Vertical takeoff and landing aircraft and its control method

Abstract

This application provides a vertical take-off and landing aircraft and a control method thereof. [Solution] The vertical take-off and landing aircraft comprises an airframe and 2N tiltrotors. Main wings are provided on both sides of the airframe, a tail fin is provided at the rear of the airframe, and a rudder is provided on the tail fin. The 2N tiltrotors are mounted symmetrically on both sides of the airframe, and some of the 2N tiltrotors are located at the tail fin. Here, N is a natural number greater than or equal to 2. In the vertical take-off and landing state, the projection of the propellers of the 2N tiltrotors in the horizontal plane is center-symmetric with respect to point B, and both point B and the center of gravity G of the vertical take-off and landing aircraft are located within the plane of symmetry of the airframe, with point B located on the side of point G closer to the tail fin. In the process of the vertical take-off and landing aircraft transitioning from the vertical take-off and landing state to the cruising state, both point G and point B move toward the nose side along the plane of symmetry, and point B is always located on the side of point G closer to the tail fin. This layout improves the trim-pitch capability of the vertical take-off and landing aircraft.

Description

Technical Field

[0001] This application claims the priority of a Chinese patent application filed on December 7, 2023, with application number 2023116776187 and title "Vertical Takeoff and Landing Aircraft and Its Control Method", and a Chinese patent application filed on December 7, 2023, with application number 2023233405592 and title "Vertical Takeoff and Landing Aircraft", and the entire contents of these are incorporated herein by reference.

[0002] This application relates to the technical field of aircraft, specifically to vertical takeoff and landing aircraft and their control methods.

Background Art

[0003] The vertical takeoff and landing fixed-wing aircraft (distributed propulsion type) combines the vertical takeoff and landing ability of a helicopter and the efficiency and high speed in horizontal flight of a fixed wing. It is quieter, more comfortable, and more economical than a helicopter, and is more efficient and has a longer endurance than a multi-rotor. Also, compared to a fixed-wing aircraft, it can vertically take off and land on an urban takeoff and landing platform, making it an optimal choice for air travel in the city. However, in conventional vertical takeoff and landing aircraft, when a rotor is installed on the tail, the airflow generated by the rotor and the airflow generated by the tail interfere with each other, making pitch control difficult and bringing unexpected technical problems to the flight control of the vertical takeoff and landing aircraft. Summary of the Application

[0004] In view of the above drawbacks of the prior art, this application provides a vertical takeoff and landing aircraft and its control method that improves the problem that the airflow interference between the tilt rotor and the tail in the tail of a conventional vertical takeoff and landing aircraft is large and pitch control is difficult.

[0005] To achieve the above-mentioned objectives and other related objectives, the present invention provides a vertical take-off and landing (VTOL) aircraft comprising an airframe and 2N tiltrotors. The airframe is provided with wings on both sides, a tail fin on the tail end of the airframe, and a rudder fin on the tail fin. The 2N tiltrotors are mounted symmetrically on both sides of the airframe, and some of the 2N tiltrotors are located on the tail fin. Here, N is a natural number greater than or equal to 2, and in the VTOL configuration, the horizontal projections of the propellers of the 2N tiltrotors are center-symmetric with respect to point B, and both point B and the center of gravity G of the VTOL aircraft are located within the plane of symmetry of the airframe, with point B located on the side of point G closer to the tail fin, and during a mode change of the VTOL aircraft, both point G and point B move along the plane of symmetry, with point B always located on the side of point G closer to the tail fin.

[0006] In one embodiment of the vertical take-off and landing aircraft of the present invention, during flight, the projections of the rotation axes of any of the tiltrotors at the tail and any of the tiltrotors at other positions onto the plane of symmetry of the aircraft are not parallel.

[0007] In one embodiment of the vertical takeoff and landing aircraft of the present invention, in the cruising state and / or the vertical takeoff state and / or the mode change state, there is a first difference between the tilt velocity of any of the tiltrotors at the tail fin and the tilt velocity of any of the tiltrotors at other positions, and the first difference is not zero.

[0008] In one embodiment of the vertical takeoff and landing aircraft of the present invention, in the cruising state and / or the vertical takeoff state and / or the mode change state, there is a second difference between the rotational speed of one of the tiltrotors on the tail and the rotational speed of any of the tiltrotors at any other position, and the second difference is not zero.

[0009] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft is, During flight, the pitch control ratio between the rudder rotor and the 2N tilt rotors is distributed according to the current airspeed or dynamic pressure. Pitch control is performed to achieve pitch trim and maneuvering by individually controlling the rudder rotor and the 2N tilt rotors according to the pitch control ratio.

[0010] In one embodiment of the vertical take-off and landing aircraft of the present invention, controlling 2N tiltrotors according to the pitch control ratio is: The pitch moment is differentially adjusted by the difference in tilt angle between the tilt rotor on the tail fin and any other tilt rotor, thereby performing pitch trim and control. and / or, differential adjustment of the pitching moment by the difference in rotational speed between the tiltrotor and any other tiltrotor in the tail fin, thereby performing pitch trim and control. and / or, by differential adjustment of the pitching moment, pitch trim, and control based on the difference in tilting speed between the tilt rotor and any other tilt rotor at the tail fin.

[0011] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft comprises four tiltrotors symmetrically mounted on both sides of the aircraft body, two of which are symmetrically mounted to the main wings with respect to the aircraft body, and the remaining two tiltrotors are symmetrically mounted to the tail fin.

[0012] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft comprises six tiltrotors symmetrically mounted on both sides of the aircraft, of which four tiltrotors are symmetrically mounted to the main wings with respect to the aircraft, and the remaining two tiltrotors are symmetrically mounted to the tail fin.

[0013] In one embodiment of the vertical takeoff and landing aircraft of the present invention, the tail fin is a V-tail fin, the two tiltrotors on the tail fin are fully tiltable tiltrotors, and the two fully tiltable tiltrotors are each attached to the wingtips on both sides of the upper part of the V-tail fin.

[0014] In one embodiment of the vertical take-off and landing aircraft of the present invention, two of the tiltrotors are provided at the wingtips of the main wing, and the tiltrotors located at the wingtips of the main wing are fully tiltable tiltrotors.

[0015] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft further comprises 2M fixed rotors, where M is a natural number of 2 or more, the 2M fixed rotors are symmetrically mounted on the main wings on both sides of the aircraft and located outside the tilt rotor, and in the vertical take-off and landing state, the horizontal projections of all the fixed rotors are all center-symmetric with respect to point A, point A is located within the plane of symmetry of the aircraft, and during mode changes of the vertical take-off and landing aircraft, point G is located on the side of point A closer to the nose, or overlaps with point A, and point B is always located on the side of point A closer to the tail fin.

[0016] In one embodiment of the vertical take-off and landing aircraft of the present invention, the angle between the rotation axes of the 2M fixed rotors and the plane of symmetry of the aircraft is -15° to +15°, and / or, in the process of the 2N tilt rotors being tilted, the angle between the plane formed by the rotation axes and the plane of symmetry of the aircraft is -15° to +15°.

[0017] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft comprises four tiltrotors and four fixed rotors. If the distance from the center of gravity G to point A is L1 (L1≧0), the distance from point A to point B is L2 (L2>0), the spacing between the four fixed rotors in the extending direction of the aircraft body is L3, and the spacing between the four tiltrotors in the extending direction of the aircraft body is L4, then 0.1(L3+L4)≧4L1+2L2≧0.01(L3+L4).

[0018] In one embodiment of the vertical takeoff and landing aircraft of the present invention, the tail is a V-tail, the two tiltrotors on the tail are fully tiltable tiltrotors, and the two fully tiltable tiltrotors are each attached to the wingtips on both sides of the upper part of the V-tail.

[0019] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft is, During flight, the pitch control ratios of the rudder rotor, 2N tilt rotors, and 2M fixed rotors are distributed according to the current airspeed or dynamic pressure. In accordance with the aforementioned pitch control ratio, the rudder rotor, 2N tilt rotors, and 2M fixed rotors are individually controlled to perform pitch control in order to achieve pitch trim and steering.

[0020] In one embodiment of the vertical take-off and landing aircraft of the present invention, the four fixed rotors are mounted symmetrically on both sides of the aircraft body, the four tilt rotors are located inside the four fixed rotors, two of which are mounted symmetrically to the tail fins, and the two tilt rotors located to the tail fins are fully tiltable tilt rotors.

[0021] In one embodiment of the vertical take-off and landing aircraft of the present invention, the tail fin is a V-tail fin, two tiltrotors are attached to the V-tail fin, and the two tiltrotors are each attached to the wingtips of the V-tail fin. In the vertical take-off and landing state, the distance between the rotation center of the tiltrotors on the tail fin and the leading edge of the wingtip of the V-tail fin is t1, along a direction parallel to the roll axis of the vertical take-off and landing aircraft, the chord length of the wingtip of the V-tail fin is t2, and the ratio of t1 to t2 is 15% to 40%.

[0022] In one embodiment of the vertical take-off and landing aircraft of the present invention, the tiltrotor located at the tail fin is a fully tilted tiltrotor, and the tiltrotor located at a position other than the tail fin is a partially tilted tiltrotor.

[0023] In one embodiment of the vertical take-off and landing aircraft of the present invention, tiltrotors are provided at the wingtips of the main wings, and both the tiltrotor located on the tail and the tiltrotor located at the wingtips of the main wings are fully tiltable tiltrotors.

[0024] In one embodiment of the vertical takeoff and landing aircraft of the present application, the fully tilted tilt rotor includes a first rotor and a nacelle, the first rotor is connected to the nacelle, the nacelle is rotatably connected to the tail wing or the main wing, and the nacelle tilts synchronously with the first rotor during the tilting process of the first rotor.

[0025] In one embodiment of the vertical takeoff and landing aircraft of the present application, the first rotor includes a propeller and a rotary drive device, the propeller is attached to the output shaft of the rotary drive device, the nacelle includes a nacelle housing and a tilt mechanism located inside the nacelle housing, and the tilt mechanism tilts the rotary drive device to tilt the propeller.

[0026] In one embodiment of the vertical takeoff and landing aircraft of the present application, the tilt mechanism includes a rocker arm, a drive arm, a tilt drive device, and a connecting rod. Both the rocker arm and the drive arm are rotatably attached to the tail wing. The base of the tilt drive device is attached to the tail wing. The drive end of the tilt drive device rotationally drives the drive arm. The connecting rod is rotatably connected to the drive arm and the rocker arm respectively. The rocker arm is fixedly connected to the rotary drive device of the first rotor.

[0027] In one embodiment of the vertical takeoff and landing aircraft of the present application, a first shaft body and a second shaft body parallel to each other are fixedly installed on the tail wing. The rocker arm is rotatably attached to the first shaft body. The drive arm is rotatably attached to the second shaft body. The base of the tilt drive device is attached to the second shaft body. The drive part of the tilt drive device is fixed to the drive arm.

[0028] In one embodiment of the vertical takeoff and landing aircraft of the present application, a holding structure is provided between the first shaft body and the second shaft body. One end of the holding structure is fixedly connected to the base, and the other end of the holding structure surrounds and holds the first shaft body.

[0029] In one embodiment of the vertical takeoff and landing aircraft of the present invention, the first shaft and / or the second shaft are hollow shafts.

[0030] In one embodiment of the vertical take-off and landing aircraft of the present invention, the rudder beta includes a rudder plate and a rudder body drive device, wherein the rudder plate is rotatably connected to the tail fin or the tail of the aircraft, and the rudder body drive device rotates the rudder plate to adjust the direction of the vertical take-off and landing aircraft.

[0031] In one embodiment of the vertical take-off and landing aircraft of the present invention, the roll axis of the vertical take-off and landing aircraft is set to 0°, the upward tilt of the tiltrotor is set to the positive direction, the downward tilt is set to the negative direction, and the rotation axis of the tiltrotor is tilted within the range of -20° to 110°.

[0032] In one embodiment of the vertical take-off and landing aircraft of the present invention, the tail fin is a tail fin that sweeps forward, and the tiltrotor on the tail fin that sweeps forward is a fully tilted tiltrotor.

[0033] In one embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft further includes a vertical stabilizer provided below the forward-sweeping tail fin and connected to the tail of the aircraft.

[0034] The present invention further provides a method for controlling a vertical take-off and landing (VTOL) aircraft. The VTOL aircraft comprises an airframe and 2N tiltrotors. Main wings are provided on both sides of the airframe, a tail fin is provided at the tail of the airframe, and a rudder is provided on the tail fin. The 2N tiltrotors are mounted symmetrically on both sides of the airframe, and a portion of the 2N tiltrotors are located at the tail fin. Here, N is a natural number greater than or equal to 2, and in the VTOL configuration, the horizontal projection of the propellers of the 2N tiltrotors is center-symmetric with respect to point B, and both point B and the center of gravity G of the VTOL aircraft are located within the plane of symmetry of the airframe, with point B located on the side of point G closer to the tail fin, and during mode changes of the VTOL aircraft, both point G and point B move along the plane of symmetry. Specifically, during the process in which the vertical take-off and landing aircraft transitions from a vertical take-off and landing state to a cruising state, both point G and point B move toward the nose side along the plane of symmetry, and point B is always located toward the tail side of point G. The control method is as follows: The pitch control ratio of the rudder better and the 2N tilt rotors is distributed according to the current airspeed or dynamic pressure. The pitch control process includes individually controlling the rudder rotor and 2N tilt rotors according to the pitch control ratio to achieve pitch trim and steering.

[0035] In one embodiment of the control method of the present invention, before allocating the pitch control ratio of the rudder better and the 2N tilt rotors according to the current airspeed or dynamic pressure, Get the current tilt position of each tilt rotor, If the current tilt position does not coincide with the set cruising position, the current airspeed or dynamic pressure at the corresponding tilt rotor's current tilt position is obtained, and it is determined whether the current airspeed or dynamic pressure is equal to or greater than a preset threshold at the current tilt position. If the current airspeed or dynamic pressure is above a preset threshold at the current tilt position, the tilt rotor is tilted to the next preset position. The process further includes a tiltrotor control process that gradually increases the rotational speed of 2N tiltrotors.

[0036] In one embodiment of the control method of the present invention, the vertical take-off and landing aircraft further comprises 2M fixed rotors, where M is a natural number greater than or equal to 2, and the 2M fixed rotors are distributed around 2N tilt rotors, and the rotor control process further includes gradually increasing the rotational speed of the 2N tilt rotors and gradually decreasing the rotational speed of the 2M fixed rotors to a set rotational speed.

[0037] In one embodiment of the control method of the present invention, before the rotor control, The 2N tilt rotors are tilted until the axis of rotation is vertically upward or diagonally upward. Tilt the rudder vehicle downwards. The takeoff control process further includes activating 2M of the fixed rotors and 2N of the tilt rotors, and issuing a horizontal flight instruction once the vertical takeoff and landing aircraft has reached a set altitude.

[0038] In one embodiment of the control method of the present invention, the vertical take-off and landing aircraft further comprises 2M fixed rotors, the 2M fixed rotors distributed around 2N tilt rotors. Before allocating the pitch control ratio of the rudder better and the 2N tilt rotors according to the current airspeed or dynamic pressure, The 2N tilt rotors are tilted until their axis of rotation is horizontal and forward-facing. Tilt the rudder vehicle downwards. The takeoff control process further includes activating 2M of the fixed rotors and 2N of the tilt rotors, and issuing a horizontal flight instruction once the vertical takeoff and landing aircraft has reached a set altitude.

[0039] In one embodiment of the control method of the present invention, a tiltrotor control process is further included, after the takeoff control process and before pitch control, in which the rotational speed of 2N tiltrotors is gradually increased and a forward flight instruction is issued, and the rotational speed of 2M fixed rotors is gradually decreased to a set rotational speed.

[0040] In one embodiment of the control method of the present invention, the rotor control process further includes controlling the rudder vater such that the rotational speed of 2M fixed rotors is gradually reduced to a set rotational speed, then returned to zero according to the current airspeed or dynamic pressure, and gradually participates in the pitch control process.

[0041] In one embodiment of the control method of the present invention, a ground preparation process is further included before the takeoff control process, the ground preparation process includes starting the vertical takeoff and landing aircraft, performing a power-on self-test, and confirming the full stroke state of the servo system.

[0042] In one embodiment of the control method of the present invention, the rudder rotor and 2N tilt rotors are individually controlled according to the pitch control ratio to achieve pitch trim and steering, which includes differentially adjusting the pitching moment based on the difference in tilt speed, tilt angle, and / or the difference in rotation speed of the tilt rotors, where the centers of gravity are located at different positions, and performing pitch trim and steering.

[0043] In one embodiment of the control method of the present invention, the control method further includes sequentially repeating a rotor control process and a pitch control process until the tiltrotor is tilted to a cruising position, takes off, and flies horizontally.

[0044] The vertical take-off and landing (VTOL) aircraft of this invention is equipped with elevators and 2N tiltrotors. In the VTOL take-off and landing state, the horizontal projection of the propellers of the 2N tiltrotors is center-symmetric with respect to point B. Both point B and the center of gravity G of the VTOL aircraft are located within the plane of symmetry of the aircraft, with point B located on the side of point G closer to the tail. During mode changes of the VTOL aircraft, point B is always located on the side of point G closer to the tail. In this layout, the center of gravity G of the VTOL aircraft and point B, the center of symmetry of the tiltrotors, do not coincide. In particular, during the transition of the VTOL aircraft from the VTOL take-off and landing state to the cruising state, both point G and point B move toward the nose along the plane of symmetry. Therefore, the moment of the traction force generated by the tiltrotors in front of the center of gravity with respect to the center of gravity G becomes smaller, and the moment of the traction force generated by the tiltrotors behind the center of gravity becomes larger. The difference in moment between the forward and rear tiltrotors allows the tiltrotor's downwash region at the tail to resist a portion of the nose-up moment generated by the tail's action, thus mitigating the difficulty of pitch control. Therefore, when the rotational speed and throttle state of the tiltrotors before and after the center of gravity G are the same, the difference in the length of the lever arms at the center of gravity G generates a nose-down moment. This nose-down moment cancels out or partially cancels out the nose-up moment generated by the tiltrotor's downwash region at the tail's action. Thus, vertical take-off and landing aircraft can trim the pitching moment when the throttle states of the front and rear rotors are matched.

[0045] The control method of the present invention allocates the pitch control ratio between the rudder rotor and the 2N tilt rotors according to the current airspeed or dynamic pressure, and enables pitch control through the cooperation of the rudder rotor and the 2N tilt rotors. [Brief explanation of the drawing]

[0046] To more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings that may be used in describing the embodiments or the prior art are briefly introduced below. Clearly, the drawings in the following description represent only a few embodiments of the present application, and those skilled in the art can obtain other embodiments based on these drawings without any creative effort.

[0047] [Figure 1] This is an axonometric view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state in one embodiment of the present invention. [Figure 2] This is a side view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state according to one embodiment of the present invention. [Figure 3] This is an axonometric view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state in another embodiment of the present invention. [Figure 4] This is a side view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state, according to another embodiment of the present invention. [Figure 5] This is an axonometric view of a vertical take-off and landing aircraft in cruising condition in yet another embodiment of the present invention. [Figure 6] This is an axonometric view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state in yet another embodiment of the present invention. [Figure 7] This is a top view of a vertical takeoff and landing aircraft in the vertical takeoff and landing state, according to yet another embodiment of the present invention. [Figure 8] This is a side view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state, according to yet another embodiment of the present invention. [Figure 9] This is a rear view of a vertical takeoff and landing aircraft in a vertical takeoff and landing state, according to yet another embodiment of the present invention. [Figure 10] This is a partial view of a fully tilted tiltrotor. [Figure 11] This is a partial view after the nacelle housing has been removed from the fully tilted tiltrotor. [Figure 12] This is a top view after the nacelle housing has been removed from the fully tilted tiltrotor. [Figure 13] Figure 12 is a cross-sectional view of the DD. [Figure 14]This is a view from a different angle after the nacelle housing has been removed from the fully tilted tiltrotor. [Figure 15] This is a three-dimensional view from a different angle after the nacelle housing has been removed from the fully tilted tiltrotor. [Figure 16] Figure 14 is a cross-sectional view of the FF. [Figure 17] Figure 14 is a cross-sectional view of the PP. [Figure 18] This is a partial view of the aircraft when the full tilt rotor on the tail fin is in the vertical takeoff position and the chord length of the rudder plate accounts for 30% of the total length. [Figure 19] This is a partial view of the aircraft when the full tilt rotor on the tail fin is in the vertical takeoff position and the chord length of the rudder plate accounts for 60% of the total length. [Figure 20] This is a partial view of the tail fin when the full tilt rotor is in the vertical takeoff position and the chord length of the rudder plate occupies 100%. [Figure 21] This is a partial view of the tail fin when the full tilt rotor is in cruising position. [Figure 22] This is a partial view of the tail fin when the full tilt rotor is in an upward-sloping position. [Figure 23] This is a diagram showing the rotational trajectory of the rudder beta on the tail fin. [Figure 24] This is a top view of the tail fin of an aircraft in a vertical takeoff and landing position. [Figure 25] These are schematic diagrams showing the chord length ratio of the tail fin's rudder plate and the deflection angle of the rudder plate. [Figure 26] This is a top view of the tail fin of an aircraft in a vertical takeoff and landing position. [Figure 27] This shows the curves (CFD simulation results) of the change in the pitch moment (dimensionless) of the entire aircraft as it changes with wind speed (dimensionless) in control methods that partially tilt the nacelle and control methods that fully tilt the nacelle. [Figure 28] This is a flowchart showing the transition from ground state to cruising state for one embodiment of the vertical take-off and landing aircraft of the present invention. [Figure 29]This is a flowchart showing the transition from takeoff to horizontal flight for one embodiment of the vertical takeoff and landing aircraft of the present invention. [Figure 30] This is a flowchart of the pitch control for one embodiment of the vertical take-off and landing aircraft of the present invention. [Figure 31] This is a flowchart of the rotor control for one embodiment of the vertical take-off and landing aircraft of the present invention. [Figure 32] This is a flowchart of the takeoff control process for one embodiment of the vertical takeoff and landing aircraft of the present invention. [Figure 33] This is a flowchart of the takeoff control process for one embodiment of the vertical takeoff and landing aircraft of the present invention. [Figure 34] This is a top view of a vertical takeoff and landing aircraft in a further embodiment of the present invention. [Modes for carrying out the invention]

[0048] The embodiments of this application will be described below through specific examples, but those skilled in the art will readily understand other advantages and effects of this application from the disclosures herein. This application can also be implemented or applied through other different specific embodiments, and various details herein can be modified or changed in various ways on a variety of viewpoints and uses without departing from the spirit of this application. The following embodiments and features thereof may be combined in any way that does not contradict each other. It should also be understood that the terms used in the embodiments of this application are for the purpose of describing specific means of implementation and are not intended to limit the scope of protection of this application. In the following embodiments, test methods for which specific conditions are not explicitly stated usually follow conventional conditions or conditions recommended by each manufacturer.

[0049] Where numerical ranges are indicated in the examples, it should be understood that any two endpoints and any values ​​between those two endpoints can be selected unless otherwise specified in this application. Unless otherwise specified, all technical and scientific terms used in this application are consistent with the understanding of those skilled in the art relating to the prior art and the description of this application, and this application may be carried out using any similar or equivalent prior art methods, apparatus, and materials as described in the examples of this application.

[0050] Furthermore, terms such as "up," "down," "left," "right," "middle," and "one" used herein are used solely for explanatory purposes and do not limit the scope of the invention. Changes or adjustments to relative relationships are also considered to be within the scope of the invention, as long as the technical content is not substantially altered.

[0051] Referring to Figures 1 to 34, the present invention first provides a vertical take-off and landing (VTOL) aircraft. In this VTOL aircraft, 2N tiltrotors are symmetrically mounted on the outer circumference of the aircraft body 10, and elevators are provided on the tail fin 30. By utilizing the positional relationship between the 2N tiltrotors and the center of gravity, it is possible to improve the control of the pitching moment of the VTOL aircraft in a complex interfering flow field and to improve the control problems caused by airflow interference between the tiltrotors and the tail fin 30 in conventional VTOL aircraft.

[0052] Referring to Figures 1 to 7, the present invention provides a vertical take-off and landing aircraft comprising a fuselage and 2N tiltrotors. The fuselage 10 has a symmetrical structure and has a plane of symmetry 60 extending along the longitudinal direction of the fuselage 10 (i.e., the vertical plane on which the straight line O1-O2 in Figure 3 is located). The other structure and shape of the fuselage 10 are not limited and the structure of the fuselage 10 of a conventional vertical take-off and landing aircraft can be referenced. Conventional aircraft operating systems such as avionics systems, flight control systems, electrical systems, and navigation systems are included in the fuselage 10 and will not be described further here. Wings 20 are provided on both sides of the fuselage 10, and the wings 20 on both sides are symmetrical with respect to the plane of symmetry 60 of the fuselage 10. The structure of the wings 20 can be referenced from the structure of a conventional fixed wing of an aircraft and will not be described further here. A tail fin 30 is provided at the tail end of the fuselage 10. The tail fin 30 is integrally molded with the fuselage 10 or mechanically connected and mounted symmetrically with respect to the plane of symmetry 60 of the fuselage 10. The tail fin 30 is provided with a rudder 31, and the mounting position and structure of the rudder 31 are diverse; for example, it may be provided at any suitable position on the tail fin 30, or it may be any suitable conventional rudder 31 structure.

[0053] Importantly, referring to Figure 7, 2N tiltrotors are mounted on both sides of the aircraft 10, where N is a natural number greater than or equal to 2, the 2N tiltrotors are mounted symmetrically with respect to the plane of symmetry 60 of the aircraft 10, and some of the 2N tiltrotors are mounted on the tail fin 30. In the vertical takeoff and landing (VTOL) configuration, the horizontal projection of the propellers of the 2N tiltrotors is center-symmetric with respect to point B, and both point B and the center of gravity G of the VTOL aircraft are located within the plane of symmetry of the aircraft, with point B located on the side of point G closer to the tail fin. During mode changes of the VTOL aircraft, both point G and point B move along the plane of symmetry, for example, during the transition from VTOL to cruising, both point G and point B move along the plane of symmetry towards the nose, with point B always located on the side of point G closer to the tail fin.

[0054] With the above layout, the center of gravity G of the vertical take-off and landing aircraft and point B, the center of symmetry of the 2N tiltrotors, do not overlap, and during the transition of the vertical take-off and landing aircraft from the vertical take-off and landing state to the cruising state, both point G and point B move toward the nose along the plane of symmetry 60. Therefore, the traction force generated by the tiltrotors in front of the center of gravity has a smaller moment relative to the center of gravity G, while the traction force generated by the tiltrotors behind the center of gravity has a larger moment relative to the center of gravity G. The moment difference between the forward and rear tiltrotors allows the unwash region of the tiltrotors on the tail fin 30 to resist a portion of the nose-up moment generated by the action on the tail fin, thereby reducing the difficulty of pitch control. Therefore, when the rotational speed and throttle state of the tiltrotors in front of and behind the center of gravity G are the same, a nose-down moment is generated due to the difference in the length of the lever arms of the center of gravity G. This nose-down moment cancels out or partially cancels out the nose-up moment generated by the tiltrotor's downwash region acting on the tail. Thus, vertical take-off and landing aircraft can trim the pitching moment when the throttle states of the front and rear rotors are the same.

[0055] In the vertical take-off and landing (VTOL) aircraft of the present invention, the rotation axes between the tiltrotors may or may not be parallel in various states, such as cruising (flying horizontally), vertical take-off (taking off and landing vertically), and mode transition states (including transitions from vertical take-off to cruising and from cruising to vertical take-off). In some embodiments, during flight, the projections of the rotation axis of any of the tiltrotors at the tail fin 30 and the rotation axis of any of the tiltrotors at other positions on the plane of symmetry 60 of the aircraft body 10 are not parallel. These non-parallel rotation axes allow different tiltrotors to provide moments in different directions, thereby controlling the pitching moment of the VTOL aircraft as a whole.

[0056] The tilt speeds between each tiltrotor may be the same, and pitch control can be performed by other means, but in one embodiment of the vertical take-off and landing aircraft of the present invention, in various states such as cruising (flying horizontally), vertical take-off (taking off and landing vertically), and mode conversion states (including conversion from vertical take-off to cruising and from cruising to vertical take-off), there is a first difference between the tilt speed of any tiltrotor on the tail fin 30 and the tilt speed of any tiltrotor at other positions, and this first difference is not zero. By setting a threshold for the first difference, the tiltrotors can be tilted to obtain a larger pitch control moment, thereby adjusting the pitch force of the vertical take-off and landing aircraft in various states.

[0057] While the rotational speed of each tiltrotor may be made the same and pitch control may be performed by other means, in one embodiment of the vertical take-off and landing aircraft of the present invention, in various states such as cruising (flying horizontally), vertical take-off (taking off and landing vertically), and mode change states (including conversion from vertical take-off to cruising and from cruising to vertical take-off), there is a second difference between the rotational speed of any tiltrotor on the tail fin 30 and the rotational speed of any tiltrotor at other positions, and this second difference is not zero. By setting a threshold for the second difference, the pitch force in various states of the vertical take-off and landing aircraft can be adjusted by the rotational speed of the tiltrotors to obtain a larger pitch control moment. Furthermore, by combining the first difference in tilt speed between the forward tiltrotor (the tiltrotor located in front of the direction of travel of the tiltrotor on the tail wing 30) and the tiltrotor on the tail wing 30 with the second difference in rotation speed between the forward tiltrotor and the tiltrotor on the tail wing 30, it is possible to achieve control of the pitching moment for multiple strategies.

[0058] The present invention may include only 2N tiltrotors, as shown in Figures 1 to 4, but preferably, referring to Figure 7, in another embodiment of the present invention, the vertical take-off and landing aircraft includes 2M fixed rotors, where M is a natural number greater than or equal to 2, and the value of M may be the same as or different from the value of N. The 2M fixed rotors are mounted symmetrically on the wings on both sides of the aircraft and are located outside the tiltrotors, and in the vertical take-off and landing state, the horizontal projections of all the fixed rotors are all center-symmetric with respect to point A, where point A is located in the plane of symmetry of the aircraft and coincides with point G, or is located on the side of point G closer to the tail fin. In the process of the vertical take-off and landing aircraft transitioning from the vertical take-off and landing state to the cruising state, points G and B move along the plane of symmetry toward the nose side, where point G is located on the side of point A closer to the nose, or coincides with point A, and point B is always located on the side of point A closer to the tail fin.

[0059] Specifically, with the nose of the vertical take-off and landing aircraft pointed forward, the center point B of the 2N tiltrotors is located behind the center point A of the 2M fixed rotors, and the distance from point A to point B is L2, where L2 > 0. As the 2N tiltrotors tilt forward, the centers of gravity of the 2N tiltrotors, the center of gravity G of the vertical take-off and landing aircraft, and the center of symmetry B move toward the nose. Throughout the entire process of tilting the 2N tiltrotors from a preset vertical take-off and landing position (e.g., tilt angle 90°) to a preset cruising position (e.g., tilt angle 0°), L2 > 0, and simultaneously, the center of gravity G of the vertical take-off and landing aircraft is located in front of the center of symmetry B of the 2N tiltrotors, and also in front of the center of symmetry A of the 2M fixed rotors. The distance from point A to point G is L1, where L1 ≥ 0. As the 2N tiltrotors tilt forward, the center of gravity G gradually moves forward, and the absolute value of L1 increases. In this layout, the center of gravity of the vertical take-off and landing aircraft does not coincide with the center of symmetry of the fixed rotor or the center of symmetry of the tiltrotor, and during the transition of the vertical take-off and landing aircraft from vertical take-off and landing to cruising, both points G and B move toward the nose along the plane of symmetry, with point G located on the side of point A closer to the nose or coinciding with point A, and point B always located on the side of point A closer to the tail fin 30. Therefore, the traction force generated by the tiltrotors and fixed rotors in front of the center of gravity has a small moment with respect to the center of gravity G, while the traction force generated by the tiltrotors and fixed rotors behind the center of gravity has a large moment with respect to the center of gravity G. The difference in moment between the forward and aft rotors allows the tiltrotor's downwash region at the tail to resist a portion of the nose-up moment generated by the tail's downwash region, thereby reducing the difficulty of pitch control. Therefore, when the rotational speed and throttle state of the fixed rotors or tiltrotors on both the forward and aft sides of the center of gravity G are the same, the difference in the length of the lever arms at the center of gravity G generates a nose-down moment, which can offset or partially offset the nose-up moment generated by the tiltrotor's downwash region at the tail.Therefore, vertical take-off and landing aircraft can better trim their pitching moment when the throttle states of the front and rear rotors are matched.

[0060] Furthermore, during the tiltrotor's tilt transformation phase, air interference generates an additional large nose-up moment for the vertical take-off and landing (VTOL) aircraft. As the tiltrotor tilts, the VTOL aircraft's center of gravity G gradually moves towards the nose as the tilting process progresses, and the tiltrotor's center of symmetry B also gradually moves forward of the nose. Because the center of gravity G is always ahead of point B, throughout the entire tilting phase, a partial nose-down moment is also generated by the moment difference between the tiltrotor forward of the center of gravity G and the tiltrotor backward of the center of gravity G, which can offset or partially offset the nose-up moment caused by air interference.

[0061] Furthermore, during the tilting and cruising phases, the center of gravity G is closer to the forward side of points A and B, and point B is always located closer to the tail fin 30 than point A. As a result, there is a large margin of static stability in the longitudinal and forward directions, the aircraft has a high ability to withstand extremely strong winds, and flight becomes safer.

[0062] Referring to Figure 26, in one embodiment of the vertical take-off and landing aircraft of the present invention, the tail fin 30 is a V-tail fin, a third tiltrotor 43 is attached to one wingtip of the V-tail fin, and a fourth tiltrotor 44 is attached to the other wingtip. The third tiltrotor 43 and the fourth tiltrotor 44 are symmetrical with respect to the aircraft's plane of symmetry 60. In the vertical take-off and landing state, the distance between the rotation center of the tiltrotor on the tail fin and the leading edge of the wingtip of the V-tail fin is t1, along a direction parallel to the aircraft's roll axis X, the chord length of the wingtip of the V-tail fin is t2, and the ratio of t1 to t2 is 15% to 40%. According to stress analysis of the wingtip structural strength of the tail fin 30, the tilt mechanism is connected to the wingtip within that range of the tail fin 30, the tail fin 30 is thick, and the stress state of the tail fin 30 is good.

[0063] In one embodiment of the vertical take-off and landing aircraft of the present invention, the tiltrotor located on the tail is a fully tilted tiltrotor, and the tiltrotors other than those on the tail are partially tilted tiltrotors. In other embodiments, the tiltrotors other than those on the tail may also be fully tilted tiltrotors, and in particular, if the tiltrotors are mounted on the wingtips, the tiltrotors located on the wingtips of the main wings may also be fully tilted tiltrotors. Many conventional fully tilted tiltrotors include a rotor and a nacelle 442, the nacelle 442 of which a motor and control devices may be mounted. In the "partially tilted tiltrotor" described above, the nacelle 442 is often cut off. During the process of the rotor tilting, the portion closer to the rotor tilts with the rotor, while the portion further from the rotor remains fixed relative to the aircraft 10. In the "fully tilted tiltrotor" described above, the entire nacelle 442 tilts with the corresponding rotor. If mounting conditions permit, all 2N tiltrotors may be fully tilted tiltrotors. However, considering that conventional tiltrotors on the fuselage 10 or main wing 20 in front of the tail fin 30 are mainly attached to arms, as shown in Figures 1 to 6, in the embodiment of the present application, the tiltrotors on the tail fin 30 are fully tilted tiltrotors, and the tiltrotors on the fuselage 10 or main wing 20 in front of the tail fin 30 are partially tilted tiltrotors. The specific position and mounting method of the partially tilted tiltrotors on the main wing 20 or fuselage 10 are not limited. For example, they may be directly attached to the main wing 20, or they may be attached to the main wing 20 or fuselage 10 via arms.

[0064] 2N tiltrotors are symmetrically arranged around the fuselage 10, and a fully tiltable tiltrotor is mounted on the tail fin 30. The fully tiltable tiltrotor is rotatably mounted on the tail fin 30 via a nacelle 442, so that when the fully tiltable tiltrotor tilts, the nacelle 442 also tilts in sync, and the nacelle 442 and rotor on the tail fin 30 are always on the same axis of rotation. This reduces the shielding area of ​​the nacelle 442 on the corresponding rotor when the vertical take-off and landing aircraft is hovering, reducing the area on the tail fin 30 where the rotor downwash strikes, and reducing the mutual interference between the airflow generated by the rotor and the airflow generated by the tail fin 30, thereby reducing the nose-up moment of the aircraft while hovering, and thus making it easier to control the pitching moment of the vertical take-off and landing aircraft in complex interfering flow fields.

[0065] The fully tilting tiltrotor structure in this application may be any conventional tiltrotor structure that enables the entire nacelle 442 to tilt synchronously with the rotor. Referring to Figures 10 and 18, in one embodiment of the vertical take-off and landing aircraft of this application, the fully tilting tiltrotor includes a first rotor 441 and a nacelle 442, the first rotor 441 being connected to the nacelle 442, the nacelle 442 being rotatably connected to the tail fin 30 or the main wing 20, and the nacelle tilting synchronously with the first rotor 441 during the tilting process of the first rotor 441. The housing of the nacelle 442 contains a power unit, which in a purely electric configuration includes, for example, a motor, an electronic control unit, an environmental control unit, and a tilt mechanism. In a fuel-driven configuration, the nacelle includes an engine, an ECU, and a tilt mechanism. Of course, it is preferable that this solution adopts a purely electric configuration.

[0066] Referring to Figure 10, in one embodiment of the vertical take-off and landing aircraft of the present invention, a fairing 4413 is provided in the center of the first rotor 441, and the fairing 4413 is mounted on the windward side of the first rotor 441 and is used to reduce airflow drag. Preferably, the projection of the fairing 4413 covers the projection of the nacelle 442 along the direction of extension of the rotation axis of the first rotor 441. This configuration reduces the influence of the nacelle 442 on the rotor's downwash region during flight. However, those skilled in the art will understand that the projection of the nacelle 442 can also be partially located within the projection of the fairing 4413 to partially reduce drag, but the effect will be weaker than when it is completely covered.

[0067] The shape of the nacelle 442 includes, but is not limited to, a body of revolution, a cube, and an ellipsoid. Preferably, in one embodiment of the vertical take-off and landing aircraft of the present invention, the nacelle 442 is a body of revolution structure, the axis of rotation of the body of revolution structure is coaxial with the axis of rotation of the first rotor 441, and the surface of the nacelle 442 is streamlined. In this way, the influence of the nacelle 442 on the downwash region of the corresponding rotor during flight can be reduced.

[0068] Referring to Figures 10 to 17, the first rotor 441 includes a propeller 4411 and a rotary drive unit 4412, the propeller 4411 being mounted on the output shaft of the rotary drive unit 4412, and the nacelle 442 includes a nacelle housing 4431 and a tilt mechanism located within the nacelle housing 4431, the tilt mechanism driving the rotary drive unit 4412 to tilt. The tilt mechanism in this application may be any suitable type of tilt mechanism capable of synchronously tilting the nacelle 442 and the rotary drive unit 4412. Preferably, in this embodiment, the tilt mechanism includes a rocker arm 4422, a drive arm 4429, a tilt drive unit 4421, and a connecting rod 4426. The tilt drive unit 4421 may be any suitable structural type with a rotary output shaft, such as a rudder servo, and a combination of a rudder servo and a gearbox. In this application, the tilt drive unit 4421 is a rudder servo. The rocker arm 4422 is rotatably mounted to the tail fin 30, and the end of the rocker arm 4422 closest to the rotary drive unit 4412 is fixedly connected to the rotary drive unit 4412. The drive arm 4429 is rotatably mounted to the tail fin 30, and the axis of rotation of the drive arm 4429 is set parallel to the axis of rotation of the rocker arm 4422. The base of the tilt drive unit 4421 is fixedly mounted to the tail fin 30, and the drive end of the tilt drive unit 4421 rotates the drive arm 4429. One end of the connecting rod 4426 is hinged to the rocker arm 4422 by a first hinge shaft 4427, and the other end of the connecting rod 4426 is hinged to the drive arm 4429 via a second hinge shaft 4428. While bearings 4430 are not required to be attached to the first hinge shaft 4427 and the second hinge shaft 4428, preferably, bearings 4430 are provided between the connecting rod 4426 and the first hinge shaft 4427, and also between the connecting rod 4426 and the second hinge shaft 4428, thereby making the tilting process more stable. The above tilt mechanism can realize a two-axis connection structure with the tail fin 30 via the rotation axis of the rocker arm 4422 and the rotation axis of the drive arm 4429.By increasing the axial distance from the axis of rotation between the rocker arm 4422 and the tail fin 30 to the axis of rotation between the drive arm 4429 and the tail fin 30, the moment borne by the tilt mechanism can be increased, reducing the force acting on a single rod, strengthening the mechanism's anti-torsion capability, and improving the mechanism's support rigidity. With this configuration, the tilt mechanism can be fixed to the aircraft using a single-sided support system, reducing the drive requirements via the four-bar linkage mechanism, and allowing for easy adjustment of the overall rigidity and natural frequency of the mechanism by adjusting the length ratio of the four-bar linkage mechanism.

[0069] Referring to Figures 11 to 17, preferably in this embodiment, a first shaft body 4423 and a second shaft body 4424 are fixedly installed on the tail wing 30, parallel to each other, with one end of the first shaft body 4423 and the second shaft body 4424 fixed to the wingtip of the tail wing 30, the other ends of the first shaft body 4423 and the second shaft body 4424 protruding in a cantilevered manner, the rocker arm 4422 is rotatably attached to the first shaft body 4423 via a bearing 4430, the drive arm 4429 is rotatably attached to the second shaft body 4424 via a bearing 4430, the base of the tilt drive device 4421 is rotatably attached to the first shaft body 4423 via a holding structure 4425 and positioned along the axial direction of the first shaft body 4423, the drive end of the tilt drive device 4421 is coaxial with the second shaft body 4424 and fixedly connected to the drive arm 4429. This configuration allows for the positioning and mounting of the tilt drive unit 4421 via the holding structure 4425 and the second shaft body 4424, while also reducing the difficulty of mounting the tilt mechanism to the housing of the nacelle 442 and improving the stability of the one-sided mounting of the fully tilting tilt rotor. In other embodiments, the tilt drive unit 4421 may be fixedly mounted to the tail fin 30, with the output shaft of the tilt drive unit 4421 protruding and fixedly connected to the drive arm 4429, thereby driving the connecting rod 4426 to tilt the rocker arm 4422. However, this configuration requires more usable space inside the tail fin 30 compared to this embodiment and is not suitable when the airfoil is thin or there are many internal components. At the same time, since the torque acting on the tilt drive unit 4421 is ultimately absorbed entirely by the mounting seat of the tilt drive unit 4421, the requirements for the mounting strength of the tilt drive unit 4421 are high.

[0070] Referring to Figures 11-17, the first shaft 4423 and the second shaft 4424 can be reinforcing each other's positions by the aforementioned holding structure 4425, thereby strengthening the structural strength of the first shaft 4423 and the second shaft 4424 installed on the tail fin 30. This ensures that the first shaft 4423 and the second shaft 4424 maintain a parallel positional relationship with each other on the tail fin 30, and that the drive unit and the rocker arm 4422 remain in the same horizontal plane during the tilt drive process, thereby enabling smooth tilt drive of the rocker arm 4422 by the drive unit. One end of the holding structure 4425 is fixedly connected to the base of the tilt drive unit 4421. Specifically, one end of the holding structure 4425 is provided around the base of the tilt drive unit 4421. More specifically, one end of the clamping structure 4425 is fixedly connected to the base of the tilt drive device 4421 and is coaxially mounted on the second shaft body 4424. The other end of the clamping structure 4425 surrounds and firmly holds the first shaft body 4423. Specifically, as shown in Figure 15, the other end of the clamping structure 4425 is interlocked around the first shaft body 4423, achieving a tight clamping with the first shaft body 4423. More specifically, a slit 44251 is provided at the other end of the clamping structure 4425, and the slit 44251 extends from the outer edge of the other end of the clamping structure 4425 to the inner wall that contacts the first shaft body 4423, with the gap of the slit 44251 extending along the axial direction of the first shaft body 4423. The other end of the clamping structure 4425 utilizes the elastic stretching properties of the slit 44251 to surround and clamp the first shaft body 4423 of different outer diameters, thereby expanding the tolerance range allowed by the interference fit between the clamping structure 4425 and the first shaft body 4423.

[0071] In this application, the first shaft 4423 and / or the second shaft 4424 are hollow shafts through which cables or pipes can pass, thereby preventing damage to the cables or pipes from being damaged by irregular shaking outside, and also preventing damage to the cables or pipes by narrowing their range of motion. Considering that the second shaft 4424 needs to withstand a greater load, in this embodiment it is preferable that the first shaft 4423 is a hollow shaft and the second shaft 4424 is a solid shaft. Of course, in other embodiments, if the second shaft 4424 can withstand a greater load, the second shaft 4424 can also be a hollow shaft, or both the first shaft 4423 and the second shaft 4424 can be hollow shafts.

[0072] Furthermore, considering that the entire tilt mechanism is mounted on one side, as the extension distance of the first shaft 4423 increases, the bending moment of the first shaft 4423 increases. Preferably, referring to Figure 12, in one embodiment, the first shaft 4423 and the tail fin 30 are attached to an axially long fitting surface 4432. By providing it in this way, the mounting length of the first shaft 4423 increases, the contact area between the first shaft 4423 and the base of the tilt drive device increases, and the bending moment received by the first shaft 4423 is balanced, thereby reducing the bending deformation of the first shaft 4423 and improving the mounting stability of the entire mechanism.

[0073] In one embodiment of the present invention, the distance from the rotation center of the tilt drive device 4421 to the axis of the second hinge shaft 4428 is a, the distance from the axis of the first hinge shaft 4427 to the axis of the second hinge shaft 4428 is b, the distance from the center of the first shaft body 4423 to the axis of the first hinge shaft 4427 is c, and the distance from the center of the first shaft body 4423 to the center of the second shaft body 4424 is d. a is smaller than b, c, and d, respectively, c is larger than b and d, respectively, and the sum of a and c is smaller than the sum of b and d. Thus, once the tilt drive unit 4421 becomes uncontrollable, the output shaft of the tilt drive unit 4421 will always be in a forward or reverse rotation state, causing the rocker arm 4422 to swing. That is, the rocker arm 4422 rotates clockwise to its limit position, then counterclockwise, and after rotating counterclockwise to its limit position, it rotates clockwise. In this way, the range of motion of the tilt mechanism is limited, preventing the tilt mechanism from rotating beyond its range of motion and colliding with other parts, and preventing the front propeller from colliding with the aircraft body or other structures due to excessive tilting, thereby preventing damage to the aircraft body or other structures. Therefore, this invention limits the tilt angle range of the tilt mechanism by adjusting the length ratio of each part in the tilt mechanism, and at the same time cleverly solves the position restriction problem of the tilt mechanism, so that the limit position of the forward rotation and the limit position of the reverse rotation of the tilt mechanism are both at the same limit position of the connecting rod. This ingenious design not only effectively ensures the safety of the mechanism, but also simplifies the design by eliminating the need to install a separate limiting mechanism to restrict the limit position, and significantly optimizes the mounting space for the transmission link mechanism during movement.Specifically, by adjusting the length ratio of each component in the connecting rod mechanism, the distance between the end axis of the drive arm 4429 (i.e., the axis of the second hinge axis 4428) and the axis of the first shaft body 4423 becomes equal in the initial angle state (i.e., minimum tilt angle) and the final angle state (i.e., maximum tilt angle). At this time, the angle between the plane formed by the end axis of the drive arm 4429 and the axis of the first shaft body 4423 in the initial angle state and the plane formed by the end axis of the drive arm 4429 and the axis of the first shaft body 4423 in the final angle state becomes equal to the rotation angle of the rocker arm 4422 ± 5°. In other words, two directions can be constrained at a single positional restraint point.

[0074] In this embodiment, the tilt mechanism has different reduction ratios depending on the tilt angle, satisfying the control accuracy requirements at different angles, while reducing space requirements through its compact structure and centralized arrangement. Furthermore, by using multiple links, the reduction ratio changes according to the angle, allowing the reduction ratio to be increased under heavy loads and decreased under light loads, thereby reducing the peak value of the drive moment and lowering the drive requirements. Moreover, the tilt drive device of this application has a rotary drive structure, and the driving force can be increased by a brake or reduction ratio adjustment device inside the tilt device. On the other hand, it overcomes the torque borne by the execution side and achieves the ability to maintain the state at any position, maintain the current state position even if the drive device fails, and continue operation after the drive device is restored. Furthermore, the arrangement of structural components can be further centralized.

[0075] As shown in Figures 1 to 7 and Figures 20 to 23, in one embodiment of the vertical take-off and landing (VTOL) aircraft of the present invention, the rotation axis of the tiltrotor is tilted in the range of -20° to 110° with respect to the roll axis X and upward as the positive direction. Referring to Figure 21, 0° indicates that the rotation axis of the tiltrotor extends forward along the direction of the roll axis X. Referring to Figure 23, 90° indicates that the rotation axis of the tiltrotor is tilted upward along the vertical direction, and referring to Figure 22, the rotation axis of the tiltrotor is tilted between 0° and 90°. When the tilt angle range is 90° to 110°, the VTOL aircraft can fly with its nose pointed forward and backward, so the flight envelope and capabilities of the VTOL aircraft are greatly expanded and the risk of the VTOL aircraft needing to change direction in mid-air is reduced. When a vertical takeoff and landing aircraft needs to take off, the tilt angle of all inner tiltrotors can be set to any angle between 0° and 90°, for example, 0°, 30°, 45°, 60°, or 90°, depending on the flight control requirements.

[0076] The number of tiltrotors on the tail fin 30 may be any even number less than 2N. Referring to Figures 5-9, preferably in this embodiment, the vertical take-off and landing aircraft comprises four tiltrotors and four fixed rotors, the four fixed rotors being mounted symmetrically on both sides of the fuselage 10, the four tiltrotors being located inside the wingspan of the four fixed rotors, two of the tiltrotors being mounted on the tail fin 30, and the remaining two tiltrotors being mounted on the fuselage 10 or the main wing 20 in front of the main wing 20. Specifically, the four tiltrotors are divided into two groups of the same number, which are denoted as the first tiltrotor group and the second tiltrotor group. A first tiltrotor group is mounted on the fuselage 10 or main wing 20 located in front of the center of gravity G of the vertical take-off and landing aircraft, and a second tiltrotor group is mounted on the tail wing 30 located behind the center of gravity G of the vertical take-off and landing aircraft. The first tiltrotor group includes a first tiltrotor 41 and a second tiltrotor 42, and the second tiltrotor group includes a third tiltrotor 43 and a fourth tiltrotor 44. The first tiltrotor 41 is mounted on the third arm 411, and the second tiltrotor 42 is mounted on the fourth arm 421 and is mounted symmetrically to the first tiltrotor 41 with respect to the plane of symmetry 60 of the fuselage 10. The third tiltrotor 43 is mounted on the tail wing 30 via a first nacelle 442, and the fourth tiltrotor 44 is mounted on the tail wing 30 via a second nacelle 442. The first nacelle 442 and the second nacelle 442 are mounted symmetrically with respect to the plane of symmetry 60 of the aircraft 10, and the fourth tiltrotor 44 and the third tiltrotor 43 are mounted symmetrically with respect to the plane of symmetry 60 of the aircraft 10.

[0077] In the vertical takeoff and landing configuration, the rotation axes of all four tiltrotors are tilted upward along the vertical direction, and the first tiltrotor 41, second tiltrotor 42, third tiltrotor 43, and fourth tiltrotor 44 are all distributed on a first circle centered at point B. The horizontal projections of the first tiltrotor 41 and the fourth tiltrotor 44 are center-symmetric with respect to point B, and the horizontal projections of the second tiltrotor 42 and the third tiltrotor 43 are also center-symmetric with respect to point B.

[0078] Referring to Figure 7, the four fixed rotors are each divided into two groups of the same number, which are denoted as the first fixed rotor group and the second fixed rotor group. The first fixed rotor group is mounted on the main wing 20 in front of the center of gravity of the vertical take-off and landing aircraft, and the second fixed rotor group is mounted on the main wing 20 behind the center of gravity of the vertical take-off and landing aircraft. The first fixed rotor group includes the first fixed rotor 51 and the second fixed rotor 52, and the second fixed rotor group includes the third fixed rotor 53 and the fourth fixed rotor 54, with the first fixed rotor 51, the second fixed rotor 52, the third fixed rotor 53, and the fourth fixed rotor 54 all distributed on a second circle centered at point A. The first fixed rotor 51 and the second fixed rotor 52 are symmetrical with respect to the plane of symmetry 60 of the aircraft 10, and the third fixed rotor 53 and the fourth fixed rotor 54 are also symmetrical with respect to the plane of symmetry 60 of the aircraft 10. The rotation axes of the four fixed rotors all protrude upward. The horizontal projections of the first fixed rotor 51 and the fourth fixed rotor 54 are symmetrical with respect to point A, and the horizontal projections of the second fixed rotor 52 and the third fixed rotor 53 are symmetrical with respect to point A. In this application, the front side refers to the direction of extension toward the nose of the aircraft, and the rear side refers to the direction of extension toward the tail fin 30.

[0079] In one embodiment of the vertical take-off and landing aircraft of the present invention, the distance between the four fixed rotors in the direction of extension of the aircraft body, i.e., the distance between the second fixed rotor 52 and the fourth fixed rotor 54, or the distance between the first fixed rotor 51 and the third fixed rotor 53, is L3. If the distance between the four tiltrotors in the direction of extension of the aircraft body, i.e., the distance between the second tiltrotor 42 and the fourth tiltrotor 44, or the distance between the first tiltrotor 41 and the third tiltrotor 43, is L4, then 0.1(L3+L4)≧4L1+2L2≧0.01(L3+L4). By providing this configuration, during the tilting and cruising phases, the center of gravity G is closer to the front of points A and B, and point B is always located on the side of point A closer to the tail fin 30. Therefore, there is a large static stability margin in the longitudinal and forward directions, the aircraft has a high ability to withstand extremely strong winds, and flight is safer.

[0080] Referring to Figure 7, a first arm 511 is attached to one side of the main wing 20 of the aircraft body 10, and a second arm 521 is attached to the other side of the main wing 20 of the aircraft body 10. The first arm 511 and the second arm 521 are attached symmetrically with respect to the plane of symmetry 60 of the aircraft body 10, and the 2M fixed rotors are attached symmetrically to the first arm 511 and the second arm 521 on both sides of the aircraft body 10, respectively, and are located at the front and rear sides of the main wing 20, and at the front and rear ends of the first arm 511 and the second arm 521, respectively. At the same time, the horizontal projections of all the fixed rotors are all approximately centrally symmetric with respect to point A, with two corresponding points.

[0081] Referring to Figures 1 to 4, in one embodiment, unlike the vertical take-off and landing aircraft in Figures 5 to 9, the vertical take-off and landing aircraft in this embodiment has only four tiltrotors and no fixed rotors. The four tiltrotors are mounted symmetrically on both sides of the fuselage 10, and two of the tiltrotors are mounted symmetrically to the main wings 20 relative to the fuselage 10. The structure of the four tiltrotors and their mounting relationship to the fuselage are the same as those in Figures 5 to 9 and Figure 34. The specific positions are adaptively adjusted depending on the aircraft and will not be described further here.

[0082] Unlike the vertical take-off and landing aircraft in Figures 1 to 4, in another embodiment of the vertical take-off and landing aircraft of the present invention, the vertical take-off and landing aircraft has six tiltrotors symmetrically mounted on both sides of the fuselage 10 and no fixed rotors. Four of the tiltrotors are symmetrically mounted on the wings 20 relative to the fuselage 10, and the remaining two tiltrotors are symmetrically mounted on the tail wing 30. The tail wing 30 is a V-tail, and the two tiltrotors located on the tail wing 30 are fully tiltable tiltrotors, and the two fully tiltable tiltrotors are mounted on the wingtips on both sides of the upper part of the V-tail, respectively. The two tiltrotors are provided on the wingtips of the wings 20 on both sides of the fuselage 10, and are symmetrical with respect to the plane of symmetry 60 of the fuselage 10. The tiltrotors located at the wingtips of the main wing 20 are preferably fully tilted tiltrotors, while the remaining two tiltrotors are attached to the front of the main wing 20 via arms and are partially tilted tiltrotors. If conditions permit, the tiltrotors attached to the front of the main wing 20 via arms may also be fully tilted tiltrotors.

[0083] In this application, the mounting position and structure of the ruddervator 31 are diverse and may be provided at any suitable position on the tail fin 30, or any suitable conventional ruddervator 31 structure may be used. Specifically, referring to Figures 5 and 17, the ruddervator 31 includes a rudder plate 311 and a rudder body drive unit (not shown), the rudder plate 311 being rotatably connected to the tail fin 30 or the tail of the aircraft body 10, and the rudder body drive unit rotates the rudder plate 311 to adjust the direction of the vertical take-off and landing aircraft. The drive unit of the ruddervator 31 includes, but is not limited to, a motor or a combination of a motor and a reduction gear.

[0084] Referring to Figure 25, in some embodiments, the ratio of the chord length of the rudder plate 311 to the chord length of the tail wing 30 is 15% to 100%, and may be any value between 15% and 100%, such as 15%, 30%, 45%, 60%, 90%, or 100%. For example, as shown in Figure 18, the chord length of the rudder plate 311 accounts for 30% of the chord length of the tail wing 30, and as shown in Figure 19, the chord length of the rudder plate 311 accounts for 60% of the chord length of the tail wing 30. As shown in Figure 20, the chord length of the rudder plate 311 accounts for 100% of the chord length of the tail wing 30, but is not limited to the ratios in Figures 18 to 20. Referring to Figure 25, the chord length is the distance from the leading edge to the trailing edge of the airfoil cross-section of the tail wing 30. In the top view, the chord length ratio is the ratio of the length of the rudder plate in the direction of travel to the length of the tail wing 30 (not the length ratio along the wingspan Y direction).

[0085] Referring to Figure 18, in one embodiment of the vertical take-off and landing aircraft of the present invention, with the roll axis X as the reference and upward as the positive direction, the deflection angle of the rudder plate 311 is -90° to 30°. For example, as shown in Figure 23, the deflection angle of the rudder plate 311 may be any angle within the range of -90° to 30°, such as -90°, -60°, -30°, -15°, 0°, 15°, or 30°.

[0086] In the present invention, the rotation axes of the 2N tiltrotors and 2M fixed rotors may be aligned vertically in the vertical takeoff and landing configuration. Referring to Figure 9, preferably, in one embodiment of the vertical takeoff and landing aircraft of the present invention, the 2N tiltrotors are mounted symmetrically on both sides of the aircraft body 10, and during the tilting process, the angle α between the plane formed by the rotation axes of the rotors (i.e., the plane formed by the rotation axes of the rotors rotating around the tilt axis) and the plane of symmetry 60 of the aircraft body 10 is -15° to +15°, and may be any angle between -15° and +15°, such as -15°, -10°, 0°, 10°, 15°, etc., and is positive from bottom to top and outward in the wingspan direction, and becomes negative when tilted from bottom to top towards the plane of symmetry 60. The angle between the rotation axis of the 2M fixed rotors and the plane of symmetry of the aircraft is -15° to +15°, and may be any angle between -15° and +15°, such as -15°, -10°, 0°, 10°, 15°, etc., and is positive from bottom to top outward in the wingspan direction, and becomes negative tilted from bottom to top toward the plane of symmetry 60.

[0087] In this application, the tail fin 30 may be any one of the following: V-tail, Y-tail, H-tail, X-tail, T-tail, H-tail, or U-tail. The tiltrotor of the tail fin 30 is mounted on the upper side of the tail fin 30 and tilts upward in the vertical takeoff and landing state. This reduces the possibility of the rotor causing harm to passengers when boarding or disembarking the aircraft. Referring to Figures 1 to 4, in this embodiment, the tail fin 30 is a V-tail, and two of the tiltrotors are mounted on the tail fin 30, with the two tiltrotors mounted on the wingtips on both sides of the upper part of the tail fin 30. In other embodiments, the tail fin 30 may be any one of the shapes described above.

[0088] In this embodiment, when the vertical take-off and landing aircraft is flying forward and in a cruising state, the rotation axis of the tiltrotor at the tail fin 30 and the rotation axis of the tiltrotor at other positions may extend along the roll axis X. However, in other embodiments, the rotation axis of the tiltrotor at the tail fin 30 and the rotation axis of the tiltrotor at other positions do not have to extend along the roll axis X, and may be within a range of ±20° with respect to the roll axis X in the vertical plane where the roll axis X is located. In this application, tiltrotors at roll axis X coordinate positions with different projections along the roll axis X direction may have their nacelles 442 controlled independently of each other, and their tilts may also be controlled independently without correlation. In this mode, the tilt angles of the front tiltrotor and the tiltrotor at the tail fin 30 do not have to be the same, and the tilt process does not have to be synchronous. For example, if we define the roll axis X as 0°, the upper angle of the roll axis X as positive, and the lower angle of the roll axis X as negative, then the inclination angle of the nacelle 442 of the tiltrotor on the front fuselage 10 or wing 20 will be 10°, and the inclination angle of the nacelle 442 of the tiltrotor on the rear fuselage 10 or wing 20 will be -10°.

[0089] Furthermore, when the vertical take-off and landing aircraft of this invention is in a vertical take-off and landing state, the rotation axis of the tiltrotor at the tail fin 30, and the rotation axes of the tiltrotors at other positions, may or may not extend upward along the vertical direction. In other words, the tilt angle of the front tiltrotor and the tiltrotor at the tail fin 30 is not limited to a tilt angle of 90°. To improve control capability, the tilt angle of the rotation axis of the front tiltrotor and the tilt angle of the rotation axis of the tiltrotor at the tail fin 30 may be any value between 70° and 110°, such as 70°, 80°, 90°, 100°, and 110°. The rotation control and tilt control of each of the 2N tiltrotors are independent of each other, and the tilt angles of the 2N tiltrotors may be perfectly identical, any two may be different, or may be partially identical. To illustrate with an example, among the 2N tiltrotors, the tilt angles of multiple tiltrotors located in front of the tail fin 30 on the fuselage 10 or main wing 20 are the same and are referred to as the first tilt angle. The tilt angles of the tiltrotors on the rear tail fin 30 are also the same and are referred to as the second tilt angle. The first and second tilt angles are not equal; for example, the first tilt angle may be 100° and the second tilt angle may be 80°. This allows for obtaining different pitch trim moments at different positions. The tilt angle is defined as the angle between the rotation axis of the tiltrotor and the roll axis X, with the center point of the tilt axis of the tiltrotor as the vertex.

[0090] Since the rear tiltrotors of the 2N tiltrotors are located on the tail fin 30, and considering that the centers of the 2N tiltrotors are in a predetermined position, the tiltrotors on the tail fin 30 need to be positioned forward. However, considering that increasing the length of the tiltrotor nacelles reduces the overall rigidity of the vertical take-off and landing aircraft and thus reduces safety, in one embodiment of the present invention, the tail fin 30 is configured to sweep forward, and the tiltrotors on the forward-swept tail fin are fully tilted tiltrotors. The angle of forward sweep of the tail fin 30 is not limited, and it is only necessary to balance the layout of the 2N tiltrotors without affecting the overall rigidity of the vertical take-off and landing aircraft.

[0091] In fixed-wing flight mode, considering that the lever arm of the tail 30 is shortened by the forward sweep of the tail 30, the volume of the horizontal stabilizer (horizontal projection of the tail 30) and the vertical stabilizer (vertical projection of the tail 30) of the tail 30 are significantly reduced. Furthermore, the lateral projection area of ​​the fuselage 10 in front of the aerodynamic center becomes larger than the lateral projection area of ​​the fuselage 10 behind the aerodynamic center, further reducing the effective area of ​​the vertical stabilizer. However, indiscriminately increasing the area of ​​the tail 30 would make the overall layout of the aircraft uncontrollable, ultimately resulting in insufficient vertical stabilizer volume. When flying in fixed-wing mode, the lateral stability of the aircraft becomes insufficient, and the direction of travel becomes statically unstable, leading to poor flight quality in lateral dynamic stability mode and making flight control extremely difficult. To solve the above problems, referring to Figure 9, a vertical stabilizer is provided below the forward-sweeping tail 30, and the vertical stabilizer is connected to the tail of the fuselage 10. The structure and number of vertical stabilizers are not limited, and the structure of the vertical stabilizers can refer to any suitable conventional vertical stabilizer structure, and may include one vertical stabilizer body or multiple vertical stabilizer bodies, as long as the tail fin 30 and vertical stabilizers have sufficient vertical tail capacity to stabilize the lateral direction of travel of the vertical takeoff and landing aircraft when flying in fixed-wing mode.

[0092] Referring to Figures 7 and 28-32, the present invention provides a control method for a vertical take-off and landing (VTOL) aircraft. The VTOL aircraft comprises an airframe and 2N tiltrotors. The airframe 10 has a symmetrical structure and a plane of symmetry 60 extending along the length of the airframe 10. Main wings 20 are provided on both sides of the airframe 10, and the main wings 20 on both sides are symmetrical with respect to the plane of symmetry 60 of the airframe 10. A tail fin 30 is provided at the tail end of the airframe 10, and a rudder 31 is provided on the tail fin 30. 2N tiltrotors are mounted on both sides of the airframe 10, where N is a natural number greater than or equal to 2, and the 2N tiltrotors are mounted symmetrically with respect to the plane of symmetry 60 of the airframe 10, and some of the 2N tiltrotors are mounted on the tail fin 30. In the vertical takeoff and landing state, the horizontal projection of the 2N tiltrotor propellers is center-symmetric with respect to point B, and both point B and the center of gravity G of the vertical takeoff and landing aircraft are located within the aircraft's plane of symmetry, with point B located on the side of point G closer to the tail fin, and during the transition of the vertical takeoff and landing aircraft from the vertical takeoff and landing state to the cruising state, both point G and point B move toward the nose side along the plane of symmetry, with point B always located on the side of point G closer to the tail fin.

[0093] In one embodiment of the vertical take-off and landing aircraft of the present invention, During flight, the pitch control ratios of the rudder 31, the 2N tilt rotors, and the 2M fixed rotors are distributed according to the current airspeed or dynamic pressure, and the rudder 31, the 2N tilt rotors, and the 2M fixed rotors are individually controlled according to the pitch control ratios to perform pitch control in order to achieve pitch trim and maneuverability.

[0094] In one embodiment of the vertical take-off and landing aircraft of the present invention, controlling 2N tiltrotors according to the pitch control ratio is: A differential adjustment process of tilt angle to differentially adjust the pitching moment and perform pitch trim and control based on the difference in tilt angle between the tilt rotor of the tail wing 30 and any other tilt rotor on the front side of the tail wing, and / or a differential adjustment process of rotational speed to differentially adjust the pitching moment and perform pitch trim and control based on the difference in rotational speed between the tilt rotor of the tail wing 30 and any other tilt rotor on the front side of the tail wing. and / or, a differential tilt speed adjustment process is included to differentially adjust the pitching moment by the difference in tilt speed between the tilt rotor at the tail and any other tilt rotor at the front of the tail, thereby performing pitch trim and control. The differential tilt angle adjustment process, the differential rotation speed adjustment process, and the differential tilt speed adjustment process described above can be performed independently, in pairs, or all three simultaneously.

[0095] Referring to Figure 28, the flight process specifically includes four stages in order: a ground preparation process S100, a takeoff control process S200, a takeoff-to-level flight control process S300, and a cruising state S400.

[0096] In S100, the ground preparation process is carried out. In the ground preparation process, the vertical take-off and landing aircraft is first started up and a power-on self-test is performed, and then the full stroke state of the servo systems, such as the tilt mechanism and rudder 31, must be checked.

[0097] The S200 performs the takeoff control process. The takeoff control process is the process by which a vertical takeoff and landing aircraft ascends from the ground to a set altitude. During this process, the tiltrotor and fixed rotor maintain a nearly constant tilt angle and rotation speed. Compared to the transition to horizontal flight, this process is more stable.

[0098] The S300 performs the control process from takeoff to level flight. During this process, changes in the tilt rotor angle and / or changes in the rotational speed of the tilt rotor and / or fixed rotor occur frequently. Therefore, the pitch impact force is significant during this process, making control of vertical takeoff and landing aircraft difficult.

[0099] In S400, the aircraft enters cruising mode. In cruising mode, vertical takeoff and landing aircraft fly horizontally and navigate in a horizontal direction, thus maintaining stability.

[0100] Referring to Figure 32, in one embodiment, the takeoff control process in step S200 includes the following steps.

[0101] In S211, the 2N tiltrotors are tilted to a vertical takeoff / landing position or an inclined position (e.g., 0° to 90°) until their rotation axis is vertically upward, thereby providing power for ascent. The vertical takeoff / landing position may be a position where the rotation axis of the tiltrotor makes a 90° angle with respect to the roll axis, and the inclined position may be a position between 0° and 90° (excluding the endpoint value) between the rotation axis of the tiltrotor and the roll axis.

[0102] In S212, the rudder control 31 is tilted downward so that it can participate in takeoff control.

[0103] In S213, 2N of the aforementioned tiltrotors are activated, and once the vertical takeoff and landing aircraft reaches the set altitude, a horizontal flight instruction is issued.

[0104] If the tiltrotors are set to 90° during the takeoff process from S211 to S213, the vertical takeoff and landing aircraft can take off vertically normally and the throttle states of the 2N tiltrotors can be matched. At this time, the tiltrotors on the tail fin 30 have the largest shielding area during the vertical takeoff and transition phases, and a large nose-up moment is generated. If the rotation axes of all tiltrotors are set to an inclination position between 0° and 90°, when the aircraft takes off, the throttle of the inner tiltrotor is activated, and the inner rotor provides forward thrust, so the aircraft climbs diagonally, and as the flight speed gradually increases, the inclination angle of the inner tiltrotor gradually decreases, eventually tilting to 0°, and switching to a fixed-wing mode for horizontal flight. In this solution, the nose-up moment generated by the tail fin 30 shielding the tiltrotors is moderate, the difficulty of control is low, and the maximum thrust margin requirement of the power system is small.

[0105] As an example of optimization, referring to Figure 29, in this embodiment, the control process from takeoff to level flight in S300 includes the following process.

[0106] In S310, the aircraft responds to level flight instructions. Level flight instructions can be issued by the pilot, or by the vertical take-off and landing aircraft itself when it determines that the pre-set flight conditions are met.

[0107] In S320, the rotor control process is performed. This process is heavily influenced by the takeoff control process of S200 and varies greatly depending on the state of the tiltrotor.

[0108] In S330, the pitch control process is performed. During the transition from takeoff to level flight, the tiltrotor tilts multiple times to a predetermined cruising position where the axis of rotation of the tiltrotor is parallel to the roll axis. Considering that a vertical takeoff and landing aircraft experiences a constant pitch impact force during each tilting process, the pitch control process in S330 can be performed after each rotor control process in S320 when transitioning from takeoff to level flight.

[0109] In S340, the rotor control process and pitch control process are executed sequentially until the tiltrotor is tilted to the cruising position, the aircraft takes off, and flies in level flight. The cruising position may be, for example, the position where the rotation axis of the tiltrotor is parallel to the roll axis when the tilt angle is 0°.

[0110] Referring to Figure 31, in one embodiment, the rotor control process in step S320 includes the following steps.

[0111] In S321, the current tilt position of each tilt rotor is obtained. In this process, angle sensors are installed on the tilt rotors to feed back the current tilt position to the central control system, or the corresponding current tilt position can be directly fed back to the central control system via the tilt drive device corresponding to the tilt rotor.

[0112] In S322, if the current tilt position does not coincide with the set cruising position, the current airspeed or dynamic pressure at the corresponding current tilt position of the tiltrotor is obtained, and it is determined whether the current airspeed or dynamic pressure is equal to or greater than the set threshold at the current tilt position. The set cruising position is a preset position for a vertical take-off and landing aircraft in level flight, and may be, for example, the 0° position where the rotation axis of the tiltrotor is parallel to the roll axis X, or any other position between 0° and 5°.

[0113] In S323, if the current airspeed or dynamic pressure is greater than or equal to a preset threshold at the current tilt position, the tilt rotor is tilted to the next preset position.

[0114] In S324, the rotational speed of the 2N tiltrotors is gradually increased. As shown in Figure 1, if a vertical take-off and landing aircraft does not have 2M fixed rotors, forward power can be obtained by gradually increasing the rotational speed of the 2N tiltrotors. However, as shown in Figure 7, if a vertical take-off and landing aircraft has 2N tiltrotors and 2M fixed rotors, in order to achieve horizontal flight, it is necessary to gradually increase the rotational speed of the 2N tiltrotors and gradually decrease the rotational speed of the 2M fixed rotors to a set rotational speed.

[0115] Referring to Figure 7, in another embodiment, the vertical take-off and landing aircraft has 2M fixed rotors, where M is a natural number greater than or equal to 2, and the value of M may be the same as or different from the value of N. The 2M fixed rotors are mounted symmetrically on the main wings on both sides of the aircraft and are located outside the tilt rotors. In the vertical take-off and landing state, the horizontal projections of all the fixed rotors are center-symmetric with respect to point A, where point A is located within the plane of symmetry of the aircraft and coincides with point G, or is located on the side of point G closer to the tail fin. During the transition of the vertical take-off and landing aircraft from the vertical take-off and landing state to the cruising state, points G and B move along the plane of symmetry toward the nose side, where point G is located on the side of point A closer to the nose, or coincides with point A, and point B is always located on the side of point A closer to the tail fin. Referring to Figure 33, in one embodiment, the take-off process at S200 of the vertical take-off and landing aircraft is different from the processes at S211-S213 in Figure 32. The S200's takeoff process is Step S221 involves tilting the 2N tilt rotors until their axis of rotation is horizontal and forward-facing, and parallel to the roll axis X. Step S222 tilts the ladder 31 downwards, Step S223 includes starting 2M of the fixed rotors and 2N of the tilt rotors, and issuing a horizontal flight instruction once the vertical takeoff and landing aircraft has reached a set altitude.

[0116] In the takeoff process from S221 to S223, the rotation axes of all tiltrotors are set to 0° (see Figure 7), and the aircraft converts to compound wing mode. The aircraft takes off vertically using multiple outer rotors. The output of the multiple outer rotors at this time is twice that of control means 1. In the transition phase from vertical takeoff to horizontal flight, the inner tiltrotors are gradually started, and the throttle is gradually increased until the transition to horizontal flight is successful. In this solution, the shielding area of ​​the tail fin 30 of the inner tiltrotor is smallest, the nose-up moment generated is smallest, and the control is simplest. It is basically a compound wing control mode, but it requires a large thrust margin for the power system and has high requirements for the power system. This control mode is reduced except in emergencies.

[0117] A person skilled in the art will understand that in the takeoff process S221-S223, the tiltrotor axis is always parallel to the roll axis X, and therefore, in the rotor control process S320 of the takeoff to level flight control process S300, the related processes S321-S323 are unnecessary. Instead, it is sufficient to gradually increase the rotational speed of 2N tiltrotors and gradually decrease the rotational speed of 2M fixed rotors to a set rotational speed.

[0118] Regardless of whether the rotor control process in S300 includes the tilt control processes in S321 to S324, vertical take-off and landing aircraft experience high pitch impact forces during the transition from take-off to horizontal flight. Referring to Figure 30, in one embodiment of the present invention, the pitch control process in step S330 includes the following steps.

[0119] In S332, the rudder 31 distributes the pitch control ratio to the 2N tiltrotors and 2M fixed rotors according to the current airspeed or dynamic pressure. In this process, the pitch adjustment force is distributed to the rudder 31, the 2N tiltrotors, and the 2M fixed rotors according to a set distribution rule based on the current speed of the vertical take-off and landing aircraft, thereby achieving more balanced pitch control through the pitch control ratio.

[0120] S333, the rudder 31, the 2N tilt rotors, and the 2M fixed rotors are individually controlled according to the pitch control ratio to achieve pitch trim and maneuvering. In this process, the central control system individually controls the rudder 31, the 2N tilt rotors, and the 2M fixed rotors based on the assigned pitch control ratio. For example, this may be differential control of rotational speed, the tilt of the rudder 31, etc., generating different pitch adjustment forces according to different positions of the vertical take-off and landing aircraft, and balancing the pitch impact force when transitioning to horizontal flight.

[0121] In one embodiment, the vertical take-off and landing aircraft of the present invention has 2N tiltrotors installed inside 2M fixed rotors, and a rudder 31 and fully tiltable tiltrotors installed on the tail fin 30. This allows the pitch impact force during the transition to horizontal flight to be balanced through the pitch control process described above, thereby achieving a smoother control process. On the other hand, in the pitch control process, as the tiltrotors on the tail fin 30 are tilted, the nacelle 442 rotates together with the rotors. During hovering, the shielding area of ​​the nacelle 442 in the corresponding rotor is reduced, so the area on the tail fin 30 where the rotor downwash strikes is reduced, thereby reducing a portion of the nose-up moment and improving the control of the pitching moment of the vertical take-off and landing aircraft in complex interfering flow fields.

[0122] Furthermore, if the vertical take-off and landing aircraft does not have 2M fixed rotors, it is not necessary to consider the fixed rotors in step S332. It is sufficient to simply allocate the pitch control ratio of the rudder 31 and the 2N tilt rotors according to the current airspeed or dynamic pressure. Similarly, in step S333, it is not necessary to consider the fixed rotors, and it is sufficient to individually control the rudder 31 and the 2N tilt rotors according to the pitch control ratio to achieve pitch trim and maneuvering.

[0123] Considering that the rudder 31 is frequently involved in the control process during the takeoff process of a vertical takeoff and landing aircraft, the pitch control process of the present invention further includes, before step S332, step S331, which controls the rudder 31 to either return to zero in a timely manner, or to operate to a trim deviation that matches the current airspeed or dynamic pressure, thereby gradually involving it in the pitch control process. For example, an airspeed threshold may be set, and if the current airspeed is greater than or equal to the set airspeed threshold, the rudder 31 is set to zero, i.e., to the initial un-tilted position, at which point the tilt angle of the control surface becomes 0°.

[0124] In this embodiment, controlling the rudder rotor and 2N tiltrotors individually according to the pitch control ratio to achieve pitch trim and steering includes differentially adjusting the pitching moment based on the difference in tilt speed and / or tilt angle and / or the difference in rotation speed of the tiltrotors at different positions, thereby performing pitch trim and steering.

[0125] Referring to Figure 27, Figure 27 is the pitching moment curve for the entire aircraft generated by comparing a partially tilted configuration of the tiltrotor nacelle on the tail wing 30 with a fully tilted configuration of the tiltrotor nacelle on the tail wing 30 through aerodynamic simulation analysis. The first curve 101 is the simulation curve of the simulation model in Figure 7, and the structure of the tail tiltrotor during the aerodynamic simulation analysis is shown in Figure 20. The second curve 102 is the partially tilted model, which differs from the vertical take-off and landing aircraft model in Figure 7 only in the tiltrotor structure on the tail wing 30; otherwise, it is the same as the vertical take-off and landing aircraft model, except for the partially tilted nacelle structure. In the simulation analysis of the two models, the rotational speeds of the four fixed rotors and the four tiltrotors are equal, and the thrust is matched to the aircraft's take-off weight (with thrust and gravity trimmed).

[0126] The first curve 101 and the second curve 102 in Figure 27 are curves showing the change in the pitching moment of the entire aircraft of the two models as it changes with flight speed (i.e., wind speed). The flight speed in the figure has been dimensionless, and the process from 0 to 1 represents the maximum speed of the aircraft flying from the hovering state to the current tilt angle (maintaining the aforementioned 90° tilt angle), i.e., the process of flying from minimum speed to maximum speed at the current tilt angle. The pitching moment in the figure has also been dimensionless.

[0127] As can be seen by comparing the first curve 101 and the second curve 102 in Figure 27, if there is no forward / backward throttle differential between the eight rotors, the aircraft generates a large pitching moment. Therefore, in order to achieve stable flight, the pitching moment must be balanced so that the angular acceleration in the pitch direction is zero. To balance the aerodynamic pitching moment in this area, a differential between the forward and rear rotor throttles is necessary so that the pitching moment generated by the rotor differential cancels out the aerodynamic pitching moment generated by the aircraft body. In the control method for a fully tilted nacelle, a maximum pitch trim and control of 0.25 units is required during flight. In the control method for a partially tilted nacelle 442, a maximum pitch trim and control of 0.8 units is required for the entire aircraft, which is much larger than in the fully tilted control method. The larger the pitching moment generated during flight, the more difficult it is to control the aircraft, and the greater the additional power required by the power system to adjust the aircraft's attitude. Furthermore, as can be seen from the simulation curve shown in Figure 27, as the flight speed increases, the amount of pitch trim and maneuvering required by the control method for a partially tilted nacelle 442 also fluctuates greatly from -0.04 in the hovering state (dimensionless wind speed is 0) to approximately 0.8 (dimensionless wind speed is 0.57), resulting in a very volatile pitching moment, which is very unfavorable for aircraft pitch control. On the other hand, the pitching moment generated by the control method for a fully tilted nacelle 442 changes from -0.25 in the hovering state (dimensionless wind speed is 0) to approximately 0.25 (dimensionless wind speed is 0.42), and the fluctuation of the pitching moment is small, which is advantageous for controlling the aircraft's pitch direction. Therefore, the vertical take-off and landing aircraft and pitch control method provided in this application are simple and effective, and can effectively improve the safety of the aircraft.

[0128] As described above, the vertical take-off and landing (VTOL) aircraft of this application is equipped with elevators and 2N tiltrotors. In the VTOL take-off and landing state, the horizontal projection of the propellers of the 2N tiltrotors is center-symmetric with respect to point B. Both point B and the center of gravity G of the VTOL aircraft are located within the plane of symmetry of the aircraft, with point B located on the side of point G closer to the tail. During the transition of the VTOL aircraft from the VTOL take-off and landing state to the cruising state, point B is always located on the side of point G closer to the tail. In this layout, the center of gravity G of the VTOL aircraft and point B, the center of symmetry of the tiltrotors, do not coincide. Furthermore, during the transition of the VTOL aircraft from the VTOL take-off and landing state to the cruising state, both point G and point B move toward the nose along the plane of symmetry. Therefore, the traction force generated by the tiltrotor in front of the center of gravity has a smaller moment relative to the center of gravity G, while the traction force generated by the tiltrotor behind the center of gravity has a larger moment relative to the center of gravity G. The moment difference between the forward and rear tiltrotors can resist part of the nose-up moment generated by the tiltrotor's downwash region acting on the tail, thereby reducing the difficulty of pitch control. As a result, when the rotational speed and throttle state of the tiltrotors on both sides of the center of gravity G are the same, the difference in lever arm length at the center of gravity G generates a nose-down moment, and this nose-down moment can cancel out or partially cancel out the nose-up moment generated by the tiltrotor's downwash region acting on the tail. Therefore, vertical take-off and landing aircraft can better achieve balance of pitching moments when the throttle states of the front and rear rotors are matched.

[0129] The control method of the present invention allocates the pitch control ratio between the rudder rotor and 2N tilt rotors according to the current airspeed or dynamic pressure, and achieves pitch control through the cooperation of the rudder rotor and 2N tilt rotors. Therefore, this invention effectively overcomes several practical problems in the prior art and has high utility and significance. [Explanation of symbols]

[0130] 10 aircraft 20 Main wing 30 tail fin 31 Ladder-vator 311 Ladder board 41 First tiltrotor 411 Third Arm 42. Second tiltrotor 421 Fourth Arm 43. Third tiltrotor 44. Fourth tiltrotor 441 First rotor 4411 Propeller 4412 Rotary drive unit 4413 Fairing 442 Nasser 4421 Tilt drive unit 4422 Rocker Arm 4423 First Axis 4424 Second Axis 4425 Holding structure 44251 Slit 4426 Connecting Rod 4427 First hinge axis 4428 Second hinge axis 4429 Drive arm 4430 Bearing 51 First fixed rotor 511 First Arm 52 Second fixed rotor 521 Second Arm 53 Third fixed rotor 54. Fourth fixed rotor 60. Plane of symmetry.

Claims

1. It is a vertical takeoff and landing aircraft, The aircraft has main wings on both sides, a tail fin at the rear, and a rudder on the tail fin, The aircraft comprises 2N tiltrotors, some of which are symmetrically mounted on both sides of the aircraft and located on the tail fin. A vertical take-off and landing (VTOL) aircraft in which N is a natural number greater than or equal to 2, and in the VTOL state, the projection of the 2N tiltrotor propellers in the horizontal plane is center-symmetric with respect to point B, both point B and the center of gravity G of the VTOL aircraft are located within the plane of symmetry of the aircraft, point B is located on the side of point G closer to the tail fin, and during a mode change of the VTOL aircraft, both point G and point B move along the plane of symmetry, and point B is always located on the side of point G closer to the tail fin.

2. The vertical take-off and landing aircraft according to claim 1, wherein during flight, the projections of the rotation axes of any of the tiltrotors at the tail fin and any of the tiltrotors at other positions onto the plane of symmetry of the aircraft are not parallel.

3. A vertical takeoff and landing aircraft according to any one of claims 1 to 2, wherein in cruising and / or vertical takeoff and / or mode conversion states, there is a first difference between the tilt velocity of any of the tiltrotors at the tail and the tilt velocity of any of the tiltrotors at any other position, and the first difference is not zero.

4. A vertical takeoff and landing aircraft according to any one of claims 1 to 2, wherein in cruising and / or vertical takeoff and / or mode change states, there is a second difference between the rotational speed of any of the tiltrotors on the tail and any of the tiltrotors at other positions, and the second difference is not zero.

5. The aforementioned vertical takeoff and landing aircraft, During flight, the pitch control ratio of the rudder better and the 2N tilt rotors is distributed according to the current airspeed or dynamic pressure. The vertical take-off and landing aircraft according to claim 1, wherein the rudder rotor and the 2N tiltrotors are individually controlled according to the pitch control ratio to perform pitch control in order to achieve pitch trim and maneuverability.

6. Controlling the 2N tilt rotors according to the pitch control ratio is: The pitch moment is differentially adjusted by the difference in tilt angle between the tilt rotor on the tail fin and any other tilt rotor, thereby performing pitch trim and control. and / or, differential adjustment of the pitching moment by the difference in rotational speed between the tiltrotor on the tail fin and any other tiltrotor, thereby performing pitch trim and control. The vertical take-off and landing aircraft according to claim 5, comprising and / or differential adjustment of the pitching moment by the difference in tilt speed between the tilt rotor and any other tilt rotor at the tail fin, thereby performing pitch trim and maneuvering.

7. The vertical take-off and landing aircraft according to claim 2, comprising four tiltrotors symmetrically mounted on both sides of the aircraft, wherein two of the tiltrotors are symmetrically mounted to the main wings with respect to the aircraft, and the remaining two tiltrotors are symmetrically mounted to the tail.

8. The vertical take-off and landing aircraft according to claim 2, comprising six tiltrotors symmetrically mounted on both sides of the airframe, of which four tiltrotors are symmetrically mounted to the main wings with respect to the airframe and the remaining two tiltrotors are symmetrically mounted to the tail.

9. The vertical takeoff and landing aircraft according to claim 7 or 8, wherein the tail fin is a V-tail fin, the two tiltrotors on the tail fin are fully tiltable tiltrotors, and the two fully tiltable tiltrotors are each attached to the wingtips on both sides of the upper part of the V-tail fin.

10. The vertical take-off and landing aircraft according to claim 7 or 8, wherein two of the tiltrotors are located at the wingtips of the main wing, and the tiltrotors located at the wingtips of the main wing are fully tiltable tiltrotors.

11. The vertical take-off and landing aircraft according to claim 1, further comprising 2M fixed rotors, where M is a natural number of 2 or more, the 2M fixed rotors being symmetrically mounted on the main wings on both sides of the aircraft and located outside the tilt rotor, the vertical take-off and landing aircraft, wherein in the vertical take-off and landing state, the horizontal projection of all the fixed rotors is center-symmetric with respect to point A, point A is located within the plane of symmetry of the aircraft, and during mode changes of the vertical take-off and landing aircraft, point G is located closer to the nose of point A or overlaps with point A, and point B is always located closer to the tail of point A.

12. The angle between the rotation axis of the 2M fixed rotors and the plane of symmetry of the machine body is between -15° and +15°. The vertical take-off and landing aircraft according to claim 11, and / or, during the process in which the 2N tiltrotors are tilted, the angle between the plane formed by the axis of rotation and the plane of symmetry of the aircraft is -15° to +15°.

13. The vertical take-off and landing aircraft according to claim 11, comprising four tiltrotors and four fixed rotors, wherein the distance from the center of gravity G to point A is L1 (L1 ≥ 0), the distance from point A to point B is L2 (L2 > 0), the spacing between the four fixed rotors in the extending direction of the aircraft body is L3, and the spacing between the four tiltrotors in the extending direction of the aircraft body is L4, such that 0.1(L3 + L4) ≥ 4L1 + 2L2 ≥ 0.01(L3 + L4).

14. The aforementioned vertical takeoff and landing aircraft, During flight, the pitch control ratios of the rudder rotor, 2N tilt rotors, and 2M stationary rotors are distributed according to the current airspeed or dynamic pressure. The vertical take-off and landing aircraft according to claim 11, wherein the rudder rotor, 2N tilt rotors, and 2M fixed rotors are individually controlled according to the pitch control ratio to perform pitch control in order to achieve pitch trim and maneuverability.

15. The vertical take-off and landing aircraft according to claim 11, comprising four tiltrotors and four fixed rotors, wherein the four fixed rotors are mounted symmetrically on both sides of the aircraft, the four tiltrotors are located inside the four fixed rotors, two of the tiltrotors are mounted symmetrically on the tail, and the two tiltrotors located on the tail are fully tiltable tiltrotors.

16. The vertical takeoff and landing aircraft according to claim 15, wherein the tail fin is a V-tail fin, the two tiltrotors located on the tail fin are fully tiltable tiltrotors, and the two fully tiltable tiltrotors are each attached to the wingtips on both sides of the upper part of the V-tail fin.

17. The vertical take-off and landing aircraft according to claim 1, wherein the tail fin is a V-tail fin, two tiltrotors are attached to the V-tail fin, the two tiltrotors are each attached to the wingtips of the V-tail fin, and in the vertical take-off and landing state, the distance between the rotation center of the tiltrotors on the tail fin and the leading edge of the wingtips of the V-tail fin is t1 in a direction parallel to the roll axis of the vertical take-off and landing aircraft, the chord length of the wingtips of the V-tail fin is t2, and the ratio of t1 to t2 is 15% to 40%.

18. The vertical takeoff and landing aircraft according to claim 1, wherein the tiltrotor located on the tail fin is a fully tilted tiltrotor, and the tiltrotor located at a position other than the tail fin is a partially tilted tiltrotor.

19. A vertical takeoff and landing aircraft according to claim 1, wherein a tiltrotor is provided at the wingtip of the main wing, and both the tiltrotor located on the tail and the tiltrotor located at the wingtip of the main wing are fully tiltable tiltrotors.

20. The vertical take-off and landing aircraft according to claim 15 or 16, wherein the fully tilting tiltrotor includes a first rotor and a nacelle, the first rotor being connected to the nacelle, the nacelle being rotatably connected to the tail or the main wing, and the nacelle tilting in synchronous with the first rotor as the first rotor tilts.

21. The vertical take-off and landing aircraft according to claim 20, wherein the first rotor includes a propeller and a rotary drive, the propeller being mounted on the output shaft of the rotary drive, and the nacelle includes a nacelle housing and a tilt mechanism located within the nacelle housing, the tilt mechanism being used to tilt the propeller by driving the rotary drive to tilt.

22. The vertical take-off and landing aircraft according to claim 21, wherein the tilt mechanism includes a rocker arm, a drive arm, a tilt drive device, and a connecting rod, wherein both the rocker arm and the drive arm are rotatably mounted on the tail fin, the base of the tilt drive device is mounted on the tail fin, the drive end of the tilt drive device rotates the drive arm, the connecting rod is rotatably connected to the drive arm and the rocker arm, and the rocker arm is fixedly connected to the rotation drive device of the first rotor.

23. A vertical takeoff and landing aircraft according to claim 22, wherein a first shaft body and a second shaft body are fixedly installed on the tail fin, the rocker arm is rotatably mounted on the first shaft body, the drive arm is rotatably mounted on the second shaft body, the base of the tilt drive device is attached to the second shaft body, and the drive unit of the tilt drive device is fixed to the drive arm.

24. A vertical takeoff and landing aircraft according to claim 23, wherein a holding structure is provided between the first shaft and the second shaft, one end of the holding structure is fixedly connected to the base, and the other end of the holding structure surrounds and holds the first shaft.

25. The vertical takeoff and landing aircraft according to claim 23, wherein the first shaft and / or the second shaft are hollow shafts.

26. The vertical take-off and landing aircraft according to claim 1, wherein the rudder beta includes a rudder plate and a rudder body drive device, the rudder plate being rotatably connected to the tail fin or the tail of the aircraft, and the rudder body drive device adjusting the direction of the vertical take-off and landing aircraft by rotating the rudder plate.

27. The vertical take-off and landing aircraft according to claim 1, wherein the roll axis of the vertical take-off and landing aircraft is set to 0°, the upward tilt of the tiltrotor is set to the positive direction and the downward tilt is set to the negative direction, and the rotation axis of the tiltrotor is tilted within the range of -20° to 110°.

28. The vertical takeoff and landing aircraft according to claim 1, wherein the tail fin is a tail fin that sweeps forward, and the tiltrotor on the tail fin that sweeps forward is a fully tilted tiltrotor.

29. The vertical take-off and landing aircraft according to claim 28, further comprising a vertical stabilizer provided below the forward-sweeping tail fin and connected to the tail of the aircraft.

30. The pitch control ratio of the rudder better and the 2N tilt rotors is distributed according to the current airspeed or dynamic pressure. A control method for a vertical take-off and landing aircraft according to claim 1, comprising a pitch control process that individually controls the rudder swivel and 2N tiltrotors according to the pitch control ratio to achieve pitch trim and maneuvering.

31. Before allocating the pitch control ratio of the rudder better and the 2N tilt rotors according to the current airspeed or dynamic pressure, Get the current tilt position of each tilt rotor, If the current tilt position does not coincide with the set cruising position, the current airspeed or dynamic pressure at the corresponding tilt rotor's current tilt position is obtained, and it is determined whether the current airspeed or dynamic pressure is equal to or greater than the set threshold at the current tilt position. If the current airspeed or dynamic pressure is above a preset threshold at the current tilt position, the tilt rotor is tilted to the next preset position. The control method according to claim 30, further comprising a rotor control process of gradually increasing the rotational speed of 2N tilt rotors.

32. The control method according to claim 31, further comprising sequentially repeating a rotor control process and a pitch control process until the tiltrotor is tilted to a cruising position, takes off, and flies horizontally.

33. The aforementioned vertical takeoff and landing aircraft further comprises 2M fixed rotors, where M is a natural number of 2 or more, and the 2M fixed rotors are symmetrically mounted on the main wings on both sides of the aircraft and are located outside the tilt rotors. The control method according to claim 31, further comprising gradually decreasing the rotation speed of 2M fixed rotors to a set rotation speed while gradually increasing the rotation speed of 2N tilt rotors.

34. Before the rotor control, The 2N tilt rotors are tilted until the axis of rotation is vertically upward or diagonally upward. Tilt the rudder vehicle downwards. The control method according to claim 33, further comprising a takeoff control process of activating 2M fixed rotors and 2N tilt rotors, and issuing a horizontal flight instruction when the vertical takeoff and landing aircraft reaches a set altitude.

35. The aforementioned vertical takeoff and landing aircraft further comprises 2M fixed rotors, where M is a natural number of 2 or more, and the 2M fixed rotors are symmetrically mounted on the main wings on both sides of the aircraft and are located outside the tilt rotors. Before allocating the pitch control ratio of the rudder better and the 2N tilt rotors according to the current airspeed or dynamic pressure, The 2N tilt rotors are tilted until their axis of rotation is horizontal and forward-facing. Tilt the rudder vehicle downwards. The control method according to claim 30, further comprising a takeoff control process of activating 2M fixed rotors and 2N tilt rotors, and issuing a horizontal flight instruction when the vertical takeoff and landing aircraft reaches a set altitude.

36. The control method according to claim 35, further comprising a rotor control process after the takeoff control process and before pitch control, which involves gradually increasing the rotational speed of 2N tilt rotors, issuing a forward flight instruction, and gradually decreasing the rotational speed of 2M fixed rotors to a set rotational speed.

37. The control method according to claim 34 or 36, wherein the rotor control process gradually reduces the rotational speed of 2M fixed rotors to a set rotational speed, then returns it to zero according to the current airspeed or dynamic pressure, and controls the rudder vater to gradually participate in the pitch control process.

38. The control method according to claim 34 or 36, further comprising a ground preparation process prior to the takeoff control process, wherein the ground preparation process starts the vertical takeoff and landing aircraft, performs a power-on self-test, and verifies the full stroke state of the servo system.

39. The control method according to any one of claims 30 to 36, further comprising controlling the rudder rotor and 2N tilt rotors individually according to the pitch control ratio to achieve pitch trim and steering by differentially adjusting the pitching moment based on the difference in tilt speed and / or tilt angle and / or rotation speed of the tilt rotors at different positions, thereby performing pitch trim and steering.

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