yaw moment generated by the plurality of non-tilt rotors during the flight of the aircraft comprises: During the flight of the aircraft, the operating parameters of each non-tilt rotor and the fault index parameters of the non-tilt rotor are obtained, wherein the fault index parameters are monitored by the fault self-checking system of the aircraft, and the fault index parameters represent the degree of the fault of the non-tilt rotor; When it is determined according to the working parameter and the fault indicator parameter that at least one of the plurality of non-tilt rotors has a fault, obtaining the plurality of non-tilt rotors The first yaw moment generated by the rotor.

25. The flight control method according to claim 24, wherein: The number of the non-tilt rotors is six, and the six non-tilt rotors are arranged in pairs, and the rotation directions of the two corresponding non-tilt rotors are opposite; When a single non-tilt rotor among the plurality of non-tilt rotors has a fault, the step of determining, based on the operating parameter and the fault indicator parameter, that at least one of the plurality of non-tilt rotors has a fault, obtaining a first yaw moment generated by the plurality of non-tilt rotors comprises: When it is determined according to the operating parameter and the fault indicator parameter that one of the plurality of non-tilt rotors is faulty, the non-tilt rotor corresponding to the faulty non-tilt rotor is controlled to stop working; A first yaw moment generated by the non-tilt rotors other than the non-tilt rotor that fails and the non-tilt rotor corresponding to it is obtained.

26. The flight control method according to any one of claims 23 to 25, wherein: The obtaining of the target yaw moment of the aircraft comprises: Obtaining a current yaw angle and a target yaw angle of the aircraft; Determining a target yaw moment of the aircraft according to the current yaw angle and the target yaw angle; The step of controlling the operation of the ducted fan according to the compensated yaw moment comprises: determining a target power of the ducted fan according to the compensated yaw moment; Determining a target speed of the ducted fan according to the target power and the size and aerodynamic efficiency of the ducted fan; The ducted fan is controlled to operate according to the target rotation speed, so that the ducted fan generates a second yaw moment.

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