Pure electric aircraft and their control methods

Abstract

This application relates to a pure electric aircraft and its flight control method. The pure electric aircraft includes an aircraft body, a frame, multiple non-tilt rotors, and two ducted fans. The aircraft body is equipped with a manned flight cabin. The frame is located on top of the aircraft body. The multiple non-tilt rotors are spaced apart on the frame along a vertical line around the aircraft's center of gravity. When the aircraft body is in flight, the rotation axis of the non-tilt rotors remains fixed relative to the aircraft body. The two ducted fans are connected to the frame and located on the left and right sides of the aircraft body, respectively. When the yaw moment provided by the non-tilt rotors is unbalanced, they provide a compensating yaw moment to the aircraft body. In this embodiment, the pure electric aircraft can provide a compensating yaw moment to the aircraft body through the two ducted fans when the yaw moment formed by the multiple non-tilt rotors is unbalanced, resulting in relatively high flight safety.

Description

Technical Field

[0001] This application relates to the field of aircraft technology, and in particular to a pure electric aircraft and its control method. Background Technology

[0002] With the increasing demands of people's lives and work, flying transportation equipment has been widely used, such as passenger planes, helicopters, and unmanned aerial vehicles (UAVs). Flying transportation equipment has very high requirements for its own flight power system. During flight, all parts of the flight power system need to operate normally to ensure the safety of passengers and flying transportation equipment.

[0003] However, existing air transport equipment may experience yaw moment imbalance when the flight propulsion system is underpowered or some equipment malfunctions, causing the air transport equipment to be unable to yaw. Existing air transport equipment cannot overcome the problem of being unable to yaw due to yaw moment imbalance. Summary of the Invention

[0004] This application provides a pure electric aircraft and its control method.

[0005] In a first aspect, this 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 cockpit for carrying passengers. The frame is located on the top of the aircraft body. The multiple non-tilt rotors are connected to the frame and are spaced apart around the vertical line where the center of gravity of the aircraft is located. When the aircraft body is in flight, the rotation axis of the non-tilt rotors is always fixed relative to the aircraft body. The two ducted fans are connected to the frame and are located on the left and right sides of the aircraft body, respectively. When the yaw moment provided by the non-tilt rotors is unbalanced, they provide a compensating yaw moment to the aircraft body.

[0006] Secondly, this application also provides a flight control method for an aircraft. The flight control method is used to control the flight of the aircraft, which includes an aircraft body, multiple non-tilt rotors, and two ducted fans. The multiple non-tilt rotors provide lift, thrust, and yaw torque to the aircraft body, and the two ducted fans are located on the left and right sides of the aircraft body, respectively. The flight control method includes: acquiring a target yaw torque of the aircraft; acquiring a first yaw torque generated by the multiple non-tilt rotors during the flight of the aircraft; determining a compensating yaw torque based on the difference if the difference between the first yaw torque and the target yaw torque is greater than a specified value; and controlling the operation of the ducted fans based on the compensating yaw torque, wherein the difference between the sum of the second yaw torque and the first yaw torque generated by the operation of the ducted fans and the target yaw torque is less than or equal to a specified value.

[0007] This 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 equipped with a cockpit for carrying passengers. The frame is located on top of the aircraft body. The multiple non-tilt rotors are connected to the frame and are spaced apart around the vertical line where the aircraft's center of gravity is located. When the aircraft body is in flight, the rotation axis of the non-tilt rotors remains fixed relative to the aircraft body. The two ducted fans are connected to the frame and are located on the left and right sides of the aircraft body, respectively. When the yaw moment provided by the non-tilt rotors is unbalanced, they provide a compensating yaw moment to the aircraft body. To address the imbalance of yaw moments generated by multiple non-tilt rotors, such as yaw moment imbalance or at least one non-tilt rotor malfunctioning, the aircraft in this embodiment also includes two ducted fans connected to the aircraft body. When the yaw moments generated by multiple non-tilt rotors are unbalanced, the ducted fans can operate and provide compensating yaw moments to the aircraft body, enabling the aircraft body to overcome the problem of being unable to yaw due to yaw moment imbalance. Attached Figure Description

[0008] To more clearly illustrate the technical solution of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0009] Figure 1 This is a schematic diagram of the structure of the flying car provided in the embodiments of this application.

[0010] Figure 2 yes Figure 1 The diagram shows the decoupling structure of the flying car and the aircraft.

[0011] Figure 3 This is a schematic diagram of the vehicle structure provided in the embodiments of this application.

[0012] Figure 4 This is a schematic diagram of the structure of the aircraft provided in the embodiments of this application.

[0013] Figure 5 yes Figure 4 The diagram shows the structure of the automatic docking device for the aircraft.

[0014] Figure 6 yes Figure 4 The diagram shows the structure of the aircraft with its arms folded and landing gear mechanism deployed.

[0015] Figure 7yes Figure 1 The diagram shows the structure of the flying car and the aircraft just after docking.

[0016] Figure 8 yes Figure 7 The diagram shows the structure of the aircraft moving towards the support platform.

[0017] Figure 9 yes Figure 7 The diagram shows the structure in which the center of gravity of the aircraft is transferred to the support platform.

[0018] Figure 10 yes Figure 1 The diagram shows a top-view structural diagram of the flying car's aircraft.

[0019] Figure 11 yes Figure 10 The diagram shows the structure of the rotor rotation surface formed by the foldable rotor of the aircraft.

[0020] Figure 12 yes Figure 10 The diagram shows the structural relationship between the foldable rotor and the aircraft body.

[0021] Figure 13 yes Figure 10 The diagram shows the structure of the aircraft with its arms folded.

[0022] Figure 14 yes Figure 10 The diagram shows the structure of the aircraft with its arms deployed.

[0023] Figure 15 yes Figure 10 The diagram shows the structure of the aircraft with its foldable rotors deployed.

[0024] Figure 16 yes Figure 10 The diagram shows the structure of the aircraft with its foldable rotor in a folded state.

[0025] Figure 17 yes Figure 10 The diagram shows the structure of the aircraft with its foldable rotors deployed.

[0026] Figure 18 This is a flowchart illustrating a flight control method for an aircraft provided in one embodiment of this application.

[0027] Figure 19 This is a functional framework diagram of a flight control device disclosed in an embodiment of this application. Detailed Implementation

[0028] To enable those skilled in the art to better understand the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are merely some embodiments of the present application, and not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present application without creative effort are within the scope of protection of the present application.

[0029] Please see Figure 1 and Figure 2 This embodiment provides a flying car 1000, which includes a land transportation device like a car, a flying transportation device like an airplane, or a transportation device like a flying car that combines land and flight functions. This application embodiment provides a flying car 1000 with both land and flight functions, comprising a vehicle 100 and an aircraft 200, which are detachably coupled.

[0030] Please see Figure 3 In this embodiment, the vehicle 100 is used to travel on the ground, enabling the flying car 1000 to have the function of traveling on the ground. The vehicle 100 includes a vehicle body 110, multiple wheel sets 120, and an automatic docking device 130. The vehicle body 110 is the main structure of the vehicle 100, bearing the load of the vehicle 100 and other functional equipment. The vehicle body 110 includes a land-based cockpit 1110 and a carrying platform 1130 connected together. The land-based cockpit 1110 is provided with seats for passengers, and the carrying platform 1130 is used to park the aircraft 200. Multiple wheel sets 120 are set on the chassis of the vehicle body 110 to enable the vehicle body 110 to travel on land. The automatic docking device 130 is set on the carrying platform 1130 of the vehicle body 110, and the automatic docking device 130 is used to dock with the aircraft 200 and tow the aircraft 200.

[0031] Please see Figure 4In this embodiment, the aircraft 200 is used to travel in the air, enabling the flying car 1000 to have the function of traveling in the air. The aircraft 200 and the vehicle 100 together construct the flying car 1000 with land-walking and flight functions. The aircraft 200 includes an aircraft body 210 and multiple foldable rotors 220 disposed on the aircraft body 210. The aircraft body 210 is the main structure of the aircraft 200, and a flight cockpit 2110 is disposed on the aircraft body for passengers to ride in. The multiple foldable rotors 220 are disposed on the aircraft body to provide the aircraft 200 with the power to travel in the air. The aircraft 200 can take off and travel in the air under the drive of the multiple foldable rotors 220. The aircraft 200 may also include two ducted fans 230 connected to the aircraft body 210. The two ducted fans 230 are used to provide a compensating yaw moment to the aircraft body 210 when the yaw moment formed by the multiple foldable rotors 220 is in an unbalanced state.

[0032] In traditional flying cars, the aircraft 200 requires a direct vertical landing on the vehicle 100's platform 1130 during the landing process. On one hand, the manned aircraft 200 is large and heavy, making precise control of its attitude and movement difficult. On the other hand, the overall size of the flying car 100 is limited by existing road infrastructure, resulting in a relatively small landing surface on the vehicle 100's platform 1130. This makes it difficult for the aircraft 200 to land on the platform 1130 quickly, and the rotating rotor of the aircraft 200 poses a risk of collision with the vehicle 100, potentially damaging both the aircraft and vehicle structures. Even if it lands on the platform 1130, the aircraft 200's position may not meet the coupling requirements. After the aircraft 200 lands on the support platform 1130, adjusting the relative positions of the aircraft 200 and the vehicle 100 is extremely difficult. It not only has to overcome the weight of the aircraft 200, but also has to consider whether the structure of the support platform 1130 can withstand the repeated deflection and movement of the aircraft 200 on the support platform 1130, and also has to avoid structural collisions between the aircraft 200 and the vehicle 100. Therefore, adjusting the position of the aircraft 200 on the support platform 1130 is extremely difficult, resulting in a low user experience.

[0033] This embodiment of the application provides a new coupling method for the aircraft 200 and the vehicle 100 by setting up an automatic docking device 130. Specifically, the aircraft 200 switches from flight mode to ground-based mode. Subsequently, the vehicle 100 travels to the vicinity of the aircraft 200 and aligns its position with the aircraft 200 to drive the automatic docking device 130 to automatically connect to the aircraft body 210 and drive the aircraft body 210 to move towards the vehicle body 110 until the aircraft body 210 is detachably coupled to the support platform 1130. When the aircraft 200 needs to take off, the coupling relationship between the aircraft body 210 and the support platform 1130 is first decoupled before takeoff. After decoupling, the aircraft 200 first docks on the ground and then takes off from the ground to ascend into the air, while the vehicle 100 remains on the land. Therefore, in this embodiment, the aircraft 200 does not require precise positioning with the carrier platform 1130 and can quickly land on the ground. Subsequently, the vehicle 100 moves to a position near the aircraft 200 and aligns with it, allowing the automatic docking device 130 to dock with the aircraft 200 and pull it onto the carrier platform 1130. The control difficulty of coupling the aircraft body 210 and the vehicle 100 is relatively low, increasing the coupling speed between the aircraft body 210 and the carrier platform 1130 and facilitating the adjustment of the relative positions of the aircraft 200 and the vehicle 100. During the coupling and decoupling processes of the aircraft 200 with the carrier platform 1130, the foldable rotor 220 stops rotating, improving the safety of the flying car 1000 during use.

[0034] In this embodiment, the degree of enclosure of the aircraft 200 by the carrier platform 1130 is not limited; the aircraft 200 can be partially or fully enclosed. In this embodiment, the shape of the vehicle 100 that can partially enclose the aircraft 200 is similar to that of an existing pickup truck. Specifically, the carrier platform 1130 also includes a carrier plate 1131 and a fence structure 1132. The carrier plate 1131 is connected to the land-based cockpit 1110, and the fence structure 1132 is arranged around the outer periphery of the carrier plate 1131. When the aircraft 200 is coupled to the carrier platform 1130, it can prevent the aircraft 200 from leaving the carrier plate 1131 in the horizontal direction. The fence structure 1132, the land-based cockpit 1110, and the carrier plate 1131 together define a first accommodating space 1133 that communicates with the outside. 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 shown in the figure). The movable fence is located on the side of the support plate 1131 away from the land-going cockpit 1110. The movable fence can rotate relative to the support plate 1131 to provide a passage for the aircraft body 210 to enter the support platform 1130.

[0035] In this embodiment, the vehicle 100, which can completely surround the aircraft 200, has an appearance similar to a conventional van. Specifically, the carrying platform includes a carrying plate 1131 and a cargo box (not shown in the figure). The carrying plate 1131 is connected to the land-going cabin 1110, and the cargo box covers the carrying plate 1131. The cargo box and the carrying plate 1131 together define a second accommodating space (not shown in the figure) that is isolated from the outside. In order for the aircraft 200 to enter and exit the second accommodating space, the cargo box includes a door, which is movably disposed on the side of the cargo box away from the land-going cabin. The door is used to open the cargo box to avoid the aircraft body 210.

[0036] In this embodiment, the vehicle 100 also includes a vehicle control system 140, which controls the wheel assembly 120. Under the vehicle movement commands of the vehicle control system 140, the wheel assembly 120 can change the motion state of the vehicle body 110, such as accelerating, decelerating, moving forward, or moving backward. Similarly, the aircraft 200 also includes a flight control system 240, which controls the foldable rotor 220. Under the aircraft movement commands of the flight control system 240, the foldable rotor 220 can change the motion state of the aircraft body 210, such as taking off, landing, yawing, or hovering. Furthermore, the aircraft 200 in this embodiment can travel on land, enabling it to run and cushion its impact.

[0037] Please see Figure 3 and Figure 4 In this embodiment, the vehicle control system 140 is located in the land-based cockpit 1110, which can be understood as both a passenger space and a driving space. The driver can operate the vehicle control system 140 within the land-based cockpit 1110 to operate the vehicle 100. The vehicle control system 140 may be one or more of the following structures: a driver's instrument panel, steering wheel, accelerator, brake, gear lever, etc. The aircraft control system 240 is located in the flight cockpit 2110, and includes one or more of the following structures: a flight instrument panel, a central control console, and an aircraft joystick. In some embodiments, the vehicle control system 140 and the flight control system 240 can communicate with each other to transmit motion data or motion status information of the vehicle 100 and the aircraft 200. In some other embodiments, vehicle 100 is an autonomous transportation device, therefore the location of vehicle control system 140 is not specifically limited. It can be located on the vehicle body 110 outside the ground-based cockpit 1110, or it can be located inside the ground-based cockpit 1110 to assist passengers in driving vehicle 100. If aircraft 200 is also an autonomous transportation device, aircraft 200 has the same configuration as vehicle 100, which will not be described further here.

[0038] The following section describes the specific structures for docking and coupling of vehicle 100 and aircraft 200, and explains how vehicle 200 and aircraft 200 achieve docking and coupling through these specific structures.

[0039] Please see Figure 5 In this embodiment, the automatic docking device 130 includes a traction mechanism 1310 and a linear motion mechanism 1320. The traction mechanism 1310 is movably mounted on the support platform 1130. When the vehicle 100 controls the support platform 1130 to correspond with the aircraft body 210, the movement trend of the traction mechanism 1310 is to move closer to or further 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 support platform 1130, and the traction mechanism 1310 is connected to the linear motion mechanism 1320. The linear motion mechanism 1320 drives the traction mechanism 1310 to move on the support platform. Correspondingly, the aircraft 200 also includes a docking mechanism 250 connected to the aircraft body 210, which is adapted to cooperate with the traction mechanism 1310. After the traction mechanism 1310 and the docking mechanism 250 are connected, the traction mechanism 1310 pulls the docking mechanism 250 to pull the aircraft 200 onto the carrying platform 1130.

[0040] Please see Figure 5 First, the specific components of the linear motion structure 1320 are introduced to reveal the mechanism by which the linear motion structure 1320 drives the traction mechanism 1310. Specifically, the linear motion mechanism 1320 includes a lead screw 1321, a rotary motor 1322, and a slider 1323. The lead screw 1321 is arranged along the forward direction of the vehicle 100. The rotary motor 1322 is driven and connected to one end of the lead screw 1321. When the rotary motor 1322 is turned on, it can drive the lead screw 1321 to rotate. The slider 1323 is threadedly connected to the lead screw 1322 and is connected to the traction mechanism 1310. When the support platform 1130 moves to a position opposite to the aircraft body 210, the lead screw 1321 rotates, and the slider 1323, along with the traction mechanism 1310, moves closer to or further away from the aircraft body 210 relative to the support platform 1130. In this embodiment, there are two linear motion mechanisms 1320. The two linear motion mechanisms 1320 and the traction mechanism 1310 are arranged in the horizontal direction to avoid increasing the height of the support platform 1130 and to prevent the overall structure of the flying car 1000 from being too tall 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 makes the driving force acting on the support mechanism 1310 greater, and on the other hand, it can keep the support structure 1310 in a roughly balanced state, making it less prone to tilting or deflection and ensuring the traction effect.

[0041] Please see Figure 5 Next, the specific structure of the traction mechanism 1310 and the specific structure of the docking mechanism 250 of the aircraft 200 will be described. 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 in the figure). The traction ring 1312 is disposed on the side of the traction body 1311 facing away from the land-based cockpit 1110, so that when the carrying platform 1130 moves to a position opposite to the aircraft body 210, the traction ring 1312 and the mating ring are opposite each other and can be engaged with each other. In some other embodiments, the traction mechanism 1310 and the docking mechanism 250 may also be hook-shaped structures.

[0042] In order to enable the traction ring 1312 and the mating ring to engage, in this embodiment, the traction ring 1312 is partially exposed on the side of the traction body 1311 away from the land-going cabin 1110. The traction ring 1312 is provided with a notch 1313. The traction ring 1312 is rotatably mounted on the traction body 1311 and can rotate around its axis. When the traction ring 1312 and the mating ring are positioned opposite each other, the traction ring 1312 can rotate and make the notch correspond to the mating ring, so that the mating ring and the traction ring 1312 can engage. Then the traction ring 1312 continues to rotate, so that the notch is hidden inside the traction body 1311, ensuring the stability of the engagement relationship between the mating ring and the traction ring 1312.

[0043] Please see Figure 5 After the aircraft 200 is towed onto the carrier platform 1130, and before the aircraft body 210 is coupled to the carrier platform 1130, the movement direction of the aircraft body 200 is kept as straight as possible relative to the vehicle body 110. Therefore, the automatic docking device 130 in this embodiment also includes two guide grooves 1330. Both guide grooves 1330 are arranged horizontally on the carrier platform 1130, and the guide grooves 13430 extend along the forward direction of the vehicle 100. Correspondingly, mating wheel sets (not shown in the figure) are provided on opposite sides of the bottom of the aircraft body 210. When the carrier platform 1130 moves to a position opposite to the aircraft body 210, the two guide grooves 1330 and the two mating wheel sets are arranged opposite each other. Each mating wheel set can slide into a corresponding guide groove 1330 and roll within the guide groove 1330 to limit the movement trajectory of the aircraft body 210.

[0044] Please see Figure 5In order to facilitate the sliding of the mating wheelset into the corresponding guide groove 1330, the guide groove 1330 in this embodiment includes a first groove 1331 and a second groove 1332 that are connected. The second groove 1332 is connected to the end of the first groove 1331 that is away from the land cabin 1110. The groove width of the second groove 1332 gradually increases in the direction away from the land cabin 1110, that is, the groove width of the second groove 1332 near the mating wheelset is wider, so that the mating wheelset can slide smoothly into the second groove 1332.

[0045] After the mating wheel assembly 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 assembly is not necessarily consistent with the extension direction of the guide groove 1330, which may cause the mating wheel assembly to get stuck in the guide groove 1330. Therefore, in this embodiment, the mating ring is rotatably connected to the side of the aircraft body 210's chassis facing the ground. The mating ring can rotate in a straight line in the vertical direction relative to the chassis of the aircraft body 210. During the process of pulling the aircraft body 210, the traction ring 1312 adjusts the movement direction of the mating wheel assembly through the mating ring, so that the mating wheel assembly and the guide groove 1330 are smoothly mated.

[0046] Please see Figure 3 In this embodiment, the automatic docking device 130 further includes a limiting mechanism 1340. The limiting mechanism 1340 is used when the aircraft body 210 is coupled to the carrier platform 1130. The limiting mechanism 1340 connects between the aircraft body 210 and the carrier platform 1130 to restrict the aircraft body 210 from moving relative to the carrier platform 1130 in any direction. Specifically, the limiting mechanism 1340 in this embodiment includes a horizontal limiting component (not shown in the figure). The horizontal limiting component can be connected between the carrier platform 1130 and the aircraft body 210, or between the vehicle body 110 and the aircraft body 210. The horizontal limiting component is used to restrict the aircraft body 210 from moving relative to the carrier platform 1130 in the horizontal direction, for example, restricting 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 component (not shown in the figure). The vertical limiting component is connected between the support platform 1130 and the aircraft body 210, and is used to limit the movement of the aircraft body 210 relative to the support platform 1130 in the vertical direction. The limiting mechanism 1340 stabilizes the position of the aircraft body 210 on the support platform 1130, ensuring the stability of the fixed connection between the aircraft body 210 and the support platform 1130. The limiting mechanism 1340 restricts the movement of the aircraft body 210 on the support platform 1130, which can reduce the impact damage to the aircraft 200 itself and the vehicle 100 caused by the movement of the aircraft body 210.

[0047] This embodiment does not limit the type and specific structure of the horizontal and vertical limiting components; they can be hooks that connect together or blocks that restrict movement. In this embodiment, an electrical limiting mechanism 1340 can be used. When the aircraft body 210 and the carrier platform 1310 are coupled, the electrical limiting mechanism connects the aircraft body 210 to restrict its movement. Before the aircraft body 210 needs to take off, the electrical limiting mechanism disconnects the carrier platform 1310 or vehicle body 110 from the aircraft body 210.

[0048] Please see Figure 3 To reduce the damage to the aircraft 200 and vehicle 100 caused by vibrations of the aircraft body 210, the automatic docking device 130 in this embodiment also includes a buffer mechanism 1350. The buffer mechanism 1350 is disposed on the support platform 1130. When the aircraft body 210 and the support platform 1130 are coupled, the buffer mechanism 1350 is located between the aircraft body 210 and the support platform 1130, and also between the aircraft body 210 and the vehicle body 110. The buffer mechanism 1350 can mitigate the impact damage to the vehicle body 110 and the aircraft body 210 caused by shaking of the aircraft body 210. This embodiment does not specifically limit the type and specific structure of the buffer mechanism 1350; it can be a damper or a hydraulic buffer.

[0049] Please see Figure 4 and Figure 6 In this embodiment, the aircraft 200 also includes a landing gear mechanism 260, which supports the aircraft body 210 when the aircraft 200 is in a stationary state. 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 towards the aircraft body 210 to a folded state, which can reduce the drag experienced by the aircraft body 210 in the air. When the aircraft 200 is about to land on the ground, the landing gear mechanism 260 can rotate relative to the aircraft body 210 to a supported state and can be supported on the ground. In other embodiments, structures such as airbags or blocks can also be used to achieve the landing of the aircraft body 210.

[0050] Please see Figure 7To ensure relative balance between the left and right sides of the aircraft body 210 along the forward direction of the aircraft 200 when the aircraft body 210 is in a stationary 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 rotate relative to the aircraft body 210 to unfold or fold. In other embodiments, the landing gear mechanism 260 can be a single, integral structure. This integral landing gear mechanism 260 can simultaneously support both left and right sides of the aircraft body 210 when the aircraft 200 is in a stationary state.

[0051] Please see 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. Each drive member 2630 is correspondingly configured with one landing gear 2610, and each drive member 2630 connects the aircraft body 210 to its corresponding landing gear 2610. The drive member 2630 can automatically drive the landing gear 2610 to rotate relative to the aircraft body 210 according to the flight status of the aircraft 200, so that the landing gear 2610 is in a folded or deployed state relative to the aircraft body 210, without manual intervention, thus improving safety. In this embodiment, the drive member 2630 can be a hydraulic cylinder or a linear motor, etc.

[0052] Please see Figure 4 In this embodiment, the landing gear 2610 includes a support rod 2611 and a support foot 2612. The support rod 2611 is connected between the support foot 2612 and the aircraft body 210. When the landing gear mechanism 260 is in a supported state, the support foot 2612 abuts against the ground and supports the aircraft body 210 upwards. The support foot 2612 is a rod structure that extends approximately along the forward direction of the aircraft 200. At least one of the front and rear ends of the support foot 2612 in the forward direction of the aircraft 200 is raised away from the ground. The raised part forms an arc edge 2613 on the side of the support foot 2612 facing the ground. During the landing process of the aircraft 200, the support foot 2612 can adapt to the ground through the arc edge 2613, and can avoid foreign objects on the ground from colliding violently with the support foot 2612, avoid abnormal vibration of the aircraft body 210, and prevent damage to the support foot 2612 and the aircraft body 210.

[0053] After the traction mechanism 1310 on the carrier platform 1130 and the docking mechanism 250 on the aircraft body 210 are docked, the traction mechanism 1310 pulls the aircraft body 210 toward the carrier platform 1130 until the aircraft body 210 is coupled to the carrier platform 1130. Therefore, the process from the aircraft body 210 being independent of the carrier platform 1130 to being coupled to the carrier platform 1130 is a gradual process. Before the aircraft body 210 has moved to the carrier platform 1130, or before the center of gravity of the aircraft body 210 has moved to a safe position on the carrier platform 1130, a landing gear mechanism 260 still needs to be set 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. These two sets of landing gear mechanisms 260 are spaced apart along the forward direction of the aircraft 200 on the aircraft body 210. One set of landing gear mechanisms 260 is closer to the support platform 1130, and the other is farther from the support platform 1130. That is, one set of landing gear 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 260 may include two landing gears 260, 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 depending on the weight and length of the aircraft body 210.

[0054] In this embodiment, during the process of the aircraft body 210 going from being independent of the support platform 1130 to being coupled to the support platform 1130, the two landing gear mechanisms 260 fold sequentially. The folding timing of the two landing gear mechanisms 260 will be described in detail below.

[0055] Please see Figure 7 and Figure 8 Initially, both landing gear mechanisms 260 are in a supported state. Under the towing of the traction mechanism 1310, a small portion of the aircraft body 210 moves onto the support platform 1130 and receives support from the support platform 1130. When the support platform 1130 and one landing gear mechanism 260 away from the support platform 1130 are able to overcome the weight of the aircraft body 210 and bring it into a balanced state, the landing gear mechanism 260 closer to the support platform 1130 folds and retracts to switch to a folded state.

[0056] Please see Figure 9The traction mechanism 1310 continues to traction the aircraft body 210 toward the support platform 1130. When most of the structure of the aircraft body 210 has moved to the support platform 1130, and the center of gravity of the aircraft body 210, the support platform 1130, and the vehicle body 110 reach a balanced state, that is, when the support platform 1310 can independently support the aircraft body 210 without the assistance of a landing gear mechanism 260 far away from the support platform 1310, the landing gear mechanism 260 far away from the support platform 1310 folds and retracts to switch to the folded state.

[0057] Finally, the traction mechanism 1310 continues to traction the aircraft body 210 toward the carrier platform 1130 until the aircraft body 210 moves to the designated coupling position on the carrier platform 1130, and then the aircraft body 210 couples with the carrier platform 1130.

[0058] Please refer to it again. Figure 7 In this embodiment, the vehicle 100 is a pure electric vehicle. The vehicle 100 also includes a vehicle electrical system 170, which is disposed on the vehicle body 110. The vehicle electrical system 170 is electrically connected to multiple wheel sets 120 to provide power to the multiple wheel sets 120, enabling the multiple wheel sets 120 to drive the vehicle body 110 on land. In other embodiments, the vehicle 100 can also be a range-extended electric vehicle, that is, in addition to the vehicle electrical system 170, the vehicle 100 also includes a fuel generator. Similarly, in this embodiment, the aircraft 200 is a pure electric aircraft or a pure electric manned helicopter. The aircraft 200 also includes a flight electrical system 270, which is disposed on the aircraft body 210. The flight electrical system 270 is electrically connected to multiple foldable rotors 220 to provide power to the multiple foldable rotors 220, enabling the foldable rotors 220 to drive the aircraft body 210 in the air.

[0059] Specifically, in this embodiment, the vehicle electrical system 170 includes a first power battery pack 1710, which is disposed on the vehicle body 110. The first power battery pack 1710 is used to store electrical energy to provide power to the land propulsion system 120 and other electrical components on the vehicle 100. The flight electrical system 270 includes a second power battery pack 2710, which is disposed on the aircraft body 210. The second power battery pack 2710 is used to store electrical energy to provide power to the foldable rotor 220 and other electrical components on 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 acidic lithium batteries or alkaline zinc-manganese batteries.

[0060] In some embodiments, the vehicle electrical system 170 and the flight electrical system 270 are independent of each other, and no power transfer occurs 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, but when the aircraft body 210 is coupled to the carrier platform 1310, power transfer can be achieved between the first power battery pack 1710 and the second power battery pack 2710. Specifically, the vehicle electrical system 170 also includes a first connector 1720 electrically connected to the first power battery pack 1710, and the flight electrical system 270 also includes a second connector 2720 electrically connected to the second power battery pack 2710. When the aircraft body 210 and the carrier platform 1310 are coupled, the first connector 1720 and the second connector 2720 are electrically connected to allow electrical energy to be transferred between the first power battery pack 1710 and the second power battery pack 2710. This allows the first power battery pack 1710 to be replenished by the second power battery pack 2710, or vice versa. Additionally, in this embodiment, both the first power battery pack 1710 and the second power battery pack 2710 are provided with power connection ports to receive electrical energy replenishment from an external power source of the transport vehicle 1000.

[0061] First, the process of the first power battery pack 1710 replenishing power to the second power battery pack 2710 is described. After the aircraft 200 finishes flight and is coupled to the carrier platform 1130, the first connector 1720 connected to the first power battery pack 1710 can replenish power to the second power battery pack 2710 through the second connector 2720, so that the flight electrical system 270 stores sufficient power for the aircraft 200's next flight. Specifically, the vehicle electrical system 170 also includes an energy management module 1730, which is electrically connected to the first power battery pack 1710. When the energy management module 1730 detects that the aircraft body 210 and the carrier platform 1310 are coupled, the energy management module 1730 can control the first power battery pack 1710 to replenish power to the second power battery pack 2710.

[0062] Next, the method of the second power battery pack 2710 replenishing power to the first power battery pack 1710 is described. The flying car 1000 in this embodiment also includes a power sensor (not shown in the figure) and a vehicle controller (not shown in the figure), with the power sensor and vehicle controller electrically connected. The power sensor is used to detect the remaining power in the first power battery pack 1710 and the second power battery pack 2710, and the vehicle controller is used to control 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 in motion, if the power sensor detects that the remaining power in the first power battery pack 1710 is low, 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 manifestations of insufficient power in the first power battery pack 1710.

[0063] In the first specific scenario, the power sensor detects the remaining power in the first power battery pack 1710. The vehicle controller estimates the driving range of vehicle 100 based on the remaining power in the first power battery pack 1710. If the driving range of vehicle 100 is lower than the target driving range, it indicates that the power in the first power battery pack 1710 is insufficient to support the flying car 1000 to reach the target location. In this case, it can be determined that the remaining power in the first power battery pack 1710 is insufficient. Then, the power sensor 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, that is, when the power in the second power battery pack 2710 is sufficient, the vehicle controller controls the second power battery pack 2710 to replenish the power to the first power battery pack 2710.

[0064] In a second specific embodiment, the flying car 1000 also includes an acceleration sensor (not shown in the figure). 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 not enough power to enable the vehicle 100 to reach the preset driving speed within a preset time. Next, the power sensor 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, that is, when the power in the second power battery pack 2710 is sufficient, the vehicle controller controls the second power battery pack 2710 to replenish the power to the first power battery pack 1710.

[0065] Both scenarios where the second power battery pack 2710 replenishes power to the first power battery pack 1710 require sufficient remaining charge in the second power battery pack 2710; that is, the remaining charge in the second power battery pack 2710 must be greater than or equal to a second preset charge level. In special circumstances, even if the remaining charge in the second power battery pack 2710 is less than the second preset charge level, the second power battery pack 2710 can still replenish power to the first power battery pack 1710. However, when the charge in the second power battery pack 2710 drops to a preset minimum charge level, the transfer of power to the first power battery pack 1710 stops, and the first power battery pack 1710 needs to find other power sources to replenish its power.

[0066] Please see Figure 7 This embodiment does not limit the positional arrangement of the first connector 1720 and the second connector 2720, but makes specific arrangements based on the actual orientation between the vehicle body 110 and the aircraft body 210. As mentioned earlier, in this embodiment, the vehicle body 110 and the carrier platform 1130 are arranged side by side in the forward direction of the vehicle 100. When the aircraft body 210 and the carrier platform 1130 are in a coupled state, 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 frontal orientation of the aircraft body 210 is opposite to and approximately parallel to the frontal orientation of the vehicle body 110. Therefore, in this embodiment, the first connector 1720 is arranged on the side of the vehicle body 110 facing the aircraft body 210, and correspondingly, the second connector 2720 is arranged on the side of the aircraft body 210 facing the vehicle body 110, so that when the aircraft body 210 and the carrier platform 1310 are in a coupled state, the first connector 1720 and the second connector 2720 can be arranged relative to each other and facilitate connection.

[0067] Specifically, in this embodiment, the vehicle body 110 has a first backplate 1120 facing the aircraft body 210, i.e., the first backplate 1120 is disposed facing the support platform 1310, and a first connector 1720 is disposed on the first backplate 1120 to correspond to the support platform 1310 and the aircraft body 210. Correspondingly, the aircraft body 210 has a second backplate 2120 facing the vehicle body 110, and a second connector 2720 is disposed on the second backplate 2120 to correspond to the vehicle body 110. When the aircraft body 210 and the support platform 1310 are coupled, the first backplate 1120 and the second backplate 2120 are disposed opposite to each other, and the first connector 1720 and the second connector 2720 are also disposed opposite to each other and electrically connected to realize the mutual transmission of electrical energy between the first power battery pack 1710 and the second power battery pack 2710.

[0068] This embodiment does not specifically limit the position 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 inside the vehicle body 110. This ensures that the relatively heavy first power battery pack 1710 has a stable state, guaranteeing the power supply to the vehicle 100 and the stability of the vehicle body 110 structure. A first high-voltage wiring harness 1740 is provided between the first power battery pack 100 and the first connector 1720. The first high-voltage wiring harness 1740 extends along the structure of the chassis of the vehicle body 110 and the first back plate 1120 and is electrically connected between the first power battery pack 1710 and the first connector 1720 to provide conditions for power transmission. Similarly, the second power battery pack 2710 is fixedly mounted on the chassis of the aircraft body 210 and located inside the aircraft body 210. A second high-voltage wiring harness 2730 is provided between the second power battery pack 2710 and the second connector 2720. The second high-voltage wiring harness 2730 extends along the structure of the chassis of the aircraft body 210 and the second backplate 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 can be located in the land cockpit 1110, and the second power battery pack 2710 can be located in the flight cockpit 2110.

[0069] Please refer to it again. Figure 3 In this embodiment, the multiple wheel sets 120 include a first wheel set 1210, a second wheel set 1220, and a third wheel set 1230, as well as a vehicle drive motor. The first wheel set 1210, second wheel set 1220, and third wheel set 1230 are all located on the side of the vehicle body 110 facing the ground. At least one of the first wheel set 1210, second wheel set 1220, and third wheel set 1230 is driven by the vehicle drive motor. In some embodiments, the number of wheel sets may also be four, five, or six, etc. The number of wheel sets 120 is determined based on a comprehensive consideration of the overall weight of the flying car 1000 and the load-bearing capacity of the wheel sets 120. In this embodiment, each wheel set 120 refers to a wheel set sharing a common axle. Each wheel set 120 may include multiple wheels, such as two wheels, four wheels, six wheels, eight wheels, etc. The multiple wheels in each wheel set 120 are symmetrically arranged at opposite ends of the same axle.

[0070] In this embodiment, the positional allocation of each wheel set 120 is determined based on the weight borne by each part of the vehicle 100 when the aircraft 200 and vehicle 100 are coupled. Specifically, the first wheel set 1210 is correspondingly arranged with the land-going cockpit 1110 of the vehicle body 110 to support the land-going cockpit 1110, and the second wheel set 1220 and the third wheel set 1230 are correspondingly arranged with the support platform 1130 to support the support platform 1130. In other embodiments, the first wheel set 1210, the second wheel set 1220, and the third wheel set 1230 may be equally spaced at the bottom of the vehicle body 110.

[0071] Please refer to it again. Figure 4 In this embodiment, the aircraft 200 also includes a frame 2210. The frame 2210 is disposed on the aircraft body 210 and serves as a connection medium between the aircraft body 210 and the multiple foldable rotors 220. The multiple foldable rotors 220 are disposed on the frame 2210 and are used to provide lift, thrust, and yaw moment to the aircraft body 210, enabling the aircraft body 210 to perform takeoff, landing, forward movement, yaw, and other maneuvers. In this embodiment, the multiple foldable rotors 220 are selected with the same output power. The multiple foldable rotors 220 with the same rated output power place relatively low demands on the overall structure of the aircraft 200, 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 various rated output powers. The advantage is that it is possible to develop more functional flight modes. For example, in such flight modes, the aircraft 200 can fly in harsh environments and weather conditions, improving the adaptability of the aircraft 200 to the flight environment and making the aircraft 200 more competitive in the market.

[0072] In this embodiment, the foldable rotor 220 maintains a stable structure relative to the aircraft body 210, which simplifies the installation structure of the foldable rotor 220. The fixed position of the foldable rotor 220 is specifically manifested in that the foldable rotor 220 has a rotation axis, which is the rotation axis of the rotor of the foldable rotor 220. During flight, the rotation axis of the foldable rotor 220 remains fixed relative to the aircraft body 210. Therefore, in this embodiment, the foldable rotor 220 can also be referred to as a non-tilt rotor.

[0073] In the level flight state of the aircraft 200, 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 exactly the same. For ease of understanding, the setting of the angle range of the rotation axis of the foldable rotor will be described in detail in some specific usage environments later. It will not be discussed in detail here.

[0074] In order to improve the ability of the aircraft 200 to respond to emergencies during flight and to improve the safety of the aircraft 200 during flight, this embodiment makes specific settings in terms of the number of foldable rotors 220, the distribution of foldable rotors 220 on the aircraft body 210, the structure of foldable rotors 220, and the compensation mechanism for abnormal situations of foldable rotors 220, which will be described in detail below.

[0075] Please see Figure 10 In this embodiment, a vertical plane defined by the forward direction of the aircraft body 210 serves as the central axis of the aircraft body 210. The aircraft body 210 is symmetrical about this central axis, and the two sides of the aircraft body 210 in its forward direction are defined as the left and right sides. The weights of the left and right sides of the aircraft body 210 are approximately the same or similar, ensuring that the two parts of the aircraft body 210 are essentially in a balanced state. The number of foldable rotors 220 can be either even or odd. In this embodiment, the number of foldable rotors 220 is set to even, and multiple foldable rotors 220 are evenly distributed on the left and right sides of the aircraft body 210, meaning that the number of foldable rotors 220 on 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, the configuration can be made according to the structural features of the aircraft body 210. For example, if the aircraft 200 is a fuel-powered or range-extended 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 according to the specific location of the fuel storage device on the aircraft 200 and the influence of the weight change of the fuel storage device during flight on the center of gravity of the aircraft body 210.

[0076] The specific number of foldable rotors 220 needs to 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 the actual situation. In traditional multi-rotor manned aircraft, although it is relatively easy to arrange six rotors, if one of the six rotors malfunctions, there will be a yaw loss of control problem. Therefore, people choose to set eight or more rotors to provide redundant power and overcome the yaw loss of control problem. However, the overall size of the flying car 1000 also needs to be considered so that the flying car 1000 can be adapted to public road settings and existing parking systems. For example, in this embodiment, when the vehicle 100 and the aircraft 200 are in a coupled state, the overall structure of the flying car 1000 has a length of no more than 5.9 meters, a height of no more than 2 meters, and a width of no more than 2.2 meters. Setting up eight or more rotors would make it difficult to meet the overall size requirements of the flying car 1000, and would significantly increase the overall weight of the aircraft 200, reduce its lightweight nature, and shorten its endurance and range.

[0077] Therefore, this application chooses to use six foldable rotors 220 as the power system of the aircraft 200. This satisfies the power requirements for the aircraft 200's movement while achieving lightweight design, improving its range, and allowing the overall size of the flying car 1000 to adapt to public road layouts and existing parking systems when the aircraft 200 is coupled with the vehicle 100. With six foldable rotors 220, a malfunction in any one of them could cause the aircraft 200 to lose yaw control. Therefore, this embodiment uses two ducted fans 2310 to overcome this safety issue, breaking through the traditional limitation of six-rotor manned aircraft lacking yaw control capabilities.

[0078] Please see Figure 10The aircraft body 210 has three foldable rotors 220 on each of its left and right sides in the forward direction of the aircraft 200, so that the aircraft 200 is roughly balanced in the left and right directions. Additionally, the aircraft body 210 is also roughly balanced in the forward direction of the aircraft. Specifically, in this embodiment, the aircraft body 210 includes a first body section 2170, a second body section 2180, and a third body section 2190, which are arranged side-by-side and connected along the forward direction of the aircraft 200. The six foldable rotors 220 are divided into three groups, with two foldable rotors 220 in each group. The three groups of foldable rotors 220 correspond to the first body section 2170, the second body section 2180, and the third body section 2190 respectively. The three groups of foldable rotors 220 bear the weight of the first body section 2170, the second body section 2180, and the third body section 2190 respectively, so that the aircraft body 210 is in a roughly balanced state in the forward direction of the aircraft 200.

[0079] Please see Figure 11 In this embodiment, to provide sufficient operating space for each foldable rotor 220, multiple foldable rotors 220 are arranged sequentially and at intervals around the frame 2210. For ease of understanding, when the aircraft 200 is in level flight, the projection of the aircraft 200 onto a horizontal plane shows the projections of the multiple foldable rotors 220 surrounding the projection of the aircraft body 210. This disperses the multiple foldable rotors 220 relative to the aircraft body 210, preventing structural collisions between adjacent or close foldable rotors 220 during transport and ensuring a high degree of safety for the aircraft 200 in flight. However, to avoid the multiple foldable rotors 220 being too dispersed, the degree of dispersion of the multiple foldable rotors 220 is limited in this embodiment. In this embodiment, the foldable rotor 220 forms a rotor rotation surface 2224 when rotating. When the aircraft 200 is in a level flight state, the projection of the aircraft 200 onto the horizontal plane will cause the projection portions of the rotor rotation surfaces 2224 of at least two adjacent foldable rotors 220 to overlap, so that the multiple foldable rotors 220 have a certain degree of integrity and correlation, and improve the aerodynamic efficiency of the foldable rotors 220.

[0080] Please see Figure 12In this embodiment, multiple foldable rotors 220 are arranged divergently, and 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 are located on the same vertical line, or the parallel distance between the vertical line containing the geometric center of polygon 2225 and the vertical line containing the center of gravity of aircraft body 210 is small, so that the aircraft 200 can be in a balanced and stable state during flight. It should be understood that the above-mentioned polygon 2225 should be understood as the polygon 2225 formed by connecting the center points of 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 multiple foldable rotors 220 being approximately a centrally symmetrical figure, so that the driving force formed by any three connected foldable rotors 220 in this embodiment is approximately the same as the driving force formed by another three connected foldable rotors 220, further improving the balance and safety of the aircraft 200 in flight.

[0081] In this embodiment, two foldable rotors 220 located on the same diagonal of polygon 2225 constitute a pair of corresponding foldable rotors 220, and the two corresponding foldable rotors 220 are centrally symmetrically arranged. When the output power of the two foldable rotors 220 is the same, 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 pair are arranged in opposite directions to generate a yaw moment.

[0082] 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, with the rotors mounted on the fixed wings so that the multiple rotors are arranged around the multi-rotor manned aircraft. However, the fixed wings cannot be folded, occupy a lot of space, and the multi-rotor manned aircraft cannot be scaled down, cannot drive on public roads, and can only take off or land on the tarmac, which seriously limits the scope of use of multi-rotor manned aircraft.

[0083] Please see Figure 13 and Figure 14In this embodiment, to achieve a divergent positioning of the multiple foldable rotors 220 relative to the aircraft body 210, the frame 2210 is configured to include a support 2211 and multiple arms 2212. The support 2211 is located on the top of the aircraft body 210, and the multiple arms 2212 are mounted on the support 2211. The arms 2212 are used to mount the foldable rotors 220, therefore the number of arms 2212 is the same as the number of foldable rotors 220. For example, in this embodiment, the number of arms 2212 is consistent with the number of foldable rotors 220, both being six. The six foldable rotors 220 are arranged 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 arranged sequentially at intervals along the circumference of the support. Arm 2212 extends away from the center of bracket 2211 relative to bracket 2211, so that multiple foldable rotors 220 are connected to the aircraft body 210 in a divergent manner. As described above, bracket 2211 in this embodiment can be understood as a structurally independent frame structure, which is set on the top of aircraft body 210 and flight cockpit 2110. Bracket 2211 can be welded to the top of aircraft body 210 to enhance the integrity between aircraft body 210 and bracket 2211. Bracket 2211 can also be fixedly connected to aircraft body 210 by fixing components (such as threaded fasteners such as bolts, or adhesives such as structural adhesives) to facilitate replacement or maintenance of bracket 2211.

[0084] In other embodiments, the support frame 2211 can serve as the overall frame structure of the aircraft body 210, roughly defining the external outline of the aircraft body 210. The support frame 2211 is formed by overlapping multiple rod-like and / or column-like structures, generally forming a cuboid structure. This cuboid structure has an internal space, part or all of which serves as the flight cockpit 2110. In this embodiment, the arm 2212 is located on top of the support frame 2211, i.e., on top of the flight cockpit 2110. The arm 2212 can be integrally formed with the support frame 2211, enhancing the overall integrity between the arm 2212 and the support frame 2211. Alternatively, the arm 2212 can be fixedly and detachably connected to the support frame 2211 via a connecting structure, ensuring the stability of the connection between the arm 2212 and the support frame 2211 and facilitating the replacement or maintenance of the arm 2212. If the arm 2212 and the support 2211 are connected by a connecting structure, the arm 2212 can be rotatably connected to the frame 2211. When the aircraft 200 is in a stationary state, the arm 2212 can be folded and brought close to the support 2211 to reduce the overall size of the aircraft body 210.

[0085] When the aircraft 200 is in a parked state, or when the aircraft 200 is coupled to the vehicle 100, in order to transform the multiple foldable rotors 220 from a divergent state to a convergent state, in this embodiment, multiple arms 2212 are rotatably connected to a support 2211. When the foldable rotors 220 are in a divergent state, the multiple arms 2212 are in an extended state. The arms 2212 can rotate based on the support 2211 in a direction close to the support 2211 to overlap the outer periphery of the support 2211, so that the multiple arms 2212 transform into a folded state, and the multiple foldable rotors 220 transform into a convergent state. This reduces the outer contour 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 conforms to specifications. Conversely, when multiple arms 2212 are in a folded state and multiple foldable rotors 220 are in a clustered state, the arms 2212 can rotate relative to the support 2211 to return to the extended state, so that the multiple foldable rotors 220 return to the divergent state in preparation for flight.

[0086] Please see Figure 15 and Figure 16 To further reduce the external profile of the foldable rotor 220 when the aircraft 200 is in a stationary state, in this embodiment, in addition to the foldable arm 2212, the foldable rotor 220 is also foldable. Specifically, the 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 the hub 2223. The two blades 2222 can approach each other to fold, and the two blades 2222 in the folded state are arranged along the length of the arm 2212, that is, the blades 2222 are stacked on the arm 2212 to reduce the external profile of the aircraft 200. To simplify the structure of the foldable rotor 220, the foldable rotor 220 in this embodiment is a single-shaft, single-blade rotor, and the foldable rotor 220 is a fixed-blade rotor, that is, when the foldable rotor 220 is rotating, the blades 2222 do not deflect or move relative to the hub 2223.

[0087] Please refer to it again. Figure 10For ease of explanation and distinction of each foldable rotor 220, the six foldable rotors 220 in this embodiment are defined as 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, respectively. The six foldable rotors 220 are arranged in a ring around each other. In the counterclockwise direction of their ring 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 at intervals. When the aircraft 200 is projected onto a horizontal plane in flight, with the forward direction of the aircraft 200 as the 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, and 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 to the left of the aircraft body 210, the second rotor 2240 is located to the left front of the aircraft body 210, the fourth rotor 2260 is located to the left rear of the aircraft body 210, the sixth rotor 2280 is located to the right of the aircraft body 210, the first rotor 2230 is located to the right front of the aircraft body 210, and the fifth rotor 2270 is located to the right rear of the aircraft body 210.

[0088] Please see Figure 10 As can be deduced from the preceding text, the first rotor 2230 and the second rotor 2240 correspond to the third body section 2190, the third rotor 2250 and the sixth rotor 2280 correspond to the second body section 2180 of the aircraft body 210, and the fourth rotor 2260 and the fifth rotor 2270 correspond to the first body section 2190. Furthermore, the first rotor 2230 and the fourth rotor 2260 are located on the same diagonal of the polygon 2225 and are paired up, the second rotor 2240 and the fifth rotor 2270 are located on the same diagonal and are paired up, and the third rotor 2250 and the sixth rotor 2280 are located on the same diagonal and are paired up. Assuming that the first rotor 2230, the third rotor 2250 and the fifth rotor 2270 are configured to rotate in a first direction (e.g., clockwise), and the second rotor 2240, the fourth rotor 2260 and the sixth rotor 2280 are configured to rotate in a second direction (e.g., counterclockwise), then the second direction is set opposite to the first direction.

[0089] Please see Figure 17As mentioned earlier, when the aircraft 200 is in level flight, projecting the aircraft 200 onto a horizontal plane ensures that the projected portions of the rotor rotation surfaces 2224 of at least two adjacent foldable rotors 220 overlap, in order to limit the divergence of the foldable rotors 220. However, in actual spatial structures, it is necessary to avoid the overlap of the rotor rotation surfaces 2224 of two adjacent foldable rotors 220. Multiple foldable rotors 220 have their rotor rotation surfaces 2224 located on at least two planes to prevent structural collisions between adjacent foldable rotors 220, ensuring the safety and sustainability of the foldable rotors. In this embodiment, the two foldable rotors 220 corresponding to the first body part 2170 and the two foldable rotors 220 corresponding to the third body part 2190 are placed on the same horizontal plane, which is defined as the first horizontal plane. The two foldable rotors 220 corresponding to the second body part 2180 are placed on another horizontal plane, which is defined as the second horizontal plane, and the height of the first horizontal plane is higher than the height of the second horizontal plane. Specifically, the third rotor 2250 and the sixth rotor 2280 are located on the side of the first rotor 2230, second rotor 2240, fourth rotor 2260, and fifth rotor 2270 facing the ground. The height of the third rotor 2250 and the sixth rotor 2280 is lower than the height of the first rotor 2230, second rotor 2240, fourth rotor 2260, and fifth rotor 2270 to avoid collisions between the third rotor 2250 and the second rotor 2240, and between the sixth rotor 2280 and the first rotor 2230 and fifth rotor 2270. To avoid collisions between the first rotor 2230 and the second rotor 2240, which are located on the same plane, and to avoid collisions between the fourth rotor 2260 and the fifth rotor 2270, which are also located on the same plane, reasonable avoidance measures can be taken based on the length of the arm 2212 and the length of the blade 2222. This embodiment will not be elaborated on further here. In other embodiments, the individual foldable rotors 220 may be disposed on different planes, and this embodiment does not impose specific limitations on this.

[0090] 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 structure of the first rotor 2230, the second rotor 2240, the fourth rotor 2260, and the fifth rotor 2270 and the airflow they generate 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 in 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 in a roughly horizontal direction, forming a roughly horizontal rotor rotation surface 2224, so as to provide vertical driving force to the aircraft body 210.

[0091] Please see Figure 11 When the aircraft 200 is projected onto a horizontal plane, there is a projection overlap between the third rotor 2250 and the second rotor 2240 and the fourth rotor 2260. To avoid excessive overlap between the airflow from the third rotor 2250 and the airflow from the second rotor 2240 and the fourth rotor 2260 in the actual spatial structure, the rotor of the third rotor 2250 is tilted in this example (e.g., Figure 17 As shown), the rotation axis and rotor rotation surface 2224 of the third rotor 2250 are both inclined. The inclined arrangement of the rotor rotation surface 2224 of the third rotor 2250 can reduce the mutual influence of airflow between the third rotor 2250 and the second rotor 2240 and the fourth rotor 2260, and improve the aerodynamic efficiency of the foldable rotor 220.

[0092] Specifically, the two blades 2222 of the third rotor 2250 are essentially 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 at a higher height than the blade 2222 farther from the aircraft body 210. This causes the rotor tip 2229 of the third rotor 2250, facing upwards, to tilt outwards relative to the aircraft body 210, and the rotor rotation surface 2224 of the third rotor 2250 is also tilted outwards relative to the aircraft body 210. This outward tilting configuration allows the radial direction of the third rotor 2250 to deviate from that of the cockpit 2110. The radial direction of the third rotor 2250 is not perpendicular to the cockpit 2110, which reduces the risk of injury to the cockpit 2110 and its occupants in the event of an accidental explosion.

[0093] 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°. Based on the preceding text, the angle between the rotation axis of the third rotor 2250 and the vertical direction in this embodiment can be determined to be [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.

[0094] Please see Figure 17In 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 approximately mirror-symmetrical with respect to the central axis of the aircraft body 210.

[0095] When the aircraft 200 is projected onto a horizontal plane, there is a projection overlap between the sixth rotor 2280 and the first rotor 2230 and the fifth rotor 2270. To avoid excessive overlap between the airflow from the sixth rotor 2280 and the airflow from the first rotor 2230 and the fifth rotor 2270 in the actual spatial structure, the sixth rotor 2280 is tilted in this example (e.g., Figure 17 As shown), the rotation axis and rotor rotation surface of the sixth rotor 2280 are both inclined. The inclined rotor rotation surface 2224 of the sixth rotor 2280 can reduce the mutual influence of airflow between the sixth rotor 2280 and the first rotor 2230 and the fifth rotor 2270, and improve the aerodynamic efficiency of the foldable rotor 220.

[0096] Specifically, the two blades 2222 of the sixth rotor 2280 are essentially 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 at a higher height than the blade 2222 farther from the aircraft body 210. This causes the rotor tip of the sixth rotor 2280, facing the sky, to tilt 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 tilting configuration allows the radial direction of the motor of the sixth rotor 2280 to be relatively offset from that of the cockpit 2110. The radial direction of the motor of the sixth rotor 2280 is not perpendicular to the cockpit 2110, which reduces the risk of injury to the cockpit 2110 and its occupants in the event of an accidental explosion of the motor.

[0097] 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 that 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 preceding text, the angle between the rotation axis of the sixth rotor 2280 and the vertical direction in this embodiment can be determined to be [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.

[0098] In this embodiment, four of the six arms 2212 are corresponding to the first rotor 2230, the second rotor 2240, the fourth rotor 2260, and the fifth rotor 2270, respectively. The remaining two arms 2212 are corresponding to the third rotor 2250 and the sixth rotor 2280, respectively. 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 located on both sides of the aircraft body 210 in the forward direction of the aircraft 200.

[0099] Please see Figure 6 To achieve the staggered positioning of the six foldable rotors 220, in this embodiment, when all six arms 2212 are in the folded state, the four arms 2212 other than the first arm 2213 and the second arm 2214 are stacked side by side on the same plane, which is approximately horizontal. The first arm 2213 and the second arm 2214 are stacked on the side of the four arms 2212 facing the ground, to satisfy the positional state 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, and to improve the space utilization of the six arms 2212 and the six foldable rotors 220.

[0100] Please see Figure 4 In this embodiment, the auxiliary yaw system 230 of the aircraft 200 includes two ducted fans 2310, which are mounted on the aircraft body 210. The ducted fans 2310 can provide auxiliary yaw torque to compensate for the yaw torque deviation caused by the failed foldable rotor 220 when the yaw torque generated by the six foldable rotors 220 is insufficient or when one of the six foldable rotors 220 fails. In this embodiment, the ducted fans 2310 maintain a stable structure relative to the aircraft body 210, simplifying the installation structure of the ducted fans 2310. The fixed position of the ducted fans 2310 is specifically manifested in that the ducted fans 2310 have a rotation axis, and the position of the rotation axis of the ducted fans 2310 relative to the aircraft body 210 remains fixed during flight. In other embodiments, the ducted fans 2310 can be movably mounted on the aircraft body 210 to increase the flexibility of the ducted fans 2310 and improve the performance of the auxiliary yaw system 230. In this embodiment, a ducted fan 2310 is used as an auxiliary yaw system 230. The ducted fan 2310 can generate a large yaw torque. When the aircraft 200 is in flight, the ducted fan 2310 experiences less resistance and has higher overall working efficiency.

[0101] Furthermore, in the level flight state of the aircraft 200, the angle between the rotation axis of the ducted fan 2310 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 the ducted fan 2310 and the horizontal plane can be adaptively adjusted according to actual conditions.

[0102] Please see Figure 4 To facilitate the explanation of the specific position of each ducted fan 2310 on the aircraft body 210 and its relative positional relationship with other structures, in this embodiment, the two ducted fans 2310 are defined as the first ducted fan 2311 and the second ducted fan 2312. Specifically, the first ducted fan 2311 is correspondingly arranged with the third rotor 2250 and is mounted on the first arm 2213. The second ducted fan 2312 is correspondingly arranged with 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 in the forward direction of the aircraft 200, respectively. When the aircraft body 210 deflects, the first ducted fan 2311 and the second ducted fan 2312 mounted on the first arm 2213 and the second arm 2214 have a large deflection torque. With the rated power of the first ducted fan 2311 and the second ducted fan 2312 remaining unchanged, the compensation adjustment range formed by the first ducted fan 2311 and the second ducted fan 2312 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 on the four arms 2212 other than the first arm 2213 and the second arm 2214.

[0103] In this embodiment, both the first ducted fan 2311 and the third rotor 2250 are 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 location has a relatively long distance from the center of gravity of the aircraft body 210, resulting in a relatively long lateral lever arm and a relatively large yaw moment, thus enabling the first ducted fan 2311 to achieve higher power efficiency.

[0104] Specifically, the first arm 2213 has a top side and a bottom side facing away from each other. The top side of the first arm 2213 faces the sky, and the bottom side faces 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, which can avoid structural collision. Furthermore, in order to reduce the influence of airflow from the tip of the blade 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 blade 2222 of the third rotor 2250. Specifically, in this embodiment, the first ducted fan 2311 has a rotation center (i.e., a rotation axis). When the hub 2223 of the third rotor 2250 is rotating, it forms a first maximum rotation surface, which projects the aircraft 200 onto 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 surrounding structures are set to correspond to the hub 2223 of the third rotor 2250. From a macroscopic structural perspective, the rotation center (i.e., the rotation axis) 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 smaller and weaker, and has a smaller impact on the first ducted fan 2311.

[0105] Similarly, in this embodiment, the second ducted fan 2312 has a rotation center (i.e., rotation axis). When the hub 2223 of the sixth rotor 2280 is rotating, it forms a second maximum rotation surface. Projecting the aircraft 200 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 are set corresponding to the hub 2223 of the sixth rotor 2280. From a macroscopic structural perspective, the rotation center (i.e., 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 smaller and weaker, and has less impact on the second ducted fan 2312.

[0106] To monitor the operational status of each foldable rotor 220 and control the operating nodes of the ducted fan 2310 based on their operational status, 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 rotors 220 and the ducted fan 2310 to enable information exchange and command execution. The control mechanism for the failure of one of the foldable rotors 220 is as follows: when any one of the foldable rotors 220 fails, the yaw control system controls two ducted fans 2310 to operate, providing a compensating yaw moment to the aircraft 200. The magnitude of the compensating yaw moment is approximately equal to the difference between the actual yaw moment generated by the remaining five foldable rotors 220 and the target yaw moment. The yaw control system enables the aircraft 200 to cope with the failure of the foldable rotors 220, ensuring the safety of the pilot and the aircraft 200.

[0107] In addition, the yaw control system has a control mechanism for the six foldable rotors 220 working normally but with insufficient yaw torque. The yaw control system controls the two ducted fans 2310 to operate to provide the aircraft 200 with a compensating yaw torque. The magnitude of the compensating yaw torque is approximately equal to the difference between the actual yaw torque formed by the six foldable rotors 220 and the target yaw torque.

[0108] As mentioned above, the first rotor 2230 and the fourth rotor 2260 are paired, the second rotor 2240 and the fifth rotor 2270 are paired, and the third rotor 2250 and the sixth rotor 2280 are paired. Each pair of foldable rotors 220 can keep the aircraft body 210 balanced on the corresponding diagonal. To simplify the control of the yaw control system over the six foldable rotors 220, the yaw control system in this embodiment is further configured to: when any foldable rotor 220 fails, control the foldable rotor 220 corresponding to the failed foldable rotor 220 to stop working, and control the two ducted fans 2310 to operate to provide compensating yaw torque to the aircraft body 210. The advantage of this setup is that, assuming the first rotor 2230 fails, the yaw control system stops the fourth rotor 2260 from operating and activates the two ducted fans 2230 to compensate for the deviation in yaw torque caused by the failure of the first rotor 2230 and the fourth rotor 2260. During this process, the yaw control system only needs to ensure the two ducted fans 2310 provide the yaw torque required for the aircraft body 210 to yaw, thus eliminating or reducing the need to consider the fourth rotor 2260, such as differences in position, output power, and drag experienced by the fourth rotor 2260 and the two ducted fans 2230. This simplifies the calculation steps of the yaw control system, improves efficiency, and allows the aircraft 200 to resume normal operation in a shorter time during flight.

[0109] Furthermore, assuming the first rotor 2230 fails, the yaw control system will stop the fourth rotor 2260 from operating, while the second rotor 2240, fifth rotor 2270, third rotor 2250, and sixth rotor 2280 will still maintain the balance of the aircraft body 210. The yaw control system only needs to control the four normally operating foldable rotors 220 to increase their output power to compensate for the vertical traction force and achieve the desired traction force. There is no need to comprehensively consider the balance adjustment among the second rotor 2240, third rotor 2250, fourth rotor 2260, fifth rotor 2270, and sixth rotor 2280.

[0110] In summary, this application provides a fully electric aircraft 200, which includes an aircraft body 210, a frame 2210, and multiple non-tilt foldable rotors 220. The aircraft body 210 is provided with a cockpit 2110 for carrying passengers. The aircraft 200 uses multiple foldable rotors as its power system, wherein the frame 2210 is located on top of the aircraft body 210, and the multiple non-tilt rotors 220 are connected to the frame 2210. The multiple non-tilt rotors 220 are spaced apart 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 non-tilt rotors 220 remains fixed relative to the aircraft body 210. To address the imbalance of yaw moment generated by multiple non-tilt foldable rotors 220, such as yaw moment imbalance or at least one malfunction of the multiple foldable rotors 220, the aircraft 200 is also equipped with two ducted fans 2310. The two ducted fans 2310 are connected to the frame 2210 and are located on the left and right sides of the aircraft body 210, respectively. When the yaw moment provided by the non-tilt rotors 220 is unbalanced, they provide a compensating yaw moment to the aircraft body 210. Compared with traditional multi-rotor manned aircraft, the aircraft 200 in this embodiment does not have fixed wings to disperse the multiple foldable rotors 220. Instead, it uses a partially foldable frame 2210 to install multiple foldable rotors, allowing the overall structure of the aircraft 200 to retract or unfold. When the aircraft 200 is coupled to the vehicle 100, the size of the flying car 1000 can adapt to the road. Furthermore, the aircraft 200 is equipped with two ducted fans 2310 on the left and right sides of its forward direction. When the yaw moment formed by the multiple non-tilting foldable rotors 220 is in an unbalanced state, the fans can operate and provide a compensating yaw moment to the aircraft body 210. This allows the aircraft body 210 to overcome the problem of not being able to yaw due to the imbalance of yaw moment when it is equipped with six foldable rotors 220.

[0111] Please see Figure 18Based on the aforementioned aircraft and its yaw control system, this application also provides a flight control method for an aircraft, which includes the following steps:

[0112] S10: Obtain the target yaw moment of the aircraft.

[0113] The target yaw moment refers to the yaw moment required for an aircraft to deflect at a certain angle (e.g., yaw the target yaw angle, which can be set according to actual needs). The yaw control system obtains the target yaw moment required for the aircraft to achieve deflection based on the target yaw moment and the aircraft's parameters. Specifically, the aircraft parameters include the aircraft's weight, the drag experienced by the aircraft, the air pressure of the air layer in which the aircraft is located, and the air velocity.

[0114] In some embodiments, the step of "S10: obtaining the target yaw moment of the aircraft" includes: obtaining the current yaw angle and the target yaw angle of the aircraft; and determining the target yaw moment of the aircraft based on the current yaw angle and the target yaw angle.

[0115] S20: During the flight of the aircraft, acquire the first yaw moment generated by the foldable rotor.

[0116] The yaw control system acquires the actual output power of 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 satisfy the aircraft to yaw to the target yaw angle.

[0117] S30: If the difference between the first yaw moment and the target yaw moment is greater than a specified value, determine the compensation yaw moment based on the difference.

[0118] 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. In this embodiment, the specified value refers to the difference between the upper and lower limits of the yaw moment range that enables the aircraft body to achieve the target yaw angle, where the target yaw moment lies 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 can also be a numerical range. The minimum value within the range indicates that when the difference between the target yaw moment and the first yaw moment is the minimum value within the range, the first yaw moment can sufficiently support the aircraft to yaw to the target yaw angle. The maximum value within the range indicates that when the difference between the target yaw moment and the first yaw moment is the maximum value within the range, the first yaw moment can basically support the aircraft to yaw to the target yaw angle.

[0119] The yaw control system then calculates the compensation yaw moment that the auxiliary yaw system needs to provide based on the first yaw moment, the target yaw moment, and the specified value.

[0120] S40: Control the operation of the auxiliary yaw system based on the compensated yaw moment. The difference between the sum of the second yaw moment and the first yaw moment generated by the auxiliary yaw system and the target yaw moment is less than or equal to a specified value.

[0121] The yaw control system controls the operation of the auxiliary yaw system, in which the ducted fan operates 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 folding rotor and the auxiliary yaw system.

[0122] In some embodiments, controlling the operation of the auxiliary yaw system based on the compensating yaw moment includes: determining the target power of the ducted fan based on the compensating yaw moment; determining the target rotational 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 rotational speed to generate a second yaw moment.

[0123] Therefore, in this embodiment, by introducing an auxiliary yaw system, the auxiliary yaw system can operate and provide compensating yaw torque to the aircraft body when the yaw torque generated by the foldable rotor is unbalanced, so that the aircraft body can overcome the problem of not being able to yaw due to the imbalance of yaw torque. Specifically, the power of the auxiliary yaw system can change in real time with the flight attitude and flight dynamic parameters. When the foldable rotor failure is detected, the auxiliary yaw system starts to work. When the aircraft yaws actively or passively, the aircraft controller or flight control system can control the aircraft yaw according to the predetermined control principles (such as heading maintenance or nose turn). If the aircraft controller or flight control system finds that the control capability is insufficient, the auxiliary yaw system is used for supplementary control to meet the yaw control requirements of the aircraft.

[0124] In some embodiments, the step of “S20 during the flight of the aircraft, obtaining the first yaw moment generated by the foldable rotor” in the above method includes the following steps S21 to S22.

[0125] S21: During the flight of the aircraft, acquire the operating parameters of each foldable rotor and the fault index parameters of the foldable rotor. The fault index parameters are obtained by the aircraft's fault self-checking system and characterize the degree of failure of the foldable rotor.

[0126] Specifically, the overall controller or yaw control system can acquire the operating parameters of each foldable rotor, which may include the output power of the foldable rotor, the drag experienced by the foldable rotor, etc. The aircraft also includes a fault self-diagnosis system, which is electrically connected to the yaw control system. The fault self-diagnosis system is used to detect the degree of failure of the foldable rotor (such as the degree of damage or severity of the failure), record the degree of failure of the foldable rotor as a fault index parameter, and transmit the fault index parameter of the foldable rotor to the yaw control system. Taking the rotor assembly as an example of a foldable rotor, as one 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 conditions of the motor drive circuit (e.g., output power, current magnitude, voltage magnitude, at least one of these), thereby determining the degree of rotor motor failure; as another example, the fault detection system may also include a speed sensor, which may be connected to the rotor shaft, and is used to detect the rotor rotation state (e.g., rotation direction, rotation speed, rotation acceleration, at least one of these), and determine whether the rotor is operating normally based on the rotation state, thereby determining the degree of rotor failure.

[0127] In some examples, the fault indicator parameters may carry information such as the specific faulty component and the extent of damage. For instance, fault indicator parameters can be represented by codes, such as fault code EA2, where one code EA indicates the specific location of the fault (e.g., the motor or rotor) and the other code indicates the severity of the fault (e.g., a level 1 or level 2 fault). Therefore, the overall controller can determine the specific location and severity of the fault by reading the fault indicator parameters.

[0128] S22: Based on the operating parameters and fault index parameters, if at least one of the multiple foldable rotors is faulty, obtain the first yaw moment generated by the foldable rotor.

[0129] Specifically, since the fault indicator parameters can carry information including specific faulty components and the extent of damage, the overall controller or yaw control system can determine the number of faulty foldable rotors among multiple foldable rotors based on the fault indicator parameters. When at least one of the multiple foldable rotors is faulty, the overall controller or yaw control system can obtain the first yaw moment generated by the foldable rotor based on its operating parameters. In practice, in some examples, it is not necessary to specifically obtain the first yaw moment generated by the foldable rotor when it is not faulty. Instead, it is necessary to obtain the first yaw moment generated by the foldable rotor when the fault indicator parameters indicate that the foldable rotor is faulty, meaning that its generated yaw moment may deviate from the target yaw moment. Therefore, in this embodiment, the system resources consumed in control and calculation are relatively small, which is beneficial for the smooth operation of the overall controller.

[0130] In some embodiments, when a single foldable rotor among a plurality of foldable rotors is faulty, the step “S22: Based on the operating parameters and fault index parameters, determine that at least one of the plurality of foldable rotors is faulty, and obtain the first yaw moment generated by the foldable rotor” includes the following steps S221 to S222.

[0131] S221: Based on the operating parameters and fault index parameters, if it is determined that one of the multiple foldable rotors has failed, control the foldable rotor corresponding to the failed foldable rotor to stop working.

[0132] The yaw control system, taking into account the power system distribution described above, controls the corresponding folding rotor to cease operation when one of the multiple folding rotors fails, thus maintaining the aircraft's horizontal balance. In this embodiment, the multiple folding rotors are arranged symmetrically along their central axis and in pairs. Therefore, the control process that stops the operation of its corresponding folding rotor when one fails ensures that the lift and thrust provided by the remaining rotors remain essentially balanced. Operating parameters are also easier to control, and the computational burden on the overall control system is relatively small.

[0133] S222: Obtain the first yaw moment generated by all foldable rotors except the malfunctioning foldable rotor and its corresponding foldable rotor.

[0134] Taking six foldable rotors as an example, if two of the 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 compensation for the yaw moment.

[0135] Please see Figure 19 , Figure 19 This is a functional framework diagram of a flight control device disclosed in an embodiment of this application. Based on the above-described aircraft and its flight control method, this application also 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 compensation yaw module A40.

[0136] The target moment calculation module A10 is used to obtain the target yaw moment of the aircraft, and the actual moment calculation module A20 is used to obtain the first yaw moment generated by the foldable rotor during the flight of the aircraft. The compensation moment calculation module A30 is used to determine the compensation yaw moment based on the difference between the first yaw moment and the target yaw moment if the difference is greater than a specified value. The compensation yaw module A40 is used to control the operation of the auxiliary yaw system based on the compensation yaw moment. 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. The specific working process of each of the above functional units / modules can be referred to the corresponding process in the aforementioned method embodiments, and will not be repeated here.

[0137] In the embodiments provided in this application, the coupling between units / modules can be electrical, mechanical, or other forms of coupling. Furthermore, the functional units / modules in the various embodiments of this application can be integrated into one processing unit, or each unit can exist physically separately, or two or more units can be integrated into one unit. The integrated unit can be implemented in hardware or as a software functional unit.

[0138] Based on the aforementioned aircraft and flight control method, embodiments of this application may also provide an aircraft that includes 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 methods described in the foregoing method embodiments. The memory may be independent or integrated with the processor.

[0139] The processor may include one or more processing cores. The processor can connect to various parts of the entire landmass using various interfaces and lines. It can perform various functions and process data by running or executing instructions, programs, code sets, or instruction sets stored in memory, and by calling data stored in memory. Optionally, the processor can be implemented using at least one hardware form of digital signal processing (DSP), field-programmable gate array (FPGA), or programmable logic array (PLA). The processor 1601 can integrate one or a combination of central processing unit (CPU), graphics processing unit (GPU), and modem. The CPU primarily handles the operating system, user interface, and applications; the GPU is responsible for rendering and drawing the displayed content; and the modem handles wireless communication. It is understood that the modem may also not be integrated into the processor and can be implemented separately through a communication chip.

[0140] The memory may include random access memory (RAM) or read-only memory (ROM). The memory can be used to store instructions, programs, code, code sets, or instruction sets. The memory may include a program storage area and a data storage area. The program storage area may store instructions for implementing an operating system, instructions for implementing at least one function (such as touch functionality, sound playback functionality, image playback functionality, etc.), and instructions for implementing the various method embodiments described above. The data storage area may also store data created during the use of the application (such as phonebook data, audio and video data, chat log data, etc.).

[0141] When the computer program instructions stored in the memory are executed, the processor can be used to perform various operations performed by the land vehicle in the above method embodiments. Specific implementations of these operations can be found in the preceding embodiments and will not be repeated here.

[0142] This application also provides a computer-readable storage medium. This computer-readable medium stores computer program code, which can be called by a processor to execute various operations in the above method embodiments. Specific implementations of the above operations can be found in the preceding embodiments and will not be repeated here.

[0143] Computer-readable storage media can be electronic storage devices such as flash memory, electrically erasable programmable read-only memory (EEPROM), erasable programmable read-only memory (EPROM), hard disk, or ROM. Optionally, computer-readable storage media can include non-transitory computer-readable storage medium. The computer-readable storage medium has storage space for program code that performs any of the method steps described above. This computer program code can be read from or written to one or more computer program products. The computer program code can be compressed, for example, in a suitable form.

[0144] In this application specification, certain terms are used to refer to specific components. Those skilled in the art will understand that hardware manufacturers may use different names to refer to the same component. The specification and claims do not distinguish components based on differences in name, but rather on differences in function. The term "comprising" throughout the specification and claims is an open-ended term and should be interpreted as "including but not limited to"; "generally" means that those skilled in the art can solve the technical problem within a certain margin of error and basically achieve the technical effect.

[0145] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "inside", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the purpose of simplifying the description of this application and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0146] In this application, unless otherwise expressly specified or limited, the terms "installation," "connection," "linking," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or merely surface contact. Those skilled in the art can understand the specific meaning of the above terms in this application according to the specific circumstances.

[0147] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., refer to specific features, structures, materials, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of this application. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0148] 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 technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this application, "multiple" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0149] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of this application, and are not intended to limit them. Although this application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features. Such modifications or substitutions 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 this application.

Claims

1. A pure electric aircraft, characterized in that, include: The aircraft body, which is provided with a flight cabin for carrying the driver and passengers; A frame is located on top of the aircraft body. The frame includes multiple arms that are arranged in a series of intervals around the aircraft body. Among the multiple arms, there is a first arm and a second arm. The first arm and the second arm are located on both sides of the aircraft body, respectively. Multiple non-tilt rotors, each corresponding to one of the multiple arms, with each non-tilt rotor connected to its corresponding arm. The multiple non-tilt rotors are spaced apart along a vertical line around the center of gravity of the aircraft. When the aircraft is in flight, the rotation axis of each non-tilt rotor remains fixed relative to the aircraft body. Two ducted fans, including a first ducted fan and a second ducted fan, the first ducted fan being connected to the first arm and the second ducted fan being connected to the second arm, provide a compensating yawing moment to the aircraft body when the yawing moment provided by the non-tilt rotor is in an unbalanced state; In level flight, the first arm has a top side and a bottom side that are opposite to each other, with the bottom side of the first arm facing the ground. The first ducted fan is mounted on the bottom side of the first arm, and the non-tilt rotor on the first arm is mounted on the top side of the first arm. The first ducted fan has a center of rotation, and the hub of the non-tilt rotor on the first arm has a first maximum rotation surface when rotating, which projects the aircraft toward the horizontal plane. The projection surface of the first maximum rotation surface covers the center of rotation of the first ducted fan.

2. The pure electric aircraft as described in claim 1, characterized in that, The aircraft body is symmetrical about a predetermined central axis plane, which 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 non-tilt rotors is six, and the six non-tilt rotors are arranged at intervals around the vertical line where the center of gravity of the aircraft is located. Three non-tilt rotors are provided on each side of the central axis.

3. The pure electric aircraft as described in claim 2, characterized in that, When the aircraft is in flight, the rotation axis of the ducted fan remains fixed relative to the aircraft body.

4. The pure electric aircraft as described in claim 2, characterized in that, The non-tilt rotor has a rotor rotation surface; in flight, multiple rotor rotation surfaces of multiple non-tilt rotors are located on at least two planes; when the aircraft is projected onto a horizontal plane in level flight, the projected portions of the rotor rotation surfaces of at least two adjacent non-tilt rotors overlap.

5. The pure electric aircraft as described in claim 2, characterized in that, The aircraft also includes a yaw control system 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 to provide the compensating yaw torque to the aircraft body when any one of the non-tilt rotors fails.

6. The pure electric aircraft as described in claim 5, characterized in that, In the direction in which the non-tilt rotors are distributed around each other, the center points of the six non-tilt rotors are connected sequentially to form a polygon. Two 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 to: when any one of the non-tilt rotors fails, control the non-tilt rotor corresponding to the failed non-tilt rotor to stop working, and control the two ducted fans to operate to provide the compensating yaw torque to the aircraft body.

7. The all-electric aircraft as described in claim 6, characterized in that, 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. The six non-tilt rotors are arranged in ascending order in a counterclockwise direction in a direction in which they are distributed around each other. The third rotor is mounted on the first arm, and the sixth rotor is mounted on the second arm. The first rotor and the fourth rotor are paired together, the second rotor and the fifth rotor are paired together, and the third rotor and the sixth rotor are paired together.

8. The all-electric aircraft as described in claim 7, characterized in that, The first rotor, the third rotor, and the fifth rotor are configured to rotate along a first direction, and the second rotor, the fourth rotor, and the sixth rotor are configured to rotate along a second direction, which is opposite to the first direction.

9. The pure electric aircraft as described in claim 2, characterized in that, 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. The six non-tilt rotors are arranged in ascending order in a counterclockwise direction in the direction in which they are distributed around each other. The third rotor is mounted on the first arm, and the sixth rotor is mounted on the second arm. When the aircraft is projected onto a horizontal plane during flight, the projections of the third rotor and the sixth rotor are respectively located on both sides of the aircraft body in the direction of flight. The first rotor is located at the front right of the aircraft body, the second rotor is located at the front left of the aircraft body, the fourth rotor is located at the rear left of the aircraft body, and the fifth rotor is located at the rear right of the aircraft body.

10. The all-electric aircraft as described in claim 9, characterized in that, In level flight, 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 tip of the third rotor is tilted outward relative to the aircraft body; or / and, The rotation axis of the sixth rotor is at an angle greater than or equal to 5 degrees and less than or equal to 10 degrees with respect to the vertical direction, and the rotor tip of the sixth rotor is tilted outward relative to the aircraft body.

11. The all-electric aircraft as described in claim 9, characterized in that, The rotation axes of the first rotor, the second rotor, the fourth rotor, and the fifth rotor are arranged in a vertical direction.

12. The all-electric aircraft as described in claim 9, characterized in that, The rotor surface of the third rotor is lower than the rotor surface of either the second rotor or the fourth rotor; or / and The rotor rotation surface of the sixth rotor is lower than that of either the first rotor or the fifth rotor.

13. The all-electric aircraft as described in claim 2, characterized in that, The frame includes a support and six arms. The support is disposed on the top of the aircraft body, and the six arms are arranged sequentially at intervals along the circumference of the support. Each arm extends relative to the support in a direction away from the center of the support. The six non-tilt rotors are arranged in a one-to-one correspondence with the six arms, with each non-tilt rotor mounted on a corresponding arm.

14. The all-electric aircraft as described in claim 13, characterized in that, Each of the robotic arms is foldably connected to the support, and the robotic arm can be rotated and unfolded relative to the support to be in an extended state, and the robotic arm can be rotated relative to the support and stacked on the outer periphery of the support to be in a folded state.

15. The all-electric aircraft as described in claim 14, characterized in that, 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. The six rotors are arranged in ascending order in a counterclockwise direction in a direction in which they are distributed around each other. The third rotor is mounted on the first arm, and the sixth rotor is mounted on the second arm. In the folded state, the first arm and the second arm are stacked on the ground-facing side of the four arms corresponding to the first rotor, the second rotor, the fourth rotor, and the fifth rotor.

16. The all-electric aircraft as described in claim 15, characterized in that, In flight, the second arm has a top side and a bottom side that are opposite to each other, with the bottom side of the second arm facing the ground. The second ducted fan is mounted on the bottom side of the second arm, and the sixth rotor is mounted on the top side of the second arm. The second ducted fan has a center of rotation, and the hub of the sixth rotor has a second maximum rotation surface in the rotating state, which projects the aircraft toward the horizontal plane. The projection surface of the second maximum rotation surface covers the center of rotation of the second ducted fan.

17. The all-electric aircraft as described in claim 2, characterized in that, 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.

18. The all-electric aircraft as described in claim 2, characterized in that, Each of the non-tilt rotors includes two blades, which are rotatably arranged relative to each other and can be brought close together to be stacked or moved away from each other to be deployed.

19. The all-electric aircraft as described in claim 1, characterized in that, The aircraft also includes a docking mechanism, which is located on the side of the aircraft body opposite to the direction of travel of the aircraft. The docking mechanism is adapted to cooperate with an external traction mechanism so that the aircraft body moves under the traction of the external traction mechanism.

20. The all-electric aircraft as described in claim 19, characterized in that, The docking mechanism includes a mating ring, which is rotatably connected to the side of the aircraft chassis facing the ground. The mating ring is capable of rotating vertically relative to the aircraft chassis.

21. The all-electric aircraft as described in claim 1, characterized in that, The aircraft also includes at least two landing gear mechanisms, which are arranged side-by-side and spaced apart along the forward direction of the aircraft body. Each landing gear mechanism includes: Two landing gears, each rotatably connected to the aircraft body on either side of the aircraft's forward direction; and, Two drive units are provided, each corresponding to one of the two landing gears. Each drive unit is connected between a corresponding landing gear and the aircraft body. The drive unit drives the landing gear to rotate away from the aircraft body to unfold into a supported state, or drives the landing gear to rotate towards the aircraft body to fold into a folded state.

22. A flight control method for an aircraft, characterized in that, For controlling the flight of an aircraft, the aircraft includes an aircraft body, a frame, multiple non-tilt rotors, and two ducted fans. The aircraft body is equipped with a cockpit for carrying passengers and crew. The frame is located on top of the aircraft body and includes multiple arms arranged sequentially and at intervals around the aircraft body. Among the multiple arms, there are first and second arms, which are located on opposite sides of the aircraft body. Each non-tilt rotor corresponds one-to-one with one of the multiple arms, and each non-tilt rotor is connected to its corresponding arm. The multiple non-tilt rotors are arranged at intervals around the vertical line where the aircraft's center of gravity is located. When the aircraft body is in flight, the rotation axis of each non-tilt rotor remains fixed relative to the aircraft body. The ducted fan includes a first ducted fan and a second ducted fan. The first ducted fan is connected to the first arm, and the second ducted fan is connected to the second arm. When the yaw moment provided by the non-tilt rotor is unbalanced, it provides a compensating yaw moment to the aircraft body. In level flight, the first arm has a top side and a bottom side that are opposite to each other. The bottom side of the first arm faces the ground. The first ducted fan is mounted on the bottom side of the first arm, and the non-tilt rotor on the first arm is mounted on the top side of the first arm. The first ducted fan has a rotation center. The hub of the non-tilt rotor on the first arm has a first maximum rotation surface when rotating, which projects the aircraft toward the horizontal plane. The projection surface of the first maximum rotation surface covers the rotation center of the first ducted fan. The flight control method includes: Obtain the target yaw moment of the aircraft; During the flight of the aircraft, the first yaw moment generated by the multiple non-tilt rotors is acquired; If the difference between the first yaw moment and the target yaw moment is greater than a specified value, a compensation yaw moment is determined based on the difference. The ducted fan is controlled 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 ducted fan and the target yaw moment is less than or equal to the specified value.

23. The flight control method as described in claim 22, characterized in that, During the flight of the aircraft, the acquisition of the first yaw moment generated by multiple non-tilt rotors includes: 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 acquired. The fault index parameters are obtained by the fault self-checking system of the aircraft and represent the degree of failure of the non-tilt rotor. Based on the operating parameters and the fault index parameters, if at least one of the multiple non-tilt rotors is faulty, the first yaw torque generated by the multiple non-tilt rotors is obtained.

24. The flight control method as described in claim 23, characterized in that, The number of non-tilt rotors is six, and the six non-tilt rotors are arranged in pairs, with the two pairs of non-tilt rotors having opposite rotation directions; When a single non-tilt rotor among the plurality of non-tilt rotors malfunctions, the step of determining, based on the operating parameters and the fault index parameters, that at least one of the plurality of non-tilt rotors is malfunctioning, and then obtaining the first yaw moment generated by the plurality of non-tilt rotors, includes: Based on the operating parameters and the fault index parameters, if it is determined that one of the non-tilt rotors among the multiple non-tilt rotors has failed, the non-tilt rotor corresponding to the failed non-tilt rotor is controlled to stop working. Obtain the first yaw moment generated by all the non-tilt rotors except the one that malfunctioned and its corresponding non-tilt rotor.

25. The flight control method as described in claim 22, characterized in that, The acquisition of the target yaw moment of the aircraft includes: Obtain the current yaw angle of the aircraft and the target yaw angle; The target yaw moment of the aircraft is determined based on the current yaw angle and the target yaw angle. The step of controlling the operation of the ducted fan based on the compensated yaw moment includes: The target power of the ducted fan is determined based on the compensated yaw moment. The target rotational speed of the ducted fan is determined based on the target power, the size of the ducted fan, and its aerodynamic efficiency. The ducted fan is controlled to operate according to the target rotational speed, so that the ducted fan generates a second yaw torque.

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