Tilt rotor aircraft and aircraft control method

Abstract

The invention provides a tilt-rotor aircraft and an aircraft control method, and relates to the technical field of aircrafts. The aircraft fuselage is hung below the aircraft wings through the tilting shafts, vector thrust is generated by the multiple sets of propelling devices located on the upper surfaces and the lower surfaces of the wings to directly achieve the wing tilting function, and the damping devices are used for restraining swinging of the fuselage in the wing tilting process. On the basis that the gravity self-stabilization characteristic of the suspension structure is utilized to weaken the fuselage attitude influence of the pendulum bob effect, it is ensured that the fuselage attitude stability maintaining control link and the wing tilting angle control link are independently controlled, and kinematics control decoupling of fuselage attitude control and wing tilting control is achieved; and the damping device can be utilized to realize mechanical coupling cooperation of the two control links so as to jointly guarantee the overall stability of the attitude of the aircraft. According to the damping device, the aircraft operation energy consumption required for maintaining a specific flight mode can be remarkably reduced through wing dip angle locking operation.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft technology, in particular to a tilt-wing aircraft and an aircraft control method. BACKGROUND

[0002] With the continuous development of aviation technology, tilt-wing aircraft gradually become an important research direction in today's aviation technology because they can realize free switching between helicopter mode and fixed-wing mode by controlling wing tilting, so as to take into account the high-speed cruising ability of fixed-wing aircraft and the vertical take-off / short take-off ability of helicopters.

[0003] At present, the traditional tilt-wing aircraft mainly realizes the take-off and level flight conversion operation of the aircraft (including the conversion operation from the take-off stage of the aircraft to the level flight stage of the aircraft, and the conversion operation from the level flight stage of the aircraft to the landing stage of the aircraft) through the wing tilting action. During the take-off and level flight conversion process of the aircraft, the inertia moment generated by the wing tilting action will induce a serious pendulum effect (i.e., a violent swing) of the fuselage, which often needs to use a single-loop control algorithm considering the dynamic control coupling relationship of the wing tilting control and the fuselage attitude control to compensate the motion, so as to take into account the wing tilting effect and the fuselage attitude stabilization effect as much as possible. However, it is worth noting that this aircraft control mechanism cannot achieve a balance between the control algorithm complexity and the control system response speed, and often cannot suppress the fuselage swing in complex flight conditions, and the overall aircraft control stability is poor. In addition, this aircraft control mechanism needs to be equipped with a large-power tilting driving device to take into account the realization of the wing tilting angle adjustment function and the fuselage attitude stabilization function, which often leads to a large overall weight and size of the aircraft, and at the same time, due to the limitation of the fuselage-wing dynamic control coupling relationship, the wing tilting control precision and stability are poor. SUMMARY

[0004] Therefore, the application aims to provide a tilt-wing aircraft and an aircraft control method, which can suspend the aircraft body below the aircraft wings through a tilting shaft, directly realize the wing tilting function by using a plurality of sets of propulsion devices located on the upper and lower surfaces of the wings, and focus on suppressing the body swing by using a damping device mechanically connected with the aircraft body and the tilting shaft respectively during the wing tilting process, so as to ensure that the body attitude control link and the wing tilting angle control link are independently controlled respectively on the basis of weakening the influence of the body attitude by the pendulum effect by using the gravity self-stabilization characteristics of the suspension structure, realize the kinematic control decoupling between the body attitude control and the wing tilting control (i.e., realize the body-wing double-loop decoupling control effect), and realize the mechanical coupling of the two control links by using the damping device to co-operate to jointly ensure the overall stability of the aircraft attitude, without additionally equipping a large-power tilting driving device to realize the wing tilting angle adjustment function and the body attitude stabilization function, thereby effectively reducing the overall weight and size of the aircraft, and fundamentally avoiding the problems of phase lag, self-excited oscillation and overshoot caused by the body-wing dynamic control coupling, so as to improve the wing tilting control precision and stability.

[0005] In order to achieve the above-mentioned purpose, the technical scheme adopted by the embodiments of the application is as follows: In a first aspect, the application provides a tilt-wing aircraft, which comprises a body, a first inertial measurement unit, a main control unit, a damping device, two wings and a plurality of sets of propulsion devices, wherein the upper and lower surfaces of each wing are respectively provided with a set of propulsion devices, the two sets of propulsion devices on the same wing are symmetrical to each other, and the first inertial measurement unit is used to detect the wing tilting attitude information. The two wings are symmetrically arranged on the two sides of the body and are rotationally connected with the body through a tilting shaft, wherein the mounting position of the tilting shaft on the body is above the body gravity center position, and the mounting position of the tilting shaft on the corresponding wing is near the aerodynamic center of the wing. The main control unit is electrically connected with the first inertial measurement unit and the plurality of sets of propulsion devices respectively, and is used to regulate and control the thrust size of each set of propulsion devices in real time according to the wing tilting attitude information detected by the first inertial measurement unit during the take-off, landing and cruising conversion process of the aircraft, so that the wing tilts around the tilting shaft. The damping device is mechanically connected with the body and the tilting shaft respectively, and is electrically connected with the main control unit, and is used to lock and maintain or release the wing tilting angle under the control of the main control unit, or provide a damping force to the body to suppress the body swing during the wing tilting process.

[0006] In an optional embodiment, the distance between the mounting position of the tilting shaft on the corresponding wing and the wing leading edge ranges from 20% to 40% of the wing chord length. The mounting position of the tilting shaft on the fuselage is 0-10% of the wing chord length away from the vertical line of the center of gravity of the fuselage.

[0007] In an optional embodiment, the mounting position of the tilting shaft on the fuselage is on the vertical line of the center of gravity of the fuselage. The center of mass of the two wings and the plurality of sets of propulsion devices is at the center point of the tilting shaft.

[0008] In an optional embodiment, the aircraft further comprises a tail wing. The tail wing is mounted at the tail position of the fuselage, and the tail wing is provided with an elevator and a rudder, and each wing is provided with aileron structure at the trailing edge position of the wing. The main control unit is electrically connected with the elevator, the rudder and the aileron structure, respectively, for controlling the movement state of the elevator, the rudder and the aileron structure respectively during the level flight stage of the aircraft, so as to adjust the pitch attitude of the aircraft, the yaw attitude of the aircraft and / or the roll attitude of the aircraft, wherein the elevator is used to realize the pitch attitude adjustment function of the aircraft, the rudder is used to realize the yaw attitude adjustment function of the aircraft, and the aileron structure is used to realize the roll attitude adjustment function of the aircraft.

[0009] In an optional embodiment, the aircraft further comprises a second inertial measurement unit mounted on the fuselage, wherein the second inertial measurement unit is used to detect the pitch attitude information of the fuselage. The main control unit is electrically connected with the second inertial measurement unit, and is used to dynamically control the damping force provided by the damping device in order to suppress the swing of the fuselage in real time according to the pitch attitude information of the fuselage detected by the second inertial measurement unit during the take-off, landing and level flight conversion process of the aircraft, so as to converge the pitch angular velocity of the fuselage to zero.

[0010] In a second aspect, the present application provides an aircraft control method, which applies the main control unit included in the tilting wing aircraft according to any one of the preceding embodiments, and the control method comprises: Obtaining a take-off, landing and level flight conversion instruction, and determining a desired wing tilt angle matched with the take-off, landing and level flight conversion instruction; Controlling the damping device to release the locking of the wing tilt angle, and controlling the thrust of each set of propulsion device according to the desired wing tilt angle, the actual airspeed of the aircraft, and the wing tilt attitude information detected by the first inertial measurement unit in real time, so as to make the wing tilt around the tilting shaft to the desired wing tilt angle; Controlling the damping device to provide damping force to the fuselage to suppress the swing of the fuselage during the wing tilt process; When the wing tilts to the desired wing tilt angle, the damping device is controlled to lock and maintain the wing tilt angle.

[0011] In an optional implementation, the wing tilt attitude information detected in real time by the first inertial measurement unit includes the actual wing tilt angle and the actual wing tilt angular velocity. Then, the step of adjusting the thrust of each of the multiple propulsion devices based on the desired wing tilt angle, the actual airspeed of the aircraft, and the wing tilt attitude information detected in real time by the first inertial measurement unit, so that the wing tilts around the tilt axis to the desired wing tilt angle, includes: Based on the actual wing tilt rate, calculate the theoretical thrust difference between the propulsion devices corresponding to the upper and lower surfaces of the same wing. Based on the actual airspeed of the aircraft and the actual wing tilt angle, the wing tilt thrust difference during safe flight of the aircraft is calculated based on the theoretical thrust difference. Calculate the actual tilt angle difference between the desired wing tilt angle and the actual wing tilt angle, and calculate the desired propulsion speed of each of the multiple propulsion devices based on the wing tilt thrust difference, with the goal of converging the actual tilt angle difference to zero. According to the calculated desired propulsion speed of each of the multiple propulsion devices, the speed of each of the multiple propulsion devices is adjusted.

[0012] In an optional implementation, the step of calculating the wing tilt thrust difference for safe flight based on the theoretical thrust difference, according to the actual airspeed and the actual wing tilt angle, includes: Detect whether the actual airspeed of the aircraft and the actual wing tilt angle meet the dangerous conditions for wing rollover. If the actual airspeed of the aircraft and the actual wing tilt angle meet the dangerous conditions of wing tilt, the theoretical thrust difference is weighted by a preset safe tilt coefficient to obtain the wing tilt thrust difference, wherein the preset safe tilt coefficient is a positive decimal less than 1. If the actual airspeed of the aircraft and the actual wing tilt angle do not meet the dangerous conditions for wing tilt, the theoretical thrust difference is directly used as the wing tilt thrust difference.

[0013] In an optional implementation, when the tiltrotor aircraft includes a second inertial measurement unit, and the fuselage pitch attitude information detected by the second inertial measurement unit includes the actual fuselage pitch angle and the actual fuselage pitch angular velocity, then the step of controlling the damping device to provide damping force to the fuselage to suppress fuselage sway includes: Based on the actual fuselage pitch angle and the actual fuselage pitch angular velocity, the damping force is predicted with the aim of the fuselage pitch angular velocity converging to zero, and the corresponding sway suppression damping force is obtained. The damping device is driven to operate according to the swing suppression damping force.

[0014] In an optional implementation, the control method further includes: When the wing tilt angle maintained by the damping device is within the preset takeoff and landing tilt angle, the aircraft takeoff command or aircraft landing command is obtained. Based on the desired takeoff speed included in the aircraft takeoff command or the desired landing speed included in the aircraft landing command, the thrust of each of the multiple propulsion devices is adjusted so that the tiltrotor aircraft takes off at the desired takeoff speed or lands at the desired landing speed.

[0015] In this case, the beneficial effects of the embodiments of this application may include the following: This application symmetrically positions a pair of aircraft wings on either side of the aircraft fuselage and connects them to the fuselage via a tilt axis. The tilt axis's fuselage mounting position (i.e., its position on the fuselage) is above the fuselage's center of gravity, while its wing mounting position (i.e., its position on the wing) is near the wing's aerodynamic center. This allows the aircraft fuselage to be suspended below the wings. Multiple propulsion devices located on the upper and lower surfaces of the wings generate vector thrust to directly achieve wing tilting. A damping device is mechanically connected to both the fuselage and the tilt axis. During wing tilting, the damping device focuses on suppressing fuselage sway. This ensures stable fuselage attitude control (which relies solely on the damping device) by leveraging the gravity-based self-stabilizing characteristics of the suspension structure to mitigate the pendulum effect's influence on fuselage attitude during wing tilting, while also utilizing the damping device's independent power supply. The damping force control and wing tilt angle control loop (i.e., a closed-loop loop that controls wing tilt by adjusting the thrust difference of the propulsion device based on the wing tilt attitude information detected by the first inertial measurement unit) are controlled independently, realizing kinematic control decoupling between fuselage attitude control and wing tilt control (i.e., achieving the effect of fuselage-wing dual-loop decoupling control). At this time, the damping device can realize the mechanical coupling and coordination of these two control loops to jointly ensure the overall stability of the aircraft attitude during takeoff, landing and level flight transitions. Therefore, there is no need to equip an additional high-power tilt drive device to take care of both wing tilt angle adjustment and fuselage attitude stabilization functions. This can effectively reduce the overall weight and size of the aircraft and avoid the phase lag / self-excited oscillation / overshoot problems caused by fuselage-wing dynamic control coupling from the root, thereby improving the accuracy and stability of wing tilt control. Furthermore, when an aircraft needs to maintain a specific flight mode (e.g., level flight mode, vertical takeoff and landing mode, ultra-short takeoff and landing mode, hovering mode, etc.), the damping device can directly lock the wing tilt angle at the desired tilt angle state (i.e., wing tilt angle locking operation) without driving the propulsion device to continuously output a specific thrust difference to maintain the aforementioned desired tilt angle state, thereby significantly reducing the aircraft operating energy consumption required to maintain the aforementioned specific flight mode.

[0016] To make the above-mentioned objectives, features and advantages of this application more apparent and understandable, preferred embodiments are described below in detail with reference to the accompanying drawings. Attached Figure Description

[0017] To more clearly illustrate the technical solutions of the embodiments of this application, the accompanying drawings used in the embodiments will be briefly introduced below. It should be understood that the following drawings only show some embodiments of this application and should not be regarded as a limitation of the scope. For those skilled in the art, other related drawings can be obtained based on these drawings without creative effort.

[0018] Figure 1 This is an overall assembly diagram of a tiltrotor aircraft provided in an embodiment of this application; Figure 2 for Figure 1 A schematic diagram showing the assembly between the two wings and multiple propulsion units. Figure 3 for Figure 1 A schematic diagram showing the distribution of the fuselage mounting positions of the tilt-wing aircraft. Figure 4 for Figure 1 A schematic diagram showing the wing mounting positions of the tilt-axis in the tilt-wing aircraft shown. Figure 5 for Figure 1 A schematic diagram of the control rudder distribution of the tiltrotor aircraft shown. Figure 6 One of the schematic flowcharts of the aircraft control method provided in the embodiments of this application; Figure 7 This is a second schematic flowchart of the aircraft control method provided in the embodiments of this application.

[0019] Icons: 10- Tiltrotor aircraft; 11- Fuselage; 12- Wing; 13- Propulsion system; 14- Tiltrotor shaft; 15- Tail fin; 16- Aerodynamic control rudder; 17- Aileron structure. Detailed Implementation

[0020] To make the objectives, technical solutions, and advantages of the embodiments of this application clearer, the technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. The components of the embodiments of this application described and shown in the accompanying drawings can generally be arranged and designed in various different configurations.

[0021] Therefore, the following detailed description of the embodiments of this application provided in the accompanying drawings is not intended to limit the scope of the claimed application, but merely to illustrate selected embodiments of the application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without inventive effort are within the scope of protection of this application.

[0022] It should be noted that similar labels and letters in the following figures indicate similar items. Therefore, once an item is defined in one figure, it does not need to be further defined and explained in subsequent figures.

[0023] In the description of this application, it should be understood that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, or the orientation or positional relationship commonly used when the product is in use, or the orientation or positional relationship commonly understood by those skilled in the art. They are used only for the convenience of describing this application and simplifying the description, and are not intended to indicate or imply that the equipment or component 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.

[0024] In the description of this application, it should also be noted that, unless otherwise expressly specified and limited, the terms "set up," "install," "connect," and "link" 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; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this application based on the specific circumstances.

[0025] Furthermore, it is understood in the description of this application that relational terms such as "first" and "second" are used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element. Those skilled in the art will understand the specific meaning of the above terms in this application based on the specific circumstances.

[0026] The following detailed description of some embodiments of this application is provided in conjunction with the accompanying drawings. Unless otherwise specified, the following embodiments and features can be combined with each other.

[0027] Please refer to the reference. Figure 1 , Figure 2 , Figure 3 and Figure 4 ,in Figure 1 This is an overall assembly diagram of the tiltwing aircraft provided in the embodiments of this application. Figure 2 yes Figure 1The diagram shows the fuselage mounting locations of the tilt-wing shaft in the tilt-wing aircraft. Figure 3 yes Figure 1 The diagram shows the wing mounting positions of the tilt-axis in a tilt-wing aircraft. Figure 4 yes Figure 1 The diagram shows the wing mounting positions of the tilt-wing axis in the tilt-wing aircraft. In this embodiment, the tilt-wing aircraft 10 can switch between multiple flight modes (including level flight mode, vertical takeoff and landing mode, ultra-short takeoff and landing mode, hovering mode, etc.). The tilt-wing aircraft 10 may include a fuselage 11, a first inertial measurement unit (FMU), and a vertical takeoff and landing (VTOL) system. Figure 1 (not shown), main control unit ( Figure 1 (not shown), damping device ( Figure 1 (Not shown), two wings 12 and multiple propulsion units 13.

[0028] In this embodiment, each wing 12 is equipped with a set of propulsion devices 13 on its upper and lower surfaces. The two sets of propulsion devices 13 on the same wing 12 are symmetrical to each other. The two wings 12 are symmetrically arranged on both sides of the fuselage 11 and are rotatably connected to the fuselage 11 through a tilting shaft 14. One end of the tilt shaft 14 is fixedly connected to the end face of one wing 12 away from the wingtip, and the other end of the tilt shaft 14 passes through the fuselage 11 and is fixedly connected to the end face of another wing 12 away from the wingtip, so that the two wings 12 maintain the same wing tilt attitude in real time. At this time, the two sets of propulsion devices 13 installed on each of the two wings 12 are symmetrical about the fuselage axis of the fuselage 11 (i.e., the axis passing through the nose and tail). The axis of the tilt shaft 14 is located on the plane of symmetry between the two sets of propulsion devices 13 on each wing 12, so as to ensure that the multiple sets of propulsion devices 13 can be evenly and symmetrically distributed on the tilt-wing aircraft 10. The first inertial measurement unit can be set on any wing 12 to detect wing tilt attitude information (including actual wing tilt angle and actual wing tilt angular velocity, etc.). Each set of propulsion devices 13 is used to provide thrust, and each set of propulsion devices 13 includes at least one thruster, which can be, but is not limited to, a rotor, ducted fan, or other fluid propulsion structure.

[0029] Optionally, in the first embodiment of this example, each of the multiple propulsion devices 13 includes multiple propellers, wherein the multiple propellers in the same group of propulsion devices 13 are respectively equipped with a connecting bracket and are installed on the corresponding wing 12 through the connecting bracket, so that the installation positions (i.e. the connecting bracket setting positions) of the multiple propellers included in the same group of propulsion devices 13 on the same wing 12 are distributed at intervals.

[0030] Optionally, in the second embodiment of this invention, each group of propulsion devices 13 includes multiple propellers, wherein at least one propeller in the same group of propulsion devices 13 is equipped with the same connecting bracket and is mounted on the corresponding wing 12 via the connecting bracket, so that the aforementioned at least one propeller is mounted in the same position on the wing 12. The number of connecting brackets equipped in the same group of propulsion devices 13 is at least one, and the number of connecting brackets in the same group of propulsion devices 13 is less than the total number of propellers; the different connecting brackets are spaced apart from each other on the corresponding wing 12.

[0031] In this embodiment of the application, the tilting shaft 14 is mounted on the fuselage 11 at the following position (i.e. Figure 2 The fuselage mounting position shown is above the fuselage center of gravity, and the mounting position of the tilt shaft 14 on the corresponding wing 12 (i.e. Figure 3 The wing mounting position shown is located near the aerodynamic center of the corresponding wing 12 (approximately within 25% ± 5% of the wing chord length from the leading edge of the wing; for example, the distance from the aerodynamic center of the NACA0015 airfoil to the leading edge of the corresponding wing is 28% of the wing chord length). This allows the fuselage 11 to be suspended below the two wings 12, enabling the gravity self-stabilizing characteristics of the suspension structure to achieve self-balancing coupling between the wing aerodynamic torque and the fuselage gravity torque during wing tilting. This weakens the pendulum effect on the fuselage attitude and facilitates the decoupling of kinematic control between fuselage attitude control and wing tilting control, fundamentally avoiding problems such as phase lag / self-excited oscillation / overshoot caused by fuselage-wing dynamic control coupling. The distance between the mounting position of the tilt axis 14 on the corresponding wing 12 and the leading edge of the wing (i.e., the distance between the mounting position of the tilt axis 14 on the corresponding wing 12 and the leading edge of the wing) is within a certain range ... Figure 3 "D1" in the figure represents 20% to 40% of the wing chord length (for example, the tilt axis on the NACA0015 airfoil is mounted at a distance of approximately 28% ± 5% of the wing chord length from the leading edge of the wing); the distance from the mounting position of the tilt axis 14 on the fuselage 11 to the vertical line of the fuselage center of gravity ranges from 0% to 10% of the wing chord length.

[0032] Optionally, in one embodiment of this invention, the tilt axis 14 is mounted on the fuselage 11 at a position on the vertical line of the fuselage's center of gravity, and the overall center of gravity of the two wings 12 and the multiple sets of propulsion devices 13 is located at the center point of the tilt axis 14, thereby effectively preventing the overall center of gravity of the aircraft from shifting during wing tilting.

[0033] In this embodiment, the axis of each of the two sets of propulsion devices 13 on the same wing 12 can form a negative angle (ranging from -3° to -8°) with the chord direction of the wing 12, in order to maintain the wing's cruise angle of attack during level flight and reduce thrust loss. In one embodiment, the axis of each of the two sets of propulsion devices 13 on the same wing 12 can form an angle of -5° with the chord direction of the wing 12.

[0034] In this embodiment, the main control unit is electrically connected to the first inertial measurement unit and multiple propulsion devices 13, respectively. During the aircraft's takeoff and landing transition (including the transition from ultra-short takeoff / vertical takeoff mode to level flight mode, and from level flight mode to ultra-short takeoff / vertical landing mode), the main control unit adjusts the thrust of each of the multiple propulsion devices 13 in real time based on the wing tilt attitude information detected by the first inertial measurement unit. This causes the wing 12 to tilt around the tilt axis 14 (i.e., the wing tilt function is achieved through the vector thrust generated by the multiple propulsion devices 13), thereby adjusting the wing tilt angle to the desired wing tilt angle state corresponding to the desired flight mode (i.e., the aforementioned level flight mode or ultra-short takeoff / vertical landing mode). The main control unit can be deployed on the fuselage 11.

[0035] In this embodiment, for the tilt-wing aircraft 10, the desired wing tilt angle for the aircraft's vertical takeoff and landing modes (including vertical takeoff and vertical landing modes), ultra-short takeoff modes, and hovering modes requires the propulsion device axis to be as close to vertical as possible (i.e., parallel to the vertical line of the fuselage's center of gravity), and the thrust direction to point upwards on the fuselage; the desired wing tilt angle for the aircraft's ultra-short takeoff mode requires the propulsion device axis to form a specific angle (e.g., 30°) with the vertical direction, and the thrust direction to point diagonally upwards on the fuselage; the desired wing tilt angle for the aircraft's level flight mode requires the propulsion device axis to be as parallel as possible to the nose direction, and the thrust direction to point as much as possible towards the nose direction. Since the axis of the propulsion device 13 forms a negative angle with the chord of the wing, when the aforementioned negative angle is -5°, the expected wing tilt angle range for the aircraft's vertical take-off and landing mode, ultra-short take-off mode, and hovering mode can be 75°~95°, the expected wing tilt angle range for the aircraft's ultra-short take-off mode can be 40°~60°, and the expected wing tilt angle range for the aircraft's level flight mode can be 3°~10°.

[0036] When the tiltrotor aircraft 10 is in vertical takeoff and landing (VTOL) mode, ultra-short takeoff and landing (USTOL) mode, or hovering mode, the main control unit can adjust the thrust of each of the multiple propulsion devices 13 to achieve the aircraft pitch attitude control function, aircraft yaw attitude control function, or aircraft roll attitude control function. Specifically, the main control unit can achieve the aircraft pitch attitude control function by adjusting the thrust difference between two propulsion devices 13 on the same wing 12; the main control unit can achieve the aircraft roll attitude control function by adjusting the overall thrust difference between the two wings 12; and the main control unit can achieve the aircraft yaw attitude control function by adjusting the thrust difference between two propulsion devices 13 on at least one diagonal line.

[0037] When the tilt-wing aircraft 10 is in level flight mode, the main control unit can directly adjust the thrust of each of the multiple propulsion devices 13 to achieve the level flight cruise acceleration and deceleration function of the tilt-wing aircraft 10. At this time, the lift of the aircraft is provided by the two wings 12, and the forward propulsion of the aircraft is provided by the multiple propulsion devices 13.

[0038] In this embodiment, the damping device can be mechanically connected to the fuselage 11 and the tilt axis 14 respectively, so as to achieve mechanical coupling and coordination between the fuselage 11 and the two wings 12. The damping device is electrically connected to the main control unit and is used to directly lock the wing tilt angle at the desired wing tilt angle state corresponding to the specific flight mode when the tilt-wing aircraft 10 needs to maintain a specific flight mode. This eliminates the need for the propulsion device 13 to continuously output a specific thrust difference to maintain the desired wing tilt angle state, thereby significantly reducing the aircraft's operating energy consumption in the aforementioned specific flight mode.

[0039] Furthermore, under the control of the main control unit, the damping device can focus on providing damping force to the fuselage 11 during wing tilting to suppress fuselage sway caused by the pendulum effect. This allows for independent control of the fuselage attitude stabilization control loop (which relies on the damping force provided by the damping device alone) and the wing tilt angle control loop (i.e., a closed-loop loop for wing tilt control based on the wing tilt attitude information detected by the first inertial measurement unit and by adjusting the thrust difference of the propulsion device). This achieves a decoupled control effect between the fuselage and wing, facilitating the mechanical coupling and coordination of these two control loops through the damping device. This ensures the overall stability of the aircraft attitude during takeoff, landing, and level flight transitions, eliminating the need for an additional high-power tilt drive device to simultaneously achieve wing tilt angle adjustment and fuselage attitude stabilization. This effectively reduces the overall weight and size of the aircraft and avoids problems such as phase lag, self-excited oscillation, and overshoot caused by fuselage-wing dynamic control coupling, thereby improving the accuracy and stability of wing tilt control.

[0040] In this embodiment, the damping device can switch normally between a locked state and an operating state under the control of the main control unit. When the damping device is in the locked state, it will fix the current actual tilt angle of the wing. When the damping device is in the operating state, it will provide a damping force to the fuselage 11 to suppress swaying. The damping device can be a fixed damper (which can be implemented using damping materials such as rubber pads or high-viscosity grease), which suppresses fuselage swaying by providing a large fixed damping force to the fuselage 11 without dynamic adjustment of the damping force. Alternatively, the damping device can be a variable damper (such as a magnetorheological damper, a mechanical friction damper, or a hydraulic damper), which suppresses fuselage swaying by providing dynamically changing damping force to the fuselage 11, ensuring that the fuselage 11 quickly and smoothly returns to a stable attitude.

[0041] Therefore, the tilt-wing aircraft 10 provided in this application is essentially a free-wing aircraft configuration that combines a suspension structure with a fuselage-wing attitude control decoupling mechanism. Unlike traditional tilt-wing aircraft configurations, it can fully leverage the advantages of the suspension structure and the fuselage-wing dual-loop control decoupling mechanism, freeing the aircraft from the constraints of the pendulum effect and the requirement for high-power tilt drive devices in traditional configurations. This can effectively reduce the overall weight and size of the aircraft and fundamentally avoid problems such as phase lag / self-excited oscillation / overshoot caused by fuselage-wing dynamic control coupling, thereby significantly improving the degree of freedom of wing tilt control and the flexibility of fuselage sway suppression, facilitating the achievement of free flight.

[0042] Optionally, please refer to the following: Figure 1 and Figure 5 ,inFigure 5 yes Figure 1 The diagram shows the control rudder distribution of the tiltrotor aircraft. In this embodiment, the tiltrotor aircraft 10 may further include a tail fin 15. The tail fin 15 may be implemented using any of the following tail fin structures: an inverted V-shaped tail fin structure, a V-shaped tail fin structure, or a conventional vertical tail fin structure.

[0043] In this embodiment, the tail fin 15 is mounted at the tail of the fuselage 11, and an aerodynamic control rudder 16 (including a pitch rudder and a rudder) is mounted on the tail fin 15. Each wing 12 has an aileron structure 17 mounted at its trailing edge. The main control unit is electrically connected to the pitch rudder, the rudder, and the aileron structure 17, respectively, and is used to control the motion states of the pitch rudder, the rudder, and the aileron structure 17 during the level flight phase of the aircraft (i.e., the flight phase in which the tiltrotor aircraft 10 is in level flight mode), so as to adjust the aircraft's pitch attitude, yaw attitude, and / or roll attitude. The pitch rudder is used to implement the aircraft's pitch attitude control function, the rudder is used to implement the aircraft's yaw attitude control function, and the aileron structure 17 is used to implement the aircraft's roll attitude control function. Optionally, in one embodiment of this invention, each wing 12 is equipped with an airflow control device (such as a micro vortex generator) at its leading edge to delay airflow separation when the wing 12 is at a large tilt angle.

[0044] Optionally, in this embodiment of the application, when the damping device is implemented using a variable damper, the tilt-wing aircraft 10 may further include a second inertial measurement unit mounted on the fuselage 11. Figure 1 (Not shown), wherein the second inertial measurement unit is used to detect fuselage pitch attitude information.

[0045] In this embodiment, the main control unit is electrically connected to the second inertial measurement unit (IMU) and is used to dynamically adjust the damping force provided by the damping device in real time during the takeoff, landing, and level flight transition of the aircraft, based on the fuselage pitch attitude information (including the actual fuselage pitch angle and the actual fuselage pitch angular velocity) detected by the second IMU, with the aim of converging the fuselage pitch angular velocity to zero, in order to suppress fuselage sway. The dynamic damping adjustment method for the damping device allows for a certain static tilt angle in the actual fuselage pitch angle (i.e., the actual fuselage pitch angle may not be zero), requiring only that the fuselage pitch angular velocity converge to zero. It does not require the fuselage attitude to be completely consistent with or absolutely aligned with the wing attitude, thus significantly differentiating it from the stringent requirements of traditional tiltrotor aircraft for "fuselage attitude level," substantially relaxing the tolerance for changes in the fuselage center of gravity due to dynamic factors such as fuselage load changes and aircraft attitude transitions.

[0046] In this application, to ensure that the tiltrotor aircraft 10 can achieve kinematic control decoupling between the fuselage attitude stabilization control and the wing tilt angle control during takeoff, landing, and level flight transitions, thereby ensuring overall aircraft attitude stability, this application provides an aircraft control method to achieve the aforementioned objective. The aircraft control method provided in this application will be described in detail below.

[0047] Please refer to Figure 6 , Figure 6 This is one of the flowcharts illustrating the aircraft control method provided in this application embodiment. In this application embodiment, the aircraft control method is applied to the main control unit included in the tilt-wing aircraft 10, and the aircraft control method may include steps S210 to S240.

[0048] Step S210: Obtain the aircraft takeoff and landing level flight conversion command and determine the desired wing tilt angle that matches the aircraft takeoff and landing level flight conversion command.

[0049] In this embodiment, the aircraft takeoff and landing level flight conversion command can be a first conversion command instructing the tiltrotor aircraft 10 to switch from the aircraft's vertical / ultra-short takeoff mode to the aircraft's level flight mode. In this case, the desired wing tilt angle corresponding to the first conversion command is the desired wing tilt angle pointed to by the aircraft's level flight mode. Alternatively, the aircraft takeoff and landing level flight conversion command can be a second conversion command instructing the tiltrotor aircraft 10 to switch from the aircraft's level flight mode to the aircraft's vertical / ultra-short takeoff mode. In this case, the desired wing tilt angle corresponding to the second conversion command is the desired wing tilt angle pointed to by the aircraft's vertical / ultra-short takeoff mode.

[0050] In step S220, the control damping device is released from locking the wing tilt angle, and the thrust of each of the multiple propulsion devices is adjusted according to the desired wing tilt angle, the actual airspeed of the aircraft, and the wing tilt attitude information detected in real time by the first inertial measurement unit, so that the wing tilts around the tilt axis to the desired wing tilt angle.

[0051] In this embodiment, when the main control unit receives the takeoff and landing level flight conversion command from the aircraft, it controls the damping device to switch from a locked state to an active state to release the lock on the wing tilt angle, allowing the wing 12 to tilt around the tilt axis 14 to the desired wing tilt angle under the coordinated action of the multiple propulsion devices 13. The tilt-wing aircraft 10 can detect its actual airspeed in real time by using an airspeed meter.

[0052] Based on this, the above step "adjusting the thrust of each of the multiple propulsion devices according to the desired wing tilt angle, the actual airspeed of the aircraft, and the wing tilt attitude information detected in real time by the first inertial measurement unit, so that the wing tilts around the tilt axis to the desired wing tilt angle" may include sub-steps A to D.

[0053] Sub-step A: Based on the actual wing tilt rate, calculate the theoretical thrust difference between the propulsion devices 13 corresponding to the upper and lower surfaces of the same wing 12.

[0054] In this embodiment, the theoretical thrust difference between the two sets of propulsion devices 13 of a single wing 12 can be expressed by the formula " "Calculations show that, among which..." This is used to represent the theoretical thrust difference between the lower wing surface propulsion device and the upper wing surface propulsion device of the corresponding wing 12. Used to represent the thrust differential control gain coefficient. This is used to represent the actual wing tilt rate. Specifically, when the aircraft's takeoff / landing / level flight transition command is the first transition command, the actual wing tilt rate is much greater than the desired wing tilt angle, and the actual wing tilt rate is in a positive tilt state (i.e., wing 12 tilts around the tilt axis 14 towards the fuselage nose). At this time, the corresponding theoretical thrust difference is negative (i.e., the thrust of the upper wing surface propulsion device needs to be increased, and the thrust of the lower wing surface propulsion device needs to be decreased). Conversely, when the aircraft's takeoff / landing / level flight transition command is the second transition command, the actual wing tilt rate is much less than the desired wing tilt angle, and the actual wing tilt rate is in a reverse tilt state (i.e., wing 12 tilts around the tilt axis 14 towards the fuselage tail). At this time, the corresponding theoretical thrust difference is positive (i.e., the thrust of the lower wing surface propulsion device needs to be increased, and the thrust of the upper wing surface propulsion device needs to be decreased).

[0055] Sub-step B: Based on the actual airspeed of the aircraft and the actual wing tilt angle, calculate the wing tilt thrust difference for safe flight of the aircraft based on the theoretical thrust difference.

[0056] In this embodiment, after the main control unit obtains the actual airspeed and actual wing tilt angle of the tilt-wing aircraft 10, it determines whether the tilt-wing aircraft 10 will face serious flight risks (e.g., aircraft crash risk, wing stall risk, etc.) during wing tilting based on the actual airspeed and actual wing tilt angle. Based on the corresponding determination result, the calculated theoretical thrust difference is numerically corrected to obtain the wing tilt thrust difference that ensures the tilt-wing aircraft 10 maintains a safe flight state. Sub-step B may include: Detect whether the actual airspeed of the aircraft and the actual wing tilt angle meet the dangerous conditions for wing rollover. If the actual airspeed of the aircraft and the actual wing tilt angle meet the dangerous conditions of wing tilt, a preset safe tilt coefficient is used to weight the theoretical thrust difference to obtain the wing tilt thrust difference, wherein the preset safe tilt coefficient is a positive decimal less than 1 (e.g., 0.8). If the actual airspeed of the aircraft and the actual wing tilt angle do not meet the dangerous conditions for wing tilt, the theoretical thrust difference is directly used as the wing tilt thrust difference.

[0057] The dangerous conditions for wing tilting can be, but are not limited to, "wing tilt angles within the range of 35° to 60° and aircraft airspeed less than 15 m / s", "wing tilt angles within the range of 70° to 80° and aircraft airspeed less than 4 m / s", and "wing tilt angles within the range of 20° to 30° and aircraft airspeed less than 23 m / s".

[0058] Furthermore, it is understood that, in order to avoid serious risks caused by wing tilting maneuvers when the tilt-wing aircraft 10 meets the dangerous conditions for wing tilting, the actual wing tilting angular velocity used in calculating the theoretical thrust difference can be constrained to below 0.3 rad / s to improve the flight safety of the tilt-wing aircraft 10.

[0059] Sub-step C: Calculate the actual tilt angle difference between the desired wing tilt angle and the actual wing tilt angle, and calculate the desired propulsion speed of each of the multiple propulsion devices 13 based on the wing tilt thrust difference, with the goal of the actual tilt angle difference converging to zero.

[0060] In this embodiment, after obtaining the wing tilt thrust difference that ensures safe flight of the aircraft, the main control unit will use the formula " "Calculate the current desired compensation wing tilting moment of the tilt-wing aircraft 10 (i.e. ),in Used to represent the wing tilt thrust difference Used to indicate the dynamic torque arm (i.e., the effective distance from the point of application of the thrust to the tilt axis 14). The initial lever arm used to represent the point of application of the thrust difference Used to represent the actual wing tilt angle This is used to represent the maximum lateral offset distance of any propulsion device 13 relative to the tilt axis 14. Then, the main control unit employs a PID (Proportion-Integral-Differential) control algorithm or a deep learning method (such as DQN (Deep Q-Network) learning method) to predict the expected output thrust of each of the multiple propulsion devices 13, with the aim of "the actual tilt angle difference converging to zero and completing the torque compensation operation for the expected wing tilt moment." For each propulsion device 13, the unit calculates the expected propulsion speed matching its own expected output thrust according to a preset thrust-speed conversion relationship.

[0061] Sub-step D: Adjust the rotational speed of each of the multiple propulsion devices 13 according to the calculated desired propulsion speed.

[0062] Therefore, by executing the above sub-steps A to D, this application ensures that the corresponding tilt-wing aircraft 10 achieves a safe and stable wing tilt control effect.

[0063] Step S230: During wing tilting, control the damping device to provide damping force to the fuselage to suppress fuselage sway.

[0064] In this embodiment, during the process of controlling the multiple sets of propulsion devices 13 to drive the wing 12 to tilt to the desired wing tilt angle, the main control unit will simultaneously drive the damping device to provide damping force to the fuselage 11 in the working state to suppress the fuselage sway caused by the pendulum effect. This allows the fuselage attitude stabilization control link and the wing tilt angle control link to be controlled independently, facilitating the mechanical coupling and coordination of these two control links through the damping device, so as to jointly ensure the overall stability of the aircraft attitude during the take-off, landing and level flight transition.

[0065] Optionally, if the damping device is implemented using a variable damper and the tiltwing aircraft 10 includes a second inertial measurement unit, the second inertial measurement unit will detect the fuselage pitch attitude information of the fuselage 11 in real time. The fuselage pitch attitude information may include the actual fuselage pitch angle and the actual fuselage pitch angular velocity of the fuselage 11. In this case, the above step S230 may include sub-steps E to F.

[0066] Sub-step A: Based on the actual fuselage pitch angle and the actual fuselage pitch angular velocity, predict the damping force with the aim of the fuselage pitch angular velocity converging to zero, and obtain the corresponding sway suppression damping force.

[0067] In this embodiment, the main control unit can use a PD (Proportion-Differential) control algorithm to predict the damping force based on the actual fuselage pitch angle and the actual fuselage pitch velocity, and use the corresponding predicted damping force value as the oscillation suppression damping force to drive the damping device to adjust the damping force until the damping force control convergence condition "the actual fuselage pitch velocity is set to zero" is met. The aforementioned damping force prediction operation can employ a control law... "Achieve, among which" Used to represent the predicted value of the damping force. Used to represent the proportional control coefficient of damping output. Used to represent the differential control coefficient of damped output. Used to represent the actual pitch angle. This is used to represent the actual fuselage pitch rate. In one embodiment of this invention, to improve the reliability of damping force control and enhance the fuselage sway suppression effect, the aforementioned damping force control convergence condition can be set to "the actual fuselage pitch rate is set to zero, and the actual fuselage pitch acceleration of the fuselage 11 is set to zero", wherein the actual fuselage pitch acceleration can be detected by the second inertial measurement unit.

[0068] Sub-step F: Drive the damping device to operate according to the swing suppression damping force.

[0069] Therefore, by executing the above sub-steps E~F, this application can dynamically adjust the damping force provided to the fuselage 11, so that the fuselage 11 can quickly and smoothly restore its attitude stability during wing tilting.

[0070] In step S240, when the wing tilts to the desired wing tilt angle, the control damping device locks and maintains the wing tilt angle.

[0071] In this embodiment, after the main control unit drives the wing 12 to tilt to the desired wing tilt angle using the multiple sets of propulsion devices 13, it controls the damping device to switch from the working state to the locked state, so as to lock and maintain the actual wing tilt angle at the desired wing tilt angle, ensuring that the tilt-wing aircraft 10 successfully enters the desired flight mode pointed to by the aircraft take-off and landing level flight conversion command (for example, the aircraft level flight mode corresponding to the first conversion command, or the aircraft vertical / ultra-short take-off landing mode corresponding to the second conversion command), and can realize normal flight attitude control function in the corresponding desired flight mode.

[0072] Therefore, by executing the above steps S210 to S240, this application can achieve kinematic control decoupling between the fuselage attitude stabilization control link and the wing tilt angle control link during the take-off, landing and level flight transition of the aircraft, thereby ensuring the overall stability of the aircraft attitude.

[0073] Alternatively, please refer to Figure 7 , Figure 7 This is a second schematic flowchart of the aircraft control method provided in this application embodiment. In this application embodiment, [the method is]... Figure 6 Compared to the aircraft control method shown, Figure 7 The aircraft control method shown may also include steps S250 to S260 to ensure that the tiltrotor aircraft 10 can take off and land normally in the aircraft's vertical / ultra-short takeoff and landing mode.

[0074] Step S250: When the wing tilt angle maintained by the damping device is within the preset takeoff and landing tilt angle, obtain the aircraft takeoff command or the aircraft landing command.

[0075] In this embodiment, the preset takeoff and landing tilt angle can be the desired wing tilt angle for the aircraft's vertical takeoff and landing mode or ultra-short takeoff mode, or the desired wing tilt angle for the aircraft's ultra-short takeoff mode. Specifically, when the wing tilt angle maintained by the damping device is the desired wing tilt angle for the aircraft's ultra-short takeoff mode, the takeoff command received by the tilt-wing aircraft 10 from the user is an ultra-short takeoff command. At this time, the desired takeoff speed recorded by the ultra-short takeoff command is the minimum safe lift speed (approximately 60% of the preset vertical takeoff speed). When the wing tilt angle maintained by the damping device is the desired wing tilt angle for the aircraft's vertical takeoff and landing mode or ultra-short takeoff mode, the tilt-wing aircraft 10 can receive takeoff commands (which are vertical takeoff commands) or landing commands (which can be vertical landing commands or ultra-short takeoff commands) from the user. The desired takeoff speed recorded by the vertical takeoff command is the preset vertical takeoff speed, the desired landing speed recorded by the vertical landing command is the preset safe vertical landing speed, and the desired landing speed recorded by the ultra-short takeoff command is the preset safe taxi landing speed. In addition, the aforementioned aircraft takeoff or landing commands can all record aircraft pitch / roll / yaw attitude adjustment commands used to drive the propulsion device to adjust thrust.

[0076] Step S260: Based on the desired takeoff speed included in the aircraft takeoff command or the desired landing speed included in the aircraft landing command, adjust the thrust of each of the multiple propulsion devices to enable the tiltrotor aircraft to take off at the desired takeoff speed or to enable the tiltrotor aircraft to land at the desired landing speed.

[0077] Therefore, by executing the above steps S250 to S260, this application can ensure that the tilt-wing aircraft 10 can take off and land normally in the vertical / ultra-short takeoff and landing mode.

[0078] The above descriptions are merely various embodiments of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A tilt-wing aircraft, characterized in that, The aircraft includes a fuselage, a first inertial measurement unit, a main control unit, a damping device, two wings, and multiple sets of propulsion devices. Each wing has a set of propulsion devices installed on its upper and lower surfaces, and the two sets of propulsion devices on the same wing are symmetrical to each other. The first inertial measurement unit is used to detect the tilt attitude information of the wings. Two wings are symmetrically arranged on both sides of the fuselage and are rotatably connected to the fuselage via a tilt axis. The tilt axis is installed on the fuselage above the fuselage's center of gravity, and the tilt axis is installed on the corresponding wing near the wing's aerodynamic center. The main control unit is electrically connected to the first inertial measurement unit and multiple propulsion devices respectively, and is used to adjust the thrust of each of the multiple propulsion devices in real time according to the wing tilt attitude information detected by the first inertial measurement unit during the take-off, landing and level flight transition of the aircraft, so as to make the wing tilt around the tilt axis. The damping device is mechanically connected to the fuselage and the tilt axis respectively, and electrically connected to the main control unit. It is used to lock, maintain or unlock the wing tilt angle under the control of the main control unit, or to provide damping force to the fuselage during wing tilting to suppress fuselage sway.

2. The aircraft according to claim 1, characterized in that, The distance between the installation position of the tilt axis on the corresponding wing and the leading edge of the wing ranges from 20% to 40% of the wing chord length. The distance from the mounting position of the tilt axis on the fuselage to the vertical line of the fuselage center of gravity is within the range of 0 to 10% of the wing chord length.

3. The aircraft according to claim 2, characterized in that, The tilting shaft is mounted on the fuselage at a position on the vertical line of the fuselage's center of gravity; The center of mass of the two wings and the multiple propulsion devices is located at the center point of the tilt axis.

4. The aircraft according to claim 1, characterized in that, The aircraft also includes a tail fin; The tail fin is mounted at the rear of the fuselage, and the tail fin is equipped with pitch and rudder. Each wing is equipped with an aileron structure at its trailing edge. The main control unit is electrically connected to the pitch control, the rudder, and the aileron structure, respectively, and is used to control the motion state of the pitch control, the rudder, and the aileron structure during the level flight phase of the aircraft, so as to adjust the pitch attitude, yaw attitude, and / or roll attitude of the aircraft. The pitch control is used to realize the pitch attitude control function of the aircraft, the rudder is used to realize the yaw attitude control function of the aircraft, and the aileron structure is used to realize the roll attitude control function of the aircraft.

5. The aircraft according to any one of claims 1-4, characterized in that, The aircraft also includes a second inertial measurement unit mounted on the fuselage, wherein the second inertial measurement unit is used to detect the fuselage pitch attitude information; The main control unit is electrically connected to the second inertial measurement unit and is used to dynamically adjust the damping force provided by the damping device in real time during the take-off, landing and level flight transition of the aircraft, based on the fuselage pitch attitude information detected by the second inertial measurement unit, with the aim of converging the fuselage pitch angular velocity to zero, so as to suppress fuselage sway.

6. An aircraft control method, characterized in that, The control method, which utilizes the main control unit of the tiltwing aircraft according to any one of claims 1-5, comprises: Obtain the aircraft takeoff and landing level flight transition command, and determine the desired wing tilt angle that matches the aircraft takeoff and landing level flight transition command; The control damping device releases the lock on the wing tilt angle, and adjusts the thrust of each of the multiple propulsion devices according to the desired wing tilt angle, the actual airspeed of the aircraft, and the wing tilt attitude information detected in real time by the first inertial measurement unit, so that the wing tilts around the tilt axis to the desired wing tilt angle. During wing tilting, the damping device is controlled to provide damping force to the fuselage to suppress fuselage sway; When the wing tilts to the desired wing tilt angle, the damping device is controlled to lock and maintain the wing tilt angle.

7. The control method according to claim 6, characterized in that, The wing tilt attitude information detected in real time by the first inertial measurement unit includes the actual wing tilt angle and the actual wing tilt angular velocity. The step of adjusting the thrust of each of the multiple propulsion devices based on the desired wing tilt angle, the actual airspeed of the aircraft, and the wing tilt attitude information detected in real time by the first inertial measurement unit, so that the wing tilts around the tilt axis to the desired wing tilt angle, includes: Based on the actual wing tilt rate, calculate the theoretical thrust difference between the propulsion devices corresponding to the upper and lower surfaces of the same wing. Based on the actual airspeed of the aircraft and the actual wing tilt angle, the wing tilt thrust difference during safe flight of the aircraft is calculated based on the theoretical thrust difference. Calculate the actual tilt angle difference between the desired wing tilt angle and the actual wing tilt angle, and calculate the desired propulsion speed of each of the multiple propulsion devices based on the wing tilt thrust difference, with the goal of converging the actual tilt angle difference to zero. According to the calculated desired propulsion speed of each of the multiple propulsion devices, the speed of each of the multiple propulsion devices is adjusted.

8. The control method according to claim 7, characterized in that, The step of calculating the tilt thrust difference for safe flight of the aircraft based on the actual airspeed and the actual wing tilt angle, and on the theoretical thrust difference, includes: Detect whether the actual airspeed of the aircraft and the actual wing tilt angle meet the dangerous conditions for wing rollover. If the actual airspeed of the aircraft and the actual wing tilt angle meet the dangerous conditions of wing tilt, the theoretical thrust difference is weighted by a preset safe tilt coefficient to obtain the wing tilt thrust difference, wherein the preset safe tilt coefficient is a positive decimal less than 1. If the actual airspeed of the aircraft and the actual wing tilt angle do not meet the dangerous conditions for wing tilt, the theoretical thrust difference is directly used as the wing tilt thrust difference.

9. The control method according to claim 6, characterized in that, In the case where the tiltwing aircraft includes a second inertial measurement unit, and the fuselage pitch attitude information detected by the second inertial measurement unit includes the actual fuselage pitch angle and the actual fuselage pitch angular velocity, then the step of controlling the damping device to provide damping force to the fuselage to suppress fuselage sway includes: Based on the actual fuselage pitch angle and the actual fuselage pitch angular velocity, the damping force is predicted with the aim of the fuselage pitch angular velocity converging to zero, and the corresponding sway suppression damping force is obtained. The damping device is driven to operate according to the swing suppression damping force.

10. The control method according to any one of claims 6-9, characterized in that, The control method further includes: When the wing tilt angle maintained by the damping device is within the preset takeoff and landing tilt angle, the aircraft takeoff command or aircraft landing command is obtained. Based on the desired takeoff speed included in the aircraft takeoff command or the desired landing speed included in the aircraft landing command, the thrust of each of the multiple propulsion devices is adjusted so that the tiltrotor aircraft takes off at the desired takeoff speed or lands at the desired landing speed.

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