Free wing aircraft and aircraft control method

Abstract

The invention provides a free wing 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, and the wing tilting function is directly achieved through vector thrust generated by the multiple sets of flight propelling devices on the aircraft wings. Meanwhile, a fuselage active stabilizing device on the aircraft fuselage provides vector thrust to actively inhibit fuselage pitching motion caused by the pendulum bob effect, so that the independent control effect of a fuselage attitude stability maintaining control link and a wing tilting angle control link is achieved based on a suspension structure; the kinematics control decoupling effect and the physical structure decoupling effect between fuselage attitude control and wing tilting control are ensured, and the stability of the overall attitude of the aircraft can be effectively guaranteed on the basis of reducing the overall weight and size of the aircraft through the vector thrust synergistic effect of the flight propelling device and the fuselage active stabilizing device.

Description

TECHNICAL FIELD

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

[0002] With the continuous development of aviation technology, the tilt wing aircraft gradually becomes an important research direction of current aviation technology because it can realize free switching between the helicopter mode and the fixed wing mode by controlling the wing tilt to balance the high-speed cruising ability of the fixed wing aircraft and the vertical take-off / short take-off ability of the helicopter.

[0003] At present, the traditional tilt wing aircraft mainly realizes flight mode switching operation (including switching from the aircraft level flight mode to the aircraft hovering mode or the aircraft vertical descent mode, switching from the aircraft hovering mode to the aircraft level flight mode or the aircraft short landing mode, switching from the aircraft short take-off mode to the aircraft hovering mode, etc.) through the wing tilt action, and in the flight process (including flight mode switching operation) of the tilt wing aircraft, the pendulum effect caused by serious transition coupling easily affects the attitude stability of the fuselage, so the single-loop control algorithm considering the dynamic control coupling relationship of the wing tilt control and the fuselage attitude control is needed for motion compensation to balance the wing tilt effect and the fuselage attitude stability effect as much as possible. However, it is worth noting that this aircraft control mechanism cannot balance between the control algorithm complexity and the control system response speed, and often cannot suppress the fuselage swing caused by the pendulum effect under 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 tilt driving device to balance the realization of the wing tilt angle adjustment function and the fuselage attitude stability function, resulting in a large overall weight and size of the aircraft, and the corresponding wing tilt control precision and stability are poor due to the limitation of the fuselage-wing dynamic control coupling relationship. SUMMARY

[0004] Therefore, the free wing aircraft and the aircraft control method can suspend the aircraft body below the aircraft wing through the tilting shaft, directly realize the wing tilting function through the vector thrust generated by the multiple sets of flight propulsion devices on the aircraft wing, and actively inhibit the body pitching motion caused by the pendulum effect through the vector thrust provided by the body active stabilizing device on the aircraft body, thereby realizing the independent control effect of the body attitude stability control link and the wing tilting angle control link based on the suspension structure, ensuring the kinematic control decoupling between the body attitude control and the wing tilting control (i.e., realizing the body-wing double-loop decoupling control effect), completely eliminating the need for the special mechanical tilting drive mechanism which is necessary, heavy and complex in the conventional tilting wing aircraft, significantly reducing the weight and simplifying the structure of the aircraft, and physically decoupling the wing tilting control and the body attitude control (i.e., realizing the physical structure decoupling effect between the body attitude control and the wing tilting control), which can fundamentally avoid the problems of phase lag, self-excited oscillation and overshoot caused by the body-wing dynamic control coupling, and effectively improve the wing tilting control precision and stability. Meanwhile, the vector thrust of the physically decoupled flight propulsion device and the body active stabilizing device can be used to ensure the stability of the overall attitude of the aircraft.

[0005] To achieve the above object, the technical solutions adopted by the embodiments of the present application are as follows: In a first aspect, the present application provides a free wing aircraft, which comprises a body, a first inertial measurement unit, a second inertial measurement unit, a master control unit, two wings, a body active stabilizing device and multiple sets of flight propulsion devices. The two wings are symmetrically arranged on both sides of the body and are rotationally connected to the body through a tilting shaft, wherein at least one set of flight propulsion devices is separately installed on each wing, and the flight propulsion devices installed on the two wings are symmetric about the body axis of the body, the first inertial measurement unit is used to detect wing tilting attitude information, and the second inertial measurement unit is used to detect body pitching attitude information. The master control unit is electrically connected to the first inertial measurement unit and the multiple sets of flight propulsion devices, respectively, and is used to regulate and control the vector thrust generated by each set of flight propulsion devices in real time according to the wing tilting attitude information detected by the first inertial measurement unit during flight mode switching, so that the wing tilts around the tilting shaft. The master control unit is electrically connected to the second inertial measurement unit and the body active stabilizing device, respectively, and is used to regulate and control the vector thrust generated by the body active stabilizing device according to the body pitching attitude information detected by the second inertial measurement unit, so as to inhibit the body pitching motion caused by the pendulum effect, wherein the body active stabilizing device is installed on the body.

[0006] In optional embodiments, at least one set of flight propulsion devices is evenly installed on the upper and lower surfaces of each wing, wherein each set of flight propulsion devices provides vector thrust by at least one fixed-pitch rotor or variable-pitch rotor.

[0007] In optional embodiments, at least one set of flight propulsion devices is evenly installed on the leading edge of each wing, wherein each set of flight propulsion devices provides vector thrust by cyclic variable-pitch rotor.

[0008] In optional embodiments, at least one set of flight propulsion devices is evenly installed on the leading edge of each wing, and aileron structure is installed on the trailing edge of each wing, wherein each set of flight propulsion devices provides vector thrust by fixed-pitch rotor; The main control unit is also electrically connected with the aileron structure, and is configured to control the deflection angle and / or deflection direction of the aileron structure during flight mode switching, so that the aileron structure assists the multiple sets of flight propulsion devices to adjust the tilting torque acting on the wing, to drive the wing to tilt around the tilting axis.

[0009] In optional embodiments, the aircraft further comprises a wing tilting locking device; The main control unit is electrically connected with the wing tilting locking device, and is configured to control the wing tilting locking device to lock the tilting angle of the wing or release the tilting angle of the wing.

[0010] In optional embodiments, the aircraft further comprises a tail; The tail is installed at the tail position of the fuselage, and the tail is provided with a hydrodynamic control surface, and the trailing edge of each wing is provided with aileron structure, wherein the hydrodynamic control surface comprises a pitch control surface and a yaw control surface; The main control unit is electrically connected with the pitch control surface, the yaw control surface and the aileron structure respectively, and is configured to control the motion state of each of the pitch control surface, the yaw control surface and the aileron structure, to adjust the pitch attitude of the aircraft, the yaw attitude of the aircraft and / or the roll attitude of the aircraft.

[0011] In a second aspect, the application provides an aircraft control method, which is applied to the main control unit included in the free-wing aircraft according to any one of the preceding embodiments, and the control method comprises: obtaining a flight mode switching instruction, and determining an expected wing tilting angle corresponding to a target flight mode pointed by the flight mode switching instruction; controlling the vector thrust generated by each set of flight propulsion devices according to the expected wing tilting angle and the wing tilting attitude information detected by the first inertia measurement unit in real time, to make the wing tilt around the tilting axis to the expected wing tilting angle; According to the body pitch attitude information detected by the second inertial measurement unit in real time, the body active stabilizing device is controlled to provide vector thrust to the body to suppress the body pitch motion caused by the pendulum effect.

[0012] In optional embodiments, the wing tilt attitude information includes an actual wing tilt angle and an actual wing tilt angular velocity, and the step of regulating the vector thrust generated by each of the plurality of flight propulsion devices according to the desired wing tilt angle and the wing tilt attitude information detected by the first inertial measurement unit in real time, so as to tilt the wing around the tilt axis to the desired wing tilt angle, includes: calculating an actual angle difference between the desired wing tilt angle and the actual wing tilt angle, and calculating a desired tilt torque acting on the wing based on the actual wing tilt angular velocity, with the purpose of converging the actual angle difference to zero; adjusting the vector thrust generated by each of the plurality of flight propulsion devices according to the desired tilt torque.

[0013] In optional embodiments, the step of controlling the body active stabilizing device to provide vector thrust to the body according to the body pitch attitude information detected by the second inertial measurement unit in real time, so as to suppress the body pitch motion caused by the pendulum effect, includes: determining a desired body pitch angle and a desired body pitch angular velocity corresponding to the target flight mode; calculating a desired pitch torque required for the body to achieve the desired body pitch angle and the desired body pitch angular velocity according to the body pitch attitude information, the desired body pitch angle and the desired body pitch angular velocity; regulating the vector thrust provided by the body active stabilizing device according to the desired pitch torque.

[0014] In optional embodiments, the body pitch attitude information includes an actual body pitch angle and an actual body pitch angular velocity, and the step of calculating the desired pitch torque required for the body to achieve the desired body pitch angle and the desired body pitch angular velocity according to the body pitch attitude information, the desired body pitch angle and the desired body pitch angular velocity, includes: calculating an actual angle difference between the desired body pitch angle and the actual body pitch angle, and calculating an actual angular velocity difference between the desired body pitch angular velocity and the actual body pitch angular velocity; adopting a PID (Proportion Integral Differential) algorithm to solve the pitch torque with the purpose of converging the actual angle difference to zero and converging the actual angular velocity difference to zero, so as to obtain the desired pitch torque.

[0015] In this case, the beneficial effects of the embodiments of the present application can include the following: The present application can suspend the aircraft fuselage below the aircraft wings by symmetrically arranging a pair of aircraft wings on both sides of the aircraft fuselage and connecting the aircraft fuselage with the aircraft wings through a tilting shaft, at this time, the wing tilting function can be directly realized by a plurality of groups of flight propulsion devices located on the aircraft wings through the generation of vector thrust, and the fuselage active stabilization device deployed on the aircraft fuselage can actively suppress the fuselage pitch motion caused by the pendulum effect through the provision of vector thrust, thereby achieving the independent control effect of the fuselage attitude control link (i.e. the closed loop of the fuselage attitude control through the adjustment of the vector thrust generated by the fuselage active stabilization device based on the fuselage pitch attitude information detected by the second inertial measurement unit) and the wing tilting angle control link (i.e. the closed loop of the wing tilting control through the adjustment of the vector thrust generated by the flight propulsion device based on the wing tilting attitude information detected by the first inertial measurement unit), ensuring the kinematic decoupling between the fuselage attitude control and the wing tilting control (i.e. achieving the double-loop decoupling control effect of the fuselage-wing), completely eliminating the need for the traditional tilt-wing aircraft for the special mechanical tilting drive mechanism that needs to simultaneously consider the wing tilting angle adjustment function and the fuselage attitude control function, not only significantly reducing the weight of the aircraft and simplifying the structure of the aircraft, but also physically decoupling the wing tilting control from the fuselage attitude control (i.e. achieving the physical structure decoupling effect between the fuselage attitude control and the wing tilting control), which can effectively improve the wing tilting control precision and stability by avoiding the problems of phase lag / self-excited oscillation / overshoot caused by the fuselage-wing dynamics control coupling. In addition, the present application can ensure the stability of the overall attitude of the aircraft through the synergistic effect of the vector thrust of the flight propulsion device and the fuselage active stabilization device (including but not limited to the vector thrust component in the wing tilting plane or the fuselage pitch plane perpendicular to the axis direction of the tilting shaft).

[0016] In order to make the above-mentioned purposes, features and advantages of the present application more obvious and easy to understand, the following preferred embodiments are described in detail below, and the accompanying drawings are described as follows. BRIEF DESCRIPTION OF DRAWINGS

[0017] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following will briefly introduce the drawings needed in the embodiments, it should be understood that the following drawings only show some embodiments of the present application, therefore should not be regarded as a limitation on the scope, for those skilled in the art, without creative labor, other related drawings can also be obtained from these drawings.

[0018] Figure 1Figure 1 is a schematic diagram of the overall assembly of a first embodiment of a free wing aircraft according to the present application; Figure 2 Figure 2 is a schematic diagram of the overall assembly of a second embodiment of a free wing aircraft according to the present application; Figure 3 Figure 3 is a schematic diagram of the overall assembly of a third embodiment of a free wing aircraft according to the present application; Figure 4 Figure 4 is a schematic diagram of the overall assembly of a fourth embodiment of a free wing aircraft according to the present application; Figure 5 Figure 5 is a schematic diagram of the overall assembly of a fifth embodiment of a free wing aircraft according to the present application; Figure 6 Figure 6 is a schematic diagram of the overall assembly of a sixth embodiment of a free wing aircraft according to the present application; Figure 7 Figure 7 is a schematic diagram of the overall assembly of a seventh embodiment of a free wing aircraft according to the present application; Figure 8 Figure 8 is a schematic diagram of the overall assembly of an eighth embodiment of a free wing aircraft according to the present application; Figure 7 Figure 9 is a schematic diagram of the flow of sub-steps included in step S230.

[0019] Figure 10: 10 - free wing aircraft; 11 - fuselage; 12 - wing; 13 - flight propulsion device; 14 - fuselage active stabilizing device; 15 - tail; 16 - aerodynamic control rudder; 17 - aileron structure. DETAILED DESCRIPTION

[0020] In order to make the objectives, technical solutions, and advantages of the embodiments of the present application clearer, the following will be combined with the accompanying drawings for the embodiments of the present application to make a clear and complete description of the technical solutions in the embodiments of the present application. Obviously, the described embodiments are some but not all of the embodiments of the present application. The components of the embodiments of the present application described and shown in the accompanying drawings can be arranged and designed in various different configurations.

[0021] Therefore, the following detailed description of the embodiments of the present application provided in the accompanying drawings is not intended to limit the scope of the claimed present application, but only represents selected embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative work are within the scope of protection of the present application.

[0022] It should be noted that: similar reference numerals and letters represent similar items in the following drawings, therefore, once an item is defined in one drawing, it does not need to be further defined and explained in subsequent drawings.

[0023] In the description of the application, it needs to be understood that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer" and the like indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, or the orientation or positional relationship commonly understood by those skilled in the art, and are only used to facilitate the description of the application and simplify the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation of the application.

[0024] In the description of the application, it also needs to be explained that, unless otherwise explicitly specified and limited, the terms "set", "install", "connect", "connect" should be understood broadly, for example, it can be fixedly connected, or it can be detachably connected, or integrally connected; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium, or it can be the communication inside two elements. For those skilled in the art, the specific meaning of the above terms in the application can be understood according to the specific circumstances.

[0025] In addition, in the description of the application, it can be understood that the relationship terms such as "first" and "second" and the like are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between the entities or operations. Moreover, the terms "include", "contain" or any other variants thereof are intended to cover non-exclusive inclusion, so that the process, method, article or device including a series of elements not only includes those elements, but also includes other elements not explicitly listed or inherent to such process, method, article or device. Without more limitation, the element defined by the sentence "including a…" does not exclude the presence of other identical elements in the process, method, article or device including the element. For those skilled in the art, the specific meaning of the above terms in the application can be understood according to the specific circumstances.

[0026] Some embodiments of the application will be described in detail below with reference to the accompanying drawings. The following examples and features in the examples can be combined with each other without conflict.

[0027] Please refer to Figure 1 , Figure 2 , Figure 3 , Figure 4 , Figure 5 and Figure 6 , wherein Figure 1 is one of the overall assembly schematic diagrams of the first free wing aircraft 10 provided by the embodiments of the application, Figure 2This is the second overall assembly schematic diagram of the first type of free-wing aircraft 10 provided in the embodiments of this application. Figure 3 This is one of the overall assembly schematic diagrams of the second type of free-wing aircraft 10 provided in the embodiments of this application. Figure 4 This is the second overall assembly schematic diagram of the second type of free-wing aircraft 10 provided in the embodiments of this application. Figure 5 This is one of the overall assembly schematic diagrams of the third type of free-wing aircraft 10 provided in the embodiments of this application. Figure 6 This is the second overall assembly schematic diagram of the third type of free-wing aircraft 10 provided in this application embodiment. In this application embodiment, the free-wing aircraft 10 can switch between multiple flight modes (including level flight mode, vertical takeoff / landing mode, ultra-short takeoff / landing mode, hovering mode, etc.), and actively compensates and suppresses the fuselage sway phenomenon caused by the pendulum effect during the switching of different flight modes or during the maintenance of the desired flight mode, so as to effectively ensure the stability of the overall attitude of the aircraft. The free-wing aircraft 10 may include a fuselage 11, a first inertial measurement unit (attached... Figures 1-6 (Not shown), main control unit (attached) Figures 1-6 (None shown), Second Inertial Measurement Unit (attached) Figures 1-6 (Not shown), two wings 12, multiple sets of flight propulsion devices 13 and active fuselage stabilization device 14.

[0028] In this embodiment, each wing 12 is individually equipped with at least one set of flight propulsion devices 13. The two wings 12 are symmetrically arranged on both sides of the fuselage 11 and connected by a tilt axis (attached). Figures 1-6 (Not shown) is rotatably connected to the fuselage 11, allowing the fuselage 11 to be suspended below the two wings 12, thereby achieving a physical decoupling effect between fuselage attitude control and wing tilt control through the suspension structure. All flight propulsion devices 13 mounted on each of the two wings 12 are symmetrical about the fuselage axis (i.e., the axis running through the nose and tail of the fuselage) of the fuselage 11 (i.e., all flight propulsion devices 13 mounted on one wing 12 are symmetrical about the fuselage axis of the other wing 12). One end of the tilt shaft is fixedly connected to the end face of one wing away from the wingtip, and the other end of the tilt shaft passes through the fuselage 11 or is fixedly connected to the end face of the other wing 12 away from the wingtip through a pivot housing structure fixedly mounted on the fuselage 11, so that the two wings 12 maintain the same wing tilt attitude in real time at the free-wing aircraft 10.

[0029] The first inertial measurement unit can be arranged on any one of the wings 12 to detect the wing tilting attitude information (including actual wing angle and actual wing tilting angular velocity information) in real time. Each set of flight propulsion devices 13 can independently generate a vector thrust. The thrust size and / or thrust direction of the vector thrust generated by different flight propulsion devices 13 can be the same or different. The thrust size and / or thrust direction of the overall vector thrust corresponding to each of the upper and lower wing surfaces of a single wing 12 is adjusted to enable all flight propulsion devices 13 to cooperatively generate flight vehicle flight power (for example, flight vehicle lift, flight vehicle yaw power, flight vehicle roll power, or flight vehicle forward power) for adjusting the flight vehicle pose or wing tilting torque for driving the two wings 12 to tilt about the tilting axis. Each set of flight propulsion devices 13 includes at least one propeller, which can be but is not limited to a fluid propeller structure such as a rotor or a ducted fan.

[0030] Optionally, in the first implementation of the embodiment of the present application, Figure 1 and Figure 2 For example, each wing 12 of the free wing flight vehicle 10 is uniformly provided with at least one set of flight propulsion devices 13 on the upper and lower wing surfaces respectively. The flight propulsion devices 13 arranged on the upper and lower wing surfaces of the same wing 12 are symmetrical to each other, and the axis of the tilting axis is located on the symmetrical plane between the flight propulsion devices 13 on the upper and lower wing surfaces of each wing 12. Each set of flight propulsion devices 13 can provide a vector thrust by using at least one fixed-pitch rotor or variable-pitch rotor. When a single set of flight propulsion devices 13 includes multiple fixed-pitch rotors or variable-pitch rotors, the multiple fixed-pitch rotors or variable-pitch rotors can be fixedly installed on the corresponding wing 12 by using the same connecting bracket.

[0031] In this process, when the flight propulsion devices 13 in Figure 1 or Figure 2 are constructed by using fixed-pitch rotors, the wing tilting torque for driving the corresponding wing 12 to tilt about the tilting axis towards the tail or the head of the fuselage can be generated by adjusting the difference between the vector thrust on the upper wing surface and the vector thrust on the lower wing surface of a single wing 12. If the vector thrust on the upper wing surface of a single wing 12 is larger than the vector thrust on the lower wing surface of the wing 12, the corresponding wing tilting torque is used to drive the corresponding wing 12 to tilt about the tilting axis towards the head of the fuselage. If the vector thrust on the upper wing surface of a single wing 12 is smaller than the vector thrust on the lower wing surface of the wing 12, the corresponding wing tilting torque is used to drive the corresponding wing 12 to tilt about the tilting axis towards the tail of the fuselage.

[0032] When the flight propulsion devices 13 in Figure 1 orFigure 2 When the flight propulsion device 13 is in use, the wing tilting moment that drives the corresponding wing 12 to tilt around the tilting axis to the tail or nose of the fuselage can be generated by adjusting the magnitude and / or the difference in the vector thrust difference between the upper and lower wing surfaces of a single wing 12. For a single wing 12, the blade pitch of each set of flight propulsion devices 13 on the corresponding wing 12 can be adjusted to increase the blade angle of attack of all flight propulsion devices 13 on the upper wing surface and decrease the blade angle of attack of all flight propulsion devices 13 on the lower wing surface, so that the generated wing tilting moment can be used to drive the corresponding wing 12 to tilt around the tilt axis towards the nose of the fuselage, or the blade angle of attack of all flight propulsion devices 13 on the upper wing surface and increase the blade angle of attack of all flight propulsion devices 13 on the lower wing surface can be used to drive the corresponding wing 12 to tilt around the tilt axis towards the tail of the fuselage.

[0033] Optionally, in a second embodiment of this application, using Figure 3 and Figure 4 For example, each wing 12 of the free-wing aircraft 10 has at least one set of flight propulsion devices 13 evenly installed on its leading edge. In this case, each set of flight propulsion devices 13 can use at least one periodic variable-pitch rotor to provide vector thrust. When a single set of flight propulsion devices 13 includes multiple variable-pitch rotors, these multiple periodic variable-pitch rotors can be equipped with the same connecting bracket and fixedly installed on the corresponding wing 12 through the connecting bracket.

[0034] When using a periodic variable-pitch rotor... Figure 3 or Figure 4 When the flight propulsion device 13 is in use, the wing tilting moment that drives the corresponding wing 12 to tilt around the tilting axis to the tail or nose of the fuselage can be generated by adjusting the magnitude and / or the difference in the vector thrust difference between the upper and lower wing surfaces of a single wing 12. Specifically, for a single wing 12, the blade pitch of each of the flight propulsion devices 13 on the upper and lower wing surfaces can be periodically adjusted. This increases the angle of attack of the rotor blades of a single flight propulsion device 13 when it rotates to the upper wing surface and decreases the angle of attack of the corresponding rotor blades when it rotates to the lower wing surface, so that the resulting wing tilting moment can be used to drive the corresponding wing 12 to tilt around the tilt axis towards the nose of the fuselage. Alternatively, the angle of attack of the rotor blades of a single flight propulsion device 13 when it rotates to the upper wing surface can be decreased and the angle of attack of the corresponding rotor blades when it rotates to the lower wing surface can be increased, so that the resulting wing tilting moment can be used to drive the corresponding wing 12 to tilt around the tilt axis towards the tail of the fuselage.

[0035] Optionally, in the third implementation of the embodiments of the present application, each wing 12 of the free-wing aircraft 10 is uniformly installed with at least one set of flight propulsion devices 13 at the leading edge of the wing 12, and each wing 12 is installed with aileron structure 17 at the trailing edge of the wing 12, wherein each set of flight propulsion devices 13 provides vector thrust by fixed-pitch rotor, and the aileron structure 17 on the same wing 12 can cooperate with the corresponding flight propulsion devices 13 to adjust the difference in the magnitude and / or direction of the vector thrust on the upper and lower surfaces of the wing 12 during the flight mode switching process, so as to generate a wing tilting moment for driving the corresponding wing 12 to tilt around the tilting axis towards the tail or the head of the fuselage. For example, Figure 6 when the flight propulsion devices 13 installed on the two wings 12 maintain the same vector thrust, the aileron structure 17 can be controlled to deflect towards the lower surface of the corresponding wing 12 (at this time, the airflow on the upper and lower surfaces of the single wing 12 is as shown by the blue thin solid arrows in Figure 6 ), so that the corresponding generated wing tilting moment (as shown by the red wide solid arrows in Figure 6 ) is used to drive the corresponding wing 12 to tilt around the tilting axis towards the head of the fuselage, or the aileron structure 17 is controlled to deflect towards the upper surface of the corresponding wing 12, so that the corresponding generated wing tilting moment is used to drive the corresponding wing 12 to tilt around the tilting axis towards the tail of the fuselage.

[0036] In the embodiments of the present application, the main control unit is electrically connected with the first inertia measurement unit and all flight propulsion devices 13, respectively, for real-time regulation of the vector thrust generated by each set of flight propulsion devices 13 (including thrust magnitude regulation and / or thrust direction regulation) according to the wing tilting attitude information detected by the first inertia measurement unit during the flight mode switching process (including the switching process from the aircraft ultra-short take-off / vertical take-off mode to the aircraft hovering mode, the switching process from the aircraft hovering mode to the aircraft level flight mode, the switching process from the aircraft level flight mode to the aircraft hovering mode, the switching process from the aircraft hovering mode to the aircraft ultra-short take-off / vertical landing mode, etc.), so that the wing 12 tilts around the tilting axis (i.e., the wing tilting function is realized by the vector thrust generated by the multiple sets of flight propulsion devices 13) to regulate the wing tilting angle to the desired wing tilting state corresponding to the desired flight mode (i.e., the aforementioned aircraft level flight mode, the aircraft vertical take-off / landing mode, the aircraft ultra-short take-off / landing mode, or the aircraft hovering mode, etc.).

[0037] Wherein, for the free-wing aircraft 10, the desired wing angle of attack requirements of the aircraft vertical take-off and landing mode (including the aircraft vertical take-off mode and the aircraft vertical landing mode), the aircraft super-short take-off mode and the aircraft hovering mode are that the device axes of each group of flight propulsion devices 13 are as close to vertical as possible (i.e. parallel to the vertical line of the center of gravity of the fuselage), and the thrust direction points upward; the desired wing angle of attack requirement of the aircraft super-short take-off mode is that the device axes of each group of flight propulsion devices 13 are as close to a certain angle (for example, 30°) as possible, and the thrust direction points obliquely upward; the desired wing angle of attack requirement of the aircraft level flight mode is that the device axes of each group of flight propulsion devices 13 are as close to the heading direction as possible, and the thrust direction points to the heading direction.

[0038] In the embodiment, the main control unit can be deployed on the fuselage 11; when the free-wing aircraft 10 adopts the aileron structure 17 to realize the wing surface vector thrust differentiation effect in the flight mode switching process (i.e. the free-wing aircraft 10 adopts the third embodiment described above for structural deployment), the main control unit can control the deflection angle and / or deflection direction of the aileron structure 17 in the flight mode switching process, so that the aileron structure 17 assists the multiple groups of flight propulsion devices 13 to adjust the tilting moment acting on the wing (including moment size adjustment and / or moment direction adjustment), to drive the wing 12 to tilt around the tilting axis, at this time the main control unit is also electrically connected with the aileron structure 17.

[0039] In addition, it can be understood that the first inertial measurement unit can also detect the wing roll attitude information (including the actual wing roll angle and the actual wing roll angular velocity) and / or the wing yaw attitude information (including the actual wing yaw angle and the actual wing yaw angular velocity) at the same time as detecting the wing tilting attitude information, so that the main control unit further considers the wing roll attitude information and / or the wing yaw attitude information when regulating the vector thrust of each group of flight propulsion devices 13, so that the finally adjusted vector thrust can effectively consider the attitude stability effect of the wing 12 in the roll direction and / or the yaw direction while realizing the wing tilting function.

[0040] In the embodiment of the present application, the fuselage active stabilizing device 14 is fixedly installed on the fuselage 11 (wherein the installation position of the fuselage active stabilizing device 14 on the fuselage 11 can be at the head position or at the tail position of the fuselage), wherein the fuselage active stabilizing device 14 is configured to provide a vector thrust, and based on the provided vector thrust, actively generate a fuselage pitch moment acting on the fuselage 11 in the fuselage pitch plane, so as to suppress the fuselage pitch motion caused by the pendulum effect through the fuselage pitch moment, and ensure that the corresponding fuselage 11 can be stably maintained at a desired pitch angle state in a desired flight mode (i.e. the pitch angle that the fuselage 11 needs to continuously maintain in the aforementioned desired flight mode).

[0041] In the embodiment of the present application, the fuselage active stabilizing device 14 can be implemented by at least one variable-pitch rotor, wherein the blade rotation plane of each variable-pitch rotor is parallel to the axis direction of the tilt shaft. The fuselage active stabilizing device 14 can also be implemented by a plurality of fixed-pitch rotors, wherein the blade rotation plane of each fixed-pitch rotor is parallel to the axis direction of the tilt shaft, and wherein the vector thrust direction of a first part of the fixed-pitch rotors can be opposite to the vector thrust direction of a second part of the fixed-pitch rotors, and the first part of the fixed-pitch rotors and the second part of the fixed-pitch rotors are switched to operate.

[0042] In the embodiment of the present application, the second inertial measurement unit is installed on the fuselage 11 and configured to detect the fuselage pitch attitude information (including the actual fuselage pitch angle and the actual fuselage pitch angular velocity). The main control unit is electrically connected with the second inertial measurement unit and the fuselage active stabilizing device 14, and configured to regulate and control the vector thrust generated by the fuselage active stabilizing device 14 according to the fuselage pitch attitude information detected by the second inertial measurement unit, so as to suppress the fuselage pitch motion caused by the pendulum effect.

[0043] In addition, it can be understood that, while detecting the fuselage pitch attitude information, the second inertial measurement unit can also synchronously detect the fuselage roll attitude information (including the actual fuselage roll angle and the actual fuselage roll angular velocity) and / or the fuselage yaw attitude information (including the actual fuselage yaw angle and the actual fuselage yaw angular velocity), so that the main control unit further considers the fuselage roll attitude information and / or the fuselage yaw attitude information when regulating and controlling the vector thrust of the fuselage active stabilizing device 14, so that the finally adjusted vector thrust can effectively take into account the attitude stabilization effect of the fuselage 11 in the roll direction and / or the yaw direction while realizing the fuselage pitch motion modulation function.

[0044] Therefore, the free-wing aircraft 10 provided by the application can realize independent control effects of a body attitude stability control link (i.e., a closed loop of body attitude control through the vector thrust generated by the body active stabilizing device 14 based on the body pitch attitude information detected by the second inertial measurement unit) and a wing tilt angle control link (i.e., a closed loop of wing tilt control through the vector thrust generated by the flight propulsion device 13 based on the wing tilt attitude information detected by the first inertial measurement unit), ensures kinematic control decoupling (i.e., realizes double-loop decoupling control effects of the body and the wing) and physical structure decoupling effects between the body attitude control and the wing tilt control, and can ensure the stability of the overall attitude of the aircraft through the vector thrust (including but not limited to the vector thrust component in the wing tilt plane or the body pitch plane perpendicular to the axis direction of the tilt axis) of the flight propulsion device 13 and the body active stabilizing device 14.

[0045] That is, the free-wing aircraft 10 substantially belongs to a new aircraft configuration combining a suspension structure and a body-wing attitude control decoupling mechanism, which is different from a conventional tilt-wing aircraft configuration, can fully exert the advantages of the suspension structure and the body-wing double-loop control decoupling mechanism, can completely free the aircraft from the constraints of the pendulum effect of the conventional configuration and the requirement of a large-power tilt driving device, can effectively reduce the overall weight and size of the aircraft, and can avoid problems such as phase lag, self-excited oscillation and overshoot caused by the kinematic control coupling of the body and the wing from a physical root, so as to significantly improve the wing tilt control freedom and the body swing suppression flexibility, and facilitate the realization of the free flight effect of the aircraft.

[0046] Optionally, in the embodiment of the application, the free-wing aircraft 10 can further include a wing tilt locking device (not shown) electrically connected to the main control unit and used for controlling the wing tilt locking device to lock or release the wing tilt angle. Figures 1-6 When it is necessary to stably maintain the free-wing aircraft 10 in a specific flight mode (for example, an aircraft level flight mode, an aircraft hovering mode, etc.), the wing tilt angle is directly locked in the expected tilt angle state corresponding to the specific flight mode by the wing tilt locking device (i.e., a wing tilt angle locking operation), without the need to continuously maintain the expected tilt angle state by driving the flight propulsion device 13, so that the energy consumption of the aircraft required for maintaining the specific flight mode is significantly reduced.

[0047] Optionally, in the embodiment of the application, the free-wing aircraft 10 can further include a tail wing 15.

[0048] In the embodiment, the tail fin 15 is installed at the tail position of the fuselage 11, and a pneumatic control rudder 16 (including a pitch rudder and a yaw rudder) is installed on the tail fin 15, and aileron structures 17 are installed at the trailing edge positions of each wing 12. The main control unit is electrically connected to the pitch rudder, the yaw rudder and the aileron structures 17 respectively, for controlling the movement state of each of the pitch rudder, the yaw rudder and the aileron structures 17, to adjust the pitch attitude of the aircraft, the yaw attitude of the aircraft and / or the roll attitude of the aircraft. Among them, the pitch rudder can realize the pitch attitude adjustment function of the aircraft in the aircraft level flight mode; the yaw rudder can realize the yaw attitude adjustment function of the aircraft in the aircraft level flight mode; the aileron structures 17 can realize the roll attitude adjustment function of the aircraft in the aircraft level flight mode, and the aileron structures 17 can also realize the yaw attitude adjustment function of the aircraft in the aircraft hovering mode. Optionally, in an embodiment of the embodiment, airflow control devices (such as micro vortex generators) are installed at the leading edge positions of each wing 12, for delaying airflow separation when the wing 12 is in a large inclination state.

[0049] In the present application, in order to ensure that the above-mentioned free wing aircraft 10 can realize kinematic control decoupling between the fuselage attitude stabilization control link and the wing inclination angle control link during the mode switching / maintenance process of the flight mode specified by the user (i.e. the target flight mode), and ensure the overall stability of the aircraft attitude, the present embodiment realizes the foregoing purpose by providing an aircraft control method applied to the above-mentioned free wing aircraft 10. The aircraft control method provided by the present application will be described in detail below.

[0050] Please refer to Figure 7 , Figure 7 is a flowchart of the aircraft control method provided by the present embodiment. In the present embodiment, the aircraft control method is applied to the main control unit included in the above-mentioned free wing aircraft 10, and the aircraft control method can include steps S210-S230, wherein step S220 and step S230 are executed in parallel.

[0051] Step S210, obtain a flight mode switching instruction, and determine a desired wing inclination angle corresponding to the target flight mode pointed by the flight mode switching instruction.

[0052] In the embodiment, the flight mode switching instruction is used to instruct the free wing aircraft 10 to switch the current aircraft flight mode to the target flight mode for flight operation. The main control unit can be pre-configured with the working wing inclination angle corresponding to each of all known flight modes, so as to find out the target working wing inclination angle matched with the target flight mode through data query, and then take the found target working wing inclination angle as the desired wing inclination angle.

[0053] Step S220, according to the desired wing angle of attack and the first inertial measurement unit real-time detection of the wing tilt attitude information, the vector thrust generated by each of the plurality of flight propulsion device is regulated, so that the wing around the tilt axis to the desired wing angle of attack.

[0054] In this embodiment, when the master control unit receives the flight mode switching instruction, the wing tilt locking device will be unlocked to control the wing tilt angle, so that the wing 12 can be tilted around the tilt axis under the coordination of the plurality of flight propulsion devices 13 to the desired wing angle of attack.

[0055] Wherein, the wing tilt attitude information can include the actual wing angle of attack and the actual wing tilt angle velocity, then the above step S220 can include: Calculate the actual angle of attack difference between the desired wing angle of attack and the actual wing angle of attack, and based on the actual wing tilt angle velocity, calculate the expected tilt torque acting on the wing 12, so that the actual angle of attack difference converges to zero. According to the expected tilt torque, the vector thrust generated by each of the plurality of flight propulsion devices 13 is adjusted.

[0056] Wherein, the adjustment operation of the vector thrust acting on the upper and lower surfaces of the wing 12 by each of the plurality of flight propulsion devices 13 can refer to the related description of the above Figures 1-6 , here will not be described one by one.

[0057] In this process, it can be understood that the master control unit can first detect whether the desired wing angle of attack and the actual wing angle of attack detected by the first inertial measurement unit are consistent before executing the step S220, and when the desired wing angle of attack and the actual wing angle of attack are consistent, it is determined that the current aircraft flight mode of the free wing aircraft 10 belongs to the target flight mode, at this time the master control unit will stop executing the above step S220 but continue to execute step S230, or when the desired wing angle of attack and the actual wing angle of attack are not consistent, continue to execute step S220 and step S230 in parallel, so as to effectively reduce the calculation resource occupation.

[0058] Step S230, according to the second inertial measurement unit real-time detection of the body pitch attitude information, control the body active stabilizing device to provide vector thrust to the body to suppress the pendulum effect caused by the body pitch motion.

[0059] In the embodiment, the main control unit can drive the fuselage active stabilizing device 14 to actively suppress the fuselage pitching motion caused by the pendulum effect in the mode switching / maintaining process for the target flight mode, so as to ensure the stability of the overall attitude of the aircraft through the vector thrust of the flight propulsion device 13 and the fuselage active stabilizing device 14.

[0060] Optionally, please refer to Figure 8 , Figure 8 is Figure 7 The flowchart of the sub-steps included in step S230 is shown in FIG. 8. In the embodiment, the fuselage pitching attitude information can include the actual fuselage pitching angle and the actual fuselage pitching angular velocity of the fuselage 11, and step S230 can include sub-step S231 to sub-step S233 to ensure that the fuselage active stabilizing device 14 can actively suppress the fuselage pitching motion caused by the pendulum effect in the mode switching / maintaining process for the target flight mode, while maintaining the fuselage 11 in the expected pitching angle state corresponding to the target flight mode.

[0061] In sub-step S231, the expected fuselage pitching angle and the expected fuselage pitching angular velocity corresponding to the target flight mode are determined.

[0062] In the embodiment, the main control unit can be pre-configured with the working fuselage pitching angle and the working fuselage pitching angular velocity corresponding to each of all known flight modes, so as to find out the target working fuselage pitching angle and the target working fuselage pitching angular velocity matching the target flight mode through data query, and then take the target working fuselage pitching angle and the target working fuselage pitching angular velocity as the expected fuselage pitching angle and the expected fuselage pitching angular velocity respectively.

[0063] In sub-step S232, the expected pitching moment required for the fuselage to achieve the expected fuselage pitching angle and the expected fuselage pitching angular velocity is calculated according to the fuselage pitching attitude information, the expected fuselage pitching angle and the expected fuselage pitching angular velocity.

[0064] In the embodiment, the main control unit can use a PID (Proportion-Integral-Differential) control algorithm or a deep learning method (such as a DQN (Deep Q Network) learning method) to predict the fuselage pitching moment for the fuselage 11 in order to keep the fuselage 11 at the expected fuselage pitching angular velocity when achieving the expected fuselage pitching angle, and obtain the expected pitching moment.

[0065] Optionally, in an implementation form of the embodiment, the main control unit executes the above-mentioned sub-step S232 by using a PID algorithm, and the sub-step S232 can include: calculating an actual angle difference between the desired fuselage pitch angle and the actual fuselage pitch angle, and calculating an actual angular velocity difference between the desired fuselage pitch angular velocity and the actual fuselage pitch angular velocity; converging the actual angle difference to zero and the actual angular velocity difference to zero, and using a PID algorithm to solve the desired pitch moment, so as to obtain the desired pitch moment.

[0066] Sub-step S233: regulating the vector thrust provided by the fuselage active stabilization device according to the desired pitch moment.

[0067] Therefore, by executing the above-mentioned sub-step S231 to sub-step S233, the fuselage active stabilization device 14 can actively suppress the fuselage pitch motion caused by the pendulum effect while maintaining the fuselage 11 in the desired pitch angle state corresponding to the target flight mode during the mode switching / maintenance process for the target flight mode.

[0068] By executing the above-mentioned steps S210 to S230, the free wing aircraft 10 can realize kinematic control decoupling between the fuselage attitude stabilization control link and the wing tilt angle control link during the mode switching / maintenance process for the target flight mode, and ensure the overall stability of the aircraft attitude.

[0069] The above-mentioned is only various implementation forms of the present application, but the protection scope of the present application is not limited thereto, any person skilled in the art can easily think of changes or replacements within the technical range disclosed in the present application, which should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A free wing aircraft, characterized in that, The aircraft comprises a fuselage, a first inertial measurement unit, a second inertial measurement unit, a main control unit, two wings, a fuselage active stabilizing device and multiple sets of flight propulsion devices; The two wings are symmetrically arranged on both sides of the fuselage and are rotationally connected to the fuselage through a tilting shaft, wherein at least one set of flight propulsion devices is separately installed on each wing, the flight propulsion devices installed on the two wings are symmetric about the body axis of the fuselage, the first inertial measurement unit is used to detect wing tilting attitude information, and the second inertial measurement unit is used to detect fuselage pitching attitude information; The main control unit is electrically connected with the first inertial measurement unit and the multiple sets of flight propulsion devices, and is used to control the vector thrust generated by the multiple sets of flight propulsion devices in real time according to the wing tilting attitude information detected by the first inertial measurement unit during flight mode switching, so that the wings are tilted around the tilting shaft; The main control unit is electrically connected with the second inertial measurement unit and the fuselage active stabilizing device, and is used to control the vector thrust generated by the fuselage active stabilizing device according to the fuselage pitching attitude information detected by the second inertial measurement unit, so as to suppress the fuselage pitching motion caused by the pendulum effect, wherein the fuselage active stabilizing device is installed on the fuselage.

2. The aircraft of claim 1, wherein, The upper and lower surfaces of each wing are uniformly provided with at least one set of flight propulsion devices, wherein each set of flight propulsion devices provides vector thrust by using at least one fixed-pitch rotor or variable-pitch rotor.

3. The aircraft of claim 1, wherein, The leading edge of each wing is uniformly provided with at least one set of flight propulsion devices, wherein each set of flight propulsion devices provides vector thrust by using a cyclical variable-pitch rotor.

4. The aircraft of claim 1, wherein, The leading edge of each wing is uniformly provided with at least one set of flight propulsion devices, and the trailing edge of each wing is provided with aileron structure, wherein each set of flight propulsion devices provides vector thrust by using a fixed-pitch rotor; The main control unit is also electrically connected with the aileron structure, and is used to control the deflection angle and / or deflection direction of the aileron structure during flight mode switching, so that the aileron structure assists the multiple sets of flight propulsion devices in adjusting the tilting moment acting on the wings, so as to drive the wings to tilt around the tilting shaft.

5. The aircraft of any one of claims 1-4, wherein, The aircraft further comprises a wing tilting locking device; The main control unit is electrically connected with the wing tilting locking device, and is used to control the wing tilting locking device to lock and maintain or release the wing tilting angle.

6. The aircraft of claim 5, wherein, The aircraft further comprises a tail wing; The tail wing is installed at the tail position of the fuselage, the tail wing is provided with a pneumatic control surface, and the trailing edge of each wing is provided with aileron structure, wherein the pneumatic control surface comprises a pitch rudder and a yaw rudder; The main control unit is electrically connected with the pitch rudder, the yaw rudder and the aileron structure respectively, and is used to control the motion state of each of the pitch rudder, the yaw rudder and the aileron structure, so as to adjust the aircraft pitching attitude, the aircraft yawing attitude and / or the aircraft rolling attitude.

7. An aircraft control method, characterized in that, The control method is applied to the main control unit of the free-wing aircraft in any one of claims 1-6, and the control method comprises: acquiring a flight mode switching instruction, and determining a desired wing angle of attack corresponding to a target flight mode pointed by the flight mode switching instruction; controlling vector thrusts generated by a plurality of flight propulsion devices according to the desired wing angle of attack and wing tilt attitude information detected by a first inertial measurement unit in real time, so as to tilt the wing around a tilt axis to the desired wing angle of attack; controlling a vector thrust provided by a fuselage active stabilizing device to the fuselage according to body pitch attitude information detected by a second inertial measurement unit in real time, so as to suppress body pitch motion caused by a pendulum effect.

8. The control method according to claim 7, characterized by, The wing tilt attitude information includes an actual wing angle of attack and an actual wing tilt angular velocity, and the step of controlling the vector thrusts generated by the plurality of flight propulsion devices according to the desired wing angle of attack and the wing tilt attitude information detected by the first inertial measurement unit in real time, so as to tilt the wing around the tilt axis to the desired wing angle of attack, includes: calculating an actual angle difference between the desired wing angle of attack and the actual wing angle of attack, and calculating a desired tilt torque acting on the wing based on the actual wing tilt angular velocity, so as to converge the actual angle difference to zero; adjusting the vector thrusts generated by the plurality of flight propulsion devices according to the desired tilt torque.

9. The control method according to claim 7, characterized by, The step of controlling the vector thrust provided by the fuselage active stabilizing device to the fuselage according to the body pitch attitude information detected by the second inertial measurement unit in real time, so as to suppress the body pitch motion caused by the pendulum effect, includes: determining a desired body pitch angle and a desired body pitch angular velocity corresponding to the target flight mode; calculating a desired pitch torque required for the fuselage to achieve the desired body pitch angle and the desired body pitch angular velocity according to the body pitch attitude information, the desired body pitch angle and the desired body pitch angular velocity; controlling the vector thrust provided by the fuselage active stabilizing device according to the desired pitch torque.

10. The control method according to claim 9, characterized by, The body pitch attitude information includes an actual body pitch angle and an actual body pitch angular velocity, and the step of calculating the desired pitch torque required for the fuselage to achieve the desired body pitch angle and the desired body pitch angular velocity according to the body pitch attitude information, the desired body pitch angle and the desired body pitch angular velocity, includes: calculating an actual angle difference between the desired body pitch angle and the actual body pitch angle, and calculating an actual angular velocity difference between the desired body pitch angular velocity and the actual body pitch angular velocity; adopting a PID algorithm to solve the pitch torque so as to converge the actual angle difference to zero and the actual angular velocity difference to zero, and obtaining the desired pitch torque.

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